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FAMILIARIZATIONMANUAL:
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AIRCRAFT _?
__ .... . NOV_ 1961
............_ ,_,-_:
PAGE ....
REPORT SEDR 10_
MODEL_ 133
REVISEDST. LOUIS, MISSOURI
REVISED..
INDEX OF EFFECTIVE PAGES
The pages of this Manual currently in effect are
listed below in numerical order.
Page No. Issue
A Basic
B Basic
C Basic
i-i thru 1-29 Basic
2-i thru 2-28 Basic
3-i thru 3-54 Basic
_-i thru4-63 Basic
5-I thruS-22 _'_ ....6-1 thru6-11 _,_ BasicBasic
7-1 thru 7-5 Basic
8-i thru8-10 Basic •
9-1 thru 9-32 Basic
10-1 thrulO-35 Basic
ii-i thru ii-52 Basic
12-1 thru12-29 Basic
13-i thru 13-76 Basic
.}
DATE ,1 November 1961
REVISED
REVISED
ST. LOUIS 3, MISSOURI
PAGE ....... .1_
REPORT 8'_R i0_
MODm. 133
FOREWORD
The purpose of this document is to present a clear,operational description of the various capsule systems
and major components. Two types of usage for the man-ual are visualized. The first is an orientation-in-
doctrination type document. The second use is as a
reference document containing relatively detailed In-
formation on all systems and components.
The manual is divided by capsule systems. The first
part of each section is devoted to the description and
operation of "specification compliance capsule" system.
Capsules numbered 18 and 19 are the specification
compliance capsule. They are manned orbital capsules
and are representative of the Mercury Program. Imme-
di&tely following the specification compliance sys-
tem coverage is the Test Configuration Capsule Cov-
erage. _his area compares the other capsules to the
specification compliance capsule or to a prior cap-
sule. The capsules are compared on a, "like specifi-cation compliance capsule except as follows" bases.
The Test Configuration Capsules are compared to the
specification compliance: capsule system or to anyother preceding capsule depending on which reference
causes the least duplication. The reader will not be
required to refer to more than two prior capsules sys-
team including the specification system. Separate in-
formation is provided for each capsule test configura-
tion, when the information is the same for each cap-sule it will not be repeated. All capsules will be
covered in this manual for one revision after the par-
ticular capsule has been launched successfully. Afterthat date they will be dropped from future issues.
Capsules numbered 2, 3, 4, 5, 6, 7, 11 and 14are
covered in the i February 1961 issue of SEX_ 104
revised I August 1961. Capsules numbered 12, 15, 17
and 20 have been assigned an eighteen orbit mission
and will be covered in a latter publication.
All capsule configurations are not finalized as of
this printing, additional information will appear in
subsequent revisions to th_s document, ::reflecting
changes as they are incorporated in the capsule.
¢
SECTION INDEX
PAGE
SECTION I
INTRODUCTION ........ ...................................................................... 1-1
SECTION II
MAJOR STRUCTURAL ASSEMBLIES .......................................................... 2-1
SECTION II!
ENVIRONMENTAL CONTROL SYSTEM .................................................... 3-1
SECTION IV
STABILIZATION CONTROL SYSTEM ....................................................... 4-1
SECTION V
SEQUENCE SYSTEM, LAUNCH THROUGH RETROGRADE OR ABORT .................. 5-1
SECTION V!
ESCAPE AND JETTISON ROCKET SYSTEMS ............................................... 6-1
SECTION VII
POSIGRADE ROCKET SYSTEM ................................................................ 7-1
SECTION VIII
RETROGRADE ROCKET SYSTEM .............................................................. 8-1
SECTION IX
SEQUENCE SYSTEM, LANDING THROUGH RECOVERY ................................. 9-1
SECTION X
ELECTRICAL POWER AND INTERIOR LIGHTING SYSTEMS .............................. 10-1
SECTION XI
COMMUN ICATION SYSTEM ................................................................ 11-1
SECTION Xll
NAVIGATIONAL AIDS ....................................................................... 12-1
SECTION XIII
I NSTRUME NTATi ON SYSTEM ............................................................... 13-1
illillllil_i IPIIIIIIpIL lnPI • I
1-1
SECTION I
INTRODUCTION
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TABLE OF CONTENTS
TITLE PAGE
MISSION DESCRIPTION ........................ 1-3
CAPSULE DESCRIPTION ........................ 1- 3
BOOSTER DESCRIPTION ........................ 1-19
CAPSULE RECOVERY ............................ 1-19
CREW ................................................... 1-19
TEST CONFIGURATION CAPSULES ..... 1-23
•wv m "1 I I 5/Ll_l I I/ii I-
PAGE
REPORT
MODEL.
1-_2
SEDR 104
133
ST.LOUIS. MISSOURI
DATE 1 NOVEMBER 1961
REVISED
REVISED ,,
ESCAPE ROCKET_
PYLON JETTISON ROCKET_
ANTENNA
CAPSULE ADAPTER (REF)
FIGURE 1-1-CAPSULE PREL_AUNCH CONFiGURATiON
SPIKE
PYLON
ET ROROCKET AND
POSIGRADE ROCKETS
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DATE i November 1961
R£VISED
REVISED
McDONNE ST. LOUIS 3, MISSOURI
_- ..... 1/" lu7 l'll f ,,
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mm_omv SEDR 104
,4oom. 133
I. INTRODUCTION TO PROJECT MER_
i-i. MISSION DESCRIPTION
The possibility of man vesturing into space has shifted quite recently
from the fantasy of "science fiction" to the realm of actuality. Scientific
progress has slowly but surely loosened man's ties to the earth, and recent
technological advances have promised to release him completely. Today, space
flight is considered well within the range of man's capabilities.
Initiated by the National Aeronautics and Space Administration, a space
flight program is now underway. Through the research, design and production
facilities of McDonnell Aircraft Corporation and their many sub-contractors,
an _nerican will venture into space. The program that will put him there
is Project Mercury.
Fundamentally, the mission of Project Mercury is the projection of a
manned capsule into a semi-permanent orbit about the earth, the study of man's
capabilities in space flight, and the subsequent safe return of the capsule and
its occupant to the earth's surface. It is immediately obvious that the mis-
sion, while simply stated, is of tremeadous scope and magnitude, and requires
exceptional coordination of maupower and facilities. The date contained in
this and succeeding sections will provide detailed information on the equipment
and procedures utilized to accomplish that mission.
1-2. CAPSULE DESCRIPTION
1-3. General
See Figures I-I, 1-2, 1-3 and i-_. The Project Mercury capsule is basi-
cally a conical structure containing a pressurized area suitable for human
occupation during launch, orbit, and recovery phases of the mission. The
"base" of the cone contains provisions for attachment to the ATLAS booster,
PAGE
REPORT
MODEL
1-!4
$EDR 104
133
Mc'DONNELL ' ......... ST.LOUIS, MISSOURI
DATE 1 NOVEMBER 1961
REVISED, ,
REVISED
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1 NOVEMBER 1961 Mc,DONNE_ __ PA_E '-_ST.LOUIS, MISSOURI REPORT .SEDR 104
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PAGE I-6
REPORT SEDR104MODEL 133
Mc'DONNEL_L_ __ST.LOUIS. MISSOURI
DATE
REVISED
REVISED
1 NOVEMBER 1961
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DATE 1 NOVEMBER 1961
REVISED ....
REVISED
Mc'DONNE ST.LOUIS, MISSOURI
PAGE
REPORT
MODEL
1-7
SEDR 104
,133
/--
ROLL
Z
+
X
PITCH
/
Y
+
Y
YAW
NORMAL ORBITAL FLIGHT
PITCH
PITCHIS DEFINEDASTHE ROTATION OF THECAPSULEABOUTITS X-AXIS
THE PITCH ANGLE IS ZERO DEGREES t0 °) WHEN THE Z-AXIS LiES IN A
HORIZONTAL PLANE. USING THEASTRONAUTS RIGHT SIDE ASA REFER-
ENCE, POSITIVE PITCH ISACHIEVED BY COUNTERCLOCKWISE ROTATION
FROM THE ZERO DEGREE t0 °) PLANE. THE RATE OF THIS ROTATION IS
THE CAPSULE PITCH RATE AND IS POSITIVE IN THE DIRECTION SHOWN
AS ARE THE CONTROL MOVEMENTS WHICH CAUSE IT. THE CONTROL
HANDLE MOVES TOWARD THE ASTRONAUT AND THE POSITIVE + PITCI"
REACTION JET FIRES.
YAW
YAW IS DEFINED AS ROTATION OF THE CAPSULE ABOUT ITS Y-AXIS.
CLOCKWISE ROTATION OF THE CAPSULE, WHEN VIEWED FROM ABOVE
THE ASTRONAUT, IS CALLED RJGHT YAW AND IS DEFINED AS POSITIVE
(+).THIS MOVEMENT IS PRODUCED BY POSITIVE CONTROL MOTION. THE
CONTROL HANDLE IS ROTATED CLOCKWISE {AS VIEWED FROM ABOVE
THE ASTRONAUT) AND THE POSITIVE t +) YAW REACTION JET FIRES.
YAW ANGLE IS CONSIDERED ZERO DEGREES (0 °) WHEN THE CAPSULE IS
IN NORMAL ORBITAL POSiTiON (BLUNT END OF CAPSULE FACING LINE
OF FLIGHT). WHEN THE POSITIVE Z-AXIS OF THE CAPSULE IS DIRECTED
ALONG THE ORBITAL FLIGHT PATH tRECOVERY END OF CAPSULE FAC-
ING LINE OF FLIGHT), THE YAW ANGLE IS 180%
ROLL
ROLL IS DEFINEDAS THE ROTATION OF THE CAPSULE ABOUT ITS Z-AXIS.
CLOCKWISE ROTATION OF THE CAPSULE, AS VIEWED FROM BEHIND THE
ASTRONAUT, IS CALLED RIGHT ROLL AND IS DEFINED AS POSITIVE (+).
THIS MOVEMENT IS INITIATED BY MOVING THE CONTROL HANDLE TO
THE RIGHT THEREBY FIRING THE POSITIVE (+) ROLL R EA C TIG N JET.
WHEN THE X-AXIS OF THE CAPSULE LiES IN A HORIZONTAL PLANE, THE
ROLL ANGLE IS ZERO DEGREES (0°)..
ACCELEROMETER POLARITY WITH RESPECT TO GRAVITy
WITH THE CAPSULE IN THE LAUNCH POSITION THE Z-AXIS WILLBE PER-
PINDICULAR TOTHE EARTH'SSURFACEAND THE Z-AXISACCELEROMETER
WILL READ +1 "G".
WITH THE CAPSULE IN AN ATTITUDE SUCH THAT THE Z AND X-AXIS
ARE PARALLEL TO THE EARTH*S SURFACE AND THE ASTRONAUT IS IN A
HEAD UP POSITION, THE Y-AXIS ACCELEROMETER WILL READ +1 "G".
WITH THE Z AND Y-AXIS IN A PLANE PARALLEL TO THE EARTH'S SUR-
FACE AND WiTH THE RIGHT SIDE OF THE ASTRONAUT UP, THE X-AXIS
ACCELEROMETER WiLL READ +1 "G".
FIGURE1-5 CAPSULE POLARITY ORIENTATION WITH RESPECT TO ASTRONAUT
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PM45-233
PAGE i-8
REPORT _DR iO_I.
MODEL 133
ST. LOUIS, MISSOURI
i, , . "
I Novem1:_r i C_lDATE
REVISED
REVISED,,,
through use of special adapters. The "apex" of the cone contains the devices
for recovering the capsule at the conclusion of a mission, end equipment which
would allow the astronaut to escape in the event of an emergency during the
launch phase. Provided in the capsule proper are systems which regulate
environment, flight attitude, data recording and telemetering, and capsule
recovery.
When in place on the nose of the booster, the s_all end of the capsule
is up. The Astronaut is on his back in a sitting position. During launch and
acceleration phase, the Astronaut faces forward with respect to capsule flight
path. When the booster-capsule combination reaches a specific altitude, atti-
tude and velocity, they separate. The booster slows, and returns to the
earth's atmosphere where it is destroyed. The capsule is stablized momentarily,
then rotated 180 ° about its yaw axis. Throughout the remainder of the flight,
whether orbital or ballistic, the Astronaut faces aft with respect to capsule
flight path.
i-4. Cabin
1-5. Arrangement
The equipment within the capsule cabin interior, Figure 1-6, is arranged
so that all operating controls and emergency provisions are accessible to the
Astronaut when in the normal restrained position. Cabin equil3nent basically
consists of an Astronaut's support couch, a restraint system, instrument and
display panels, navigational aids, flight and abort control handles, food and
water supply, waste container, survival kit, cameras, and electronic equipment
required to operate ccmmnunication system.
1-6. Support Couch
The Astronaut support couch (Figure i-7) is designed to firmly support the
Astronaut's body during the capsule launch, re-entry and landing phase_.
+
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SUPPORT C
WATE RCONTAI N ER5
INSTRUMENTATIONEQUIPMENT
FLIGHT CONTROLHANDLE
FOOD EONTAII_
CONIMUN_CATIONSEQUIPMENT
RATE DAMPER
\
INSTRUMENTATION
EQUIPMENT
LARGEBULKHEAD
AMERA COEI_ELkTtONCLOCK
PANELOBSERVER CAMERA
RESTRAINTHARNESS
ABORT CONTROLHANDLE
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COOLING Duct
ELECTRONICEQUIPMENT
FACE LENS SEALSUPPLY BOTTLE
BAAX.ALT.SENSORBATTERY
ENVIRONMENTALSYSTEM EOUIPMENT
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MAIN PANEL
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ESCAPE HATCH
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DATE 1NOVEMBER1961
REVISEDREVISED
M_'DONNELL__"_ __ST.LOUIS. MISSOURI
PAGE ,,, 1-1]
REPORT SEDR 104
MODEL 133
"D" SPECIAL INSTRUMENTATION
"B" PRIMATE COUCH
HOIST ASSEMBLY
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"C* ASTRONAUT.% SUPPORT COUCH
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FIGUIRE I-7 _S.IC MlSSION EQUIPMENT
_.._.,_[.- _;fr-I_ll i I/'_ IL
P)A45- I_50
PAGE
REPORT
MODEl.
McDON E SE R REV,SED
ST. LOUIS, MISSOURI
133 ,.__ __ REVISED
The support couch also protects the Astronaut from loss of consciousness during
capsule peak acceleration and from possible injury at capsule impact. The
support couch is centrally located adjacent to the large pressure bulkhead.
The couch is constructed of a crushable, honeycomb material, bonded to a Fiber-
glass shell, and lined with a comfort liner. The support couch is molded to
the contour of a specific Astronaut's body to provide maximum body support
during capsule flight. The couch is fabricated in sections to enable couch
installation through the capsule entrance hatch.
1-7. Restraint System
The Astronaut's restrain system (See Figure 1-8) is designed to firmly re-
strain the Astronaut in the support couch during capsule maximum deceleration.
The restrain system consists of shoulder and chest straps, leg straps, aI
crotch strap, a laB belt, and toe guards. The shoulder straps may be adjusted
to restrain or release the Astronaut, by a harness reel control handle, located
forward of the upper left side of the support couch. The lap belt and crotch
strap firmly support the Astronaut's lower torso, and the chest strap and
shoulder straps restrains the Astronaut's upper torso. The leg straps and toe
guards firmly restrain the Astronaut's legs and feet. The Astronaut's hands
and arms are restrained by gripping the abort and flight control handles,
located near the ends of the support couch arm rests.
1-8. Instrument Panels (See Figures 1-9 And 1-10)
The capsule instruments and controls are located on a main instrument panel,
a left console, and a right console. The main instrument panel is located
directly in front of the Astronaut's support couch, as viewed By the Astronaut,
and is ettached to the periscope housing. The main panel is designed so that
the periscope display scope forms the lower control section of the instrument
-,° -,°- ,,, ,,, ,., C N F:D iii,i
DATE
REVISED
REVISED
1 NOVEMBER 1961 Mc'DONNELL, ST.LOUIS, MISSOURI
PAGE , 1-13
,REPORT. $EDR 104
MODEL 133
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PAGE
REPORT
MODEL
1-1i4
SEDR 104
133"
MIDONNE_ _.__ST.LOUIS o MISSOURI
DATE ....
REVISED
REVISED
1 NOVEMBER 1961
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MODEL
1.15 Mc,I)0NNELL_H._ __ DATE.SEDR 104
ST.LOUIS, MISSOURI REVISED
133 . . / REVISEDL_
1 NOVEMBER 1961
+-,.
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FIGURE 1-9 MAIN INSTRUMENT PANEL,(CAPSULE 10, 13, 16/t8 & 19i) (SHEET20Fi2)
I[ l lpl I |"
DATE
REVISED
REVISED
1 NOVEMBER 1961 Mc'DONNELL_L_ _ __ST.LOUIS. MISSOURI
PAGE 1-16
REPORT SEDR 104
MODEL 133
_ 7-"°' @
HOT
@_ .._oo_o@
RIGHT CONSOLE
[D CAPSULE 10,13,16, 18& i9
[_[_> CAPSULE 13 ONLY
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LEFT CONSOLE
FIGURE 1-10 RIGHT AND LEFT CONSOLE (CAPSULE 9 10,13,16,18 & 19)PM45-1G115
panel. Navigational instruments are located in the left and center sections
of the main panel. Environmental system indicators and controls are located
in the right upper section of the main panel. Electrical switches, indicators
and communication system controls are located in the lower right section of the.
main panel. The left hand console is located on the left side of the main panel
and is arranged to provide accessibility and visibility to the Astronaut when
in the fully restrained position. The console includes a telelight sequence
and warning panel, and indicators and controls for the capsule autc_atic stabi-
lization control system, environmental control and landing system. The right
hand console, located below the capsule entrance hatch, includes controls for
the environmental control system. A window pole, located adjacent and to the
left of observation window, enables the Astronaut to actuate controls in the
fully pressurized suit.
1-9. Navigational Aids
The capsule is equipped with navigational aids and instruments to enable
the Astronaut to compute factors relative to his flight or landing. The naviga-
tional aids consist of the periscope, satellite clock, earth path indicator,
altimeter, angular rate indicator and the map case. All navigational aids are
located directly in front of the Astronaut, on or adjacent to the main instrument
panel.
i-i0. Controls
Capsule controls are located forward of each arm rest of the support couch.
An emergency escape handle is located forward of the support couch left arm
rest. The escape handle is utilized to initiate the abort sequence. To pre-
vent inadvertent actuation of the escape system, the escape handle is provided
with a manual lock. The manual control handle, located forward of the support
couch right arm rest, is utilized to control fligh t attitude of the capsule in
PAGE 1-18
REPORT ,q_.n_ 10h
MOOa_ 133
Mc'DONNE ST. LOUIS, MISSOURI
DATE i November 1961
REVISED
REVISED
the event the automatic system failed. This handle is also normally locked.
i-Ii. Food, Water and Waste Storage
All manned capsules will be provided with food and water sufficient for the
particular mission. (See Figure 1-6). The food will provide approximately
3,000 calories. The 6 pound water supply is contained in two flat bottles) each
fitted with an extendable tube. A waste container is located on the interior
of the entrance hatch.
1-12. Survival Equipment
The survival kit (See Figure 1-11), stowed at the left side of the couch,
contains the following :
1 Life Raft
1 Desiring Kit (for8 pts.)
2 Shark Repellant Packages
3 Dye Markers
1 Tube Zinc Oxide
1 First Aid Kit
1 Bar Soap
3 Morl_hine and Anti-
Seasickness Injectors
i Signal Mirror
i Ultra SARAH Rescue Beacon
! Survival Ration
i Container of Matches
i Whistle
I cord (ioft.)
i Signal Light
i Pocket Knife
A knife, installed in a case, is located on the entrance hatch interior. (see
Figure 1-6). A flashlight is located adjacent and to the left of the observa-
tion window.
1-13. Cameras
One 16 ram. camera is mounted in the lower left corner of the main instrument
panel for viewing the Astronaut's head and shoulders. A second 16 _m. camera is
positioned above and to the left of the Astronaut's couch to record instrmnent
panel readings. These csmeras operate continuously during launc_ and descent,
_ _ /5̧ /
=A_ i November 1961
R_ISED.
REVISE:D.
Me'DONNEST. LOUIS 3, MI_URI
PAGZ, , 1-19
mm_o_ _nR i0_
13MODEL ........
and at programmed intervals during orbit.
1-14. BOOSTER DESCRIPTION
The launch vehicle, or booster, used to project the Project Mercury capsule
into orbit is the ATLAS "D" missile. Capsule adapters replace the nose comes
of the missiles. The capsule "base" is them attached to the adapter with a
segmented clamp ring. At the proper time, explosive bolts in the clamp ring
are fired, releasing the capsule. The adapter remains with the booster.
1-15. CAPSULE RECOVERY
A normal mission is intended to terminate with the capsule landing in a
predetermined area of the ocean. Under normal circumstances, ships and heli-
copters will be standing by with provisions to pick up the buoyant capsule
immediately after landing. Considering the possibility that the capsule could
land in other than the intended area; numerous devices, both electronic and
visual, are automatically energized or deployed at the time of lending to aid
in locating the capsule. Depending upon the weather, possible capsule damage
and the Astronaut's physical condition, etc., he may either stay in the capsule,
or take to the life raft which is provided as part of the survival equipment.
1-16. CREW
1-17. Re%ulrement s
The capsule crew consists of one man representing the peak of physical and
mental acuity, training and mission indoctrination. Much more will be required
of the crewman than is normally required of the modern aircraft test pilot.
The crewman must not only observe_ control and ca_nent upon the capsule system,
but must scientifically observe and comment upon his own reaction while in a
new, strange environment.
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REVISED
REVISED
1-18. Selection
.M!DONN ST. LOUIS 3, MISSOURI
PAGE 1-21
REPORT SRI_R 10_
Moore. 133
From the large number of men who volunteered for Project Mercury, a
relatively small group has been selected. Each man in the group has undergone
extensive testing and examination which has proven conclusively that he
possesses the intelligence, stamina and mental stability required for a project
of this type.
1-19. Training
An extensive training program is being conducted for the Astronauts and
other designated personnel associated with Project Mercury. The program will
provide detailed descriptions and operation of all capsule components in such
a manner that the trainee will fully understand the function of each component
and the reasons for selecting a particular design. Supplementary briefings will
be held so that current design decisions can be made known. Initial training
will be of the "group discussion" type, progressing to procedural trainers.
Training aids and equipment will be designed to train the Astronauts to achieve
the highest attainable degree of proficiency in all normal and emergency
procedures. The following objectives will be sought :
a. The Astronaut will be indoctrinated in the general purpose
and plans of the space program.
b. He must be completely familiar with all normal and emergency
procedures. Emphasis will be placed on this Point, so that
normal procedures are performed almost automatically.
c. He must be indoctrinated as far as possible in the environ-
mental and physiological aspects of the mission.
d. Since the Astronaut himself has the highest utility value
and is the most flexible component in the capsule_ he must
be able to handle the normal work load in the capsule and
..... "DE ......
PAGE 1-22
REPORT _ lOk.
MODEL_. 133
McDONNE _._,Ltguls _ M_SSOUm
1 November i(:_iDATE
REVISED
REVISED
still function as an efficient scientific observer.
The completion of an adequate training program will assure a far higher level
of reliability with respect to crew function, and of course, increase the
probability of a successful mission.
1-20. Physiological Preparation
To minimize the possibility of the Astronaut having to pass body waste
solids for the duration of a mission, strict dietary control will be maintained
for a considerable period prior to flight. This will allow a nutritional and
physical buildup in anticipation of the stringent demands which will be placed
upon the Astronaut's physical and mental facilities, and at the same time,
control the type of solid waste which will remain in the digestive and elimina-
tion systems. Finally, low residue or non-residue type food will be supplied
for the Astronaut's consumption during flight.
1-21. Aer_nedical Instrumentation
It is extremely important that certain bio-physical functions be measured
and recorded during all phases of the mission. Such measurements will assist in
monitoring the Astronaut's mental acuity and physical fittness, and will
contribute significantly to aeremedicat research. The blood pressure microphone,
respiration rate and volume 3 temperature tranducers and the F2_ pickups'are used
to register the Astronaut's physical reactions. Leads are routed from the
Astronaut's body to teminals which extend through the suit. Capsule wiring
will attach to the suit at these points. The instrumentation will be accomplished
in laboratory facilities at the launching site prior to donning the pressure suit.
The data thus derived is supplied intermittently to the capsule tape recorder
and continuously to the telemetry equipment.
,,.R=VlSED ST. LOUlS 3, Mmsoum nS_,ORT _DR iO_
.wls=o_ ,_ _ ,4oo_ 133
: '_s j
1-22. Astronaut' s Apparel
The Astronaut's apparel will consist of a completely enveloping pressurized
suit with helmet, suitable undergarments, and boots. The helmet face plate
can be opened while the capsule interior is pressurized although normal pro-
cedure will be to keep the face plate closed. Each Astronaut is specially
fitted and trained in the use of his suit. Oxygen, regulated as to temperatu/e,
pressure and humidity, is supplied to the suit for breathing and ventilation.
For Astronaut comfort, ventilating air should be supplied to the suit at all
times.
1-23. TEST CONFIGURATION CAPSULES
The data contained in paragraphs I-i through 1-22 pertains to the
Specification Compliance Capsules. Deviations from this data, as applicable
to the TEST CONFIGt_%ATION CAPSULES, is explained in paragraphs i-2_ through 1-39.
If no reference is made tO a particular system or item it is the same as the
Specification Compliance Capsule.
I-2_. TEST CONFIGbTJtTION NO. 8 CAPSULE
Capsule No. 8 is similar to the specification capsule except in the follow-
ing areas. Refer to Section II for structural differences.
1-25. Mission Description
The flight objectives of Capsule No. 8 combined with an Atlas missile, will
be a single eliptical orbit around the earth, with Re-entry initiated at a pre-
set point on the orbital trajectory. Capsule No. 8 will not be manued_ but will
be equipped with a cre_aan simulator to artifically contaminate the capsule
environmental system to a degree, similar to that which would be experienced
by an Astronaut during an extended flight. This flight will qualify the
following systems:
_A_.mlII-- .m.. °
PAGE.
REPORT
MODEL
1._., MCDONNELL__ _ D,+E....SEDR 104
ST.LOUIS, MISSOURI REVISED
133 __ REVISED,
1 NOVEMBER 1961
m
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Qz
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FIGURE 1-',12 MAIN INSTRUMENT PANEt/(CAPSULE 8)PM45-11qC
WIll! •
DATE
REVISED
REVISED
! NOVEMBER 1961
M 'DONNE ST.LOUIS. MISSOURI
PAGE 1-25
REPORT SEDR 104
MODEL 133
EMER O:
41-EMER -_NORM
RIGHT CONSOLE
LEFT CONSOLE
FIGURE.1_13 RIGHT AND LEFT CONSOLES (ICAPSULES_ 8 )P_J,4S- I _.OB
_""_'_''''_''_''1 & II I Ii i llJli 1 J Ii'il
PAGE 1-26
REPORT SE_ 10_
MODEL. 133
M 'DONNE ST. LOUIS, MISSOURI
DATE i November 1961
REVISED
REVISED
a. Capsule-booster separation.
b. Capsule and escape systems.
c. Environmental control system.
d. Automatic stabilization system.
e. Landing re-entry and recovery system.
f. Posigrade and Retrograde rockets systems.
g. Instrumentation end telemetry systems.
In addition the following objectives will be achieved:
a. Demonstrate the ability of the capsule and ground complex to
initiate an acceptable re-entry after an orbital flight.
b. Demonstrate the ability of ground range stations to perform the
necessary monitoring and control functions during flight.
c. Establish the adequacy of the location and recovery procedures
associated with re-entry.
d. Qualify the capsule primary systems for exit and re-entry flight.
e. Determine the capsule full-scale motions and afterbody heating
rates during re-entry.
f. Evaluate the capsule environment for a one orbit flight, using
the crewman simulator. This test will determine the capsule's
environmental systems capabilities, for human inhabitation during
the entire mission frcm pre-launch through the post-landing phase.
1-26. Crewman Simulator
The support couch is not installed in Capsule No. 8. Instead, an instrument
package and a crewman simulator are installed. (See Figure 1-7. ) The crewman
simulator is a box-like structure containing a carbon dioxide tank, water tank,
strip heaters, valves and controls. This device simulates the carbon dioxide
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/.
output, perspiration output, and oxygen consumption of a human being. The
simulator is calibrated prior to flight and is activated automatically when
the capsule special instrumentation package is energized.
1-27. Instrument Panels
Capsule No. 8 instrument panels are similar to the Specification Compliance
Capsules. See Figures 1-12 and 1-13 for Capsule No. 8 instrument panel configura.
tions.
1-28. Food_ Water and Waste Storage
Capsule No. 8 will not contalnfood, water and waste containers.
1-29. Cameras
The camera installation in Capsule No. 8 will consist of an Instrument
Observer Camera, an Earth and Sky Camera and a Periscope Observer's Camera.
(Refer to Instrumentation Section XIII).
1-30. TEST CONFIGURATION NO. 9 CAPSULE
Capsule No. 9 is similar to the specification capsule except in the follow-
ing areas. Refer to Section II for structural differences.
1-31. Mission Description
Capsule No. 9 Mission will be an Atlas_ Three Orbit Flight. The primary
objective of this flight will be aerQmedical instrumentation of the primate and
the primate's behavior in a space environment. This flight will further qualify
the systems listed in Paragraph 1-25 and in addition achieve the following
objectives :
a. Evaluate the effects on a primate and test the capsule environment
for satisfactory conditions during prelaunch period thr_h the
post landing phase of the mission.
b. Determine in detail the heating effects of the launch and re-entry
phases on capsule environment.
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PAGE
REPORT
MODEL_
_..T_ 104 REVISED.... ST, LOUIS, MISSOURI
133 __=_]_ REVISED.
c. Demonstrate the integrity of the capsule structure, ablation
shield and afterbody shingles for a normal re-entry from orbit.
d. Determine the capsule motions during a normal re-entry fr_ orbit.
1-32. Support Couch
The primate support couch utilized in Capsule No. 9 (See Figure i-7) is
designed to contain, sustain and support a medium-sized primate (chimpanzee)
during an unmanned capsule mission. The couch contains an instrument panel and
controls to test the primate's reactions during capsule flight. This unit is
essentially a two-section container. The aft section is the actual couch that
supports and restrains the primate; the forward section contains the instrument
panel, controls, and observation window. The primate couch, including occupant,
is installed just prior to capsule launch. A Hoist Assembly packed in a plastic
bag is attached to the primate couch support structure. This hoist assembly
is used to assist in removing the primate couch from the capsule during recovery
operations. It is located aft of, and to the right of the prlmate's couch and
is accessible through the escape hatch opening.
1-33. Instrument Panels
The instrument panels on Capsule No. 9 are the same as used on the
Specification Compliance Capsule. (See Figure 1-9,)
1-34. Food; Water and Waste Storage
Capsule No. 9 will have food and water dispensers within the couch for the
primate. The waste containers are not installed.
1-35. Cameras
Capsule No. 9 has an Instrument Panel and Primate Observer Cameras installed.
Also an Earth and Sky Observer Camera is installed which will photograph a portion
of the earth and sky through the observation window. In addition a periscope
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DATE ! November 1961
REVISED
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ST. LOUIS s, M,SSOURJ REPomr S_.oR lob
- MOme. 133
-L;. 7
camera is installed which will photograph the view displayed on the periscope.
Refer to Instrumentation Section XIII.
1-36.
unit.
TEST CONFIGURATION CAPSULE NO. i0
Capsule No. i0 is an orbital training simulator and altlt1._lechamber test
The configuration of the capsule is subject to change during various
phases of the test program; therefore, this capsule cannot be accurately
described in this manual. When Capsule No. lO is assigned to a flight mission
the final configuration will be included during a subsequent revision or
reissue of this publication.
1-37. TEST CONFIGURATION CAPSULES NO. 13 AND 16
Capsule No. 13 is the same as the Specification Compliance Capsule.
(Refer to paragraphs i-i thru 1-22).
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i
SECTION I_
iii.
MAJOR STRUCTURALASSEMBLIES
TABLE OF CONTENTS
TITLE PAGE
INTRODUCTION ....... . ........................... 2-.3
CAPSULE FOREBODY ........................... 2-4
CAPSULE AFTERBODY ........................... 2-5
ENTRANCE HATCH .............................. 2-7
OBSERVATION WINDOW .................... 2-9
RECOVERY COMPARTMENT ................ 2-11
ANTENNA FAIRING ............................. 2-14
IMPACT LANDING SYSTEM ................... 2-16
ESCAPE TOWER .................................... 2-:>0
PYLON-CAPSULE CLAMP RING ........... :>-20
ATLAS MISSILE ADAPTER ..................... 2-21
CAPSULE-ADAPTER CLAMP RING ......... 2-24
TEST CONFIGURATION ..................... :>-26
PAGE
REPORT
MODEL
2-2
SEDR 104
133
ST.LOUIS, MISSOURI
DATE
REVISE_
REVISED
1 NOVEMBER 1961
HEAT
SHIELD
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FOI_EBODY AREAVENT (42)
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DATE_ i November 1@61
REVISED
IREVISIED
,Mc'DONNE ST. LOUIS 3, MISSOURi
PAGK 2-3
MOOL 133
II. MAJOR STRUCTURAL ASSEMBLIF_S
2-1. INTRODUCTION
The Project Mercury capsule, Figure 2-1, is designed to contain an
Astronaut, primate, or crewman simulator during capsule ballistic or orbital
flight, Capsule payload will depend upon mission purpose. (See Figure 1-7,
Section I.) The capsule will also contain recording equipment, environmental
provisions, and equipment necessary to control the capsule durin_ flight.
The capsule is basically of a conical configuration consisting of a forebody
and afterbody. During orbital flight, the capsule forebody is forward with
respect to the capsule flight path. The capsule forebody is the large, dish-
shaped structure forward of the cabin area. The capsule afterbody consists of
a conical mid-section attached to a small cylindrical section. The capsule is
of a eonventional semi-monocoque construction utilizing titanium for the primary
structure. Capsule construction is designed to protect the internal cabin from
excessive heating, noise and meteorite penetration. Provisions are incorporated
in the capsule to permit cabin entry, exterior viewing, normal and emergency
exit.
Prior to capsule flight, an escape tower and antenna fairing are attached
to the capsule afterbody cylindrical section. The escape tower, designed to
aid in capsule-missile emergency separation, consists of a pylon framework
equipped with rockets. The antenna fairing is a cylindrical shaped structure
containing the capsule radio main receiving and transmitting antenna. The
escape tower is Jettisoned during the capsule launch phase or during an escape
sequence. During the capsule landing phase the antenna fairing is ejected
and serves to deploy the capsule main chute.
PAGE 2-4
REPORT S_DR 10_
MODEL lq_
MCDONNE REVISED
_ kouls ..M SSOU I_t'_ _" J.._, E _'_ REVISED
2-2. FOREBODY
The capsule forebody, Figure 2-i, mainly consists of a large, blunt, dish-
shaped structure that is supported by the large pressure bulkhead and adjoins
the afterbody conical section. The large pressure bulkhead internally
separates the forebody from the afterbody. The forebody dish-shaped structure
is an ablation heat shield that is designed to protect the capsule from extreme
thermal conditions during re-entry flight. It is also designed to prevent
capsule damage upon landing impact. The heat shield is attached to the heat
shield attach ring, which in turn is riveted to the capsule conical section
inner skin. The heat shield attach ring incorporates elongated holes, for the
installation of the heat shield to the capsule and to allow for thermal expan-
sion. The ablation shield is designed to ablate heat and is constructed of
fiberglass shingles, radially laminated to form a smooth contour. A retrograde
package assembly is attached to the heat shield by means of three straps. The
retrograde package is Jettisoned from the capsule following retrograde rocket
firing, which initiates capsule re-entry.
The forebody area, between the large pressure bulkhead and the heat shield,
is vented to atmosphere through a series of vents located around the periphery
of the capsule forebody, adjacent to the forebody and afterbody Junction. Two
toroldal shaped hydrogen peroxide tanks and six reaction control nozzles, each
covered wlth Min-K heat insulation, are located In the forebody area. The fore-
body area also houses the heat shield release pneumatic system. A landing
impact skirt is also stored in the capsule forebody area. The rubber-cloth
impact skirt, attached to the capsule heat shield attach ring and the heat
shield, is designed to absorb high energy shock loads encountered during a
capsule landing on land or water; and also to stabilize the capsule during
/-- ,
Astronaut's egress, following a capsule landing in water. During the capsule
landing phase, the heat shield is released, and extends the full _length of the
impact skirt. Upon heat shield contact with land, air within the impact skirt
is forced out through a series of holes located in the impact skirt wall which
in turn provides a cushion-like effect. To prevent damage to the large pressure
bulkhead in the event the heat shield strikes the capsule during landing, the
large pressure bulkhead incorporates a reinforced laminated fiberglass shield
assembly. The fiberglass shield assembly is attached to the torus tank support
brackets. Sandwiched between the fiberglass shield and the large pressure bulk-
head are sections of honeycomb. Fabricated of stainless steel and located about
the periphery of the impact landing skirt, are 24 straps which prevent tearing
of the impact landing skirt during high horizontal velocity water landings.
Located inside the impact landing bag and alternately located in relation to
the steel straps, are 2A stainless steel cables. The cables retain the heat
shield to the capsule in the event strap failure should occur. The afterbody
conical section exterior shingle arrangement extends beyond the large pressure
bulkhead, to the forebody heat shield, and encloses the equipment located
between the large pressure bulkhead and the heat shield. Located adjacent to
the capsule forebody and afterbody Juncture, and bolted to the heat shield attach
ring, is a fiberglass attach ring. During capsule-adapter installation, the
fiberglass attach ring and the adapter attach flange are clamped together with
a segmented clamp ring. Receptacles for the capsule retro-package, adapter, and
the clamp ring pneumatic and electrical connectors are located under the forebody
shingles adjacent to fiberglass attach ring. Six spring loaded access doors,
for the receptacles are incorporated in the shingles.
2-3. A_DY
The capsule af_erbody conical mid-section mainly consists of a pressurised
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REPORT 8EDR lO_ .... "ST. LOUIS, MISSOURI " REVISED
MODEL_ 133 _z_'_- "_ REVISED
cabin that is Supported between a small pressure bulkhead and the large pressure
bulkhead. The cabin interior wall is lined with channeled frames to provide
additional structural strength and equipment attach points. The mid-section is
constructed of a conically formed inner and outer titanium shell, seam welded
together. The outer skin is beaded to form small sealed pressure panels capable
of withstanding high pressures and structural loads. The outer conical skin is
reinforced with longitudinal hat stringers. A blanket of thermoflex insulation
is bonded, in between the hat stringers, to the outer (beaded) conical skin.
Min-K insulation is also installed over the hat sections and covered with a
shingle arrangement. The shingle arrangement is similar to the shingle installa-
tion used on the recovery system compartment. The forward end of the conical
section is attached to the forebody heat shield. The combination of the conical
section beaded outer skin, the hat section reinforcements, thermoflex insulation,
and external shingle installation provide the capsule with adequate heat, noise
and meteorite protection.
Located in the bottom of the conical section, as viewed during capsule
normal flight attitude, is a retractable door that encloses the periscope lower
lens flange and the capsule ground checkout umbilical receptacle. The door,
mechanically linked to the periscope housing, automatically opens and closes
with periscope extension and retraction. Two auxiliary hoist fittings, attached
to left and right side of the capsule, provide ground handling attach points. The
hoist fittings are removed prior to capsule launch. An explosive door, namely
the snorkel explosive door, is provided in the capsule shingles, between the
small pressure bulkhead and the capsule conical-cylindrical sections Juncture.
This door is exploded from the capsule during capsule landing. The exploding
of the door from the capsule enables cool air to be drawn into the capsule
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through the snorkel valve when the cabin air inlet valve opens.
2-_. Entrance Hatch
An entrance hatch, Figure 2-2, is located on the right side of the after-
body conical section as viewed from the capsule crew member station. Entrance
hatch construction, similar to the conical section construction, consists of
an inner and outer (beaded) skin seam welded together and reinforced with hat
stringers. A waste container and th_ Astronaut's knife are attached to the
hatch. An explosive charge, moulded in the hatch sill, is provided to quickly
release the hatch and enable the Astronaut to egress rapidly from the capsule.
An explosive charge initiator, located in the upper aft corner of the hatch, is
linked to an internal release control initiator. Prior to capsule launch, the
hatch is bolted and _sealed into position with bolts, and two corrugated shingles
are installed over the hatch. The bolts are inserted through the entrance hatch
sill, which incorporates the explosive charge, and threaded into the capsule
sill. A magnesium gasket, with inlaid rubber, forms the hatch seal when the
hatch is bolted into position° Two hatch shingles are attached to the hatch
stringers, but in no manner are they attached to capsule shingles. (This
enables the hatch to separate cleanly, upon ignition of hatch explosive charge. )
Following capsule impact, the Astronaut removes the initiator cap from the
initiator, and the safety pin from the initiator plunger. By depressing the
initiator plunger, the inltiator's two spring-loaded firing pins strike the
explosive charge percussion caps and detonate the explosive charge. This
action explodes the hatch from the capsule. An exterior hatch release control
is also provided to enable ground personnel to explode the hatch in the event
the Astronaut is unable to do so. Hatch retention springs, secured by pip plus,
are incorporated ca the inner side of the entrance hatch to prevent injury to
ground personnel in the event the initiator plunger is accidemtly depressed.
2-8PAGE
REPORT SEDR 104
MODEL 13 :_
Mc'DONNE ST cLLOU IS, MISSOURI
DATE
REVISED
REVISED
1 NOVEMBER 1961
ENTRANCE MATCH%
EXPLOS|VE. CHARGE.
FIGUREZ-Z CA_P5ULE
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SEAL
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R_V,SE_ _ _oomu 133
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Two pressure valves, located in the hatch, enables pressurization and purging
of the capsule, during capsule ground checkout operations.
2-5. Observation Window
An observation window, Figure 2-3, located on the afterbody conical section,
provides the Astronaut with external viewing. The window, located above the
main instrument panel, consists of an inner and outer assembly. The inner
window assembly is made of three glass panes and a fourth image suppression
filter pane. The three glass panes are sealed in a titanium frame that is
attached to the cabin wall. Each glass pane is independently sealed to provide
a pressure seal between the panes. The image suppression filter pane eliminates
secondary (reflected) images. The filter pane is not pressure sealed. The
outer window assembly consists of a glass pane sealed in a titanium frame, that
is attached to the capsule outer skin. The outer window assembly is sealed
separately, from the inner window assembly, to provide a complete seal. The
outer window conforms to the curvature of the capsule conical section. The
observation window is equipped with filters and door llds, enabling the
Astronaut to regulate external light entering the cabin. The observation
window includes a mirror assembly which increases the Astronaut's angle of
observation.
2-6. Small Pressure Bulkhead
The small pressure bulkhead internally separates the cabin pressurized
area from the recovery system compartment and structurally supports the aft
conical section. A sealed escape hatch, Figure 2-4, internally actuated, is
provided in the small pressure bulkhead to enable the Astronaut's exit follow-
ing capsule landing. The dish-shaped escape hatch is constructed of a beaded
aluminum skin spotwelded to an inner skin, that is reinforced with structural
"Z" shaped members. The hatch outer flanged edge fits into the small pressure
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P"_E _.1o Mc,DONNELL_,S__ __ D,,TEREPORT SEDR 104 ST.LOUIS, MISSOURI REVISED
MODEL 133 .... ,- REVISED ,
1 NOVEMBER 1961
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FI&URE 2-3 CAPSULE OBSERVATION WINBOW ph_4.5-14.5.,
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bulkhead sill and is held in place with a retaining ring. Expanding the re-
tainer r_g by rals_ the hatch hsoldle, wedges the retainer ring between the
bulkhead sill and the hatch flanged edge and forces the hatch f_e aft to
provide a sealing action. The titanium small pressure bulkhead is seam welded
to the conical section inner skin and bolted to the conical hat stringer flanges.
2-7. Large Pressure Bulkhead
The large pressure bu3_khead supports the forward end of the conical section
and internally separates the pressurized cabin from the forebody heat shield.
The large pressure bulkhead is constructed of a combined inner and outer
titanium skin. The outer skin is beaded and seam welded to the inner skin.
The bulkhead is reinforced with horizontal channels installed on the outer skin.
The bulkhead inner skin is provided with two vertical channels, centrally
located and spaced, that furnish structural attach point for the Astronaut
support couch. Honeycomb shelves are provided on the bulkhead inner skin, out-
board of the two vertical channels, for equipment installation. The bulkhead
outer flange ring is bolted to the conical section inner skin and the bulkhead
is also bolted to the conical section inner attach ring. Vents are provided
in the large pressure bulkhead to enable overboard venting of the capsule
battery vapors and environmental control system exhaust steam.
2-8. RECOVERY COMPARTM_qT
The capsule afterbody, Figure 2-1, basically consists of the short cylin-
drical section and the truncated cone shaped structure. The cylindrical section
is referred to as the capsule recovery system compartment and contains the laud-
ing parachutes, recovery aids, and the reaction control nozzles. The truncated
cone shaped structure, referred to as the capsule afterbody conical section,
encloses the pressurized cabin. The recovery compartment is connected to the
pressurized cabin by a small pressure bulkhead. The recovery system compartment
PAGE
REPORT
MODEL
2-12
SEDR 104
133.... gy:L_llS :_Mrs so UR+._
DATE
REVISED
REVISED
1 NOVEMBER 1961
//
//
RETAIN INg . RETAINING r--t-lATCH _ [,\I_[N_'-_
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"_ RELEA%E HANDLE
HATCH P_ELF.ASED HATCH CLO-_ED A_s_YPRING"--_ _,,_/
5ECTIOM : (_" B HATCH P,ELEASEI2
FIGURE 2-4 CA_ULE ESC_,PE HATCH
I
DATE
REVISED
REVISED
1 NOVEMBER 1961 Mc'DONNEL_L_ __._ST.LOUIS, MISSOURI
11"
PAGE 2-13
REPORT SEDR 104
MODEL 133
CHUTE
POST
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PAGE 2-14
REPORT SEDE lO_
MODEL_ 133
MCDONNEL LST. LOUIS, MISSOURI
"_--CC,i'_[':D _ Y:TI_
DATE i November 1961
REVISED
REVISED
is a cylindrical formed titanium skin structure, reinforced with longitudinal
hat stringers, and covered with a corrugated beryllium shingle arrangement. A
layer of thermoflex insulation is installed between the hat stringers and the
external shingles to prevent excessive heating within the compartment. The
shingles are individual panels bolted to the hat sections with allowances for
thermal expansion. A set of reaction control exhaust nozzles are internally
located every 90 ° , between the compartment inner skin and the external shingle
installation. The recovery system compartment interior is structurally divided
into a left and right section. The compartment left section houses the recovery
aids, electrical wiring and plumbing routed through the compartment. The right
section of the compartment houses a fiberglass container, structurally divided
into two sections that contain the main and reserve parachutes. The container
can be removed by the Astronaut from the cabin following capsule landing, to
permit egress through the recovery compartment.
2-9. ANTENNA FAIRING
The capsule antenna fairing, Figure 2-5, is a cylindrical shaped structure
that houses the pitch and roll horizon scanners, and the main receiving and trams-
mitting antenna. The antenna fairing basic structure is of titanium construction
and is covered with '%qene-_l" shingles. An 8 inch window assembly is located
around the outer base of the fairing and acts as a dielectric between the top
of the fairing and capsule. The window assembly consists of a silicone base,
fiberglass insulation, vycor glass and teflon strips. In line with the three
teflon strips and attached to the antenna fairing shingles, are three laminated
fiberglass guides. The fiberglass guides and teflon strips prevent damage to
the antenna fairing when the escape tower is jettisoned. An aluminum bi-conical
horn is internally located at the base of the antenna fairing. An electric
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REPORT SEDR 104
MODEL 133ST. LOUIS, MISSOURI
DATE,,.,t November 1961
REVISED
REVISED
insulator and lockfoam, located above the bi-conical horn, aid lu antenna fair-
ing insulation. A pitch horizon scanner is located at the top of the antenna
fairing. A roll horizon scanner is located lu the side of the fairing, In line
with the pitch horizon scanner. The fairing is attached to the capsule by a
mortar gun located in the capsule recovery compartment. A steel post located in
the center of the fairing is used as a guide when the fairing is Jettisoned.
Three index pins and six support clips, in the antenna fairing lower matlng
flange, align with three holes and six brackets in recovery compartment mating
flange. The antenna fairing also houses the drogue chute. Three cables retain
the drogue chute risers to the fairing when the chute is deployed.
2-10. De-Stabilizer Fla_
A spring loaded de-stabillzer flap is attached to the top of antenna fair-
ing, opposite the pitch horizon scanner. The de-stabilizer flap ensures capsule
correct re-entry attitude during capsule abort and re-entry phases. During cap-
sule launching phase, and up to the capsule-tower separation, the spring loaded
de-stabillzer flap is held flat against the antanna fairing by means of a nylon
cord routed through two de-stablizing flap reeflmg cutters contained within the
antenna fairing housing. Jettisoning of the escape tower actuates the cutters
which sever the cord releasing the spring loaded flap to the outboard position.
When the capsule descends to lO,O00 feet altitude, the antenna fairing is auto-
matically Jettisoned from the capsule by the firing of the fairing mortar gum.
2-11. IMPACT LANDING SYSTEM
The capsule impact landing system, Figure 2-6, is designed to absorb high
energy shock loads encountered during landing; and also to stabilize the capsule
following a landing in water. The impact system basically consists of a heat
shield release mechanism, heat shield retaining straps (24), heat shield reten-
MAC 2:_ICL (27 APR Sg) _.JVJ._I_I. AU4.aJ._I Jl. I_'l. JLI
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REVISED ST.LOUIS , MISSOURI REPORT
REVISED " _ MODEL
2-17
SEDRI04
133
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MAIN gQUII3 _u5
FIGURE Z -7 IMPACT LANDI'NG SYSTEM SCHEMATIC MH45 -.t._l C
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PAGE .2-18
REPORT SEDR 104MCDONNE
REVISED .ST. LOUIS, MISSOURI
MODEL_ 133 _,._ . REVISED
tion cables (24), a rubberized cloth impact skirt and a fiberglass shield
assembly. The impact skirt is stored in the capsule forebody area. During the
capsule normal landing phase, the capsule 10,O00 feet barostats energizes the
Main Deploy Relay to eject the antenna fairing, which in turn deploys the main
parachute. (See Figure 2-7. ) Ejection of the antenna fairing closes the
antenna fairing separation sensing switch, which in turn directs 2_ V D-C
electrical power to the Main Inertia Switch Relay #l (time delay relay). Twelve
seconds later, the Main Inertia Relay #1 energizes to direct electrical power to
the heat shield release system limit switches and also to energize the Landing
Bag Relay. Energizing the Landing Bag Relay directs 24 V d-c electrical power
to ignite the 2 heat shield release explosive squib valves.
Ignition of the squib valves allows 3,000 psig nitrogen pressure to flow
to the 2 heat shield release mechanism actuators. This action moves the heat
shield from the capsule. Simultaneously with the actuation of the release
mechanism, the mechanism 2 limit switches close to energize the Landing Bag
Extension Signal Relay. Energizing the signal relay, directs electrical power
to illuminate the Landing Bag Telellght (green), indicating a safe condition.
When the actuator piston fully travels to the open limit, the actuator is locked
by a spring loaded lock pin.
In the event the heat shield mechanism failed to actuate, and release the
heat shield, the two limit switches will remain open and the Landing Bag Warn-s
ing Light Relay will energize within 2 seconds. This in turn directs power to
illuminate the Landing Bag Telelight (red), indicating an unsafe condition to
the Astronaut. The Astronaut should place the LANDING RAG SWITCH to the "MAN"
position. Placing the landing switch to the manual position energizes the
Emergency Landing Bag Relay, which In turn provides power to ignite the 2 heat
shield release explosive squib valves. Ignition of squib valves will actuate
DATE 1 NOVEMBER 1961
REVISED
REVISED
ST.LOUIS, MISSOURI
PAGE 2-1'9
REPORT SEDR 104
MODEL 133
DEF..STABILIZING
(CAPSULE 8 ONLY) FIGUI::2E 2-t_
PAGE • 2-20
REPORT SEDR lob,.
MODEL_ 133
Mc'DONNEL LST. LOUIS, MISSOURI
DATE i November 1961
REVISED
REVISED
the mechanism to release the heat shield, and in turn the limit switches will
close to illuminate the telelight (green).
2-12. ESCAPE TOWER
The escape tower (Figure 2-8), designed to aid in Capsule-missile emergency
separation, consists of a pylon framework equipped "with rockets. The pylon is
a triangular shaped structure that is designed to support an escape rocket and
a Jettison rocket. The pylon is constructed of 4130 tubular steel and is
approximately lO feet in length. The base of the pylon structural tubing is
bolted to a steel flanged attach ring. A four foot escape rocket casing is
bolted to the top (apex) of the pylon. Bolted to the bottom of the escape
rocket casing is a Jettison rocket. Electrical wiring is routed through the
structural tubing, from the rockets to connectors, located on the pylon attach
ring. Pylon tubular structure is covered with heat protective material. Prior
to capsule launch the pylon is installed onto the capsule, by clamping the
pylon attach ring to the capsule recovery system compartment with a chevron
shaped, segmented clamp ring. Explosive bolts connect the clamp ring segments
in tension. The bolts are fired to separate the clamp ring when the pylon is
Jettisoned from the capsule. During capsule normal launch the escape rockets
are fired to separate the pylon from the capsule. In the event the capsule
escape system is activated, during launch phase, the escape tower is fired to
propel the capsule away from the missile and then the Jettison rocket is fired
to separate the pylon from capsule.
2-13. PYLON-CAPSULE CLAMP RING
The clamp ring consists of three chevron shaped segments that clamp the
pylon attach ring to the capsule recovery system compartment flange. Three
explosive bolts, with dual ignition provisions, connect the ring segments in
, -77,,¸
MAC 231G (Rev 14 Oct. 55)
DATE i November i_I
REVISED
REVISED
M DONNE ST. LOUIS 3, MISSOURI
PAGz .... 2- 21
REPORT _EDa 104
MOD_ 133
l
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tension. The clamp ring is basically the same in design as the capsule-adapter
clamp ring (Figure 2-11), but considerably smaller in size. The clamp ring
retains the pylon to the capsule until the clamp ring explosive bolts are fired
which in turn separates the clamp ring. The exterior of the clamp ring is
covered with a heat shield to protect the clamp ring and explosive bolts from
excessive heating during capsule launch. A layer of thermoflex insulation is
bonded to the interior of the heat shield. The heat shield is attached to the
clamp ring with screws. An aerodynamic stability wedge attached to pylon heat
shield, aids in stabilizing the capsule during the launch phase. Six cable
straps, bolted to the pylon structure and clamp ring stability wedge, aid in
capsule-pylon separation by retaining the clamp ring segments of the pylon when
the explosive bolts are fired.
2-14. ATLAS MISSILE ADAPTER
The Atlas missile adapter I Figure 2-9, is a slightly tapered, cylindrical
shaped structure that is designed to mate the capsule with the Atlas missile.
Upon adapter and capsule •installation to the missile 3 the adapter is bolted to
the missile and the capsule is attached to the adapter. The adapter is of
semi-monocoque construction and is approximately 4 feet in height. The adapter
basically consists of an outer corrugated titanium skin assembly, riveted and
Seam welded to an inner titanium skin assembly and internally reinforced with
two titanium support rings, riveted between the ends of the adapter. A steel
flanged ring is riveted to the bottom, inner surface of the adapter. The
flanged ring is provided with holes to enable the attachment of the adapter
to the missile with bolts. Alignment marks are provided on the ring for proper
adapter-missile alignment. Riveted to the top, inner surface of the adapter
is an aluminum flanged ring. The adapter aluminum ring mates with the capsule
forebody fiberglass attach ring, during capsule to adapter installation.
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An alignment mark on the adapter ring BY axis enables proper alignment of
capsule to adapter. The top of the alum lnum ring is slotted at 120 ° intervals,
to provide adequate clearance for the capsule retrograde rocket assembly attach
straps, when the capsule is attached to the adapter. A metal striker bracket
is riveted internally, every 120 °, to the adapter skin. When the capsule is
attached to the adapter, these striker brackets depress (open) the capsule-
adapter separation sensing switches, located on the bottom of the retrograde
rocket assembly attach straps. The capsule is attached to the adapter by
installing a chevron shaped, segmented clamp ring over the mated flanges of the
capsule forebody fiberglass attach ring and the adapter upper ring.
A retainer assembly, attached to the adapter interior skin, is provided to
prevent the retro-package and the explosive bolt fragments from striking the
Atlas missile adapter LOX tank. The retainer assembly is a cup shaped structure,
that fits over the retro-package dome, and is supported by three metal straps
that are attached to the adapter with cable assemblies. A vent port, located
in the adapter skin, receives the missile boil-off valve tube and enables the
relieving of liquid oxygen from the missile. Opposite the liquid oxygen boil-
off port, is an adapter door installation. The door installation provides
access to the booster and Capsule heat shield area while on the pad. A fiber-
glass shield attached above the vent port opening, streamlines the adapter and
shields the boil-off tube. Two stretch fittings, located 180 ° apart at the
upper section of the adapter, provide a means of supporting (stretching) the
missile while in the vertical position following adapter installation. Six
cable assemblies, attached to fittings spaced around the adapter outer corrugated
skin, are attached to the clamp ring that attaches the capsule to the adapter.
The cables retain the clamp ring to the adapter following capsule-adapter
separation.
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REPORT _DR lO b" ST. LOUIS, MISSOURI REVISED.
MODEL_ 1_3 J _.T_,A-_ ....... _ REVISED
2-15. CAPSULE-ADAPTER CLAMP RING
The capsule-adapter clamp ring, Figure 2-10, is provided to attach the
capsule to the adapter. The clamp ring secures the capsule to the adapter
throughout the capsule launching phase until the clamp ring is separated by means
of explosive bolts, which in turn allows the capsule to separate from the adapter.
The clamp ring consists of three chevron shaped segments, that when installed,
mate with the capsule forebody fiberglass attach ring and the adapter upper
support ring. Three explosive bolts, with dual ignition provisions, connect
the 3 clamp ring segments in tension. A metal striker bracket is bolted 3 every
120 °, to the inside of the clamp ring. When the clamp ring is installed, the
striker brackets depress the capsule ring separation sensing switches, located
in the outer periphery of the capsule forebody.
The exterior of the clamp ring is covered with a heat shield that protects
the explosive bolts from excessive heating. The heat shield consists of three
fairing assemblies which are located directly over the explosive bolts and three
segmented fairing assemblies which cover the remainder of the capsule adapter
clamp ring. The fairing assemblies which locate directly over the explosive
bolts are a three piece installation. The top piece is fabricated from aluminum
and the two bottom pieces are made from titanium. These fairing assemblies are
fastened to the clamp ring support fittings. The interior of the fairing
assemblies is insulated with thermoflex. The three segmented fairing assemblies
are of a titanium construction whose interior is insulated with thermoflex. The
segmented fairing assemblies are bolted to the adapter clamp ring. Six cable
straps are bolted to the capsule adapter cable fitting. These straps aid in
capsule-adapter separation, by retaining the clamp ring to the adapter when the
explosive bolts are fired. An electrical cable, clamped around the interior of
the adapter, is connected to each of the clamp ring explosive bolts, to two
/- •
DATE 1 NOVEMBER 1961
REVISED
REVISED
SECTION'(: -C"
FIGUP, E 2 -I0 CAP5ULE ADAPTER CLAMP P,ING PM_5- 26
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REPORT BE.DR 101¢ ST. LOUIS, MISSOURI REVISED
MODEL 133 ......... T_nT.._ffm,.. REVISED-- _.4_,..I J. 11 4L" JL.L_ JLr_W-'-.I=--._"'L"I_" ....
receptacles in the capsule forebody area and to two receptacles on the missile.
A pneumatic line is also connected to one end of the explosive bolt and to a
quick disconnect in the capsule forebody.
2-16. TEST CONFIGURATION CAPSULES
2-17. TEST CONFIGURATION CAPSULE NO. 8
Capsule 8 differs from the specification capsule in the following manner.
2-18. Forebod[
Capsule 8 does not contain an impact landing skirt. Instead, the heat
shield is bolted directly to the heat shield attach ring (See Figure 2-1).
Capsule 8 capsule adapter clamp ring heat shield has a large frontal area in
comparison to the specification capsule adapter clamp ring heat shield. Capsule
8 does not incorporate a large pressure bulkhead fiberglass shield protector.
Capsule 8 Atlas missile adapter is the same as the specification capsule (See
Figure 2-10).
2-19. Afterbodz
Capsule 8 does not have an explosive entrance hatch (Figure 2-12). The
hatch can only be removed externally. Hatch removal is accomplished by removing
the attach bolts that secure the hatch to the capsule. Hatch sealing is similar
to the specification capsule. The entrance hatch stringers are interlocked
(bolted) with the capsule stringers when the hatch is bolted in place. Capsules
8 and 12 do not contain entrance hatch retention springs.
2-20. Windows
Capsule 8 contains two cabin windows, but does not contain an observation
window (Figure 2-1). One window is located on the upper left side as viewed
from the capsule crew member station, to permit Astronaut's exterior viewing.
Located in the lower right side of the capsule is a window that enables the
. Y
........_ r__,T"__"_"_Wr',_",_--"-"_ -_'-'-''".... A .,.,.a.,_ ._a a,'t A A2"IkJ_
DATE
REVISED
REVISED
1 NOVEMBER 1961
Mc'DONNE ST.LOUIS, MISSOURI
PAGE 2-27
REPORT SEDR 104
MODEL 133
/
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HATCH SHINGLE
[ INSULATION
INSULATION
HATCH SEAL
PI_2-:ll. C_PSG_E E_I'RAlqC_ HEI"CH
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PM,_,r_2A
REPORT _EDR 10[I - REVISEDST. LOUIS, MISSOURI
MODEL- 133 _ REVISED,_r , , ii I i i II .... _
photographing of the earth and sky during capsule flight. A camera is internally
located adjacent to this window. Each cabin window consists of an inner and outer
window assembly. The cabin inner (main) window assembly consists of four glass
panes sealed in a titanium frame attached to the cabin wall. The glass panes are
spaced and independently sealed to provide a pressure seal between the panes.
The outer cabin window assembly consists of a glass pane sealed in a titanium
frame that is attached to the capsule outer skin. The outer pane conforms to the
curvature of the capsule conical section. Capsule 8 does not incorporate a
snorkel explosive door.
2-21. Antenna Fairing
The antenna fairing destabilizing flap on Capsule 8 does not contain a
horizon scanner cover. (See Figure 2-1).
2-22. TEST CONFIGURATIONS CAPSULES 9, i0, 13 and 16
Capsules 9, I0, 13 and 16 major structural assemblies are basically the
same as that of the specification capsule.
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ENVIRONMENTAL CONTROLSYSTEM
TABLE OF CONTENTS
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TITLE PAGE
DESCRIPTION 3-3eoeoeqsoeeeeoeeooeooeeoeeeeeeooooeeeeooeoeoe
CABIN ENVIRONMENTAL CONTROL ......... 3-S
SUIT ENVIRONMENTAL CONTROL ............ 3-6
SUIT EMERGENCY CONTROL, .................... 3-.7
OXYGEN SUPPLY ................ , .................... 3--8
COOLING CIRCUIT .................................... 3-8
OPERATION ............................................. 3-9
SYSTEM UNITS ......................................... 3-29
TEST CONFIGURATION..., ....................... 3-53
II_'ILI 1 I lfmlim
PAGE
REPORTMODEL
3-2
SEDR104133
C
M,DONNELL ST.LOUIS, MISSOURI
DATE 1 NOVEMBER 1961
REVISED ,
REVISED
4p 0
FIGURE _-t EMVIP, ONMENTAL CONTROL SYSTEM
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REVmEO, ST. -outs 3, MiSSOURi REPORT _EDR 104
.EV,SEO. . MO0_ 133
III. ENVIRONMENTAL CONTROL SYSTEM
3-i. DESCRIPTION
The environmental control system Figure 3-!, provides the Capsule Cabin and
the Astronaut with a 100 per cent oxyge n environment to furnish the breathing,
ventilation and pressurization gas required during the capsule orbital flight and
ventilation for a 12-hour post-landing phase. The environmental control system
also functions in the following manner: Removes odors, CO2 and moisture from the
Astronaut's suit circuit; provides an "emergency rate" usage of oxygen in the
event suit circuit pressure drops below 4.0 + .1 psia; maintains cabin and suit- "3
temperature at approximately 80°F; provides an emergency fire extinguishing
capability and operates in "weightless" or "high g" conditions. System function-
ing will be automatically controlled during all phases of flight. In the event
the system automatic control malfunctions, manual controls are provided to insure
system operation.
The environmental control system is designed to be operated in either the
suit mode, cabin mode or emergency mode. The system primary mode, suit environ-
mental control, is normally utilized and enables the Astronaut to function in
the closed suit circuit during cabin pressurized and emergency (depressurized)
conditions. In the event one control mode malfunctions, the remaining control
mode will continue to operate. The emergency mode, suit environmental control,
insures Astronaut survival in the event both the suit and cabin environmental
control modes malfunction. (see Figure 3-2. )
The environmental control system provides a primary and secondary oxygen
supply for both the cabin and suit circuits. Primary and secondary oxygen systems
are basically the same, except the secondary oxygen regulated pressure is lower
than the primary oxygen regulated pressure. A manually controlled cooling circuit,
_vz_li il,d'l_lNI I 1/_ I-"
PAGE 3-4
REPORT SEDR 104
MODEL 133
Mc'DONNELL ' ....... -ST.LOU IS. MJ,SSQUR!,_.8
CASIN AII;:I I
INLET J
VALVE J
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DATE II1 NOVEMBER 1961
REVISED II
REVISED
PRIMARY
OXYGEN OXYGEN
BOTTLE BOTTLE
FLAPPER J
tVENTIL?TIONI
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SUIT
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I:'IGU_E 3'_. EI,qV_i2.0NI_ENTAL CONT_0L 5"YSTE_ BLOCV.. OIA, GIZ/X, lv_
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DATE i November 1961
REVISED
REVISED
Mc'DONNE ST. LOUIS 3, MISSOURI
PAGE 3"5
REPORT ,q_l_ 10_ ......
MOCk. .......13_
/
/
\\
for suit and cabin systems, is provided to control suit and cabin temperatures
during capsule flight. The capsule environmental control system components are
located below the Astronaut's support couch adjacent to the large pressure bulk-
head and also on the interior of the small pressure bulkhead adjacent to the
capsule escape hatch. System manual controls are located on the left and right
consoles; system instruments and warning lights are located on the main instrument
panel.
3-2. CABIN ENVIRONMENTAL CONTROL
During capsule normal orbital flight, the environmental control system is
normally operated in both the cabin and suit environmental control mode. Opera-
tion in the cabin control mode permits the Astronaut to open his helmet face-
plate for short periods of time. The capsule primary and secondary oxygen supply
furnished the cabin with pressurization, breathing, and ventilation gas. The
cabin is equipped with automatic and manual controls for cabin ventilation, de-
compression, pressurization, temperature control, landing and post landing
ventilation.
The cabin is cleared of contaminants and a lO0 per cent oxygen environment
is made available by purging the cabin prior to launch. The purging operation
is accomplished by utilizing the Capsule Leakage Tester. During orbital flight,
cabin pressure is automatically controlled by a cabin pressure control valve. The
cabin pressure relief valve prevents excessive pressure buildup within the cabin
and provides a manual means of decompressing the cabin in the event of a fire or
buildup of toxic contaminants. A water coolant supply tank, common to both cabin
and suit circuits heat exchangers, provides cabin cooling. In addition to its
cooling capability, the coolant tank is a source of drinking water for the
Astronaut. Cabin temperature is controlled by a manually controlled selector
valve, which regulates the amount of water entering the cabin heat exchanger, and
_v. mm m_al A |f_k I--
3-6 M ,DONNELL oA, REPORT R_,D_ ].Oh ST. LOUIS, MISSOURI REVISED
MODEL_ 133 _ REVISED
in turn provides cabin cooling by means of water evaporation. The cabin fan,
located on the inlet side of the heat exchanger, forces cabin air through the
exchanger to provide cabin cooling and ventilation. Cabin air inlet and outflow
valves, located on the small pressure bulkhead, provide ventilation during the
"capsule landing and post landing phase.
3-3. SUIT ENVIRONMENTAL CONTROL
During capsule normal orbital flight, the capsule common o_jgen supply
furnishes oxygen simultaneously to the suit and cabin environmental control
circuits. If a cabin circuit malfunction, such as cabin decompression, should
occur at a time when the Astronaut has his faceplate removed, the Astronaut
should immediately close his faceplate. Closing the faceplate initiates the
suit environmental control mode and confines the Astronaut to the closed suit
control circuit.
While operating in the suit environmental control mode, the suit pressure
regulator controls the suit pressure to approximately 5 psia and replenishes
oxygen consumed by the Astronaut, duringnormal suit control circuit operation.
The suit pressure regulator is supplemented by a constant oxygen bleed system
as a supplier of oxygen to the suit circuit. Suit circuit pressure is also
utilized as a secondary means of pressurizing the water coolant tank, in the
event the coolant quantity system fails. The suit environmental control circuit
incorporates compressors, filters, absorbers and a temperature control to insure
Astronaut's maximum comfort. Suit circuit temperature is controlled by means
of water evaporation. A water separator utilizes the common oxygen supply
pressure, to remove moisture from the suit circuit oxygen supply. A compressor,
located on the upstream side of the suit circuit heat exchanger, forces the suit
circuit oxygen supply throughout the circuit, providing suit circuit ventilation.
.... _. =._=, JL JL.-_-1LII, _
DATE ,1 November l_l "J_--_M_DONN_,_k//_/_ __ PAG_ 3-7
REVISED ST. LOUIS 3, MISSOURI REPORT _DR 104
R_,mSED MOZ_. 133
/i'
/
During the capsule landing and post landing phase, atmospheric air is drawn in
through the cabin air inlet valve to provide suit circuit ventilation.
3-4. SUIT EMERGENCY CONTROL
While operating in the suit environmental control mode during orbital flight,
the suit emergency control mode automatically activates when suit circuit pressure
decreases below 4.0 + .1 psia pressure. A control handle is provided to enable.3
manual selection of the emergency mode. During the capsule landing phase, the
emergency mode is automatically activated to increase Astronaut's cooling. The
environmental system oxygen rate valve and the suit circuit shutoff valve actuate
simultaneously to switch the environmental system frum the suit environmental
control mode to the suit emergency control mode. Actuation of these valves may
be either automatic or manual.
Operation in the suit emergency control mode, during orbital or landing
phases, basically consists of eliminating suit circuit oxygen flow through the
suit CO 2 and odor absorber, heat exchanger and water separator units. During the
landing phase, oxygen flow through the Compressor is maintained. During orbital
phase, oxygen flow is eliminated by deactivation of the Compressor. Elimination
of oxygen flow through these accessory components, while operating in the suit
emergency control mode, reserves the oxygen supply to remove the Astronaut's
generated heat, pressurize the Astronaut's pressure suit, and provide a breathing
source for the Astronaut.
An 02 EMERG light, located on the main instrument panel, and a tone generator
indicates when the environmental system is operating in the suit emergency mode.
Manual provision, to activate the suit emergency control, is located on the right
console.
_8 aA,.Jmm*A • • as *ram
Mc,DONN.ELLL z REPORT ,_ ] Oh ST. LOUIS, MISSOURI REVISED
MODEL. 133 " _ REVISED,
3-5. OXYGEN SUPPLY
The environmental system is supplied with oxygen, from primary and secondary
bottles. The primary and secondary oxygen bottles are directly interconnected
by a supply line, that forms a common oxygen supply to the cabin pressure control
valve, suit pressure regulator, emergency oxygen rate valve, and the suit circuit
water separator. The primary and secondary oxygen supply lines incorporate shut-
off valves, pressure transducers, pressure reducers, and check valves. The
pressure transducers transmit oxygen pressure, present in the primary and second-
ary oxygen bottles, to a dual quantity indicator, tape recorder, and to a tele-
metry unit. The primary oxygen bottle pressure is reduced to lO0 + 10 psig, by
a primary oxygen pressure reducer. Two primary oxygen pressure reducers are
provided to insure oxygen pressure reduction, in the event one pressure reducer
fails closed. The secondary oxygen bottle pressure is reduced to 80 psig by a
secondary pressure reducer. The primary oxygen supply reduced pressure, being
greater than the secondary oxygen supply reduced pressure, permits the primary
oxygen supply to be utilized during normal conditions with the secondary oxygen
supply in reserve. The oxygen supply line check valves prevent the loss of
oxygen, in the event either the primary or secondary oxygen pressure reducers
malfunct ion.
3-6. COOLING CIRCUIT
During normal capsule orbital flight, the environmental system cooling circuit
furnishes the cabin and the suit circuit with provisions for independently con-
trolling the cabin and suit circuit temperatures. Water is supplied, under
oxygen pressure, from the capsule water tank to the cabin and suit circuit heat
exchangers which in turn provides cooling by evaporation. The heat exchangers
absorb heat from the cabin and suit circuit oxygen and boil it off as steam.
CC_,_FIDELT_aL
DATE i November 1961
REVISED,
REVIS£D,
ST. LOUIS 3, MISSOURI
P^S_ 3-9,
mm:,OmT _I_R 10b, , ,
• -_ ..... Mo_m.. 133
The cooling circuit contains an independent oxygen bottle that provides
pressure necessary to pressurize the water coolant tank, enabling water flo_ to
the heat exchangers. The oxygen bottle pressure is used as a reference pressure
for the coolant quantity indicator. A llne, interconnected between the suit
circuit line and the coolant tank pressurization llne, insures water coolant tank
pressurization, in the event the coolant quantity indicating system malfunctioned
The cooling circuit basically consists of a water tank, cabin and suit tempera-
ture control valves, heat exchangers, and indicators. A coolant quantity indi-
cator, EXEESS CABIN H20 and EXCESS SUIT H20 lights are located on the main panel.
Temperature control valves are located on the right console.
3-7. OPERATION
The environmental control system is designed to sequentially operate auto-
matically during the launch, orbit, re-entry and post-landing phases of capsule
flight. The mode in which the environmental system is operated is dependent
upon the existing conditions within the cabin and suit circuits.
During the pre-launch phase of operations, the capsule oxygen and water
supply are fully serviced and a capsule preflight is performed. Refrigerated
air is ducted through the capsule hatch to pre-cool the capsule cabin and struc-
ture during capsule preflight. The refrigerated air supply is removed and an
external supply of freon coolant is directed to the cabin and suit circuit heat
exchangers, through the umbilical, to continue pre-cooling the capsule structure
and cabin equipment. The oxygen supply bottles shutoff valves are opened and
the Astronaut is connected to the capsule suit circuit by attaching the suit
circuit personal leads (flex hoses) to the Astronaut's pressure suit. The suit
compressor and cabin fan are activated. The suit circuit is purged with an
external source of low pressure oxygen applied through the suit circuit purge
valve. Following the purging operation, a suit circuit leakage check Is per-
.. .m i _ i am am
s- o .MC,DONNELL REPORT SED_ lO_ ST. LOUIS, MISSOURI REVISED
MODEL. 1_ _ _ REVISED
DATE i November 1961
formed with the Astronaut's faceplate closed. The capsule entrance hatch is
bolted into position and the capsule cabin is then checked for leakage and purged
with oxygen. The suit circuit incorporates provisions for obtaining launch purge
oxygen samples.
Forty-five seconds prior to launch, the ground umbilical plug is dlsconnectec
and freon coolant supply to the capsule ceases. During launch and orbit, the
cabin pressure relief valve maintains cabin pressure at approximately 5.5 differ-
ential (cabin/ambient) psia. During capsule launch, the suit circuit pressure
regulator maintains the suit circuit pressure approximately equivalent to cabin
pressure. The suit circuit oxygen is kept free of contaminants by a solids trap
and a CO2 and odor absorber. The solids trap removes foreign particles such as
food particles, nasal excretions, hair, etc. The CO 2 and odor absorber filters
odors and CO 2 from the circulating oxygen. Moisture from the suit circuit
oxygen is removed from the system by a water separator. The pneumatically
activated water separator deposits the moisture into a condensate tank. Cabin
and suit circuit temperatures are controlled by manually operated metering valves,
that regulate the water flow rate from the water coolant tank to the cabin and
suit circuit heat exchangers. Upon reaching altitudes where the saturation
temperature of water is lower than the cabin and suit circuit gas temperatures,
the cabin and suit circuit heat exchangers will provide cooling by water evapo-
ration.
Prior to capsule re-entry from orbital flight, the Astronaut positions the
temperature control valves to a COLD setting. When the capsule descends to an
altitude of approximately 21,0OO feet, the snorkel explosive door is ejected.
(Door is located on capsule exterior.) At an altitude of approximately 17,0OO
3000 feet, the cabin air inlet and outflow valves open barometrically venting
the cabin to the atmosphere. Operation of the suit circuit compressor draws out-
CO_F_'DEL'TIAL
1 ovember M!DONNELL 3-nREVISED. ST. LOUIS 3, MISSOURI REPORT ._R.T_R 1 04
REVISED • MOD_EL- 1_3
side air into the suit circuit through the ejected snorkel door opening, the
snorkel valve and the open cabin air inlet valve. The air, circulating through
the suit circuit, is relieved into the cabin and in turn flows out through the
cabin air outflow valve. Simultaneously, with the opening the cabin air inlet
and outflow valves, the environmental system mode of operation switches to the
suit emergency mode, but the suit compressor continues to operate. Switching
to the emergency mode provides a greater cooling capacity for the Astronaut.
When the capsule descends to approximately I0,000 feet altitude, the antenna
fairing is ejected. Ejection of the antenna fairing directs capsule power to
ignite explosive squibs, and in turn open the cabin air inlet and outflow valves°
(Opening of the cabin air inlet and outflow valves, when the antenna fairing is
ejected, is provided to supply ventilation air into the cabin in the event of a
capsule low altitude abort (below 17,000 feet) during capsule launch.) An inlet
air snorkel valve and an outflow air diaphragm flapper ventilation valve located
on the unpressurized side of the small pressure bulkhead -- opposite the cabin
air inlet and outflow valves, prevent water from entering into the cabin in the
event the capsule submerges after landing in a water environment. A vacuum
relief valve, located in the flexible ductlng between the cabin air inlet valve
and suit circuit, enables suit circuit ventilation whenever the inlet snorkel
valve closes. During the post-landing phase, the Astronaut may continue to
operate his suit circuit compressor to provide suit circuit ventilation. The
suit circuit compressor draws atmospheric air into the suit circuit, through
the cabin air inlet valve.
3-8. CABIN ENVIRONMENTAL CONTROL
Operation of the environmental control system in the cabin environmental
control mode, (Figure 3-3), after the capsule has entered the orbital flight
path; permits the Astronaut to open his helmet faceplate and be exposed to cabin
.J.J'_/_k IIPILI'_PIL l'l'll A I
REPORT S]EDR 104. REVISEDST. LOUIS, MISSOURI
133 ..........-- -" _" REVISEDMODEL. -,ovJ1z zJ_,.',-_ z *._
DATE i November 1961
environment for short periods. The cabin circuit also provides a manual method
for decampressing and repressurizing the cabin. The cabin pressure relief valve
relieves cabin pressure in excess of 5.5 psia. In the event cabin pressure tends
to exceed the 5.5 psia differential (cabin over ambient), the relief valve will
open to relieve the excessive pressure. In the event the cabin pressure decrease_
below 5 psla, the cabin pressure control valve will sense the pressure drop and
open. Opening of the cabin pressure control valve allows oxygen to flow into the
suit circuit. The suit pressure regulator will sense the increase in suit pres-
sure, and relieve excess pressure into the cabin. Routing the cabin pressure
control valve oxygen supply through the suit circuit, provides a constant purgiug
of the suit circuit. Cabin pressure control valve maintains cabin pressures to
5.1 +. "2 psia..3
During orbital flight, cabin gas is circulated throughout the cabin by the
cabin fan, located at one end of the cabin heat exchanger. The cabin fan forces
cabin gas through the cab_u heat exchanger. The cabin gas circulating through
the cabin, absorbs the heat generated by the cabin electronic equipment and in
turn is cooled when the gas passes through the cabin heat exchanger. Water from
the water tank circulates through the heat exchanger and absorbs the heat from
the cabin gas passing through the heat exchanger. The heated water evaporates
and passes overboard through the large pressure bulkhead steam vent. Regulating
the amount of water entering the heat exchanger provides cabin temperature con-
trol. A cabin temperature control valve, located on the right console, is
manually operated by the Astronaut to control cabin temperature.
In the event of a fire or a buildup of toxic contaminants, within the cabin,
the Astronaut may manually decompress the cabin by actuating the DECOMPRESS "T"
handle, located on the left console. The decompression handle is connected to
the cabin pressure relief valve with a cable. During decompression of the cabin,
, j
2;/¸.
DATE 1 November 1961
REVISED,
REVISED
ST. LOUIS 3, MISSOURI
PA6E 3-13
REPORT SEDR 104
MODEL 133
the cabin pressure control valve closes when cabin pressure decreases to 4.1
psia. Following fire extinguishment, or the removal of toxic contaminants, the
Astronaut may repressurize the cabin by closing the DECOMPRESS "T" handle and
actuating the REPRESS "T" handle. The REPRESS "T" handle is connected to the
cabin pressure control valve with a cable. When the cabin has been repressurized
to 5.0 psia, the REPRESS "T" handle must be manually closed, in the event of a
cabin decompression, due to a metorite penetration or excessive cabin leakage,
the cabin pressure control valve will close automatically and prevent oxygen flo_
to the cabin, after the cabin pressure decreases to 4.1 psia. Closing of the
cabin pressure control valve reserves the remaining oxygen supply for the suit
environmental control circuit, enabling the Astronaut to continue the mission.
Prior to capsule re-entry, the Astronaut should assure that his helmet
faceplate is closed, and pre-cool the cabin structure and equipment by position-
ing the cabin temperature control valve and suit temperature control valve to
the cold settings. During capsule descent, cabin pressure is maintained at
approximately 5 psia pressure, until the cabin altitude is approximately 27,000
feet. At 27,000 feet altitude the cabin pressure relief valve begins to open,
allowing atmospheric air to enter the cabin and equalize capsule internal and
external pressures. When the capsule reaches 17,000 + 3000 feet altitude, the
cabin air inlet and outflow valves open and the cabin fan ceases operation.
Opening of the cabin air inlet valve provides outside air ventilation for the
suit circuit. Suit circuit air is then vented to the cabin and out through the
cabin outflow valve. If the cabin air inlet and outflow valves fail to open
at 17,000 + 3000 feet altitude, the Astronaut should actuate the SNORKEL pull
ring to open the valves. In the event the Astronaut fails to open the cabin
air inlet and outflow valves, the valves will open automatically at lO,000 feet
when the antenna fairing is ejected. Ejection of the antenna fairing directs
-T_9_b.... _-::T:AL
PAGE,
REPORT
MODEL
3-14
SEDR 104
133
M,_DONN_L/z__ ___ST.LOUIS. MISSOURI
DATE ...... 1 NOVEMBER 1961
REVISED ....
REVISED
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REVlSEO. ST. LOUIS 3, MISSOURI R_ORT S_R 104
REVISED _: • " Moore. 133
electrical power to ignite the cabin air inlet and outlet valves explosive
squibs, which in turn mechanically open the valves. A snorkel valve, provided
on the inlet side of the cabin air inlet valve and a diaphragm flapper venti-
lation valve provided on the outlet side of the cabin air outflow valve, prevent
water from entering the cabin when the capsule lands in the water. Following
capsule landing, the Astronaut may operate suit compressor for ventilation.
A cabin pressure indicator and a cabin temperature indicator are provided
on the_ main instrument panel to indicate cabin pressures and t_ratures. A
humidity indicator, with an incorporated sensor and amplifier, is also provided
on the main instrument panel to indicate relative humidity content of the cabin
gas. A CABIN PRESS light and tone generator, located on the main panel, indi-
cate when cabin pressure has decreased below _.0 psia pressure.
3-9. SUIT ENVIRONMENTAL CONTROL
The suit environmental control circuit, Figure 3-_, is supplied oxygen from
the environmental system oxygen supply, through the suit pressure regulator and
the cabin pressure control valve. During capsule launch and re-entry phases,
when the Astronaut's helmet faceplate is closed, the suit pressure regulator
utilizes cabin pressure as a reference to control the suit circuit pressure.
While operating in the suit environmental control mode, (helmet faceplate closed):
oxygen from the suit pressure regulator flows through the suit compressor, CO2
and odor absorber, suit heat exchanger, water separator, Astronaut's pressure
suit, and the suit circuit solids trap. In the event the suit circuit oxygen
pressure decreases, (more than 2.5 to 3.5 inches of water below cabin pressure)_
the suit pressure regulator will sense the pressure drop and open, allowing
oxygen to flow into the suit circuit to maintain the suit circuit pressure
within 2.5 to 3.5 inches - water of cabin pressure. When not in operation, the
suit environmental control system is guarded against contamination by a neoprene
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PAGE 3 -17
REVISED ST.LOUIS, MISSOURI REPORT SEDR 104
REVISED w . _ MODEL. |33
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PAGE 3-18
REPORT _ lO4
•'__ 133
REVISEDST. LOUIS, MISSOURI
•n_'_T_Irlr_nnlr • T._ REVISED
coated nylon closure coupling. The closure coupling is attached to the ends of
the suit environmental control system inlet and outlet ducts.
The suit circuit incorporates two compressors that are installed parallel
to each other. During normal suit circuit operation the #I suit compressor
circulates the suit oxygen from the pressure suit outlet, throughout the suit
circuit. A differential pressure switch is vented to the inlet and outlet duct-
ing of the #I suit compressor. In the event the #I suit compressor malfunctions,
or fails to operate, the differential switch senses the pressure drop across the
#1 suit compressor, and in turn directs power to operate the #2 suit compressor.
A SUIT FAN switch is provided on the main instrument panel, to enable selection
of either compressor. Oxygen flowing from the compressors passes through the
CO2 and odor absorber. The absorber is divided into individual sections that
contain activated charcoal and lithium hydroxide removing odors and carbon
dioxide from the oxygen, to prevent any disccmfor_ to the Astronaut. Filters,
incorporated in the absorber, remove charcoal or lithium hydroxide dust from
entering the suit circuit oxygen.
Suit circuit temperature is controlled by a suit heat exchanger, that
removes heat from the suit circuit oxygen flowing through the heat exchanger.
Waterflow to the heat exchanger is controlled by a suit temperature control
valve, located on the right console. Suit circuit oxygen pressure is also
utilized to pressurize the water coolant tank in the event the coolant quantity
oxygen bottle, normally utilized to pressurize the water tank, malfunctioned.
Suit pressure and temperature sensors, located in the suit circuit, transmit
suit circuit pressure and temperature to the SUIT ERVIRON_ indicator. The
SUIT ENVIRO_ indicator, located on the main panel, is a dual face indicator
and indicates SUIT PRESS and SUIT TEMP. The water separator is a filter-type
sponge, that collects moisture from the suit circuit oxygen flOWlZ_ through the
':....._ C2NFID=NTIA, --
l ove ber Mc,DONNELL C , ,ffi 3-19REVISED ST. LOUIS 3, MISSOURI REPORT _R 10_
REVISED. _<"_ "" ' .... MODEL 133
i _" _ "*"iI
separator. At timed intervals, the sponge is pneumatically compressed to remove
the water from it. Water removed from the sponge is drained into a condensate
storage tank, located adjacent to the water separator. The sponge is compressed
by a piston that is actuated by oxygen pressure. The capsule programmer pro-
rides 24V D-C electrical power to energize the water separator solenoid valve for
30 seconds every 30 minutes. Energizing the solenoid valve, opens the valve to
allow oxygen pressure (lOO psi) to actuate the water separator piston. The
water separator piston incorporates two magnets. The magnets aid in determining
the position of the piston before and after actuation. Oxygen, flowing through
the water separator, flows to the closed emergency oxygen rate valve and through
the Astronaut's pressure suit. Oxygen, from the pressure suit, then passes
through a solids trap, that is provided to remove any foreign matter such as food
particles, hair, nasal excretions, etc., from the suit circuit oxygen supply.
The solids trap incorporates a relief feature, to prevent the possibility of
foreign matter blocking suit circuit flow. The suit circuit shutoff valve,
located downstream of the solids trap, is mechanically locked in the open posi-
tion during operation in the suit environmental control mode.
During capsule pre-launch phase, the suit circuit is purged and saturated
with oxygen from an external low pressure source. Suit heat exchanger is also
supplied with a freon coolant, from an external ground supply, to provide suit
circui¢ cooling. The suit circuit oxygen circulates throughout the suit circuit,
during the suit environmental control mode operation. During capsule flight the
pressure within the suit circuit is autamatlcally maintained at approximately 5
psia by the pressure regulator. During the landing phase, the capsule 21,000
feet barostats close at a capsule altitude of approximately 21,000 feet. Closing
of the 21,000 feet barostats directs 24V D-C electrical power to energize the
inlet air door relay. Energizing the inlet air doo_ relay directs power to
PAGE 3-20
REPORT .c_,DR lo4
MODEl 133
McDONNE ST. LOUIS, MISSOURI
F!
/
DATE ' 1 NoVember l_l
REVISED ,
REVISED, ,
ignite an explosive squib, which in turn ejects the snorkel explosive door.
(Door is located on capsule exterior. ) When the capsule cabin descends to an
altitude of 17,0OO + 3000 feet, the cabin air inlet and outflow valves open
barometrically. The suit compressor draws atmospheric air into the suit circuit
through the ejected snorkel door opening, the snorkel valve and the open cabin
air inlet valve. In the event the cabin air inlet and outflow valves fail to
open, the Astronaut may manually open the valves by actuating the SNORKEL pull
ring, located on the left console. If the Astronaut does not open the valves
manually, the valves will open automatically at I0,000 feet when the antenna
fairing is ejected. Ejection of the antenna fairing directs electrical power to
ignite the cabin air inlet and outflow valves explosive squibs, which in turn
mechanically open the valves. Opening of the cabin air inlet valve automatically
switches the environmental system mode of operation to the suit emergency mode,
but the suit circuit compressor continues to operate to provide suit circuit
ventilation. Also, opening of the cabin air inlet valve directs electrical power
to close the suit circuit shutoff valve, which in turn mechanically opens the
emergency oxygen rate valve and provides electrical power to illuminate the O2
EMER light and operate a tone generator. Electrical power is also/_rovided to
energize the emergency air inlet door relay, which in turn directs electrical
power to ignite the snorkel explosive door squib and eject the s_orkel explosive
jdoor. (This provision insures the ejection of the snorkel explosive door, in
the event the door failed to eject at 21,OOO feet. ) Air circulating through the
suit circuit is vented through the suit pressure regulator to the cabin, and in
turn is vented out of the capsule through the cabin outflow valve. During the
capsule post-landlng phase, ventilation is provided by operating the suit circuit
compressor, which in turn draws outside air in through the cabin air inlet valve
and vents cabin air out the cabin outflow valve. In the event the capsule sub-
C ll-i ,_l&TI'_ lr n m---_-- ...... _-. A_A, A lz_klJ
RI:_VISED. ST. LOUIS 3. MISSOURI REPORT _.'_R I0_
R=v,sED. _..................... _oo=L 133
merges momentarily, following a water landing, the ball float in the cabin air
inlet valve and the diaphragm flapper valves in the cabin air outflow valve
will seat. Seating of the valves (snorkel and flapper), prevents water from
entering into the suit circuit and cabin, through the open cabin air inlet and
outflow valves. Operation of the suit circuit compressor with the snorkel valve
closed will create a vacuum in the flexible ducting, located between the cabin
air inlet valve and the suit circuit. The vacuum relief valve, located in the
flexible ducting, will open when the pressure differential between the cabin and
flexible ducting is 10-15 inches of water. Opening of the vacuum relief valve
allows cabin pressure to enter into the suit circuit flexible air inlet ducting,
and unseat the air inlet snorkel valve ball float, if the capsule snorkel valve
is above the water. This action in turn allows outside air to enter into the
suit circuit to continue ventilation. During capsule submersion, cabin air
entering the open vacuum relief valve provides suit circuit ventilation,
3-i0. SUIT E_GENCY CONTROL
The suit emergency control, Figure 3-5, is provided to insure Astronaut's
survival in the event the cabin and suit environmental control circuits mal-
function. Operation in the suit emergency control mode basically consists of
opening the emergency oxygen rate valve, to supply oxygen at a rate greater
than normal; and closing of the suit circuit shutoff valve, which in turn
eliminates oxygen flow through the temperature control and impurity removing
units. !S_luminatiom of the 02 EMERG light and the movement of the EMERG 02
rate handle to EMERG position, indicates environmental system operation in the
suit emergency control mode.
When operating in the suit environmental control mode, during capsule
normal orbital flight, the emergency oxygen rate valve is closed, the suit
circuit shutoff valve is open, suit compressors are operative, and the suit
_T_'7 '_--.._'T'.;L
PAGE 3-22
REPORT b-'_DR 10/4•
MODEL_ 133
MCDONNE DATE 1 November 1961REVISED .....
. ST. LOUIS, MISSOURI• " REVISED
circuit pressure regulator is controlling oxygen flow to the suit circuit. The
emergency oxygen rate valve remains closed as long as suit circuit pressure
remains at approximately 5 psia pressure. In the event the suit circuit pres-
sure drops to 4.0 +. .1 psia, the rate valve internal anerlod extends, to offseat.J
a poppet, and allows oxygen from the oxygen supply to flow through the rate valve
and into the suit circuit. The extension of the rate valve aneriod, due to low
pressure, actuates a limit switch that provides electrical power to energize the
suit circuit shutoff valve solenoid and the suit fan cut-off relay, illuminate
the 02 E_RG light, and operate a tone generator. Energizing the suit fan cut-
off relay removes the llSV A-C electrical power to operate the suit circuit com-
pressor. (At an altitude of 17,000 +. 3000 feet, the cabin air inlet relay will
open. Opening of the cabin air inlet relay de-energizes the suit fan cutoff
relay. The de-energized suit fan cutoff relay routes power to the #1 suit cir-
cuit compressor. If the #1 suit circuit compressor fails to operate within 12
seconds, the suit fan selector relay will energize and allow power to be directed
to the suit fan cutoff relay and then on to the #_ suit circuit compressor.)
Energizing the shutoff valve solenoid releases the shutoff valve shaft arm, and
mechanically moves the EMERG 02 handle, right console, to the EMERG position.
Movement of the EMERG 0 2 handle moves a cable, that is connected to the emergency
oxygen rate valve shaft arm, and mechanically actuates the emergency rate valve
to maintain the rate valve in the open position. With the emergency oxygen rate
valve open and the suit circuit shutoff valve closed, oxygen from the oxygen
supply flows into the pressure suit and is discharged through the suit pressure
regulator relief valve.
Actuating the E_ERG 02 handle to the NORM position resets the shutoff valve
to the open position, the emergency oxygen rate valve to the close position,
starts suit compressor operation, extinguishes the 02 EMERG light, and in turn
REVISED ST. LOUIS 3, MISSOURI REPORT _ 10_
REV,sz. MOOE. 133
/
switches the suit circuit operation to the suit environmental control mode. The
suit emergency mode is also automatically selected durin_ capsule landing phase,
when the capsule has descended to an altitude of 17,000 + 3000 feet. At 17,000
+ 3000 feet the cabin air inlet valve opens. Opening of the cabin air inlet
valve actuates a limit switch that provides electrical power to operate the suit
circuit compressor and close the shutoff valve, which in turn mechanically opens
the emergency oxygen rate valve. An inlet power switch, located on the main
instrument panel to the right of the satellite clock, is incorporated in the
environment control system. The inlet power switch allows operation in the
suit environmental control mode in the event the cabin air inlet valve prematurel_
opens (See Figure 3-_). Premature opening of the cabin air inlet valve deacti-
vates the cabin fan and closes the suit circuit shutoff valve which in turn opens
the emergency oxygen rate valve. The suit circuit is now operating in the
emergency mode. To initiate transition back to the suit environmental control
mode, the inlet power switch is placed in the BY-PASS position. With the inlet
power switch in the BY-PASS position, the cabin fan is activated (See Figure 3-3)
and the suit circuit shutoff valve is deactivated. The E_BR 02 handle, right
hand console, is now placed in the NORM position; placing of the EM_ 02 handle
to the NORM position opens the suit circuit shutoff valve which in turn closes
the e_ergency oxygen rate valve. The environmental control system is now operat-
i_ in the suit envi_ntal control mode. To prevent snorkel door separation
upon premature opening of the cabin air inlet valve, the emergency inlet air
door relay is interconnected to the antenna fairing separation relay during
descent. After opening of the cabin air inlet and outflow valves, the inlet
power switch is placed in the NORMAL position.
3-ii. OXYGEN SUPPLY
During the capsule pre-launch phase and prior to installation of the capsule
C9::-:------_: .--:AL
PAGE _-24
REPORT ,_EDR 104
MODEL_ 133
Mc'DONNEL L ST. LOUIS, MISSOURI
DATE 1 November 1961
REVISED
REVISED,,,
entrance hatch, the capsule oxygen supply shutoff valves are manually opened, by
ground crewmen, to activate the oxygen supply. Opening of the primary and second
ary oxygen supply shutoff valves, Figure 3-6, provides oxygen to the cabin pres-
sure control valve, suit pressure regulator, suit emergency oxygen rate valve,
constant bleed orifice which routes oxygen directly to the suit circuit inlet
duct and the suit circuit water separator solenoid valve.
During operation, when the primary oxygen bottle pressure drops below
approximately 200 psig, due to near depletion of the primary oxygen supply; the
secondary oxygen supply line pressure will override the primary oxygen supply
line pressure and continue to supply the environmental system with oxygen. Two
oxygen flow sensors, one located in the primary oxygen supply llne and one
located in the secondary oxygen supply llne, detect environmental control system
transition from primary oxygen supply to secondary oxygen supply. Used in con-
Junction with the oxygen flow sensors is an oxygen flow switch (See Figure 3-6).
The oxygen flow switch, located on the main instrument panel to the right of the
satellite clock, is a two position switch marked "SEC" and "PRIM". When the
primary oxygen supply flow rate drops to approximately zero pounds per minute
at 80 psig, the oxygen flow sensor will direct a 24 volt d-c signal to illuminate
@
the 02 QUAN light and operate the tone generator. The 02 QUAN light, located
on the main instrument panel, and the tone generator are provided to indicate
to the Astronaut that the secondary oxygen supply is being utilized. The G2
flow switch should now be positioned to the "SEC" position. Positioning the 02
flow switch to "SEC" removes the 24 volt d-c signal from the 02 QUAN light and
tone generator circuits. A quantity indicator gage, located on the main instru-
ment panel, is provided to indicate remaining oxygen supply. Two transducers,
primary and secondary supply, are provided to enable telemetering of oxygen
quantity remaining.
DATE
REVISED
REVISED
I NOVEMBER 1961 Mc,DONNELL_H_,__ST.LOUIS, MISSOURI
PAGE 3-25
REPORT SEDR 104
MODEL 133
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FIGURE 3"_, PP-.IMAR'Y AND SECONDAR_ OXYGEN SUPPLY
PAGE 3 -26
REPORT _!_ID_ !C'/I.
MODEL 133
MCDONNE oA, REVISED.
ST. LOUIS, MISSOURI
........ ............ . " _. _ REVISED
3-12. COOLING CIRCUIT
The cooling circuit, Figure 3-7, basically consists of a water coolanttank,
cabin and suit temperature control valves, cabin and suit heat exchangers,
inverter heat exchangers, coolant quantity indication circuit, and an excessive
water warning circuit. During capsule pre-launch, after the entrance hatch has
been installed, cabin, suit circuit and inverter cooling is achieved by supply-
ing freon (F-f14) through the capsule umbilical connector and into th_ cabin,
suit and inverter heat exchangers. The freon coolant absorbs heat from the cabin
and suit air flowing through the heat exchangers, and boils overboard through
the environmental system steam vents, located in the large pressure bulkhead.
Freon metered to three inverter heat exchangers absorbs and dissipates heat from
the inverters in the same manner as the cabin and suit circuit heat exchangers.
Forty-five seconds prior to capsule launching, the freon coolant supply is dis-
continued. When the capsule reaches approximately 115,0OO feet altitude, in-
verter, cabin and suit circuit cooling is achieved by water evaporation, that
occurs within the inverter, suit and cabin heat exchangers. The coolant quantity
indicating system pressure relief valve remains open, during capsule launch,
until the pressure decreases to approximately 5-9 psia. The relief valve closes
at 5.9 psia pressure and in turn enables the coolant quantity indicating system
to operate.
Water from the water coolant tank is supplied, under a 5-5 psia pressure,
through the temperature control valves, to the suit and cabin heat exchangers.
Water is also supplied and me'cereal to the main and standby inverter heat
exchangers. Oxygen, stored under a 5OO psig presmxre in the coolant quantity
oxygen bottle, is utilized to pressurize the water coolant tank. A pressure
regulator decreases the coolant oxygen supply 500 psig pressure to 5.5 psia
pressure. A cooling circuit pressure relief valve relieves cooling circuit
C2NI'ID Li',;_-ta,_ .......
RL='VISED ST. LOUIS S, MISSOURI REPORT SEDR 10_
. MOO=. .... .
\
J
pressures in excess of approximately 6 psia. Oxygen pressure within the water
coolant tank tends to move the tank diaphragm, which in turn forces the water
supply out of the tank at a rate dependent upon the position of the temperature
control valves. In the event the coolant quantity oxygen supply should deplete
or malfunction, oxygen, at 5 psla pressure from the suit circuit, will continue
to pressurize the water coolant tank. The temperature control valves control
the amount of water entering the heat exchangers, and in turn controls cabin and
suit temperatures. Water within the heat exchangers absorbs heat from the cabin
and suit oxygen, flowing through the heat exchangers; The heated water evaporates
and flows out through the steam vents, located in the large pressure bulkhead.
Water metered to three inverter heat exchangers absorbs and dissipates heat from
the inverters in the same manner as the cabin and suit circuit heat exchangers.
In addition, heat generated by the Inverters is drawn into the cabin heat
exchangers by means of the cabin fan. Indicator lights are provided on the main
instrument panel to indicate extreme cold conditions in the cabin and suit heat
exchangers exhaust ducts, which could possibly freeze and plug overboard steam
vents. If the heat exchangers exhe_st ducts temperatures drop below 42°F, a
thermal switch, located in each of the exhaust ducts actuates close. Closing of
either thermal switch directs 24V D-C electrical power to illuminate either the
EXCESS CABIN H20 or the EXCESS SUIT H20 light, and operate the tone generator;
thus indicating to the Astronaut of extreme cold temperatures in the heat
exchangers exhaust ducts. The Astronaut must then position the cabin or suit
temperature control valve to a warmer setting, in order to reduce the possibility
of water freezing in the exhaust duct.
A coolant quantity indicator, main instrument panel, is provided to indicate
the quantity of water coolant remaining in the water coolant tank. The indicator
operates in direct relationship to the oxygen remaining in the coolant quantity
PAGE 3-28
REPORT SEDR 104
MODEL 133
Mc'DONNEL ST.LOUIS, MISSOURI
DATE ......
REVISED
REVISED
1 NOVEMBER 1961
_,_ TE.ST
_W iTCH
_4v DC
0
DIM._ NI BRT,
E NVI_ON MENTAL WARNING
CONTROL LI6NT SWITCH WATER
24V DC COOLANT TANK
CABIN TEMPERATURECONTROL VALVE
ORIFICE PREON CHECKVALVE
METERING VALVE
(WATER)
WATER CHECK
VALVE.
H
H
[1 METERING 'VALVE
(FREON)
FREON CHECKVALVE
COOLANT QUANTITY
OXYC_ N BOTTLE
(5ooPSi)
FREON CHEC}KVALVE
THERMAL
I SWITCH
I
I
TNERMALSW_TCH I
['
TOPRESSL RE
ORIFICE
'ILARGE PRESSURL
BULKHEAD
NOTESCHEMATIC DEPICTS COOLING
CIRCUIT OPERATION OURIN_
NORMAL CAPSULE FLIGHT
HEAT HEAT HEAT
EXCHANGER E_CHANGER EXCHANGER
_IIN INVERTER (5TBY INVERTER MAtH INVERTER250 _A.) Zso_A_ 150_A.)
I I ICAPSUCEUMBILEALPLU6
FIGURE 3-7 COOLING CIRCUIT
SUIT
CIRCUIT
FLOW
I'{'FREONSUPPLIEDDURING "_
_,GROUNDOPERATION ONLYJ
SUIT HEAT
EXCHANGER
PM4.5 -36B
P/"_ I_1E" I f"_ P" " " -- " -....... • .,.,_1 I I/all, l[--
,.oo....REVISED ST. LOUIS 3, MISSOURI MI_ORT. _D_ 10_.
R_'VISED , . ,i, MODEl.. 133
oxygen bottle, through a pressure transducer and instrumentation package. When
the coolant quantity oxygen bottle is full, (500 psi), the coolant quantity
indicator will read 100%. As the coolant oxygen bottle pressure decreases, as
a result of water being utilized, the coolant quantity indicator reading will
decrease accordingly.
3-13. SYSTEM UNITS
3-i_. PRIMARY AND SECONDARY OXYGEN BOTTLES
The primary and secondary spherical shaped oxygen bottles are located
beneath the Astronaut's support couch adjacent to the capsule conical section
and large pressure bulkhead. Each bottle has a capacity of 4 pounds oxygen,
stored under a 7500 psig pressure at 70°F temperature. Reduction of pressure
for utilization in the environmental control system is accomplished with pressure
reducers. The primary supply is reduced to approximately I00 psig; the secondary
oxygen supply is reduced to approximately 80 psig. Servicing of the oxygen
bottles is accomplished through a quick disconnect filler coupling.
3-15. SUIT CIRCUIT PRESSURE REGULATOR
The suit circuit pressure regulator, Figure 3-8, is provided to regulate
oxygen pressure to the suit circuit and to replenish suit circuit oxygen consumed
by the Astronaut, absorbed by moisture or carbon dioxide or lost through leakage.
The regulator is a demand type diaphragm operated regulator that controls suit
circuit pressure in reference to cabin pressure. Suit circuit pressure is main-
tained approximately 2.5 - 3.5 inches of water below cabin pressure during normal
system operation, under ideal (no cabin leakage) conditions. Cabin pressure is
sensed on the upper side of the regulator control diaphragm and suit circuit
pressure is sensed on the lower side of the diaphragm. The regulator also con-
tains a resilient type diaphragm that is used to relieve excessive suit circuit
PAGE 3-30
REPORT SEDR 104
MODEL_ I_
Mc,DONNELL __ DATE !" November l_l
REVISED• ST. LOUIS, MISSOURI
REVISED,,
pressures. Two anerolds are provided to shut off cabin vent port of regulator in
the event cabin pressure decreases below 4.6 + .2 psla. In the event of cabin
leakage, not decreasing below 4.0 + .2 psia, the cabin pressure control valve
will open to replenish cabin pressure to 5.1 + .2 psia. Make up oxygen from the- ,i
cabin pressure control valve will flow through the suit circuit and out through
the suit pressure regulator relief valve sad into the cabin. At this time, suit
circuit pressures will exceed cabin pressures due to pressure differential (suit
above cabin) required to open the suit circuit pressure regulator relief valve.
The cabin leakage rate will determine the amount the suit circuit pressures will
exceed the cabin pressures.
During normal capsule ascent, cabin pressure decreases, and the regulator
relief diaphragm relieves suit circuit pressure to within 2 - 9 inches _0 above
cabin pressure. During capsule normal orbital flight, the control diaphragm
will regulate suit circuit pressure in relationship to cabin pressure. An in-
crease in cabin pressure will act on the diaphragm to offseat a poppet valve and
allow suit Circuit pressure to increase to within 2.5 - 3.5 inches of H20 below
cabin pressure. In the event cabin pressure decreases below 4.6 + .2 psla, the
aneroids will extend sad close off cabin vent port of regulator. Two 60 cc/min
bleed ports will then bypass the poppet valve to the cabin sensing side of the
control diaphragm sad regulate suit circuit pressure to 4.6 + .2 psia. Two
aneroids and two bleed ports are provided to insure regulator operation, in th_
event either aneroid or either bleed port fails to function. Descent operation
of the regulator is the same as an increase in cabin pressure during capsule
normal orbital flight.
3-16. SUIT CIRCUIT SHUTOFF VALVE
The suit circuit shutoff valve, Figure 3-9, is designed to shut off oxygen
,<.
CCNFIDE- _;Tia,.
±IN_
t, Ol llQ_l$
"I3QOIN
J._lO,:t_ld lttFIOSSiiN ' SI FIO-I'.LS
a3Sl^31:l
Q3SIA31:i
_.LVCI
PAGE 3-32
REPORT SEDR 104
MODEL 133
Mc'DONNELL ' ST,LOUIS, MISSOURI
_%../I_II IL'/L'-i_i I I/-'ti=' _
DATE
REVISED,
REVISED
1 NOVEMBER 1961
SUIT CIRCUIT
FI(_URE 5-9. SUIT CIRCUIT SHUTOFF .VALVE
5HUTOFP VALVE
PIA45-152 A
pl_Llkl P I _ P IL • -- - " --•.._ v I ". I I I._ I.. i 1 i-'i rt I..
REVISED ST. LOUIS 3. MISSOURI REPORT _DR ].04
aEvlsED ........ MOD_',' .........133
flow to the suit environmental circuit accessory components, whenever the suit
circuit is operating in the emergency mode. Closing of the suit circuit shutoff
valve reserves the remaining oxygen supply for the Astronaut's pressure suit.
The shutoff valve, spring loaded to the close position, is latched in the open
position during normal suit circuit operation. Valve is maintained in the open
position by a solenoid controlled detent pin engaged into the valve spoon arm.
A micro switch, depressed by the valve arm, completes the solenoid circuit when
the valve is latched open. Opening of either the emergency oxygen rate valve
or the cabin air inlet valve directs an electrical signal to energize the shut-
off valve solenoid. Energizing the solenoid retracts the detent pin and allows
the valve spring to rotate the valve spoon to the close position. Closing of
the valve opens the solenoid circuit and opens the emergency oxygen rate valve,
through an inter-connecting cable. The shutoff valve is mechanically opened by
the EMER 02 control handle, located in capsule. The shutoff valve is inter-
connected to the emergency rate valve, so that when the emergency rate valve
closes, the shutoff valve opens.
3-17. EMERGENCY OXYGEN RATE VALVE
The emergency oxygen rate valve, Figure 3-10, is provided to supply a
regulated amount of oxygen directly into the Astronaut's pressure suit, in the
event malfunction occurs in the suit circuit operation. The rate valve is
designed to operate automatically and contains provisions for manual operation.
The valve, closed during normal suit circuit operation, contains an aneroid that
senses suit circuit pressure. Whenever suit circuit pressure drops below 4.0
+ .i psia, the aneroid extends to offseat a spring loaded poppet and allow oxygen.3
pressure to enter the diaphragm chamber. The pressure in the diaphragm chamber
increases and fully strokes the poppet, allowing oxygen to flow into the
Astronaut's suit at a fixed flow of .049 to .051 #/_Lin. Simultaneously with the
_Im,_mLL irmml imll, l,,mL |,mira I At 11
PAGE
REPORT
MODEL
3-34
SEDR 104
133 '
McDO__ __ST.LOUIS, MISSOURI
DATE
REVISED
REVISED
1 NOVEMBER 1961
/
----E.ME.R_F_.NO..YO_ f_ATE VALVE
F_O_ CAPSULE
,_...... ,/--MANUAL CONTROL
POPPET-__ _ CONTROL SWITCH
_o_ -_ _
FIGURE 5-10 EMERGENCY O_ RATE VALVE_h_45-i53
//
MCDONNE , 3-35REVISED ST. LOUIS 3, MISSOURI REPORT., _DR 10_
REVISED of . - ' MODEL 133
offseatlng of the poppet, a control switch is actuated through a lever mechanism,
and directs electrical power to close the suit circuit shutoff valve, illuminate
the 02 EMERG light, and stop suit circuit compressor operation during capsule
orbital flight. Suit circuit shutoff valve is interconnected with emergency
oxygen rate valve. Therefore, closing of the shutoff valve actuates the emergency
oxygen rate valve manual control shaft to close off oxygen flow to valve poppet
inlet. Oxygen then flows directly into suit circuit through the valve aneroid
chamber.
Emergency rate valve control shaft actuation actuates a pin to depress con-
trol switch and also moves EMERG 02 control handle (right console) to EMERG
position. Emergency oxygen rate valve may be opened manually by selecting EMER
position with EMERG 02 control handle. Operation will be same as control shaft
am operation, described above. Whenever the EMER 02 control handle is moved
to NORM, the suit circuit shutoff valve opens and emergency oxygen rate valve
close s.
3-18. SUIT CIRCUIT COMPRESSORS
The suit circuit environmental control system utilizes two electric motor
driven, single-stage, centrifugal compressors (See Figure 3-4). One compressor
is for the normal circulation of gases within the suit circuit; the other is a
standby compressor used in the event of normal compressor failure. If the
normal compressor fails, the standby compressor is activated by a pressure dif-
ferential switch which directs power to the standby compressor electrical connec-
tions. The only time the suit compressor is inoperative is during orbital flight
when the Astronaut is utilizing oxygen from the emergency oxygen rate valve.
When supplementary oxygen from the emergency oxygen rate valve is being used
below 20,000 feet, the suit circuit compressor will continue to operate to
circulate ambient air to the Astronaut.
3-36 Mc,DONNE ., DA, Uow= rREPORT _ I011" ST. LOUIS, MISSOURI REVISED
MODEL_ 13_ L,vJ.,,_. , .... , , REVISED
3-19. CO2 AND ODOR ABSORBER
The CO2 and odor absorber, Figure 3-I1, is provided to remove Astronaut
emitted odors and carbon dioxide frun the suit circuit. The absorber is basl-
cally a metal cannister divided into two sections. The inlet section contains
activated charcoal that removes obJectlonal odors from the suit circuit oxygen.
Lithium hydroxide, located in the center sections removes carbon dioxide. The
outlet section is an exit filter, provided to prevent charcoal and lithium
hydroxide dust frem entering the suit circuit oxygen flow. The charcoal and
lithium hydroxide granules are compressed by a spring force. The CO2 and odor
absorber has an operating life of approximately 31 hours, and to insure proper
absorber operation the absorber should be replaced prior to capsule mls sion.
3-20. SUIT CIRCUIT HEAT EXEHA_ER
The suit circuit heat exchanger (Figure 3-12) is of a plate fin construction
with rectangular offset fins, double sandwich, one pass on the oxygen side and
two pass, single sandwich on the water side. The function of the heat exchanger
is to cool the gases circulating throughout the suit circuit. Water from the
water cooling tank is routed to the inlet side of the heat exchanger at which
point it is directed to a high density woven felt pad. The function of the felt
pad is to evenly distribute the water through the heat exchanger. As water
passes through the felt pad, it comes into contact with the heat transfer sur-
faces on the water side of the heat exchanger. The water then absorbs the heat
from the circulating gases and the water is then boiled off as steam and dumped
overboard.
3-21. WATER SEPARATOR
The water separator, Figure 3-13, is provided to remove moisture, condensed
as a result of suit heat exchanger operation, from the suit circuit oxygen. The
CO _,'FID Z_A'A,A L
DATE
REVISED
REVISED
1 NOVEMBER 1961
Mc'DONNELjL-__ __ST,LOUIS, MISSOURI
PAGE 3-37
REPORT, $EDR 104
MODEL. 133
i I i /
1 II
FiGUP,E.5"I[. CO_ AND ODOR AIBSORBEP,
PAGE
REPORT
MODEL
3-38
SEDR 104
133
Mc'DONNE ST.LOUIS, MISSOURI
DATE ........
REVISED
REVISED ,
1 NOVEMBER 1961
L
_UIT HEAT EXCHA_C_ER
m
[_. ox'_ C___NOX",f C_E N F----Jx, x OUTLET
INLET
(FELT PA, D
FI_, E. ON _N LE.T
_'TE AIv_ _OUTLET
FIGURE 3-12 %BIT HEAT EY,CHAN_EP,
SECTIONSHOWIN_' OPERATION
PM45- _._G
' 'q,.*v m,_ m ! !'
l ov er M¢,DONNE , 3-39R_ISED ST. LOUIS 3, MISSOURI REPORT. _:_ 10t
REVISE=) _ MOOre. 133
separator is basically a filter type sponge that collects moisture from the
oxygen passing through it. The sponge, pneumatically compressed, removes the
condensate from the sponge and deposits it into a storage tank. During suit
circuit operation, the sponge filters moisture from the oxygen flowing through
the sponge. Once every 30 minutes, for a duration of 30 seconds, the capsule
programmer supplies electrical power to energize the water separator solenoid
valve. Energizing the normally closed solenoid valve opens the valve and directs
oxygen from either the primary or secondary supply to the piston stem and the
piston plate chambers.
Due to the difference in area, on each side of piston, the piston raises
the sponge out of the suit circuit oxygen flow and is compressed against the
separator housing plate. Water squeezed out of the sponge drains into the con-
densate tank. Following the termination of squeezing (30 seconds), the water
separator solenoid valve is de-energized and the solenoid valve closes. Oxygen
below the separator piston is vented to cabin through the separator solenoid
valve. Oxygen above the piston, entrapped by a check valve, forces the piston
down, thus returning the sponge into suit circuit oxygen flow. Suit circuit
pressure is also supplied to top of separator to_aid in forcing the sponge down.
Two check valves, located on top of separator, prevent water from entering the
suit circuit and also prevents water backflow into the top of separator. A check
valve is supplied to prevent oxygen depletion in the event separator mechanisms
developed a leak. During squeezing operation suit circuit oxygen flow will not
be affected, as oxygen will continue to flow through area normally occupied by
the sponge. The water separator piston incorporates two magnets. The magnets
aid in determining the position of the piston before and after actuation.
3-22. SOLIDS TRAP
The suit circuit solids trap, Figure 3-14, is located in the pilot's suit
PAGE
REPORT
MODEL.
3-4_0
, SEDR 104
133
M_'DONNELL_'•_ ____ST.LOUIS. MISSOURI
DATE1 NOVEMBER 1961
REVISED ........
REVISED
SPONGE
VRLVE
III
FROMOXY&EM,_UPPLY
OXYGEN SUPPLY
FIGUP.E _-IR W/kTE_ "3EPAtZATOD.
C _ I_1 !_" I I"_ l" _ "-- -
PM,_5-157B
DATE
REVISED
REVISED
1 NOVEMBER 1961 M_'DONNE_"4_ __ST.LOUIS, MISSOURI
PAGE
REPORT ......
MODEL
3-41
SEDR 104
133
/ .,
.FIQIjRE
-- PRIMARY FLOW
-- -- 5ECONDARY FLOW
SOLIDS TRAPpM 4_j- Z'_lh
--Will Ii./r-:-il I lnL
..o. - M_,DONNE___ °_......REPORT, SEDR 10,4
ST.LOUIS, MISSOURI REVISED
MODEL ] 3 3 r m-_J__ l_ P- L i -P_L.td _ REVISED
1 NOVEMBER 1961
W^$HERS I.,
" _NLI CAP:WE_
FIGURE 5-15 CALVIN HEAT EXCHANGEP-, AND FAN PM_,S-Z_TA
7 j/,'
CC"
REVISED. " ST. LOUIS 3o MISSOURI REPORT SEDR 10)4
REVmEO _ MOOEL l_,!D " ' ' '
oxygen outlet duct. The trap consists of a 40 micron mesh screen filter which
incorporates an integral bypass to insure operation in the event the trap would
become choked with collected solids.
3-23. CABIN HEAT EXEHANGER
The cabin heat exchanger (Figure 3-15), cools the cabin gas in the same
manner as the suit circuit heat exchanger except the cabin heat exchanger employs
an electric driven motor fan to draw in and reclrculate cabin gas. Internal
structure is the same as the suit circuit heat exchanger.
3-24. CAPSULE WATER TANK
The capsule water tank is a pressurized compartmented cylindrical shaped
container (Figure 3-15). Water is displaced from the tank by means of a rubber
bladder which is activated by oxygen from the Coolant Quantity Oxygen bottle.
Water from the tank is directed to two manual control valves which control the
water supply to the suit circuit and cabin heat exchangers. The capsule water
tank also provides the Astronaut with a source of drinking water.
3-25. CABIN PRESSURE CONTROL VALVE
The cabin pressure control valve, Figure 3-17, is provided to maintain
cabin pressure to 5.1 + .3 psla. The control valve contains two aneroids that
sense cabin pressure. Whenever cabin pressure drops below 5.1 + .3 psia, the
aneroids partially expand and unseat the spring loaded metering pins, which in
turn permit oxygen to flow into the suit circuit. The suit pressure regulator
senses the increase in pressure, and relieves suit circuit pressure to the cabin.
Directing oxygen flow through the suit circuit provides constant purging of suit
circuit. When cabin pressure increases to 5.1 +- .3 psia the aneroids contract,
allowing the metering pins to seat and shut off the oxygen flow. In the event
of cabin deccmpression, or whenever cabin pressure drops below _.0 + .2 psia,- .i
....... "_ "I"i #i I
PAGE,
REPORT
MODEL
3-44
SEDR 104
133
_ST,,LC_U_S, M_SSOURI
- qlmJ_.Jl_lli Ii._tJl_l i IX-_t.,.
DATE
REVISED ,
REVISED
1 NOVEMBER 1961
TANK
I
F IGURF. '3-16 CAPSULE WkTE.P, T_NK_M 4-5- Z58
CC.;r;GF_J_i J IAL
3-4_5
REVISED ST.LOUIS, MISSOURI REPORT SEDR 104
REVISED _ MODEL, 133
CABIN PRESSURE CONTROLVALVE
SPRING - LOA£)EDPOPPET
FIGURE 3-17 CABIN 9RESSURE CONTROL. VALVE pM4s-158 A
_vl111 II,,,,/bl"ll i I.r'l.l,,,,_
REPORT Sg'DR 10_ REVISEDST. LOUIS, MISSOURI
MODEL. 133 REVISED
the aneroids fully expand and seat against the inlet port. This prevents oxTgen
flow, through the cabin pressure control valve, and reserves the remaining o_ger
supply for the suit circuit. Two aneroids are provided in the valve to insure
operation in the event that one aneroid fails. A manual control is also pro-
vided to enable cabin repressurlzation in the event cabin depressurization was
manually selected. Actuation of the REPRESS "T" handle, in the cabin, offaeats
a spring loaded poppet in the valve and allows oxygen to flow directly into the
cabin. REPRESS "T" handle should then be pushed in, followin E cabin repressuri-
zation, to enable cabin pressure control valve automatic operation.
3-26. CABIN PRESSURE RELIEY VALVE
The cabin pressure relief valve, Figure 3-18, automatically controls cabin
pressure relative to ambient pressure durin 6 launch, orbit, re-entry and landing
phases. In the event of a water landin6, the valve incorporates provisions to
keep water from entering the cabin. The valve also features means for manually
deconpressing the cabin. The cabin pressure relief valve consists of a cali-
brated spring control section and e. poppet valve control section. The calibrated
spring control section incorporates ambient and cabin seusing chambers separated
by a sensing dtaphra_, spring loaded metering valves and calibrated sprin6s.
The poppet valve control chamber incorporates a manual control am, a check
valve, poppet stem orifices, spring loaded poppet valves, poppet chamber
diap_ and poppet chambers.
After the cabin pur_m 6 operation, cabin pressure will be the same as the
existing ambient pressure. During launch, as ambient pressure begins to de-
crease, the cabin pressure relief valve will relieve cabin pressure until the
pressure differential (cabin/ambient) reaches 5-5 psia. The cabin pressure
relief valve will then prevent cabin pressure build up in excess of 5-5 ps:l.a
• ._::/
DATE 1 NOVEMBER 1961
REVISED
REVISED
McDONNELL__ __ST.LOUIS, MISSOURI
PAGE 3-47
REPORT .............. SEDR 104
MODEL 133
• /
CABIN P_F-SSU_E _F.L_EFVALVE
l I
f..' - '.N C
F_'Y_" //////////////__V///////////////px/ _'_""--------cH,_B_R"CABIN . _ _ '_. _ _" _ _MANUALDIAI:_RA_M / I_1 21_1 I _ I _ , Y_I _ _CoNTROL"CHAMBER'-'_- _._ - L_I I _ ' I r_l I M _ _RM
_EEP_F0"_ _,_ o_PFPI_Z STEM
CAISIH . F/J I_I' Y//J ,_/_" '11 _ _ _""---CHECKVALVE
- . _-_'_'/_ _ ________._ __ __ CABIN.
F_C)PPET _ _ _'- POPPET
_'_ ......._2_ ,;_:/_........................v,_v_FIGURE _-i8 CABIN P_ES_J_E ._ELIEF VALVE
t_'_Jlt_lllt. Jlllllltllllil|llllll i •
_111_ "II1_ l ]1 l I II/ i111 | I I J--'_ IIII
PM45- _
PAGE
REPORT
MODEL.
McDONNELL z_T_R 104 REVISED
ST. LOUIS, MISSOURI
throughout the orbit, re-entry and landing phases. Cabin pressure will be
vented, through the poppet stem orifices, into the poppet valve chamber. Cabin
pressure will also be vented, through the cabin air filters, into the cabin
sensing chamber. Ambient pressure will be vented, via the ambient port, into
the ambient sensing chamber. The calibrated springs are designed to respond to
differential pressures in excess of approximately 5.5 psi (cabin/ambient). When
the pressure differential between the cabin sensing chamber and ambient sensing
chamber exceeds approximately 5-5 psla (cabin/amblent), the calibrated springs
will retract. The metering valves will then be lifted from their seats, allowing
differential pressures in excess of 5.5 psia to escape through the ambient port.
Due to the ambient port being larger than the poppet stem orifices, the dissipa-
tion rate of the excessive differential cabin pressure (inside the poppet valve
chambers) will exceed the rate of buildup in the poppet valve chambers. This
will momentarily cause the cabin pressure to be greater than the poppet valve
chamber pressure. The greater cabin pressure _-ill act against the cabin dia-
phragm, unseating the poppet valves. The poppet valves will then aid in the
relieving of excessive differential pressure. If the Astronaut executes a manual
decompression of the cabin, the check valve acts as an exhaust for poppet valve
chamber pressure.
During orbit, the cabin pressure relief valve will prevent cabin pressure
buildup in excess of approximately 5.5 psia. Cabin pressure in excess of
approximately 5-5 psia will be exhausted to the outside atmosphere. Upon re-
entry, when the ambient pressure becomes 15 inches of water greater than cabin
pressure, the poppet valves will commence to open allowing ambient pressure to
enter the cabin. Valve relieving operations will then be similar to those dur-
ing launch. In the event the capsule makes a water landing, the poppet valves
will not open until water pressure exceeds cabin pressure by 15 inches of water.
.,j
• k.,IV_.'t,l. Jl. Jl.Jrl:,{'l-_'_l.''_._-_ -
REVISED ST.''f ------_- ---7"-_LOUIS 3, MISSOURI REPORT _..DR 10_
REV,S_O -" _ " ..... Moo_ 133
/
A manual control located on the left hand console and marked "PRESS MEG" manually
closes the poppet valves to prevent water frQm entering the cabin through the
valve.
3-27. SNORKEL AND DIAPHRAGM FLAPPER VALVES
The cabin inlet air snorkel valve and the cabin outflow diaphragm flapper
valve act as water check type valves. During the landing and post landing phases
(often reaching a pressure altitude of approximately 17,0OO ft.) ambient air is
circulated through the valves. In the event the valve parts were under water,
the valves would seat and prevent water from entering the cabin.
3-28. CABIN AIR INLET VALVE
The cabin air inlet valve, Figure 3-19, provides ventilation and cooling
for the suit circuit and cabin during capsule landing and post landing phases.
It is a spring loaded close, spoon type valve and is barometrically controlled.
Prior to capsule launch, the valve is manually latched closed so that one
mechanism spring loaded detent pin rides on the large diameter of the aneroid
plunger (maximum allowable pull to set detent pin is five (5) pounds); and the
valve arm is engaged by the release link, which is engaged by the spring loaded
aneroid locking pin. During capsule launch the aneroid expands due to decreasing
cabin pressure, and forces the aneroid plunger down. The valve mechanism detent
pin then slips off the plunger large diameter onto the plunger small diameter.
During capsule landing phase, when the capsule descends to an altitude
of approximately 17,000 +. 3,000 feet, the aneroid retracts as cabin pressure
increases. Retraction of the aneroid moves the aneroid plunger upward, engaging
the detent pin against the plunger larger diameter, which in turn compresses
the aneroid locking pin spring. This action raises the locking pin from release
llnk and allows spring loaded valve to open. The valve arm is attached to valve
PAGE
REPORT
MODEL.
3-50 MCDONNELL '_ _//_ DATE,, 1 NOVEMBER 1961
ST.LOUIS. MISSOURI REVISED
REVISED
SEDR 104
i33-- vv • |w
EIC-I.URE _>r:19 _CABIIq AIR INLET VALVE
CC;;.;GZN, _A,'PM 4.5-/$4 I_,
:.Z.Y
o mber Mc,DONNE , , 3-51REVISED ST. LOUIS 3, MISSOURI Rli_OR'i" _-_)R 10)4
REVISED . t-' MODEL 133
shaft and moves with closing, thereby disengaging micro-switches. Disengagement
of micro-switches directs electrical power to stop cabin fan operation and close
the suit circuit shutoff valve, which in turn opens the emergency oxygen rate
valve. A manual control arm is provided to enable valve opening in the event
valve failed to open at specified altitude. Actuation of the manual control arm,
mechanically, contacts the locking pin spring and disengages locking pin from
release link, allowing valve to open. In the event of a capsule low altitude
abort, an explosive squib will force the locking pin up, to enable valve opening.
Valve must be manually reset to close position. Opening of the valve enables
suit ccmpressor to draw ambient air into suit circuit to provide suit circuit
and cabin ventilation.
The cabin air outlet valve is basically of the same construction and
functions in the same manner as the cabin air inlet valve.
3-29. VACUUM RELIEF VALVE
The vacuum relief valve, Figure 3-20, is designed to open at a pressure
differential of lO to 15 inches of water, to provide suit circuit ventilation
whenever the inlet snorkel valve closes (ball float seats). The relief valve is
located in the flexible ducting, between the cabin air inlet valve and the suit
circuit inlet duct. In the event the capsule submerges momentarily, following
a water landing, the snorkel valves ball floats will seat (close) and prevent
water from entering into the suit circuit and cabin. The operation of the suit
circuit compressor and the closed air inlet snorkel valve will create a vacuum
in the suit circuit air inlet duct (flexible ducting). When cabin pressure
exceeds the flex duct pressure, by lO - 15 inches of water, the vacuum relief
valve will open. As the valve opens, cabin pressure acting on the valve poppet
surface will be great enough to hold the valve open until the pressure differ-
ential (between cabin and duct) is approximately 2 inches of water or less. Suit
via .u Ita,,Ji-..INI I l_b'IL t-.-*
PAGE
REPORT
MODEL
_'_ Mc'DONNE_ C__ DATESEDR 104 ST.LOUIS, MISSOURI REVISED
133 C,_T_-_ i _ I,,_,,__[_ REVISED
1 NOVEMBER 1961
VACUUM RELIEF" VAL'_..
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FIGURE 3-20 _/ACUUM RELIEF VAWE
AA _ I =
mATE 1 November 1961
REVISED
REVISED,
ST. LOUIS a, MISSOURI REPORT SE_R 10_
MOOEL 133
circuit ventilation is provided by the cabin air, entering the opened vacuum reli_f
valve, whenever the inlet snorkel valve ball float is seated (closed). Also, the
opening of the relief valve removes the vacuum in the flex duct to enable the
snorkel valve ball float to unseat (open) whenever the snorkel valve is above
water.
3-30. TEST CONFIGURATION CAPSIK_S
3-31. TEST CONFIGURATION CAPSULE NO. 8
Capsule 8 environmental control system differs from the specification
capsule in the following manner.
Capsule 8 environmental control system is the same as Capsule 2 and 3
except for the following: Capsule 8 will incorporate a cabin 02 pressure warn-
ing light but this warning light will not be supplemented with a tone generator
circuit. There is no indicating system to reflect 02 emergency operation.
Capsule 8 will incorporate a circulation fan to aid in cooling the special
instrumentation package. The air inlet and outlet valves open at 17,000 .+
3,000 feet. The air outlet valve is the ssme as Capsule 2 and 3. Capsule 8
does not contain heat exchangers to cool the main and standby inverters, nor
does it contain manual provisions for closing the cabin pressure relief valve
poppet valves.
3-32. TEST CONFIGURATION CAPSULES NO. 9, !0 AND 13
Capsules 9, I0 and 13 environmental control systems are basically the same
as the specification capsule except as follows: Capsule 9 solids trap is
designed for a primate; it does not contain heat exchangers to cool the main
and standby inverters; the water separator is deactivated; and the water tank
is not utilized as a source of drinking water for the Astronaut. Capsule 9
suit circuit does not incorporate a suit 02 partial pressure system; and the
-__D;.F;D=;4T;AL
PAGE , 3-5_
REPORT SEDR 104
MODEL_ 133
Mc,DONNELL __ DATE 1 November l_lREVISED
._T..Aou!s,. MmSOUm-_v,, ..... _ REVISED
suit circuit does not incorporate a constant bleed oxygen system.
3-33. TEST CONFIGURATION CAPSULE NO. 16
Capsule 16 environmental control system is basically the same as the
specification capsule.
4-1
SECTION IV
STABILIZATION CONTROLSYSTEMS
TABLE OF CONTENTS
/, ,.
i I¸ i),.;'j'
TITLE PAGEIII ii
AUTOMATIC STABILIZATIONCONTROL SYSTEM
SYSTEM DESCRIPTION ....................... 4-- 3
SYSTEM OPERATION .......................... 4-12
SYSTEM UNITS .................................. 4-19
RATE STABILIZATIONCONTROL SYSTEM
SYSTEM DESCRIPTION ...................... 4"-24SYSTEM OPERATION ......................... 4-25SYSTEM UNITS .................................. 4-25
REACTION CONTROL SYSTEM
SYSTEM DESCRIPTION ...................... 4-2'7
SYSTEM OPERATION ........................ 4-29
SYSTEM UNITS ................................. 4-36
HORIZON SCANNER SYSTEM
SYSTEM DESCRIPTION ...................... 4-40
SYSTEM OPERATION ........................ 4- 41
TEST CONFIGURATION CAPSULES ........ 4-55
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El'. LOUIS 3, MISSOURI
PAGE _"-3
Rli_PORT. _EDR 104
MOC,_L.....,133
z¸
IV. STABILIZATION CONTROL SYSTEMS
4-1. GENERAL
Stabilization of the capsule in space is accomplished by the Automatic
Stabilization Control System in conjunction with two sub-systems, the Horizon
Scanners and the Reaction Control System. These systems establish and maintain
a stable platform with four basic automatic modes; Damper, Orientation, Attitude
Hold and Re-entry. A redundant rate "back-up" system, the Rate Stabilization
Control System (RSCS), is also provided. The RSCS provides the Astronaut with
an emergency method of controlling the capsule with a "rate-stick" in the event
of a failure in the Automatic Stabilization Control System. In addition, a visual
indication of yaw, roll, and pitch attitude is provided. The following paragraphs
4-2 through 4'56 briefly describe the individual systems and functions involved
for the specification compliance capsule.
AUTOMATIC STABILIZATION CONTROL SYSTEM
SYSTEM DESCRIPTION
The Autcmatic Stabilization Control System (ASCS) is composed of a
Directional Gyro, Vertical Gyro, .05g accelercmeter switch, Rate Gyros (yaw,
roll and pitch), a rate damper, and an Amplifier Calibrator Unit. Location of
the individual components within the capsule is shown in Figure 4-i. Total
weight of the ASCS is approximately 59 pounds.
Three switches are provided in conjunction with the ASCS. The GYRO switch
AD_O/RATE COMD switch, and the NORM-AUX DAMP-FBW switch are located on the
Astronaut's left console. With the NORM-AUX DAMP-FBW switch in the NORMAL posi-
tion and the AUTO-RATE COMD switch in AUTO, stabilization is accomplished in a
completely automatic manner, requiring no assistance from the Astronaut. In the
FLY-BY-WIRE position, the automatic feature is disabled and 24 V d-c power is
_ ' _T r_ ,r, mll, j L ,i
REPORT SEDR 10_ ..... ST:..LOUI$,.. MISSOURI REVISED
MODEL. 13_3 f,_ __n_____.4 REVISED
connected to the Fly-By-Wire limit switches on the Astronaut's control stick.
In this position stabilization is accomplished through an electro-mechanical
arrangement (See Figure 4-9) by movement of the Astronaut's control stick in the
desired plane. Low and high thrust actuation occur at approximately 30% and 75%
of full travel, i.e., 3.9 ° and 9.8 ° for yaw, pitch, or roll. The AUX DAMP posi-
tion disables both the automatic and fly-by-wire function, permitting rate
damping as a singular feature. The GYRO switch is a three position switch in-
corporating a CAGE, FREE, and NORMAL position. In the CAGE position the Attitude
gyros are mechanically caged and the Horizon Scanner slaving function is dis-
abled. In the FREE position theAttitude gyros are uncaged; the Horizon Scanner
slaving function remains disabled. The NORMAL positlon uncages the attitude
gyros and permits Horizon Scanner slaving. The AUT0/RATE COMD switch provides
a method of energizing either the RSCS or ASCS systems as desired. In the RATE
COMD position, the attitude gyros and slaving circuits remain energized although
they are not used to control the capsule. (See Figure 4-12. )
4-4. ASCS Sequencing
The following paragraphs, 4-5 and 4-9, describe the ASCS sequential opera-
tion under normal and abort conditions. Figures 4-2, 4-3 and 4-4 are provided
for clarity and should be followed closely in conjunction with the text concern-
Lug the various modes of operation.
4-5. Normal Sequencing
In Figure 4-2, the progress of a normal orbital mission is shown divided
into eight phases appropriate to the following discussion.
The ASCS is in the "ready" status prior to separation of the escape tower,
its gyros are running and all circuits except the final 12 output relays are
fully energized. RSCS operation is prevented by the AUTO/RAT_ COMD switch being
in the AUTO position. Phase (A), involving gyro slaving to the Horizon Scanner
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PAGE 4-6
REPORT, _DR 104
M ,DONNELL REVISED
ST. LOUIS, MiSSOURi
MODEL- 1,_3 ""_ _'"_: _, ..__._' _""._ REVISED_ _;""_2_2J" J, _l " " _ _ rr_ ,
pitch and roll outputs during ascent s is to minimize _o errors which may
accumulate while the capsule is being boosted.
Phase (B) starts after capsule separation when a brief, flve-second signal
commands the ASCS to provide rate damping to stop any tendency to tumble.
Phase (C) is initiated at the completion of five seconds of rate damping.
The ASCS is placed in the orientation mode, capsule turn around (180 ° counter-
clockwise Yaw Rotation) is accomplished, and the capsule is pitched-down to the
retrograde firing angle within 30 seconds. Pitch, roll and yaw gyro slaving to
the Horizon Scanners is provided during phase (C) and for the first 4_ minutes
of phase (D) to yield a good yaw-angle reference prior to settling down in orbit.
In phase (D) the capsule is in orbit. An orbit pitch attitude of -34 °
(small end down) is held so that the capsule is ready for an immediate abort.
The attitude gyros are slaved to the Horizon Scanner at a cycle of 8.5 minutes
on, 9-1.5 minutes off, throughout phase (D). During the orbit phase manual con-
trol and fly-by-wlre control may be utilized as desired. Rate damping becomes
optional under manual control conditions by positioning of the ASCS MDDE SELECT
switches and RCS controls. By switch manipulation, rate damping is provided by
either the ASCS or RSCS. See Paragraph 4-3. Rate gyro run-up is continued
throughout phase (D). Another feature utilized in Phase (D) is an automatic
return to the orientation mode. If the capsule drifts (from orbit attitude)
beyond the limits of the retro-interlock sector switches, automatic return to
orientation mode will occur at + 12° pitch, + 30° yaw and roll. The Astronaut
can also place the ASCS back in the orientation mode by manipulation of the
AUTO/RATE COMD switch.
In phase (E) of Figure 4-2, rate gyro run-up is automatically assured by
relay switching lO minutes prior to retrograde attitude. Horizon Scanner slaving
operation becomes constant at this time to obtain a good reference prior to
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REPORT SEDR 104 ....... ST. LOUIS, MISSOURi REVISED
MODEL 133 ....... ' REVISED
achieving retro-attitude. The Astronaut may change any one or all three of the
capsule attitudes maintained by the ASCS by changing the space reference plane
or planes of the attitude gyros. To maintain the new reference plane or planes,
the Horizon Scanner slaving cc_mand must be stopped by placing the gyro switch
in the Free position. New reference planes may be established by the Astronaut
while the ASCS is in operation by placing the gyro switch in the Free position,
manually turning off the ASCS fuel in the axis or axes affected, utilizing
manual control to position the capsulej and then caging and uncaglng the gyros.
The ASCS may then be returned to fully automatic operation in all three axes
with the exception of Horizon Scanner slaving. To utilize Horizon Scanner slav-
ing, the capsule attitudes must be within the observation range of the scanners
and the gyro switch must be placed in the Normal position.
The ASCS receives Tr signal and maintains capsule in high torque retro-
grade attitude (phase F). Horizon Scanner slaving is discontinued at this time.
Thirty seconds after retrograde attitude command, the retro rockets are fired.
During the period of retrograde rocket firing the ASCS utilizes h_gh torque
action to hold the capsule within one degree of the ideal angles. Retrograde
rocket firing command and ASCS high torque switching command occur simultaneously
Rocket firing is completed in 20 seconds and the high torque switching command is
held for 23 seconds.
Upon completion of retro package Jettison, the ASCS automatically pitches
the capsule to the post-retro fire attitude (phase G) in preparation for re-
entry drag. The ASCS returns to orientation mode with constant scanner operation
to accurately maintain the re-entry attitude.
Finally, when re-entry is sensed by the .05g accelerometer switch, the
eighth and last phase (H) of the ASCS performance starts with the turning off
of the attitude gyro power. During this period the ASCS initiates and maintains
REVISED ST. LOUIS 3. MISSOURI REPORT,,,,I., _F_R 104
.Ev,s_o _ MOrnS. ,,, 133
r'
a constant roll rate of i0*%o 12° pe_ second to minimize touchdown dispersion.
Rate damping is provided to stabilize the re-entry trajectory. ASCS operation
in this phase continues until main chute deployment, at which time all ASCS
power is removed.
Pilot-overrlde provisions permit interruptions of the preceding "normal"
sequence by manual, fly-by-wlre or RSCS stick-steering control manipulation and.
return to the "normal" ASCS MODE. Thus to a significant degree the Astronaut is
the intelligent "back-up" for the ASCS, Full utilization of this reliability
augmentation principle has led to gyro caging and other switching features which
are intended to make the capsule manually controllable. The following table
lists the switch and control positions necessary to achieve the four basic modes
of control after attaining orbit. Variations of the various modes can be
obtained by further switch manipulation. A more detailed discussion of the con=
trol modes available is contained in SEDR 109, Capsule Flight Operations M_nual.
RCS "T" HAttDLE POSITIONSWITCH POSITION
CONTROL MSEE Auto/Rate Cored Norm-Aux Damp-FBW Auto Fuel M_nual Fuel
sys. Sys.
AUTOMATIC AUTO NORMAL PUSH ON RATE COMD
FLY BY WIRE AUTO FLY BY WIRE PUSH ON RATE COMD
RATE-STICK RATE COMD NORMAL PULL OFF RATE COMD
DIRECT AUTO AUX. DAMP PUSH ON DIRECT(With Rate
4.6. Abort Sequencing
In general, abort sequencing (See Figure 4-4) is programmed to correspond
to the safest procedures at all times. The possible abort situations can be
divided into three types, namely (I) abort, before tower separation when ASCS
rate damping is required; (2) abort after tower separation but before the
trajectory is truly orbital; and (3) abort fr_a orbit. The following paragraphs,
¥'_-, .a il./r--i_l | I/_i. -
Mc,DONNELL
REPORT $EDR 104 ST.LOUIS. MISSOURI REVISED
MODEL 133 --_ _'_ _ _ e._i.._P" _ T _J-'_" REVISED
1 NOVEMBER 1961
FIGU_E4-¢ .ASCS EMEI_GENCY OPEI2/kTION DI44_" 39£
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ST. LOUIS 3. MISSOURI
DATE 1 N_z_bel 4 1_1 PAGE_ 4-11
REVmED R_oxr SEDR 104
RSV,SED _ eOO_ 133
4-7 through 4-9_ dlscuss _CS sequencing in each of the abort conditions.
4-7. Abort Before Tower Separation
If an abort mission Is started durln 6 the perto_ when the booster and sus-
talner eng/nes are burnlngj the ASCS Is utilized for rate damping only after the
following external operations have been achieved.
(1) Booster and sustainer engines cut-off.
(2) Capsule separation from adapter.
(3) Escape tower rocket firing.
(4) Retro rocket separation from capsule.
(5) Timed arrival at approximate peak of trajectory.
(6) Separation of escape tower from capsule.
Upon completion of the latter operation, the ASCS is coezuu_led to provide rate
dampinS, using the rate gyros which are continuously energized dur'l]_g the normal
ascent and "abort trajectory" flight. A constant roll rate of 10 ° to 12 ° per
second is employed. Rate dsmping ceases upon deployment of the main chute.
4-8. Abort After Tower Separation
The flrst operation is engine cut-ofT. This is followed immediately by
capsule separatlon_ posl_de firing, and the normal mission post-separation
signal sequence tO the ASCS. The effect Is i_nedlate dsmplng of any capsule
tendency to tumble. After 5 seconds of rate damping, the automatic sequence
co,ands capsule turn around an an attitude angle of 34 degrees. Then either
the Astronaut or Ground Command must initiate retrograde sequencing. Upon
achievSug the proper roll_ pitch and yaw an61es within rather wide "permission"
bounds (See _ph _-Sj Page 4-6)_ the ASCS enables rapid-sequence retro
rocket _ to proceed.
NOTE
ASCS "permission interlock" during retro flre can be over-
i
PAGE _-12
REPORT _. _R _0_
MODEL. 1_
McIDONN_ _4_ DATE ! __ovember l_lREVISED
.....Slr. LOUIS, MISSOURI
_---'_--JCI_T 1C_TT_ ]_ REVISED
ridden at any time by the Astronaut. It is also noted that
the Astronaut may switch to the Rate Stabilization Control
System at any time should a malfunction occur in the Automatic
Stabilization Control System.
After retrograde operation, the abort mission in this case proceeds as in the
normal mission post-retrograde sequence (except for the difference in trajectory
time and distance intervals).
4-9. Abort From Orbit
Whenever an abort from orbit is initiated, the normal automatic or manual
retrograde operations will apply. However, if manual retrograde operations are
utilized the pre-retrograde period of gyro slaving to the Horizon Scanners
("last look") will be eliminated.
4-10. SYSTEM OPERATION
Overall system operation is best explained by Figure 4-5. The Amplifier
Calibrator receives inputs from sensors on the left side of the page and generate
outputs to Display sad Reaction Control devices on the right. The four basic
operations are slaving, repeating, mode switching and torque switching. Data
flow pertaining to the individual Yaw, Roll and Pitch channels is illustrated
in Figures _-6, 4-7 sad 4-8. In general, these diagrams are straightforward
and require no explanation. However, the method utilized in deriving Directional
information is unique to a degree and warrants the following discussion.
The Pitch gimbal (vertical gyro) is processed continuously during the
orbital phase of the normal mission, so that the capsule "local vertical "_
reference revolves 360 degrees during each orbital cycle. The gyro slaving
principles which permit Directional (yaw) information to be derived are as
follows: After initial slaving and settling of the roll sad pitch loops, the
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PAGE 4-16
REPORT SEDR 104
MODEL 133
Mc'DONNEL.__.L_j_ C-_,,_,_ST.LOUIS, MISSOURI
DATE
REVISED
REVISED
1 NOVEMBER 1961
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_SCS controls the capsule to the command pitch attitude, and to level roll
attitude. Initially, after separation and capsule turn-around, some yaw error
(as great as I0 degrees) may exist due to directional drift during boost. Since
the Roll gimbal of the vertical gyro is the inner gimbal, yaw misalignment of
the capsule causes the Roll glmbal output to contain an error component due to
the constant orbital _itch) angular rate. Thus a comparison of the Roll Horizon
Scanner and vertical gyro roll indications will provide an error signal producing
a roll gimbal torquing rate. This torquing rate which is a direct function of
yaw error is used to slave the yaw gimbal of the directional gyro.
Another area that warrants discussion is that of torque switching, i.e.,
the thrust output of the Reaction Control System in conjunction with the various
modes of ASCS operation.
Figure _-i0 serves as an introduction to the torque switching behavior of
the ASCS. For maximum conservation of control fuel, the behavior varies accord-
Ing to the ASCS mode appropriate at a given time. A so-called "phase-plane" plot
of angular rate vs. angle is shown in the lower right corner of Figure 4-i0
adjacent to a typical Pitch time-history for the "Orbit" mode. Current ASCS
design permits a plus or minus three degree oscillation about the nominal orbital
attitude, which in turn is referenced to the Horizon Scanner's sensed "Horizontal".
The oscillation is non-simusoidal because of the discontinuous torque program;
pitch rate is a square wave, and pitch angle a sawtooth, both having a
characteristic period of 240 seconds. Portrayed on the Phase-plane, the "Orbit"
mode oscillation is a gentle drift from -3 degrees relative pitch angle to +3
degrees relative pitch (-S_ to -Sl ° degrees, referenced to true horizontal).
This drift lasts for approximately one-half-period of two minutes. When the
error becomes +S degrees, a low torque pulse causes the angular rate to reverse
from _2 to _ deg/sec., where upon the second half-period is spent drifting
_" -_|r'%r"lh .m. - -
PAGE
REPORT ....
MODEL
4-18!
SEDR 104
133
Mc'DONNELLL ' .,_J,_.U IS_, MISSOURI
DATE , 1NOVEMBER1961
REVISED
REVISED , ,
,SCS AMP- CAL
O_TPUT RELAYS L
i _
FIGUI2E 4-9 ASCS
//
//
//
NOTESONLY THE YAW MICRO SWI"T'CH_
AND ASSOCIATED CIRCUITRY'
A_E _i40WN, _0LL AN{:) PITCHARE _I_ILA_
[E)_V_/ITCHE_ SktO'cc'N IN %"riCKNEUTRAL POSiTiON LOWTORQUE 5WITCHES CLO_EFIRST
[_ RELAYS ,DE- ENERGIZED I kl
P_W F_OC_l"riON
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CHAMBER GYP.')
FLY-BY- WII2E CONTI_'OL PH,HS-_o_
OAT_ I NoVember I C_l
REVISED.
ST. L,OUIS 3. MISSOURI RE:PORT_ ,,,_--'_[ 104"
Moo=, 133elMmem_,_, ,,,
slowly through zero to -3 degrees error.
As another example of ASCS torque-switching, Figure 4-10 shows the
"Retrograde-Hold" torque logic phase-plane diagram. In this case high-torque
nozzles are utilized instead of the low-torque nozzles which were adequate to
control the gentle orbit oscillation. A series of attitude gyro repeater sector
switches and rate-gyro pickoff sector switches are used to define step-like
boundaries within the phase-plane. A typical contour is shown to illustrate the
motion resulting from a large disturbance torque while in this mode. When the
capsule motion results in a pitch rate value above the right-hand stair step_
high negative torque is applied until the capsule attains a negative rate and
rotates into the "no-torque" region. The inverse occurs if the retro-rocket
thrust eccentricity or other disturbances force the capsule into a situation
calling for positive thrust. The net effect of the torque-switching logic shown
is to maintain rapid and reliable control during the important operation of
retrograde firing.
Other modes of operation requiring torque switching logic are "Orientation"
and "Rate Damper". During orientation mode both high and low torquing is uti-
lized to rotate the capsule to new preset attitudes. Both high and low torque
is also applied during rate damper mode but only rate gyro signals are needed
as a basis for switching logic. In this case 3 torque switching boundaries are
horizontal lines on the phase-plane.
_-Ii. SYSTEM UNITS
_-12. Amplifier Calibrator
The Amplifier Calibrator unit can be "functionally" divided into four
sections. These functional sections are slaving, repeating, mode switching
and torque switching.
PAG E .
REPORT
MODEl.
4-20
SEDR 104
133
/oMCDONNELL___d# _,_
ST.LOUiS, MISSOURI
¢" ¢"__.i_ I_-_,r..L.,-,-,__L..,--
DATE ,
REVISED
REVISED
1 NOVEMBER 1961
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FIGUI_E 4-10 TO_'QUE L0454C PN/_SE PLANE
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4-13. Attitude Gyro Slaving
This section contains amplifiers anm summing netvorks which accept roll
and pitch information from the Horizon Scanners and generate currents to torquers
in the attitude gyros. Thus, upon command frcm an external timing device, the
Gyros' Roll, Pitch and Yaw gimbals are aligned with corresponding directions in,
or perpendicular to the orbit plane. (Ref. Para. 4-10. )
_-l_. Repeater Section
The repeater section •is a group of servo-mechanlsms (four in present design,
including two for pitch angle repeating). Attitude gyro outputs, which are
received at the calibrator in proportional or "analog" form, are amplified and
used to drive shafts which serve as roll, pitch and yaw signal sources for both
internal (torque switching) and external (display and telemetry) purposes. The
on-off reaction control of the Mercury Capsule makes it desirable to use con-
ductlve sectors on the shafts of three of the repeaters. The sectors serve as
attltude-level references for torque switching.
4-15. Mode Switching Section
This section of the Calibrator establishes the proper attitude angle bias,
torque switching status, and interlock signals corresponding to the ASCS mode
commanded by external devices.
NOTE
The sum of all such external devices is, for ASCS design
purposes a 'Smaster sequencer" which coordinates all automatic
functions.
The mode-switching section uses compact, solid-state switching circuits. Al-
though these circuits contain many transistors, diodes, and other electrical
components, they are of a class that i8 not critically dependent upon reference
voltage or temperature levels.
_'/_ T_'T!DE;'T. :_
Mc,DONNELL zREPORT _DR lOLl. REVISED
ST. LOUIS, MISSOURIMODEL. !33 * ' _ REVISED
4-16. Torque Switching Section
The torque switching section contains transistor and diode circuits similar
to those in the mode-switchlng section. Torque switching circuits receive the
step-functlon outputs of the attitude gyro repeaters, plus the outputs of the
rate gyros. The latter (rate) signals come from sector switches replacing the
usual proportional rate gyro pickoffs. Using these step-wlse indications of
attitude and rate conditions, along with the mode switching section output defin-
ing the current phase of the mission, "decisions" are made which result in
energizing of the appropriate Reaction Control valves.
4-17. Accelerc_eter Switch
The acceleration switch is a hermetically sealed instrument. The basic
mechanism consists of a centrally located mass supported by a cantilever spring.
The mass is damped by the viscous shear action of the fluid which fills the case.
Switch actuation is caused by the displacement of the mass element. An Acceler-
ation force of .05g, in the axis normal to and in the direction away from the
base, is required to close the circuit. M._chanical stops are provided to re-
strain the mechanism and to protect against damage when subjected to excessive
acceleration.
_-18. Attitude G_ros
The function of the attitude gyros (vertical and directional) is to determinq
attitude augles between a set of fixed axes in the moving capsule and the refer-
ence ax_s which are fixed in the orbital plane but which are moving with the
local vertical. Both attitude gyros are "free" gyroscopes with slaving capabil-
ity. A means is incorporated for caging and for obtaining electrical signal
(synchros) outputs which define the attitude of the gyros with respect to two
mutually perpendicular axes. The attitude gyros possess unrestricted mechanical
freedom in the outer axis and + 83° (minimum) of mechanical freedom in the inner
REVISED. ST. LOUIS 3, MISSOURI REPORT. SF.LDR lO_
REVISED. MODEL lB3
!
_
axis. It is noted that the degree of gyro freedom does not necessarily reflect
the attitudes permissible by manually steering the capsule in orbit. Due to
limitations in the Horizon Scanner system and the repeater section of the
Amplifier Calibrator, manual control of the capsule should be limited to + 30°
in all axes. However, barring equipment malfunction, exceeding these limits
will not prejudice the success of a mission. If these limits are exceeded, it
is recommended that the gyro switch be placed in the FREE position. Input
power requirements are 115 volt 400 cps single phase (gyro motor), and 26 volt,
400 cps (synchro and torque motor).
4-19. Attitude and Rate Indicator
The Attitude and Rate Indicator is mounted on the upper portion of the Main
Instrument panel. The indicator provides visual indications of Capsule Rate and
Attitude in the Yaw, Pitch and Roll planes. The attitude indicators are driven
by the attitude gyro synchro outputs (through the Amplifier Calibrator). The
attitude indicators are calibrated to indicate capsule attitude within a range
of + 180 ° except for Yaw which shall indicate 0°, 80°, and 270 ° In a clockwise
direction. The rate portion of the indicator is driven by the miniature rate
transducers (See Para. 4-26) which also serve as sensing elements for the Rate
Stabilization Control System. The range of rate indication is O to + 6°/sec.
for all three indicators. The roll rate indicator has the additional capability
of being externally switched to a range of 0 to 15°/sec. in order to monitor
re-entry roll rate.
0
4-20. Rate Gyros
The rate gyros perform electrical circuit switching functions at specific
rates of angular velocity about an axis perpendicular to the base of each unit,
referred to as the "input axis". Rate gyros are used in the pitch, roll and
yaw m_es, respectively. Each rate gyro consists of a high speed rotor, mounted
u. 4 MC,DONNEL L . REPORT _DR lO_ REVISED
"_;ST. '_OUIs _ MI$SOURL,.
MODEL- l_ _u-',..=. ....... REVISED
in a gimbal ring, in such a manner that it is free to process about one axis
only (the output axis) which is perpendicular to the spin axis of the rotor.
The output signals are generated by the motion of wipers, attached to the gimbal
ring, moving across the contacts of sector switches. Input power requirements
are met by ll5 volts, _OO cps.
4-P_l. RATE STABILIZATION CONTROL SYSTEM
_-22. SYSTEM DESCRIPrION
The Rate Stabilization Control System provides an excellent alternate means
of capsule attitude control in the event of failure in the Automatic Stabiliza-
tion Control System. It has been shown by flight simulation studies of the manual
control problem that the Astronaut should, by utilizing the Rate Stabilization
Control System, be able to approximate retrograde attitude error performance of
the Automatic Stabilization Control System. In addition, the Rate Stabilization
Control System provides a completely redundant rate-damper and programmed roll
rate during re-entry.
The Rate Stabilization Control System consists of three miniature rate
gyros, three (signal pickup) potentiometers, three channels of electronics (rate
damper) contained in a 300 cubic inch box, one switch, and six solenoid control
valves, which utilize the manual reaction control fuel and thrust chambers.
Figure 4-1 shows the location of the major components within the capsule. Total
weight of the Rate Stabilization Control System is approximately 25 pounds.
Power requirements for the Rate Stabilization Control System are met by 24
volt d-c and ll5 volt, 400 cycle a-c. Power is connected directly to the rate
damper box through the AUTO/RATE COMD switch mounted on the left console. See
Figure 4-12.
REVISED. ST. LOUIS 3, MISSOURI REPORT SEDR 10_
REVISED. .... i MOOE.. _.33
4-23. SYST_q OPERATION
Figure 4-ii is a functional block d/agram of the Rate Stabilization Control
System. A typical channel is shown. In general, the Rate Stabilization Control
System provides the Astronaut with a redundant rate damping and "rate-stlck"
steering feature. The outputs of the rate transducers and the control stick
potentlomenters are combined in the summation networks. The summation net-
works compare the capsule,s angular rate errors and, if errors exceed preset
dea_-zomes, 2 de_sec in Roll and 3 de_sec in Pitch and Yaw, the appropriate
off-on corrective torque is commanded by energlzimg solenoid control valves in
the Mamua_ Reaction Control System. With the control stick at zero deflection
(Rate Stabilization Control System operational) an autcmatic three axis rate-
damper is achieved, including an automatic 7.7 + 2 de_sec constant roll rate
when re-entry is sensed. By manipulation of the control stick, steady-state
angular rate____sother than zero (approximately proportional to stick deflection)
may be attained if desired. This is in contrast to the proportional accelera-
tion (Torque) response which remains as an alternative in event of malfunction
in the Rate Stabilization Control System.
4-24. SYSTEM UNITS
4-25. Rate Damper Box
The rate damper box provides three channels of transistorized electronic
unit caaprlslng the rate-damper portion of the Rate Stabilization Control
System. Each channel contains a summation network, preamplifier and demodulator,
and two trigger circuits. The 5½ x 6 x 9 in. box weighs approximately seven
pounds and is mounted below an_ immediately aft of the control stick. The rate-
damper box mounts two AN type electrical connectors, one for GSE and one for
capsule lnterconnectlon. The dead band adjustments which are provided are a
2 de_sec to 4 deg/see in pitch and yaw axis and a 1.5 de_sec to 3 de_sec in
_,,,, ^ ,k, I--I 11_ Plk I"!" 1 A !
PAGE.
REPORT
MODEL
ST.LOUIS, MISSOURI .REVISED
REVISED
SEDR 104
133 'w.,_./l r ,L./'-:--;T_IInL
. _i_ _II l
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CONTROL STICK POTENTIOMETER I
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REVISED ST. LOUIS 3, MISSOURI REPORT _EDR 104
R_m=o. _ MOO_ .... 133
the roll axis. Signals from the rate transducers and stick potentic_eters are
sent to the rate damper box where the two signals are summed together forming
an error signal. The error signal is sent to a two stage voltage amplifier
followed by a double ring diode demodulator. In accordance with the magnitude
and phase of the error signal the demodulator will select one of two transistor
operated relays which applies a 24 volt d-c output signal to appropriate sole-
noid control valve such that corrective torque is applied.
4-26. Miniature Rate Transducers
Each of the miniature rate transducers consists of a gyroscope, an ampli-
fier, and a demodulator. These components function together to produce an a-c
output signal proportional to input rate of change of attitude. All three rate
transducers are identical except for gyro orientation in the transducer base.
A special indexing feature prevents installation in the wrong location. Input
power utilized by the rate transducers is ll5 volts, 400 cycle a-c.
4-27. Control Stick Potentiometers
The three identical 1000 ohm 120 turn potentiometers are connected to the
manual control stick linkage in such a manner that output signals are produced
proportional to stick deflection in the yaw, roll, and pitch planes. Active
sector is equal to 40 degrees with a minimum of lO degrees of over-travel
without electrical discontinulty.
4-28. REACTION COB"I_OL SYSTEM
4-29. SYSTEM _ESCRIPTION
The Reaction Control System is used for capsule yaw, pitch and roll control.J
The system is a pressure fed monopropellant, catalyst bed design. The right
angle thrust chambers obtain thrust by decomposition of 90% hydrogen peroxide
(H202). The system is divided into two individual systems; one for automatic
control (ASCS), and one for manual control (control stick and RSCS). The auto-
P_E _'_ Mc'DONNELL,_ __ D_T_REPORT SEDR 104 ST.LOUIS, MISSOURI REVISED
MODEL 133 REVISED
1 NOVEMBER 1961
II
I
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___GYRO CAGE COMD.
V LOGIC DC
P AFT G'Yr_O PWR
J _.OG|C AC
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H_' AC PWR
D DC PWR
_ Sc ,,_ CAL
FI GURE4- I_, .POWER DISTRI _LITION DIAGRAMPM 4,5-159
,,2J
C C t'; -; F_.,.F_N T i,kL-
DATE i November 1961
REVISED
REVISED
Mc'DONNELL . ST. LOUIS 3, MISSOURI
PAeE_ 4-29
REPORT, SED_ I0_
MODEL 133
matic system consists of a pressurization system, an "electrically" controlled
solenoid fuel distribution system, and twelve thrust chambers. The manual
system is similar to the automatic system except that it consists of only six
thrust chambers. The manual system also utilizes proportional "manually" con-
trolled fuel distribution valves in addition to the electrically operated
solenoid control valves. Figure 4-13 shows the location of all system compon-
ents within the capsule.
4-30. SYSTEM OPERATION
The following paragraphs, 4-31 through 4-35, briefly describe the operation
of the automatic and manual systems. Figure 4'18 and 4-19 should be followed
closely in conjunction with the following text.
4-31. Automatic System
The automatic system consists of twelve hydrogen peroxide monopropellant
thrust chambers of fixed thrust levels and their associated valves, lines, H202
tank, pressure regulator and pressurization bottle. (See Figure 4-14.) The
automatic system can be essentially divided into three sections; pressurization
and fuel supply, distribution, and propulsion units. The fuel supply is un-
stable hydrogen peroxide (H202) contained inside a flexible bladder which in
turn is contained in a half toroidal tank. The flexible bladder has a fuel
capacity of approximately 32 pounds of liquid H202. Helium, under pressure,
surrounds the bladder containing the H202 and acts as the pressurization agent.
The Spherical helium tank, pre-servlced to 2250 psi, has a capacity of 265 cubic
inches.
The following sequence of events occurs in producing a thrust output.
Assume the bladder is serviced with H202 and the helium sphere pressurized to
2250 psi. Upon opening the helium regulator manual shutoff valve, helium is
allowed to pass through the filter, regulator, checkvalve, and finally surrounds
_Z;;;-;D--;gY;AL
CI4EC_
HELIUMRELIEF VALVES
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PRESSURE _WI
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FlU-DRAIN 5HUT-OF_" VALVE
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THRUST CHAt_BERS
-PRES;URE TRAN,_I)_W:ER
-HELIUM REGULATORS
ELIUM MANUAL SHUT OFF VALVES
TRAN_DUCEI?
MANUAL VENT VALVE_
-HELIUM (Hz)TANK 0,1ANUAL')
(Hz)TANK ('AUTO)
MANUAL PUSH-PULLOFF VALVE
MANUAL PUSI4-PULLSHUT OFF VALVE
C_ECK V_LVE_
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///I
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// /
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SOLE|NIOIDVALVES
)ITCHTHRUST CHAMBER(o TO24 LBS.)
THRUST CHAMBER(_ANO2# LBS.)
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THRUSTCNAMBER .(i ANP ?4.LB_{,)
YA_N SOLENOIDi_ALVE_
//
VENT VALVES
YAW -_0LENO|D
YAW THRUST(!A_P 24 LBS,)
YAW THRUST.{0 TO Zar LBS)
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PITCHTHRUSTCHAMBERSf AND 24 LBS)
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l November Mc,DONNELLREVISED ST. LOUIS 3, MISSOURI REPORT _ 10t
REVISEO. _ MO0_ 133
the bladder pressurizing it to 480 psi. The helium pressure forces the H202
out of the bladder through the perforated transfer tube and into the down-stream
lines. By opening the manual push-pull shutoff valves the H202 becomes available
at the "electrically" operated solenoid shutoff valves. Upon receiving a 24V
d-c signal from the ASCS or fly-by-wlre control system, the appropriate solenoid
valve opens. H202 enters the solenoid valve through an integral 165 micron
filtration screen and passes into the corresponding thrust chamber where it is
decomposed and produces the desired thrust. See Figure _-20.
System components not directly associated with the preceding explanation
of a thrust output are explained as follows. The helium pressure transducer
provides a means of monitoring (by proper calibration) the percentage of H202
present in the bladder. The perforated tube in the propellant tank (external
of the bladder) is used to prevent the possibility of trapping helium pressure
while servicing the H202 bladder.
_-32. _snual System
The manual system (See Figures 4-15 and _-16) consists of six thrust
chambers of the same configuration as those in the automatic system with pro-
portional thrust output added. The fuel flow in the manual system may be con-
trolled in either of two ways: (1) by manually controlling the proportional
control valves, or (2) by electrical solenoid control valves. A two position
selector valve is provided such that the method of control may be selected.
See Figure 4-19. The manual control valves have a dead band of + 1/16 of an
inch from theoretical neutral and a total stroke of 3/8 of an inch from
theoretical neutral for each thrust chamber. The throttle valve arm assemblies
or bellcranks which rotate these proportional control valves are designed to
shear at less than full Astronaut effort on the manual control system. See
Figure _-29. In the event a proportional control valve should Jam and immobilize
-n
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SOLENOID VALVES
RUST CHAMBERS 0 AND i_ LBS_
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HELIUM REGULATOR
I_AANUAL SHUT-OFF VALVE
HELIUM MANUAL VENT VALVE
,/- HELIUM FILTER
MANUAL PUSH- PULL
SHUT OFF VALVE
MANUAL PUSH-PULL
SHUT OFF VALVE
E_ "VALV E_
SOLENOID
VALVES
i-IELiUl_ RLL AND
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YAW 50LENOID
VALVE 5
THRUST CHAMBER (
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PAGE 4-34
SEDR 104.REPORT ......
MODEL 133
M_'DONNF__f__ST.LOUIS, MISSOURI
//
/'
\\
//
DATE
REVISED
REVISED
1 NOVEMBER 1961
\
//
///
/
FIGURE 4-16 RC5 ,CONTP, OL L%NKAGE PM45-SZ C
!,
(
D,TE _NO_._., MCDONNELL_ __ P,_E.•REVISED ST.LOUIS, MISSOURI REPORT .....
REVISED . MODEL
"1_-- _AFET_'
4-35
SEDRI04
133
\
\
i \ \\\
PITONCONTROLLINKA6F.
YAW
FIGUI_E 4-17 TI-I_EE AXIS HAND CONTROLLEI_
__.all_ _li. IL I _ I II1_ IIIm Ik I i •
REPORT SEDR 104 ST. LOUIS, MISSOURI REVISED
MODEL_ 133 -,_,,.,""7_';DETUTIAL ......... REVISED
DATE I November 1961
the manual control system in one axis, increased Astronaut effort would break the
shear pin in the bellcrank and free the system. The manual control system could
then be utilized for automatic control system fly-by-wire switch operation or
manual control system "rate stick" potentiometer deflection. See Figures 4-9 and
4-11. The remainder of the manual control system is similar to the automatic
system except for fuel capacity, which is 23,_ lbs. of H202 for the manual system.
4-33. SYSTEM UNITS
Due to the simple nature of the system components, a discussion of each is
considered unnecessary. However, two items (thrust chambers and propellant fuel)
do warrant brief explanations.
4-34. Propellant Fue ! (H202)
Hydrogen peroxide is a clear, colorless liquid soluble in all proportions in
water and most substances which are miscible with water. Hydrogen peroxide when
catalytically decomposed releases water vapor, oxygen gas, and heat. H202 de-
composition when properly contained and controlled is capable of producing usable
thrust. One pound of H202 solution (90%) when properly decomposed will produce
approximately 60 cubic feet of gas. Hydrogen peroxide (90%) freezes at ll.3°F,
and boils at 286°F.
_-35. Thrust Chamber
The thrust chamber assemblies (See Figure _-20) consist of a stainless steel
chamber that contains a metering orifice, a distribution disc followed by a
catalyst bed and then a nozzle. The catalyst bed contains a stack of removable
nickel screen wafers. The screen gauge resembles common household screen. The
screen is covered with an electrolytically deposited coating of 99_ silver and
1% gold (called drexite) that enhances the catalytic properties of the nickel.
The open area between the catalyst bed and the right angle nozzle forms a short
/f.,
|1 .
DATE I NOVEMBER 1961
REVISED ,, ,
REVISED
MCDONNELL_I_ __ PA_E.... _-_ST.LOUIS. MISSOURI REPORT. $EDR 104
.............. " MODEL 133
i ":'_';
PAGE
REPORT
MODEL
4-381s
SEDR 104
133
MCDONNELL_ ___/_Z DATE
ST.LOUIS, MISSOURI REVISED
____ REVISED
1 NOVEMBER 1961
°/I_o_
I
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SCI-.IEMA'r:I C . OlAGI:::2AM
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/
REVISED ST.LOUIS, MISSOURI REPORT SEDR 104
REVISED ..... ., . -v MODEL. 133
/
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DREXITE COATED NICKEL SC2EEN_- PO;ZOUS STAINLESS STEEL FLOW/-- DISTRIBUTION PLATE -- _ILTEI_
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DII,_EXITE COATED NICKEL SCREENS- /--_LO#OU_STS_I_UNLIE# _ #LTEAIEL /FILTERS /'-_"L'TEI_.
c_LWS_ cups _ / ,/ _ SCREEN
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24 LB. T_I_UST
FIGUI_E 4-20 12EACTION CONTIL:k3L TN_UST CHAMBERPP14_-giB
PAGE b,.-Jl.O
REPORT .q_._ 1 Oh
MODEL. !_3
Mc,DONNELL DATE 1 November l_lREVISED
ST. LOUIS, MISSOURI
...... _._, REVISED
plenum chamber to smooth out the flow prior to reaching the nozzle throat.
H202 enters the thrust chamber through a metering orifice upon actuation
of the solenoid valve. The stainless steel porous plate distributes the flow
and presents the catalyst bed with a uniform input. Upon entering the first
stage of the catalyst bed, a violent reaction takes place. Expanding gases
rush through the remainder of the catalyst bed resulting in a thrust output in
the right angle nozzle. The majority of the decomposition (and most violent)
takes place within the first two catalyst cups. Temperatures of approximately
1400°F can be expected in this area. The remainder of the catalyst cups are
to assure a complete decomposition process and to prevent any liquid form of
H202 from reaching the nozzle.
_-36. HORIZON SCANNER SYSTEM
4-37. SYSTEM DESCRIPTION
The Horizon Scanner System incorporates two identical scanning units. The
purpose of the Horizon Scanner system is to provide a roll and pitch reference
during the orbital phase of the normal mission. The scanners produce an output
signal that slaves the ASCS attitude gyros to the proper angles upon command
/
from an external programmer.
4-38. Basic Construction
Figure 4-21 is a photograph of a Horizon Scanner Unit. _ll major components
and "subassemblies are mounted from the large circular plate and include the
scanning prism assembly, prism drive system, infrared detector, electronics,
synchronous switches, electrical connector and cover. The circular plate is
flange mounted so that the scanning prism compartment projects into the space
outside of the vehicle. The elctronics system is completely transistorized and
the various functional sections are fabricated on separate printed circuit
_v _-q T le_ m-_
•,_vJ.'t.a. Am.IJL_I • .ILJ"Ik.JL'r _'*'_"
:M 'DONNE REVISED_ s'r. LOUIS 3, MISSOURI REPORT. _J_DR 104.
RSvisEo _--_ MOm_ 133
:.C._ /
boards. Three of these printed circuit boards are enclosed in the shielded
housing fastened to the circular plate. The remaining boards are fastened to
the four posts mounted on the circular plate. For rapid servicing the four
posts with attached boards can be replaced as a single unit, or individual
boards can be replaced as required.
4-39. S_ecial Features
The Horizon Scanner has a number of special features. It is compact in
size (6 5/32" long x 5 7/8" diameter over-all) and •light in weight (3.02 lbs).
The scanner is equipped with a centrifugally - activated shutter. The shutter
prevents solar radiation from dwelling upon the detector and resulting in
probable damage during those periods when the scanning prism is not rotating.
Another feature is a special circuit which can be used to disconnect the error
signals from the vehicle reaction devices during those periods when the presence
of the sun in the scan path or the loss of horizon would result in erroneous
error signals. The final feature of significance is that only a single power
source providing llO volts, 400 cycles, 3.2 va is required to operate the entire
system. The highly regulated power supply in the system eliminates the need
for the bulky batteries usually required to bias the infrared detector.
4-40. SYSTEM OPERATION
OPeration of the Horizon Scanners depend s upon infrared radiation received
from the earth as compared to the essentially zero radiation from space. These
differences in radiation levels provide a sharp radiation discontinuity at the
horizon. The Scanner system uses this discontinuity for both day and night
vertical reference sensing. When the capsule is oriented so that the earth is
present in its scanning path, there will in general be two points where the
scan intersects the earth's horizon (See Figure 4-22). The scanner detects
the thermal discontinuity, or change in radiation level, between the earth and
i
............. |Iv_._.s_a a/md,b.d_ql JL Adr-ka_
PAGE 4-42
REPORT SEDR 104
MODEL 133
Mc'DONNEL_LH'_ __ST.LOUIS, MISSOURI
DATE .... I NOVEMBER 1961
REVISED
REVISED
FIGURE 4-211 WIDE ANGLE HORIZON SCANNER
.- : , .,
PM4_-___8 A
'i
'_vi',illl =idliil I I_
DATE 1 NOVEMBER 1961
REVISED
REVISED
Mc,DONNEL_LH._ __ PA_EST.LOUIS, MISSOURI REPORT
MODEL
4-43
SEDR 104
133
s-
.=
SCAN
\\ \
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//
/I
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FIG. 4-2.2 I-IOP,J ZON SENGOI:E... SC.$wN PATTIZP,,I_
I
PAGEv- 7-
REPORT b_DR 104 ST. LOUIS', MISSOURi REVISED
MODEL_ i_ __='--_ _ REVISED
space at the two horizon points. The Scanner then bisects the included angle
from itself to the Horizon points, compares the direction of the bisector with
that of a fixed reference in the capsule and generates linear error signals
proportional to the angle between the bisector and the fixed reference. As
previously stated, these error signals (roll and pitch) are used to slave the
ASCS attitude gyros.
Figure 4-23 shows a simple block diagram of the Horizon Scanner. The
following discussion entails a brief explanation of the functioning of each
block as related to the over-all operation.
4-41. Radiation Gradient At Horizon
There is a large difference in the radiation which the detector receives as
it scans across the boundary between space and the upper atmosphere (troposphere)
This change is approximately equal to that from black bodies at 0°K and 200°K.
respectively, and the radiance difference is approximately 0.003 watts/cm 2 -
steradian. The location of this gradient is sharply defined, and it is much
larger than any others that can be encountered during the scan cycle.
4-42. Correction For Reflected Solar Radiation
Sharp radiation gradients do exist because of reflected solar radiation.
Such gradients are found at cloud edges, topographical irregularities on the
earth's surface and the terminator line between night and day. These radiation
changes can be filtered out so that the horizon gradient is the only one that is
detected by the system. Selective filtering can be accomplished since most of
the reflected solar radiation falls in the spectral region between 0.2 and 2.0
microns, while the radiation emitted by the earth and troposphere is at wave-
lengths longer than 5 microns. The filtering is accomplished by a germanium
prism and field lens in front of the detector. As a filter, germanium sharply
cuts off all radiation at wavelengths shorter than 1.8 microns while transmitting
REVISED ST. LOUIS 3, MISSOURI REPORT _DR 10t
R_VlSED MO_EL 133
very uniformly radiation from 1.8 to 20 microns. The use of this filter remove_
over 90 percent of the reflected solar radiation. Signal clipping techniques
in the electronics remove any residual effects.
_-_B. Scanning and Radiation Detection
Details of the scanning prism assembly can be seen in Figure 4-21. The
infrared detector is fixed to the_center of the circular plate and its field of
view extends through the circular opening in the center of the scanning assemblyo
The detector field of view is 2° by 8° and the presence of the scanning prism
has the effect of deflecting it 55° "from the normal. Thus the apex angle of
the scanning cone is llO°. In operation 3 the drive system rotates the scanning
prism and the detector field scans the field of view through the conical pattem_
described previously. Different amounts of radiation strike the detector during
various portions of the scan cycle: and the amplitude of the detector output
changes accordingly. The detector output signal is processed by the electronics
system and the error signal produced is available at the electrical connector.
4-_. Synchronizing Generator
Closely associated with the prism drive system is the reference signal
generator. The output of this generator Is a square wave signal at a frequency
of 30 cycles per second. This signal is the fixed reference against which the
detector horizon signals are compared. The reference signal Is triggered by the
interaction between a magnetic pickup coil and a semi-clrcular steel vane. The
vane is imbedded in a slot cut into the surface of the scanning prism assembly
gear. A pickup is mounted so that the end of its magnetized core comes close
to the surface of the vane. As the scanning prism assembly turns, the ends of
the vane pass by the end of the magnetized pickup •coil core, generating the
reference pulse. A subsequent electronic network converts the pulse to a phase I
locked 30 cycle square wave. The use of this signal will be considered later[
-=::;:-;7.,:;;TIAL
PAGE _-4_
REPORT SEDR 10_-
MODEL., 133
in this section.
4-45. Sun Shutter
REVISEDST. LOUIS, MISSOURI
REVISED
The sun shutter consists of a pair of spring loaded metal slides which fit
into opposed transverse slots through the tube section of the scanning mirror
assembly. When the scanning mirror assembly is not rotating spring tension
pulls the two slides together and the detector field is obstructed. When the
scanning mirror is turning, the centrifugal force on the slides is sufficient
to open the shutter.
_-46. Infrared Arrangement
Infrared radiation frcm the field of view strikes the infrared detector and
produces the electrical signal which is processed by the electronics system. The
infrared detector is a thermistor bolometer with its active element immersed in
the germanium lens.
The active element is a rectangular flake of thermistor material and is con-
nected in a bridge circuit with a similar compensating flake which is shielded
from radiation. The two flakes are oppositely biased and their Junction is con-
nected to the input of the preamplifier which follows. By immersing the active
element in the rear surface of the germanium lens the over-all detectivity can
be increased by a factor of about 3-5 over an unimmersed detector having the
same field of view. The material in the thermistor flake has a high negative
_emperature coefficient of resistance. That is, when the temperature of the
material is raised, the flake resistance decreases. Since the surface of the
thermistor flake is blackened, it absorbs impinging radiation and its resistance
is decreased. When the shutter is closed, both flakes in the detector bridge
are at the same temperature. Since both flakes have the same linear character-
istics, their resistances are the same. Gradual variations in ambient temperatu_
change the resistances of both flakes by equal amounts and the voltage of their
l ember MCDONNELL REVISED ST. LOUIS 3, MISSOURI REPORT SEDR 104
REVISED ............... MODEL_ 13._
Junction remains the same. When the shutter opens, incoming radiation is
focused on the active element; the compensating element is shielded from out-
side radiation. The temperature of the active element is changed, and its
resistance becomes different from that of the compensating element. As a
result, there is a voltage change at the Junction of the two flakes and this
change is connected to the electronics system. As the scanning prism turns and
causes the detector field of view to cross the horizon, there is a sharp change
in the radiation level striking the detector. The result of the radiation
changes during a complete scan cycle is the generation of an approximate square-
wave signal at a frequency of 30 cycles per second.
Electronics system is physically arranged so that functionally related
parts are located close to each other. The electronics system is divided into
eight major circuits, located on individual printed circuit boards. In some
cases, the requirements of compact and economical construction have resulted
in two or three sub-circuits being located on one board. Thus, the functionally
related booster amplifier, signal centering circuit and phase inverter-limiter
are located on one board which the block diagram (Figure 4-23) shows as divided
by dotted lines. Although the power supply and reference generator circults
are not closely related in function they are both located on the same printed
circuit board.
The paragraphs that follow describe the functions of the major sections
and sub-sections of the electronics system. The description is made with
reference to the waveforms generated by system operation, and these are shown
in Figure 4-24. Functional description will be made at the level of the major
circuits and sub-clrcuits shown in the block diagram Figure 4-23.
_-47. Immersed Detector
The radiation falling upon the detector determines the waveshape of the
REPORT ISEDR lOLl • REVISED.S_!" LOUIS, MISSOURI
detector output signal which is to be processed by the electronics system. The
radiation characteristics are determined by the scanning cycle described previ-
ously, and are shown as the first _aveform, WF-1, of Figure 4-24. The waveform
shows that earth radiation is higher than space radiation and that there is an
abrupt shift from one level to the other as the detector scans across the horizon
The change in radiation requires 200 microseconds to take place because the
detector is of finite size, and this time is required for a complete shift of
radiation level across the entire surface of the detector. Thirty complete
cycles of radiation change take place in one second.
_F-2 shows the detector output signal which results from the radiation
changes taking place at the detector. This signal resembles the radiation signal
with the exception that the shift between the two levels takes a longer time.
The reason for this is that approximately 2 milliseconds is required for the
active detector flake to reach the half-level of its now stabilized output.
The detector output signal has an amplitude in the order of 2 millivolts.
4-48. Pre-Am_llfier and Booster Amplifier _ircuits
The Junction of the two thermistor flakes is direct-coupled to the input of
the pre-amplifier. The pre-amplifier has a voltage gain of 400 at 30 cycles
per second. Direct-coupllng is used between l_m-ampllfier stages to provide good
low-frequency response and to prevent phase shift. Negative feedback is used
_-Ithin the pre_Ifier to provide stable gain, and the RC coupling network in
the feedback loop provides a high-frequency boost to compensate for the long
detector time constant. WF-3 shows the effect of this boost. The rise time of
the waveform has been reduced to approximately 350 microseconds at the half-
level point. The booster amplifier provides an additional voltage gain of 5 at
30 cycles per second. WF-4 shows the output of the booster amplifier. The
peak-to-peak signal amplitude is in the order of 5 volts.
DATE
REVISED
REVISED
I NOVEMBER 1961 Mc'DONNEL LST, LOUIS , MISSOURI
PAGE 4-49
REPORT ..... SEDR 104
MODEL 133
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PAGE
REPORT
MODEL
4-50
SEDR 104
133
Mc'DONNELL ST.LOUIS, MISSOURI
-_Vl_lr il./r--I_l I i ....
DATE
REVISED
REVISED
1 NOVEMBER 1961
WF I-RADIATION ON DETECTOR
WF 2 - DETECTOR OUTPUT
5O%
WF 5- PREAMPUF!ER OUTPUT
WF4- BOOSTER AMPLIFIER OUTPUT ( 0" HORIZON)
V.
WF4A- BOOSTER AMPLIFIER OUTPUT (2,5" DROPPED HORIZON]
__#_ /-_A---4_-_/2-_v.
WF 5- SIGNAL CENTERING CIRCUIT OUTPUT
WF &- LIMITER AMPLIFIER OUTPUT
WF 7-PHASE INVERTED LIMITED OUTPUT ._IGNAL
50-150 _ _EC.
WF 8- IMPUT_ TO PHASE DETECTOR
_RCE VIBRATING ARH NO,7 AT_-SOURCE -_-
AT R 801 _V
AT R 80Z _OURCE
LOAD
WF 9-CHOPPER INPUT (AT K!O.I CONTACT)
I-- LOAD "--'F" 5CURCE--I
KIOI _--_20V
I-.-_0uRCE--I.-- LOAD----I
KID& I
WF I0- CHOPpEP, OUTPUTS (AT NO. 7 CONTACT)
I--LO4D---p SOURCEH -_sv
KI02
INTERNAL SIGNAL WAVEFORM5
o, NO HORIZON
B_ 15 VIX: T t__j_ I_'E 5"8V P/P
0
_. LOW ALTITUDE HdRIZON
I
I I I VOLTS
t i _ -0
_) INSUFFICIENT 'GAIN OR PARTIAL SIGNAL LOSS
J _ _ jIF VOLT_--i - 0I I
, t I i i
c_ NORMAL _IPM_IAL DROPPEDHORIZON
a. NORMALAPPROX. ,
36 _ SEC_
| '/I-ayocl,.g_l tjJITTER MUST aE L-_S
..... •_--__J/THAN 5 U 6EC WITH
,i o RESPECT TO LEADING EDGE
_. MOTOR ._TOPPED(LOSS OF M.,,E-.Nr'T_CF'_CKUPREFERENCE PUL.S,E)
C, PARTIAL FAILURE OF SWITCHING TRANSISTOR
I) DE_c_IVE r_,_ _,,._ . 14. v NORMAL
0
_') DEFECTIVE I I-'' J-- I + I4 V NORMAL
Q 305 rtoy==R= 3vi:x-., __n',_;
J%P-L__J-L-O
] I I,I-7, ,F-l, IF "_°I " I I I U VOLT.5
_, L L _.LJ L___, _0 _. TYPICAL LOCK- OUT SIGNAL
i raT'pM_,-'I I I '>
'_" SUNIN SCaNII)_N AT I'-- IJ1 ir_____[]_.iI'-INCREASED VIDI"H CAUE'ED BY "SUN PUL'_Ei-'--÷?'0 j I LOCK-0UT T0 GROUND
I I vo.TsMORIZ.0N _
__----_ ,_------d, -O
NORMAL H0mZON D_SCONTINUrFY _') li II cI I . ' I U LJ L_ K- OUT TO +14 V2.)BUN IN _ -4,,tI-,--.$UN PUI.._E , _
_J _i_ _L----J _00LTG e.,. IMPERFECT _YNCHRONISM
O T OTATLIMITED SIGNAL OUTPUT AT P-IOtA
FIG-,4-24. HORIZON SCANNER WAVE FORM._ PM_-Z20 A
:i':U
, (:;
l Ii_ .m- - --
CC,;;";,..,_,,, , ,,-,,--
_ IIL/
OAVE I November 1961
REVISED
REVlSL=ID
ST. LOUIS 3, MI_URI RmPORT _DR i0_
Moo,,, 133
4-49. Signal Centerin_ Circuit
In the signal processing considered previously there has been no particular
interest in the voltage level of the average signal. However, this average
signal level is important in system operation. The reason for this is that the
error signal must be determined only by the phase angle between the horizon
and the fixed reference in the vehicle, and amplitude variations in the signal
should have no effect. Amplitude variations will take place because of changes
in earth temperature at different parts of the trajectory or orbit. When these
amplitude variations are combined with the rise-time characteristic of the
detector there is a difference in phase between different portions of the lead-
Lug edge and the fixed reference signal. Error signals would also be affected
by amplitude changes due to changes in amplifier gain and supply voltage.
Limiting can be used to eliminate the amplitude variations but the limiting
slice must be taken at a point of minimum phase variation. These variations
are greatest at the peaks of the wave and least at the center. Using an RC
circuit to couple the signal to the llmiter would balance equal areas of the
signal waveform above and below ground. Changes in the angle of horizon de-
pression would cause a shift in the d-c level of the signal. Hence, a signal
centering circuit is employed before the limlters to assure that the same center
slice is sampled for phase shift under all conditions. The signal centering
circuit consists of two diodes connected back-to-back as d-c restorers. The
diodes conduct on opposite peaks and thus permit the associated capacitors
to charge up to opposite peak values of the signal. The two levels are then
summed in a resistive divided network, and half the sum is sampled by tapping
the divider at its midpoint. An emitter follower couples this signal which is
shown in WF-5 to the llmiter circuit.
REPORT SEDR lOLl • St. LOUIS, MISSOURI REVISED
MODEL. 133 ............. _ _ _ • REVISED
DATE 1 November 1961
4-50. Limiting and Phase S_litting Circuits
The signal next enters the first of a pair of cascaded feedback amplifiers,
each of which acts as a limiter and phase inverter. The amplifier consists of
a grounded emitter stage which performs the phase inversion s and an emitter
follower. The feedback ratio is about 50:1 and the over-all gain of the section
is about 30. The output swing is lO volts each side of the fixed lO-volt level.
A low output impedance is maintained during the time when the emitter follower
is in cutoff by feeding the signal from the collector of the grounded emitter
stage directly through a shunt diode. The first section of the feedback
amplifier is fed by the signal centering circuit. Its output signal is the
"limited signal output" and sho_u in WF-6. Part of this signal is fed to
another limited amplifier substantially the same as the first, where it under-
goes a second inversion to become a mirror image of the output of the first
section. The output of the limlter section is thus the dual signal shown in
_F-8. While either signal carries the signal information s the presence of the
image signal will be found useful in cancelling out undesirable ripple components
in the rectified signal.
4-51. Phase Detector
A pair of symmetrical s limited signals enter the detector section (WF-8).
From these the detector derives a d-c signal which is proportional to the phase
difference between the reference pulse and the midpoint of the two horizon inter-
cepts. The phase sensitive rectifier consists of two SPDT polarized relayss or
choppers, driven in phase opposition by the reference signal. These are desig-
nated KlO1 and Kl02 (WFg). The use of two choppers provides the advantages of
full wave rectifications notably low ripple.
To understand the action of the synchronous rectifier, it is essential to
know the relative phasing of the drive and horizon signals. Since the two
._j/
_bul_ r l_l_ A A.ak L -
,,ooREVISED ST. LOUIS 3, MISSOURI REPORT _DR 10_
REV,SEO _ MOOEL 133
choppers are driven 180 ° out of phase, the arm of one connects its capacitor
to the source while the arm of the other connects its capacitor to the load.
Swltch-over takes place when the radial sector of the scanning beam crosses the
vertical reference mark of the sensor and switching back occurs 180 ° later.
Thus, when it is connected to the source each storage capacitor receives part
of the sky pulse and part of the earth pulse. The capacitor is charged positively
during the svltched in portion of the earth pulse and negatively during the
switched in portion of the sky pulse. If the sensor horizontal is parallel to
the horizon, eachcapacitor is negatively charged an amount equal to the posi-
tive charge. Therefore, the net charge is zero. If the sensor tilts with re=
spect to the horizon the amount that each capacitor charges positively is not
equal to the amount it charges negatively. The net charge is, therefore, no
longer zero. The net signal at the input of the d-c amplifier is thus positive
for a positive tilt of the sensor (cw as viewed from the sensor along the scan
axis) and negative for a tilt in the opposite direction. WF-9 and WF-IO indi-
cate respectively the voltage of each chopper and the uncombined output.
4-52. D-C Amplifier
The output of the phase detector is combined and filtered in an R-C net-
work at the input of the d-c amplifier. The signal at this point varies
approximately i00 mv per degree of tilt and the average level is - O.15 volts.
The amplifier input is at high impedance to maintain a low ripple factor. With
a gain of approximately three, the output of the amplifier is 286 mv per degree
tilt of the sensor, reversing polarity at zero tilt. Part of the output is
fed back to the emitter of the input stage. The balanced circuit configuration
minimizes the output drift vith temperature fluctuations.
4-53. SP-HL Detection Circuit
There are two conditions under vhich unwanted error signals are generated,
-"-C._'Y:ZZ_:T'AL
PAGE 4-54
REPORT SEDR 104
MODEL_ 133
Mc,DONNE REVISED
_-, -_'r._',-.'- -,_?_W_ REVISED,,
namely, when the sun appears in the scan and when the horizon is lost. Signals
produced under these conditions trigger a logic circuit which indicates by its
output that the sun is present or the horizon is lost (hence the designation
SP-HL circuit). This output can be used to disconnect the d-c error output from
the vehicle guidance system. The effect of sun presence is shown graphically in
the waveforms of Figure 4-24. The sun pulse introduces an unsymmetrical element
into the signal train 3 and the horizon information derived from it is likely to
be false. The presence of the sun's radiation is perceived at the detector.
The sun radiation is hundreds of times greater than that of the earth. The stars
and other bodies produce negligible signals. When a sun pulse occurs, the second
stage of the pre-amplifier puts out a negative pulse with a peak amplitude of
three to four volts. This pulse causes the Void Signal Circuit to produce an
output. When a horizon is present in a normal scan a signal of 5 of 6 volts from
the signal centering circuit suffices to keep the void circuit amplifier shut off.
The absence of the signal when the horizon is lost has the same effect as a sun
pulse -- it causes the void amplifier to conduct with a consequent output current
of 4 ma into a load of 2000 ohms or less.
4-54. Phase Reference Sisnal Circuit
A phase reference signal is produced by the scanning system whenever it
passes through its highest point with respect to the Sensor. The reference
signal is generated in the scanning system. It consists of two equally spaced
pulses, one positive and one negative, for each revolution of the scanning
system. These pulses trigger the bistable multivibrator. The two-level detector
section is in synchronism with the scan cycle. The output of the reference
generator under various operating conditions is shown in Figure 4-24.
4-55. Power Su_l 7
All the power required to operate the sensor is derived from the llO volt,
./
..../
/"
DATE ,1,,November 1961
REVISED,
REVISED.
ST. LOUIS 3, MISSOURI
PA_ L55
REPORV SEDR 104
MOI_ZL .....I_
400 cps llne by a built-in power supply. Input to the supply is through the
transformer. The primary of this transformer is tapped like an auto-transformer
to provide low voltage a-c to operate the scanning motor. The transformer
secondary output is full wave rectified to produce, -30, +30 and +16 volts d-c
with respect to ground. The +30 and -30 v01t outputs are fed to the Regulator.
The +16 supplies the reference generator, and void signal circuits. Part of
the transformer secondary voltage is rectified separately to produce unregulated
power for use in the reference generator and the void signal output current
amplifier.
_-56. Power Regulator and super Regulator Circuits
The Regulator circuits convert the outputs of the power supply into
regulated voltages for use in the sensor. Most of the voltages are regulated
by cascaded zener diodes which maintain a substantially constant voltage across
their terminals by an effect similar to break-down in a gas discharge regulator.
The regulator also contains a symmetrical arrangement of transistors connected
as emitter followers. Since the base potential of each transistor is fixed by
zener action, the output voltage is accordingly regulated with reasonably low
noise. This output is filtered and further regulated in the Super Regulator
circuit to provide the highly regulated voltage required by the detector and
pre-smpllfler. This voltage is extremely stable and its noise content is
essentially transistor noise. The zener diodes used in these circuits are 1/4
watt units which regulate within 5%.
4-57. TEST CONFIGURATION CAPSULES
4-5@. TEST CONFIGURATION NO. 8 CAPSULE
The Automatic Stabilization Control System for Capsule No. 8 is identical
to the Specification Compliance capsule with the exceptions noted in the follow-
Ing Paragraphs 4-59 through 4-62.
--1 i& iIiiiwllli _1 • i/-lIJImm
: f
i
"T!
CR]m
4_I
p.}Lrl
HELIUM LOW PRESSURERELIEF VALVES
HYDROGEMPEROXIDE zTA_K (AUTOMATICSYSTEM) /
/
MANUAL PITH- PULLSHUT OFF VALVE
NOTEDUMtt Y I_ISTALLATION
[_ HYOROGEN PEROXIDE{HzOz)(MANUAL SYSTEM) HELIUM FILL ANDVENT VALVES
HELIUM FILTERS t-b.Oz3£TTISOIq V4LVE
SQUIg ACTUATED
L
ii /
' _
__AW S_OLENO_D VALV
/ YAW THRUST CHAMBERSJ(I AND 24 LBS,')
PITCH SOLEHOIIPJVALVES
O !11 >
_ m
.4
m
_B _B 0Pl Pl ),
°°10 0
IzI;
-n
G_CX3rn.pI
rao-
c:
HELIUM LOW PRESSURE-
RELIEF VALVE
.J/
HYDROGEN PEROXIDE (HzOt)
TANK /
,' /
ROLL MANUAL FU_;H-PULI.,
SHUT OFF VALVE
RELIEF WLVE (Y,_)
FILL OR_.IN IDL_-..
RELIEF VALVE (_}
-ROLL SOLENOID VALVES
-THRUST CHAMBERS (I AND 6 LB '_')
TRANSDUCER
REGULATOR
MANUAL SHUT-OFF VALVE
IJELIUM MANUAL VENT VALVE
- HELIUM FILTER
/ / MANUAL PUSH- PULL
/I ' SHUT OFF VALVE
/
_0 / / MANUAL PUSH-PULL
SHUT OFF VALVE
-PITCH Hz_ RELIEF VALVE
' YAW H=Oz RELIEF VALVE
P* PITC P. SOLENOID
ii / VALVES
/ -1 CHAmB_ (..B.)@w SOLENOID
i VALVESi
_. _ / CHA_BE_"-" (ILB)
13 /
cn ; '
rn
.//_. e_,._BER
" " " _ " -" {24- LB.)
/ MELIUh_ FILL AND /
CHECK VALVE / VENT VALVE
HELIUM (H_ TANH --/ YAW SOLENOIVALVES
HEAT • _ARRIEP,THRUST CHAI_BER (ILB) CIr. PLACES)
' PITCH SOLENOID THQUST CHAMBER
"_ VALVES (I LB.)
:;0 O
g g m
E,C1
r- 11-t
m
u1
4_
4-58 McDONNE REPORT S_EDR 104 REVISED
ST. LOUIS, MISSOURI
MODEL 133 _ REVISED .......
4-59. Automatic Stabilization Control System
In general the discussion contained in Paragraphs 4-2 through 4-20 applies
to Capsule 8. Two exceptions are the addition of a yaw data converter to permit
telemetry monitoring of vertical roll and yaw torquer command signals, and the
removal of ASCS power after the antenna fairing has been separated.
4-60. Rate Stabilization Control System
Capsule No. 8 is not equipped with a Rate Stabilization Control System,
therefore Paragraphs 4-21 through 4-27 do not apply.
4-61. Reaction Control System
Capsule No. 8 is not equipped with a Manual RCS System with the exception
of the fuel tank which has been installed and connected to the Automatic RCS
System to provide additional fuel. All text contained in Paragraphs 4-28 through
4-35 pertaining to the automatic system applies with the exception of component
location. For automatic system component location and tubing configuration, see
Figure 4-25 and 4-26. Additional differences between Capsule No. 8 and
Specification Compliance include helium regulator output pressure and thrust
chamber composition, helium regulator output pressure for Capsule No. 8 is 460
psi. Capsule No. 8 thrust chambers do not contain the filter and screen shown
in Figure 4-20.
4-62. Horizon Scanner
Same as Specification Compliance. Refer to Paragraphs 4-37 through 4-56.
4-63. TEST CONFIGURATION CAPSULE NO. 9
The Automatic Stabilization Control System for Capsule No. 9 is the same
as the Specification Compliance capsule with the exceptions noted in the follow-
ing Paragraphs 4-64 through 4-67.
\
. j
r__
|
!_
r_
o
ROLL 5HLJT-OFF VALVE.('AUTO
ROLL RSC550LE_Ot DCONTROL
CVALVE5
TO PtTCH AI'_D YAWTHROTTLE VALVE5
(MAN)
I
_B,C¢_PITC_ AND YAW_OL_Om CONT_O_VALV rr-S CMAN)
EAAI4UAL RC_ HYDROGENPErOXiDE (H=O_ TANK
m Ii1< <
m mo o
==
0
.<
=;
;0 "0m >
II-t
m
_E ,-6o M_,DONNE_ __ D_TE,REPORT SEDR 104 ST.LOUIS. MISSOURI REVISED
MODEL 133 ="_..ti"lt =L,,,tI.,, "Ti = ,_,_ REVISED
I NOVEMBER 1961
AUTOMATIC RC_ HYDF_OCtEN PE.P.,O'Y,IDE. (,H'z.Oz,")T,kNK.
MANIJAL RC._SHYbP.OC1EN PEP.OXIDE (HzOz) TANK
FI_LIRE _.Z8 INTERCONNECTION OF AUTOMATIC- AND MANUAL P-,CS FUEL SUPPLY SCHEMATIC (CAPSULE 9)
,_.:j
i • _LI
• I Ii,,/L--i_II i I:PI r---:
r
/
OATE 1 November 1961
REVISED
REVISED
M DONNE ST, LOUIS 3, MISSOURI
PAGE 4-61
REPORT SEDR 104
4-64. Automatic Stabilization Control System
In general, Capsule No. 9 ASCS system is the same as Specification Complianc_
Capsule. Some slight differences do exist in the area of normal sequencing due
to the different missions involved. For a detailed description of ASCS normal
sequencing, refer to Paragraph 4-5 through 4-9.
Normal sequencing described in the foregoing referenced paragraph is ade-
quate for Capsule 9 with two exceptions. ASCS operation ceases at antenna fair-
ing separation when the ASCS power is removed. Also a yaw data converter has
been added to the system to permit telemetering the command signals generated by
the Automatic Stabilization Control System. Refer to Paragraph 13-145.
4-65. Rate Stabilization Control System
Same as the Specification Compliance Capsule. Refer to Paragraphs 4-21
through 4-27.
4-66. Reaction Control System
The Reaction Control System on Capsule No. 9 is the same as the specifica-
tion compliance capsule with two exceptions. The control stick is not installed
and the Manual Reaction Control System hydrogen peroxide fuel tank is connected
to the Automatic Reaction Control System to provide an additional fuel supply for
the automatic control system. See Figures 4-27 and 4-28. Refer to paragraphs
4-28 through 4-35.
4-67. Horizon Scanner System
Same as specification compliance capsule. Refer to Paragraphs 4-36 through
4-56.
4-68. TEST CONFIGURATION CAPSULE NO. 13
The Automatic Stabilization Control Systems for Capsule 13 are identical to
those discussed in Paragraphs 4-2 through 4-56 which apply to the Specification
Compliance Capsule.
_A_ ,-',_r'h|_', A,
PAGE, 4-62
REPORT _DR lO_
MODEL. 1_3
Mc'DONNE ST. LOUIS, MISSOURI
DATE i November 1961
REVISED ,
REVISED
4-69.
Same as Specification Compliance.
4-70. Reaction Control System
Same as Specification Compliance.
4-71. Horizon Scanner System
Same as Specification Compliance.
Automatic Stabilization Control System
Refer to Paragraphs 4-2 through 4-20.
Refer to Paragraphs 4-28 through 4-35.
Refer to Paragraphs 4-37 through 4-56.
4-72. TEST CONFIGURATION CAPSULES NO. IOAND 16
The Automatic Stabilization Control System in Capsules No. i0 and No. 16 is
the same as the Specification Compliance Capsule. Refer to Paragraphs 4-1 throu_
4-56.
4-73. Automatic Stabilization Control System
Same as Specification Compliance. Refer to Paragraphs 4-2 through 4-20.
4-74. Reaction Control System
Same as Specification Compliance.
4-75. Horizon Scanner System
Same as Specification Compliance.
Refer to Paragraphs 4-28 through 4-35.
Refer to Paragraphs 4-37 through 4-56.
,umMp_4m=-_
DATE , I NOVEMBER 1961
REVISED
REVISED
M DONN ST.LOUIS, MISSOURI
PAGE..
REPORT
MODEL
4-63
$EDR 104
133
THROTTLE VALVE
FIGURE 4-Z9 MANUAL RC5 THROTTLE VALVE. AND BF.LLCPh_NK PM45- 7.70
_.A:; .......vv I IIN/r--I1 I I/_ t-
F _
mSEQUENCE SYSTEM, LAUNCHROUGH RETROGRADE
S_CTION v
OR ABORT
TITLE
TABLE OF CONTENTS
PAGE_.___.
NORMAL MISSION SEQUENCE ............... 5-3
ESCAPE SYSTEM ..................................... 5-11
TEST CONFIGURATION CAPSULES ........ 5-18
C_.;&:----m -.Ill'liB • I Ilmlliill
(_) fIRE BOOSTERAND 3USTAdNER ENGINE,_.
• - (_ (_) (_ BOOSTEREN6tNE CUTS OFF AND _EPARATES UPON GROUNDCOMMAND.
1"1 _ (_ TWE_T_ _ELO_b_.kTTE.R E,OOSTER.CUTOFF, FIRE TOWER Rim SEPARATIONBOLTS.
C_/ \ _ AFTER SENSING TOWERRI_ SEPARATION.ESCAPEROCKETS
_.(-- -_ "-" FIRE, PARACHUTE L,kNOIN6 SYSTEM IS kRMEB
_- _ (_) CAPSULE ADAPTER 5EPARATIONBOLT5 FIREAFIER THRUS'I I)ECA_'TO .OZC_A,_OiSEt...TINGEDELAX
_> .
f'-_I (_) CA"UIY ROTATES 180 AND A'SL_ES "_'_-°OV"_IT ATTITU DE •
_LI (_ (_ WHEN 'I'_'PP"O_I'P"A_TT'Tt_DE'A_b kFTEP" _O _EC"T'IvIEDELKY' RETRO ROCKET'_ F'REE
V _ (_ SIXTY SECONDS AFTER RETP,.OROCKETS HAVE FIREDs RET_O ROC..KET P/_,CK_kR IS
Z _"_ JETTISONED CJkP._ULEA._S_),AE_AND _AINTA,IN_ RE-E_TP-.Y _,TTITUI)E
m_ \'_ (_ UPOI'_ DE_CENT TO ZI,O00 FEET,DROGUE CHUTE IS DEPLO'YEb .
,__ (_ AT I0,000 FEET, I)ROf_ttECHbTE,,AIITEHNA FA_P.._,_ IS JETTISOhEb t_kIN
_,. _ CHUTE DF.PLO'YEb
_ I_ UPQ_ IMPACT,Mk_N CB_TE IS bI_,CONNEC..TE|)PILOT &Hb KESERVL CHUTE EJECTED,
Q _ RECO_/ERY ._,tbS DEPLO_ED
i/_- --..
\..
RE'VISED
REVISED
ST. LOUIS 3, MISSOURI
pA(; .....:5-3
REPORT _ I0_
Mo_ 133
V. SS_ SYSTEM, _ THROUGH_ OR ABORT
5-2.
5-3.
NORMAL MISSION SEO_:E
LAUNCH T_R(XJGH ST._IIEr
_IPTIONi
The launch through sCs_l.n8 sequence establishes 'baLl.c references at time of
la_ch and then remains inactlve matll _. At sta_L_ the -_satle's
booster engine separates, resulting in the escape tower bolts be_m_ fired after
a twenty seccmds time delay. The escape rockets are fired i_ed£ate_7 after
tower bolt detonation and subsequently the la_iaE system becomes armed.
The sequence system is Initiated by tvo 28 V d-c slgnsls from the atsslle
vhlch occur st 2 inches a_er l£x'toff (See _ 5-2). This ls know: as time
zero reference and ener_zes a T_e Zero latching relay in the No. 3 Launch and
Orbit relay box located within the capsule. _.n Astronaut ocutro].led back-up
switch is provided in the event the 28 V sl_ls from the _ssLle do not reach
the cap_tle. These same sl_aols are also sent to the _ Altitude Sensor
and the Satellite Clock. The s_aal to the _ Altitude Sensor results in
e_l/shi_ the functlc_ of time lift-off versus the ti_e an abort _ay occur.
At &_tel_ 1,35 seconds _sslle sta_n_ will occur whereby the _eeh_n£cal
separation of the booster engine will cause the loss o_ capsule power to the
Booster E_lne Sel_ratlc_ Sensor rely. _awa_h this de-ener_zed rel_ power
be appl£ed to the Tower Jettlson 20 Second Time Delay rely. After the 20
second t_e delay_ power will be _plied to ener_L_e the Tower Sel_ar_lu_ Bolts
Power relay and the Tower Jettlson Warni_ Llsht 2 Second _ Delay relay.
The To_er Sel_ttun Bolts Power rely7 ls ar_ed by both the Main a_ Isolated
_-e s_u£b bus throu_ the _ ARM switch. _en ener_zed, the Isolated _lu_b
PAGE
REPORT
MODEL
5i- 4
SEDR 104
133
/
Mc,DONNELjL__ __ DATE-_-'. ST, LOUIS 'r MISSOURI REVISED
P P_w, !"_ =_I_=Ne_F_ REVISED
1 NOVEMBER 1961
5T_ql NGTWO. _- ETT.W_, L_RT
W_,RN. LT. I_E L/_"_ I
L_O S_aT,D.') I
f
leOOsTaa I I
E._GmE I ^ I
_EPA_'/_T%O_I-_ INTEP, LCX:.K _ FIREREL P<'¢ I I
_lN_ LIMIT P-OCKE_
%V_/ITC.,,_ FiRE _.E.Lk'-([ [ aOq"_'F''{l"
II
I
TW_ 5E_ I I TOWE_.i_OLTS -- _OLT5
H JETTISONRQCKE.T
_///////////////////_
///////////////////'/_
BOOSTERE.Nr=.SE.P.
NOTF_.%/_ USE IF ATLP,% FAIL% TO
PRODUCE. LIFT-OFF 51C.,,_L
/_ LO_ OF ?.B VOLT POWER
z_ IF ESCAPE P.OCKE%'5 FA_LTO F_P.E ANO T_R_T bE.CkW5
TO ,_0 q.
LIFT- OFF
AT L/_% -_B VOLT_ AT
TIME ZERO
_w_TC.H I --
H
4
%._-,T E.LLI-? E
TIME 7-E_,O
MA_MUMALTITUDE
FIGURE 5- ?_.LAUNCN THROUG_ 5"T/kQ_,NG(._LOCK DIA,GR/_'Ni_
%'-/_TEM
r
!4_t
\ -#M'45-? _,5
i_ii _
bus pover flres tvo of the flve squibs (2 boZts) and _ squib bus power flzes
three of the rite squibs (3 bolts)° As the three se_ented Tower Clamp
sepazstes_ the three Tover Claap ring 1t:tt svitches return to the norasl posi-
tim and sJ.lov Iso_ted and Msln squlb bus pover through their contacts. The
Isolated squlb bus pover enerElses both the Emergency Escape Rocket Fire rela_
and the Eaergency Jettison Rocket Flre re/_y, vhtle the )_Ln squib bus p(r_er
energizes both the Escape Rocket Fire relay snd the Jettlsc= Rocket Fire relaT.
£s the contacts of the Eaers_acy JettAs_ and Jettison Rocket Fire relays are
ecanected In paza_el, either relay vl_ fire both s_Llbs Of the _ettlsaa rocket
£rcm the two different power sources. The Emergency Escape and Escape Rocket
Fire relays are cc_ected :In the Identical same :anner and _ fire both squibs
of the escape rocket _ both pover sources. Power to energlze the Eaergency
Jettison and Jettison Rocket Flre felly is routed throv_h the Capsule Separation
1 Seecad Ttae Delay relay. _s the result of the 1 second delay the Tower Clap
sel_rates_ the Escape Rockets flre and separates the rover f_m the capsule
vlth the Jettison Rockets unfired. _en this ls s_ccmpllabed, two electrical
disconnects between the tower s_t cspsule are sepaz_ted and reaove power fr_
the three Tower Separaticm Sensor relays. _ the de-energized No. 1 Tower
Sel_mtion Seuor rels_ the green JETT TOga light on the teleltght panel
l_mLinstes. When the tower and the ca_mxle sepezste_ the No. 1 and No. 2 Tower
8el_mtlc_ Relays are de-energlsed_ _ pover to energize the No. 1 mad No.
2 _ Chute S_stea Arm 2 Second _ De].a,y Rel_rs. A_er 2 seconds deist, 1;he
Chute Rela_rs a:l the 21_000 foot barosvttches and the Nstn C_te Delay 2
Second Tlae Dela_ relays. After two seconds the IO,(XX) foot _mritebes are
armed. _e power clrcult v111 ho_ st these tvo polnts watll the capsule de-
seeads down through the 21,000 foot range, at which ttae the _ sequence
is tattlated. Refer to Sectlc_ IX of thls amnu_.
PAGE 5-6
REPORT _ 10_.
MODEL..1.33
Mc,DONNE REVISED
_TLT'J'_U| _!_S_ S'O_U RI REVISED
5-5. SECOND__
5-6. _n_io_i
Second staging is initiated by sus_J_er engine cutoff at which time the
capsule adapter bolts are fired providing acceleration has decayed to .20g. The
three posl_ade rockets and the four explosive electrical disconnects are fired
lnmediately after the bolts are detonated and result in capsule separation.
Capsule separation is sensed and _uitiates five seconds of rate damping which
is followed by orbit orientation in which the capsule rotates 180 ° degrees and
Settles into a 3_ ° orbit attitude.
5-1'. OPm_TION
At approximately 285 seconds after launch, second staging will occur (See
Figure 5-3)- At thie tJ_e a 28 volt d-c signal fran the missile _-1_1 energize
the Sustainer Engine Cutoff Relay. _en the thrust drops below -_)E, the ._0_
switch in the Thrust Cutoff Sensor closes. Power is then suppl_ed through the
Capsule Separation 1 Second Time Delay and _ta_er Engine Cutoff Relays after
the 1 Second Time delay to energize the Capsule Separation Bolts Po_er and the
Capsule Separation Warning 2 Second Time Delay relays. Through the energized
contacts, power from the Main and Xsolated bus fire the Capsule Separation Bolts.
Also Main bus power is suppled through the Warning Light Time Delay relay to
l]._te the Red CapsuZe Separation Sequence Light on the Lef_ Hand Console.
the Tri-se_ented Capm_le-to-Adapter clamp ring separates it a3_lows the
Capsule Adapter Ring Limit Switch to close supply_z_ power to energize the Posi-
grade Rocket l_re, Emergency Pos_rade Rocket Fire and the Capsule Adapter Dis-
connect Squib Fire relays. The Main and Isolated busses supp_ power through the
energized contacts of these relays to fire the Pos_Erade Rockets and the four
Cap_e Adapter Explosive Disconnects, The Posi_pmde Rockets create sufficient
LT_ LT_L T_VlT_TT.L _F_------ - --
DATI_ i November ig61
REVISED
RE_ISi[D
M ,DONNE __ST. LOUIS 3, MISSOURI REPORT _ 1(_
Mo=_. 133L --
\
thrust to separate the capsule from the adapter. This allows the three Capsule
Separation Limit switches which are attached to the retrograde straps to close.
Power from the Isolated squib bus flows through the closed contacts end acti-
e
rates the #1 Capsule Separation Sensor relay which extinguishes the Red Capsule
Separation Warning Light and illuminates the Green Light. Power also flows
the Capsule Separation Sensor relay to the 5 Second Time Delay DeWing Slgnal
relay. Activation of the Dsmping Signal relay actuates the orbit orientation
relay bringing the capsule to a 3_ ° (blunt end up) orbit attitude.
5-8.
5-9. SC IPTZ0a
In order for the capsule to impact at a designated area, the re-entry
sequence must be initiated approximately 3000 nautical miles up range of the
touchdown point. The method Of initiating normal re-entry sequence is by the
closing of the Retrograde Firing Signal switch within the Satellite Clock. The
switch may be activated by the run-out of time pre-set into the clock prior to
launch for a calculated re-entry time. _ starts at booster lift-off. The
time my also be pre-set by the Astronaut or by Ground Co,and when necessary.
The sequence may be started directly by using Ground Command transmitters and the
Capsule Co,and Receivers. The final method is for the Astronaut to manually
start the sequence by pressing the Retro Sequence h_tton. The last two methods
by-pass the satellite clock. A brief res_e of the sequence starts with the
closing of the Retrograde Firing Signal switch, which energizes the Retro Sequence
Fire 30 Second Time Delay relay. After the capsule has attained the proper atti-
tude end the time delay has run out, the three Retro Rockets will fire 5 seconds
apart. _en the Retro Rockets fire end the Auto Retro Jettison switch is in the
"ARM" position the Retro Rocket Asselbly Jettison 60 Second Time Delay Relay is
enerEtzed end at the run-out of the 60 second time delay the retro package is
.......... ""f ff R!
PAGE
REPORT
MODEL
5-8
SEDR 104
133
Mc,DONNE__ PST.LOUIS. MiSSO0'Rr "_'_
C,-,-,,-:_ _,__ilT i I I
DATE
REVISED
REVISED
1 NOVEMBER 1961
®
t---I
iF--
__////////////////////
_o
II
I
_.A .
//////
//////
'//i
Zo_-
(3.
_n,O
z_O..J
bj O-JO
,Po_q o
(j c_ _ N
DJ
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{Z :"
_o 3
O. ,_,v uJ
IZ0
II z
__ 0_.
_o_
J
!
ojzo
r',,, _ o
FICIURE_-3 SECOND ,STAGIE_ (._LOC_KD(A_I_,M)
ir¢ i _r
¢ 'JI :
I
_Zq
I
/
"i
C'C._:V :='.--, _,,T_-
B//¸_ , \
DAV_ i November 1@61
RE_/ISED
R_ISED
ST. LOUIS 3, MISSOURI
Jettisoned.
.'AGE 5-9
.",'O.V S_m ]._>_,
The retro package separation is sensed and results in the separa-
tion of the three retro package electrical umbilicals.
5-10. OPER_ION (See Figure 5-4)
The satellite clock Retro Fire switch is armed by Main Squib bus power
through the Capsule Separation relay (See Figure 5.-4). With the Retro Delay
switch in the "NORM" position and the Retro Fire switch closed by any of the
three previously mentioned methods, the Retro Rocket Sequence Fire 30 Second
Time Delay relay Is energized. At the same time the Retro Sequence Indicator
Relay is activated illuminating the Green Retro Sequence Indicator light. With
the Retro Delay switch in the "INST" position power is supplied directly to the
contact of the Attitude Permission Relay No. 1 by-passing the 30 second delay.
The Astronaut may manually start the Retro sequence by pressing the Retro
Sequence switch on the lef_ hand console which will energize the Emergency
Retro Sequence Relay and allow the nodal sequence to be followed. The Retro
Interlock switch in the ASCS Calibrator closes allowing power to flow through
the Emergency Retro Fire No. 1 Relay and energize the No. 1 and No. 2 Attitude
PeDaisslon Relays when the capsule is in the proper position for Retro Rocket
firing. The Red Retro Attitude telelight is switched on when the Retro Rocket
Sequence Fire Relay_ 30 second time delay is depleted and the capsule has not
attained the proper re-entry attitude.
Normally, the Attitude Per_sslon Relay exti_u/shes the Red Retro Attitude
telelight, illuminatil_ the Green light and energizing the Retro Signal Latch
and the Retro Rocket Fire Relays. Power from the Isolated Squib bus is nov
routed through the Retro Rocket Fire Relays to the Retro Rockets in turn f_ri_
first the Left (No. 1) Rocket, after a 5 second time dell7 the Bottom (No. 2)
Rocket and 5 seconds later the Right (No. 3) Rocket. Through the Retro Sisal
Latch Relay, a circuit will be ccRpleted to energize the Retro Fire Signal
vvl_a allu_aw • a,lr-Sahm F
PAGE 5-1Q
REPORT _ l(_
MODEL. l_
Disengage 23 Second Time Delay Relay. The power to the coil of the Retro Fire
Signal Disengage Relay allows the circuit to be completed to the ASCS Calibrator
resulting in hi-torque RCS operation. This high-torque operation will last for
23 seconds, which is 3 seconds more than the duration of total retro rocket
firing. The high-torque mode holds the capsule in the 3_ O attitude while the
rockets are firing. At the end of 23 seconds, the Retro Fire Signal Disengage
23 Second Time Delay relay will energize removing power from the Retro Fire
Slgnal relay and thus removlng the high-torque signal. With the Attitude switch
in the "AUTO" position the Astronaut may press the Retro Fire switch energizing
the No. 1 Emergency Retro Fire relay allowing Isolated bus power to energize the
No. 2 Emergency Retro relay powering contacts on the No. 2 Attitude Permission
Relay firing the Retro Rockets when the capsule assumes the proper attitude.
With the Retro Attitude switch in the "BY-PASS" position, and pressing the Retro
Fire switch which energizes the Attitude Permission By-Pass relay, the Retro
Rockets will be fired regardless of the attitude of the capsule. _en the Retro
Signal Latch relay is energized, Main Squib bus Power is supplied to energize the
Retro Rocket Assembly Jettison 60 Second Time Delay relay and the Retro Fire
Warning Light 20 Second Time Delay relay. After the 20 second time delay has
run out the Red Fire Retro Telelight is illuminated. When velocity decreases to
2_0 ft./see, the velocity sensor supplies pre-impact, Main bus power to energize
the RePro Rocket Fired Relay, activating the rela_ removing Power from the Red
Fire Retro telelight and illuminating the Green light. With the Automatic Retro
Jettison switch in the "ARM" position, at the end of the 60 second time delay the
Retro Rocket Assembly Jettison 60 Second Time Delay relay allows power to ener-
gize the Retro Rocket Assembly Jettison relay and the Jettison Retro
Light 2 Second Time Delay relay. As the 2 second time delay is expired the Red
Jettison Retro telelight is illuminated. The _etro Rocket AsseRbly Jettison
J
/--- .
DAT£ 1 NoveE_er i_i PAG£ _-ii
REVISED. ST. LOUIS 3, MISSOURI REPORT _E]_ I0_
_zvlsEo. _ MO_. 133
relay directs Main and Isolated bus power to the two squibs of the Retro Rocket
Assembly Jettison Bolt. The bolt will fracture and the package will drop free
of the capsule 2 being assisted by a coil spring installed between the heat
shield and Retro package Asse_ly. T_ dropping of the retro package assembly
from the capsule will allow the three Retro Rocket Asse_7 Separation _nsors
(single pole limit switches) to return to their normal position energizing the
Retro Rocket Assembly Separation re_7. This will allow the Retro Rocket
Assembly Umbilical Separation relay to energize, firing the six squibs of the
three Retro Rocket package tm_ilicals and Jettison the electrical umbilical pluEs
milliseconds after dropping of the retro pack_e. When the Retro Rocket Assembly
Umbilical Separation Relay is activated_ it removes power from the Red Jettison
Retro telelig_t and illuminates the Green light. When the Retro Rocket Assembly
Separation relay is energized, it activates the Accelercmeter Arm 5 Second Time
Delay relay. At the end of the 5 second time delay_ the relay functions suppl_o
lng a ground for Main bus power _lch operates the .05g relays. The energizi_E
of the No. 1 and No. 2.0Sg Retro Sequence Drop, .O_g Jettison Assembly Sequence
Drop, and .O_g Retro Telelight Power Drop rele_s removes po_er from the various
relays and switches involved in the Retro Sequence as well as extin_tishi_ the
Retro telellEht s.
5-11. ESCAPE SYS_
The escape system primarily consists of a tower &ssembly desiEned to pro-
wide a safe means of abort between prelm_ch and sta_inE. By utilizing the post-
grade rocket system, escapes may still be initiated after booster staging and
throughout sustainer operation until orbit. The to_er assembly consists of a
10 foot, tubular steel structure with a _ foot escape rocket mounted to _ts
tapered end. A segmented clamp ring with 3 _xplostwe bolts secures the t_e of
SEre 104 REVISEDREPORT ..... .ST. L.OUiS ,. MISSOURI
MODEL_ I_B _; _i __,_ _j T_.._ ' REVISED
the tower to the recovery con_artment upper flange. Attached to the escape
rocket nozzle adapter plate is a Jettison rocket which is used to Jettison the
tower assembly after the escape rocket has been fired for an abort; however,
under normal launch conditions the escape rockets are fired to accomplish tower
separation at time of booster engine separation.
5-13. ESCAPE BEFORE LIFTOFF _EFORE UMBILICAL DISCONNECT
Only one ground controlled signal will energize the Mayday relays. This
signal is a direct hardline from the blockhouse abort switch through the missile
to the capsule Mayday relays. In the event that the capsule must be aborted on
the launch pad and the missile is unable to transmit the hardline abort signal,
there is one other method which may be selected. Umbilical pins _ and 45 are
abort wired and transmit 28 V power from the blockhouse to the capsule's
Ground Command Abort Signal Latching relay, energizing and locking in the relay.
Through this energized relay capsule 28 V Squib Arm Bus power is transmitted
to the pole of the Ground Test Umbilical relay; however, power will not continue
through this relay until the relay is de-energized. The only way the relay may
be de-energized is by ejecting the umbilical. Therefore, if this abort method
is required to be used it would be necessary for the blockhouse conductor to
first select the Abort switch _ower to pins _ and 45) and then milliseconds
thereafter eject the umbilical.
5-1_- ESCAPE _ LIFTOFF AFTER UMBILICAL DLSCOKIECT
During countdown there will be approximately 50 to 90 seconds between time
of capsule umbilical eject and time zero, which is two inches liftoff. During
this period, the three available methods of abort are: (1) The blockhouse to
missile hardline abort signal as explained in the previous paragraph; (2)
Ground cc_Aud receiver abort signal; (3) Astronaut's Abort handle. These three
/
REVISED ST. LOUTSs, MISSOUR; REPORT _R1:_ 104
REVISED MOI:_L- 133
methods all result in energizing the Mayday relays.
5-15. ESCAPE AFTER LY__FF _EFC_E TO_R SEPARATION
After liftoff, (Time Zero), there are three methods by which an abort may
be initiated. They are: (I) Ground command receiver abort signal and (2) Astro_
naut Abort handle, both of which were possible methods in the previous paragraph,
(3) The Booster Catastrophic Failure Detection system. This third method has
been non-effective in the two previous paragraphs due to the Time Zero relay
being de-energized. However, the Time Zero relay is energized two inches after
liftoff and completes a circuit to the M_ay relays if the Catastrophic Failure
Detection relay is de-energized by loss of power from the missile.
5-16. Operation
When the Mayday relays are energized, the abort sequence (See Figures 5-5
and 5-6) will occur as follows: The ABORT light on the left hand console will
illuminate, the Capsule Separation Bolts Power relay is energized, the Capsule
Separation Warning Light Time Delay relay is energized, and the Tower Jettison
20 Second Time Delay Relay is energized. After 20 seconds have elapsed the Abort
Relay in the Maxim,-- Altitude Sensor is energized. The M_ Altitude Sensor
computes the time delay required for the capsule to reach a safe dynsmic pressure
hefore Jettisoning the escape tower. The capsule separation bolts squibs will
be fired, releasing the capsule-adapter clamp ring and allowlng the three limit
switches to return to their normal positions energizing the Emergency Escape
Rocket Fire relay, the Escape Rocket Fire relay and the Capsule Adapter Dis-
connect Squib Fire relay, firing the escape rocket and the four capsule adapter
explosive disconnect squibs. The e_cape rocket's 56,000 pounds of thrust rill
separate the capsule from the missile and carry it away from the sustainer and
at a small angle. The capsule separation sensor limit switches also energize
the Capsule Separation Sensor relays, which turns on the Green capsule Separation
PAGE
REPORT
MODEL
5-!14
SEDR 104
133
Mc'DONNE
_VI11 I_IJlIJIil I Ig'_ iI-
DATE
REVISED
REVISED
1 NOVEMBER 1961
P_E-SETBY qRo_%
PP.,E-SET _'f FAS"T P,ON/_T
AUTOMAR_C
PEE-SET
BEFORE
' SATELL_.TE CLOCK I
_£T RO C_RA,DE
S_G_KL S\N _TCH
,,,,, T
CAP%OLK
SE.P _,RATIO_
R£LA'Y
tMh_N j7..4-V D.C. _U S
_ RETRO SEqUEnCEREL!k'Y
(3osEc T.b.)REL _,'Y
IASCS CALIBRATORI
"--"P1R,ETP-,OI_TERLOCKl-'--
l _W'Tr"H I
LcIR R _ RETROI ATTITUDE
P E.P,M t_S_ON l IR,T__L ,l_'x'
RE'fRO R.OCKETe,, LI
TVELOCITY
SENSOR,
NOTES
[D AFTER _o SEC.T.D ¢_CORR_-CT•_,TTI'TL_D_ IS A_X_I_D
AFT_._. SO g_C..T.D.
ATTKR _,0 _EC.T.D,
[_D AFTEIK ZO SEC..T,_).
[_ PWR RKt,AOVED AFT_.R2q_ Sr..C.
ATTEIK ?_ SEC.,T.D.
RETRO F_RE
W_ L_HT_EL _,Y
R.ETRO ROCKET
FIRE R.EL&'YS
IR ET_,O g iQ_,IA,L
L A,TC_, _.EL k'f
R.ET IKO ROCKET
ASSE_BL'X"
SE_APRATiON
RF.LA,y
1,P-,ETRO ROCK_.T I
L_I_ILICAL
_EPA,R_T Io_
IK_-LAy
R_.T RO IKOCKET
LIMBIL_C&L
b I_CO_',II_ECT
_ P-,ETRO ROCKET
SEP,_RAT_OI_
SENSOP., I RETRO ROtKET
_SSEMBLY
,TE.TT_SOl,i P,ELK'(
RETRO FIEE IS_C_N_L b_SE_C_AC_EL..__.
(.5o SECT.b) I_
R ELik_l" ]
RETP, O ROCKET __
S E PAP-,A,Tiot,i
BOLT
_ R,ETRO ROCKET l
ASS EMILY I_--
3"ETTI _o_,i I -
_'_.TT_SON _TROWARN _NC_ L IC_HT
(z sEcT.b)RELLY
,IETTI SOR,ETR.O
FIgUrE 5-4- NORMAL P-,E-EI'4TRYS£QL_ENCE PM_5-Z3BA
!: " i
i" C_ _, _", ,", _" '_' "" "_ ._..... . I I I _ I I
/
!
/
/
®
_ (_
'_Y \ .-//
®
_®
UPON RECEIPT OF ABORT SIGNAL...
_ SHUT OFF BOOSTER, SUSTAINERENGINES.FIRECAPSULE ADAPTER BOLTS
_2_ FIREESCAPE ROCKET
(_ SENSE CAPSULEADAPTER SEPARATION,JETTISONRZTRO-PACKABEAND JETTISON RETRO-ROCKET UMBILJCAL_
_._ MAXIMUM ALTITUDE 3ENSO_ RUN3 OOT FiRE TOWER 3EPAI_AT/ON BOLT3I
SENSE TOWER RING SEPARATION, FIRETOWER JETTISON ROCK_.T
_ SENSE TOWER SEPARATIONTHROUGR ELECTRICALDISCONNECT COMMAND _'ATE DAMPING, 3' 5E.C.TIME DELAY EJECT ANTENNA FAIRINr_ANI) DEPLOY MAIN CHUTE-
RATE PAMI_IN6 .STOP3 AT CHUTE PEPLOYMENT
_ mm 111o o
z0<m
m
,w
m >
II
PN
REPORT _I_ 10_ REVISEDST. LOUIS, MISSOURI
MODEL-] 33 _ - _ REVISED_f .....
telelight, and also energizes the Tower Separation Abort Interlock Latching
relay. The abort interlock relay energizes the Retro Rocket Assembly Jettison
relay through the"ARM" position of the Astronaut's Auto Retro switch and fires
the two squibs of the retro rocket assembly jettison bolt. The bolt will
fracture and the package will drop free of the capsule, being assisted by a coil
spring installed between the heat shield and retro package assembly for this
purpose. When the capsule reaches a maximum altitude, contacts in the MAX ALT
SENSOR will close and energize the Tower Separation Bolts Power relay firing
the bolts. As the three tower bolts are fractured, the se_nented tower clamp
ring: separates allowing the three tower ring limit switches to return to their
normal position energizing the Emergency Jettison and Jettison Rocket Fire re-
lays. Through these relays and their parallel contacts Main and Isolated bus
power will fire the two squibs of the Jettison rocket. The tower will be
_ettisoned clear of the capsule resulting in separating the two tower to capsule
electrical disconnects. The separation of either disconnect will de-energize
the Tower Separation Sensor relays energizing the Abort Rate Duping relay.
This relay will send a signal to the A_CS cosmmuding rate damping until time of
main chute deployment. De-energizing the Tower Separation relays will also
start the two 2 second timers in the recovery sequence which will arm the 21,000
foot and i0,000 foot baroswltches after 2 seconds.
5-17. ESCAPE A_ TO_ER SEPARATION
The methods of initiating an abort after staging are identical to the
methods named for the escape after liftoff and are: (I) Ground command receiver
abort signal; (2) Astronaut Abort handle; (3) Booster Catastrophic Failure De-
tection system. Any of the three methods will energize the Mayday relays. The
sequence which occurs by the energizing of these relays is described in the
d-I f'l, "_T 1"_ T T_ T_ "ILl' rr, T "1
DATE
REVISED
REVISED
1 NOVEMBER 1961 Mc'DONNE ST.LOUIS. MISSOURI
PAGE ,,,
REPORT
MODEL
5-17
SED, R 10,4
133
.s-f- ,,,
: ,. \
MAXIMUM
ALT%TUDE
5ENSO_
BLOCK
HOu 5E
CAPSULE. 5EPAP_TtON TgLELIGHT
' HJ---. CAPSULE
.__J " ,.SEPARATIGN_ I SENSOR
/ I RELAYS
L
TOWER
_EPARATION
ABORT
_NT_LOCK
RELAY._
JE'ITt._3N
RETRO
WARNING
LIGHT
2,0 SEE.
TIME. bELAY _@_LAM
CAPSULEUMBILICAL
H GROUND H
COMM,
A_ORT
SIGNAL
• RELAy,(ENERG.)
ABORT LIGHT
CAPSULE
SEPARATION
SENSOR
LIMIT
SWITCHES
TO'WE _
JETTI SON
?-0 .SEE. TIME
DELAY
TOWER,
JE[TISON
WN. LIOBT
TIME DELAY
RELAY
ESCARE
ROCKET
4 CAPSULE.
ADAPTER
EXPLOSIVE
DISCONNECT
V ill
TowH H5EFARATION TOWER
BOLTS RINC9POWER BOLTSRELAY
H HREIRO ROCKET ROCKET
JETTISON ASSEMBLY ASGEMP_YSWITCH JETTISON JETTtSON
(ARM POSITION) RELAY E::)LT
LANDINO
SYSTEM
Z SE.C.
TIME 0ELAYRELAY
RATE SEPARATION
PAMPItJG AB3K["tNTE.R- LOCK
RELAY RELAY
i
L J RATEASCS
_ DAMPING/ RELAYAISW,TCH
),
GROUND
TEST
UMBILICAL
RELAY
DE-ENERG.)
,i
CAPSULE
SEPARATION
WARNINOLIGHT
TIME DELAY
RELAY
ANTENNA
FAIRING
EJECTOR
ESCAPE
ROCKETFIRE
RELAYS
CAPSULE
ADAPTER
DISCONNECT
SQUIBFIRE RELAY
F ,TOIVER
R II,,IG
uMIT
5_)ITCHES
MAY'DAY
RELAY5
CAPSULE
SEPARATIONBOLTS
POWERIRELAY
CAPSULE
SEPARATIONBOLTS
CAPSULE
ADAPTERCLAMP RING
LIMIT
SWITCH
JETTISON
ROCKET
FIP.E
RELAYS
)
H RETKO H RETR0
PACKAC_ ROCKET
ASSEMBLY ASSEMgLY
l LIMIT SEPARATION
5Wi TCHES , RELAYS
RETRO JETTISON TELELICq-IT
I R I G I-R i- I _T_o II I _ J l J I I _ C K E T I
_ I ASSEMBLY, IIf' " " _ "l [ I UMBILICAL " I
I / / I SYSTEM II RETRO / / I RELAY II ROCKET I_, g I i
I PACK_I uM_,_A_ I
J LEJgg,_L
/ SEPARATION I. ; I /
/ RELAYS I "
-- - I _ MUST _E ENEREdZE0 _Y UMSIU-
CAL RELEASE FOR C.I_CUIT TOGET THRU,
//_ SWITCH CLO_ED _LOW E)_OOOFt
FIGUI2E5:-G. ESC.-A.PE BEFOI_E STA61N(5 BLOCK DIA, roI2AAA P_,;s-_;B
PAGE 5-18
REPORT _]_ 104
MODEL 133
following paragraph.
5-18. O_eration
MCDONNE REVISED
ST. LOUIS, MISSOURIREVISED
The signal which energizes the Mayd_7 relays also is transmitted to the
missile to shut down the sustainer engine (See Figure 5-7 and 5-8). Through
contacts of the energized Mayday relays, a power circuit is completed to the
ABORT light on the main instrument panel and the .20g contacts of the thrust
cutoff sensor are armed. As thrust decays to .206 the contacts close and
energize the Capsule Separation Bolts Power relay firing five capsule separation
bolts squibs and separating the capsule adapter clamp ring. The sequence follow-
in6 clamp ring separation is the same as the normal sequence (Refer to Paragraph
5-7)- Re-entry may be acccmplished by any of the emergency procedures (i.e.,
Astronaut or ground initiated). Refer also to Paragraph 5-9. If the abort is
initiated before the capsule has obtained the correct velocity for orbital
flight and it is not desired to fire the retro rockets, the retro package must
be Jettisoned manually. It should be noted that even if the capsule does not
attain orbital velocity the quickest way for re-entry is by emergency firing
of retro rockets.
5-19. TEST CORFIGURATION CAPSULES
5-20. TEST CORFIGURATION CAPSULE NO. 8
5-21. General
Capsule No. 8 is the same as Specification Compliance Capsule in the se-
quence of events that takes place; however, the circuitry of the Launch Orbit
and Escape and the Retrograde Relay panels differ in the following manner.
5-22. Launch Throughsecond
Same as Specification Compliance Capsule, except that immediately upon
firing the Tower bolts_ the Tower Ring limit switches spring to their actuated
P,G'f, YIr_, ;'..N =:
"11
c
111
--4I
•., f 0
Z
UPON RECEIPT OF AISoR'r SIGNAL...
Q SHUT OFF 5USTAINER ENGINE
Q SF-NSE T_RUST DECAY TO .?.DO AND FiRE CAPSULE ADAPTER I_OLT5
O SENSE CLAMP RiNG RELEASE, i:iRE CAI_SULE TO ADAPTER EXPLOSIVE.DISCONNECT5 AND File I:IO_IGIRADE I:I.OCKET_.
(_ SENSE CAPSULE-ADAPTE._ SEPARATION. START 5 5EC RATE DAMPINQ,AND JETTISON RETRO-PACKA_E,
Q ]_PTER S .SEC. GATE DAMPlN6 CAPSULE TURN,_ AROUND.
Q ABTRONAUT C_ROUNO COMMAND _NiTIAT'..5 EMEP,6E_CY PROCEED-ORURE5 TO PERFORM RE-ENTRY AT 34" ATTITUDE,
?",4
i-OE
(I)0
!
D m
rtlr" zl
"4
¢/1fln
4_
O>-IIll
'1>QIii
PAGE 5-2'0
REPORT SEDR 104
MODEL 133
Mc'DONNEL LST.LOUIS, MISSOURI
DATE
REVISED
REVISED
1 NOVEMBER 1961
GROUND
P,EC EIV_.
ABORT
NANDLE
MISf:JLF_
CATASTROPHIC
DETECTOR
AROI_T LIGHT I
4 CAPSULE
ADAPTEREXPLOS_VE
DIBCONNEC/5
L_ CAPSULE SEPARATION
DAMP(NG
51C-NALSSEE.
T.D._LAY Z_
CAPSULE
SEPARATION
SENr-_R,
P,ELAYS
,(.___
CAPSULESEPARA_]0N
BOLTS
POWER.
R.ELAY
CAPSULESEPARATION
WARNIN&
UONT
I .SEC. T,D.RELAY
CAPSULE
ADAPTER
DISCONNECT
SQUIB
FIRE RELAY
CAPSULE
SEPARATION
SENSOR '
LIMIT
5WITCHES
CAPSULE
S£PARATI_
_OLTS
CAPSULE
ADAPTER
5I_PARATION
P, IN(a
LIMIT 5WITCMES
iPOSIG_R/C)E
ROCKETS
FiRE
RELAYS
POSIGRADEROCKETS
//_ REMOVED AFTER 5 SE.C.
//_ APPLIED AFTER 5 5£C.
FIGURE 5-8 ESCAPE SYSTE.N,1 AFTER TOWER SF..PAP,ATION(BLOCIX,DIA_P.AM) PM4s-c,os,I
/
< _'_
DATE ! November I_i _ __A_ PAGE _-21
R_'iSEID, ST. LOUIS 3. MISSOURI REPORT _ 10_
.............. ., MOD_ 133L ,
REVISED,
position and complete a circuit to the Emergency Jettison and Jettison Rocket
Fire relays and to the Emergency Escape and Escape Rocket Fire relays. _ough
these four relays which are connected in parallel for each rocket circuit, the
rockets are fired and the tower separates from the capsule.
_-23. Re-_atz 7
The stone as Specification Capsule except that the velocity sensor is re-
placed by three pressure switches, one on each Retro Rocket. Closing of these
pressure switches energize three Retro Rocket Gone Relays which extinguish the
Red Retro Telelight and illuminate the green Retro Telelight.
5-24. TEST CONF_ION CAPSULE NO. 9
uener
Capsule No. 9 is the same as the Specification Con_liance Capsule in the
sequence of events that take place; however, the circuitry in the Retrograde
Relay panels differ in the following manner:
5-26. Re-Entry
The Retro Rocket Fired Relay which extinguishes the Red Retro Fire telelight
and illuminates the Green light is energized by a pressure switch located on the
Right (No. 3) Retro Rocket in place of the Velocity Sensor. All the Retro
telelights are extinguished through a series of relays except the Retro Jettison
light which remains on.
5-27. TEST CCMFIGURATION CAPSOLE NO. 13
e,e r.Z
Capsule No. 13 is the _ as the Specification Compllance Capsule in the
sequence of events that take place; however, the circuitry in the Retrograde
relay panels differ in the following manner:
Mc,DONNE REPORT .¢_1_ 1(_ REVISED
MODEL 3.33 .o_TJ.OUI$_MIS$OURI REVISED.
DATE i November 1961
5-_. _-Entr,/
The Retro telelights are extinguished through a series of relays except that
the .05g Retro Telellght Power Drop relay is not in the circuit and the Retro
Jettlson light remains on.
5-30. TEST COBF_RATION CAPSULES NO. iO AND 16
Capsules No. IO and 16 are the same as Specification Capsules.
Q
6-I-
/'4_- _,
S£CTION V_
ESCAPE AND JETTISON ROCKET
::::::'_--_.._
'::_ii_iiiiiiiiiiiiiiiiiiiiiHiiii_'":
SYSTEMS
TABLE OF CONTENTS
TITLE PAGE
SYSTEM DESCRIPTION ......................... 6-3
SYSTEM OPERATION 6 3eeee.ee.eoeoeoeeee.eeeeeee.
ABORT HANDLE .0. ............................... 6-3
ESCAPE ROCKET ....... , ......... , .............. 6-S
JETTISON ROCKET. .............................. 6-7
TEST CONFIGURATION ....................... 6-11
_ALIIPlRPILIqI_N ,il, I
• n_ ¥ m li 8 IkllP n I I I I/_1_ b
PAGE
REPORT
MODEL.
6-2
SEDR 104
133
MCDONNEL_L_ __ST.LOUIS, MISSOURI
DATE ... 1 NOVEMBER 1961
REVISED ...
REVISED .
zS uJ
v'3
.,%
',O
,.oh" ,._ 0 --
x
FIGUPE ¢o'1 ESCAPE SYSTE_ (SI-IEE_I'I OF;?..)
QLY
Z
o_,i-.. c_
'-_ IJJ
.P_¢O
,J3
i
• .,,,_-_ il"i,L
• • i_/'
l o ember Mc,DONNE 6-3REVISED ST. LOUIS 3. MISSOURI RI_ORT _ 10_1 •
VI. ESCAPE AND JETTISON ROCKET SYSTEM
6-1. SYSTEM DESCRIPTION
The escape system consists of the capsule pylon, escape rocket and pylon
Jettison rocket. The system also utilizes items illustrated on Figure 6-1.
These items are not part of the escape system, but assist in the sequence of
events to complete the escape function. During normal flight conditions where
an abort is never initiated, all of the associated items perform their same
function but in different sequences. For a written description of" the escape
system and diagrams of the items under discussion, refer to Section V.
6-2. SYSTEM OPERATION
The capsule electrical system provides for an abort any time after the
gantry is removed. When an abort is initiated, the capsule adapter clamp ring
is released, the escape rocket ignited, and the capsule is carried to an
altitude of approximately 2500 feet. At this time, by means of the altitude
sensor, the capsule tower clamp ring is released, the tower Jettison rocket
ignited, and the escape tower Jettisoned. Two seconds after tower jettisoning,
the drogue chute is deployed. Two seconds after drogue chute deployment, the
antenna fairing is released. Twelve seconds after antenna fairing release, the
heat shield is released, extending the landing impact skirt. During a normal
mission where the escape system is not utilized, the escape rocket is used to
separate the escape tower from the capsule.
6-3. ABORT HANDLE
The abort handle's primary function is to initiate the abort sequence.
The handle is also used as a restraint handle during launch. Location of the
abort handle is forward of the Astronaut's support couch left arm rest. For
-CAP3ULE TO ADAPTERSEPARATION RIN_
SENSIN6 SWITCH
--CAPSULESEPARATIONSENSIN6 SWITCH
"----STRIKER (LOCATEDIN ADAPTER)
ADAPTER RING SEPARATION
3ENSOR AND CAPSULE-ADAPTERSEPARATION SENSOR
ESCAPE AND _-JETTISON [Z.
ROCKET -,_
u)
rn
H
0-rlr_
v
CAPSU LEITO ADAPTER.
SE PARATIOIq RIN6
"D
J>.
i
RECOVERY--_COMPARTMENTCAPSDLE
CLAMP RING
SEHSING SWITCH _
CAPSULE RECOVERY_COMPARTMENT
TOWER SEPARATION
RING SENSING SWITCH-
0 m >
Plr" •
-t
!
m Z
o _C
0
;o _o oI11 Ill ),<: < ,-.I
rrl I11El 0
Z
rl
DATE 1 November 1961 pA_,, 6-5
RUrVlSED ST. LOUIS 3, MISSOURI REPORT SRDR l0 b,
._,sEo _omm. 133, , J,
an Astronaut initiated abort, the release button located in the top of the
handle must be depressed, allowing the handle to be rotated outboard. When
moved to the abort (outboard) position, an electrical switch is actuated,
which acts to detonate the capsule-to-adapter clamp ring bolts. The escape
sequence is then initiated, providing that the main umbilical has been dis-
connected. Before umbilical release, the abort handle is inoperative. (See
6-2.)
6-b,. ESCAPE ROCKET
The escape rocket consists of an electrically actuated igniter assembly, a
I inch b,130 steel case, rocket nozzle assembly, plenum chamber and a solid fuel
propellant (See Figure 6-b,). The length of the escape rocket is approximately
70 inches. The diameter of the rocket case is approximately 15 inches. The
weight of the motor prior to firing is approximately 350 pounds. For aero-
dynamical stability, ballast is added to the top of the rocket (See Figure 6-1
Sheet i of 2). The nozzle assembly incorporates three equally spaced exit
cones. The exit cones are canted at 19 degrees from centerline of rocket case
to centerline of exit cones so as to direct the rocket blast outward and away
frQm the tower and capsule. The plenum chamber incorporates a Jettison motor
boss which facilitates for the installation of the Jettison rocket motor. The
Jettison motor boss also provides for the attachment of the thrust alignment
mirror. The optical sighting of the resultant thrust vector is accomplished
by the thrust alignment mirror. For off the pad escapes the rocket will obtain
a maximum capsule altitude of approximately 2500 feet.
The escape rocket propellant is a polysulfide ammonium perchlorate formula-
tion which is widely used in the rocket industry. The United States Bureau of
Explosives classifies the propellant as a "Class B Explosive". The propellant
is sensitive to pressure and a spark or flame may easily ignite it. The
• _ left1 A •
PAGE 6-6 r
REPORT SEDR 104
MODEL. 133
Me'DONNELL__ __ST.LOUIS, MISSOURI
l -b
DATE
REVISED
REVISED
I NOVEMBER 1961
\\
J
DAVE I November 1961
REVISED.
REVISED.
ST. LOUIS 3, MISSOURI
pA_z 6-7
Rm,ORV SEDR i0_
.om,_ 1_3
'< ••
propellant grain is an internal burning nine point star which is cast directly
into and bonded to the case. With the nlne-polnted port design, the possibility
of thrust mlsalignment is reduced. This is due to the improved alignment between
the star ports and the exhaust nozzles. The nominal resultant axial thrust at
70 degrees F is 52,000 pounds for 0.78 of a second; it then drops off uniformly
to 5000 pounds in the next 0.6 of a second. The thrust will then diminish at a
reduced rate to zero. The total impulse of the motor, at sea level, is
approximately 56,500 pounds - second.
The escape rocket igniter is a head mounted dual unit with two completely
independent initiation systems to each unit. The dual initiation •system to
each unit has independent circuitry from different batteries. This igniter
is cylindrical in shape, and is a central dynaflow type of long duration. This
igniter is essentially a miniature rocket motor. It incorporates a small pro-
pellant grain which can be ignited by either of two squibs surrounded by
boronpotasslum nitrate pellets. Surrounding this is an annular plastic tube
filled with a metal-oxidant mixture in which are located two sets of four-
squibs. Either set is capable of initiating the igniter. Each igniter has
two initiation systems, either of which can start the igniter in the event some
of the squibs are inoperatlve. The igniter is a Class A Explosive.
6-5. JETTISON ROCKET
The Jettison rocket is a qualified Thor retro unit and is manufactured by
the Atlantic Research Corporation. The rocket consists of an electrically
actuated igniter, a motor case and a tri-nozzle assembly (See Figure 6-5). The
nozzle's cones are canted at a 30 degree angle (from canterline of motor case
to centerline of exit nozzle). It weights 19.5 pounds, has a length of 18 inches,
a diameter of 5.5 inches, and produces 785 pounds of thrust for 1.4 seconds at
70° F. in vacuum. The rocket has been successfully fired from -7_ O F. to 175 ° F.,
PAGE
REPORT
MODEL ,
6-!8
SEDR 104
133
ST.LOUIS, MISSOURI
DATE
REVISED
REVISED
1 NOVEMBER 1961
ESCAPE
• I_OC KET
"METAL CONDUIT
_TI.L,; _
DATE
REVISED
REVISED
1 NOVEMBER 1961
' Mc'DONNE_ _,,_ST.LOUIS , MISSOURI
PAGE , 6,;9
REPORT SEDR 104
MODEL 133
SGNITER FLIGI4 T
I-4 E_, D
PROPELLANT
PAGE 6!-1 O:
SEDR 10:4REPORT
133MODEL
Mc'DONNEL_L_ C__ST.LOUIS, MISSOURI
DATE 1 NOVEMBER 1961
REVISED
REVISED
/i
L,_ C E DEbFJ
l << c_ ,'61'Ai ,A,I_I
P OT T IN C-i
SQUIB _
I_N\TION PKLLE-TS _
Nl IGNITER F_E Iv__LW, /J
__ \I W/- _ov_,.,•
j-L, NEa
'b"mN6_...
NOT..ZLE TNSE_IT J (_"_,_J _/_ "-_" N0 ZT. L _: A, f) EM,_LY
ELECTRICAL DIAG, RAM
FIGUI:2E 6-5 JETTISON _OCI¢,F_TI" pM4s-_7-¢
j___7"
DATE l _°vemb_r l_ 1 _ _ PAGE 6roll
REVISED , ST. LOUIS 3. MISSOURI REPORT ._,nR IO).l.
REV'SED " " MOC_. 133J
and from sea level to vacuum.
The Jettison rocket igniter is a head mounted unit with dual ignition
capabilities. This unit is cylindrical in shape with a hexagonal head and
threads into the top of the jettison rocket. The igniter contains approximately
7 grams of USF-2D ignition pellets which are ignited by either of four squibs.
Each pair has independent circuitry from a different power source aud either
palr is capable of igniting the pellets.
6-6. TEST CONFIGURATION CAPSULES
6-7. TEST CONFIGURATION CAPSULES NO. 8, i0 AND 16
The above capsules differ from the specification capsule in the following
manner. For a normal launch, the escape tower will be Jettisoned from the
capsule by the simultaneous firing of the escape and jettison rockets.
64. •EST CONFIGURATION CAPSULES NO. 9 AND 13
Same as specification capsule.
7-1
l -"........"_.....................................$£CTJON
mPOSIGRADE ROCKET SYSTEM
VB_
TITLE
TABLE OF CONTENTS
PAGE
SYSTEM DESCRIPTION. ....................... 7-3
SYSTEM OPERATION ............................ 7-.3
POSIGRADE ROCKET ............................ 7-3
TEST CONFIGURATION ...................... 7-5
.-::::. .::::::"
-A AIIIPllll_lPLl_lel A |
_V|l| |flJimll • Io"_b_
PAGE
REPORT
MODEL
7-2
SEDR 104
133
MCDONNEL_L__"__ST.LOUIS, MISSOURI
DATE
REVISED
REVISED
t NOVEMBER 1961
.FIGURE -7-I
\\
\
POSIG_ADE I_OCKET SYSTEM
DATZ I November 1961
REVISED
REVISED
ST. LOUIS 3, MISSOURI ._oR'r ._.I_ 104
MOO_ .... 1_
VII. POSIGRADE ROC_ SYSTEM
7-i. SYSTEM _ESCRIPTION
The poslgrade rocket system consists primarily of the three poslgrade
rockets and igniters mounted in the retrograde package and the associated wiring
necessary to ignite the rockets at the proper time. (See Figures 7-1 and 7-2. )
7-2. SYSTEM OPERATION
The purpose of the poslgrade rockets is to accomplish separation between
the capsule and booster at a rate of 15 feet per second when orbital velocity
has been achieved. They also perform the same separation function during an
abort after tower separation. The three rockets are fired simultaneously;
however, should two of them fail, the remaining unit would successfully affect
separatlon.
7-3. POSIGRADE ROCKET
The poslgrade rocket primarily consists of a nozzle assembly and •case, a
solid propellant and an electrically actuated igniter. It is a cylindrical
device measuring approximately 14.7 inches in length, 2.8 inches in diameter
and weighing approximately 5.24 pounds. This rocket is basically an Atlas
retro-rocket with minor changes for increased reliability. Reliability has
been gained by two methods; (a) Dual ignition of the igniter squibs from two
different buses, (b) only one of the three rockets is necessary to accomplish
successful separation. Due to the wide temperature range of these rockets,
a temperature control system is not required. The propellant utilized in the
posigrade rocket is Arcite 377 which provides an average thrust of 370 pounds
for one second in a vacuum.
..... ° "_ n"_ l"Jl" I AI_,_1 11 I WIm,*l • , .....
PAGE
REPORT
MODEL
7-4
SEDR 104
133 _
Mc'DONNEL__ __ST.LOUIS, MISSOURI
DATE
REVISED
REVISED
1 NOVEMBER 1961
(0 A
PO,StGR_,DE
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EM ERGEMP..Y
PQSIGRAOE
ROCK ET5 BUS
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EM ERGI_t_.YPOSIC_RADE _OCKE'r
FIRE RELAY "
w
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F'IGURE 7-_ POSlGRAOE ROCKET IGNITION 5Y'STEM _CHEMA'I'IC
"1"0
POSIGI_,DE ROCKET¢_
2 AND "_
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PM4S'G6A
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....... _='m • |4f-t. L--,
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L_
• r•:,•
OAT_ i November I_i
REVISE:D
RE[VISED
i
7-4.. Rocket Igniter
VONNE ST. LOUIS 3. MISSOURI
i
PAGE T-5
REPOnT SEDR 104
.o_ 133
The posigrade rocket igniter is a head mounted unit with dual ignition
capabilities. The igniter is cylindrical in shape with a hexagonal head for
threading it into the top of the posigrade rocket (see Figure 7-1). This unit
contains approximately 3 grams of ignition pellets which are ignited by either
of two pairs of squibs. Each pair has independent circuitry from a different
power source and either pair is capable of igniting the pellets.
7-5. TEST CONFIGURATION CAPSULES
7-6. TEST CONFIGURATION CAPSULES NO. 8, 9, i0, 13 AND 16
Ssme as specification capsule.
8"1
RETROGRADE ROCKET SYSTEM
S[CTION VJl_
TABLE OF CONTENTS
TITLE
SYSTEM DESCRIPTION .......................... 8-3
ROCKET MOUNTING • 8-3oeeeeeleeoeleeleeeee eeeee.
RETRO PACKAGE ELECTRICAL WIRING. 8-3
SYSTEM OPERATION ................... ......... 8-6
RETROGRADE ROCKET ........................ 8-8
TEST CONFIGURATION ....................... 8-9
PAGE
_..I A i I i. i l.k I.. L i .i_'i A i _'_ ............. _ ....
_Vlll l_ll I ini--
PAGE 8- 2
REPORT SEDR 104
MODEl, 133
Mc'DONNE_ __ST.LOUIS, MISSOURI
DATE
REVISED
REVISED
1 NOVEMBER 1961
ADAPTEP.
@
BLANPCETP, EC F-PTA C LE5
FIGUI_E 8-1 P, ETI_,OGI_ADE ROCKET SYSTEM _M4.5-65A
DATE,, 1 No'ulCer 1_ 1 , _ __#_, PAGE 8-?
REVISED. -,.,_- --_--_----7----'-ST,LOUIS 3, MISSOURI REPORT _]_ 10_"
REv,sEo ........... Moron. 133
VIII. REZRCORAIZ ROCKET SYSTEM
8-1. SYSTEM IESCRIPTIONQ
The retie rocket system consists primarily of the three retro-rockets,
their pyrogen i_£tere, and the associated wiring necessary for rocket _tlon.
The retro-rockets are housed in the Je_tlsonable retrograde packa_ along with
the ro ts. 8-i.)
8-2. ROCKET MDURTIm
The retro-rockets are mounted in the retro-package which, in turn, Is
mounted to the capsule by means of three straps _oined at the bottom center of
the package by an explosive bolt. Sixty seconds follovi_ retrograde firing
signal 2 the bolt detonates, the straps are released, and a toll spring e_ects
the package away fr_ the capsule. To protect the rockets, particularS7 from
micro-meteorites, each rocket has a metal cover over its exposed nozzle end.
The cower is blown off by rocket blast at time of light-off. Mount_n6 of the
rockets Is so designed as to direct the resultant thrust vector towards the
capsule's predetermined center of grawlty st time of firiug.
8-3. REPRO PACKAGE _CAL WlRI_
The retro l_ckage Is supplied electrical power through the three electrical
explosive disconnects that are equally positioned around the base of the capsule
Electrical wire bun_les from the dtsccanecte follov the three retrograde package
retenslon straps down to the retro package, where they enter this unlt through
rubber grc_mts. Within the packase all wiring for the retro rockets, sensors
and heaters are routed to the outside of the package face end then into each
retro rocket through the slotted metal shield. Postgrade rocket wiring and
explosive bolt virin8 remain within the package. (See Figure 8-_.)
PAGE
REPORT
MODEL.
8-4
SEDR 104
133
Me'DONNELL_L.L.L.L.L.L.L.L____._ST.LOUIS, MISSOURI
DATE 1 NOVEMBER 1961
REVISED
REVISED
Pg, E- 5£T
BY GROUND
COMMAND
AUTOMATIC
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t_E_QRE
LA.UN C_
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UMBiLiCAL
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FIGURE 6-2 RETROGRADE ROCKET AUTOMATIC FIRING 5YST£'_ PM45-_4-
J," r'l N ',=!m E ._:T. _-=-
DATE
REVISED
REVISED
1 NOVEMBER 1961 Mc'DONNELL H ST.LOUIS, MISSOURI
PAGE.
REPORT
MODEL
8-5
SEDR 104
133
(
¢g r-" -"ISOLkT_D
_ATT.
AN_ r_ECOVE_..Y ¢, _,A_I
IKIO. rZELAY
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PAGE 8-6
REPORT SEDR 104
MODEL_ 1_
McDONNEL LST. LOUIS, MISSOURI
DATE _. 1 November I_i
REVISED
REVISED
8-4. SYST_4 OPERATION
The purpose of the retrograde rocket system Is to slow the capsule prior to
re-entry. Actual firing of the rockets is preceded by a 30-second period during
which the capsule is positioned to the "retro" attitude. "Retro" attitude Is
defined as follows: 34 ° + 12.5 ° Pitch, 00 + 30 ° Roll or Yaw. The firing
sequence will not begin, nor_Llly, until the "retro" attitude limits have been
attained, and will be temporarily interrupted should the capsule exceed the
attitude 1/mits after the sequence has begun. Should the need arise 3 however,
the above ltmits may be manually overridden. Firing of the rockets, which occurs
at five-second intervals can be initiated by any of the following: (i) Satellite
clock runout; (2) Astronaut selection; (3) Ground command. Any one of the
rockets wall effect a satisfactory re-entry in the e#ent the other two fail to
fire. Sixty seconds after slgnal for retro fire, the entire retrograde p_ckage
is Jettisoned.
8-5. ROCK_ FIRING
All three rocket flre relays receive 2_ V d-c siemltaneous_7; however, the
No. 2 and No. 3 rocket fire relays have a five- and ten-second time delay, respec-
tlve_y. Therefore, the following table shows _hen each rocket receives its flre
signal and the length of burning time thereafter. Note that the asterisk lndi-
cares time of firing.
Left No. 1 R_ e
Bottom No. 2 RK_
Right _o. 3 R_
+
+
÷
123k 5 6 7 8 910 ll1213141516171819 20.
SECOImS
The retro rockets are fired sequentis_ to avoid the Inei_ectlve results
DATI_
REV,S_O -- "_" "_'a""'_rr.Loum s, MiSSOUm R_'OnT. _ 10_
133
\ /
from a failure of either of the first two rockets. Consequently, if No. 1 roc_t
failed to a degree which would disrupt the retro attitude position, the ASCS
attitude interlock would remove power from the attitude permission relay and
also the No. 2 rocket fire relay. The capsule could then be reposltloned auto-
matically or manually by the RCS, end upon regaining the 34° position s the No. 2
rocket fire relay would receive po_er and fire. The same sequence would occur
if the No. 2 rocket were to fail. (See Figure 8-2. )
8-6. RETRO ROCKET EMERGERCY OVERRIIE
There are four telellghts on the Astronaut's left console which concern
the retrograde system. The first one is RETRO SEQ. and is a green function
light. This light will illmninate when the retro sequence is started, either
by the satellite clock or by the button adjacent to the light. The purpose of
the button is to initiate re-entry prior to satellite clock runout or failure
of (seen re 8-3.)
The next two telelights in the retrograde sequence are REdO ATE. and
FIRE RETRO. For a normal flight, the RETRO ATTITUDE switch adjacent to the
RETRO ATT. light should always be in AUTO. The RETRO ATT. light will illum_te
green as soon as the capsule reaches the 34 ° retro attitude position. Approxi-
mately forty seconds later the FIRE REdO telelight should illmninate green. If
the _ ATT. telelf_ht 411,--tn-tes red, the Astronaut must check the capsule
attitude in order to determine if the capsule is in correct retro fire Position.
If the capsule is found to be in the correct attitude, then the Astrcnaut should
position the RETRO ATT. switch in the BY PASS Position, and also push the FIRE
REdO button. Ten seconds later the FIRE REdO light should illuminate green.
However, if the Astronaut determined that the capsule was not in the correct
Position s the fly-by-wlre system should be employed in order to correctly posi-
tion the capsule in the 34 ° attitude. (See Section IV. ) When this is accem-
..... ''"_"_I _ I
PAGE 8-8 MCl DONI_.h_._ ___T_TT A I"9 DATE 1 November 1961
REPORT _ 104 _ _'_--_" "_ REVISEDST. LOUIS, MISSOURI
MODEL. 133 _I* ....... REVISED,,
pllshed the RETRO ATT. telellght should illuminate green.
If the EETR0 ATT. telelight illuminates green, but the FIRE EETRO telelight
illuminates red, the button adjacent to the FIRE RETRO telelight should be pushe_
leaving the RETRO A_T. switch in the AUTO position.
The fourth telelight is the JETT. RETRO. This telellght will illuminate
green 60 seconds after the illumination of the FIRE RETRO telelight. In the
event this light illuminates red, the adjacent button should be selected to
supply an alternate source of power to the Jettison bolt. If the retro package
cannot be Jettisoned by the automatic or override method, it will be ejected
sometime during re-entry when the extreme heat encountered will detonate the
explosive bolt or burn the retention straps to allow the coil spring to eject
the package.
8-7. RETBO ROCEET HEA2ERS
The retrograde rockets are equipped vlth blanket type heaters operated by
means of ground power only. A resistor type thermostat maintains heater tempera-
ture between 75° + 5°F and 95 ° + 5°F preventlmg moisture from collecting on the
retro-rocket nozzle closures.
8-8. RETROGRAnE ROCKET
The retrograde rocket, manufactured by Thiokol Chemical Corporation, is a
variation of the Model TE-236. Leading particulars are : Total weight approxi-
mately 69.55 pounds, length 15.4 inches, diameter ]2 inches. These rockets
have a total impulse of approximately 13,000 pound seconds, providing an average
thrust of 992 pounds each for 13.2 seconds action time. Due to the importance
of the retrograde system to the overall mission, a redundant rocket firing system
has been employed. Dual ignition to all igniters has been provided from separate
electrical sources. Heaters are bonded to motor case and nozzle closure and are
MAC 23TCL (27 APR 59)
DAT_ I November 1961
R£VISED
IR£%rIS£D
ST. LOUIS 3, MISSOURI
P^G_ 8-9
MOm_ 133
.4--'-,
thermostatically controlled.
8-9. TEST CO_FIELTATION CAPSUI_S
8-10. T_ST COIFIGURATION C_ NO. 8, i0, 13 AND 16
The above listed capsules are basically the same as the speclficatlon
capsule.
_A_E '-,0 MC,DONNELL__.__¢_'_j,'ez._ t, _ DATE
REPORT, SEDR 104 ST.LOUIS, MISSOURI REVISED
MODEL.. 133__F:_ S".'T :_ L REVISED
I NOVEMBER 1961
\
: /
/
/
/
F_URE 8-,4--RE-FRO PACKAQE EL.EC-rRICALINS,TALLATION; _ ,,_,
.t_kll=..ll_lm.li.,-r- - : :
pH45-'13
• /
_lm_. a_.:.,p. ,._,,.,m,_ _ - •
S£CTJON JX
SEQUENCE SYSTEM LANDINGTHROUG RECOVERY
\
TITLE
TABLE OF CONTENTS
PAGE
AUTOMATIC SEQUENCE DESCRIPTION .. 9-3
AUTOMATIC SEQUENCE OPERATION .... 9-3
EMERGENCY SEQUENCE DESCRIPTION... 9-6
EMERGENCY SEQUENCE OPERATION ... 9-6
SYSTEM COMPONENTS., ....................... 9-11
TEST CONFIGURATION CAPSULES ........ 9-27
_^ilPII'_IP|I'TI A I .
tv i 11 i IBw bl 1 I Idr_ll
I_1315A5 ,K_I'_AOD3_I (3blW _bllq_IW70IIW_OIOW I-b SB_I3
.; ii:
03SLASH IHROSSIIN 'SlOOl'/s _ol llOlq /HOd'ai:l
lg6l _I_RW:IAON l
DATE i November i_i
REVISED
REVISED
Mc'DONNE ST. LOUIS 3, MISSOURI
p^_E 9-3
REPORT SEDR 104
MODEL I_
IX. SEQUENCE SYSTEM LANDING THROUGH RECOVERY
9-I. AUTOMATIC SEQUENCE DESCRIPTION
The landing and recovery sequence system provides automatic electrical
sequencing to land the capsule safely after an abort or a/_er re-entry, and to
initiate locating aids for assistance in the subsequent recovery. The primary
(completely automatic) system incorporates a drogue parachute, used initially
to decelerate and stabilize the capsule in the initial phase of recovery and a
main parachute which further decelerates the capsule. Capsule landing is
accomplished by a 63 foot diameter parachute which is deployed at I0,000 feet.
In the event of a main chute failure, a 63 foot diameter reserve chute may be
deployed by the Astronaut's manual selection. Both main and reserve chutes
are reefed to limit shock loads at initial opening. Automatically, the reefing
line is severed after a predetermined time delay an_ the chute will open fully,
lowering the capsule at the prescribed landing speed. After main chute deploy-
merit, the landing impact bag is extended, providing a cushioning effect for the
landing impact. Immediately'after impact, the main chute is autcma_ically dis-
connected and the reserve chute ejected. The Astronaut will then egress normally
taking with him the survival kit which contains a life raft and other survival
aids.
9-2. AUTOMATIC SEQUENCE OPERATION
On tower separation, power is removed from the Tower Separation Relays
allowing the Main Chute System Arm 2 Second Time Delay relays to be energized
(See Figure 9-I). After the 2 second time delay has run out power flows through
the closed contacts energizing the Main C_ute Delay 2 Second Time Delay relays
which in turn after a 2 second delay arms the 10_000 ft. Baroswitch. The 21,000
ft. Baroswitch is armed through the closed contacts of the Main Chute System Arm
PAGE 9-_
REPORT _I_R 1OZ_
MODEL_ l_
REVISEDST. LOUIS, MISSOURI
I REVISED
2 Second Time Delay relays. Upon descent to 9_I,0OO ft. the _1,000 ft. ]_o-
switch is actuated, energizing the Drogue Deploy relay, firing the Drogue Chute
Mortar, deploying the Drogue Parachute, and ejecting the chaff package. The
Drogue Chute stabilizes and decelerates the capsule; the chaff package disperses
finely cut _etal foil that provides an enlarged radar reflection area that can
be picked up by search planes hundred of miles distant. At approximately I0,000
ft. the I0,000 ft. Baroswitch actuates the Main Deploy relay resulting in the
removal of ground circuits and firing of all four squibs of the Antenna Fairing
Ejector. Also, the Main Deploy Warning Light 2 Second Time Delay relay is
actuated, and at the end of 2 seconds delay the Red Main Deploy Telelight is
illuminated. The Antenna Fairing Separation Sensor Arm relays are energized
through the closed contacts of the Main Deploy relays arming the two Antenna
Fairing Separation Sensor Switches.
The firing of the four Antenna Fairing Ejector Squibs causes the Antenna
Fairing to separate from the capsule. A lanyard, connected from the Antenna
Fairing to the Main Chute, extracts the Main Chute frca the chute compartment.
The Main chute opens initially in the reefed condition to limit shock loads.
Four seconds after the chute is deployed the reefing line is severed by a
pyrotechnic charge in the reefing line cutters allowing the parachute to open
fully.
The separation of the fairing from the capsule allows the Antenna Fairing
Separator Sensor Switches to function. Through the switches, power is routed
to energize the Main Ejector Relay firing the Main Ejector Bag Squibs. When
the squibs fire, they generate a gas, filling the Ejector B_ at the bottom of
the _ Chute cc_r_ent aiding the ejection of the chute. At the same time
the Antenna Fairing Separation Signal'relay is energized, illuminating the Green
Main Deploy Telelight and removin6 power from the Red Telelight. Power is also
CC_!FIDENTIA,.
directed through the Antenna Fairing Separation Sensors to energize the Main
Inertia Switch Arm 12 Second Time Delay relays. After the 12 second time delay
the energized contacts allow power to be supplied to energize the Landing Be_
Extend and Landi_ Bag Warning Light 2 Second Time Delay relays as well as power
to the Landing Bag Unlock Signal Limit Switch and Tnertia switch. The closed
contacts of Landing Bag Extend relay fire the squibs of the La_iing Ba_ Valve
releasing the heat shield, and extending the impact landing ba_. As the 2 second
time delay runs out the Landing Bag Warning Light relay illuminates the Red
Landing Bag telelAght. Upon heat shield separation the T_na4nZ Bag Unlock Signal
Limit switch is actuated and through its closed contacts power is directed to
energize the Landing _k_ Extend Si_ relay illuminating the Green Landing _ag
Telelight and extinguishi_ the Red light. The force of impact on landing
operates the Inertia Switch and in turn energizes the Electric Switch which pro-
wades a ground for power flo_ to the Impact relay coil. Through the closed
contacts of the Impact relay, power is supplied to energize the Post Landing
System Relay. Through the activated contacts of the Post T_na_nZ System Relay,
power is transmitted to energize the Impact Signal, and the Capsule Stabilization
1 Second Time Delay relays. Through the closed contacts of the PoSt Landing
System relay the self-powered Flashing Recovery Light circuit is con_leted
setting the light in operation. When the _pact Signal relay is energized the
Green Main Deploy and Landing Bag Telelights are extinguished and the Red Rescue
Aids Telelight is illuminated. At the end of the 1 Second time delay the Capsule
Stabilization relay is activated allowing the Main Disconnect and Reserve Dis-
connect relays to be energized. Through the energized contacts of Main Dis-
connect relay, the Main Chute Disconnect squibs are fired, releasi_ the main
chute f_ the capsule. The Reserve Disconnect Relays fire the Reserve Chute
Disconnect Squibs releasin_ the Reserve Chute and energizes the Reserve Deploy
REPORT ,SEDR lOLl• ST. LOUIS, MISSOURI REVISED
MODEL_ 133 ' ' "* REVISED
relay. The Reserve Deploy relay fires the Reserve Chute Deploy Gum Squibs de-
ploying the Pilot Chute and the Reserve Chute Ejector Bag squibs. The Reserve
Chute Ejector Bag squibs activates the gas generator which has a one second delay
in ignition time before inflating the ejector bag expelling the Reserve Chute
and the dye marker. When the Rescue Aids switch is placed in the "MAN" position,
the Rescue Aids Switch Signal, and the Post Landing System Power Drop Hold relays
are energized. The energized Rescue Aids Switch Signal relay removes power from
the Red Rescue Aids Telelight and illuminates the Green light. The Post-Landlng
System Power Drop Hold relay energizes the Post-Landing System Power Drop 30
Second Time Delay relay. At the end of the 30 second time delay the Whip Antenna
Extend relay is energized firing the Whip Antenna Extend Squibs activating the
gas cartridge extending the active element of antenna to its full length. When
the Post Landing System Relay is energized on impac% power is removed from a
number of components in the circuit. After a 30 second delay the Post Landing
System Power Drop relay is energized removing power from the remainder of the
components, except the Whip Antenna Extend relay and the Rescue Aids Switch
Signal relay which leaves the Green Rescue Aids Telelight illuminated.
9-3- EMERGENCY SEQUENCE DESCRIPTION
The Emergency provisions of the landing system basically consists of
manually operated back-up systems initiated by the Astronaut. The appropriate
button, pull-ring, and switches are located on the Left Hand Console. The
emergency system controls manually initiate deployment of the Drogue, Main and
Reserve Chutes, extend the Landing Bag, and initiate rescue aids.
9-4. F._R_ENCT SEQUENCE OPERATION (See Figure 9-3)
On descending to 21,000 ft. and Drogue Chute failure is detected by lack
of opening shock and by a visual check through the window, depress the Drogue
/
DATE
REVISED
REVISED
1 NOVEMBER 1961 Mc'DONNELL____ST.LOUIS, MISSOURI
PAGE 9-7
REPORT SEDR 104
MODEL 133
i,i '
\\
F WHIP ANTENNA
PILOT CHUTE LANYARD
TO TOPOF PILOT CHUTE
\
EtGU!_E 9-2, _MAIN AND 12ESEI_VE PAI2ACI4UTE SYSTEM(SHEET I OF
i
PAGE 9-_8
REPORT SEDR 104
MODEL 133
MCDONNELL ST.LOUIS, MISSOURI
DATE
REVISED
REVISED
I NOVEMBER 1961
\\ . ',P . , \
\
'\
I)I_.o _'J E -._
_/SO_k__o_,_
"_--__ __--__ __-__RK_ _ __
FIGUP.E g-2 MAIN AND 12ESEI2VE PAP..ACHU'ITESYSTEM(_SN, EET Z OF Z) PM,_s--7_,_
}
l .......MCDONNELL 9-9DAT£
REVISED ST._" _'_'LOUIS 3, MISSOURI REPORT _D_:_ i(_
REVISED .... x MODEL 1BB
button. Depressing the button allows the Emergency Drogue Deploy relay to be
energized and the Drogue Chute Mortar squibs to be fired deploying the Drogue
Chute and dispersing the chaff. If the Green Main Deploy telelight fails to
illuminate, failure of the Main Chute deployment may be detected by a lack of
opening shock, a visual check and no decrease in rate of descent. Upon determin-
\
ing that the Main Chute has not deployed, operating the Main Deploy Pull Ring
energizes the Emergency Main Deploy relay, firing the Antenna Fairing Ejector
Squibs, ejecting the Antenna Fairing and deploying the Main Chute through the
normal automatic sequence. When the Green Main Deploy telelight is illuminated
and the rate of descent is above 32 feet per second, check the chute visually for
damage. If the chute is damaged or did not deploy, actuate the Reserve Deploy
Pull Ring which energizes the Reserve Deploy relay. Through the energized con-
tacts, power is applied to the _ Chute Disconnect firing the squibs disconnect-
ing the chute from the capsule. At the same time, the Reserve Chute Deploy Gun
Squibs are fired deploying the Pilot Chute, and the Reserve Chute Ejector Bag
squibs are fired generating a gas after a i second delay inflating the ejector
bag which aids in deploying the Reserve Chute.
Twelve seconds after the Main Chute is deployed the Green Landing Ba 6 tele-
light should be illuminated. If the light does not come on, place the _-d_ng Bag
switch in the "MAN" position, energizing the Emergency Landing Bag Extend relay,
firing the Emergency Landing Bag Valve Squibs, releasing the heat shield and
extending the impact landing bag. Ten minutes after impact the Rescue Aids Switch
By-Pass relay is energized. When activated, the relay by-passes the Rescue Aids
switch and energizes the relays that supply power to fire the Main Chute Dis-
connect squibs as well as the squibs for the Reserve, Disconnect, Deploy Gun, and
the Ejector Bag in the same manner as would be done if the switch was placed in
the "MAN" position.
C_::y:___:'T' AL
PAGE , ' 9-10,
REPORT SEDR 104
MODEL 133
Mc'DONNELLj ST.LOUIS, MISSOURI
_ ,,p_ m- I]P" II_" L I "_l" .
DATE 1 NOVEMBER 1961
REVISED ,, ,
REVISED
EMERGENCY
DROGUE DEPLOY
RELAY
£_ER_ENCX MM_
DEPLOY
RELAY
MAIN
CNUT _- FAIL5
"TO DF.PLOY C,R 15
DAN1/kG&O
.FAIL ¢' TO r=."_TEt,4D
RF..SCUr=A_05 sw.Nor _N "MA_"PO'_TION
FIGURE 9-5
R_VE F
_EPLOY
RELAYS
DP.OGU E CHUTE
MORTAP.
_CtUI B-_
CH_.FF __
q ncE
._J MAIN C_UTE,
E)I5CONNE.CT
SQ_I B._
__ _E._E_NE CHUTe.
0 _:I_LOY GU l,,t
5QLIIB5
____ _E._E I:l.Nt E CHUTE
EJ ECTOR
J
PILOT CHUTE
J --"-_ ELECTRICAL SI6klAL j
, MOVEMENT II
l _.MER_ENCy EM E R _ENC"('
._._.J_LAND_NG BAG L._N_INC_ _,,_C_
I "IEXTENO RELAY VALVE %q_IBS LANI_ING BAG,
/ _',,_ c_,u_-_ Ir--IN _sco_c'r I
_---_ - _¢ONN_.CT I
J I s_,_ I
IT _ DEPLOY Gu_ k\\\\\_ .>, . , °
EaECTOR _AG
LANDING AND P, ECOVERY EMERGENCY SYSTEM iSLOCK PM45-z6o_
• !
DATE 1 _ove_er l_61
REVISED
REVISED,
ST. LOUIS 3, MISSOURI
PAG _"
REPORT
MODEL
9-11
/ ",
"LI.J
9-5. SYSTEM COMPOHE_TS, , ,
9-6. PARACHUTE
The drogue parachute assembly (See Figure 9-5) consists of a conical ribbon-
type drogue canopy with integral r_ser, drogue deployment bag, drogue mortar,
sabot, chaff packet, and drogue mortar cover. The drogue parachute canopy t s a
conical ribbon parachute having 8 gores of 2-inch wide, 460-lb. tensile strength
ribbons and 8 tubular nylon suspension lines of 1,O00-lb. tensile strength each.
The parachute is constructed to a diameter of 6.85 feet end permanently reefed
(restricted) to an effective diameter of 6.0 feet by means of pocket bands. The
constructed total porosity is 27._ and the effective porosity (through reefing)
is 36.3%. The SO-ft. long integral riser is made from three layers of 3,000-Ib.
tensile strength low-elongation hot-stretched Dacron webbin 6. The drogue para-
chute stabilizes and decelerates the capsule. The canopy weighs 2.9 lbs. without
riser and 5.9 lbs, includi_ the 30-ft. Dacron riser. The drogue parachute de-
ployment bag serves a dual function of (i) protecting the drogue parachute during
ejection and (2) providing means for orderly deployment of the drogue parachute.
The bag is manufactured of cotton sateen fabric reinforced with nylon webbing and
covered at the upper end with a heat insulator of glass cloth. The bag is
weighted at the upper end with a 0.5 lb. lead disc which assists in stripping
the bag from the canopy at the completion of line and riser stretchout. Inside
of the bag are _ cotton tapes to which the riser is secured during packing in
order to provide orderly riser deployment. The mouth of the bag is closed with
a light cotton cord.
9-7. m0uuE CN_E M_a A_D _
The drogue parachute ejection mortar is a device for positively deploying
the drogue parachute with sufficient energy to overcome local pressure gradients
and gravitational forces. The drogue parachute is packed in a protective bag
-C__":".":DZ;_T; A:.
PAGE
REPORT
MODEL
9-12
SEDR 104
133
McDONNELLST.LOUIS, MISSOURI REVISED
_ I_ I_r'__ ! _ I1_'_._..r_!"_P Lz.__.., REVISED
1 NOVEMBER 1961
I, ANTENNA FAIRING EJE'CTOR CBUN
2. PILOTCHUTE D_PLOYMP-N'T FdJN
5. PILO-r CHUTE
4. _OFAR E_4_ ( 2 )
5. SEA MARKER
6. SHARK REPELLANT
7.RF_ERVE CHUTE ANO BA_
8. EJECTION BAG_ (2) ,
9. GAB G_NERATO_S L2)
I0. INERTIA %WITCH
• fl. 5EQuENCE CONTROLLER BO×ES(,Z) /
12. c'_JRVIVAL KIT I
f3. MAIN CHUTE ANO _AG I
14. EIAROSWITCHE3 (4) I
t5. ANTENNA LANYARD I16. CHAFF PACKET I
t7. RECOVERY LIGHT I
I& WHIP ANTENNA I
19. PARACHUTE. OlSCONNEC'r_(2) !
20. DROGUE I MORTAR I II I_
NOTEt CAPSULE 8 ONLY
FIGURE 9-4 LANDING AND RECOVERY SYSTEM INSTALLATIONPM45-74C
pj_l_llpIlitpIl--- - ---vi I In il_ll_I1 i I_ L
DATE,, ,1NOVEMBER1961
REVISED
REVISED
Mc'DONNELL_'J_.___ST.LOUIS, MISSOURI
PAGE
REPORT .....
MODEL
9-1'3
SEDR 104
133
FIGURE 9- 5 DROGUE PARACHUTE
"_ L ' e"s r'_P"_'i"PI A |
PAGE 9-I_
REPORT S_I_ i04
MODEL_ 133
REVISEDST. LOUIS, MISSOURI
._,,_,T.T_ _._Tr_Z_JL,=-- _ REVISED
and stowed in the mortar tube on top of a lightweight sabot (See Figure 9-6).
The sabot functions as a piston to eject the parachute pack 2 when pressured from
below by ga se s generated from a pyrotechnic charge. The propellant charge is
initially fired into a breach chamber of small volume, to produce high pressure
which is subsequently vented through a small orifice and into the main chamber
at relatively lower pressures. In this manner_ reaction loads are kept to a
minimum, since the pressure energy is _ot expended instantaneously. The pressure
sealing quality of the sabot is derived from an "0" ring t installed in a groove
near the base. Two small holes are located in the "0" ring groove to vent air
trapped in the mortar tube underneath the sabot on installation. For proper
operation, the "0" ring and the inner wall of the mortar tube_ which is always
in contact with the "0" ring, are lubricated before installation. The drogue
parachute pack is retained in its stowed position within the mortar tube by
a thin metal (Rene "_1) cover which is attached to the upper surface of the
antenna housing. Three cut-out sections, provided in the sides of the cover,
permit routing of the steel cable risers into the drogue chute can. The cover
is designed to constrain the chute in its compartment against negative decelera-
tions end also to require minimal forces to break loose from its attachments at
the time of deployment. Pressure of the chute pack causes the cover to deflect
such that attachment tabs pull out from under attaching screw heads through a
slotted hole designed for this purpose. The energy required to expel the drogue!
chute from its compartment is provided frcm high pressure gasest generated by
ignition of a pyrotechnic charge. The cartridge is loaded with 66 grains of
powder_ contained in a propellant can attached to a steel body which houses the
ignition wiring, and teminates in an electrical connector. The ignition cir-
cuitry consists of two separate and individual bridges, either of which is
capable of igniting the Lxr_der charge upon application of the proper current.
DATE
REVISED
REVISED
1 NOVEMBER 1961
ST.LOUIS, MISSOURI
PAGE
REPORT
MODEL
9-i5
SEDR 104
133
\
1. ['O" RING
2. ',SABOT
3. ICHAFF PACKAGE
4. INSULATION
5. ICOVER
6. ',C_TRIDGE
7. [CHAMBER
FIGURE 9-6 DROGUE CHUTE MORTAR ASSEMBLY
PAGE,, 9-16
REPORT _ I0_.
MODEL. I_B
9-8.
MCDONNELLREVISED
ST. LOUIS, MISSOURI
_'_ Ik.Jck,_l 41, ,ILU.IL,4J._I A
i November 1961
REVISED .....
CHAFF PACKET
The chaff packet is a locating aid which performs its function by distribut.
ing fine divided metal foil having radar reflecting capability. The foil strips
act as miniature antennae and are capable of providing an average echo area of
600 square feet, covering the S-_ L- and C-band radar frequencies. The package
is ejected on deployment of the drogue parachute. The chaff is packed in a
cardboard package of rectangular shape with dimensions approximately i x 3 x 5
inches. When Jettisoned into the airstreem, the package spills open and dis-
perses its contents.
9-9. _._ P_,_Jl_l_.
The main parachute assembly consists of: Main parachute canopy, riser,
deployment bag, and parachute disconnect. The main parachute canopy is a 63
foot nc_inal diameter ringsail type. The ringsail parachute is a slotted canopy
similar in design to the ringslot parachute. The parachute is fabricated from
2.25- and 1.1-ounce per square yard nylon parachute cloth into _ gores with
suspension lines of _50-pound tensile strength. The main parachute is
packed in a deployment bag which provides a low snatch force and orderly deploy-
merit (See Figure 9-7)- The bag is manufactured from cotton sateen fabric, re-
inforced with nylon webbing and covered at the upper end with Thermoflex and
glass cloth insulation. Inside the bag, midway along its length, is a pair of
transverse locking flaps. Their function is to separate the canopy fabric from
possible entanglement with the lines and to cause full time stretch-out before
canopy deployment.
9-10. PARACHUTE DISCORRECT
Both main and reserve parachutes are attached tO the capsule by a device
designed to sustain the parachute loads during descent and to disconnect the
J
_L'VtSEO ST. LOUTSa, M,SSOUm nZPORV _ 10_
Ev,s=o Moo . 133
parachute on ground impact (See Figure 9-8). The disconnect function is neces-
sary to prevent capsule upset or damage by dragging in surface winds after
touchdown. The assembly consists of 5 separate details installed in a mounting
structure which is an integral part of the capsule. The parachute riser is
looped around the arm which transmits the load to the structure through the
piston. The shear pin restrains the piston from any motion tending to displace
it. On ground impact, an electrical impulse from the inertia switch reaches the
squib cartridge causing it to fire. The gas pressure, thus generated, forces
the piston forward into the arm recess, cutting the shear pin in the process.
Full displacement of the piston removes parachute load transmission to structure,
allowing the arm to rotate around the pivot pin. The loop of the parachute riser
slips off the arm and the disconnect function is complete. The lead buffer
serves to absorb energy of moving piston and prevents rebound of the piston
back into the locked position.
9-ii. RESERVE PARACHUTE
The reserve parachute assembly consists of: The pilot chute deployment gun
and lanyard, pilot parachute, reserve parachute canopy, reserve parachute deploy-
ment bag, and reserve parachute disconnect. The reserve parachute deployment
bag is similar to the main parachute deployment bag with the addition of flaps
at the upper end of the bag to contain the packed pilot chute. The reserve
parachute disconnect is identical to that used to disconnect the main parachute.
The reserve parachute canopy is identical to the main Parachute canopy.
9-12. PILOT PARACI_
The pilot parachute is a flat, circular type, 72 inches in diameter with a
30 ft. bridle. It is manufactured of 3.5-ounce per square yard fabric in the
canopy and 2.25-ounce fabric in the vanes.
PAGE
REPORT,
MODEL
9-1S
SEDR 104
133
Mc,DONNF,LLLL DATE 1 NOVEMBER 1961
REVISED,,
REVISED ,
.FIGURE 9-'7 MAIN PARACHUTE AND PACKINGBOX
. s_UCTUR_ s.s._R PIN
SHORTING WIRE 6. PISTON
',31 SQUIB CARTRIDGE 7. LEAD BUFFER
• BUSHING S. ARM .
FIGURE 9-8 MAIN AND RESERVE CHUTE DISCONNECT
J
\, J
DAT _" i November 1961
REVISED.
REVISED
ST. LOUIS 3, MISSOURI REPORT _ lO_
MODEL 133
9-13. PILOT CHUTE _PLOYMEBT GUN
The pilot chute deployment gun (See Figure 9-9) initiates the first step
in the sequence of reserve parachute deployment. Either gas pressure or an
electrical impulse will cause the gun to fire, thus expelling a 12-ounce pro-
Jectile to which is attached the reserve parachute pilot chute. The pilot chute
inflates and in turn pulls out the reserve landing chute, ccmpleti_ the sequence
Whether fired electrically or pneumatically, a one-second time delay is provided
between receipt of the Impulse and detonation of the main charge. This delay
permits the main parachute (if deployed and damaged) to separate from the capsule
to avoid entanglement with the reserve parachute to be deployed. The gun is
basically a tubular body which contains the main firing cartridge and the pro-
Jectile assembly. The projectile assembly is held in place by a pin which is
sheared when the projectile is expelled. The main cartridge, which generates
the gas pressure to eject the projectile, is fired as follows: (i) Gas pressure,
through the gas flrln 6 mechanism (supplied when RESERVE-PULL-RIX} is operated),
drives a firing pin into the primer cap at the base of the main cartridge,
initiating a time delay train, causing a subsequent detonation of the charge.
A minimum of 750 psi gas pressure is required for pneumatic operation. (2) An
electric impulse is received at the time delay igniter installed through the
side of the gun. After a one-second delay, the igniter fires through the wall
of the main cartridge and detonates it instantaneously. Firing characteristics
of the igniter cartridge are as follows : All Fire Current 2.5 emps per bridge,
All Fail Current 0.5 amps per bridge. The ignition circuit consists of two
individual bridges terminating in a _-pin receptacle. _tzzle velocity of the
projectile is 2r)o-_X) ft/sec.
9-i_. PARACHUTE EMPTOR
The ejector bags are inflatable air cells made of lightweight rubberized
vva_a I_a_a Jd'TKJhne
PAGE
REPORT
MODEL
9-201
SEDR 104
133
M 'DONNE ST.LOUIS, MISSOURI
DATE 1 NOVEMBER 1961
REVISED
REVISED, ,
II
FIGURE 9-9 PILOT CHUTE DEPLOYMENT GUN
FIGURE 9-10 PARACHUTE EJECTOR BAG
J
DAVE I November 19_i
REVISED
REVISED.
nylon fabric (See Figure 9-10). The design inflated shape is that of a cylinder 3
ii inches in diameter and approximately 35 inches in height. The upper end of
the bag is slanted at full inflation to promote Jettison of the parachute pack
overboard on landing impact.
9-15. PARACH_ EJECTOR GAS GEaR
This is a device to provide a rapid and sufficient volume of gas to inflate
the main and reserve parachute ejector bags (See Figure 9-11). The reserve
parachute gas generator is similar to that used for the main parachute except
the additional feature of one second delay in ignition time. The generator
functions to produce gas by the relatively slow burning of a solid powder pro-
pellant in the main chamber. The gas is directed from the main chamber into
the ejector bags through a 3/8 in. diameter stainless steel tube. The tube
serves also as a heat exchanger to reduce temperatures to within tolerable
values prior to entry into the ejector bag. The generator body is equipped with
lugs for mounting to the parachute container with four bolts. Ignition circuit
characteristics are as follows: All Fire Current 2.5 amps, All Fail Current
0.5 ups.
9-16. SOFAS
A post-landin_ recovery aid. SOFAR is an abbreviated form for "sound fixing
and ranging" (See Figure 9-12). This component performs its function when it
detonates by hydrostatic pressure at a predetermined water depth. Shock waves
from the explosion are received by sound detection devices sboanl picket ships
or shore bases and a position fix on the capsule is thus made. The
range of the Mercury SOFAR bombs is 3000 miles. Two Bombs are carried aboard
the capsule; one set to detonate at _ feet depth and one set for 3500 feet.
The b_b weighs approximately 2 pounds. One 3500 foot bomb is tossed over-
_ v i Ii I !11 I i ....
1 November 1961
REPORT _ 10_ REVISED-.4;T_LOUI$ ,_,MISSOURI
MODEL 133 ,.,v,,.-.,.uL_,, .... REVISED.
board at main chute deployment. The second SOFAR bomb remains with the capsule
and only serves to notify the search party that the capsule is sufficiently
(See Figure 9-13below the water's surface to render it non-recoverable.
for Operation. )
CAUTION
The bombs are relatively safe to handle, but it should
be remembered that they are high explosive devices and
hence common sense precautions should be practiced.
9-17. DYE MARKER PACKET
The dye marker packet (See Figure 9-14) is a post landing recovery aid
which performs its function by dissolving in water, thus producing a highly
visible yellow green patch. Approximately 1 pound of fluorescein dye is packed
into a soluble plastic bag, which in turn is packed into an outer aluminum con-
tainer. The entire packet assembly is ejected overboard, at the time of reserve
chute ejection. The fluorescein dye forms a spot on the ocean surface which is
visible from an airplane i0,000 feet high at a distance of i0 miles on a clear
day.
CAUTION
The dye marker package should be stored in a dry
place and not be exposed to water in any way.
9-18. _ECO_ LIGHT
To aid in the visual location of the capsule al_er landing, a flashing
light is installed in the recovery compartment. The intensity of the light is
such that it will be visible in normal darkness for _O nautical miles and up to
an altitude of 12,000 feet. The flashim_ rate is approximately 15 flashes per
minute. Powered by self contained, dry cell batteries, the light's circuit will
f_ g'a w.v r,_,','_ _ ------- •_ v J. , A A A.F A_ J. , £ £ ._A. JL
DATE
REVISED
REVISED
1 NOVEMBER 1961 M*,DONNE ST.LOUIS, MISSOURI
PAGE
REPORT
MODEL
9-23
SEDR 104
133
I 1 . ELECTRICAI_ CONNECTOR
2. MAIN CHAMBER
3. PROI_ECTIVE CAP
4. TUBE FITTING
5. ATTACHMENT LUGS
6. SHORT!NG PLUG
FIGURE 9-11 MAIN AND RESERVE CHUTE GAS GENERATOR
FIGURE 9-112 SOFAR BOMB
(,..m._ ...... _"1 A !VVIll iJb/Vllllll • ia 'l.I
PAGE 9-2_4
REPORT $EDR 104
MODEL 13 3
McDONNELL ST.LOUIS. MISSOURI
• •,- . ,_,eat#l
DATE
REVISED
REVISED
1 NOVEMBER 1961
/
OPERIqTIO_q
DETONRTION OFTHE UAIN CHARGE 15RCCOMPLISHEDIN TWO ST/::IGE,5:
WATER PRE.%_URE ON .SURFACE ",q'CREATES /_FORCE 3UFFICIENT TO BREAK SHEAR PIN',q"PERI_IITTING THE INTERRUPTER BLOCI_ TO.t_OVE UPW/_RD /_GRIk/_TTHE 3TOPPING
SHOULDER. WHEN IN THIS POSITION_ THEPRIMEg CHRRGE 13 IN LINE WITH THE FIR-ING PIN.
B. W/_TER PRE53URE ON SURFACE 'B' CREATES
FORCE TO BREAK SHEAR PIN 'B' ANDDRIVE THE FIRING PtN INTO THE PRIMERCHRRGE. THE PIPIMER CHRRGE BLASTS IN-TO THE BOOSTER CHFIRGE VIR THE LERDIN ORIFICE_ fiND THE I_OoSTER CHARGECRUSE3 THE MPdN CHRRGE TO DETON-ATE, STRENGTH OF SHEAR PIN "B' IS PRE-DETERMINED FOR DESIRED DETONFITIONDEPTH. STRENGTH OF 5HERR PIN "R' I_SUCH THRT IT _IILL SHEAR RT A DEPTHOF RPPROXIMATELY ONE H/:ILF THE DEPTHREQUIRED TO 5HERR Plkl "B'.
SHEFIR PIN
SURFACE
DETON _TO R
LERD IN
BOO3TER C
MAIN CHRRGE
(O.? LB. HBX
FIGURE. 9-1.5 SOFRR BOMB 5CHE_FITICPM45-16;
/n_/_ILIPiii_IpI, I'q_u- - -=
_J
_J
R_VIS_D s'r. -ou,s -_.MISSOUR, nm=ORV. _DR 10_
R_VlSED _ MODe. 133
, f::\,
be closed through an energized contact of the Post Landing System Relay which
is activated by the closing of the Inertia switch on impact and energizing
the Impact relay. The light will operate for approximately 28 hours.
9-19. WHIP A_rE_mA
To provide operation of the HF voice receiver-transmitter and HF recovery
beacon, a Whip Antenna is used. The active element is stowed in a collapsed
condition in the recovez-j compartment and when extended is approximately 16
feet long. The antenna is extended by a gas cartridge which is activated when
the Post-Landing System Power Drop 30 Second Time Delay relay energizes the
Whip Antenna Extend relay. When it is extending a galling action takes place
between the segments of the active element holding it rigid in the extended
position.
9-20. BAROSWITCHES
There are two pair of Baroswitches used in the recovery system (See Figure
9-16)' In these switches an over-center spring is included in the design to
minimize chatter during vibration and shock and to prevent contact oscillation.
One pair of switches is set to close at 21j000 feet and the other is set to close
at 10,600 + 750 feet. The switches are located in the recovery compartment where
they are connected to a plenum chamber. The plenum chamber collects ambient air
pressure through four static pressure ports equally located around the recovery
compartments outer surface.
9-21. INERTIA SWITCH
The inertia switch is essentially a spri_ device actuated by mass (See
Figure 9-17). A landing shock of 7.5 plus or zLtnus 1.13g's _ will produce
momentary closing of two electrical contacts, thus completing an electrical
circuit. This switch is used in conjunction with a latching relay which receives
• m ! P'_ Im'lk |_lp, | I!
_it_llwa Ii aml,Na 11 • ia ._r--=--.
PAGE.
REPORT
MODEl,.
9 -26
SEDR 104
133
Mc'DONNEL LST.LOUIS, MISSOURI
--_ L. a. :--_, I it1, ir, k ! ,,I- !. ,i_jm_
DATE ,,
REVISED
REVISED
1 NOVEMBER 1961
FIGURE9-14 DYE MARKER PACKETP_ 4-5-i?O A
a,.._ a,._ _. p. ..... ---v,._-i"li aL/r---I_l I I/_L.,.
_AT_ 1 November 1961
REVISED
REVISED
PAG- 9-27
sT. Loum 3,MlSSOUm RZPORT _ IC_
__C__ _!V_ MOnm. 133
an electrical pulse and, by latching into a latched position, provides continuous
electrical continuity. The inertia switches used consist of four Selmrate snap-
action switches and two separate masses, all housed in a common case.
9-22. TEST CONFIGURATION CAPSULES
9-23. TEST CONFIGURATION CAPSULE NO. 8
The normal mission _equence and escape sequence for Capsule No. 8 differs
from the Specification Compliance Capsule as described in the following para-
graphs.
The reserve parachute is replaced with a flotation pack. This flotation
pack will be attached to the main parachute in such a manner as to act as a buoy
after main chute disconnect to enable recovery of the main parachute. Three dye
marker packets are installed in the antenna, fairing, which remains attached to
the main chute. These markers will assist in locating the floating main chute
and sn_nna fairi_.
9-2_. Operation
The capsule's landing system is armed by 28 volt d-c power at time of escape
tower separation. Both isolated and main battery power circuits are applied
throu6h the de-energized impact relay No. 1 located in the No. 1 recovery relay
box, through the de-energized No. 2 and No. 3 orbit attitude relays and the de-
energized No. 2 and 3o. 3 tower separation sensor relays located in the No. 3
launch and orbit relay box, to the two 3 second time delay relays located in
sequence controllers, units A and B. After the 3 second time delay both relays
are energized and complete two separate circuits to the four baroswitches. The
Baroswitch contacts will be open due to altitude being in excess of 21,000 ft.,
and therefore the power circuit will stop at the open contacts of the switches.
Upon re-entry from orbit, at an altitude of 21,000 feet the two 21,000 ft. baro-
switches are actuated and the drogue parachute is deployed and a chaff package
PAGE
REPORT
MODEL
9-28
SEDR 104
,133
Mc'DONNEL_L_ __._....ST.LOUlS o MISSOUR,/
DATE
REVISED
REVISED
1 NOVEMBER 1961
FIGURE 9-15 BAROSWITCH
FIGURE 9-16_ INERTIA SWITCH
r
! • r
"_±j
DATE 1 November l_l MCDONNELL,,_' '...7_ PAG_ 9-29ST. LOUIS 3. MISSOURI REI=ORT. _ 10_
Moola. 133i ,,
is ejected. The drogue chute stabilizes and decelerates the capsule; the chaff
package disperses finely cut metal foil. As capsule descent passes throu_ the
I0,000 ft. level, the two remaining baroswltches will close causing dual circuits
through the small pressure bulkhead disconnects to the system A and B sequence
controllers. These two circuits energize a main squib short (auto) relay within
each controller unit, resulting in the removal of ground circuits for all four
squibs of the antenna fairing ejector. At the same time, and by the same l_wer
that energized the relays, power is applied to the pre-grounded circuits of the
four antenna fairing squibs through the de-energized main over-ride relay. As
the circuits are completed to the four squibs, two other branch circuits are
applied to the two antenna fairing separation sensor limit switches. The firing
of the four squibs causes the antenna housing to separate fr_a the capsule. A
lanyard, connected frcm the antenna housing to the main chute, extracts the
main chute from the main chute c_npartment. The separation of the housing from
the capsule allows two antenna fairing separation sensor limit switches to spring
to their actuated position. Through these two double pole, actuated limit
switches, three signals are returned, two entering the system A and B sequence
controllers where they start two 12 second time delay relays, and the third
signal entering the No. 1 recovery relay box where it energizes the four antenna
fairing separation relays. Through the energized No. 3 antenna fairing separa-
tion relay, in the No. 1 recovery relay box, a 28 V si_ml is completed to the
Main Deploy telelight, illu_tuating the light Green. Power is also applied
through the contacts of the No. 3 antenna fairing separation relay to fire the
two Main Chute E_ector Bag Squibs. The firing of the squibs allow the eject
bag to be gas inflated end therefore, simultaneously eject the main chute as
the chute is being extracted by the antenna fairing lanyard. At the run-out of
the two twelve second time delay relays, the relay coils are energized and sa-maing
, _= lli,==llT=_ .......wwa., alllvL--i_l ! I¢'_ill.-.
PAGE 9-30
REPORT _E_R 104
MODEL 1"_:_
REVISED---t_r..muls,..Mlssg.uR_,[._,
REVISED
the contacts of the inertia switch. At the time of landing the inertia switch
will sense the impact and complete two 28 V signals to the two inertia switch
slave latching relays located in system A and B sequence controllers. These
two relays when energized will in turn energize the two reserve and main Jettison
relays In their respective sequence controller boxes. The energized reserve
and main Jettison relays will fire the two squibs of the main chute disconnect
thereby releasing the main chute from the capsule. The reserve chute disconnect
will not be fired at this time. The main battery power circuit that energized
the system A reserve and main Jettison relays through the inertia switch slave
relay is now continued out of the system A sequence controller to the No. I
recovery relay box. This is the first time in the overall sequence that a dual
redundant circuit is not provided. This 28 V circuit to the No. i recovery
relay box passes through this relay box to the No. 2 recovery relay box where
it again passes through the de-energized air shutoff relay and out of the No. 2
recovery relay box after branching into two circuits. One circuit is applied
directly to the Astronaut's rescue aids switch and the other is returned to
the No. I recovery relay box where it energizes the four impact relays. Through
the energized No. i impact relay the self-powered recovery flashing light circuit
is closed starting the flashing light to operate. Through this same relay, power
is removed frcB the two 3 second time delay relays and the Main Deploy telelight
will go out. With the rescue aids switch clo_ed, a 28 V circuit is completed
to the No. 1 recovery relay box where it energizes the rescue aid relay. Through
the energized relay the Torus Bag Deflate 5 second time delay relay is energized
and starts timing. Power circuits to the two squibs of the reserve chute dis-
connect, and also to the two squibs of the reserve chute deploy gun and the two
squibs of the reserve chute eject "bag are-cc_rpleted. Simultaneously_ the reserve
chute is discounected_ the reserve pilot chute is deployed and the eject bag be-
.... _ : , -.: _. _-_ _-2L_-
DATE
REVISED
REVISED
1 NOVEMBER 1961
ST.LOUIS, MISSOURI
PAGE
REPORT
MODEL,
9- 3_1
SEDR 104
133
FIGURE 9-117 SEQUENCE CONTROLLER SYSTEM A AND B
REPORT _:_ 1011" REVISEDEl'. LOUIS, MISSOURI
MODEL. 133 • _ REVISED
neath the reserve chute is inflated forcing the chute out. After the runout of
Torus Bag Deflate 5 Second Time Delay relay the Whip Antenna Extend relay is
energized allowing power to flow to the whip antenna squib. The energized Gas
Cartridge extends the active element its full length.
9-25. Emergency Sequence
9-26. Description
The landing and recovery manually operated emergency system is the same as
the Specification Compliance Capsule but Capsule No. 8 is unmanned.
9-27. Components
9-28. Sequence Controller Assembly
There are two landing and recovery system sequence controllers located
within the pressurized section of the capsule (See Figure 9-18). These sequence
controllers in conjunction with three relay boxes accomplish all the system
sequencing and provide other capsule systems with initial connnands. The sequenc_
controller assembly contains the relays_ fuses_ and timers needed to accomplish
deployment of the chutes in proper sequence.;
9-29. TEST CONFIGURATION CAPSULES NO. 9 AND 13
Basically the same as Specification Compliance Capsule.
9-30. TEST CONFIGURATION CAPSULES NO. i0 AND 16
Same as Specification Compliance Capsule.
iJ
- 10-I
H
SECTION X
ELECTRICAL POWER AND INTERIORLIGHTING SYSTEMS
TABLE OF CONTENTS
TITLE PAGE
ELECTRICAL POWER SYSTEM
SYSTEM DESCRIPTION .....................
SYSTEM OPERATION ......................
SYSTEM UNITS ...............................
TEST CONFIGURATION CAPSULES.
_° ..................
............... °....... ...........
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10-3
10-5
10-11
10-12
INTERIOR LIGHTING ............................ 10-27
SYSTEM DESCRIPTION ..................... 10- 27
TEST CONFIGURATION CAPSULES. 10-30
_ A i Illll_ll'lZlllWll_l • I
_li_ V I I ! I liz_ lira I 1 i IN_l lira
7I
A
DIODE POW|
5WITCH FOSE PANEL
/ t/
i
\
[1_ C,t_PSULE16 NASI_00 WATt- HOUP_15_,TTE P..IE5 INTHESE LDCATIOI4 S
--- 1500 WAT THOUR BATTERY
BATTERYON*OFF
5',N ITCH
X REMOTE
7\
I
DETAIL'A"BATTEP._' ON'O.CF 5WITCH
ZSO YA _N_/ERTER5
PAklEL
OETAIL'C =
I_ELAY5
\\5NUNT
"" DE.TAIL'B'
__ ELF_CTRICAL POWEI_ CONTROL 'SECTION
(MAIN IN_'t'RCJMF-.NT _AI4EL)
/ Ii
/
"_ DE.Th, I I. _'C"
i:U 5E HO"H"_DEI:_'(TOP TO BOTTOM) /
_I P,E.'I'ROGRADE _P.ECO_ERY CSA) 1 ' / /
COMMUk_CATION5 (SA) _ \- l /
LAUNCH, 0_T _' ESCAPE (SA) --
MISCELLANEOLI5 (_0 _ ZS/_I 05_, ACCELEg, OMETEli:Itv_L5 C E LL A NEOLI'5 (_A)
I_ 0
m m
REVISED ST. LOUIS 3, MISSOURI REPORT _ 104
RSVISED _ MOO,=. 133
X. ELECTRICAL POWER AND INTERIOR LIGHTING SYSTEMS
I0-I. ELECTRICAL POWER SYSTEM
10-2. SYSTEM _ESCRIPTION
The capsule power supply consists of three main batteries, two standby
batteries and one isolated battery which supply 6, 12, 18 and 24 volts d-c
through control circuitry to the d-c power distribution busses. Distribution
is accomplished through high priority and low priority busses. Low priority
busses may be switched off to conserve battery power. External d-c power is
supplied through a diode panel prior to launch. (See Figures I0-i and 10-2. )
The capsule ammeter is used to indicate bus current when capsule power is
applied. Capsule bus voltage may be read on the capsule d-c voltmeter when
external power or battery power is supplying the busses.
The d-c electrical loads are supplied through fuses, except the "Abort"
Control which incorporates a solid conductor in place of a fuse. Some of the
d-c circuits utilize two fuses in parallel for redundancy, with a two position
switch which permits operation in the normal (No. i) Position or the emergency
(No. 2) Position. In addition s to prevent possible loss of power during an
emergency, a solid conductor is installed in place of a fuse in the emergency
(No. 2) position of the following circuits.
(a) Emerg. Capsule Separation Control
(b) Tower Separation Control
(c) Emerg. Main Chute Deploy
(d) Retro. Jett. Control
(e) Retd. Manual control
(f) Reserve Chute Deploy
(g) Emerg. Reserve Chute Deploy
.... _ .... ---_a 1row A |
PAGE
REPORT
MODEL
10-4
SEDR 104
, 133
, ,0
M 'DONNELL: , l" _'_ST_LOUIS. MISS_UR| _"
F " =""lmm_V a I u Jb,r _
DATE
REVISED
REVISED
1 NOVEMBER 1961
i RECOVERY IRELAY PANELNO. 2
I I_ELAY PANEL NO. 1III
,-T'_E ._s I_;IiS'_E " ', I_"__ I
I AF "_ it OI sQutts_ I
ARM _ JI RELAY --I1=.L
_40"T EE> t5oo w. _, _
C.APS. _G
r.._u _ _ _ ................
981
24V MAIN lPRE-IMPAET + I0 M.
9814.1 i
S4, i
94-1qq _
imo_
UI_glLIChLDISCONNECT
IS01
_us "IISOL
II,,
24- I_ *b -
;ISOLATED _ATT EEYIcjO0 W-H
FIOURE I0-2 D-C POWER CONTROL SCHEM_IE (CAPSULES IG,I_ 19)PM45-_ F
i_,,i p ili.i IE. 1 i,,,,,I p i _ _ - =...... .,_=,..= _= I IJ"=ll, IL-
The a-c loads are not fused because of Luherent overload protection in the
imverters.
The a-c power is obtained by using two main imverters and one standby
inverter to transform battery voltage to I15 volt, 400 cycle, single phase
a-c. The a-c voltage is fed through control circuitry to the a-c distribu-
tion busses.
10-3. SYSTEM OPERATION
10-4. D-C POWER CONTROL
The main 24 volt d-c power supply consists of three 3000 watt-hour
batteries. (See Figure 10-2). Each battery is connected in parallel to the
circuit by an ON-OFF switch on the battery. Individual reverse current diode
protection prevents discharge through a faulty or low battery. The 24 volt
d-c power from the three main batteries, or from the external power source
prior to launch, is supplied to the main d-c bus. The main 24 volt d-c bus
is connected directly to the 250 and 150 volt inverter filter inputs.
Two 1500 watt hour standby batteries are installed in the capsule to
supply communications bus voltage and to act as standby power for the main
system. The standby batteries incorporate diode reverse current protection
and are connected to the circuit through ON-OFF switches located on the
batteries.
Standby-battery taps, through diodes, supply 6, 12 and 18 volts d-c to the
various system busses. Prior to launch these circuits are energized by external
power through the umbilical disconnect and external power diode package. Stand-
by battery 24 volt d-c power application is controlled by the STDBY BATT.
switch on the main instrument panel. This switch may be placed in the OFF or
ON position. In the OFF position, the main batteries supply 24 V d-c power
to the capsule electrical systems through the main bus. In the "ON" position,
T__ ._:.": D -";.T :A L
REPORT _R1_ 104 "'t'IP._'tOUt$_'MI$$QURL._ REVISED
_"" _T_TL-" E _T_I_dIL_I_ REVISEDMODEL 1_3 _...
the standby 24 V d-c bus is connected directly to the main bus and both battery
groups supply power to the capsule systems in parallel. (See Figure 10-2).
An emergency hold circuit is utilized in event a HOLD command is initiated after
umbilical separation. To reduce cabin heating and provide external observation
through the periscope the emergency hold circuit removes power from the non-
essential circuits and also applies power to the cabin vent squibs and to the
extend motor of the periscope. Circuit switching is accomplished as follows:
The hold signal from the booster is applied through normally closed contacts
of the ground test umbilical relay to the solenoid of the No. 1 emergency hold
relay. Power from the main d-c bus is then applied through the normally open
contacts of No. 1 emergency hold relay to the solenoid of No. 2 and No. 3
emergency hold relays. Actuation of these relays_ removes power from some of
the nan-essential circuits, and energizes the impact sensor relay. At the
same time, power is applied through the normally open contacts of the emergency
hold relays to extend the periscope, and fire the cabin vent squibs.
A 1500 watt-hour Isolated battery is installed to provide emergency audio
bus and squib firing voltages and to supply emergency voltages to other circuits
in event the main and standby batteries are depleted. The isolated battery also
incorporates reverse current protection and is connected to the isolated bus
through an ON-OFF switch. Isolated battery taps supply 6 and 18 volts d-c
through self-contained diodes, to selected system busses. The 24 volt isolated
battery output is available through the ARM position of the squib arming switch
and through the EMERG position of the AUDIO BUS switch to the associated busses.
The Isolated 2_ volt d-c output may also be connected in parallel with the 24
volt output of the standby batteries through the STDBY position of the ISOL-BTRY
switch.
External d-c power is supplied through the umbilical cable to capsule
"k
:L ¸ i
_./
DATE 1 November 1961,I
REVISED
REVISED,
ST. LOUIS 3. MISSOURI REPORT ,_..T'_ i0_
Moo,,,, 133
circuitry. This power is used for pre-launch operations in order to conserve
the capsule battery supply. Normally 6, 12, 18 and 24 volts d-c are supplied
through seven inputs to the external power diode panel, however, only four
external power supplies are used to supply the 6, 12, 18 and 2A volt require-
ments. A d-c voltmeter and selector switch permit the Astronaut to read
individual battery and main bus voltages as desired. The d-c ammeter is used,
with the as_eter switch in the NORM position to indicate d-c load current from
the batteries in the circuit. A zener diode panel is incorporated connecting
zener diodes between ground and each of the Main and Isolated 24 d-c Busses
which provides protection for the transistorized equipment from transient spike
voltages (See Figure i0-2).
10-5. A-C POWER AND CONTROL
10-6. Main Inverter
Main 115 volt 400 cycle a-c power is supplied by two inverters of 150 volt-
amperes and 250 volt-amperes. The a-c load is divided into two groups namely
the ASCS bus and the FANS bus. The 250 VA inverter supplies the ASCS bus and
the 150 VA inverter supplies the FANS bus. The main d-c bus Powers the 150 VA
(fans bus) inverter through a filter circuit and a 25 ampere fuse. The 250 VA
inverter (ASCS bus) is also powered from the main d-c bus through a lime filter
and 25 ampere fuse. The d-c power is controlled through the NCRM. position of
the ASCS, AC B_s, and Fans AC Bus switches on the Main Instrument Panel. The
outputs of the FANS and ASCS inverters feed the solenoids of the Fans Bus Relay
and the ASCS Bus Relay. These energized relays feed the inverter output through
the closed contacts of the relays to Power the fans and ASCS busses. (See
lO-3.)
10-7. Standby Inverter
Standby IISV _ cycle a-c power is supplied by one 250 volt ampere standby
.... --;'-"''" a lv sis _aml _ a .l_m
PAGE
REPORT
MODEL
10-8
SEDR 104
133
M_,DONNELL___,_,_ (_,_,_......... ST.LOUIS, MISSOURI
DATE , ,
REVISED
REVISED
1 NOVEMBER 1961 ,,
[ MAIN INSTRUMENT PANEL!
FAt4S •
T_ SCS
SELECT CWITCH
_ 1Z/26VDC
• LT (AUTO) |prL
------otl
o
STBYOFF
r--o* 1
A_CS AC BUS
SW ITCH
STaY _
OFF'
No_,_o' ', Il
t-o'FANS AC BUSSW ITCH
STAN0BY
!NVEP.TEP,
Z50 VA
j,_ll(iMAIN IINV ERTEP. I
150vA IF_NS)__J
MAIN I
I zsovA IL_'"'_PLJ
i SI A_CS
5E ,_. D.C
III
_J
Ii
It
I
II
j ,
....,, _: _ _I T ,-J MAIM STBY _ MAIN J
I zsovA Iso v_ IL i.._,v..._|,'rER =.v _L.1,_E¢._j
, j
riIII
l' I
I
II
I
'iIII
POWEP-, SYSTE.M CONTROL
P,ELAY PARCEL NO. I
L.
m
! --
JhSCS BUSRELAY
P---O_ A
IrD
I"
FA-JN BUS_ RELAY
O'_D
IA
_--'_STAN DBY
T INV. POWER,
F.,ELAY
TIMER.C_TcHDIsCONI4ECT)
FIGURE 10-3 A-C POWER '.CONiROL .SCHEMATIC (,CAPSULES. I/o,I_.F))PM4-5- ZI5 IB
'_'_ I _11,_ I_11 I I
/_ -4
OATE i November 196_
REVlSE;D
REVISED,
ST. LOUIS 3, MISSOURi
inverter.
R_PORT SEDR 104
MODm. 133
The standby inverter will supply a-c power to either, or both, ASCS
and Fans A-c busses by selecting the STANDBY position on the respective ASCS
or Fans Bus switch located on the mai_ instrument panel.
In event of failure of either the 250 VA or 150 VA Main inverters the
respective Fan Bus relay or ASCS Bus relay will be de-energized. This action
will automatically energize the Standby Inverter Relay which in turn will apply
d-c power from the filter to the standby inverter. The a-c output from the
standby inverter is then directed through contacts in the de-energized ASCS or
Fans Bus Relays to their respective busses. A warning light on the Main Instru-
ment Panel indicates when the Standby Inverter is switched into operation by
reason of failure of either of the main inverters. (See Figure 10-3. )
10-8. POWER DISTRIBUTION
i0-9. D-C POWER DIS_I_JTION
D-C power is taken from three separate battery groups, namely the main
battery, standby battery and isolated battery. Various sub busses which operate
from these sources and the bus separation method are as follows:
(a) Main d-c bus directly to the main batteries.
(b) Main 2_ V d-c antenna separation bus from main bus through
separation relay.
(c) Pre-impact plus i0 minute from main bus through impact relay.
(d) Main 2_ V squib bus through SQUIB ARM SW from main bus.
(e) Pre-impact main bus through impact relay from main bus.
(f) Audio bus from main bus or isolated bus through AUDIO BUS SW
(3 position, center OFF).
(g) Isolated d-c bus directly to isolated battery.
(h) Standby d-c bus directly to standby battery.
(i) Isolated 6 and 18 volt Busses directly to taps on isolated battery.
......... ii-r_| - •--_ _m#" Jl _11 Jl J[ &mr &m| 11i i | J--_ J&m
PAGE
REPORT
MODEL,
10-10
Mc'DONNE :SEDR 104 ST. LOUIS , MISSOURI REVISED
133 ' _"_._,m_-- P4P_kll_-ir%L'_'_ _ REVISED
DATE 1 NOVEMBER 1961
NOTE,S
_>:!/-Y _E fN3TALLEDIN CAPSULE FOR r3,ROUND q EST_ OR FLIG, FIT_h,lAY _:,E !NST?,LLE_ IN CAPSULE,FOR GRQUN.b TESTS3L___bONOT INSTALL Ikl CAPSULE, MAY _P_. ACTIVATE.) AND USED FOR #OWER.
._Y,._T=----...'N_TESTS/G;'-_OUND CA_T FOR."I'HFa,U THE HATCH POWE.I_"AND B4TTEeY CHARGERTE:ST,_
i
1-24 v,
OI::F
ON
A A
C, _4 D E._ev I_e
C
!
I •i i
I
I 24V Z4V
[_EAGLE PICHER MAR 4.02B-I, MAC 45-79707-15 (ISO0 W FI)
EAGLE PICHER 1,4AR4027-5, MAC 45-79707-21 ('..5000 WH)AGLE PlCHER _AIA 4028-A, MAC 4.5-79707-19 (I.500 WN)A&LE PICHER MAR4028-8, MAC 4S-7_707-19 (.ISO0 WFI)A&LE PIe.HER _AR4028-1_,,I'4DE4567020-5 (.1_00 WFt)
OFF
I "{EAGLE PICHER MP_R 4027C, MAC 45-79707-/7EAGLE pICHER MAR 4027-1/ MAC 45-79707-13_AG.LE PlCHER MAR 4027A/ MAC 45-797o7-17EAGLE. PIqHF_.P,_AR, 4027C/ 1"4DE 4587020-._
(SOO0 W/O(_ooo WH)(3o0o W_)C3(_o t4//4)
FIG 10-4- BATTERY' INTERklAL WIRING, DIAGRAMPM 45-93A
: \
i _ _i• i_J'
CO II, ! llll - . . - __•., ,P._.-'-;'.;T,r,.L
MC,DONNE REPORT _xE]_ 104 REVISED
.....St. ,_Outs, MISSOU_IMODEL. 1_ ....... _, " - " '_ REVISED
DATE i November 1961
put in an ambient atmosphere of 160°F or at 80°F at 5 psia 100% oxygen. The
output is ll5 volts a-c + 5%, single phase to ground, with a frequency of 400
cycles + 1.0% and essentially sinusoidal in wave-form. A short circuit across
the output of an inverter will not damage the inverter or the wiring involved
in short circuit.
10-14. D-C AMMETER 0-50 AMPERE
The d-c ammeter is located on the main instrument panel and provides the
Astronaut with an indication of total current drain from all batteries. The
basic ammeter movement has a 50 millivolt sensitivity. A shunt of suitable
resistance is connected across the input of the meter providing a low resistance
path to ground with proper voltage drop for the meter movement.
10-15. D-C VOLTMETER 0-30 VOLTS
A d-c voltmeter, and its selector switch, are located on the main instrument
panel. Approximate battery condition can be determined by placing the D-C VOLTS
switch to the appropriate positions and reading the individual battery voltages.
lO-16. TEST CONFIGURATION CAPSULES
Deviations from data applicable to the Test Configuration Capsules are
explained in the following paragraphs and if no data is presented for a particula]
item, this item is the same as that used on the Specification Compliance Capsule.
lO-17. TEST CONFIGURATION CAPSULE NO. 8
The Electrical Power System on Capsule No. 8 is the same as the Specifica-
tion Compliance Capsule except for dlfferencea as noted in the following
paragraphs. (See Figure 10-6 for major components installed.)
lO-18. System Description
The Capsule Power Supply in Capsule No. 8 consists of three 3000 watt-hour
_-_v_.- _ .=..JJ.It,.4_.l JI. J_"L llj-
DATE
REVISED
REVISED
1 NOVEMBER 1961
McDOtNNEL LST.LOUIS, MISSOURI
PAG E 10-13
REPORT SEDR 104
MODEL | 33
L. /
PRE-LAUN.CNZ. HRS.
LAUNCH
.IHK
oRBIT_5 HRS.
RE-ENTR_(.5 HRS
POST
PERIOD
I?..H RS.
1oo ZOOI I
WATT&
300 400 500 _oOO 700 800 _E)O IC_)O {I00 (ZOO
I I | I I I I I I f
SZ4.q.- WATER %EF. 30 SEC/Bo ML_q.+.
I _51._- PROt%_AFAM£._ "A" 5OMIL-SEC/17.5 SIP_.I
To-30 MIN.---- _ -- T__LE.METV<'.(-ONi
I_ |QHF T'P,ARC_EI_VER, _8OOST ST_'I'.
Im /_¢. BAND BEACON-ON,To-IZ MIN. i "_ TAPE REC.ORb-O_
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PAGE
REPORT
MODEL
I0-I 4
SEDR 104
133
Mc'DONNEL LST.LOUIS, MISSOURI
-:.t'_:d_,, Ig ! P ! I'% I_",1_ I "riP'1, , ALl_lPV I li i ;imw, a ¢ ! i IN ILB _liL_
DATE
REVISED
REVISED
1 NOVEMBER 1961
\
/
_[O o
,,40:: 0
_// _
\\
\
j
DAT__ i November 1961-
REVISED_
REVISED--
main batteries, one 1500 watt-hour Main battery, one 1500 watt-hour standby
battery, and one 1500 watt-hour isolated battery. (See Figure 10-7.)
D-c electrical loads are supplied through fuses except to the (1) Retro-
grade Seq. control, (2) Emerg. Capsule Separation, (3) Abort Control, end
(4) Auto Landing System A circuits - these circuits incorporate solid conductors
in place of fuses. S_ne of the d-c circuits utilize two fuses in parallel for
redundancy, with a two position switch which permits operation in the normal
(No. l) position, or the emergency (No. 2) position. In addition, to prevent
possible loss of power during an emergency, a solid conductor is installed in
place of a fuse in the emergency (No. 2) positions of the following circuits:
Retro Manual Control (No. 1 and No. 2)(a)
(b)
(c)
Reserve chute deploy - Systems A and B
Tower separation control.
A-c loads are not fused because of inherent overload protection in the inverters.
i0-19. D-C Power Control
Capsule special instrumentation uses two of the four main batteries. The
normal connector plugs on the No. 3 and No. 4 main batteries are tied back at
bundles 800A and 8013. Special instrumentation bundles 867 SI and 836 SI,
respectively, are connected to the No. 3 and No. 4 batteries.
The main 24 volt d-c Power supply consists of only two main batteries.
(See Figure 10-7.) Each battery is connected in parallel to the circuit by an
ON-OFF switch on the battery. Individual reverse current diode protection
prevents discharge through a faulty or low battery. The 24 volt d-c power from
the two main batteries, or from the external power source prior to launch, iis
applied to the main d-c bus. The main d-c bus is connected through the secondary
bus relay to the secondary bus.
The 1500 watt-hour standby battery is installed in the capsule to supply
196l _I3_W:IAON t
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91-01.
DATE i November 1961
REVISED
REVISED
ST. LOUIS 3, MISSOURI REPORT R_T)R i0_
_oom.., 133a_LL , ......... ,,
communications bus voltage and to act as standby gower for the main system.
The standby battery also incorporates diode reverse current protection and is
connected to the circuit through an ON-OFF switch located on the battery.
Standby battery taps through diodes, supply 6, 12 and 18 volts d-c to the
various system busses. Prior to launch these circuits are energized by exteraal
power through the umbilical disconnect and external power diode package. Stand-
by battery 24 volt d-c power application is controlled by the STDBY BATT switch
on the main instrument panel. This switch may be placed in the AUTO or MANUAL
position. In the AUTO position, a decrease in main battery voltage to a level
somewhat below standby battery voltage causes the No. i secondary bus control
relay to energize. Power is applied through normally open contacts of this
relay to energize the No. 2 and No. 3 secondary bus control relays. These
relays have holding contacts which lock their solenoids to the main d-c bus.
With the No. 2 and No. B control relays energized 3 power is applied to the
solenoids of the secondary bus relays. Energizing the secondary bus relays
removes voltage from the secondary bus and connects the standby battery 24
volt output to the main bus in parallel with the reduced main battery output.
The secondary bus relay also energizes the STANDBY D-C AUTO indicator light
indicating automatic use of the standby battery. Signal voltage indicative of
STANDBY D-C AUTO lamp operation is supplied to the instrumentation system.
The _N position of this STDBY BATT switch connects the standby battery
24 volt output directly to the main 24 volt d-c bus and through the secondary
bus relay to the secondary 24 volt d-c bus. No visual indication is made of
this use of the standby battery other than Astronaut checks of current and
voltage. An emergency hold circuit is utilized in the event of a "hold"
command after umbilical separation. To reduce cabin heating the emergency hold
circuit removes power from the secondary and the main ASCS busses and applies
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DATE i November 1961
REVISED.
REVISED,
s-,-.Lou,s 3. M,sSOURI ,,_ORV _n 1O4
_' DV;_T;AL ,.,oom. 133
power to the cabin vent squibs end to the extend motor of the periscope. Cir-
cuit switching is accomplished as follows. The '_old" signal from the booster
is applied through normally closed contacts of the No. 3 ground test umbilical
relay to the solenoid of the No. 1 emergency hold relay. Power from the main
d-c bus is applied through normally open contacts of the No. 1 emergency hold
relay to the solenoids of the No. 2 and No. 3 emergency hold relays. Normally
open contacts of the No. 2 emergency hold relay apply power to the solenoids
of the secondary bus relays, impact sensor relay and to the extend motor of the
periscope. Other normally open contacts of the No. 3 emergency hold relay
apply power to the cabin vent squibs and remove power from the main 2A volt
Ascs bus. (see Figure lO-?.)
10-20. A-C Power sad Control
10-21. Main Inverters
Main, 115 volt, 400 cycle a-c power is supplied by one 250 volt ampere
inverter and one 150 volt ampere inverter. (See Figure 10-8.) The a-c load
is divided between the ASCS bus sad the fan bus. During flight when the ASCS
is the primary a-c load, the 250 Volt ampere inverter powers that bus while the
150 volt ampere inverter powers the fan bus. During re-entry, at approximately
300,000 feet, the fan system becomes the primary a-c load. Therefore, the In-
verters are switched such that the higher capacity inverter powers the fan bus.
The main d-c bus supplies input power to the inverter feeding the fan bus while
the secondary d-c bus supplies input power to the inverter feeding the ASCS bus.
The output from one inverter energizes the ASCS BUS RELAY while the output of
the other inverter energizes the FAN BUS RELAY. Inverter output is then
supplied through the energized bus relays to the appropriate bus. An A-C VOLT-
METER is provided with a spring loaded A-C VOLTMETER SWITCH which normally
closes the voltmeter circuit to the fan bus sad must be manually operated to
read ASCS bus voltage.
REPORT SEDR 10/4 . REVISED....sT. LOUlS, MJssou!_! _
MODEL l_ ........ ,,-,_nn,, • = , REVISED_ IL.4%.I.L, JL' .IL./I_4L_.I., 4L 4L.C'JkJL./-
10-22. S..t_sadby Inverter
Standby, 115 volt, 400 cycle a-c power is supplied by one 250 volt ampere
inverter. The STANDBY INVERTER switch determines which d-c bus shall supply
the standby inverter input voltage and whether the a-c output voltage shall
power the ASCS bus or the fan bus or shall automatically power either bus in
the event of main inverter failure. With the STANDBY INVERTER SWITCH in the
ASCS ONLY position, the standby inverter is energized by 24 volts d-c from
the secondary d-c bus. The standby inverter output energizes the STANDBY ASCS
BUS RELAY. The energized STANDBY ASCS _JS RELAY removes the main inverter out-
put from the llne nnd applies the standby inverter output.
When the STANDBY INVERTER switch is in the FANS position the standby in-
verter is energized by 24 volts d-c from the main d-c bus. The standby inverter
output energizes the STANDBY FAN _JS RELAY. The energized STANDBY FAN BUS RELAY
removes the main inverter output from the llne and applies the standby inverter
output.
The AUTO position of the STANDBY INVERTER CONTROL SWITCH allows the in-
verter to power either a-e bus should the main inverter feeding one of them
fail. Failure of a main inverter causes the associated bus relay to de-energize,
This connects the standby inverter input to the d-c bus used by the failed main
inverter and feeds the output to the proper a-c bus. The STANDBY A-C AUTO
light illuminates during standby inverter operation while in the automatic mode.
Should both main inverters fall while in automatic mode, the standby inverter,
operating from the main d-c bus, will power the fan a-c bus. Signal voltage
indicative of STANDBY A-C AUTO light operation is supplied to the instrumenta-
tion system.
No Emergency A-C Power switch is provided for use in Capsule No. 8. In-
stead, 24 volts d-c is connected directly through the No. 10.05g relay into the
/
J
MAC 231C (Rev 14 Oct. 65)
1, owber Mc,DONNEg. ..o. lo-2 nev,sED ST.LOU,S s, u,ssoum _e_oRv SEDR 104
REV,SeO . MOOre. 133
150 VA Main Inverter.
10-23. D-C Power Distribution
On Capsule No. 8, d-c power is taken from three separate battery groups,
namely the main battery, standby battery and isolated battery. Various sub
busses which operate frcm these sources and the bus separation method are as
follows. See Figure 10-7 for special instrumentation battery usage.
(a) Secondary d-c bus supplied by main d-c bus through secondary
bus relay.
(b) Secondary ASCS bus supplied by secondary bus through ELEC PWR RELAY.
(c) Pre-impact secondary bus through impact relay.
(d) Main 24 V squib bus through SQUIB ARM SW from main bus.
(e) Pre-impact main bus through impact relay from main bus.
(f) Audio bus from main bus or isolated bus through AUDIO _JS SW
(3 position, center OFF).
(g) Isolated d-c bus directly to isolated battery.
(h) Standby d-c bus directly to standby battery.
(i) Isolated 6 and 18 volt busses directly to taps on isolated
battery.
(J) Standby 6, 12 and 18 volt busses directly to taps on standby
battery.
10-24. D-C Power Loading
See Figure 10-13 for graphical summary of the Primary d-c Power Loading
on Capsule No. 8.
10-25. TEST CONFIGURATION CAPSULE NO. 9
The Electrical Power System on Capsule No. 9 is the same as the Specifica-
tion ,ompliance Capsule except for differences as noted in the following
paragraphs. (See Figure 10-6 for major components installed. )
A A - -- ---
PAGE
REPORT
MODEL
10"22
SEDR 104
133
M 'DONNELjL_---_T_.OUlS,,.MISS_I..,,,
DATE
REVISED
REVISED
1 NOVEMBER 1961
-13
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DATE 1 November 1961
REVISED.
REVISED.
ST. LOUIS 3, MISSOURI REPORT _E_R 104,,
• " - MOOW. 133 ,
10-26. S_stem Description
The capsule power supply for Capsule No. 9 is similar to that used on the
Specification Compliance Capsule.
i0-27. D-C Power Control
The d-c power control on Capsule No. 9 is similar to that used on the
Specification Compliance Capsule (Refer to Paragraph IO-A), except for the
following differences. (See Figure i0-7).
On Capsule No. 9 the Secondary bus control relays No's. I, 2 and 3, and
the Secondary bus relays No's. 1 and 2 are not used. The standby battery power
is connected directly to the Main d-c bus, in parallel with the main batteries,
with the standby battery switch placed in the "OR" position. The standby d-c
light (AUTO), or automatic switching of the standby bus to the Main bus is not
used. Operation in the "ON" position of the standby battery switch as well as
all other controls is the same as the Specification Compliance Capsule. The
hold relay operation is the same as used on the Specification Compliance Capsule
except additional contacts are used in the No. 2 hold relay to control the
Power to the electrical power relay from the umbilical discomnect.
10-28. A-C Power and Control
10-29. Main Inverters
Main 115 volt _00 cycle a-c power is supplied by two inverters of 150 volt-
amperes and 250 volt-amperes. The a-c load is divided into two portions namely
the ASCS bus and the FANS bus. The 250 VA inverter supplies the ASCS bus and
the 150 VA inverter supplies the FANS bus. The main d-c bus powers the 150 VA
(fans bus) inverter through a filter circuit and a 25 ampere fuse. The 250 VA
inverter (ASCS bus) is also powered from the Main d-c bus through a line filter,
a 25 ampere fuse and the antenna separation relay. The Main ASCS bus is
energized through the EL_ 1_ relay.
............ |!"'_¥1_ili i ll./_ll_ll • it-ilil--
REPORT SEDR 10_ REVISED..... ST, LOUIS, MISSOURi ......
MODEL_ 133 C,_ _'_ _i_) [1 _ TI._._ REVISED,
(Other ASCS d-c loads are fed from the ASCS bus.) The outputs of the Main
150 VA and Main 250 VA inverters feed the solenoids of the Fans Bus Relay and
the ASCS Bus Relay. These energized relays feed the inverter output through the
closed contacts of the relays and through the normally closed contacts of the
Standby ASCS Bus Relay to Power the fans and ASCS busses. (See Figure i0-ii).
The 250 VAC invez_ers are deactivated by the antez_,za fairing separation relay
during descent.
10-30. S_t_and_byInverter
Standby 115V _00 cycle a-c power is supplied by one 250 volt ampere standby
inverter. The STDBY INVERT switch determines the mode of operation and/or which
a-c bus shall be supplied by the standby inverter.
The EMER A-C position of the switch manually energizes the standby inverter
from the ASCS bus. Standby inverter a-c power then energizes the solenoid of
the Standby ASCS Bus Relay. A-c power then flows through the energized Standby
ASCS Bus Relay's contacts to Power the ASCS a-c bus.
The FANS ONLY position of the STDBY INV switch manually energizes the
standby inverter from the main d-c bus. Standby inverter a-c power then energlm_i
the solenoid of the Standby Fan Bus Relay. A-c power then flows through the
energized standby fan bus relay's cOntacts to Power the FANS bus. (See Figure
lO-n.)
The AUTO Position of the STDBY INVERT switch allows the inverter to Power
either a-c bus should the main inverter feeding one of them fail. Failure of
a main inverter causes the associated bus relay to de-energize. This connects
the standby inverter input to the d-c bus used by the failed main inverter and
feeds the output to the proper a-c bus. The STDBY A-C AUTO light illuminates
during standby inverter operation while in the automatic mode. Should both
main inverters fail while in automatic mode, the standby inverter operating from
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DATE 1 November 1961
REVISED.
REVISED
ST. LOUIS 3, MISSOURI
p._ 10-27
the main d-c bus, will power the fan a-c bus. Signal voltage indicative of
STDBY A-C AUTO light operation is supplied to the instrumentation system.
10-31. D-C Power Distribution
The d-c power distribution on Capsule No. 9 is the s_ne as used on the
Specification Compliance Capsule (Refer to Paragraph 10-8), except for differ-
ences as shown on Figure 10-7.
10-32. D-C Power Loading
See Figure i0-i0 for graphical summary of the primary d-c power loading
on Capsule No. 9.
10-33. TEST COHF]DORATION CAPSUI3S NO. i0, 13 AND 16
The electrical power system on Capsules No. i0, 13 and 16 are the same
as the Specification Compliance Capsule. (See Figure 10-5 for graphical
summary of the primary d-c power loading on Capsules No. I0 and 16, and Figure
No. 10-13 for Capsule No. 13. )
iO-3_. INTERIOR LIDHTING
10-35. SYST_q _gSCRIPTIONi
Interior lighting for the Specification Compliance Capsule consists of
four fluorescent flood lights and a series of warning telelights. See Figure
i0-I_ for iocatic.1 and arrangement of cabin and system telelights.
10-36. Cabin Flood Li_ts
Two flourescent flood lights are mounted on bracgets to the right and
left and above the Astronaut. Power for the cabin lights is supplied from the
115 V a-c Capsule invezCer Fans _ts and controlled by a three position switch
located on the lef_ console, the switch positions are marged "_>I_" "L.H. 0nly"
and "OFF". The cabin flood lights are of high actinic value_ especially suit-
able for camera usage. The lights produce little heat and have a low wattage
.-b_= ---- .... _w • |
PAGE
REPORT
MODEL
I0-:2,8_
SEDR 104
133,
McDONNELLjH__o_ST.LOUIS, MISSOURI
_Vlli I'llJ_lll i 18--_ l'-T'-
DATE
REVISED
REVISED
1 NOVEMBER 1961
0 /00
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ORBIT
1
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F/GURE I0:1_ D-Cw,_T_-HOU_LO._/NGCAPSULE NO.,9)
PM45 - _55
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PAGE 10-70 ' McIDONNE_ %_r_ DATE 1 November 1961
REPORT SEDR lO_ ST. LOUIS, MISSOURI REVISED
13_ _a_l _.T _ _.__ REVISEDMODEL_
mtlng of four watts each.
10-37. Photo Flood Lights
nm='e 1o-15).
Two flourescent flood lights are mounted on brackets to the left and right
of the cabin flood lights described in Paragraph 10-35. The photo flood lights
are identical to the cabin flood lights and are controlled by a PHOTO LIGHTS
switch, which has an ON and OFF position. The PHOTO LIGHTS switch is mounted
on the left console.
1o-38. TeZeli hts
Warning telelights are proviaed for various capsule systems and are mounted
on the main instrument and left console panels as shown on Figure 10-14. The
telelights are connected to the various systems to notify the Astronaut of mal-
function of a particular system or verific•tion of an individual function.
Power for the telellghts is supplied by the Capsule 12V d-c or 24V d-c bus
through • five ampere IbASe as shown on Figure 10-15. A Dim-Brlght Switch is
provided for daylight or dark operations.
10-39. TEST CORF_ION CAPSULES
The data contained in Paragraphs 10-34 through 10-38 applies to the
Specification Capsule. Deviations fr_n the data applicable to the Test Con-
Figuration Capsules are explained in the following paragraphs. If data is not
presented for • particular item, then the item is the same as that used on the
Specification Compliance Capsule.
i0-_0. TZST CONFIGURATION CAPSULE NO. 8,, ,,,
1o- 1. hood m ts
The, cabin flood lights on Capsule No. 8 is the same as the Specification
CoEpliance Capsule, except the control switch on Capsule No. 8 has "ON-OFF"
positions only. (Refer to Paragraph 10-43. ) No photo flood lights are used on
DATE
REVISED
REVISED
1 NOVEMBER 1961 Mc'DONNE #ST.LOUIS , MISSOURI
• __ . " ,,
PAGE 10-31
REPORT SEDR 104
MODEL 133
NOTE_
EE> _AME PL_E.
[_ _IOT II_-T'_t.t.F._ OI4 CAPeJ,_LE ¢)
......... I • •
_vilr Iur-iN I In_--'
PAGE ,, 10-32
REPORT S_ 104
MODEL. l_
Capsule No. 8.
i0-_2. Warning Telelights
REVISEDST. LOUIS, MISSOURI
REVISED
The warnin6 telelight installation on Capsule No. 8 is the same as the
Specification Compliance Capsule (Refer to Paragraph 10-38 and Figure 10-16),
except for the following differences.
(a)
(b)
(o)
(d)
(e)
(f)
The Periscope retract telelight is not used on Capsule No. 8.
The Lending Bag telelight is not used on Capsule No. 8.
A Drogue Chute telelight is used on Capsule No, 8.
The Cabin Pressure telelight is not used on Capsule No. 8.
The Fuel _antity telelight is not used on Capsule No. 8.
Mayday and Ready telelights are used On Capsule No. 8 in place
of the abort lights.
(g) A recording telelight is used on Capsule No. 8 to indicate
operation of the tape recorder.
(h) A Standby d-c Auto telelight is used on Capsule No. 8.
lO-_3. _ST co__o_ c_ _o. 9t lOt 13 A_ 16
The interior lighting system on these Capsules is the same as the Specifica-
tion Compliance Capsule (Refer to Paragraphs 10-35 through i0-_) (See Figure
lO-I_ _ lO-15).
_mw • AJ'q _'_ ILT IF'_ T _r_L _
DATE 1 NOVEMBER 19611
REVISED
REVISED
Mc,DONNELL__ __ .PA_E ,ST.LOUIS. MISSOURI REPORT
_ l . MODEL
10_33,
SEDR 104
133
L:: • :,;
P IC_.,UIEEIO-.IBI_IEIEIOP.,LkC_HTS AND WARNIi',I_LIGHTS. scHE_ATIC(,cA_._I_I_i_._;)._v_._n=
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11-I
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SECTION XI
COMMUNICATION SYSTEM
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TITLE
TABLE OF CONTENTS
PAGE
SYSTEM DESCRIPTION ......................... 11-5
SYSTEM OPERATION 11-9eee.eeeo oeeeoeoeo el oeeeeeeee
SYSTEM UNITS .................................... 11-28
TEST CONFIGURATION CAPSULES ...... 11-36
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ANTENNA ML_TI PLEXER
MAIN UHF RECEIVE:R-TRANSMITTER
IJHF- DE:,
SWITCH _T_ANSMIT SWITCH
VOLUM LCONTROLS/_ L X
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PANEL (ANT SWITCH) DETAIL ASELECTOR
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AUDIO CONTROL PANELSEE DETAIL k
XI _ 0
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CO14_IffICATIONI_'NCTIONS
CALL CAPSULES).
FJRCTION
m_' (_,, _,_,_:,) R/_
m,' (_) R/_
_ (_) RI_
HF (Recovery) R/T
C_and Receivers
Telemetry Transmitter
C and S Band Beacons
C and S Band Beacons
UHY Recovery Beacon
H.F. Recovery Beacon
Aux. UHF Rescue Beacon
Ultr. Sarah Rescue Beacon
Unmanued Capsules - No. 's 8, 9Manned Capsules - No. 's 10, 13,
LAUNCH ORBIT RE-ERTRY POST LANDINGIMPACT
Astronaut Selection on All Manned CapsulesI I !Pre-Launch Selection on All Unnmnned Capsules
i IAstronaut Selection on Manned
Capsules
iPre-Launch Selection _--
on Unmanned Capsules
-_ -- All CapsulesI
._ --All Capsules
I-_Manned Capsules - On Continuous
or Controlled by Astronaut and/orGround. CcmnaudI I
_-U_d Capsules - On Continuousor Controlled by Ground Command
I
Bicone Ant. Separation
Bicone Ant. Separation
I16, 18 & 19
-_---Al.lC_sules
v
=-i0 Minutes After Impact
, All Capsules
All Capsules
ii0 Minutes After--_-Capsules I0_13,16,18 and 19
Impact I
Capsule 8 & 9 only
MCDONNE ll-5R_-VtSEO. ST. LOUlSa. M,SSOUm _ORT SE_ 104
.... \
--f,
XI. COMMUNICATIONS SYSTEM
ii-i. SYSTEM DESCRIPTION
ii-2. VOICE COMMUNICATION
The Astronaut Is provlded with volc_ communlcatlons throughout the entire
mission (See Table ii-i). A dual headset and microphone contained within the
Astronaut's helmet, operate through the audio control circuits to the selected
voice communications set (See Figure 12-2). A capsule-pad Luterphone system is
available prior to umbilical cable disconnect.
HF reception is available through the main KF voice communication set during
launch and orbit. HF voice transmission may be used only after capsule separa-
tion by Astronaut selection of the HF position of the Tranmnit Switch. The main
set is disabled during re-entry as the antenna fairing is Jettisoned. It will
be de-energized and replaced by the recovery HF set upon landing. The recovery
HF voice communications set provides reception and transmission, during the post
landing phase of the mission.
UHF reception is available throughout the entire mission by the Cc_m. UHF
voice communications set and its UHF Booster Amplifier. Transmissions over this
setmay be made when the "UHF" position of the TRANSMIT switch is selected by
the Astronaut. A backup (low power) UHF voice cce_unication set identical to the
main set, but without the UKF Booster Amplifier, may be placed in operation by
the Astronaut at any point during the mission.
The selected transmitter may be energized by operation of a push-to-talk
switch_ or by a voice operated relay when the VOX switch is in the "ON" position.
By speaking into the microphone, the selected transmitter is automatically
energized. Normally the selected UHF transmitter will automatically be energized
upon landing to provide a direction finder signal. This automatic feature may be
_J,J,;_D; D-,_.i4T;AI.
l.....
REVISED ST. LOUIS 3, MISSOURI RI[PORT _ 10_
REV,S¢. ................. MODm. , 133
overridden by the Astronaut.
The Cc_mmnd Receivers provide an emergency ground station-to-capsule voice
communications channel throughout the mission until capsule impact. Power for
the voice communications systems is supplied through fuses located in the
CQmmunications and Ccmnunications ASCS fuse holders (See Figure ii-2).
11-3. COMMAND RECEIVERS
Two separate sets of receiver-decoder and auxiliary decoder units are used
for reception and decoding of ground ccmnand signals. These signals are for th_
purpose of activating various capsule control circuits.
Power for the command receivers is supplied through the fuses located in
the Cc_unications and the Communications ASCS Fuse Holders (See Figure 11-10).
ii-4.
Telemetry transmitters are provided for ccm_mieating capsule information
to ground stations. Information is picked up throughout the capsule in the form
of voltages from vol_a6e divider circuits. These voltages are modified by coding
circuits to sul_ly suitable inputs to the telemetry transmitters. (Refer to the
Instrumentation Section XIII of this manual). Two transmitters are used for
transmission of the telemetry information, each having a power output of 3.3
watts. Their frequencies are slightl_ separated. Transmitters are operated
continumasly frum launch, until i0 minutes after impact. The power outputs of
the telemetry transmitters are fed to either the main or the UKF Recovery
Antenna. Power for the system is obtained fr_n fuses located in the Instrumenta-
tion Fuse Holders (See Figure 11-_ and Figure 11-11).
_-5. z_co_s
The beacons provided in the capsule to aid tracking by ground stations are
a C-Band end an S-Band beacon, a UHF recovery beacon, energized during re-entry,
_, .... -- w _"lr"% !'1" ! AItWINIi" i lll*llli . . ...__
,....
U'_.C:
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ANTENNA
SWITCH
El
-TELEMETRY
ANTENNAL POWER SUPPLY
MULTIPLEXERTELEMETRY
TRANSMITTERS
\INSTRUMENTATION
PACKACE"C"
MAIN INSTRUMENT PANELDETAIL A
RL'VmED. rr. Louts 3. M,ssoum Rm:,owr ,Sg_, lOk,
.SVmED. --- ..... ,,,or,,- 133
an aux. UEF beacon energized 10 rain. after impact, and an HF recovery beacon,
energized upon land_. These beacons provlde signals compatible with direction
finding equil_nent used by the recovery crews. The U_ voice communications
transmitter may be keyed upon landing to provide an additional signal for
direction finders. A flashing light is installed for visual location of the
capsule after landing. (See Section IX of this manual. )
Capsule power for the beacons system is supplied through fuses located
in the communications and ASCS fuse holders. (See Figures II-_ and 11-12).
The voice communications, telemetry and beacon receivers and transmitters,
with their various frequencies and types of outputs require an antenna system
with wide capabilities. Therefore, four types of antennas are used to fulfill
the entire mission requirements. A main HF-UEF antenna is used for the major
portion of the mission. During re-entry this antenna must be Jettisoned to
allow main parachute deployment. To replace the UEF function, a compact UHF
antenna is automatically placed in operation. Upon landing, an HF Recovery
Antenna is extended to permit HF operation. Throughout the entire mission, C
and S band antennas are provided for operation of the radar beacons. Antenna
switching and multiplexing are Performed autcmatlcally by the RF circuitry.
(See Figures 11-6 and 11-13). Power requirements for antenna switching are
supplied through a switch-fuse located on the left console switch-fuse panel.
(see 11-13).
11-7.
11-8.
11-9.
OPERATION
VOICE COMMUNICATIONS
Audio Control and Ground Inter_hone ST_em
HF and UEF receiver outputs are routed to the control panel. This panel
1_*Vil_lla la,..__. _ Jl _ ___
11-10 DATE 1 NOVEMBER 1961PAGE
REPORT
MODEL
SEDR 104
133
ST.LOUIS , MISSOURI REVISED
REVISED ,
/
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REVISED. ST. LOUIS 3, MISSOURI REPORT... _ ]-0_
Revis_ __,'_ _ I _, ,-_,- L ..... 133_ Rm_R_Y_'_._ _ __,_, __-_ ,_.-- MODEL
' i
_-_:_
provides one volume control for HF audio and one volume control for UHF audio.
Outputs from the two c_and receivers are connected to the Co-_unication Control
Panel for mixing. Separation of command and voice audio signals by a low pass
filter, and amplification of resulting voice audio is done in the control panel_
(See
Communication audio signals from the volume controls, the interphone audio
from the pad-to-the-pilot and alarm tones from the satellite clock are supplied
to the tape recorder relay and the two headset amplifiers in the audio center.
The headset amplifiers serve to amplify the audio signals and feed them to the
individual earphones in the Astronaut's helmet. The de-energized position of
the tape recorder relay supplies a path for receiver audio to the main tape
re corder.
Audio from the microphones is fed to two separate microphone amplifiers
in the audio center. These two amplifiers serve to amplify microphone output
' to a level sufficient to supply modulation circuits of the voice transmitters.
The microphone smplifer output is also fed to the input of the VOX (voice
operated relay circuitry). The voice circuits are energized by use of the Push
to Talk switch on the abort handle, or by the VOX circuit when the VOX switch
on the instrument panel is in the "ON" position.
ii-iO. KF Voice Communications
11-11. Main HF Ccmnunications
The main HF voice cc_mnications set is an AM receiver-transmitter unit
designed to operate on a frequency of approximately 15 MC.
Power from the main pre-impact 24 volt d-c bus is fed directly to the
receiver section of the set. The transmitter is fed 2It volts through the EF
Position of the Transmit Switch and the closed contacts of the tower separation
relay, after tower separation. Audio input to the transmitter Portion of the
....... .,_,..a i_ ! _, ,..
P'_E. ,,.'1_ MCDONNE_ __ D_TEREPORT SEDR 104 ST.LOUIS. MISSOURI REVISED
_' ".-_ m •
MODEL 133 ___-__ REVISED
1 NOVEMBER 1961
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ANTEN N,_
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R_VISED ST. "OUm s. MiSSOURi ma:ORT. SE_ 10_
REV:sEo. . ........... _oom. 133
unit is from the microphone amplifier in the audio center. The transmitter is
energized either automatically through the VOX circuit or manually by thei
Astronaut's use of the push to talk switch. (See Figure ll-8).
The antenna connection from the set is through the antenna multiplexer and
the antenna switch to either the bicone or the UHF Rescue antenna. HF raddatiau
from the descent autenna is negligible. Audio output from the receiver, includ-
Lug sidetone during transmission is routed to the HF volume control in the
control panel.
ll-12. Recovery HF Ccemm_uications
The recovery HF voice comnmicatioms set is basically the same as the main
HF unit but with the transmitter having a lower output.
The power input to the recovery HF unit is supplied upon laudimg, through
the impact relay, from the audio 24 volt d-c bus. Audio is supplied and keying
of the transmitter is in the same manner as the maim set.
The antenna connection is through the recovery dIplexer to the extended
HF Rescue antenna. Audio output from the receiver, including sidetome, is
routed through the HF volume control in the control panel through the audio
center and to the headsets.
11-13. UHF Voice C_icatlons
ll-14. Maim UHF
The main UHF voice camumicatious set is am AM receiver-tramsmitter unit
designed to operate on a frequency of approximately 299 MC. The transmitter
output is increased by a separate _THF booster amplifier.
Power fram the audio 2_ volt d-c bus is fed through the HI-PWR position
of the UHF Selector Switch directly to the receiver section of the set. (See
Figure 11-9). This power is also fed to the Transmit Switch. Power for the
tramsmitter section of the set is then taken from the UHF position of the
!1-16 ..... _ f3__ DATE 1 November 1_961
ST. LOUIS, MISSOURIMODEL,, 1B3 ............_,.,_,_"..............,.,_, _._-_.._ REVISED,
Transmit Switch. At bicone antenna separation the bicone separation relay con-
tacts assume the same function as the UHF contacts of the Transmit Switch thus
providing a continuous UHF signal for DF purposes. Audio input to the trans-
mitter portion of the unit is from the microphone amplifers in the audio center.
The transmitter is energized either automatically or manually by the Astronaut.
When the HI-POWER set is selected with the UHF Selector Switch, it will be
energized automatically at bicone separation to provide a UHF signal for direc-
tion finding equipment. This feature may be overriden by operation of the UHF
DF Switch on the control panel to the OFF or the UHF position.
Antenna connection from the set is through the UHF booster amplifer coax
switch, antenna multiplexer, and the antenna switch to either the main bicone
or UHF rescue antenna. Operation of the microphone switch or energizing the
voice operated relay while in the UHF mode causes the booster amplifier to be
inserted in series with the coax line. Transmitter output is then boosted by
this amplifier to 2 watts. The booster is also available after landing. The
multiplexer output is connected through the antenna switch to either the main
bicone or the UHF rescue antenna. Audio output from the receiver, including
sidetone during transmission, is routed to the UHF volume control in the control
panel.
11-15. Backup UHF
The Lo power USF voice communications set is identical to the main set but
without the booster amplifier. It may be energized by Astronaut operation of the
UHF Selector Switch to the "LO Ir_R" position. Power input, audio input, receiver
output, and the control of the transmitter is in the same manner as used for the
main UHF set.
Antenna connection is routed through the UHF coax switch, which has been
energized by selecting the "LO PO_R" position of the UHF Selector Switch, throu_
1
MAIN• POWER :UH_ VOICE
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PAGE 11-18
REPORT SEDR-104
MODEL. 133
McDONNE DATE 1 November 1961
REVISED-,,_ST. LOUIS.,. MISSOURI
"/'_'_T ...... n.,.,- • _."_ REVISED_,_,_._ 4,_I.IL .iL,LJ41.JJ. ll JL JL,Z-LJ.._
the UHF booster's coax switch, the antenna multiplexer and the antenna switch
to either the main bicone or UHF rescue antenna.
11-16. Command Receivers
The receiver-decoder unit consists of an FM receiver operating in the
frequency range of 406 to 450 MC. The received signal may be modulated with a
maximum of six of a possible twenty audio frequencies. The receiver reduces
the input signal to the modulation frequencies which operate individual control
relays. (See Figure 11-10).
Each control relay provides contacts for a normally open or a normally
closed control channel. Ten channels are provided in the "A" recelver-decoder
with an additional ten available in the "A" auxiliary decoder. These channels
are paralleled by the output of the "B" receiver-decoder, and auxiliary decoder
units. Command channel assignments are not disclosed for security reasons.
Emergency voice communications may be had from the ground station to the capsule
through the command receivers. Receiver outputs are supplied through a filter
and amplifier in the audio center circuits to the Astronaut's headset. Power
for the Hi frequency command set is supplied from the isolated 18 volt d-c bus
while power for the Lo frequency command set is supplied from the standby 18 volt
d-c bus. Both power circuits are routed through sections of the impact relays
in order to de-energize the set upon landing.
Antenna input is from the bicone or UHF rescue antenna through the antenna
switch and antenna multiplexer to an impedance match which supplies both re-
ceivers.
-17. Telemetr v
11-18. Low Frequency Telemetry
The low frequency telemetry set is an FM transmitter operating on a fre-
quency of approximately 228 MC.
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REVISED _lr, LoUIs 3, MISSOURI REPORT ,_, 10_ ,
_v,sED ,'___!F! ___._'._._._L_ MOD,_ IS3
Before umbilical drop, the low frequency telemetry transmitter and its power
supply receive 24 volts d-c from the M_in Pre-impact bus through the energized
Ground Test Umbilical Relay, normally open contacts. This relay's solenoid
is energized through the umbilical until the umbilical is dropped. (See Figure
ll-ll).
To silence the TM transmitter, the solenoid of the Instrument M,c_leand
R-F Silence relay may be energized through the following three methods. First
by poSitioning the GROUND TEST switch in the block house to the "TEST" position,
the Instrument Mode and R-F Silence Relay will be energized. By a second
method the solenoid may be energized through the de-energized Orbit Attitude
Relay, thus causing the silencing of the TM transmitters. The third method of
energizing the Instrument Mode and R-F Silence Relay is through the normally
open contacts of the energized Retro Command Relay #2. After the umbilical is
dropped, the ground test umbilical relay is de-energized and power is directed
through the closed contacts to the telemetry Power supply. (See Figure 11-11).
A LO FREQ. TLM, 0N-OFF switch located on the left console, when in the "OFF"
Position, will break the Power source to the LO FREQ. TM.
Coded instrumentation information is supplied from the instrumentation
package "D", and used tO frequency modulate the transmitter. (See the
Instrumentati_ Section XIII of this manual).
RF power output is fed to the antenna multiplexer where it is routed througk
the antenna switch to the main bicone or UHF recovery antenna.
11-19. High Frequency Telemetry
The hi frequency telemetry set operates on a frequency of approximately
260 MC. The 2_ volts d-c to the hi frequency telemetry is supplied by the same
method as described under the low frequency telemetry description, with the
exception that the ON-OFF switch provided in the capsule is in the form of a
REPORT ,S_DR 10_ ST. LOUIS, MISSOURI REVISED
i'_ --'_ ...... "_ _"' .... REVISEDMODEL. ,._.... -_.- .......
switch-fuse (HI FREQ. ), located on the left hand switch-fuse panel. Input power
for the set is from a separate telemetry power supply operating from the second
pre-impact 24 volt d-c bus and supplying filament and B+ voltage. The B+ volt-
age supply is routed through the key on the control panel. This allows the
Astronaut to interrupt the circuit transmitting code, in the event the voice
communications fail. (See Figure ll-ll).
Coded instrumentation information is supplied from the instrumentation
package "D" to frequency modulate the transmitter. (See the Instrumentation
Section XIII of this manual).
RF power output is fed to the antenna multiplexer where it is routed through
the antenna switch to the main bicone or UHF recovery antenna.
ll-20. Beacons
ll-21. C-Band Beacon
The C-Band beacon is a transponder unit consisting of a receiver and trans-
mitter operating on a frequency of approximately 5_O0 to 5900 MC. The beacon is
double pulsed and is compatible with the FPS-16 radar when the ground units are
modified for this type of operation. Upon ground command, through the command
receivers, or by Astronaut selection of the "CORTIN." position of the Beacon
Switch, the beacon receiver is energized. Interrogation by ground radar will
then result in a coded reply from the beacon transmitter. Input power is from
the main pre-impact 24 volt d-c bus through the beacon relay controlled by the
command receivers, or, for continuous operation, through the Beacon Switch.
(See Figure ll-12. )
Beacon antenna connection is through a phase shifter and the C-Band power
divider to the three C and S baud beacon antennas.
11-22. S-Band Beacon
The "S" Band Beacon is a transponder unit consisting of a receiver and trams.
C_JA'q J. A_-I_A_ A £A_MLR _"
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PAGE ! 1-2_
REPORT _]_ 10_
MODEL_ 133
Mc,DONNE DATE 1 November 1_1REVISED
ST. LOUIS, MISSOURI
_, _. _'__1;.____. _ 'i' _. ,_.._r_ REVISED
_Ltter. (Bee Figure Ll-12). The unit operates on a frequency of approximately
2700 to 2900 MC and is double pulsed to reduce possibilities of unauthorized
interrogation. This unit is compatible with ground based Verlort Radars and
operates at a positive acceptance tolerance of -+ 0.5 micro-seconds and a positive
rejection tolerance of +- 1.8 micro-seconds.
Power circuits, interrogation and reply are the same as the C-Baud Beacon.
(Refer to paragraph 11-21).
Beacon antenna connection is through the S-Band Power Divider to three "C"
and "S" Band Beacon antennas.
11-23. HF/UHF Recovery Beacon
Two recovery beacons are combined into one unit. One beacon operates on
high frequency, while the other operates on ultra high frequency. Both are
energized to provide radio signals for recovery direction finder equipment.
(See FJ.g_e Ii-12).
The HF recovery beacon operates on a frequency of 8.364 MC with a tone
modulated output. It is powered by the 12 volt standby bus through the impact
relay and is energized upon landing. The RF power output is fed through the
rescue diplexer to the elevated HF recovery antenna.
The UHF recovery beacon operates on a frequency of 211.3 _ with pulse
modulation. It is powered by the 6 volt isolated bus through the antenna fair-
img separation relay. This circuit is energized during re-entry when the
antenna fairing is Jettisoned. The RF power output is fed through the antenna
multiplexer and the antenna switch to the UHF recovery antenna.
ii-2_. Auxiliary Rescue Beacon
The Aux. Rescue Beacon operates on a frequency of 243 MC with pulse modula-
tion. It is powered by the 6 volt standby bus through the emergency hold relay
#2 and the 10 minute delay impact relay. These circuits are energized 10 minutes
r
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REPORT SEDR 104 REVISEDST. LOUIS, MISSOURI
MODEL. 1_ _r,,rt.l_ T_ lr n _ __L_... REVISED
after impact. The RF power is radiated from the Aux. Rescue Beacon Antenna.
(see Figure n-12).
11-25. Antennas
11-26. Main Bicone (HF and UHF)
A biconical antenna is used for pre-launch, launch, orbit and initial re-
entry phases of the mission. This antenna is an integral part of the antenna
fairing and is located over the open end of the recovery system compartment of
the cylindrical capsule a_erbody. The blconlcal antenna serves the main HF
and UHF voice receiver-transmitters, the command receivers, and the telemetry
transmitters. The active element of the biconical antenna forms the upper
portion of the antenna fairing while the lower portion of the fairing and the
capsule body forms the ground plane for the antenna. (See Figure If-IS).
II-L_'/. _ Recovery Antenna
A UEF antenna is used for the final phase of re-entry, landing and rescue.
It is a compact antenna located on the open surface of the recovery systems
compartment. The antenna is folded when the antenna fairing is installed. Six-
teen seconds after the fairing is jettisoned the UHF recovery antenna is erected
and serves the UHF voice receiver-transmitters, the UHF portion of the recovery
beacon, the command receivers, and the telemetry transmitter. The main EF voice
receiver-transmltter is connected and operating but radiation from this antenna
is negligible. (see Figure 11-6).
11-28. Main Blcone and UHF Recovery Antenna Feed
The various radio systems are connected to the bicone antenna or the UH_
recovery antenna in the following manner: (see Figure 11-13).
(1) The main HF voice receiver-transmltter antenna lead is connected to the
antenna multiplexer.
(2) The IUHF Selector Switch determines whether the hi-power or the lo-power
,:i ¸'
- C,CNF:DENYiX-,.-
OAVE i November 1961
ST. LOUIS 3. MISSOURI REPORT
UEF receiver-traasmitter is used. It also energizes the EEF coax switch to
connect the operating UHF set to the antenna multiplexer.
(3) The two ccmmmnd receiver antenna leads are connected to an impedance match-
ing network. This enables both receivers to share a single antenna lead from
the impedance match to the antenna multiplexer.
(4) The Hi and Lo frequency telemetry tranm_Itters each feed directly to the
antenna multiplexer.
11-29. Antenna Multiplexer
The antenna multiplexer enables simultaneous or individual operation of the
radio systems using one antenna. Effectively this is a radio frequency Junction
box. Final connection to the antenna is through the antenna switch to either
the biconical antenna, or the UHF recovery antenna. The antenna switch is
operated by the antenna fairing separation relay to cause the automatic shift
from the maim antenna to the UEF recovery antenna upon antenna fairing Jetti-
son. (See Figure II-13).
ll-30. Recovery Antenna
An antenna is provided to permit HF radio transmission and reception after
landing. The antenna is a telescoping whip antenna which is automatically
extended by gas pressure after impact. Once extended, the antenna is used for
the HF rescue voice receiver-transmitter and the HI _ portion of the recovery
beacon.
Antenna leads from the HF recovery voice receiver-transmitter and the HF
recovery beacon are connected to the HF recovery diplexer. This diplexer allows
simultaneous or individual operation over the single lead to the antenna. (See
nm, -e ll-Z3).
ll-31. C and S Band Antennas
Three C and S band antenna units are installed in the capsule structure
lz-2?
._.nR, Io4
.....,..,.,, ..r" LL',E_MTI A h.
PAGE, ii- 28
REPORT ._.T_R 104
MODEL l_
•,MC'DONNE oA, REVISED
.... ST. LOUIS, MISSOURI _,-- _"_ _lIT1_ _r-r_ _r_,, ,._ -7_3 B REVISED
_J.l._l.l, ,l..IL_l.4.l. ll .IL
for the C and S-baud beacons. These units are equally spaced about the circum-
ference of the conical section. Each antenna unit consists of one helix as a
C-band antenna and one helix as an S-band antenna.
Antenna leads from the C-band and S-band beacons are routed through indi-
vidual power dividers to the three associated helix antennas. (See Figure ll-13)
Ii-32. SYSTEM UNITS
11-33. AUDIO CENTER
The audio center provides trausistorized audio amplifiers, a "voice operated
relay" (VOX), an audio filter, tape recorder control circuitry and transmitter
control circuitry. (See Figure ll-1). All comments are contained in a light
weight, foam encapsulated unit.
Two fixed gain headset amplifiers are used to bring audio signals up to
headset level and feed the headsets separately. Two fixed gain amplifiers are
provided to increase the dynamic microphone output to a level suitable to be
used with the various transmitters.
A low pass filter, with a cutoff for frequencies above 300 cps, filters the
audio supplied from the command receivers. Outputs from the filter is fed to a
variable gain, ccmmaud audio amplifier.
The "voice operated relay" is a transistorized amplifier with separate
adjustable threshold level and release time controls. The amplifier operates a
relay to provide a grounding circuit for transmitter control. This unit parallelJ
the external microphone switch.
The audio center furnishes a circuit to apply the transmitter control ground
I_tential to the various transmitters. Each circuit is protected from the rest
by a crystal diode.
A relay is installed in the audio center for supplying Imwer and audio
; 7 r_\
//
REVmEO ST. LOUIS S, MmSOUm n_ORT _R lO_
REVISED, __#"*_ _I I_'I r_eli-r I A 1__
signals to the tape recorder. In the de-energized condition, the relay closes
a circuit to the tape recorder input, thus audio received by the capsule is
recorded whenever instrumentation programs tape recorder operation.
When the microphone switch or V0X is operated, the tape recorder relay
is energized. One set of closed relay contacts now completes the recorder
power circuit independent of instrumentation programming, while a second set
of contacts routes signal from the microphone amplifiers to the recorder input.
The circuits in the audio center operate directly from the capsule 2_
volt d-c inputs with no further regulation or voltage increase.
ii-34. CORTROL PANEL
The audio control panel provides controls and _ircults for the audio
signals of the various capsule receivers. (See Figure 11-7).
The two HF and two UHF circuits are routed through individual T-pads to
volume controls. The two EF circuits share a single volume control, the same
is true of the two UEF circults_ while separate volume control is provided
for the command audio circuit. Fixed inputs are used for the alarm tone and
ground interphone circuits.
The panel also contains a switch override for the impact Eeying feature
used with the UHF transmitters, and heylng button on the panel to interrupt
the Hi-Frequency telemetry B+ supply during emergency keying. (Refer to
Figure n-lO).
11-35. MAIN HF VOICE RECEIVER-TRANSMITTER
The main HF voice set is an AM receiver-transmitter designed as a small,
light weight unit operating near 15 MC. (See Figure 11-8).
The receiver section of the unit is a transistorized circuit using a
crystal filter, crystal diode detector and class B audio amplifier. The final
%e_q_ l "l I i li_lmai it i aE =i
PAGE II-_Q
REPORT S_DR 104 ....
MODEL. J._
m
audio amplifier is used for sidetone during transmissions.
The transmitter section of the unit utilizes vacuum tube stages for the
crystal controlled oscillator, driver and power amplifier. The power amplifier
may be modulated up to 90% by a transistorized speech amplifier and modulator.
These audio stages are also used for sidetone. Transmitter output is 5 watts.
Capsule power, 2_ volts d-c, is supplied to the unit, with voltage regulator
provided by a Zener diode, trausistor circuit. Power is routed through an
external switch and contacts of the capsule separation relay which controls
transmitter filament power, relay operation and a transistorized power converter.
High voltage from this converter is used for the transmitter power amplifier.
Antenna switching is accomplished by a solid state circuit which blocks the
receiver during transmission. D-c voltage is also removed from the receiver
RF stages.
11-36. RECOVERY HF VOICE RECEIFER-TRANSMI_I_R
The HF Recovery voice set is similar to the main set. (See Figure 11-8).
The main difference liem in the transmitter section. This section consists of
a crystal controlled oscillator and power amplifier delivering i watt output.
11-37. HIGH POWER(MAIB) Am) LOW PO_ (BACKUP) b'EF VOICE E_CEIVJ_-TZ_NS_
The main and backup sets are identical. (See Figure 11-9). They consist
of an AM receiver-transmitter designed as a mall 3 lightweight unit operating
near 297 MC. Transmitter output is .5 watt. The main receiver-transmitter is
boosted in output by a final booster amplifier.
The receiver section of the unit is a transistorlzed superheterodyne circuit
using a crystal controlled local oscillator, crystal filter and crystal diode
detector. The audio section of the receiver also serves as the speech amplifier,
modulator and provides sidetone for the transmitter. The transmitter section of
(._ ir._ _T T.q= .=..,.-,,.., --..'t._ _.aV _k I A" & 41=F 4K4 g I A A_J4d
OATE i November 1961
REVISED
REVISED,
st. LOUmS S.M'SSOUm nEPOnV SEDR 104
_oo_ 133
the unit utilizes a crystal controlled oscillator, tri_ler and power amplifier.
The RF section uses vacuum tubes while the modulation circuits are transistor-
ized.
Capsule power, 24 volts d-c, is supplied to the set. This voltage is
applied to the receiver, audio circuits and back through an external transmit
switch to an internal power converter. This transistorized converter supplies
B+ voltage to the transmitter RF sections. Transmitter filament voltage is
also applied by the external transmit switch or the bicone separation relay
after bicone separation.
Switching from receiver to tlansmltter operation is accomplished when
ground potential is applied to a switching relay and a blocking circuit. The
relay provides antenna and power converter switching. The blocking circuit
removes receiver voltage.
11-38. UHF 200STER AMPLIFIER
A booster amplifier is used prior to landing to increase the .5 watt out-
put of the Main UHF transmitter to 2.0 watts. The higher power is also avail-
able after landing.
Signal input to the booster is routed through a double pole, double throw
relay. When the relay is de-energized, the signal is routed through the con-
tacts to the output Jack. Energizing the relay feeds the signal through the
amplifier and takes the amplifier output to the output Jack.
11-39. RECEr - CO ERS
The cc_aand receiver-decoder is a transistorized unit consisting of an
FM receiver designed to operate in the frequency range of _O6 to 450 MC, and a
decoder unit to operate control circuits. (See Figure 11-10).
The receiver section of the unit is a dual conversion superheterodyne
PAGE 11-32
REPORT ,_EDR 1011•
MODEL_ 1"_
Mc'DONNEL L..... ST. LOUIS, MISSOURI
DATE 1 November 1961
REVISED
REVISED
e
circuit. The first local oscillator is crystal controlled and uses two stages
of frequency multiplication. Two sta_es of amplification are used for the first
IF, 78 ME signal. The second local oscillator is also crystal controlled, mix-
ing with the first IF and giving a resultant second IF of 10.75 ME. Output from
the IF strip is through a limiter to the discriminator. Audio amplifiers boost
the discriminator output for the co_nand voice channel and the decoder driver.
The driver in turn supplies the ten decoder channels in the set.
The individual decoder channels each provide filters for their specific
ccmmand frequency and amplifiers to operate a double pole, double throw relay
for each channel. The ten relays thus make available normally open and normally
closed contacts for external control circuit operation.
Capsule power, 18 volts d-c, is used to power the set. A Zener diode
circuit, within the unit, is used for voltage regulation.
11-40. AUXILIARY E_CODERS
An auxiliary decoder operates with each of the two receiver-decoder units,
allowing an additional ten channel capability.
The decoder channels in the auxiliary decoder are identical to the decoder
channels of the receiver-decoder, with the exception of the command frequencies
at which they operate.
The auxiliary decoder operates from capsule 18 volt d-c power. No further
voltage regulation or increase is required.
11-_i. _=EMETRY PO_ER SUPPLIES (See Figure ll-ll)
The telemetry power supplies generate voltage used in the telemetry trans-
slitters. The unit is transistorized and uses crystal diodes. Capsule Power,
2_ volts d-c a is applied to a transistor switching circuit operating into the
primary of a power transformer.
• !
'k
/
OATE 1 November 1961 PAOE ii-33
.EVmED ST.LOUm 3,MISSOUm R_o_r S_nR i0_
RL_mED, _ Moo,- 133
A full wave, crystal diode rectifier is used on one secondary, with volt-
age regulation, to provide 200 volts d-c.
11-42. TELEMETRY TRANSMITTERS
The two telemetry transmitters are essentially identical (See Figure ll-ll).
They must be ground adjusted for an output of 3-3 watts. One set is operated
at approximately 228 MC while the other is tuned to approximately 260 MC.
The transmitter is an FM unit using modulation inputs from instrumentation
circuits. Modulation signals are applied to the oscillator which through four
stages of doubling feeds a power amplifier stage. All stages, with the excep-
tion of the final doubler and the power amplifier, are transistorized. Fila-
ment voltage, 6.3 volts and B+ voltage, 200 volts d-c, are obtained from the
separate power supply. Capsule power, 24 volts d-c, is applied to the trans-
mitter which provides voltage regulation.
Ii-43. C-BAND BEACON
The C-Band transponder is a pressurized superheterodyne receiver and pulse
modulated, 400 watt peak output transmitter, operating in the frequency range
of 5400 to 5900 MC (See Figure ll-12). With the exception of the magnetron, the
unit is transistorized. The receiver consists of a pre-selector, local oscil-
lator, 40 MC IF amplifier strip, pulse detector, pulse amplifier and decoder.
Resonant cavities are used for the pre-selector and local oscillator.o
The transmitter section accepts decoder outputs and pulse modulates the
transmitter output.
The unit contains a power supply for converting capsule 24 volt d-c input
to filtered 24 volt d-c and regulated, 115 and 150 volts d-c outputs. Antenna
switching is through an internal dIplexer.
PAGE II-_ZI.
REPORT SEDR IO_
MODEL 133
McDONNE REVISED
ST. I.OUI$, MISSOURI
REVISED_k_l_A1JL JLJt_w41.4J. 1 _L A_LJI_
1 November l_l
iI-_. S-BAND _F_ACON
The S-Band transponder is a pressurized superheterodyne receiver and pulse
modulated, lO00 watt peak output tranmnitter operating in the frequency range of
2700 to 2900 MC. (See Figure ll-12). Like the C-Band beacon, the unit is
transistorized except for the output tube.
Receiver and transmitter circuits are the same as those used in the C-Band
beacon with the exception of the pre-selector, local oscillator and transmitter
which are designed for S-Band frequencies.
xx- 5. /U RECOV YZ ACON (SeeFigu -X2)
The recovery beacon combines an HF, tone modulated, 8.364 MC transmitter and
a UHF, pulse modulated, 243 MC transmitter into one small, foam encapsulated unit.
The UHF section of the beacon is a one tube circuit with a pulse coding network.
The KF section of the beacon is a transistorized crystal oscillator and two stage
power amplifier with tone modulation supplied from a power converter. The beacon
utilizes 6 and 12 volts d-c from the capsule power system. The UNF section is
energized by applying the 6 volt d-c to a transistorized power converter. A
full wave, crystal diode circuit is used to rectify the power converter output
which is applied to theUHF stage. Applying 12 volts d-c energizes the HF sec-
tion of the beacon. No power converter is required for the L_ volt input.
M_dulation for the _ section is provided by routing the 12 volt supply to
the power amplifier stages through a secondary winding of the power con_rter.
11-_6. AUXILIARY UHF RESCUE _ACON
The Aux. UHF Rescue Beacon, consists of a pulse modulated transmitter and
power supply which is enclosed in a foam-encapsulated case. (See Figure 11-12).
The unit is connected to the 6 volt standby bus and has an output of 91 watts.
• /
_vA*_ A _e.f ae._ A
i_!:_¸
oA_ ,! _ovember 1961
REVISED
R L:/3/lSlED
ST. LOUIS S, M,SSOUm R_"ORT _]:_ 104
__f-t-_._____5_v_i "----._T:;,: Moom.__ 133
_U.J__. J I .............. . ......
11-_7. AWI'lm_A MJLTIPLEX_
The antenna multiplexer allows reception and transmission of the many
capsule frequencies over one line to the bicone or UHF recovery antenna. The
unit consists of a number of filters arranged so that all capsule frequencies
between 15 and 450 MC can be multiplexed on the single feed line. Each input
channel is provided 60 db of isolation.
iI-_. RECOVERY DIPLEXER
The recovery diplexer unit is used for the HF recovery voice receiver-
transmitter and HF section of the recovery beacon. One low pass and one high
pass filter is used to diplex 8.364 and 15 MC on one feed line to the HF
recovery antenna.
ll-49. COAXIAL s_n_s (_ s_)
RF switching is accomplished with motor driven SPDT switches. Application
of capsule 24 volts d-c through external circuits drives the switch to the
appropriate RF position and opens the power circuit for that position.
11-50. BICOHE ARTENNA
The capsule is electrically divided in two sections. (See Figure ll-6).
The antenna fairing structure at the Junction of these sections resembles a
discone antenna. This Junction is center fed by a coaxial cable from the
communications sets. At frequencies between 225 and 450 MC the antenna fairing
acts like discone antenna. A lower frequency of 15 MC causes the unit to re-
semble an "off center fed" dipole. Between the upper and lower limits, at 108
MC, the unit behaves as a composite dipole-discone antenna.
Thus the bicone antenna serves all capsule frequencies, with the exception
of C and S-bands, allowing reception and transmission within limits of the cap-
sule system.
_"_ = lll"l II'% I_"I_.I"I" | X I_,n_Vl ,mi . .........
PAGE 11-36
REPORT SEDR 104
MODEL_ 133
Mc'DONNE REVISED
.....St, LOUIS, 1MISSOURL_
ll-51. BICOHE ISOLATOR
Au isolator is provided to shield electrical wires that pass through the
bicone antenna fairing s_ructure. The isolator is formed into a tube which is
curved to allow mounting beneath the periphery of the antenna fairing.
11-52. UHF _ESCENT AND RECOVERY ANTERNA
The UHF descent and recovery antenna takes over the UHF functions of the
bicone antenna when the antenna fairing is Jettisoned. (See Figure ll-6). The
UHF descent and recovery antenna is a fan shaped3 vertically polarized monopole
located on the top of the recovery compartment.
Ii-53. HF RECOVERY ANTENNA
Upon landing, impact circuits initiate a sequence for the HF recovery
antenna. (See Figure 11-6). The elevated antenna acts as a vertically polarized
monopole for HF frequencies.
Three antenna units serve the C and S band beacons. (See Figure 11-6). Each
unit consists of a C and a S band radiator. Each radiator is a cavity mounted
helix antenna.
ii-55. TEST CONFIGURATION CAPSULES
The data contained in Paragraphs ll-i through 11-54 applies to the specifi-
cation compliance capsule. Deviations from this data as applicable to test
configuration capsules are covered in the following paragraphs. Differences
mainly involve such things as UHF Transmitter power output, pre-recorded tape
method of voice modulating transmitters, sequencing differences in transmitter
operation, umbilical control of C and S band beacons. Certain items common to
the manned capsule will not be utilized in some other capsule, such as the
. j/
dp ir_ _v lr:_ lr TI _ !-'--" v '
_/_%••
REVISED. _ro LOUIS 3. MI_|OUIR| REPORT ..... _ ZO_
.w,s_ .oom. 13_
telemetering key, push to talk switch, umbilical interphone, headset amplifier
output and beacon switch modes of operation. (For operation of systems see
Communications Table ll-1). If no reference is made to a particular item, it
is the same as the specification compliance capsule.
11-56. TEST CONFIGURATION CAPSULE NO. 8
Cc_mnication systems on Capsule No. 8 are the same as the Specification
Compliance Capsule except that Capsule No. 8 is intended for an unmanned mission
therefore data pertaining to Astronaut control is either progrmmned within the
capsule or remotely controlled by ground command through the Command Receivers
and De-coders. Differences are explained in paragraphs 11-57 thru 11-67.
(see Fi_res ll-15 t_ru 11-25).
L£-57. Voice Communications
Voice ccemm_nication for Capsule No. 8 is similar to the Specification
Compliance Capsule, except Capsule No. 8 is un-manned and is equipped with a
playback tape recorder containing a pro-recorded tape to simulate the Astronaut 'E
voice which will modulate the NF and UHF transmitters in the capsule. (see
Figures 11-15, i1-19 and 11-22). This playback tape recorder is provided with
a dual track recording. The output is staggered as shown on Figure I_I-14.
Starting of the tape recorder is controlled at the beginning of the mission by
the CAMERA AND TAPE RECORDER Switch in the blockhouse or the additional tape
record relay (see Instrumentation section XIII). The playback tape recorder
operation begins prior to launch and continues throughout the_mission until
tape depletion. Duration of tape recorder running time is approximately 45
minutes. Single ended output from the playback tape recorder is fed to the
two separate microphone amplifiers in the audio center which energizes the
voice operated relay circuit in the Audio Center. The voice operated relay in
turn completes a ground circuit to the tape recorder relay thus causing the
PAGE iI-_8 M_DONN__ __ DATE, 1 November 1961REPORT _ 104, REVISED
._,.tguns, MISSOUmMODEL_ 133 _,u,-_v.,_ .,,.u _ _._,._ .=_-_r.,J REVISED
modulation signal to be routed to the main tape recorder as well as modulating
the HF and UKF transmitters. A push to talk switch is not utilized.
20 Sec. i00 Sec. 20 Sec. i00 Sec.
Audio Silence Audio F"--- Silence
Silence "_' Audio '_-'7 I_-"I00 Sec. 20 Sec.
Silence -----_ Audioi00 Sec. 20 Sec.
FIGURE 11-14 PRE-RECORDED TAPE SEQUENCE
11-58. HF Voice Communications
The H_ C_unication Tape Recorder will transmit the playback recorder
signal alternately 20 seconds "on" and i00 seconds "off" until blcone separation
or until the pre-recorded tape is exhausted, whichever occurs first. Audio out-
put from the HF Receiver is directed into the main tape recorder only, since the
headsets and amplifiers are not utilized. After impact HF Communications is
conducted through the HF Recovery Tape Recorder. (See Figures 11-15 and 11-20).
11-59. UHF Voice Communications
Operation of the UHF _in Tape Recorder (T/R) is similar to the Specifica-
tion Capsule (refer to Paragraph 11-I_) except the UHF Booster Amplifier is not
installed. The transmitter is modulated by the playback recorder and the
receiver output is fed into the main tape recorder similar to the HF Comm. T/R
except that an extra channel inthe main tape recorder is utilized to record the
UHF receiver audio. The UHF Voice Communications set is not utilized unless the
UHF Selector Switch should be placed in the "UHF" position prior to launch. UHF
,i;j
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NOTES
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RANSMIT
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[_ INSTALLED BUT NOT NORMALLY OPERABLE UNLESS
THE UHF SELECTOR ,'SWITCH IS IN THE RESERVE
POSITION.
[_. CAPSULE 8 HAS AN EXTRA TAPE RECORDER INSTALLED.
"" .-"_--_'. / /'-"-UHF DES-'---'--C,ENT--_-
/ C''-- /_ /_/,, / RECOVERY ANTENNA
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PAGE
REPORT
MODE[
11-40
SEDR 104
133
Mc'DONNELL__ __ST.LOUIS. MISSOURI
liiee _ I 1 I i liJ lille i _i • la ilk
DATE 1 NOVEMBER 1961
REVISED
REVISED
¢o_11
_-2 _
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\
P-___
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/
FIGURE li-16 COMMAND RECEIVERS SYSTEM IN3TALLATION (:CAP3ULES 8 _ 9)PM415-t_/_
DATE ,
REVISED
REVISED
I NOVEMBER 1961
ST.LOUIS, MISSOURI
PAGE = 11- 411
REPORT, SEDR 104
MODEL 133
/,
t '
/7.[7; _
FIGURE tl'17 TELEMETRY- SYSTEM (CAPSULE% 8 _ 9)I:_14.5. 189£
l/ _ II I p i I,_lll_, kl_l,, Iii i..lllil alil I • i_ii
PAGE ,, 11-42.
REPORT SEDR 104
MODEL 133
McDONNE_ __ST.LOUIS, MISSOURI
/'_,_ki.lE", _, _Ira"p_ T' A_L__V I I I • I _ t IMIil
DATE
REVISED ,,
REVISED
1 NOVEMBER 1961
\
/
/
..... /
REVISED ST. LOUIS 3, MISSOURI REPORT _'_T_R 104
REVISED MODEL 1_
transmission and reception is available throughout the mission, after impact
the receiver will continue to feed its output to the main tape recorder, end
the transmitter will emlt voice signals until the pre-recorded tape is depleted
after which a CW signal will be transmitted. The CW signal also serves as a
direction finder signal. (See Figure ll-15 and ll-21).
ll-60. HF Recoverer Taps Recorder
HF Recovery Tape Recorder (T/R) operation is similar to the Specification
Compliance Capsule (Refer to Paragraph ll-12), except that after impact the
transmitter will emit the pre-recorded tape voice signal until the tape is
depleted, after which a CW signal will be emitted. The HF Recovery T/R is
connected to an automatically raised whip type antenna. (See Figures ll-15
and ll-20).
ll-61. Command Receivers
The Ground Command Receiver operation is the same as the Specification
Compliance Capsule (Refer to Paragraph ll-16), except the audio is routed into
the main tape recorder instead of into the Astronaut's headset which is not
used. The C and S Band Beacon operation is controlled through the command
receivers. (See Figures ll-16 and 11-22).
11-62. Telemetr_
The two telemetry transmitters perform their functions in the same manner
as in the Specification Compliance Capsules (Refer to Paragraphs 11-18 and Ii-19)
except that operation is continuous from launch to impact. (See Figures 11-17
and 11-23) •
11-63. C and S Band Beacons
Operation of the C and S Band Beacons is the same as the Specification
Compliance Capsule (refer to Paragraphs 11-21 and ii-_2), except that both
beacons operate continuously throughout the mission. The beacon switch must
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PAGE
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MODEL
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REVISED•
REVISED,
Mc'DONNE _A_ _--_1 r
ST. LOUIS _, MISSOURI R_ORT _DR 104
MODEL 133
be placed in the "Continuous" position prior to launch. The S-Band Beacon is
double pulsed as on the Specification Compliance Capsule but the C-Band Beacon
is single pulsed which requires different ground coding for interrogation.
(seeFi e ii-24).
ii-6_. Auxiliar_ UEF Rescue Beacon and Antenna
Capsule No. 8 does not have the Aux. UHF Rescue Beacon installed, however
it is equipped with an ultra Sarah Rescue Beacon which is a self-contained unit,
energized by bicone antenna separation and operation is automatic. (See Figure
ii-65. Antennas
The antennas used on Capsule No. 8 are the same as the Specification
Compliance Capsule (Refer to Paragraphs 11-25 thru 11-31).
11-66. Main Instrument Panel
Capsule No. 8 does not use a tone generator.
11-67. Special Instrumentation
Capsule No. 8 is equipped with special instrumentation and incorporates
an extra capsule tape recorder. This tape recorder is located in the special
instrumentation pallet and consists of a power supply and seven channels for
recording. The extra tape recorder is controlled through the communicati@ns
control panel (Refer to Section XIII Instrumentation. ).
11-68. _ST CONFIGURATION CAPSULE NO. 9
Capsule No. 9 is the same as the Specification Compliance Capsule except
for differences as noted in Paragraphs 11-69 through 11-71. (see Figures 11-15
t rough11-25).
11-69. Voice Ccurunications
Voice communications for Capsule No. 9 is similar to the Specification
Compliance Capsule except a playback tape recorder is installed to simulate the
_J_::V:D=;;Y_AL
PAGE l ] -5?
REPORT ,?_EDR 1OL_
MODEL. 13B
REVISED_**_T,, LOUIS, MISSOURI
Astronaut's voice, as in Capsule No. 8. A Special Instrumentation Relay is
installed to provide playback tape recorder control. (Refer to Paragraph 11-57).
11-70. C and S Beacons
Operation of the C and S Band Beacons is the same as the Specification
Compliance Capsule (Refer to Paragraphs 11-21 and ii-22), except that both
beacons are on continuous. The beacons switch must be placed in the "Continuous"
•position prior to launch.
11,71. Auxiliary UHF Rescue Beacon and Antenna
Capsule No. 9 does not have the Aux. UHF Rescue Beacon installed, however,
it is equipped with an Ultra Sarah Rescue Beacon, which is a self contained unit_
energized by bicone antenna separation and operation is automatic. (See Figure
11-25).
ii-72. TEST CONFIGURATION CAPSULES NO. i0, 13 AND 16
Communications on Capsules No. I0, 13 and 15 are the same as the
Specification Compliance Capsule (Refer to Paragraphs 11-1 thru 11-54).
12-1
S_CTJON Xl_
NAVIGATIONA.L AIDS
TABLE OF CONTENTS
\4
i i :
:::::::::::::::
..................;!iiiiiiiii!i!_!_iii!
::'::,
':i:.
TITLE PAGE
GENERAL ............................................. 12-4
PERISCOPE .......................................... 12-4
SATELLITE CLOCK ................................ 12-14
EARTH PATH INDICATOR .................... 12-17
ALTIMETER ........................................... 12-18
LONGITUDINAL ACCELEROMETER ....... 12-18
HAND COMPUTER ............................... 12-18
ATTITUDE RATE INDICATOR ................ 12-26
NAVIGATIONAL AID KIT ................... 12-27
TEST CONFIGURATION ...... ._................ 12-27
PAGE 12-2
REPORT $EDR 104
MODEL, 13 3
ST.LOUIS, MISSOURI
CC_ .".' F ' _ "- "''_-'_"
DATE 1 NOVEMBER 1961
REVISED
REVISED
CAUTIONI.ALLOW 5 M_NUTE_ COOL-
IN_ TIk_E FOR MOTOR AFTEt
E/kC_ ETEND-RETRACT CYCLE
Z, IN_URE MOTOR _% '_TOPPED
BEFORE RE_ERS_ I_
EXTE_D-RE_RACT CYCLE
WARNINGINSURE ENGAGE LEVER
UP BEFORE PERISCOPE
POWER APPLIED.I I "
UMBILICAL (REF.)
FIGUR£ I?.-IPER_SCOPL ASSEmbLY (SHEET I:OF2)
_ ,_llr I RI I
PM4,5-110C-I
, -.- )
DATE
REVISED
REVISED
1 NOVEMBER 1961 " MCDONNELL,>_ __ST.LOUIS, MISSOURI
___"
PAGE
REPORT
MODEL
12-3
SEDR 104
133
r
_.TURAL
FIGURE 12-'iPERISCOPF..ASSEMBLY (_)HEE'T2 OF 2)Dk445- __O-;:'
PAGE 12-_
REPORT SEDR 101_
MODEl 133
MCDONNE REVISED
ST. LOUIS, MISSOURI
__ ,_,_ _T _..-, _ .... REVISEDirl
XII. NAVIGATIONAL AIDS AND INSTR_S
12-1. GENERAL,,, ,
Normally the Astronaut will not find it necessary to compute any of the
factors relative to his flight or landing. In the event the need should arise,
however, the Astronaut is provided with all the equipment required to compute
altitude, course, velocity and landing data, and to attain and maintain the
proper attitude for each phase of the flight.
12-2. PERISCOPE DESCRIPTION
The periscope is a compact navigational instrument designed to withstand
loads up to 100 G's. The periscope consists of three major assemblies: The
display assembly, the upper housing assembly, and the lower housing assembly.
(See Figure 12-1.) The periscope is mounted in the capsule so that the objective
cartridge may be extended and retracted through a periscope door opening in the
bottom of the capsule. The opening and closing of the periscope door is
synchronized with the extension and retraction of the objective cartridge by
means of a mechanical link. Periscope progra_ning, during a normal mission,
begins with the periscope extended while the capsule and booster are still on
the pad. Upon capsule umbilical disconnect, the periscope retracts and remains
retracted until capsule separation. At this time the periscope extends and
remains extended during orbit. Thirty seconds after retro-package separation,
the periscope is again retracted. At approximately 10,000 feet altitude, the
periscope is extended for the last time and remains extended thronghout impact.
Following umbilical disconnect and Just prior to launch, if the "HOLD" sequence
is initiated, the periscope will automatically extend. Located on the left
hand console is the RETRACT SCOPE telelight and a switch which has AUTO and MAN
!/ !
-i'Ir .... m m_v_, _ _AD_, _ l_ _
DATE 1 NOVEMBER 1961
REVISED
REVISED
Mc'DONNEL_L-_ __ST.LOUIS , MISSOURI
PAGE
REPORT
MODEL
12-5
SEDR 104
'133
;ROLL
®
®
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KNOB
FIGU_'E:]Z'Z T_.I_cSC--OPEDISPLAY (,SHEET iIOF"7--SHEETS)i
PM4$-II ZA'I
PAGE.
REPORTMODEl
12_!6
SEDR104i ,
i 133i
Mc,DONNE_ST.LOUIS. MISSOURI
DATE 1 NOVEMBER 1961
REVISED
REVISED
©
7 °4
P-t ,--._
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5* LEFT ROLL
EXAMPLE NO.1.
©
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CAPSULE 5MALLENDPITCHED 5 Q BELOW ORI_IT ATTITUDE
E_AMPLE NO.,_.
FIGUI:2E IZ-?_
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CAPSULE LEVEL
EXAMPLE NO.2.
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CAPSULE_ IN _ETROGIRADEATTITUDE
EXAM F:::)LE NO. 4-.
PEQISCOPE DISPLh,_( (SNEF_.T. 2 OF 2)
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RL'VlS_O ..... ST. :.OUlS 3, MISSOUm RmPORV SEDR 104
positions. If the periscope does not retract _0 seconds after retro-package
separation, the RETRACT SCOPE telelight will go on. The switch is then placed
in the MAN position and the periscope is retracted manually. The following
navigational information can be obtained concerning the capsule's position
relative to the earth: Drift, altitude, pitch, roll, true vertical, retro
angle, field of view of the earth-sky camera and relative bearings of the sun
and moon.
12-3. DISPLAY ASSEMBLY
The display assembly (See Figure 12-1 and 12-2) includes the following:
Scales, indicators, altitude reticles, attitude reticles and the controls for
their manipulation, earth image, horizon image and the retract pilot light. It
is mounted at the top of the upper housing assembly. The display assembly pro-
vides the Astronaut with a visual indication of the capsule's altitude and
attitude relative to the earth. It also provides him with the necessary indi-
cators and scales for obtaining bearings relative to the sun and moon. When
the earth is centered in the display, the portion of the image bordering the
altitude reticles will depict the earth's horizon. The center of the view
(crossline reticle) represents a vertical line to the center of the spherical
earth (see Figure 12-3). Image erection is such that the horizon and straight
down views are true.
12-_. True Vertical and Optlcal Vertical
In the true vertical attitude (see Figure 12-3), as compared to the
optically vertical attitude, the longitudinal axis of the capsule is perpendicular
to a line through the center of the spherical earth. The capsule is optically
vertical with respect to the earth when the capsule is pitched nose down at l_
degree amgle. _-
L w W,m g mR. _ ..... ._ . .. i_.JL--|_l Jl. |f'_im
PAGE,
REPORT
MODEL
12-8
SEDR 104
133
MCDONNE ST.LOUIS. MISSOURI
DATE
REVISED
REVISED
I NOVEMBER 1961
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REVISED ST.LOUIS3,MISSOURI REPORT,SEDR 10_
REVISED MODEL 10_
12-5. Altitude
When the capsule is in the optically vertical attitude, the altitude of the
capsule may be determined by means of the altitude mechanism. This mechanism
includes four pairs of altitude reticles, which form a square within the circular
display area, the altitude knob and the altitude indicator (see Figure 12-2).
Turning the altitude knob moves the reticles to change the size of the square
and also rotates the scale of the altitude indicator. To determine altitude,
the knob Is turned to change the size of the square until the earth's image is
inscribed within the inner square formed by the reticles. The altitude is then
read from the altitude indicator. The point on the earth's surface vertically
below the capsule is indicated by the crossline reticle which appears at the
center of the display area. The crossline reticle is illuminated by a lamp,
the brightness of which is controlled by the reticle knob. The altitude indica-
tion range extends from 50 to 250 nautical miles. The altitude mechanism is
calibrated to indicate altitude within +5 nautical miles within the range of lO0
to 140 nautical miles, within +lO nautical miles at altitudes below lOO nautical
miles and within +50 nautical miles at altitudes above l_O nautical miles.
12-6. Pitch and Roll
The pitch and roll degree of the capsule with respect to its optical vertical
attitude may be approximated by use of the altitude mechanism (see Paragraph 12-5)
The distauce between the parallel curves which form each side of the altitude
reticle square is equivalent to a 5-degree angle at an altitude of ll5 nautical
miles. If the vehicle is off pitch, it will be impossible to adjust the reticle
square so that both the fore and the aft reticle curves (reticle curves relative
to the 180 ° and 0° positions on the drift scale) are tangent to the earth image.
If the vehicle is off in roll, it will similarly be impossible to make the left
and right reticle curves tangent to the earth image. To determine the approximate
--_AL ,,-, ,-_z-_l-_ i A !
REPORT SEDR 10_ REVISED"$T:"_UI$,, ,_MlUQU_I_. _,,,
MODEL I_ . -- C/___._.__T_W_.TTT._.T. REVISEDF
pitch angle, the reticle square is adjusted so that the horizon image lles across
either the fore or aft (depending upon the direction of pitch) inner reticle curve
by approximately the same amount that the opposite horizon edge lles inside the
opposite inner reticle curve (see Figure 12-2, Sheets 1 and 2). The degree of
pitch is estimated by comparing the distance that the one side of the horizon
image extends across the inner curve with the distance between the inner and
outer curve. The approximate roll angle is determined Im the same manner as
pitch angle.
12-7. Drift
Vehicle drift is determined by reference to five parallel drift reticles
etched on the face of the drift plate covering the circular display area (see
Figure 12-2 3 Sheet i). The drift plate is rotated by means of the drift knob.
A clear plastic drift scale is mounted under the drift plate. The drift scale
covers plus and minus 5 degrees of drift in increments of one degree. To
determine drift, the periscope is set for high power magnification. The Astro-
naut then rotates the drif_ knob until the ground track of the capsule in the
central high power view appears to parallel the drift lines. The drift of the
capsule is then indicated by the position of the center drift line with respect
to the drift scale.
12-8. RetroGrade An_le
The retrograde reticles provide an indication of whether or not the capsule
is in the proper retrograde attitude prior to re-entry. To achieve proper retro-
grade angle starting from the optically verticle attitude at an altitude of 115
nautical miles aud with zero drift angle, the nose of the capsule must be pitched
down until the horizon image at the bottom of the display is taugent to the three
retrograde reticles which form an arc across the lower half of the display area
(see Figure 12-2_ Sheet 2).
i[
: 7::_
LT_ LT-_ 1LT 5r_ T _ _ ..... _
-... vJ._ A" A_A_i, A A_L
DAT_
REVm_D. Srr.'--_ "'-'0""_LOUlS3, MmSOUm R_POnV _g 104
REV,SED .... MODm.. 133
.+
12-9. Sum-Moon Index
A ring shaped sun-moon index scale, used to indicate relative bearings of
sighted objects, is mounted so as to frame the circular display area (see Figure
12-23 Sheet l). The index ring is mauually rotated by means of a finger-grlp
control. The index scale is calibrated from zero to 360 degrees. A sum index
indicator, which enables sighting on the sun to obtain bearings, is mounted at
the zero degree position of the sun-moon index scale. A moon index indicator,
which enables sighting on the moon for bearings 3 is mounted at the 180 degree
position of the sun-moon index scale. Bearings are read off the sun-moon index
scale at the sun-moon bearing index mark located on the display assembly directly
in front of the Astronaut.
12-10. Earth-Sky Camera View
Four camera reticles located at the right center of the display area give
the approximate outline of the field-of-view of the earth-sky camera which is
included in the capsule equipment (see Figure 12-2, Sheet 1).
12-11. Orbit Velocity
The crossllne reticle is used to aid in the computing of orbital velocity
(see Figure 12-2, Sheet 1). A stopwatch is started as the first check-point
passes under the crossline reticle. When the second checkpoint passes under
the crosslime reticle, the watch is stopped. With this elapsed time, the orbit
velocity is computed by employing the satellite hand computer.
12-12. UPPER HOUSI_D ASSEMBLY
The upper housing assembly will incorporate the following functional com-
ponents: Two periscope mirrors, a filter installation, a magnification change
control mechanism, a manual extension-retraction control, a housing exhaust
valve, a desiccator assembly which includes the housing intake valve and the
housing purging valve.
_-"_""'"_"'*''! A I,_ll ma • i,=_'llll 'ql • il .__
PAGE, _,2-_2
REPORT SEDR 104
MODEL_ l°,_
_ DATE I November 19(31
St. LOUIS, MISSOURI
12-i3. Periscope Mirrors
The mirrors are so situated in the upper housing assembly as to present the
earth's image at a convenient angle to the observer (see Figure 12-_).
12-14. Filters
The upper housing assembly contains a clear, a red, a yellow and a medium
neutral density filter. The filters are mounted in a rack which can be rotated
to position the desired filter in the optical path. The filter rack is manually
driven through a system of pulleys and cables by means of the filter selector
which is located on the left side of the upper housing assembly (see Figure 12-1,
Sheet I of 2). Detents accurately position the filters with respect to the
optical path.
12-15. Ma_ification Chan6e Control
The periscope optic system is capable of high and low magnification. The
change in magnification is brought about by manipulation of the magnification
changer (see Figure 12-1, Sheet I of 2).
12-16. Manual Extension-Retraction Control
The manual extension-retraction mechanism enables the objective cartridge to
be extended and retracted manually (see Figure 12-1, Sheet i of 2). The mechanism
consists of a manually driven gear system which couples to the gear box and motor
drive assembly of the lower housing assembly through a manually operated clutch.
The gear system is driven by the ratchet handle on the right side of the upper
housing assembly. The ratchet handle is pulled up to place it in operating
position and pushed down to place it in the stowed position. For manual manipu-
latic_, the manual engage lever will be in the down position. The ratchet pawl
arm will determine the direction of motion (see instructions affixed to naviga-
tional aid kit).
12-17. Intake and Exhaust Valve
The intake and exhaust valve (see Figure 12-1, Sheet 2 of 2) allows the
upper housing to "breathe" without the passage of dirt to the inside of the
upper housing. The intake valve is mounted in an intake valve body to which the
intake end of a desiccant tube is attached. When the ambient cabin pressure
exceeds that inside the upper housing' the intake valve admits air to the upper
housing through a desiccant (silica gel). Moisture and dirt is thus removed
from the entering air to minimize condensation and dust deposits on optical
components. A purging valve (see Figure l, Sheet 2 of 2) provides an inlet and
an attachment point when purging the upper housing assembly with dry nitrogen
gas.
12-18. LOWER HOUSING ASSEMBLY
The lower housing assembly consists of a structural shell housing the
objective cartridge. The following items are mounted on the structural shell:
Extension-retraction motor, gear box, lower limit switch assembly and a reticle
illumination assembly (see Figure 12-1, Sheet 2 of 2).
12-19. Objective Cartridge
The objective cartridge houses the objective lens, power change lens and
the collective lens (see Figure 12-4). It is mounted inside the structural shell.
A connecting link between the objective cartridge and the periscope door allows
the door to move in unison with the telescopic motion of the objective cartridge
(see Figure 12-1).
12-20. Objective and Collective Lens
Wide-angle objective lens provides a wide field of view (approximately 180°).
Light gathered by the objective lens passes through the collective lens, where
the first image is formed (see figure 12-_).
MCDONNE REPORT SF.,DR'lOb.. REVISED
ST. LOUIS, MISSOURIMODEL 1,_,_ '_-_Jr%_,T117T T_T:_...... REVISED
12-21. Power Chaz_e Lens
The power change lens (see Figure 12-4) can be moved in and out of the
optical path by means of a mechanical linkage.. The mechanical linkage is actuated
by the magnification changer (see Figure 12-1). For low power periscope operation
the power change lens is moved out of the optical path; while for high power
operation, the lens is moved into the optical path.
12-22. SATELLITE CLOCK
The specification satellite clock is an electro-mechanical timing device
located above and to the right of the periscope display assembly. The satellite
clock will indicate time of day, TIME FROM LAUNCH, TIME TO RETROGRADE and RETRO-
GRADE TIME (see Figure 12-5). The time of day will be reflected by a manually
wound spring-driven movement watch. The manually wound watch is located in the
upper left-hand corner of the satellite clock. Time From Launch, Time To Retro-
grade, and Retrograde Time will be displayed on drum counters (digidial). The
drum counters will indicate time in hours, minutes and seconds. The time
elements will move in one step increments. The Time To Retrograde digidial will
be supplemented by a telelight, located in the upper right-hand corner of the
satellite clock, which will illuminate 5 minutes prior to retrograde time; in
addition to the telelight, an aural signal to the Astronaut's headset is
initiated i0 seconds prior to retrograde time. The satellite clock is auto-
matically started by 28V d-c power at liftoff. Should this not occur, a push
button switch is provided above and adjacent to the clock to allow the Astronaut
to energize the clock (aud maximum altitude sensor) manually. The retrograde
time is normally computed and set prior to flight, but the retrograde time can
be manually changed by the Astronaut (retrograde time reset handle) or remotely
set through the cce_nand receivers. Ten minutes prior to retrograde time, the
)_J
DATE
REVISED
REVISED
1 NOVEMBER 1961
ST.LOUIS. MISSOURI
PAGE.
REPORT
MODEl
12'15
SEDR 104
133
DISPLAY ASSEMBLY
UPPER HOUSING ASSEMBLY
MIRROR
LOWER HOUSING ASSEMBLY
OBJECTIVE CARTRIDGE
POWER CHANGE LENS
FIGURE 12-4 OPTICAL SYSTEM SATELLITE PERISCOPE
DRIFT LINE PLATE
FIELD LENSES
MIRROR
FILTERS
ERECTOR LENS
COLLECTIVE LENS
OBJECTIVE LENS
PM45-262
PAGE
REPORTMODEL
12-16
SEDR 104
133
ST.LOUIS, MISSOURI
DATE 1 NOVEMBER 1961
REVISED
REVISED
' POSITION RELATIVE TO EARTH
TIME OF DAY
IEARTH PATH INDICATOR
ELELIGHT
,'TYP .)
SATELLITE CLOCK
TIME P_SET HANDLE
I FIGURE 12-_5 i EARTH PATH INDICATOR AND SATELLITE CLOCK PM4._-II3E
VVl 11 ni.tl---,I I i l/'_i,IL-
DAT_, ,iNovember 1961
R£V|SI[D,
R£VIS£D
ST. LOUIS 3, MISSOURI
PAG_ 12-17
mm=ORV, SEDR 10b,
MOmta. 133
satellite clock transmits a signal to the ASCS to start horizon scanners operating
continuously and assures rate gyro operation in preparation for retro sequence.
Upon reaching retrograde time, a set of contact points within the clock close,
initiating the retrograde sequence. Time From Launch and Retrograde Time
digidials provide outputs for telemetering.
12-23. EARTH PATH INDICATOR
The earth path indicator (see Figure 12-5) consists of a spherical map
(globe) of the earth gimballed and rotating in a manner to indicate ground
position under the capsule. The indicator is spring motor powered and is
capable of running 20 hours without re-windlng. The globe, which is approxi-
mately 3.85 inches in diameter, will display the following geographical features:
(i) All continents
(2) All bodies of water having major dimensions of 300 statute miles
(3) The sixteen largest rivers of the world
(4) All islands having major dimensions of 500 statute miles
(5) All known islands or island clusters separated from continents
by 300 statute miles and having major dimensions less than 500
statute miles shall be identified by an .020 diameter circle.
(6) The fifty largest cities of the world are identified by .020 dots.
(7) 15° latitude and longitude lines are presented and numbered.
Controls are provided on the face of the indicator to wind the spring motor
and to adjust the orbit time, adjust orbit inclination and to slew the globe
about the earth and the orbital axis. The touchdown area is displayed as a
rectangle and the luminous dot inside of the rectangle being the point of impact.
The landing area is 3040 nautical miles ahead of instantaneous orbital position
above the earth as indicated by the four ring bullseye. The instrument is
externally lighted by cabin floodlights.
T.L;4Fi_V.i,iT;AL
PAGE 12-18
REPORT S_nR Ina
MODEL_ i_
Mc'DONNE ,REVISED
ST. LOUIS, MISSOURI_,-_r%]%TW_Tw-_W_T.n. • T REVISED
12-24. ALTIMETER
The altimeter visually indicates external pressure above sea level. (See
Section I for location.) It is a single revolution type, calibrated from 0 to
i00,000 feet, with a marker at I0,000 (MAIN) and 20,000 (SNORKEL). Static
pressure is obtained from a centrally located plenum chamber which connects to
four static ports spaced equally around the small end of the capsule conical
section. The instrument is lighted externally by cabin floodlights.
12-25. LONGITUDINAL ACCELEROMETER
The accelerometer is a self-contained unit which is housed in a hermetically
sealed inclosure. (See Section I for location.) The accelerometer is designed
to indicate acceleration in the range 0 to -9 and 0 to +21 g units (acceleration
of 32.2 feet per second per second). Attached to the face of the accelerometer
are three pointers. One pointer will indicate instantaneous acceleration. The
remaining two pointers are memory pointers. One memory pointer will record
positive acceleration and the other will record negative acceleration. The
memory pointers will incorporate a ratchet device which will maintain a deflec-
tion until they are reset by means of a reset knob which is located in the lower
left hand corner of the accelerometer.
12-26. HAND COMPUTER
The satellite hand computer (see Figure 12-6), physically resembling a
circular slide rule, is provided to aid the Astronaut in solving navigation
problems. The computer consists of three discs; a basic fixed disc, a small
top fixed disc and a rotating intermediate disc. The computer is stowed in the
Astronaut's map case. The computer may be used to find orbital tangential
velocity, orbital angle, drift, true ground speed, indicated ground speed, and
for multiplication, division and proportions. Operation of the computer is as
d_ ink ?LT _ T lr'k lr_ _v r_r_ v *
DATE 1 NOVEMBER 1961
REVISED
REVISED
M 'DONNELL ' ST.LOUIS, MISSOURI
PAGE,
REPORT
MODEL
12-19
' /
SATELLITE (_
'_ HAND COMPUTER
USI_ NUM_RStN (")
FIGURE 12.-8 Z_,NE) CO_P_TF-.Q.Pk445 - 115
PAGE 12-20
REPORTSED: lo4
MODEL_ 1,3_
REVISEDST. LOUIS, MISSOURI
-_ _',-,,"_"_'r_ **'_ ...._ REVISED
follows :
12-27. Orbital Tangential Velocity
A. To compute orbital tangential velocity when ground distance, altitude
(nautical miles) and time (minutes) are known, proceed as follows:
(a) Set the time on scale B opposite the ground distance on scale A.
(b) Read orbital tangential velocity in nautical miles per minute on
scale C opposite the altitude.
(c) Read orbital tangential velocity in knots (nautical miles per hour)
on scale A opposite the altitude (at 6.0 on scale B).
I.
Be
Sample Problem
Given:
ground distance
time
alt it ude
Re quired :
= 2880 nautical miles (N.M.)
= 12 minutes
= 120 N.M.
Orbital tangential velocity in N.M. per minute and
orbital tangential velocity in N.M. per hour.
Operation:
(I) Set the time 12 (1.2) minutes on scale B opposite the ground
distance 2880 (2.88) N.M. on scale A.
(2) Read orbital tangential velocity 248.39 N.M./min. on scale C
at 120 on the altitude scale.
(3) Read orbital taagential velocity 14,903.4 N.M./hr. on scale A
opposite 12Oon altitude scale (at 6.0 on scale B).
To compute orbital tangential velocity when angular position in degrees,
time in minutes sad altitude in nautical miles are known, proceed as
follows:.
(a) Set the time on scale B opposite angular position on scale A.
v// ¸
......rr-,---.... TIA '-
REVISED. ST, LOUIS 3, MISSOURI RKPORT, SEDR 10_
REVISED ......... MODEL. 133
i
Ci
(b) Read orbital tangential velocity in nautical miles per minute
(N.M./rain.) on scale D.
(c) Read orbital tangential velocity in N.M./hr. on scale A opposite
altitude on altitude scale (at 3.6 on B scale).
II. Sample Problem
Given :
angular position
time
altitude
Required :
= 48 degrees
= 12 minutes
= 120 N.M.
Orbital tangential velocity in N.H./min. and orbital
tangential velocity in N.M./hr.
Operation :
(1) Set the time 12 (1.2) rain. on scale B opposite angular position
48 degrees (4.8) on scale A.
(2) Read orbital tangential velocity 248.39 N.H./min. on scale D.
(3) Read orbital taugentlal velocity 14,903.4 N.M./hr. on scale A
opposite 120 on the altitude scale (at 3.6 on scale B).
To compute orbital tangential velocity when ground speed in nautical miles
per minute (N.H./min.) and altitude in nautical miles are kno_ra, proceed
as follows :t
(a) Set the altitude zero (0) on the E scale opposite the ground speed
on the scale A.
(b) Read orbital tangential velocity on scale A opposite altitude on
scale E.
IIl. Sample Problem
Given:
ground speed = 239.6 N.M./rain.i
PAGE 12-22
REPORT _..oI_ 104
MODEL. 133
REVISED .....ST. LOUIS, MISSOURI
_'_ .',,,..._ REVISED ....
altitude
Required :
Operation:
= 120 N.M.
Orbital tangential velocity in N.M./min.
(i) Set altitude zero (0) on scale E opposite the ground speed
239.6 (2.396) N.M./min. on scale A.
(2) Read the orbital tangential velocity 248.00 N.M./min. on scale A
opposite altitude 120 N.M. on scale E.
NOTE
Conversion of orbital tangential velocity to
ground speed is the same as the above procedure
but in reverse. Set altitude on scale E opposite
orbit tangential velocity on scale A then read
ground speed on scale A opposite zero (0) altitude
on scale E.
Orbital _le
To compute orbital angle when grotmd distance in nautical miles (N.M.)
is known, proceed as follows:
(a) Set 1.O of scale B opposite the degree index mark (located at 1.67)
of the A scale.
(b) Read orbital angle on the A scale opposite the distance on the B scale.
I. Sample Problem
Given :
ground distance = 2880 N.M.
Required: Corresponding orbital angle
Operation:
(1) Set 1.O of B scale opposite degree index mark (located at 1.67)
of the A scale.
,L/',
DATE i November 1961
R_VlSED
RF-.'VIS£D
MCDONNELL.ST. LOUIS 3, MISSOURI
PAGE 12-23
Rm_o,_T _oR 104
MOOre. 133
,k,
_j
B.
(2) Read equivalent orbital angle of 48 degrees on scale A
opposite the distance 2880 (2.88) N.M. on the B scale.
NOTE
The procedure is the same for finding
ground distance when the orbital angle is
known. Set 1.O on B scale opposite degree
index mark on A scale then opposite to any
degree on the A scale the equivalent ground
distance is read on the B scale.
To Calculate orbital angle when latitude and orbital inclination are known,
proceed as follows:
NOTE
The G scale has two sets of numbers. The
plain numbers, increasing clockwise_ are
used for computimg drift while the numbers
in parenthesis, increasing counterclockwise,
are used for cc_puting orbital angle.
(a) Set zero (0) of G scale opposite latitude on F scale.
(b) Read orbital augle on G scale opposite orbital inclination on F scale.
If. Sample _roblem
Given:
ground latitude = 14 degrees 29 minutes
orbital inclination = 30 degrees
Required: Orbital angle
Operation:
(1)
(2)
Set zero (0) of G scale opposite latitude 14° 29' om F scale.
Read orbital amgle (30°) on G scale opposite orbital inclina-
PAGE 12-24
REPORT SFZ)R lO4
MODEL_ IBB
REVISEDST. LOUIS, MISSOURI
"-_v _ _ _,.._._,_, REVISED
C.
De
(a)
(b)
12-29.
A.
tion 30° on F scale.
To compute latitude when orbital angle and orbital inclination are known,
proceed as follows :
(a) Set orbital angle on G scale opposite orbital inclination on F scale.
(b) Read latitude on F scale opposite zero (0) on G scale.
To compute angle of inclination when latitude and orbital angle are known,
proceed as follows:
Set zero (O) of G scale opposite latitude on F scale.
Read orbital inclination on F scale opposite orbital angle on G scale.
Drift
To compute drift when orbital inclination, orbital angle in degrees and
orbital velocity in degrees per minute are known 3 proceed as follows:
(a) Set the orbital angle on the G scale opposite the angle of inclina-
tion on the F scale.
(b) Read the drift in the lower window opposite the orbital angular
velocity on the I scale.
I. Sample Problem
Given:
orbital inclination = 32°
orbital angle - 0O
orbital angular velocity = _°/min.
Required: Drift
Operation:
(I) Set orbital angle 0° on the G scale opposite 32° angle of
inclination on the F scale.
(2) Read drift 1.897° in lower window opposite 4° orbital angular
velocity on the I scale.
4
J
_J_'A JLJLJJL_I" A JL_IL]L_
"4
REVISED ST. L.OUIS 3, MISSOURI RI[PORT SEDR 101r
REVISED ._ _ MO0_. 133
12-30. Orbital Ground Speed
A. To compute orbital ground speed, we must first find the effective component
of the earth's tangential velocity and then add it to the indicated ground
speed. To find the earth's tangential velocity component, the following
must be known: Orbital inclination, orbital angle and the earth's tangential
When the above information is available, proceedvelocity at the equator.
as follows :
NOTE
In the following presentations, the F and
G scale is not used according to their desig_
nation. They are to be used as directed.
(a) Set the orbital Imclinatlon on the G scale opposite 90 degrees (zero)
degrees) on the F scale.
(b) Read effective component of earth's tangential velocity on the A scale
opposite the earth's tangential velocity at the equator on the B scale.
(c) Orbital ground speed will be indicated ground speed plus effective
component of earth' s tangential velocity.
I. Sample Problem
Given:
earth's tangential velocity at the equator =
orbital inclination
orbital angle
indicated ground speed
Required : Orbital ground speed.
Operation :
(1)
15 ..../rain.
= 30 degrees
= 0 degree
= ..M./min.
Set the orbital inclination 30 degrees on the G scale opposite
orbital angle zero degrees (90 degrees) on the F scale.
t#1_/i_iF i1.11_il • l* le,r_-
PAGE 12-26
REPORT S_ 104
MODEL_ I,,_
,MCDONNE ST. LOUIS, MISSOURI
DATE 1 November l C_l
REVISED.
REVISED
(2)
(3)
Read effective component of earth's effective tangential velocity
12.99 (1.299) N.M./mln. on the A scale opposite the earth's
tangential velocity at the equator 15 (1.5) N.M./min. on the B
scale.
Orbital ground speed-- indicated ground speed (240 N,M./min.)
plus effective component of earth's tangential velocity (12.99)
or 252.99 N.M./min.
NOTE
Effective earth velocity is constant;
orbital ground speed for all practical
purposes can also be considered constant
for any given orbit.
12-31. ATTITUDE-RATE INDICATOR
The Attitude-Rate indicator is a three axis angular rate and attitude indi-
cating system located approximately at the top center of the main instrument
panel. It is designed to indicate attitude and the rate of change of attitude.
The unit is a composite arrangement consisting of a rate indicator around which
are positioned a roll attitude indicator, a yaw attitude indicator and a pitch
attitude indicator (see Section I, Figure 1-15). The rate indicator will display
three pointers. The rate of roll pointer is the pointer which is parallel to the
pointer of the roll attitude indicator. The rate of yaw and rate of pitch
pointers are pointed towards the yaw and pitch attitude indicators respectively.
All components are completely interchangeable. The failure of one component will
not necessitate recalibration or replacement of allied components. The attitude-
rate indicator is activated by pitch rate, roll rate and yaw rate transducers.
Each transducer consists of a gyroscope, an amplifier and a demodulator. These
J._ A A_ILJIm
___ImmJeiOMbm_
REVISED ST. LOUIS 3, MISSOURI REPORT. SERE lO_
REVIs_o - L _ Moo_ 133
/,r ,
components function together to produce a d-c output signal proportional to the
input rate of change of attitude.
12-32. NAVIGATIONAL AID KIT
The navigational aid kit consists of a neoprene coated nylon case, a binder
assembly and a computer board. It is mounted to the periscope directly below
the circular display area (see Figure 12-1, Sheet 1 of 2). The binder assembly
consists of a number of index cards. The index cards will be used to file check
lists, rate cards and navigational charts that shall be provided as required by
the particular capsule mission. The following Items are attached to the binder
assembly: Pencil holder, mechanical pencil and nylon retention springs. The
pencil holder is fabricated from neoprene coated nylon and is sewn to the case.
The mechanical pencil is secured to the binder assembly by means of a nylon
retention spring. A nylon retention spring also secures the binder assembly to
the neoprene coated nylon case. A computer board, which is constructed from the
hand computer (see Figure 12-4), forms a part of the navigational aid kit.
12-33. TEST CONFIGURATION CAPSULES
12-34. TEST CONFIGURATION CAPSULES 8, 9, i0, 13 AND 16
12-35. General
Capsules 8, 9, I0_ 13 and 16 are fundamentally the same as the specification
capsule. Some differences will exist in the location of the various instruments.
(See Section I for instrument panel illustration. ) Other differences are
enumerated in the following paragraphs.
12-36. Periscope
Capsule 8 periscope system will not incorporate a RETRACT SCOPE telellght
and the associated MAN-AUTO periscope retract switch. Capsules 9_ 13 and 16 will
have provisions for telemetering periscope door closure. Before ground operation
vvx_. m_awa ,mm zm mkm
REPORT _DR 104 ST. LOUIS, MISSOURI REVISED
MODEL. 133 ---L'_ .NF'_ E N,_. REVISED
of periscope, observe operational instructions attached to the navigational aid
kit (see Figure 12-1). Equipment in the periscope upper housing assembly in
Capsule 9 has been removed to faciliate installation of an internal camera
(see Figure 12-7).
12-37. Altimeter
The altimeter used in Capsule 8, 9, i0, 13 and 16 have markings at the 20,00C
and I0,000 foot levels to indicate when the snorkel valves and main parachute,
respectively will actuate. Along the outer edge of the instrument dial starting
at 0 feet and advancing to 28,000 feet are psia indications which are as follows:
0 feet 15 psia, i0,000 feet i0 psia and 28,000 feet 5 psia.
12-38. Longitudinal Accelerometer
Capsules 8 and 9 will not have memory pointer or a reset mechanism. The
location may vary with each capsule depending on planned mission. (See Section I
for instrument panel illustration. )
i2-39. Hand Compute.r
Only manned capsules will be equipped with the hand computer.
12-40. Attitude-Rate Indicator
NOTE
External appearance of attitude-rate indicator display is the same for all
capsules except for color of pointers and as follows:
On Capsule 9, the pitch dial of the display mates with the zero pitch rate
index at -43 degrees.
On Capsules 8, i0, 13 and 16, the pitch dial of the display mates with the zero
pitch rate index at -34 degrees.
12_41. Navigational Aid Kit
Capsules 8, 9 and 13 will not contain a navigational aid kit.
.....i,i mil i i _wJk _
DATE ,
REVISED
REVISED
I NOVEMBER 1961 M_'DONNE__ __ST.LOUIS, MISSOURI
PAGE 12_29
REPORT SEDR 104
MODEL 133
\4 __
(_ ,_"_"%_ UPPER PERISCOPE
f ""d) _ f HOUSING ASSEMBLY
V %..
CLOCK---...._ / - t+_7_-,__
, ............ _,__+_-_: + ,_,,_+,
MIRROR LENS j
WINDOW ASSEMBLY _ " " " _/ ll_'_""_r+..._mr+....+. LO_/E R PE pjSC O p_
HOUSlNG ASSEMBLY
FIGURE 12-7 INTERNAL PERISCOPE CAMERA (CAPSULE 9 ONLY)
CC.;-;;_..JT:AL
PM45-272
13-1
SECTION XII
INSTR UMENTATION SYSTEM
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TITLE
TABLE OF CONTENTS
PAGE
SYSTEM DESCRIPTION ...........................
SYSTEM OPERATION ....... •......................
SYSTEM MONITORING ...........................
SYSTEM INSTRUMENTATION CONTROL ...
INSTRUMENTATION RECORDING ............
13-4
13-4
13-4
13-24
13-25
SYSTEM UNITS ........................................ 13-29
TEST CONFIGURATION ......................... 13-55
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DATE , ,, , I NOVEMBER 1961
REVISED
.REVISED
_'11 .....
Mc'DONNELL_L_ __ST.LOUIS, MISSOURI
PAGE ,,
REPORT
MODEL.
,13-3
SEDR 104
133
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--,_-_,,;.;, ;T-,_-;4T ;A.L
PAGE, 1B-_
REPORT _EDR lOl_
MODEL. 133
ST. LOUIS, MISSOURI
DATE i November 1961
REVISED
REVISED
XIII. INSTRUMENTATION SYSTEM
13-1. SYSTEM DESCRIPTION
The instrumentation system consists of the major components as shown on
Figure 13-1. These components coupled with various transducers and other pickup
devices provide a means of monitoring the physical condition and reactions of
the Astronaut as well as capsule conditions and systems performance. The data
so obtained is coded and applied to the telemetry transmitters and radiated to
ground stations for immediate analysis and evaluation; it is also recorded on a
tape recorder in the capsule for subsequent study and interpretation. Cameras
are installed and so positioned to record the Astronaut's facial expressions
and to provide a chronological record of the instrument readings on the main
instrument panel. Provisions are also provided for automatic programmed control
over some components not intended for continuous operation.
13-2. SYSTEM OPERATION
The instrumentation system is completely automatic in operation from the
time power is applied to the capsule until lO minutes after landing impact,
however, certain components may be controlled or interrogated during flight by
either the Astronaut or ground command. The instrumentation system is divided
into three groups, namely_ monitoring, control and recording. These three groups
are treated individually in paragraphs 13-3, 13-59 and 13-63.
13-3. MONITORING
Instrumentation monitoring consists of sampling values of pressures, tempera-
tures, conditions and operations of various units and functions throughout the
capsulej these samples are converted into signals composed of voltages propor-
tional to the temperature, pressure and conditions being measured. The proper-
MAC 231C (Eev 14 Oc +. GG)
DAT£ I November i_I ___- _'_' IB-5
R_VlSED ..... s'r. Lou,s s, MJssoum ms,owr SEDR 104
RWmEO. ..................... Hoom. 133
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tlonal voltages are calibrated within common maximum and minimum ranges to pro-
vide zero to full scale readings. These signals are then channeled into either
or both of the two commutators (electronic switches) designated High Frequency
and Low Frequency, which are located in Instrumentation Package "A". The
majority of the signals are fed to both "HF" and "LF" Commutators for redundancy.
Each commutator continuously samples its input Channels, combining the signal
voltage pulses into a pulse train for each commutator. The pulse train from the
HF Com_Atator and the LF Commutator is applied to separate but identical 10.5 Kc
voltage controlled subcarrier oscillators, where the changing voltage of the
pulse train vary the frequency of the oscillators. The output of this and other
voltage controlled oscillators driven by aero-medical information are mixed and
used to modulate the telemetry tram_mltters. The HF Commutator signals are
telemetered through the high frequency transmitter while the LF Commutator
signals use the low frequency set. Both the HF and LF Commutator signals are
also converted to pulse duration signals and recorded on a tape recorder in the
capsule.
Figure 13-2 is a block diagram showing the parameters of the instrumentation
that is monitored as well as the Commutator Point assignments with a brief
explanation of each parameter given in paragraphs 13-_ through 13-58.
1S-4. Capsule Electrical Power
Capsule electrical power system instrumentation consists of monitoring the
circuit illustrated on Figure 1S-2.
13-5. 400 cps Monitor
ASCS and fan bus ll5 volt a-c is applied thru two 115/6.3 volt transformers
in Instrumentation Package "A" and '_". The secondary outputs are rectified,
filtered and attenuated prior to being applied to the commutators as a zero to
three volt d-c signal. A three volt signal (full scale) represents 120 volts
PAGE, 13,_6
REPORT SEDR 104
MODEL 133
Mc'DONNELj__ __ST.LOUIS, MISSOURI
v v i ! n 8 IltW mines| 1 • | _J_lk Ikm
DATE
REVISED
REVISED
1 NOVEMBER 1961
I hiIVELOC,T'_s_._o__i1_H
l;'r=C-TI _'l E_,_.N D I i_'r'T _.. t-a U O:r 0 _.
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T _J:_, N% K^. IT'T E 0
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LANDIN_
FIGURE. 15- Z. bI_NI_OI_Ik_GIlN,ST_UMEt/TATION BLOCK DI/_GP, AM-(C,_PSULE5 ie_£)
• a,lJml..n _l • I1"_= 2
J
MAC 231C (Roy 14 Oct. 55)
ovemberl l Mc,DONNELL 1 -7
for each bus.
13-6. D-C Current
D-c current amplitude is sensed by the shunt for the instrument panel
ammeter. This shunt is in the negative lead of all capsule batteries and senses
total battery current. The voltage across the shunt is 50 millivolts when 50
amperes are flowing. This voltage is applied to two d-c amplifiers in package
"C" which amplifies it to a zero to three volt level. A three volt level (full
scale) represents 50 amperes battery current. The output of one amplifier is
applied to the HF Commutator, and the other amplifier output is applied to the
LF Comnutator.
13-7. D-C Voltage
The 24 volt d-c monitor circuit is made up of a voltage divider network
in package "A". Voltage from the main pre-lmpact 24 volt d-c bus is applied to
this divider. A three volt signal (full scale) represents 30 volts bus voltage.
The 18 volt d-c isolated bus and the 18 volt d-c standby busses are similarly
monitored through voltage divider networks located in the '_" package.
13-8. Standby Inverter
The standby inverter on signal is obtained through normally open contacts
of the standby inverter relay. This relay energizes when either of the main
inverters fails. With the relay energized, 2_ volts d-c is applied to an
attenuator in the power and control relay box. Attenuator output (2.2 volts
d-c, nominal) is applied to the commutators.
13-9. Instrumentation Power Supplies
Instrumentation power supplies instrumentation consists of the monitor
circuits for the 3 volt d-c reference, zero reference, and 7 V 400 cps power
supplies. (See Figure 13-2. )
YVlll |ia/l_mll • illl_q
PAGE,,, 17-8 MCIDONNE_ __ DATE 1 November-1961
REPORT SEDR 10_ .......... ST:'_OUIS-, _MtSSQURL. REVISED
13-10. Three Volt Reference
The 3 volt d-¢ reference power supp17 furnishes excitation for all potenti-
ometer type instrumentation pick-ups. The power suppl7 is located in package "A"
It is a zener diode regulated supply fr_n the 24 volt bus. The output from the
power supply is applied directly to the cumautators and serves as a reference
full scale si_l..
13-11. Zero Volt Reference
The zero reference si6nal is slsnal ground and is also the return for the 3
volts d-c reference power supply output. This signal is also applied to the
c_utat ors.
13-12. Seven Volt 400 cps
The 7 volt 400 cps Power supply fur_Lshes excitation for the input bridge
circuits utilized with the resistance element amplifiers. Power supply output
is rectified, filtered and attenuated to a zero to three volt level. This zero
to three volt signal is applied to the cazmutators. A three volt si_l (full
scale) represents a 7 volt output level. The power supply is a transistorized
Power inverter which operates on 2_ volts d-c to provide the 7 volt 400 eps out-
put. It is located in package "A".
13-13. Calibration On
Calibration ON instrmnentation consists of a circuit which monitors presence
of the full scale and zero scale calibration cunm_d s_ls. 1_ds_ is present
when the CALIBRATION switch in the telemetry trailer is placed to "FULL SCALE" or
"ZERO" position. When the f_l scale calibrate cc_nand is present, 24 volts d-c
is applied to an attenuator in Package "C". The output of the attenuator (2.8
volts d-e, naeLinal) is applied to the ca_utators. When the zero scale calibrate
command is present, 24 volts d-c is applied to a different input point on the
same attenuator network. The output of the attenuator (1.5 volts d-c, nce_mal)
f_ lr't _T lr_ lr "ri T_ ",_',"-" - •
MAC 231C (l_v 14 Oct. §S)
1Novem r Mc,DONNEg 13-9REVISED ST. LOUIS 3, MISSOURI REpoRT, SF_R 10)-t -
RWISE= -_- MO.m. 133
L !
is applied to the commutators.. Thus, an upper scale signal indicates presence
of the full scale calibrate command and a half-scale signal indicates presence
of the zero scale calibrate command. These command signals energize relays
which apply calibrate signals to numerous other instrumentation channels. The
'_" and "Z" calibrate function may be initiated from a ground station through
the command receivers while in orbit.
13-i_. Static Pressure
Static pressure instrumentation consists of a potentiometer type transducer
which is operated by static pressure. The potentiometer is excited with 3 volts
d-c from the instrumentation power supply and wiper voltage output is inversely
proportional to static pressure. A three volt signal (fUll scale) is representa-
tive of 0 psia.
13-15. Environmental Control System
Environmental control system instrumentation consists of circuitry which
monitors primary and secondary oxygen supply pressures, suit inlet air pressure
and temperature, cabin pressure and temperature and coolant quantity.
13-16. Suit 02 Partial Pressure
02 partial pressure is sensed by a transducer in the Astronaut's suit# the
signals are amplified and transmitted to a pressure gage on the instrument panel.
The signals are also applied to the "HF" and "LF" commutators in Instrument
Package "A" where they are converted to PAM and PDM signals. The PAM signals
are telemetered to ground and the PDH signals are recorded on the capsule tape
recorder.
13-17. Oxygen Supplies
Primary and secondary oxygen supply pressures are sensed by pressure
actuated dual potentiometers in the environmental area. One potentiometer
operates a panel indicator while the other wiper picks off a value for instru-
..... •_-L,-P! A !
PAGE .... 13-10
REPORT _nR lo_
MODEL. l_
Mc,DONNE DATE 1 November l_l
REVISEDST. LOUIS, MISSOURI
--- C___. ___T__T_____.T___T_ REVISED ,
mentatlon. Wiper voltage output is linearly proRortional to pressure. Excitation
is applied from the 3 volt d-c instrumentation power supply. The zero to 3.00
volt wiper output represents a pressure range of zero to 7,500 psi with 0 to 100%
meter indications. Outputs from the primary and secondary oxygen supply pressure
transducers are applied to both commutators.
13-18. Suit Inlet Temperature
Suit inlet air temperature is sensed by two resistance element transducers
in the suit inlet air line. Transducer resistance varies proportionally with
temperature. Each transducer is part of a bridge input circuit to an amplifier
in package "A". The zero to three volt (full scale) output from the amplifier
is representative of a temperature range of 35 ° to lOO°F. One amplifier output
is applied to the HF Commutator; the other amplifier output is applied to the LF
Commut at or.
13-19. Suit Inlet Pressure
Suit inlet air pressure instrumentation consists of a potentiometer type
transducer which is pressure actuated. The potentiometer is excited with 3
volts d-c from the instrumentation power supply and wiper voltage output varies
linearly with pressure. The zero to three volt (full scale) output represents
a pressure range of zero to 15 psia and is applied to both commutators.
13-20. Cabin Pressure
Cabin pressure instrumentation consists of a potentiometer type pressure
transducer installed in package "C". The potentiometer is excited with 3 volts
d-c from the instrumentation power supply and wiper voltage output is linearly
proportional to cabin pressure. The zero to three volt (full scale) output
from the wiper represents a pressure range of zero to 15 psia.
13-21. Cabin Air Temperature
Cabin air temperature is sensed by two platinum resistance wire transducers
i •
MAC 231C (Eev 14 Oct 55)
l ovember Mc,DONNELL l -llREVISED ST. LOUIS 3, MISSOURI REPORT _EDR lO_
REVISED _____ Hoo" ,133
mounted in package "A". Transducer resistance varies proportionally with
temperature. Each of the transducers is part of a bridge input circuit to an
amplifier in package "A". The zero to three volt (full scale) output from the
amplifier is representative of a temperature range of zero to 200°F. Amplifier
outputs are applied to the HF and LF Commutators.
13-22. Coolant Quantity
Coolant quantity is measured by sensing the pressure of the oxygen bottle
used to force water fr_n the coolant tank. This bottle supplies five hundred
pounds pressure. As coolant quantity decreases, the confined volume of the oxygez
increases with a resulting decrease in oxygen pressure. A pressure potentiometer
excited by three volts d-c from the instrumentation power supply monitors oxygen
pressure. Wi_r output is applied to the HF and LF Commutators and through an
attenuator in package "A" to the coolant quantity indicator. Zero to three
volt (full scale) covers a range of zero to 100% coolant quantity. Oxygen
pressure at 100% coolant quantity is 480 psi, oxygen pressure at 0% coolant
quantity is 230 psi.
13-23. Reaction Control STstem
Reaction control system instrumentation consists of monitors for automatic
and manual reaction control supply pressure and Astronaut hand control position.
13-2_. Horizon Scanner
Horizon scanner instrumentation monitors for the pitch and roll horizon
scanner outputs and ignore signals for each of these outputs.
The horizon scanner system utilized two identical infra-red scanning units
to provide pitch and roll reference signals. The Horizon Scanners are on con-
tinuously, from launch until re-entry at which time the Scanners are de-actlvated
by the O.05g relay but during the orbital phase the reference signals are applied
to the ASCS attitude gyros only upon command from the programmer. (Refer to
"--'_"_i A !Ill !- ! !_1 ! i ii ii
PAGE 13-12
REPORT SEDR 104
MODEL. 133
l%4c,DONNE REVISED
ST. LOUIS, MISSOURI•-I"L _,T_ T*_ • REVISED
Table 13-i. ) The signals that are applied to the gyros are monitored by instru-
mentation. The pitch and roll signals range between + lO volts d-c. These
signals are applied to a biased attenuator card to provide a zero to 2.68 volt
output which is coupled to separate channels of the HF Commutator. The signals
represent an output range of + 35°.
Occasionally a scanner sweeps across the sun. Since the scanners are infra-
red devices, sweeping of the sun introduces error voltage. To prevent utiliza-
tion of this voltage, the scanner supplies an "ignore" signal to the ASCS. This
"ignore" signal is monitored as an on-off type of signal by instrumentation.
Pitch ignore and roll ignore signals are applied to the HF Commutator only. A
half scale signal represents presence of the pitch ignore signal and full scale
level indicates presence of the roll ignore signal.
13-25. Attitude
Attitude instrumentation consists of telemetry channels which monitor capsule
pitch, roll and yaw attitudes. Each attitude is read out of a syachro actuated
potentlometer. The synchros are driven by the automatic stabilization control
system. Excitation for the potentlometers is furnished by the three volt d-c
instrumentation power supply. Signal voltage varies along a multiple slope
function with capsule attitude. Pitch and roll signals cover a range of +130 °
to -190 °. Yaw signals cover a range of +70 to -250 ° . Each of the attitude
signals is applied to separate channels of HF and LF Commutators. After retro-
grade assembly Jettison and energizing of a O.05g relay, the potentiometer-
positioning synchros become inoperative. At this time, attitude signals are
removed from the commutators and attitude rate signals are applied to the re-
linquished commutator channels.
13-26. Attitude Rate
Attitude rate instrumentation utilized signals from rate gyros. The gyros
•LAG 231C (Eev 14 Oct. 55)
oA_ i November 1961
REVISED
REVISED
ST. LOUIS 3, MISSOURI
paG= 13-13
R=Pom'. SEDR 10_
are part of the attitude rate indicating system. A zero to three volt signal
level represents a rate level of decreasing 40 ° per second to increasing 40 °
per second. Attitude rate signals are applied to the LF Commutator. Roll,
pitch and yaw rates are assigned to separate channels. In addition, attitude
rates are applied to the channels normally occupied by attitude data when
attitude data is no longer generated. (Refer to Paragraph 13-25. )
13-27. Reaction Control System Solenoids
The reaction control system solenoids control the thrust Jets used for
capsule stabilization in flight. These solenoids can be energized manually or
automatically. When a solenoid is energized, 24 volts d-c is applied through
an attenuator in package "C" to the HI_ and LF Commutators. Each of the twelve
solenoids is represented by a separate commutator channel. This on-off signal
is presented to instrumentation circuitry from the amplifier-calibrator in the
ASCS system.
13-28. Supply Pressures - H202
The monitor circuits for reaction control supply pressures are identical
in operation. A helium source of 2050 psi is utilized to expel hydrogen per-
oxide frum a bladder. As hydrogen peroxide is expelled, the confined volume
of the helium increases and helium pressure decreases. A pressure po_entiometer
senses this change in pressure. The potentiometer is excited with three volts
from the instrmaentation power supply. Wiper output voltage is applied to both
commutators and through an attenuator to an indicator. Transducer range is
600 to 2200 psi. A pressure of 2050 psi provides a reading of 100% on the
indicator. Hydrogen peroxide is exhausted at approximately 600 psi. helium
pressure. Indicator reading at this pressure is approximately 0%.
13-29. Hand Control
Astronaut hand control position is monitored by three potenticmeters. The
C:::7:::;.T:--L
PAGE ' 13-1_
REPORT ,q_.T_ 1 Oh
MODEL. l_
,MCDONNE ST. LOUIS, MISSOURI
DATE I November i_61
REVISED
REVISED
wipers of these potentiometers are driven by linkage to the hand control. Three
volts from the instrumentation power supply is utilized to excite the potenti-
ometers. Zero to three volt signal level represents + 13° hand control movement
in the roll and pitch planes and + lO ° movement in the yaw plane. Wiper outputs
are applied to both commutators.
13-30. Capsule Acceleration
Capsule acceleration instrumentation consists of circuitry which monitors
acceleration along the mutually perpendicular axes. Three accelerometers in-
stalled in package "D" provide zero to three volt d-c outputs proportional to
acceleration along the longitudinal (Nz) lateral (Nx) and normal (Ny) axes of
the capsule. The accelerometer outputs are linear with a zero acceleration
providing a 1.5 volt d-c signal. The longitudinal axis accelerometer covers
a range of + 30g to provide zero to three volt output signals. The normal and
lateral axes accelerometers operate in two ranges. During launch and re-entry
a zero to three volt signal represents accelerations between -+ _g. During orbit,
a zero to three volt signal represents accelerations between + 0.Sg. These zero
to three volt signals are applied to the commutators.
13-31. Velocity Sensor
The velocity sensor is designed to measure separation velecity increase
during posigrade rocket firing and velocity decrease during retrograde rocket
firing. During posigrade firing the integrating accelerometer will, with appli-
cation of 24 volt d-c signal, integrate for 1.5 + _ seconds after a velocity of
lO ft/sec. + 1.5 ft/sec, has been reached. At this time the msximum velocity
signal will be held for approximately six (6) minutes. This signal is applied
to the HF and LF Commutators as a zero to three volt value and is used to
determine the separation velocity of the capsule from the preceding stage. Dur-
Lug retrograde firing, integration will be started by application of a 24 volt d-c
MAC 231C (Rev 14 Oct. _5)
,,o,REVlSEO. ST._" --_'--0"_7"'_'°aOUmS, MnSSOUm _ORT $_DR 104
signal and will integrate maximum velocity for 22 + 2 seconds after a velocity
of 2_0 + 40 ft/sec, has been reached. The signal will be held for 30 minutes
or until impact. The integrating accelerometer will also actuate the 240 ft/seco
relay which wlll direct 24 volts d-c power to the Telemetry velocity sensor relay
through a four and eight minute timer; the timer in turn directs the 24 volt
power through various slze resistors to produce a signal composed of three
different values Of four minutes duration each. These timed signals together
with the velocity signals are applied to the HF and LF Commutators to be used to
determine the approximate impact area of the capsule.
13-32. Structural Temperatures
Structural temperature instrumentation consists of monitor circuits for
ablation shield, outer skin and inner skin temperatures as well as inverters,
telemetry transmitters and retro-rockets.
13-33. Ablation Shield Temperatures
The ablation shield temperatures are monltored through the telemetry system.
Four transducers are embedded in the inner face of the shield. The transducers
have a temperature range from zero to 2000 degrees F. wlth a nominal resistance
of lO0 ohms at 70°F. Input power of 7 V d-c 400 cps applied to the transducers
Is attenuated to a value dependent upon the transducer resistance. The output
from the transducers is directed Into two identical smpllflers (two transducers
to each emplifler) in package "A" where the voltage Is converted and amplified
to a zero to 3 V d-c signal and applied to the telemetry commentators. The
amplifiers are "R" and "Z" calibrated at intervals from ground command.
13-34. Outer Skin Temperature
Outer skin temperature is sensed by two resistance element transducers
welded to the inside surface of the outer skln shingles In forward and aft loca-
tions. Transducer resistance varies proportionally wlth temperature. Each of
PAGE i_-16
REPORT n1_nn io_
MODEL_ i_
REVISEDST. LOUIS, MISSOURI
these transducers is part of a bridge input circuit to an amplifier in package
"C". The zero to three volt (full scale) outpu_ from the amplifier is representa-
tive of a temperature range of -65 to 2200°F. The out1_at from the amplifier
associated with the aft transducer is applied to the HF Commutator. The forward
transducer amplifier output is applied to the LF Commutator.
13-35. Inner Skin Temperature
Inner skin temperature is sensed by two resistance element transducers
adhering to the inner skin in forward end aft positions. Transducer resistance
varies proportionally with temperature. Each of these transducers is pert of
a bridge input circuit to an amplifier in package "A". The zero to three volt
(full scale) output from the amplifier is representative of a temperature range/
of 0 to 300°F. The output of the amplifier associated with the aft transducer
/
is applied to the LF Commutator. The forward transducer amplifier output is
applied to the HF Commutator.
13-36. Inverter Temperatures
The 150 VA and the main 250 VA inverter temperatures are monitored through
the telemetry system. A zero to 300 degrees F. transducer is attached externally
on each inverter. The transducer resistance change of 133 ohms to 2_0 ohms is
representative of a temperature change from zero to 300 degrees. Input power
of 7 volts 400 cycles applied to the transducers is attenuated to a value de-
pendent upon the transducer resistance, which in turn is controlled by the
temperature,of the inverter. The output from the transducer is directed into
an amplifier in package "A", where the voltage is converted and amplified to a
zero to 3 volt d-c signal and applied to the telemetry commutators. The ampllflez
is "R" and "Z" calibrated at intervals from ground command.
13-37. Telemetry Tran_nltter Temperatures
Both HF and LF telemetry transmitter temperatures are monitored through the
.
MAC 331C (P,_v 14 0_... 66)
DATE 1 NOVember 1_1 ____i_j_ _p,_4_ PAGE I_'17
R_|SE:D
telemetry system. The transducers are the some as used on the inverters (refer
to paragraph 13-36) and ere attached externally on each telemetry transmitter.
The processing of the temperature signal is the same as for the Inverters except
that the amplifier is located in "E" package.
13-38. Retro-Rocket Temperature
The retro-rocket temperature is monitored through the telemetry system.
The transducer is the same as used on the inverters (refer to paragraph 13-36)
and is cemented to the surface of the lower retro-rocket case. The processing
of the temperature signal is the some as for the invaders except that the
amplifier is located in "E" package.
13-39. Aercmedlcal
Aercmedlcal instrumentation consists of monitor circuits for electro-
cardiograph, respiration signals_ and body temperature.
13-_O. Astronaut Blood Pressure
The blood pressure system consists of (1) an occludin 6 cuff, (2) a pulse
sensor_ (3) differential transducer, (_) pressure source_ and (5) a control
system. The occluding cuff is attached externally to the Astronaut's pressure
suit and contains a transducer which measures the differential pressure between
the cuff and the capsule cabin pressure. The pulse sensor is a small transducer
attached to the Astronaut's skin. The pressure source is a separate oxy6en
bottle containing sufficient oxygen to provide 150 cycles of operation during
the mission. The system measures the Astronaut's blood pressure_ converts the
pressure to a corresponding electrical signal which is then applied to the 2.3
Kc voltage controlled oscillators and transmitted by the Hi and Lo telemetry
transmitters.
The blood pressure system may be put into operation either by an "R"
calibrate signal initiated from ground c_maand or by the Astronaut actuating a
REPORT _xEDR 104 REVISEDST. LOUIS, MISSOURI
MODEL_ 133 ........... REVISED.
"start" switch on the instrument panel. In either case, a 24 vdc pulse of five
seconds duration, causes the system to pressurize to 4.4 psi differential pressure
from the pressure source. After pressurizing the system bleeds off at a linear
rate to 0.75 psi in approximately 22 seconds. The output signal from the pulse
sensor is routed through the pressure suit disconnect and mixed with the differ-
ential pressure signal in a superimposing manner. This combined signal is routed
through a relay and relay contacts to the two 2.3 Kc VCO's located in "D" package,
and then to the Hi and Lo telemetry transmitters. Directing the signal through
the relay is necessary in order to share the 2.3 Kc VCO's with the EE_ signals.
The first appearance of the signal indicates systolic pressure with a
minimum peak amplitude of 150 mv while the last occurrence of the pulse signal
indicates diastolic pressure with a minimum peak amplitude of 150 my. The maximum
pulse pressure is 1 volt peak.
Upon completion of the cycle the system will remain at rest (below 3/4 psi
pressure) for lO seconds, after which time the cycle will repeat. If the system
is manually (Astronaut) initiated, operation will continue for 6+1 minutes or
approximately lO cycles unless manually interrupted by the "stop" switch on the
instrument panel. A light on the panel indicates when the system is operating.
13-_l. Astronaut EEG
Electrocardiograph signals are obtained _from four transducers attached to
the Astronaut's right and left side, and on the upper and lower chest. The out-
puts from the transducers are applied to two amplifiers in "D" package (left and
right side paired to one amplifier and upper and lower chest paired to the other).
Signals from the amplifiers are then directed to the 2.3 Kc and 1.7 Kc voltage
controlled oscilators, which in turn apply their outputs to the Hi and Lo
telemetry transmitters. The 2.3 Kc VCO's input signals are divided between the
Astronaut's EI_S and blood pressure outputs.
CCrJY.'DX;_Ti_,.
_c zslc (_v 14o__ ss)
DATE 1 November 1961
REVISED
REVISlED_
srr.-oum s. Mmsoum ._om" __.._R ;L04
.,......................... .oo ,
13-42. Respiration Rate
The Astronaut's breathing rate and depth is monitored through the telemetry
system. A thermistor is mounted inside the Astronaut's helmet, adjacent to the
microphone. Input power of 3 vdc applied to the thermistor is attenuated pro-
portionally to the changing resistance of the thermistor, which in turn is
proportional to the magnitude of the Astronaut's respiration. The respiration
rate is determined by the frequency of the Astronaut's breathing. The changing
output voltage of the thez_istor varies the bias of a transistor; the transistor
output signal is then applied through a calibrate card in "E" package, to two
1.3 Kc voltage controlled oscillators in "D" package. The output of the VCO's
is then applied to the Hi and Lo frequency _mlemetry transmitters. The calibrate
card provides a means of interrupting the signal for "R" and "Z" calibrate. A
potentiemeter in the electronic assembly provides an adjustment for sensitivity.
13-43. Astronaut Body Temperature
Body temperature is sensed by a rectal temperature probe. The probe is
a 1490 + _ ohm thermistor element which is utilized as one leg of a bridge
circuit, which forms the inputs to two d-c amplifiers. The output of each
amplifier is applied to the telemetry commutators. The zero to 3 vdc output
represents a temperature range of 95° to 108 ° C. Both amplifiers are "R" and
"Z" calibrated by ground command.
13-44. Sequence System Normal Launch
Normal launch sequence instrumentation consists of monitor circuits for
tower release, capsule separation, retrograde attitude command, and retrograde
rocket assembly Jettison. These signals are -all on-off type functions and each
is applied to the HF and LF Commutators.
13-45. Satellite Clock
The satellite clock utilizes potentio-_ters to provide electrical signals
representative of elapsed time from launch and retrograde time. These potenti-
REPORT _DR 104 ST. LOUIS, MISSOURI REVISED
MODEL. 933 ..... __ " REVISED
ometers are excited with three volts from the instrumentation power supply. The
outputs for each type of time are divided in signals representative of O to 10
seconds, 0 to 1 minute, O to lO minutes, 0 to 1 hour, 0 to lO hours and 0 to 30
hours. Wiper output is linearly proportional from zero to three volts for each
time span. Wiper outputs are applied to the HF and LF Commutators. Instrumenta-
tion monitors ELAPSED TIME from LAUNCH and also EVENT TIME of retrograde. Elapsed
time from launch is the length of time capsule has been in motion. Prior to lift-
off, elapsed time will be zero. Instrumentation recording devices also will
indicate zero time. Output signals for elapsed time therefore are directly pro-
portional to time. As time increases so will output voltage, for example,
elapsed time recorded by clock is lO hours, 5 minutes and 10 seconds. Output
signals will then be as shown below:
SATELLITE CLOCK OUITUTS FOR i0 HOURS,
5 MINUTES, iO SECONDS, AFTER LAUNCH
TIME POTERTIOMETERS SIGNAL
POTENTIOMETERS WIPER TRAVEL IN _ VOLTAGE
0 - i0 Hours 100% 3 Volts
0 - 1 Hour 100% 3 Volts
O - i0 Minutes 50% 1.5 Volts
0 - 1 Minute i00_ 3 Volts
0 - i0 Seconds 100% 3 Volts
Event time of retrograde is pre-set prior to lift-off. After retrograde
time has been set, instrumentation will receive this time signal continuously
throughout the mission. Event time of retrograde can be changed at any time
during the mission by either the Astronaut or by ground command. When retro-
grade time is changed during the mission, Instrtmentation will receive this
change also. Signal output voltage is proportional to retrograde time. For
example, if retrograde is set to commence at 20 hours, 5 minutes and 5 seconds,
MAO _31C (l_v 14 Oct.. 5_)
DA'r_
remora" s_R lO4
MOO,-- 133
ST'. LOUIS 3. MlSSOUR|
instrumentation will be receiving the signal voltage outputs as shown below:
SATELLITE CLOCK OUTRYI_ FOR RETR0-FIRE AT 20 HOURS,5 _3V_ES A_D 5 SECONDS
TI_ POTENTIOMETERS SIGNAL
POTERTIOMETERS WIPER TRAVEL IN _ VOLTJ_E
O - iO Hours 100% 3 Volts
0 - 1 Hour 1OO% 3 Volts
0 - i0 Minutes 50% 1.5 Volts
0 - 1 Minute l_ 3 Volts
O - i0 Seconds 50% 1.5 Volts
13-46. Tower Separation
When the tower separates from the capsule, the No. 3 tower separate sensor
relay de-energlzes and applies three volts d-c to the commutators. This signal
is present for the remainder of the mission.
13-47. Capsule Separation
When the capsule separates from the booster, a limit switch closes and
causes the No. 1 capsule separation sensor relay to energize. While this relay
is energized a three volt d-c signal is applied to both commutators. This relay
remains energized for the remainder of the mission.
13-48. Retrograde Attitude
The retrograde attitude co_nand signal normally occurs when the retrograde
clock runs out. It may also be caused by ground command or by operation of a
bypass switch on the instrument panel. This signal remains present until the
retrograde rocket assembly is Jettisoned (approximately 90 seconds). Signal
level is approximately three volts. Normally open contacts of the retrograde
attitude ccmnnand relay in retrograde relay box No. 2 closes to route the signal
to the HF and LF Commutator.
PAGE 1_-22
REPORT S_DR lO_
MODEL_, 133
REVISEDST. LOUIS, MISSOURI
-- _'_._.T_*L-_ _':"_'_'_,._..._, REVISED, ,
13-49. Retrograde Rocket Assembly Jettison
The retrograde rocket fire occurs at five second intervals. The first fire
occurs thirty seconds after reception of retrograde clock rumout if the retro-
grade interlock is closed in the ASCS.
The retrograde rocket assembly Jettison signal occurs 60 seconds after the
initiation of the retrograde fire signal. The signal is routed through normally
open contacts of the retrograde rocket assembly separation sensor relay in retro-
grade relay box No. 1. This relay is energized by limit switches which close
when the retrograde assembly is blasted away from the capsule. The relay remains
energized until the O.05g relay drops out. (The O.05g relay de-energizes at
lO,O00 feet. ) A d-c signal of approximately 3 volts is applied through normally
open contacts of this relay to the HF and LF Commutators.
13-50. Emergenc Y Escape Sequence
Emergency escape sequence instrumentation consists of Mayday abort and tower
escape rocket fire signal monitors.
13-51. Mayday
The Mayday signal is produced by the Mayday alarm relay. This relay is
energized by any abort signal. With the relay energized, three volts d-c
(nominal) is applied to the KF and LF Commutators. Once initiated, this signal
is present for the remainder of the mission. The Mayday alarm relay is in launch
and orbit relay box No. 4.
13-52. Astronaut 's Abort Switch
Instrumentation is provided to monitor an abort signal originating from the
Astronaut's Abort Handle; this signal is applied to both commutators.
13-53. Tower Escape Rocket
The tower escape rocket signal is obtained from the emergency escape rocket
fire relay in launch sad orbit relay box No. 2. This relay remains energized
_.,_v _,,.m _..a..J4/_4.'_l JI. Jl.,Z'lkJL,I
\\
_XAC 2310 (P_v 14 Oct 55)
DAVE I November 1961
REVISED
REVISED
ST. LOUIS 3, MISSOURI REPORT _R_R 104
,._ Moore. 133 ,,
for less than one second but a capacitor is connected across the input to the
HF and LF Commutators to maintain a signal level of more than O. 3 volt for
approximately 30 seconds.
13-54. Landing System Sequence
Landing system instrumentation consists of monitor circuits for chute
deploy and Jettison and release of the antenna fairing. These signals are
approximately three volts and are applied to the HF and LF Commutators. Main
and reserve chute deploy signals are obtained from toggle switches in the chute
compartment. Lanyards from the chutes operate these switches when the chutes
deploy. The main chute Jettison signal is obtained through a limit switch in
the chute compartment. The antenna fairing release signal comes from the
antenna fairing separation relay in the communications relay box. This relay
is energized through a limit switch. All landing system signals remain on until
impact.
13-55. O.05G Relay
Instrumentation of 0.056 relay operation consists of an on-off type signal
which indicates whether the relay is energized or de-energlzed. The relay may
be energized by operation of the 0.05g sensor or by the command receiver.
When the relay is energized a d-c signal is applied to the HF and LF Commutators.
13-56. Drogue Chute
Drogue chute deployment is monitored by a three volt signal controlled by
the drogue chute sensorj through a set of contacts on the antenna separation
relay and is applied to both commutators.
13-57. Landing Bag
The landing bag operation is monitored by a voltage signal applied to the
HF and LF Commutators through two sets of unlock signal limit switches.
13-58. Periscope Retract and Door Closure
The periscope retract si_monltors the voltage applied to the retract
J_,;.;'; D_.;.T:AL
PAGE 13-24
REPORT SEDR 104
MODEL
McIDONNE_ __ DATE 1 November 1961REVISED
ST. LOUIS, MISSOURI
_ REVISED
relay. While the periscope is retracting, 24 volts d-c is applied through an
attenuator in package "C" to both commutators. Input level to the commutators
is approximately three volts.
INSTRUMENTATION CONTROL SYSTEM
13-60. C_ing
The signals applied to the HF and LF Commutators are sampled once every 0.80
seconds. Commutator outputs are square wave pulses with amplitude between -i and
+3 volts. These pulses are applied to voltage controlled oscillators and pulse
duration modulation converters.
a. The frequencies of the voltage controlled oscillators are varied
between 10.5 Kc + 6-3/4_ by the commutated pulse an_litude signals. This fre-
quency band corresponds to IRIG Channel 12. The frequency modulated outputs of
the 10.5 Kc voltage controlled oscillators are applied to two mixers.
b. Commutator pulse amplitude modulation signals are also applied to pulse
duration modulation converters. The converters reshape the pulse amplitude
wave-shapes to obtain pulse duration wave trains. These wave trains are then
applied to the tape recorder.
c. Amplified aeromedlcal signals are coupled to pairs of 1.3 Kc, 1.7 Kc
and 2.3 Kc voltage controlled oscillators. A zero to full scale signal causes a
deviation in center frequency of + 6-3/4%. The frequency bands of the oscillator=
correspond to IRIG Channels 5, 6 and 7- The oscillator outputs are applied to
mixers.
d.
combined.
In the mixers, the commutated outputs and the aercmedlcal signals are
Mixer A also accepts a signal from the compensating oscillator. (This
signal serves as a reference during data evaluation to indicate fluctuations in
tape speed. ) The composite signal from each mixer is applied to the tape
MAC _IC {l_v 14 Oct. _)
DATE:]- November 1961 _ _ PAGer 1_-'_
R_VlSE_ ST.LOUlS S,MISSOUm RIPO_r SEDR 10_
Moo.. i33
k
i
recorder, the ground test umbilical and a telemetry transmitter.
13-61. Transmission
Ground testing and control of the instrumentation system is provided through
the umbilical receptacle. Non-radiating checks can be performed to evaluate
system operation. Radiating checks are performed through the telemetry llnk.
Refer to Section XI for further information regarding telemetry.
13-62. Instrumentation Control
The instrumentation system controls and programs (see Table 13-1) power
to its own and other systems equipment by means of mode relays and programmer.
(See Figure 13-3. )
a. The ASCS horizon scanners are powered at regular intervals during the
mission to provide reference signals for the capsule gyros.
b. The water extractor in the environmental control system is also
programmed at regular intervals during the mission.
c. The pilot and instrument panel observer camera operates at a high and
low rate of speed during various phases of the mission.
d. Calibration voltages, R-calibrate for maximum readings and Z-callbrate
for minimum readings, are supplied periodically to the monitoring instrumentation
circuits. This is done prior to launch and by ground commaud at intervals dur-
ing orbit.
e. The X and Y axes accelerometer ranges are + 4 g from launch until orbit
and from retro command until landing. These ranges are changed to +_0.Sg while
the capsule is in orbit. The Z axis acceler_neter range is + 30g throughout
the mission.
13-63. INS_U_EATION RECORDING (see Figure 13-3).
Recording instrumentation consists of a tape recorder and two cameras.
:= ;:.-: :'__L:TIAL
CAMERA
ASTRONAUT-PRIMATE
9-13-16-18-19
INSTRUMENT PANEL
OBSERVER CAMERA
9, 13, 16,18,19
PERISCOPE
CAMERA 8, 9
EARTH AND SKY
OBSERVER CA_/_A
CAPSULE 8, 9
HORIZON SCANNER
ECS WATER
EXTRACTOR 8,9,13,16,18,19
PLAYBACK TAPE
RECORDER 8-9
INST. PANELOBSERVER CAMERA
TABLE 13-1. INSTRUMENTATION PROGRAMMING
Before Umb Umb.EJ.to Cap. Sep. Tr.to Retro Retro Jett.
Separation Cap Sep + Plus 60 Jettison to .OSG60 Sec. Sec.to Tr..
i frame/ 6 frames/ I frame/12 6 frames i frame
12 Sec. Sec. Sec.,Except per Sec. per 12
6 frames/ i Sec.Sec. every J30 Min.
.........................Slow Fast Slow,Except Fast SlowIF/2 Sec. 6 FPS fast 30 Sec.
every 30 Min
1800 frames*
per hour
i0 frames iI0 frames
per Min. _per Min.
!
operating O_erati_ i8.5 ran/30 Min.
1
On until tape depletion at
35.Sec./180 Sec. with 90 Min. stagger.
6 F?S 6 n_ 18oo _PH 6 FPS 6 FPS
•05G to Ht. Shld.Ht. Shld. onward
Deploy
per Sec. per 12Sec.
I
Mayl)ayonward
O
frame s
per Sec
Fast
i0 frames i0 frames
per Min. per Min.
OFF at
Re-entry
iO frame_ i0 frames
per Min. per Min.
............ f .................
30 Sec/
30 Min.
6 FPS 6FPS
F
m_m
!
g
m m
0 o
>m
t-i
:L" i
MAC 231C (Rev 14 Oct. 55)
November _MC,DONNE ,.. 1 -2"rR_vlsEo s'r.-oms s. MiSSOUm _m-o_-r ._.nR :1o_
._IsED Moo_ 133
13-64. Ta_e Recorder
A low power, lightweight tape recorder provides seven channels for data
recording. Mixer A output, voice _m,_cations, LF Commutator, pulse duration
modulation and HF Commutator pulse duration modulation signals are applied to
channels i, 2, 3, 4, 5 and 7, respectively. The tape recorder operates continu-
ously during ascent and descent, and is on 1 mlmute and off 3 minutes during
orbit. In addition, the recorder is turned on during any voice transmission
from the capsule. Recorder speed is 1-7/8 inches per second.
13-65. Instrumen_ Panel Camera
A 16 millimeter camera is installed in the capsule to observe the main
instrument panel. This camera is mounted to the left of the Astronaut's head
and views the panel from over his shoulder. The camera is supplied 2_ volts
d-c from the capsule power system. Operation of the camera is controlled by
also applying 2_ volt d-c power pulses to the camera clutch. Each pulse trips
the shutter, exposing one frame, and advances the film for the next exposure.
The camera operates at a high speed of 6 frames per second and a low speed of
One frame every two seconds. The Instrument Panel Observer Camera contains
450 feet of 16 millimeter film. Power is applied through an Imstrumentatloa
Mode relay and programmer and internal camera programmers which produce the
necessary power pulses; these are supplied during ascent, descent and orbit at
speeds shown on Table 13-1.
13-66. Astronaut Camera
A 16 millimeter camera mounted behind the lower left corner of the main
instrument panel views the Astronaut. This camera is also supplied 24 volt
d-c power and trigger pulse voltage. The Astronaut camera operates at a high
speed of 6 frames per second and a low speed of 1 frame every 12 seconds. The
camera normally contains 250 feet of 16 millimeter film. Internal clock, time
PAGE 13-28
REPORT SEDR 104
MODEL 133
ST.LOUIS, MISSOURI
_Vliln allm,_l---n I • B_e--lknm_
DATE
REVISED
REVISED
1 NOVEMBER 1961
_f,T_O _F.P. t:3.ELAY h
2P_,E- t_PACT8U3
I_A LANb_kI_ BAG
UNLOCK LthA_TSWITC_I
,L
F'I_URE 15-5 CAMERAS AND TAPE _ECORDE_ CONTROL CIP_UIS_CAP._ULE_:/_O,L_,I_/8,J_)
• +
vvnSi iikOkmlq I I/'_ l-
.,,_VL_ =,,.*..,,m.,*+_aa+mU__
51". LOUIS 3, MISSOURI
correlation is included within the camera. Operation is similar to that of the
instrument panel observer camera. Camera speeds are supplied at varying rates
as shown on Table 13-1. The cameras and tape recorder can be controlled through
the umbilical during 6round checkout. Return circuits from these components to
the umbilical provide indications of component operation.
13-67. SPECIAL INSTRUMENTATION
Not applicable to a manned capsule.
13-68. SYSTEMtmZTS
13-69. TRANSDUCERS
Potenti_neter type transducers are connected across instrumentation 3
Volt d-c power. The wiper is activated by the action to be measured. Wiper
voltage is then proportional to the action.
13-70. Control Stick Motion Potentlometers
Control stick motion is translated into rotary potenticmeter movement. One
potenticmeter is provided for each axis of motion.
13-71. Satellite Clock Potenticmeters
The satellite clock (refer to Section XII) utilized potenticmeters to
indicate elapsed time from launch and time to retrograde outputs for O-10
seconds, O-1 minute, O-lO minutes, 0-I hour, and O-lO hours.
13-72. Manual and Automatic Su_l_ Helium Pressure Potentlometers
Helium supply pressure, 2250 psig maximum, actuates the wiper of each
potentlometer transducer to a resistance position proportional to the pressure.
13-73. Attitude Potentlcmeters
The ASCS calibrator (refer to Section 17) provides synchro actuation of
potentlometers for pitch, roll and yaw. Each wiper output is then proportional
to the capsule attitude for that axis.
.... --e .... --...
PAGE 13-qO
REPORT SRJ)R lO4
MOOa_Z33
MCDONNE REVISED
._Sty, LOUIS, MISSOURI
,_,.,,., .,..,..¢., -',-,I .= *.c_..=._ REVISED
13-74. Main and Reserve Oxygen Pressure Potentiometers
Each oxygen bottle pressure actuates a dual potentiometer transducer. A
low resistance linear element is used to operate a panel indicator while a higher
resistance linear element is used for instrumentation. Wiper voltage outputs
are proportional to applied oxygen pressure.
13-75. Static Pressure; Suit Pressure and Coolant Quantity PressurePotentiometers
Each pressure transducer is used to provide a linear output proportional to
the applied pressure. Pressure ranges are 0-15 psia for static pressure, 0-15
psia for suit pressure and 0-500 psia for coolant quantity pressure.
13-76. Respiration Rate and Depth Thermistor
Pilot breathing actuates a thermistor mounted in the helmet microphone area.
Respiration depth is determined by the extremes to which the thermistor is
actuated while breathing rate determines the frequency of actuation.
13-77. Resistance Element Transducers
Resistance elements are used to measure temperatures. The resistance of
the element varies proportionally to its temperature. Mounting of the element
depends on the application. Stick-on surface temperature elements are small,
lightweight units. Other elements are mounted as an integral part of the capsule
structure. The following llst indicates the purpose and approximate temperature
and resistance ranges for each transducer.
(a) Outer Skin Temperature: -65 ° to 2200°F
(b) Ablation Temperature: -55 ° to lO00°F
(c) Suit Inlet Temperature: 35 ° to lOO°F - 249 to 300 ohms
(d) Cabin Air Temperature: 0° to 200°F - 216 to 316 ohms.
1B-78. Body Temperature Transducer
The body temperature transducer is a rectal temperature pickup which con-
COy F,D -
MAC 231C (B.ev 14 Oct,. 55)
DATE 1 November 1961
REVISED
REVISED.
ST. LOUIS 3, MISSOURI
PAGE 13-31
REPORT SEDR 104
.oow. 133
sists of a thermistor imbedded in sealing compound at the end of a flexible
pigtail.
13-79. D-C Current Shunt
The shunt resistance used for the instrument panel ammeter also supplies
voltage for instrumentation. This shunt is discussed in Section XI of this
manual.
13-80. Electrocardio_raph Pickups
Cardiac activity is sensed by four 3 one half inch square, stainless steel
wire screens. These electrodes are attached with silver metalized adhesive
to the Astronaut's body. Small connecting wires leading to the Astronaut's
suit disconnect, complete the circuit.
13-81. Oxygen Partial Pressure Transducer
The 02 Partial Pressure transducers are used to convert 02 partial pressures
to a signal compatible with high level telemetry. The voltage range of 0-3 V
d-c output is representative of 0-6 psi oxygen partial pressure.
13-82. TAPE RECORDER
A low power, lightweight tape recorder is used in the capsule to make avail-
able 7 channels of recorded data. At the present time channel assignments are
as follows: Channel i, HF VCO mixer output; Channel 3, voice c_ications;
Channel 5, LF Commutator PDM; Channel 7, HF Commutator PDM. Tape speed is
1-7/8 ips, convertible to 15 ips by a modification kit. Tape capacity is 4800
feet of ½ inch wide mylar base tape. The tape transport consists of a capstan
drive 3 supply reel and take-up reel mechanism. A d-c motor is used, through
reduction gearing, for capstan drive. A limit switch is provided to interrupt
recorder power should the tape break. Record amplifiers are incorporated in
the unit for Channels i and 3. Channels 5 and 7 utilize amplifiers incorporated
in the comratators located in instrumentation package "A".
Mc,DONNE REPORT ,-°_DR 104. REVISED
•_, STj LOUIS, MISSOURIMODEL- 133 _,tj,_._.t" xa.,, -',_', • ,.,-s.J_" REVISED
"1 Nnv_mher' ] <:k_]
13-83. INSTRUMENTATION PACKAGE "C"
The "C" package incorporates units of various functions into one compact
panel allowing convenience of mounting and of making electrical connections.
These various sub-units are discussed in the following paragraphs.
13-84. Cabin Pressure Transducer
Cabin pressure actuates the wiper of a lO,O00 ohm potentiometer located in
the "C" package. Three volt's d-c from the "A" package is applied across the
potentiometer. Wiper output voltage is then proportional to the cabin pressure.
13-85. Topic Cards
The instrumentation package utilizes a unique method of construction and
mounting of the transistorized amplifiers, power suppliers, attenuators and
monitors. Each unit consists of the necessary component parts mounted on a
printed circuit, dielectric card with printed connector contacts at the bottom
for plug-in insertion. The card is ghen covered 3 with the exception of base
connector contacts and side mounting edges, with a thin layer of epoxy resin.
This coating is used to provide moisture protection, to insure operation in a
100% oxygen atmosphere and to improve mechanical rigidity of components. These
"Topic Cards" are package mounted in boxes providing side rails, base contact
receptacles and printed circuit inter-connections.
13-86. D-C Current Amplifier Cards
Two amplifiers are used to bring the d-c current shunt voltage up to a
maximum of 3 volts d-c. One amplifier is used for the HF Commutator while the
other is used for the LF Commutator. The amplifiers are transistorized and
card mounted.
13-87. Respiration Rate and De_th Calibration and Attenuator Card
Relays mounted on this card allow "R" and "Z" calibration of the signal
supplied frcm the respiration rate and depth transducer. One relay is supplied
CONFILL':NT2 ..
MAC 2310 (Bey 14 O_ 55)
DAVE i November 1961
REVISED
REVISED
!DONNE , ! -BS ,ST. LOUIS 3, MISSOURI RII[PORT SEDR 104
MOm_ 133
for "R" calibration and one for "Z" calibration. Energizing the desired relay
breams the normal circuit and applies the calibration voltage to the commutators.
Resistors are also mounted on the card for attenuation of 24 volt d-c voltages
to proportloaal voltages compatible with the commutators.
13-88. Voltage Monitor Cards
Resistors, capacitors and circuit isolating crystal diodes are used to
attenuate reaction control solenoid valve energizing voltages, standby battery
voltage and periscope retract voltage prior to application to the cca_u_ators.
Each attenuator circuit output, a maximum of 3 volts d-c, is applied to both
commutators.
13-89. Horizon Scanner Card
The horizon scanner amplifier card provides circuitry for processing the
scanner roll and pitch signals, roll and pitch ignore signals and "Z" calibrate
prior to being applied to HF and LF Commutators. During launch and orbit the
scanners are operating continuously; however, during orbit the signals applied
to the ASCS and the c_mmutators are timed by the programmer located in "C"
package. In addition, the horizon scanner slaving signal is applied to the KF
commutator.
13-9o. PR_
The programmer contains switch contacts which operate control circuits for
specific intervals. The progra, ner is mounted on the forward side of the peri-
scope housing.
13-91. A - Section
The programmer used for orbital missions consists of two sections. When
power is applied to the progremmer, electronic controlled timers continuously
operate the following contacts:
.......... | • i_VI'qlF | JS,,_&&,dlRI It ,il4"_lS,M
PAGE 13-'_4
REPORT SEDR 10_
MODEL 133
Mc'DONNE ST. LOUIS, MISSOURI
IL_v J, 1 I. JLJLP
DATE 1 November 1961
REVISED
REVISED,
CONTACTS CLOSED
WAFER SECTION DURATION
i. Water Extraction 30 Seconds
2. Full Scale Calibrate 5 Seconds
3. Zero Calib. 5 Seconds
RATEu
1 per 30 minutes
See Table 13-2
See Table 13-2
13-92. B - Section
The B Section of the programmer is energized through the command receiver-
decoders and the electronic timers to provide full scale and zero calibrate
signal, as follows :
WAFER SECTION
Full Scale Calibrate
Zero Calibrate
CONTACTS CLOSED
DURATION RATE
3 Seconds On Command
3 Seconds On Command
le
2.
13-93. INSTRUMENTATION PACKAGE "A"
The "A" package also incorporates units of various functions into one panel.
These sub-units are discussed in the following paragraphs.
13-9_. Cabin Air Temperature Transducers
A platinum resistance wire is used to measure cabin air temperature. Tempera-
ture changes frcm 0° to 200°F cause the element to change resistance from 200 to
300 ohms. The resistance element forms a part of an amplifier circuit. Two
transducers are used in conjunction with two amplifiers to supply signal to both
commutators.
13-95. Filament Transformer
A filament transformer is used to step down ll5 volts 400 cps capsule power
to 6.3 volts for use In packages "A" and "C".
/'_ jr% I_T 1_ lr lr-i 117 _v n_Llr.-Jl-,I.d_ ....
MAC 231C (Rev 14 Oct. 55)
DATe i November 1961 PAng I_-_ ,
REVISED. $1".LOUlS 3, .ISSOUR' R[PORT SEDR 10A
R_VlSED - * _OOU. 13_
13-96. Topic Cards
The instrumentation package "A" also utilizes the Topic Card principle for
amplifiers and power supplies.
13-97. Resistance Element Amplifier Cards
The same type amplifier is used for heat shield outer skin, suit inlet air
and cabin air temperature transducer signals. Each amplifier is of dual channel
design in order to accommodate the two transducers used to measure each type of
temperature. Seven volts, 400 cps is supplied from the resistance element power
supply. This voltage is applied across a bridge circuit in each amplifier. The
transducer associated with each bridge circuit causes the voltage in the circuit
to vary proportionally to the transducer temperature. This voltage change
appears across a transformer and is rectified, using crystal diodes, to a
maximum output of 3 volts d-c. The two outputs from each amplifier are supplied
to the commutators. Two relays on each card amplifier allow full scale and zero
calibration of each channel. Calibration potentiometers are also provided for
each channel.
13-98. Resistance Element A-C Power Supply Card
Resistance element amplifier circuits require 7 volts, 400 cps a-c. Capsule
power, 24 volts d-c, is applied to a transistorized power inverter. The In-
verter, using zener dlode-translstor voltage regulation and transistor switching,
supplies a 7 volt a-c output which is monitored by an attenuator, rectifier
circuit. The monitor output, a maximum 3 volts d-c, is applied to the
commutators.
1S-9R. Body Temperature Amplifier Cards
Two transistorized d-c amplifiers are used to increase the output of the
temperature transducer to a maximum 3 volt d-c level prior to application to
the commutators. These amplifiers are of the same type as those used in instru-
REPORT aEDR lOLl• REVISEDST. LOUIS, MISSOURI
MODEL 1_ -- C'_v_?;_==_E?_-'_I_:._i P REVISED,,
mentation package "C" for d-c current amplification.
13-100. Signal Condition and D-C Su_l_ Card
This Topic Card provides four functions in the instrumentation system.
Filament transformer output is applied to a monitor circuit. This circuit
attenuates and rectifies to provide a maximum 3 volt d-c signal indicating trans-
former operation. Capsule power, 24 volts d-c, is applied to an attenuator
circuit which provides a 3 volt d-c output for the monitoring circuit. This
3 volt output Is them applied to the commutators as a monitor of the main 24
volt d-c bus voltage. The signal condition and d-c power supply card also pro-
vides meter attenuator resistors which limit current flow in the panel indicator
circuits. These circuits involve the maln and reserve oxygen supply pressures,
the automatic and manual fuel supply pressures and the coolant quantity circuit.
13-101. HF and LF Cc_mutator-Ke_er-Record Amplifiers
Two units are provided in the "A" package for commutatlng transducer data
and supplying PDM and PAM outputs. The commutator portion of each unit is a
90 x l_, solid state device which samples sequentially, 88 channels of signal
input information. The output produced is a pulse amplitude modulated signal
wave train. Each 0 to 3 volts d-c input to the commutator is sampled l_ times
per second per IRIG standards. The PAM wave train output is fed through a
buffer stage to a PAM/PDM converter. The PDM output is then applied to a record
amplifier which produces a signal capable of directly driving the recorder head
in the capsule tape recorder. The PAM output is fed through a gating circuit
which introduces a master pulse and negative pedestal pulses to operate automatic
decom_utation equipment in the ground station. A power supply is incorporated
in the unit to provide the positive and negative voltages required in the circuits
CC, L :IDZ,. -
MAC 231C (Eev 14 Oc_ 55)
1 ovember Mc,DONNELL 13.37REVISED slr'. LOUlS 3, MISSOUm RE_OR7. SEDR 104
R_,SE_ .... _OOL 133
13-102. INSTRUMENTATION PACKAGE "D"
The primary function of the "D" package is to convert capsule informatlon
to signals capable of modulating the telemetry transmitters. Transducers and
amplifiers are also contained in the package to complete capsule information
circuits.
13-103. Accelerometers
Three accelerc_eters are mounted in the "D" packs_e and used to determine
"the static lon61tudinal, lateral and normal accelerations of the capsule. Each
unit gives a d-c output which is applied to 3 channels of each commutator.
13-104. Electrocardlo_ram Am_llfler Cards
Four amplifiers are used for the E_D transducer inputs. Each amplifier
increases the transducer output to a 3 volt peak to peak signal.
13-105. Oscillators
The "D" package supplies sub-carrler oscillators to allow two channels of
instrumentation data. The A channel is associated with the HF commutator, high
frequency telemetry transmitter and tape recorder. The B channel is associated
with the low frequency telemetry transmitter. The following paragraphs describe
the sub-carrler oscillators.
13-106. Cc_nsatlm_ Oscillator Card
During playback of a tape, the recorded signal from this oscillator is
monitored to detect changes in tape recorder speed. A frequency shift indicates
a change in speed. The oscillator is of Topic Card construction and operates
at 3125 cps. Output level is adjustable.
13-107. Voltage Controlled Oscillator (V. C. O. 1 Cards
Instrumentation data voltages are applied to the sub-carrler oscillators
causing oscillator frequency shift proportional to the input amplitude. The
transistorized oscillator consists of a free running multlvibrator and filter.
The oscillator functions and frequencies are given below:
i H
PAGE 13-38
REPORT _DR 104
MODEL l_
REVISEDST. LOUIS, MISSOURI
........... REVISED
(i) C_mel A
Commutator - 10.5 kc
Right side (+) Ekg V.C.O. 2.3
Left side (Comm) Ekg V.C.O. 2.3 Kc
Lower Chest (+) Ekg. V.C.O. 1.7 Kc
Upper Chest (Co=m) Ekg. V.C.O. 1.7 Kc
Respiration rate and depth - V.C.O. 1.3 Kc
Compensating oscillator - 3.125 Kc
(2) c_el
Commutator - 10.5 Kc
Right Side (+) l_.kg V.C.O. 2.3 Kc
Left Side (Comm) Ekg V.C.O. 2.3 Ec
Lower Chest (+) Ekg V.C.O. 1.7 Kc
Upper Chest (Coma) Ekg V.C.O. 1.7 Kc
Respiration rate and depth - V.C.O. 1.3 Ec
13-108. Mixer Amplifier Card Power Supply
Capsule power, 24 volts d-c, is converted to 6 volt d-c for use by the sub-
carrier oscillators. A mixer circuit combines the sub-carrier oscillator outputs.
Solid state components for these circuits are combined on one Topic Card.
13 -109. CAMERAS
13-110. Instrument Panel Observer Camera
A standard camera is modified by the application of a special drive motor,
self contained slow and fast programmer and a special housing. Slow operating
of the motor shutter mechanism and film transport begins when the capsule camera
and tape record relay is de-energized by the block-house. Full capacity is 450
feet of 16 millimeter film.
• _ I¸
DATE I NOVEMBER 1961
REVISED
•REVISED
Mc'DONNE_ __ST.LOUIS. MISSOURI
:D _.;-, ,'7-I-A-L
PAGE "13-39
REPORT .SEDR 104
MODEL 133
" :;_
<7
_, _Q_:::.::;i_::i::::i_:_TE_/A`RSEM_LY:• _:•:::::::.:_.LEN_cmP_::_i::::M_::_7-iNE_SSEMaLY::::_iii::ii::iiiii::i::i::_:YEU.0W i : i ::i:ii_iiiiiiill_i C0V_.I:
!i ii !!iiiiiiiili!i?i i!i!,iiii!iiiiii!iiii!i i_:!!I ?iil!i i!i
FIGURE13-# EARTH AND:SKY :OBSERVER CAMERA PM45-1g&
._lil_lh. llllllllllklllLllllll | •
VVI_il lll_m,---ll I l,l"_ I,-,-
PAGE 13-140
REPORT SRDR 1o4
MODEL 10;4
__Z DATE 1 November 1961
ST. LOUIS, MISSOURI
Table 13-2.. Instrumentation Commutator Point Assignment
(High Level 0-3 Volt. )
ELECTRICAL POWER SYSTEM
18V Isol Bus
18V Stby Bus
Fans A-C Bus
ASCS A-C Bus
D-C Current
Standby Inverter ON
Standby Batteries ON
ASCS Slovlng
24 Volts D-C
INSTRUMENTATION POWER
SUPPLIES
3V D-C Reference
0%' (Zero) Reference
7V, _0CPS
STATIC __
v_zol (co_ Ao_z)
EBVIRORME_TAL CONTROL
S_t I_%et Pre=sure
l_resm_z_
BY LF
12 12
12 12
9HF LF
_l' 41
42 42
62 62
63 63
83 83
41 41
42 42
62 62
63 63
83 83
1 1
2 2
3 3
22 22
5 2749 71-64 6_
8 8
9 9
1 1
2 2
3 3
22 22
6_ 6_
8 8
l0]IF LF
41 41
42 42
62 62
63 63
63 63
83 83
1 1
2 2
3 3
22 22
6_ 6_
8 8
9 9 9 9
t
33 33
77 77
42 42
62 62
63 63
63
26 26
1 1
2 2
3 3
22 22
6464
8 8
9 9
16BY LF
41 41
63 63
83 83
1
2
3
1
2
3
9 9
8 8
6_ 6_
22 22
MAC SIC (P_v I_ Oc_ 56)
OATE, ! November 1961
REVISED
REVISED
ST. LOUIS 3, MISSOURI
_IT: D__._ITIAL.,.
._s=, 13-41
._o_ SEDR 104
MOO= 133
Table 13-2. Instrumentation Commutator Point Assignment
(HighLevel 0-3 Volt.) (Cont'd.)
8HF LF
ENVIRONMEA_AL CONTROL
SYSTEM (Cont 'd.)
Cabin Temp. i0 I0
Suit Inlet Air Temp.
Reserve 02 SupplyPressure
Cabin Pressure 82 82
Coolant Quantity 84 84
sult 02 PartlalP_sm=_ 6 6
RERCTICm COWI_OL SYSTEM
Reaction Cont.
Pre,msre (Auto.) 39 39
Reac_i_ Cont. Supply
Control P_sltl_s
Pitch
Taw
X Axis
X Azle
is _s43 13
73 73
14 4_7_
9HF LF
i0 i0
11 11
12 12
82 82
84 84
6 6
39 39
4o
13 _s43 13
73 73
7_
i0EF LF
i0 i0
11 ii
12 12
82. 82
84 84
6 6
39 39
_o
23 23
24 24
25 25
13 43&3 1373 73
10 10
11 11
12 12
82 82
8_
39 39
_o
13 43
_3 13
73 73
16HF LF
i0 i0
11 ii
12 12
82 82
8_ 84
6 6
39 39
_o _o
23 23
7_ 74
15 1545 45
PAGE 1_-42
REPORT SEDR 104
MODEL i_
"" %6_'_ DATE 1n_ D ONNE__._ REVISED
ST. LOUIS, MISSOURI
November 1961
Table 13-2. Instrumentation Commutator Point Assignment
(ugh Level 0-3 Volt.) (Cont'a.)
ACCELERATION
z Axis(cont,_.)
STRUCTURAL _ERATURE
Heat Sink
Ablation Shield Temp.
Outer Skin Temp.
Inner Skin Temp.
25OVA Inverter Temp.
15OVA Inverter Temp.
AEROMEDICAL DATA
EKD
Respiration
Body Temperature
O.OSG RELAY
HORIZON SCANNER
Pitch, Output
Roll Output
Pitch Ignore
8HF LF
19 19
2O 2O
21 21
87 87
86 86
88 88
25 25
9
19 19
2O 2O
21 21
23.
l.T&2.3KcVCO's
1.3 KcVCO's
4 4
87 87
4o
88 88
S5
i0
HF LF
19 19
2O 2o
21 21
@
1.T&2.3 KcVCO's
1.3 KcVCO' s
4 4
87 87
86
88
85
13HF LF
19 19
2o 2O
21 21
Z.T&2.3KcYCO's
1.3 KcVCO's
4 4
87 87
16HF LF
2O 2O
21 21
1,7 &2.3_YCO's
1.3 KcVCO 's
4 4
87 87
88
86
85
MAC 231C (_v 14 0¢f. 5S)
oAr¢ i November i_61
REVISED
RL=V|SED,
ST. LOUIS 3. MI_._OURI
Moore. 133
k
Table 13-2. Instrumentation Commutator Point Assignment
(HighLevel 0-3 Volt. ) (Cont'a.)
HORRIZON SCANNER
(Cont 'd)
Roll Ignore
ATTITUDE
Pitch
Roll
Yaw
ALTI_N/DERATE
Pltch AB
Roll AB
Yaw AB
RCS CONTROL SOLENOIDS
Pitch High Up
Pitch High Down
Pitch Low Up
Pitch Low Down
Roll High CW
Roll High CCW
Roll Low CW
8KF 12
59 59
16 86
17 8888 17
18 18
2.3KcVCO 's
1.7KcVCO's
1.3KcVCO's
65 65
66 66
67 67
68 68
69 69
70 70
76 76
9HF LF
85
16 %16
17 17
18 % 18
S
16 16
S17 17
65 65
66 66
67 67
68 68
69 69
7o 70
76 76
lO
HF LF
85
16 16
17 17
18% 18
S16 85
s
17 86
s18 88
65 65
66 66
67 67
68 68
69 69
7o 70
76 76
13HF LF
16 86
18 18
65 65.
66 66
67 67
68 68
69 69
70 7o
76 76
:3;ir;D:;.7;A;
16
85
S16
16 85
1717 86
S18
18 88
65 65
66 66
67 67
68 68
69 69
7o 7o
76 76
PAGE 13-I_4
REPORT _DR ZO4
MODEL 133
REVISEDST. LOUIS, MISSOURI
Table 13-2. Instrumentation Commutator Point Assignment
(High Level 0-3 Volt.) (Cont'd).
RCS CONTROL SOLENOIDS
(Cont'a.)
Roll Low CCW
Yaw High Left
Yaw High Right
Yaw Low Left
Yaw Low Right
SATELLITE CLOCK _lapsed
Time)
i0 Seconds
i Minute
i0 Minutes
i Hour
I0 Hours
0 - i0 Seconds
SATELLITE CLOCK (Retrograde
Time)
i0 Seconds
1 Minute
i0 Minutes
1 Hour
I0 Hours
0 - i0 Seconds
8HF LF
77 77
78 78
80 80
81 81
26 26
28 28
29 29
30 30
3l 3l
3e 32
33 33
34 34
35 35
36 36
37 37
38 38
77 77
78 78
80 80
81 81
26
28 28
29 29
30 30
3l 3l
32 32
33 33
34 34
35 35
36 36
37 37
38 38
10
HF LF
77 77
78 78
8o 8o
8l 8l
26 26
28 28
29 29
30 30
3l 3l
32 32
33 33
34 34
35 35
36 36
37 37
38
_gY.FIDEr:TIA_ c
?7 77
78 78
80 8o
81 81
16HF LF
77 77
78 78
8o 8o
81 81
26 26
28 28
29 29
30 30
31 31
32 32
33 33
34 34
35 35
36 36
37 37
38 38
MAC '_31C (Rev 14 Oct,, 55)
DATE. i _ovember 1961
REVISKD
REVISED
ST. LOUIS 3, MISSOURi
PAGt i_-_5
RmP,oi_-r. SEDR 104
MO.m. 133
Table 13-2. Instrumentation Commutator Point Assignment
(High Level 0-3 Volt. ) (Cont 'd.)
INTERIM CLOCK
0 - lO Seconds
O - lOO Seconds
0 - lO00 Seconds
0 - i0,0OO Seconds
TR TIMER
NORMAL LAUNCH SEQUENCE
Sync Pulse
Sync Pulse
Tower Release
Capsule Separation
Retro Attitude Cored.
Retro Rkt. Assy.Jett.
Retro Rkt. No. 1 Fire
Retro Rkt.No. 2 Fire
Retro Rkt.No.3 Fire
Ret ro Rk%. As sy.Jett.
EMERGENCY ESCAPE SEQUENCE
Pilot Abort
8HF LF
_6 46
47 47
48 48
53
50 50
52 52
53 53
9HF LF
46 46
47 47
48 48
53 53
5o 5o
51 51
52 52
53 53
lO
HF LF
_,7 47
53 53
53 53
59 59
13HF LF
26 3131 3333 26
28 3232 3434 28
29 3737 3535 29
30 3838 30
36 36
89 89
90 90
46 46
47 47
48 48
53 53
50 50
52 5e
53 53
_9 59
16HF LF
_6 46
47 47
48 48
53 53
59 59
PAGE 13 -46
REPORT SEDR 104
MODEL_ 1]_
Mc'DO1NNELL __ ST. LOUIS, MISSOURI
IL.,_.JJ.'q A AJILJJF,_.I.I .IL JLA'3L'_L4
DATE i November i_61
REVISED
REVISED,
Table 13-2. Instrumentation Commutator Point Assignment
(alghLevel0-3 Volt.) (Cont'd.)
EMERGENCY ESCAPE SEQUENCE
(Cont' d. )
May Day
Tower Escape Rocket
LANDING SYSTEM SEQUENCE
Drogue Chute Deploy.
Antenna Release
Main Chute Deploy
Main Chute Jettison
Reserve Chute Deploy
Landing Bag
PERISCOPE RETRACT
INTEGRATING ACCELEROMETER
;/_ _._re,_ a_SOR
PRINATE REACTION TOTAL
8 9 lO
_F LF EF LF HF LF
60 60
61 61
54 54
55 55
56 56
57 57
58 58
6o 6o
61 61
54 54
55 55
56 56
57 57
58 58
7_ 72
ANIMAL RESPONSE
R. H. Switch
Center Switch
L. H. Switch
3 -_-
6 _--
59 ---
Pos. Identification
COMMAND RECEIVER ALL
CHANNEL SIG.
PARACHUTE COMPT. PRESS
4 m__
--- 25
6o 6o
6]. 6l
54 54
55 55
56 56
57 57
58 58
72 72
52 52
50 50
13LF
60 60
61 61
54 54
55 55
56 56
57 57
58 58
72 72
5 5
16KF LF
60 60
61 61
54 54
55 55
56 56
57 57
58 58
5l 5l
72 72
CgNYIDZ;_TIA,_-
MAC 231C (_v 14 Oct,. 5s)
OATS, I November 1961 ]VIC'DONNEL L-ST. LOUIS 3, MISSOURI
m,A@,r 13-47
m_m1" SEDR IO_
,4oo-', 133
/ •
\
Table 13-2.
ROLL TO YAW COM' D
(m Gain)
ROLL TO YAW COM'D
(Lo Gain)
YAW TORQUER SIG.
HEAT SHIELD CAVITY PRESS
TELEMETRY INTERROGATIONMONITOR
Instrumentation Commutator Point Assignment(High Level 0-3 Volt. ) (Cont 'd.)
8HF LF
85 85
9KF I/
27 27
_9 49
71 71-
25
l0 13 16HY LF KF LF HF LF
_o
S.
Kr.
IF.
Not Ccmmutated
With O.05G Relay De-Energized
With O.05G Relay Energized and Retrograde Assembly Jettisoned
High Frequency T/M Transmitter Cosmratator
Low Frequency T/M Transmitter Commutator
=J. ;.7"; D _.;4T;A L
i 2
YAW CON'TI_. LMo'rioNPOTENTIOMETER
PITCH CONTROL
MOTIONPOTENTIOMETER
o, RESER_VE-ROLL ERROR ¢, PARTIALPRE_.Ua._ CHUTE'_W TORQUE DEPLOY
OEMOOULATOR LIMIT
TR,qNSDUCER
(_ul'r _ILET (.SUIT INLET M[IIN CHUTE
All_ PRESSL)RF.) RIR TF_.,MR) DEPLOYLIMIT 3WlTr.H
;0 "BF! >
m _
g
;0 ;0 0
g g m
°°10 0
L
rf• "L
X,Y,Z ACCELE P_OMET EIK%
INSTRUMENTATIONPACKAOE _B" --
"lAW PJkTE
DETECTOR ,_//
INSTRUMENTATION
"RA
_ RECORDERVOICE AND TELEMETERING
PDM-COIvlIvt A (HI FRE_.)PDM-COMM. B (LO FREQ.)
iI
i // j
I / /' /
I
, ', _ ,
HOLDER
(INSTRUMENTATION)
COOLANT QUANTITY
TR/kNSbI3CER
<iii_ mm 111o 13
o<m
i.
_ >m
m
_°o
PAGE 17-5o
REPORT_ SEDR, 104
MODEL 13_
REVISEDST. LOUIS, MISSOURI
..:_-_l'_ _T_ x w'__. _T_ %_][_ ... REVISED
Table 13-3. Special Instrumentation Parameters Commutator Point Assignments
(Low Level - 3MV to 13MV Ref. )
CapsulePole
Nomenclature
Zero Scale (-3MY)
Full Scale (13MV)
Ref. Junction Temp60°F to lO0°F
Ref. Junction Temp70°F to 90°F
T/M "A" TransmitterZll2
UEF Power Amp.
ZLO6.5
Near Batt. Zl05
"S" Band Beacon
Zl07.5
Envlr. Area Zl17.5
H.F. Voice Trans-
mitter
Cabin Air Temp ZlO1.5
"C" Band Beacon
Cabin Air Tamp Z145.5
Heat Exchanger Outlet
Heat Exchanger Inlet
H_O_ Line, Conicalsec_. (Auto)
H209 Line, Conicalsec_. (man)
Solenoid Valve 24"
ZlTO.5
9
Chart. Chart.
i i
2 2
3 3
5 5
6 6
7 7
8 8
9 9
i0 i0
ll
12
13
#1 #2Chan. Chan.
5 5
6
7
8
9
lO
6
ii ii
12 12
13 13
7
8
9
lO
ll
12
13
4"
MAC $31C (l_v i40c_ _;
DArE 1 November 1961 I_I)0NN__........ _.-_ rAG,, .1.3-51 .n_VlSED ST.UOUm S.M,SSOU_, REPORT SEDR 104
Rsv,sm: _ _ Moo= 133
Table 13-3. Special Instrt..entation Parameters Commutator Point Assignments(Low Level - 3_ to 13MV Ref. ) (Cont 'd.)
Nomenclature
CapsulePole
(Outer
(Outer
9_1 •Chan. Chan.
14
15
16
17
18
19
20
21
22
23
24
25
26
27
28
14
15
16
17
18
19
20
21
22
23
24
25
26
27
28
29
H202 Line to 24"Chamber
_0 2 Tank(Man)
_o 2 T_ (Auto)
Heat Shield (Blkh'dcap)
L/O Auto System RollSolenoid Valve
L/O Auto System Roll
ti202 Line
Recovery Comp.
(_er _in)
Inner Skin Below
Reaction Control Jets
Recovery Comp (Int.Structure )
Recovery Comp (Inner
skin)
Ablation Shield (inbd)
Ablation Shield (inbd)
Parachute Comp. Z179
Parachute Ccmp. Z179
Parachute Camp.
Parachute Comp.
Emerg Egress Hatch
(Outer Edge)
Parachute Comp (outer
29
__ 30 30
Chin. Cm,_.
14 14
15 15
16 16
17 17
18 18
19 19
20 2O
21 21
22 22
23 23
24 24
25 25
26 26
27 27
28 28
29 29
30
::Y:____.':T'f.L
w
PAGE 13-52
REPORT SEDR lO4
MODEL_ 133
Mc'DONNEL LST. LOUIS, MISSOURI
DATE 1 November 1961
REVISED
REVISED
Special Instrumentation Parameters Commutator Point Assignments
(Low Level - 3MV to 13MV Ref. ) (Cont 'd.)
Nomenclature
CapsulePole
Parachute Comp (Outer
Skin) Lx
Emerg Egress Hatch
(Outer Edge)
Conical Section Sides
Outer Skin
z154.5 T7zA54.5z_54.5 LxZ_26.5 TFzi17.0 Ty
Emerg Egress (Per-
cussion Ignitor)
Conical Section Sides
Outer Skin
Z126.5 ByZ126.5 By
Conical Section Sides
Outer Skin
zao7. TyZI07 ByZI07 LxZa07 Lx
Tower to CapsuleRetaining Ring (Top)
Tower to Capsule Retain-
ing (Bottom)
Tower to Capsule Retain-
ing Ring (Bottom)
Window Outside of
Inner Pane
Chan.
69
31
32333435
36
3738
394o4142
43
45
Chan.
69
31
32
333435
36
3738
394o4142
43
45
13
Chart.
31
323334
35
36
3738
39404142
43
44
45
Chan.
31
Ablation Shield
Mid.25 Glass Fiber
Mid.25 Glass Fiber
46
4748
46
4748
46
4748
323334
35
36
3738
39404142
43
46
4748
MAC 231G (Rev 14 Oct. 55)
DAVE I November 1961
REVISED
REV|SED
ST. LOUIS 3, MISSOURi
PA_- 13-53
Rm_o_rr SEDR 104
MODL .....133
ii)
/! ;,
_.[T:."TJ
Table 13-3. Special Instrumentation Parameters Commutator Point Assignments(Low Level - 355/to 13}47 Ref. ) (Coat 'd.)
Nomenclature
Capsule 9 13Pole #1 ! #2 #l __ #_
Chan. Chan. Chem. Chan.
Inbd Surface of Glass
Inbd Surface of Glass
Hat Sect. (Bottom)
#12 str.
HatSect.(Bottom)#12_r.
R.H. Console Z126. 5
Hat Sect. (Side)
#12 _r.
T/MXmitterPwr
supply
Hat Sect. (Top)
#12 Str.
Between C&S Band
Beacons
Hat Sect. (Bottom)#24 Sir.
Hat Sect. (Bottom)#24 Sir.
Transmitter Equipment
Hat Sect. (Side)
#24 sir.
Hear Calibrator
H20 Tank
Hat Sect. (Top)
#24 Str.
Near Inverters
Left Side Window
(Bolt)
49 49 49 4950 50 50 50
51 51 51 51
52 52 52 52
53 53
53 53
54 54 54 54
55 55 55 55
56 56 56 56
57 57
57 57
58 58 58 58
59 59 59 59
...... "" _'-"T'| | I
PAGE, 1_-_4
REPORT SEDR 104
MODEL_ 133
Mc'DONNEL LST. LOUIS, MISSOURI
DATE i November 1961
REVISED ,
REVISED,
Table 13-3. Special Instrumentation Parameters Commutator Point Assignments
(Low Level - 3MV to 13MV Ref. ) (Cont 'd. )
Nc_enclature
Capsule
Pole #I
Chan.
6o
61
6263
64
65
66
6?
68
7o71
72
?3
75
Chart.
60
61
62
63
64
65
66
67
68
7o?_
13' #l _2
Chsa. Chart.
6o 6o
61 61
62 62
63 63
64 64
65 65
66 66
70 70
67 67
69 69
?l 71
72 72
73 73
74 74?5 75
76 76
77 77
Left Side Window
(Inner Pa_e)
React ion Control
Thrust Chamber
Auto 24#
Auto l#
Man 24#
L/O Auto System Roll
Catalyst Bed L.H.
L/O Auto System Roll
Catalyst Bed L.H.
Heat Shield
Attaching Bolts
Horizon Scanner (Pitch)
Horizon Scanner (Roll)
Ant. Comp. (Outer Skin)
Ant. Comp. (Inner Skin)
Ant. Camp. (Outer Skin)
Drogue Chute Can
Parachute Comp. (Outer Skin)Lx
Ant. Comp. (Inner Skin)
Ant. Camp. (Outer Surface)
Retro Rocket Package
z88 _ xoz88 _ XRZ88 _ X5
Retro Rocket (Right)
Retro Rocket (Bottom)
76
77
72
?37475
76
77
/
J
_MAC 231C (Rev 14 Oct,. 55)
OATE._._ November 1961
REVISED.
REVISED
MCDONNELL, ,, ST. LOUIS 3, MISSOURI
PAG-- 13-55
R,_PORT"aEDR ,_O_
MO0_. 133
Table 13-3. Special Instrumentation Parameters Commutator Point Assignments(Low Level - 3 MY to 13MY Ref.) (Cont'd.)
Nomenclature
Capsule
Pole
Retro Rocket (Exp.Bolt)
Adapter SidesOuter Skin
z79.5 LX
Z79.5 By
Z79.5 Ty
Cap. to Adapter
Retaining Ring
Adapter RetainingRing Cover
Cap. to AdapterExplosive Bolt
Escape Tower Z257
Escape Tower (Left Leg)
Pylon Explosive Bolt
Stability Wedge
Chart.
79 798o 8081 81
82 82
83 83
8_ 84
85 85
86 86
87 87
88 88
Chart.
13
798o
81
82
83
84
85
86
87
88
78
798081
82
83
8_
85
86
87
88
13-111. Astronaut Observer Camera
The mechanism of the Astronaut observer camera is the same as that of the
instrument panel observer camera.
13-I12. TEST CONFI_ION CAPSULES
13-113. TEST CONFIGURATION CAPSULE NO. 8
13-i14_ S_stem Descriptiom
Instrumentation on Capsule No. 8 is similar to the Specification Compliance
Capsule (Refer to Paragraph 13-1) except that Capsule No. 8 is intended for an
unmanned (orbital) mission and therefore data pertaining to the Astronaut does
not apply. Major components installed are shown on Figure 13-5.
PAGE 13-56
REPORT aEDR 10_ "
MODEL. 1_3
.MC'DONNEL LST. LOUIS, MISSOURI
DATE 1 November 1961
REVISED
REVISED
13-115. S_stem OEe ration
Instrumentation on Capsule No. 8 is similar to the Specification Compliance
Capsule except is unmanned and operation is either automatic or ground controlled.
13-116. Monitoring Instrumentation
Monitoring instrumentation is the same as the Specification Compliance
Capsule (Refer to Paragraphs 13-3 through 13-58) except as noted in Paragraph
13-117 through 13-136. Figure 13-6 shows the various parameters monitored while
Table 13-2 lists the commutator point assignments for each parameter.
13-117. Standby Battery
The standby batteries, ON, signal comes from the secondary bus relay No. 2.
This relay energizes if the STANDBY BATTERY switch is in AUTO position and main
bus voltage is below 18 volts. Wlth the relay energized, 2_ volts d-c (nominal)
is applied to an attenuator in package C. Attenuator output (2.8 volts d-c
nominal) is applied to the commutators.
13-118. Environmental Control System
Environmental Control System instrumentation on Capsule No. 8 is the same as
the Specification Compliance Capsule (Refer to Paragraphs 13-16 through 13-22)
except that the emergency oxygen supply and 02 partial pressure is not monitored.
13-119. Reaction Control System
Reaction Control System instrumentation on Capsule No. 8 is the same as the
Specification Compliance Capsule (Refer to Paragraph 13-23) except that the hand
control positions are not monitored.
13-120. Acceleration
Capsule acceleration instrmmentation consists of circuitry which monitors
acceleration along three mutually perpendicular axes. The launch and re-entry
signal range is utilized throughout the flight. (Refer to Paragraph 13-30. )
DATE
REVISED
. REVISED
1 NOVEMBER 1961
M 'DONNE ST.LOUIS, MISSOURI
PAGE
REPORT
MODEL
13-57
SEDR 104
133
f- -
i 7
J CAPSULE _ AMPLIFIERS
ELECTRICAL
POWER SYSTEM RECTIFIER
AND
_ ATTENUATOR
INSTRUMENTATION:
POWER SUPPLIES
RECTIFIER
AND
SOUNDLEVEL j_ ATTENUATORPRESSURESTATIC
VIBRATION i
DATA
ANALYZER
CALIBRATIONON
ATTENUATOR
EN V_O:NT_N TA L
AMPLIFIERS i
ACCELERATION
--_ AMPLIFIERS
• r r REACTIONCONTROL
SYSTEM
AMPUFt_S L
AMPUFIERS,
!I •
J _:i ATTENUATORSR ATTENUATOR
I co _2_7_ i iAL
MF_GA ENC LAUNCH
I E Y SEQUENCE i
SEQUENCE / !' LANDINGSYSTEM
g_%_TON _ESSUREI
MONITOR j HEAT SHIEW
J CAVITYPRESSURE
\ IN_'Si _
"A" COMMUTATORO-3VINPUTSD--C J_
PDMCONVERTER
PULSE AMPLITUDE
r MODULATION
I 10.5KC VOLTAGE
CONTROLLED
OSCILLATOR J
II
COMPENSATING I
OSCILLATOR !
J MIXER "A"
2.3KC VOLTAGE
CONTROLLED
OSCILLATOR
1.7KC VOLTAGE J J
CONTROLLED
OSCILLATOR
1.3KC VOLTAGE J J
CONTROLLED VOSCI LLATOR
PULSE DURATION
MODUI.ATION q
J TAPE
RECORDER
1.3KC VOLTAGE
CONTROLLED
OSCILLATOR
2jKC VOLTAGE
CONTROLLED
OSCILLATOR
2.3KC VOLTAGE J_
CONTROLLED
OSCILLATOR
, f
MIXER "5"
J 10.5KC VOLTAGE
CONTROLLED
OSCILLATOR
LSE AMPLITUDE,
oo::,o CONVERTERJ t
..,,-co,._. I T0-3V D<
IN.ITS /_cooiNG ANDCO"TROL_'--'=_ \,,
Lv
HI FREQ
TELEMETRY
TRANSMITTER
81CONE
ANTENNA
TANTENNASWITCH
iDESCENT
ANTENNA
MULTIPLEXER _-- J)
TRANSMITTED
1. AND RECEIVED
VOICE SIGN_,LS
J LOW FREQ. J_
TELEMETRY
TRANSMITTER
TRANSMISSION
FIGURE!3.6 , MONITORING INSTRUMENTATION BLOCK DIAGRAM-CAPSULE 8 _
" |" _llm
VVlIIF lii..,BL.--ll l Ff=_T.i,
REPORT 8.EDR 104 ST. LOUIS, MISSOUR! REVISED
MODEL 133 _ C_.t_:_'_"_,_,.._. REVISED
13-121. Structural Temperature
Structural temperature instrumentation is similar to the Specification
Compliance Capsule. (Refer to Paragraph 13-32 through 13-38. )
13-122. Attitude Rate
Attitude rate signals are not commutated in Capsule No. 8. These signals
are applied to the voltage controlled subcarrier oscillators that are utilized
for electrocardiograph signals in the specification compliance capsule. A 0 volt
signal level represents a zero attitude rate. A -1.5 volt signal represents a
decreasing rate of 6 degrees per second and a plus 1.5 volt signal represents an
increasing rate of 6 degrees per second. Pitch rate is applied to the 2.3 Kc.
oscillators. Roll rate is applied to the 1.7 Kc. oscillators. Yaw rate is
applied to the 1.3 Kc. oscillators. Pitch and roll signals are routed through '
a biasing circuit to set up the proper center frequency signal level.
13-123. Parachute Compartment and Heat Shield Cavity
The parachute compartment pressure and the heat shield cavity pressure on
Capsule No. 8 is instrumented and monitored on the HF and LF Commutators.
13-124. Landing System Sequence
Landing system sequence instrumentation on Capsule No. 8 is the same as that
for the specification compliance capsule (refer to Paragraph 13-54) except clr-
cuitry is added to monitor the Betro Rocket Gone Relay #i, #2, and #3. This
signal is obtained from relays in Retrograde Relay Box #2. These signals are
monitored on the KF and LF Commutators. The Landing Bag and Astronaut's Abort
Systems are not monitored on Capsule No. 8.
13-125. Instrumentation Control
The instrumentation system controls and programs (See Table 13-I) power to
its own and other system equipment.
a. The earth and sky camera viewing from capsule window Operates at I0
_ jr=,ir,.,ib,,L...-,,v,.,,,._.,v-,,_ _v .._ _,_1t, ,J.,L,F,,IL,_dlll ,I. ,A,_"IL IIj
DATE . . , n =:A.N_V.M.ER 1961
REVISED ,,
REVISED
Mc'DONNELjL_r_ __ST.LOUIS. MISSOURI
PAGE.
REPORT,
MODEL
13-59
SEDR 104
133
SEE TABLE. t3,-Z POR
PROGRAMMER TIMING
MAiN I
F_E-I M_'ACt"D-C. _.U5
-rAPF. R_C.ORO ON' LtO.M'I"
5_CONO_Ry 1
PR.E -IMPACT
34V D-C BUS
CAMERA _ID _h;_REC.oI_oER SWl-rCH
EARTH ANO 5V_y CAMERAFILM W 1"f14 DR.AW_.
;kISTRU_NT _NE.L. CAMEI_._Flt..M wFrHoR,_WL I.._,Gh"T
_'r_.O_UT OR. P_<_r_o_P¢.
_.IC_H_" •UMG_LIC.AL
COhl N P..c:*'lr0 I_,
IIJ
TAPI= RDERI_g.c¢
rC i
III
III
i
,
IPROVISI(_I5
FOR
AST _:ON/_,UT
-¢AMIERA
IN_-r_UI_ENT
I_ANEI.
GAM5R_
11r
i ''=
I
..£___ ._L
_-I' IANO _'Y (
CAMER._l
REPORT SEDR 104 ST. LOUIS, MISSOURI REVISED
..._'"_.'.,_"x"=__ _ --',"=_,'_LL=----_ REVISEDMODEL. 133 ...............
i November 1961
bo
C.
mission.
d.
frames per minute throughout the mission.
The periscope camera operates at 180 frames per minute throughout mission
The instrument panel camera operates at 360 frames per minute during the
Calibration voltages; R-calibrate for percent full scale readings and
Z-calibrate for zero readings, are supplied from the block house before the
umbilical is dropped, just prior to launch.
13-126. Recordin_ Instrumentation
Recording instrumentation consists of a tape recorder and three cameras. The
Astronaut observer camera is omitted from Capsule No. 8 and a periscope camera is
installed. (See Figure 13-7.)
13-127. Tape Recorder
The tape recorder used on Capsule No. 8 is the same as the recorder used ca
the Specification Compliance Capsule (Refer to Paragraph 13-64), except the tape
speed is 151 inches P/S. Also a Playback Tape Recorder to simulate the Astronaut'
voice (See Section XI and Table 13-1) is installed.
13-128. Cameras
Capsule No. 8, being an unmanned vehicle, does not use an Astronaut observer
camera. The usual Instrument Panel Camera is installed, as well as an Earth and
Sky Camera and a Periscope Camera. (See Figure 13-5). Programming is received
from the capsule programmer.
13-129. S_ecial Instrumentation
A special instrumentation package is used in Capsule No. 8. The package con-
sists of fuse blocks, special instrumentation relays, five track tape recorder and
a nine channel amplifier case with associated power supply. Vibration measure-
ments are taken from various points on the capsule, amplified and applied to
separate tracks on the package tape recorder or to the capsule tape recorder. A
--_......#_._ a. A4L2a/dl_ JL JL_kJL4--
MAC 231C (Rev 14, Oct. 55)
OAVE i l_ovember 1961
RI_/ISKD
RE:V|S£D
ST. LOUIS 3, MISSOURI
rA_tt, 13-61
R_o_r 8EDR 104
MO._ 133
i _ _
_L
total of nine tracks are recorded as follows:
P.a.llet Tape Recorder
Track 2 Z123 Ring Longitudinal
Track 3 Web at LXl2
Track 4 Instrument Panel
Track 6 Z123 Ring Tangential
Track 7 Parachute Compartment
13-130. System Units
Ca_st_le Ta_e Recorder
Track 2 Package "B" Noise Level
Track 4 Z123 Ring Radial
Track 6 Shingle Strain Gage
Track 7 Escape Tower
The system units used on Capsule No. 8 are the same as the Specification
Compliance Capsule except as noted in the following paragraphs. (Refer to
Paragraphs 13-68 and 13-69. )
a. Main and Reserve Oxygen Pressure Potentiometers are not instrumented
b. Respiration rate and Depth Potentiometers are not monitored
c. Control Stick Potentiometers are not monitored
d. Body Temperature Transducers are not used
e. Electrocardiogram Pickups are not used
f. Tape Recorder is the same as specification compliance capsule.
(Refer to Paragraph 13-82. )
13-131. Sound Level and Vibration Transducers
A microphone and amplifier are located in the B package for sound level
pickup. A piezo-electric/diap_ type microphome is used to pick up a pressure
level of lO0 to l_O db with a frequency range of 37 to 9600 cps. The variable
gain, transistorized amplifier consists of am impedance matching stage, filter
network, buffer Stage and final amplifier. Amplifier output, 3 volts R_, is
used with a data analyzer. A piezo-electric crystal is used as a vibration
pickup, measuring a frequency range of lO to 2000 cps. The tramsducer is used
in com_ctioa with vibration amplifier.
v_n Is mlimwm_a_ i aa_
Mc,DONNEL L REPORT SEDR i0_ ST. LOUIS, MISSOURI REVISED
MODEL. i_ _,z_ *" z.u_J._ * ,._L_:r" REVISED.
DATE i November 1961
13-132. Vibration Amplifier and Data Analyzer
A transistorized amplifier is used to increase the output of the vibration
transducer to a level compatible with an accoustical and vibration analyzer.
13 -133. Programmer
The programmer contains switch contacts which operate control circuits for
specific intervals of time. The unit for Capsule No. 8 consists of one section
of wafer contacts. When power is applied to the programmer the electronic con-
trolled timers continuously operate the following wafer contacts.
Wafer Section Duration Rate
i 30 Seconds I Per 30 Min.
2 ii0 Milliseconds 6 Per Second
3 90 Milliseconds i0 Per Minute
llO Milliseconds 3 Per Second
13-13_. Instrumentation Package "C"
Function
Water Extractor
Instr. Panel Camera
Earth & Sky Camera
Periscope Camera
The horizon scanner output amplifier card in package "C" provides a bias
voltage to keep the horizon scanner pitch and roll signals within limits compatiblq
with other system components. Calibrate control relays are installed on the card.
Roll ignore and pitch ignore signals are Jumpered through the card. The Horizon
Scanners operate continuously from launch to re-entry.
13-135. Instrumentation Package "A"
The instrumentation package "A" is the same as the Specification Compliance
Capsule except for the removal of the body temperature amplifier card.
13-136. Instrumentation Package "B"
The components used in package "B" of Capsule No. 8 are the same as in the
package "D" in specification compliance capsule except for the following units.
(Refer to Paragraph 13-102. )
a. Attitude Change Rate Filter and Calibrate Cards
__/_/'_lLTT'?ITT'%lf_lLVlSqlT • -- --
._=4.,_r 4. I .A ._.._.r,.A.&, .._ 41.,_LJL.4
MAC 231C (Rev 14 Oct 55)
1No ber Mc,DONNEg, • ,... 13-63REVISED ST. LOUIS 3. MISSOURI REPORT. _DR 104
R_V,S_D t-*_'-__ ........ Homa. 133
/
'\ : _j!
Attitude change rate filter and calibrate cards replace the electrocardiograph
amplifier cards. Three attitude change rate cards are used. They are for
pitch, roll and yaw parameters. The cards associated with pitch and roll rate
provide for application of calibrate signals to these channels. The yaw rate
card provides this function and also contains a bias battery to set up center
frequency on the voltage controlled oscillator when yaw rate equals zero degrees
per second.
b. Voltage Controlled Oscillator Cards
The oscillator cards used in Capsule No. 8 are similar to those used in the
specification compliance capsule. (Refer to Paragraph 13-107. ) However, attitude
rate signals are applied to the oscillators instead of aeromedlcal signals.
Oscillator functions and frequencies are indicated in the following list.
HF Commutator
KF Commutator output is applied to the 10.5 Kc oscillator.
Yaw Rate Is applied to the 1.3 Kc oscillator.
Roll Rate is applied to the 1.7 gc oscillator.
Pitch Rate is applied to the 2.3 Kc oscillator.
Compensating Oscillator frequency is 3.125 Kc.
LF Ccv_nutator
LF Commutator output is applied to the 10.5 Kc oscillator.
Yaw Rate is applied to the 1.3 Kc oscillator.
Roll Rate is applied to the 1.7 Kc oscillator.
Pitch Rate is applied to the 2.3 Ec oscillator.
13-137. TEST CONFIGURATION CAPSULE NO. 9
13-138. S_stem Description
The instrumentation system on Capsule No. 9 is similar to the Specification
Compliance Capsule (Refer to Paragraph 13-1) except that the capsule will be
ii
Ii
o ,,,
_ m_
INSTRUMENTATIOn"
PACKAGE'_"7
I_ADIAI'DETECTOR
6
ETER.
RECORDERVOICE AND TELEMETERINGPDM- COMM. A(ttI_,H FREQ)PDM- COMM. B (.LOWFREQ) '
XI 0
g g mm mo o
I -I Z' 0
<I m
I"m
I
I
r
3 m >
_ m
m
PAGE 13-66
REPORT _EDI_ 104
MODEL 13_
_ DATE i November 1961
ST. LOUIS, MISSOURI
•_'_ _, *" z _ . . ,_ REVISED
unmanned and a special instrumentation pallet and primate couch replaces the
Astronaut's couch. Major components installed are shown on Figure 13-8.
13-139. Monitoring Instrumentation
Monitoring Instrumentation differs from the Specification Compliance Capsule
in that Capsule No. 9 is instrumented for primate occupancy. Figure 13-9 shows
the parameters in block diagram form while Table 13-2 lists the commutator point
assignments for all telemetered instrumentation. Deviations from the Specifica-
tion Compliance Capsule are explained in Paragraphs 1S-l_O through 13-157. )
13-140. Standby Battery
In Capsule No. 9 the standby battery power is connected directly to the
main d-c bus, in parallel with the main batteries. The standby d-c auto light
or automatic switching of the standby to the main bus is not used.
13-141. Special Instrumentation Pallet
The Special Instrumentation Pallet is installed in place of the Astronaut's
Couch. The pallet contains an array of circuits for instrumenting the various
parameters as listed on Table 13-3. Location of the sensors are indicated on
Figure 13-11 and the point assignments for each parameter are listed on Table
1B-3. The values measured by each sensor is encoded by the multicoder and
recorded on a tape recorder. Power for the pallet is obtained from the 150 volt-
ampere inverter and controlled by a multicoder "ON" - "OFF" switch.
13-142. Blood Pressure Oscillograph
The oscillograph-tape recorder, which records the venous and arterial blood
pressure of the chimpanzee is mounted beside the Special Instrumentation Pallet.
The oscillograph is powered by its own, self contained, wet-cell battery. The
power is controlled by an external ON-OFF switch. A transducer is attached to
the Solar-Plexis of the chimp, which converts blood pressure values to electrical
signals. These signals are routed through an amplifier in Instrument Package "A",
j:---\
DATE
REVISED
REVISED
1 NOVEMBER 1961
McDONNE__ __ST.LOUIS, MISSOURI
-LC .... ,.,,, , ,....,,,r-..,,'w
PAGE. 13-67
REPORT • SEDR 104
MODEL 133
.f'?-\
/ " .
PAGE 13;'68
REPORT SEDR 104,
MODEL 133
Mc'DONNE ST.LOUIS, MISSOURI
_ "_,,uJ i,,,/ L-
DATE .
REVISED
REVISED
1 NOVEMBER 1961
TAPE _ECOROE_ PP.E-
IMPACT MAiN D:C._,US
P.ETIRO - P.ECOV E RYZ4V D-C BUS
_-_l SCOPE CAMERA_.4V MAIN D- C _U5
CAMERA At_D TAPE_E.CG_E)E _ SW_TCH
"ON" L_GNT
EA_.TN A_D SKY CAMERA_LM XN _T_ b_ AWAL
L_GMT
PERt 5COPE CAMERA
_'_,M W|T_D_AYVAL
-/l{- .%E._. TABLE _,3-?-PROQ_AMER TI MIt_G
CAMEP, A AN(::) TAPE
I
CAMERA ANO TAPEI::::_,ECOI:_DE _, P,EL AY
I
T
TA_E
_.ECO_.bE _
COMBINEb, C_IVES THE
qOt4F!C_U_ATION.OF CAPSULE 9
PER|_COPP. CA_E_A E.A_:_,TH At4O _KY CAMERA
FIGURE _3-10CAMERA5 AND TAPE RECOP, DE_. CONTg, OL C_RCLI|T CCAPSULE.'9) .
p t_l IM =' I I"l =' I. _''JL-_l-"--_ --- --. _ • in II-T-
DATE
REVISED
REVISED
1 NOVEMBER 1961
M;DONNELLST.LOUIS, MISSOURI
"-_Vl _qll lug,,,, i'qF- I A I
PAGE, 13-_69
REPORT SEDR 104
MODEL 133 *
ITEM DESCRIPTION LOCATION
/ ' RETRO ROCKET PACKACE _ETEO ROCKET - RIG_) RX rY2 ADA_1ER SLOES - OUTER SKIN Z79.S
3 RETRO ROCKET PACKAGE Z88.0 RX BY4 RETRO ROCKEt EXPLOSIVE BOLl zsl 6 XO YO
5 ABt_.TION SHIELD - BETWEEN .06 AND .E FIBERGLASS RX33 TY36 RETRO ROCKET PACKAGE Z88.0 XO _Y
HEAl SHtELDATTACH POINT ZIC,4,5 XO TYCAPSULE IO ADAPTER EXPLOSIVE BOLT ZIO0,O XO TY
0 H202 lANK. AUTOMATIC Z104, 5 XOCONICAL SECTION SIDES - OUTER SKIN ZI07.0 TY
It ADAPTER SIDES - OUTER SKIN Z79.5 LX12 ABLA¥1ON SHIELD - BETWEEN .06 & .2 FIBERGLASS LXI TY2
13 CABIN AIR IEMPERATURE ZIDI.5 LXI8 _o14 ?]AT SECTION - 124 STRINGER, 9OTTO*_ z127.0 TY
IS CONICAL SECTION SIDES - OUIER S KIN ZI26.5 XO TY_6 HAT SECTION - #E4 STRINGER, TOP, Z127.0 XO TY
t 7 fIAT SECTION - 124 STRINGER, SIDE TYZI27.0IS I_T SECIIOH. e24 STE_NGER. BOTTOM {7 V2 o) zl27,0 IY
19 CONICAL SECTION SLOES - OUTER SKIN Z107.0 LX2O WINDOW - BOLTEl WINDOW - INSIDE OF OU?ER PANE22 WINDOW - ZNNEA PANE
23 L/O AUTO SYSTEM ROLL - F'_O2 LINE Z_Oe. 8 LX DYI24 IJO AUTO SYSPEM eOLL. SOLENOID VALVE ZIIO.2 LX BYI
25 I_#O AUIO SYSTEM ROLL.CATALYS T BED (L. H.I ZII4.0 LX 9YI
26 H202 LtNE CONICAL SECTION - AUTOMATIC Zl26,527 CONICAL SECTION SIDES - OUTER SKiN Z154. 5 1_f29 CONICAL SECTION SIDES - OUTER SKIN Zt54.S IX
SOLENOID VALVE TOF 24 IJ3. ZITO.5 XORFJ_CTION CONTROL THRUST CHAMBER - TOP (AUIO t24I Z178.0 XO
31 PARACHUte COMPARTMENT - OUTEe SKIN Zi78.0 XO T_32 REACTION CONTROL THRUST CHAMBEX - TOP (AUTO 01) Z179.O LX
33 TOWER TO CAPSULE RETAINING RING (TOP)
34 FARACHUTE COt,APARTMENT - OUTER SKIN
5 R_COVERY COt_PART_ENT INTERNAL SleUCIUKE6 ANTENNA COMPA_Tt_ENI INNER SKiN
37 HORIZON SCANNER (ROLE)
3e ANTENNA CO_AelMENI Lm
39 ANIENNA COMPARTMENT Oute_ SKIN (_OTTOM)
40 _ECOVERY COMPARTMENT iNNER SKIN
41 PYLON EXFLOSIVE BOLT
42 PYLON EXPLOSIVE _OLT COVER
_3 IOWE_ IO CAPSULE RETAINING RING (SOTIOM),W TC_R TO CAPSULE REIAINING RING - BOTTOM
45 PARACHUTE COMPARTMENT SIDE
46 INNER SKIN BENEAIH JET SUPPCeT
47 RECOVERy COMPAeTMeNI tNNEe SKIN
48 H202 LiNE - TOP 24 LB.
49 CONICAL SECTION SIDES - OUTEe SKIN
_0 CABIN AiR TEMP_RAIUXE
_I HAT SECTION BOITOM - 7 _/_" #)2 SInINGER
52 HAl SECTION BOTTOM = J12 STRINGE_
HAP SECTION SIDE - r_2 STR_NGERHAT SECTION IOP - _12 StrINGER
SS CONiC._L SECTION SIDES - OUTEe SKIN
5_ _EAI EXCHANGE (OUTLET)
57 HEAl _XCHANG_R 0NLEI)
se CONICAL TECIION SIDES - OUTER SKJN
_OAFtEe REIA_NING RINGAD_PIER REIA_N_NG X_NG COV_
61 ABLATION SHIELD - SHIELD INBOARD SURFACE
62 A_LATION SHIEL_ eETWEEN.2FtBEeGU_SS AND HONEYCO_
RETEO ROCKET PACKAGERETeO ROCKEt PACKAGE _REIRO ROCKET - 9OTT(_)
65 AOAPTER SIDES - OUTE_ SKIN¸
66 ABLATION S}IIELI)o SHIELD iNbOArD SURFACE
67 HEAl SHIELD - BULKREAO CAP
_ ESCAPE TOWERESCAPE TOWER
ZI84.6 XO TY
zlTa.O IX YOZI_.6 IX6 BY12
Z)91 XO _Y
Z_.l XO BYE.5Z191 xo BY
Z}79,0 RX5 BYI3Z184.6 RX85 BY
ZI84.6 XO BYZI85.7 LX.5 BY
Zl78.0 xo
Zl;7.0 _2 TY_3ZI 7"/.0 TY
Z168.5 xO TYz_.o BY
Z_45.5 xo YO
Zl2/,0 ByZI27.0 BY
ZI27.0 8YZl27,0 _Yz126 XO BY
Z107.0 BYZ)04.5 XO OY
ZI04.5 XOxo _2
z_.0 sY
xO eYZ/%5 BY
Z79.5 0Y
RX33 TY2
Z257.0 xO TY_
zt(_.s LX t_"
ABLATION SHIELD BEIWEEN,2 FIBERGLASS!AND'HONEYCOM RX33 TY3REACTiON CONTROE THRUST CHJU_ER -TOP (MANUAL e24 Zl 7S.O IX TY
Z104.5 XO7Z HE_ z TANK - _AI_UAL
_3 _O2 UNE CONICAL SECTION - MANUAL Z_2_.S
ZI23.O74 EMERGENCY EGeESS HATCH - OUTER EDGE
_5 E_ERGENCY EGRESS HATCH - OUTE_ Em3E Z_46,_
7_ EMEeG_NCY EGeESS HATCH - PERCUSSION IGNITEe. TOP ZIEB.5
S _ND BFACON ZI07.S
_0 EGRESS HATCH Z124,25EG_SS HATCH Z147.2581 TM TRANSMITTER
82 TM TRANSMITTER83 )N'_AERTER
B4 INVERTER
FIGURE 13- 11 THERMOCOUPLE LOCATIONS (CAPSULE 9)
P/A4_-2'JEA ,
PAGE, 1_-_0
REPORT _EDP, 104
_.MODEL, 1-_._.
REVISEDST. LOUIS, MISSOURI
• _ REVISED
where the signal strength is sufficiently increased to be recorded on the
oscillograph. These signals are not telemetered to the ground.
13-143. Parachute Compartment aud Heat Shield Cavity
The parachute compartment and heat shield cavity pressure is monitored on
the HF and LF Commutators. The pressures are measured by transducers, the out-
puts of which are liaear from 0 to 3 volts for a pressure range of 0-15 psia.
13-144. Inactive Parameters
The following circuits are not monitored on Capsule No. 9:
a. Control Stick Position
b. Abort Switch
c. Landing Bag
13-145. Yaw Data Converter
A Yaw Data Converter is installed on Capsule No. 9- The fumctlon of the
converter is to provide calibrated conversion of the ASCS slaving, output signals
to compatible input signals for the Telemetry Commutators. Two d-c signals are
produced which are proportional to the "Roll to Yaw-Low Gain" and "Roll to Yaw-Hi
Gain". Also a third d-c signal is produced waich is proportional to the 'Maw
Torquing" a-c signal of the ASCS calibrator, and is designated as the "Yaw
Torquer" signal. Zero yaw attitude is represented by 1.75 volts output for the
low-gain signal and 1.5 volts for the hi-galn signal. Meter ranges for all three
signals is 0-3 volts d-c and represents yaw commands of -45 to 45 degrees for
lo-gain, -lO to lO degrees for hl-gain and -6.5 to 6.5 degree/minute slaving
rate, for the yaw torquer signal. These signals are telemetered on both the HF
and LF Commutators. The Yaw Data Converter is powered by ll5V 400 cps a-c, 3.0
volts d-c and 23.5 volts d-c supplied from the capsule power circuits.
13-146. Horizon Scanners
The horizon scanners are instrumented on both HF and LF Commutators.
DATm 1 November 1961
REVISED
REVISED
W,DONNE , ST. LOUIS 3. MISSOURI
PAGE 13-71
REPORT _nR ] 02
MO._ l_
i
C_
13-147. Retro Rocket Firing - Velocity Sensor
The velocity sensor is not instrumented, instead, each of the three Retro-
Rocket firing signals are applied to the KF and LF Commutators.
13-148. Body Temperature
The chimp body temperature is monitored and applied to the EF and LF
Commutators.
13-149. Command Receiver
The cammand receivers all channel signal is instrumented on Capsule No. 9.
13-150. Aeromedical
Aeromedical instrumentation on Capsule No. 9 is the same as that for the
Specification Compliance Capsule (Refer to Paragraphs 13-39 through 13-43), except
that since a primate is being used the scale factors for EED and respiration data
are different and only three EED transducers are used on the primate.
13'151. Primate Instrumentation
The primate couch is mounted on the special instrumentation pallet and is
instrumented to monitor the prlmate's reactions and response. Since Capsule No.
9 is assigned to a specific mission, the primate instrumentation is programmed
differently than other primate test capsules. The primate Programmer consists
of an electronic panel assembly and a liquid feeder. The instrument panel con-
tains three display units, the left-hand unit consists of six symbols and a cyan
blue disc display, the center unit has six symbols, a yellow disc and a white disc
display and the rlght-hand unit has six symbols, a red disc and green disc display.
Three switch levers are provided for primate actuation, with each switch closure
recorded on a four digit counter mounted near the left switch lever. A pellet
dispenser is mounted on the panel and will dispense, on command of a 28 volt pulse,
reward pellets through a hole in the lower rlght-hand corner of the panel. A
liquid dispenser is mounted in the vicinity of the right-hand side of the animal's
I November 1961
head. It will dispense a drink of liquid when actuated by a 28 volt pulse and
triggered by the animal's lips. Power is supplie_ for the prlmate's couch from
the special instrumentation fuse block through the special instrumentation record-
er relay. The animal's response during the various program sequences is recorded
and telemetered as listed on Table 13-2. The programmed psychomotor presents the
sequence of test functions the animal must perform and are as follows:
TEST TIME
1. RS-CS - Regular and classified shock avoidance 15 min.
2. TO - Time Out 2 min.
3. DRL - Differential reinforcement of low rates 5 rain.
4. TO - Time Out 2 rain.
5. FR - Fixed ration reward 5 min.
6. TO - Time Out 5 rain.
7. I_PM/NRPM Position and Negative reinforcement
Perceptual monitoring
(a) 18 Combinations of symbols 5 rain.
(b) 18 Combinations of l, -3 X symbols
8. TO - Time Out (Note: If the last 36 combinations are
completed before the 5 minutes has
elapsed the remaining time is out
and the entire program starts over).
A description of the tests in the sequence they occur and required telemetered
data is as follows:
a. RS-CS (time 15 rain.). The telemetered outputs are as follows:
(1) Program identification through indication of shock.
(2) Cumulative response of left lever.
(3) Cyan blue light on and duration.
,,,,CCL'F.'DZT_TIAL
DATE i November 1961
REVISED
REVISED
bl
Co
Mc'DONNE Er. LOUIS 3, MISSOURI
_--_r_=vlii ii_i--i_l i l/'lill
(4) C_ative response of right lever.
TO (Time 2 rain.) No displays.
Telemetered output s:
(i) Identification (zero input to TM)
(2) Cure. response of left lever
(3) Cure. response of center lever
(4) Cure. response of right lever
(mime 5
d8
e.
pAG_II 13-73
REPOR_ SEDR 104
MOO , 133
Green light in right display unit is activated. If animal waits 20 seconds
before pulling right hand lever, a green light appears beside the liquid
feeder and a drink may be obtained. If R/H lever is pulled before 20
seconds time span, the animal must wait an additional 20 seconds for
another opportunity for a drink.
Telemetered outputs :
(1) Program identification
(2) Cum. response left and center levers
(3) Liquid feeder actuation
(4) Cure. response right lever
TO (Time 2 min.) NO display.
(Time5
Yellow light in center display unit is activated. If animal presses
center lever 50 times, the pellet dispenser is actuated and he gets a
pellet as a reward.
Telemetered outputs :
(1) Program identification
(2) Ck_m. response of left lever
(3) Cure. response of center lever
..... "ill_ll I_PI /_l' "__i_li.ollr llli I . =. /I
REPORT SEDR 104 ST. LOUI,S ,.. MI$$OIJRI_,.. REVISED
MODEL_ 133 _ _: _: _ _ _: T :._ _ REVISED
f.
g.
(4)
(a:ime 2 sin. )
Cure. response of right lever
(t Line 5 rain. )
NRPM mode of operation - If correct lever is pulled, 15 seconds of time-out
results, ,lights are out. The lights then come on with new display and there
is 20 seconds allowed for another decision. Correct lever sets beneath
odd display. If no lever or wrong lever is pulled_ shocks occur every
five seconds after 20 second decision time, until correct answer is made.
PRPM mode of operation - The lever beneath the odd display provides a pellet
reward. Depression of either of the other two levers results in no reward
and provides 15 second time-out after which same display appears.
Telemetered output s :
(1) Program identifier
(2) Cure. response of left lever
(3) Ctnm. response of center lever
(4) Cure. response of right lever
h. Ten minutes after impact the nipple on the liquid feeder is armed and the
green light appears allowing the chimp to obtain a drink.
13-152. Instrumentation Control
Instrumentation Control Circuitry for Capsule No. 9 is similar to the Speci-
fication Compliance Capsule except for the progra_ner which provides controlled
power (See Table 13-1) for the following circuits.
(a) The environmental control system water extractor is supplied a 30
second pulse every 30 minutes.
(b) The Instrument Camera is the same as ia the Specification Compliance
Capsule.
©
(c) The Earth and Sky Observer Camera is provided a 90 millisecond pulse,
ten times per minute, during the entire flight.
(d) The Primate Observer Camera is programmed the same as In the Specifica-
tion Compliance Capsule.
(e) The Periscope Camera is provided a llO millisecond pulse, 30 times per
second during re-entry aud 1800 times per hour while in orbit. The camera is
operable only when the periscope is extended.
(f) '_" & "Z" calibrate is the same as the Specification Compliance Capsule.
13-158. Recording Instrumentation
Capsule No. 9 Recording Instrumentation (See Figure 13-11) consists of a tape
recorder and four cameras, the tape recorder, Instrument Camera and Astronaut
Camera are the same as used in the Specification Compliance Capsule (See Para-
graphs 13-65 and 13-66), except that the Astronaut Camera is adjusted to view the
primate. A 70 millimeter Earth and Sky observer camera is mounted at the obser-
vation window. This camera is sighted to record a field of view through the
window. The image of a timer is superimposed on one corner of each frame. The
camera is supplied 24 volt power and pulsing voltage from the capsule power
system. Each pulse operates the shutter, exposing one frame and transporting the
film for the next exposure. Trigger pulses are applied to the camera at the rate
of lO per minute by the programmer. In addition, a periscope camera is used on
Capsule No. 9. The camera is mounted inside the periscope, to view the earth as
it would be seen through the periscope, by an observer inside the capsule. Power
is applied to the camera, only when the periscope is extemded. (See Figure 13-10. )
Pulses for the camera are supplied from the capsule programmer. (See Table 13-1. )
13-15_. System Units
System Units in Capsule No. 9 are the same as used om the Specification
Compliance Capsule, (Refer to Paragraphs 13-68 through 13-81), except the units
_l_i/ lJlm/L--'l _1 • I,'m .
PAGE 1._-76
REPORT _--D_ l N_.
MODEL. ]-33
MCDONNEL L C-__ DATE l November 1961REVISED
ST._:LOUIS, MISSOURI ....
REVISED
added in Paragraphs 13-155 thru 13-157.
13-155. Yaw Data Converter
The Yaw Data Converter is a small compact unit consisting of transformers,
demodulators, relays and summing circuitry.
section of the capsule near the periscope.
13-156. Periscope Camera
The unit is mounted in the lower
(See Figure 13-8.)
The Periscope Camera used on Capsule No. 9 is the same as the Instrument
Camera except that the periscope camera is mounted inside the periscope. (Refer
to Paragraph 13-110. )
13-157. Blood Pressure Oscillograph
The blood pressure oscillograph consists of a motor timer and tape recorder.
The oscillograph also has a self-contained battery mounted with the recorder.
The battery is vented in the same manner as the capsule's main batteries. The
oscillograph operates at a speed of 8 inches per minute and is on continuous
from the time the recorder's power switch is turned on, prior to launch, until
it is turned off, after landing.
13-158. TEST CONFIGURATION CAPSULES NO. I0, 13 AND 16.
13-159. General
The Instrumentation on Capsules No. lO, 13 and 16 is the same as the Speci-
fication Compliance Capsule (Refer to Paragraphs 13-1 thru 13-110. ) (See Table
13-2 for commutator point assignments. )
MAC 231CL (27 APR 59)