Technical Information Summary AS-501 Apollo Saturn V Flight Vehicle

8/6/2019 Technical Information Summary AS-501 Apollo Saturn V Flight Vehicle

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This document is prepared jointly by the Marshall Space FlightCenter Laboratories R-ASK-P, R-ASTE-S, and R-P&"E-VN. The documentpresents a brief and concise description of the AS-501 Apollo SaturnSpace Vehicle. Where necessary, for clarification, addirional relatedinformation has been included.

It is not the intent of this document to completely define theSpace Vehicle or its systems and subsystws in detail. The infomacianpresented herein, by rext and sketches, describes launch preparationactivities, launch faci lities, and the space vehicle. This informationpermits the reader to follow the space vehicle sequence Of events begin-ning a few hours prior fo li ftof f to its journey into space.

1. Mission Purpose:

The purpose af the AS-501 mission is to develop the Saturn Vlaunch vehicle for manned flights and to veri fy the adequacy of theApollo Command Module heat shield at lunar reentry velocities.

The AS-501 mission is an unmanned, elliptical earth orbitalflight.

2. Mission Objectives:

a. Demonstrate structure and thermal integrity, and cmpati-biliry of the launch vehicle and spacecraft , and determine structuralloads and dynamic characteristics during powered and coasting flight.

c. Ver ify launch supporr equipment compatibi lity, and missionsupport capability for launch and mission operations to high post-injec-tion altitudes and Command Module recovery.

d. Demonstrate the S-XC and S-II stage propulsion systems anddetermine in-flight system performance parameter.

e. Dwonstrate the launch vehicle guidance and control systemduring powered flight; achieve guidance'cutoff and evaluate system ac-curacy.

f . Demonstrate S-X/S-II dual plane separation and S-II/S-IVSseparation.

g. Demonstrate launch'vehicle sequencing sys tem.

h. Evaluate performance of the emergency detection system(FDS) in an open-loop configuration.



i. Demonstrate S-IVB stage restart capab ility.

j. Verify adequacy of the Command Module beat shie ld for xe-

entry at lunar ixYcur* cond itions.

3. Mission Profiles:

AS-501 will be launched from Launch Complex 39, Pad A, Kennedy

Space Center (KSC); at a launch azimuth of 90"E of N. Shortly after

liftoff (approximately 12 set) the vehicle begin s a roll maneuver to at-

tain a flight azimuth of 72% of N and main tains a near zero-lift

(gravity turn) trajectory through the maximm dynamic pressure region.

Afrer S-K burn and separation; S-II burn and separation, the first burn

of rhe S-I"% will propel the S-IVB/IU/Space craft into a lOO-nautical-

mile parking orbit using the Iterative Guidance Made (IGM). The vehicle

will remain in this orbit for approximately two revolutions with its'

longitud inal axis in the orbital plane and paralle l to the loca l horizon.

During the secon d revolution, when the vehicle is within tracking range

of KSC, the S-ITS en gine will be re-started to boost the vehicle into

an ellip tical atmosphere - intersecting waiting orbit with an apogee of

approximately 9,000 nautic al miles. Spa cecraft separation occurs ap-

proximately 590 seco nds after injec tion into waiting orbit. The coas t

time in waiting orbit between S-ITS cutoff and Command Module (CM) re-

entry is approximately 4.6 hours. Shortly after spacecra ft separarion

a service propulsion system (SPs) burn and navigational corrections

will be performed to achieve lunar retnrn velocity and the proper re-

entry corridor. Following the secon d SPS burn, Cormand Module/Service

Module (CMM/SM) separation will occur and the Ww ill be reoriented for

a guided lifting reentry which will produce the heat load desired to

test the CM heat shield at lunar returning velocity. Splash-down will

be near Hawa ii.

csn in

B...t to Earth &bit-S-K,



LIST OF FIGURES

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4

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GENERAL

AS-501 Spa ce Vehic le. ............

Launch Complex - 39. ... , ........ . .

LC-39 Pad “A” Configuration ........

SPACE/LAUNCH VEHICLE

Secure Range Safety System .........

Emergency Detection System (EDS) .....

S-K/S-II Stage Flight Sequencing ...... .

S-II/S-1VR Stage Flight Sequencing .....

S-IVB Stage Flight Sequencing ........

Trajectory Information (Boost Phase) ...

Guidance and Control System Block Diagram

Vehic le Tracking Systems ..........

Space Vehicle Weight YS Flight Time .... .

S-IC STAGE

S-IC Stage Config uration ...........

F-I Engine system. .................

s-x stage Propellant system ............

S-IC Stage Tk.mst Vector Control System .......

S-IC Stage Measuring System ............

S-IC Stage Telemetry System ............

S-IC Stage Electr ical Power and Distribution System

S-II STAGE

S-II Stage Configuration ...............

J-2 Engine System - S-II Stage ...........

S-II stage Propellant System ............

S-II Stage Propellant Management System ......

S-II Stage Thrust Vector Control System ......

S-II Stage Measuring System .............

S-II Stage Telemetry System .............

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LIST OF FIGURES

Title m

S-IVB STAGE

S-IVB Stage Configuration .... .... .... .... ....

J-Z Engine System S-IVB Stage .................

S-IVB Stage Propellant System. .................

S-IVB Stage Propellant Management System. ..........

S-IVB Stage Thrust Vector Control System . .. ... .. ...

Auxiliary Propulsion System ...................

S-IVB Stage Measuring System. .... .... .... .... .

S-IVB Stage Telemetry System. .... .... .... .... .

S-IVB Stage Electrical Power and Distribution System .....

INSTRUMENT UNIT

Instrument Unit Configuration .... .... .... .... ..

Instrument Unit Measuring System. ...............

Instrument Unit Telemetry System. ...............

Instrument Unit Electrical Power and Distribution System . ~

W/S-IVB Environmental Control Syst& ............

SPACECRAFT

Spacecraft 017 Configuration. .... .... .... .... ..

SC 011 Telemetry, USB & lJpaata systems. .... .... ...

Spacecraft Eleotrical Power 2nd Distribution System. .....

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A-

S-U 5+a9a (NAA]

- 396 ‘3

-- A-

A-

-396"

s-IC S+.9c(B.ein9)



.^^ I H-P Gas Facility

\

ICr*wlerwny

riquroz/I



rlobi,. *cr”/c.Strucfun

LsAnch

Umbilical Tower

Pad



-.--~-~__ F,,.. - ..__.

n



The secure range safety systems on the S-K, S-II and S-I"Bstages provide a communications link to transmit coded comands from

ground starions to the vehicle during powered flight, providing apositive means of terminating the flight of an erraric vehicle byinitiafing emergency engine cutoff and if necessary, propellant dis-persion.

Each powered stage contains tro UHF radio receivers. Both com-mand receivers on each of the thrre stages respond fo the same com,andsignals, each providing a backup sysrem for the other.

The safety and arming device located on each stage is armed by asignal from the blockhouse before v;thicle ignition. Pollowing S-IVBcutoff the S-I"B range safety system is "safed" by a cornand fromRange Safety Control to preclude accidental dertruct.



Chrg. I” ‘OX ranti~.,rr.r.rr.r..,...,.rr.....r,,r,r,,,r., ,,,,,,, u ,,,,,Iw,z



?zmeGmNcY DETECTlON SYSTEM (EDS)

The purpose of the EDS is to sense onboard emergency situationswhich arise during the boost phase of the fltght. On AS-501 the EDS

will be flown in an open loop configuration which precludes autamaticabort.

The EDS is comprised of sensors which detect malfunctions andlogic circuitry which initiate spacecraft displays and, in two cases,automatic abort o f the M. With the exception of the g-ball, mountedon top of the LET, the EDS sensors are located in the launch vehicle.The system's relay logic is located primarily in the IU EDS Distribu-tor and the CM MissLo* Events Sequence Controller.

The EDS has two modes of operation; 'hamal", which generatesabort cues and "automatic" which inifiates firing of the LES, and CMseparation in the case of tw S-IC engines out or angular overrates

during S-K powered flight. Figure is simplified black diagram ofthe AS-501 EDS. The automafic abort initiating portion of the systemconsists of ehe launch vehicle's rate sensing subsystem, the stage

thrust sensing subsystem and the signal distribution and processinghardware which services these devices.

The angular overrate sensors (3 per axis in pitch, yaw, and roll)will initiate automatic abort o f the CM during the period they areenabled (l if to ff to about 136 seconds) whenever two sensors in any oneaxis simultaneously indicate excessive rates. Detection qf the over-rates is made by the se~-mr switch circuirry of-the Control SignalProcessor in the I". The settings for these angular rate detectorsare 5 degrees per second in pitch and yaw and 20 degrees per second in

roll. The majori ty voting of the three switch outputs in each axis isdone by relay logic in the FBS Distributor. A valid excess rate deci-sion is forwarded by the EDS Distributor to the CM Mission Events Se-quence Controller for abort initiation.

The S-IC stage engine thrust OK sensors (three per engine on allfive engines) will also initiate abort during the period they are en-abled (l if to ff to about 135 seconds) when the voted output of the sen-sors from any two engines indicates that the thrust of those enginesis below the 89% level. Thes e sensors monitor the F-l engine’s fuel

inlet manifold pressure. Majority voting of the three sensors foreach engine is done in the EDS Distributor. A valid two engines antdecision is sent to the Mission dents Sequence Controller for CM

abort initiation.



Li

Si

0 dutohmt’ic. Abort Capabilifq (Not active ou $A-501)

t. 5-W Two ehlqiwes out (134.2+134..4)

f!. Excessive aNqular rstes(13+.6+L34.a)I

$ $-IC INboard Ewir-w Cutoff - 135

II

A S-SC Outhard Ewqiwe Cutoff - 152

I

ia s-11 ullaqe Triqqer-153

IA S-fC / s-11 SeParatiON-153

Iote :&par. tiu\esshowu ate iNseconds wa-swed f row I Udisco~r4ect

1 A s-11 mqiNe start-154

I

A Activate P. U. system -159

IA Water CoolaNt Vdlve OWN-160

r----l *2* 3 *of start Of smrt Pf

Time Bose 1 The Base 2 Time BQSP 3

The BaSe DiViSiONS

Rqure 6



A Separa+io? S-II SeCONd PlaNe--L83I

A JettiSON l)?~--,s, i

A Cutoff S-it E&i!qiues--519

IAullaqe ~qwi+!L-~~~

, A S-II/S-IV0 !L.epamtiou-520

[I

A IqNite S-IV B ENqi+m-523

I A~Acti~a+e’P~.SvsteM-528

$ CutQff S-IVB ENqiNe-665

p S-IVB UIlaqeENqiws ON-665

late : I

4pprox. tiMe5 1ihoww are i*l

jecONd5wasuped

1A Activate P.U. System-665I

-3-+-4-L-----5‘start Of Start of

Time Bose 4 TiMe Bose 5

Tiqure 1



I I

A S-IVB 1l.I.~ orJcl Lox ChilldowN i’u~ps .!w - 10,913I I

A S-IVB Restart Preparations-11,235

bs-IVB Ulkqe Ew+les ON; - 11.235

~AS-IVBLH~r4Loxved valvesCIOSS11,235

A~s-;m.y2 a,; Lox chiildoWN puyp of f .

I

I

I

A S-IVB Ullaqo ENqines off -11,565

I

A 24 bp!rtiow S-IVB E~qi~e-11.570

A Acttivo{e I? U. system -11.515

I A S-IVB Tliqht Co&ml

k&e:computer BUP~

Rppro%.tir4esI

Mocle off -11.890

itmWN are iN1ec0tds Mea-iured +ro~TlJiiSCONHect.

---S---M 6 W 7-stall Of Start af

titm? bare 6 tit.%?bo5a 7

Fiqure 6

16



1 Tilt .Arrcsti i

.- 1

!

&unch Data

Launch C.mp:e#.- 39 Pad A

Launch Azimuth-90’5 of N

Flight Arlmuth-7Z0 Eof N

s-E73 24 Burn--

Start- 3hr, llmin., 44s

Cutoff Q 3 hr., 17min .,~7s

hrkinq Orbit,&an Altitude-- 185 Km

iWaitinq Orbit

IApogee-- 4.9666 km.

I, Cutoff Arminp

I.s

, IECO-- Switch S&&r command

Plus sufficient downran9a

velocity.

OECO-- 145 ’.ls; Ou tbd. ‘~“9 . propellan t

depletion cutoff enable.

ECO- - 483.9~ LOX deple tion

sensor-cutoff enable. LHzdcoletion sensor-cutoff



GU1DANCE AND CONTROL SYSTEM (G&C)

Function and De*cription

The G&C sys tem provides these basic functions during flight:(1) stable positioning of the vehicle to the command position with a

minimum amount of sloshing and bending, (2) a first stage ti lt programwhich gives a near zero li ft trajectory through the atmosphere and pro-duces reasonable end conditions at Outboard Engine Cutoff (OECO),(3) reduction of wind toads during the high dynamic pressure region,(4) steering commands during S-ID burn which guide the vehicle to apredetermined set of end conditions while maintaining a minimum propel-lant crajectary, (5) final cutoff signal, (6) attitude signals for thesensing, computing, and actuation elements of the G&C system. A blockdiagram of the hardware used to implement these functions is shown inFigure .

The Stabilized Platform (D-124&%) is a three gimbal configurationwith gas bearing gyros and accelerometers mounted on the stable element.Vehicle accelerations and rotations are sensed relative to this stableelement. Gimbal angles are measured by redundant resolvers and iner-tial veloci ty is obtained from accelerometer head rotation in the formof encoder outputs (also redundant).

The Launch Vehicle Data Adapter (LVDA) is an input-output devicefor the LVDC. These two components are digital devices which aperatein conjunction to carry out the flight program. This program performsthe following functions: (1) P ecesses the inputs from the ST-124M,

(2) performs navigation calculations, (3) provides firs t stage tiltprogram, (4) calc ulate s IGH steering comm ands, (5) resolves gimba l

angles and steering commands into the vehicle sys tem for attitude er-ror comm ands, (6) issu es cutoff and sequencing signals.

The Control/EDS Rate Gyro Package contains 9 gyros (triplex re-dundant in 3 axes). Their outputs go to the Control Signal Processor(CSP) where they are voted and sent to the Control Computer for damp-ing vehicle angular motion.

The Control Computer sends commands to the S-IC, S-II and S-IVBengine actuators and to the Control Relay Packages based on signalsfrom the LMA and rate gyros. These signals are filtered and scaled(see Figure IO), then smmed in magnetic amplifiers. 'I:& computerprovides redundant operation during S-WB burn and coast.

The Control Relay Packages accept Conrral Camp"'-er commands andrelay these comands to operate propellant valves in the AuxiliaryPropulsion System CAPS). All relays and valves are redundant.

The 8 hydraulic actuators of the S-K and S-II stages are used togimbal the outboard engines and provide control in all axes. Th e twohydraulic actuators of the S-IVB stage are used to gimbal the J-2engine and provide control in pitch and yaw axes. The APS is used forcontrol in all axes during S-ZVB coast.



The Switch Selectors are used to relay sequencing commands from theLWA to other locations in the vehicle. The cutoff signal and time

based events are issued through the Switch Se1cctors.

The vehicle is erected on the launch pad with position I at a 90'Eof N azimuth. The Stabilized Platform is aligned to 72' azimuth durFngcountdown and held in an earth fixed positioti, perpendicular to thegravity vector. A roll presetting of 18' is used to eliminate the at-titude error which would result from rhis difference in azimuth. TheLWC operates in ground routines prior to GRR and attitude error signalsare set to null. At the instant of cm, the platform becomes space

fixed to establish the guidance coordinate sysrem and the LWC entei~the flight mode. 1x1 this mode accelermeter processing, steering, navi-gation, telemetry, and other functions are performed exactly as rh6y

are after lif taff .

The Li ftof f signal (IU lrmbilical disconnect) initiates time base 1.Eleven seconds later, the tilt program starts and the initial roll pre-setting is reset to zero. The vehicle rolls into alignment with theplatform at l'/second. There is no active parh guidame during S-IC burn.Control is maintained by ginballing the four outboard engines on commandfrom the Control Computer. Sequencing signals are issned to the variousswitch selectors to perform time dependent functions.

19



DIGITAL COMMAND SYSTEM CAPABILITY:

The fallowing summary describes the AS-501 Digital Command Systems'overall camand capability:

Function

Inhibit

Time base update

Tine base update

Navigation updatt

Generalized switchselector

Sector dump

Telemeter singlememory location

Terminate

Change the time tostart coast phaseattitude maneuver

Time base time isadvanced or re-tarded at the nexftelemetry loss

Navigation quanti-ties are reset atthe time specified

Specified switchselector function isissued at the firs topportunity

Content.3 of speci-fied memory sectorare telemerered

Stop DCS processingand reset for a newcommand

Use S-US to achieveorbit

Inhibits C-Bandtransponder switch-

ing

Periods of Acceptance

FX~ TV + 100 SECO~S umiiT6 and from T, + 10 secondsuntil end of life

From Tg + 100 seconds untilT6 and from T7 + 10 secondsuntil end of life

Pram Tg + 100 seconds untilTg and from T, + 10 secondstill end of life

Frw Ti + 100 seconds untilT6 and from T-/ + 10 secondstill end of life

From Tz, C 115 seconds untilT6 + 317 seconds and fromT7 + 10 seconds till edd oflife

From Tg + 100 seconds untilT6 and from T7 + 10 secondstill end of life

From Tg + 100 seconds untilTg and from TT + 10 secondstill end of life

From Tg + 1.3 seconds untilTb; from TI, + 115 secondstill Tc + 317 seconds andfrom T7 + 10 seconds tillend of life

From Tq + 1.3 seconds till

=I

From Tg + 100 seconds untilTg and from T, + 10 secondsuntil end of life



ST- 124MStabilized -

Platfclrm

Control Actuators7 ,//,,l, Illll,rll,-fIf,,,

sraqc-nTo C.nTrol ActuatorspIIt- Engiwb 1,1,3f4 *

( P,V,b R)



-““““““““““““““““““““““““““z

S-ICC stag*



Propellant consmq?ion during S-IC Stage flight (approximately 152seconds) is approximately 4,320,OOO pounds. Propellant consumptionduring S-II Stage flight (approximately 365 seconds) is.approximately938,700 pounds and during S-IVB Stage flight , including firs t and secondburns, :approximately 459 seSopds$'is approximately 232,700 pounds.

In evenf of one engine of the S-X Stage malfunctions and is cutoffduring flight, the remaining engines will consume the propellant intendedfor the "dead" engine. Burning time of the stage would increase, and theoverall vehicle perfomance loss would be minimized.

VEHICLE WEIGHT DATA (Approximate)

Total af S-IC ignitionTotal at li ftof fTotal at s-IC O.E.C.0Total at S-I* ignition 1;415;700Total at S-II E.C.0 463,200

Total at S-IVB first ignition 352,000

Total at S-IVB first E.C.0 277,000

Total at S-IVB 2nd ignition 273,900

Total at S-IL'S Bnd E.C.0 119,100Total at S-US/SC separation 67,100



4oo,ooa

100.00~

30,ooc

6,220.700 Ibs (Total \lekcle at IqNit;oN)< 6,121.300 lbs (Total VehLcle at Liftoff )@ set)

/S-XC PropellaM Corrsumptioti-28,4oCJ Ib+/sec.

S-lt Staqe Outboard Euqira Cutoff e 152 Sec.)

Separatiohl (-153 Sec.)S-II Staqe 1qNitiot.I (~154 Sec.) (1,415.700 \bs)

S-II PmpellaNt CohlsUMptoN - 2,590 Ibs/sec.

S-II Staqe LqiNe Cutoff (-523Sec.)

Staqe SePartiiON (-520 Sec.)

Space Vehicle Weiqhtvs lliqht TiM e

Fi&-



The 8-X stage is approximately 138 feet long and 33 feet in di-

ameter and has five liquid-fueled Rocketdyne P-l engine s which generate

a total thrust of 7,500,OOO pound s. The engin es are supp lied fuel by abi-propellant system of liquid oxygen (LOX) as the oxidizer and W-1 as

the fuel.

The S-TIC stage structure consists of a thrust structure to which

the engine s attach, an W-1 fuel tank, a LOX tank, an interrank SYNC-

ture separating the LOX and fuel ranks, and a forward skirt structure

which provides an interface surface for the Saturn g-11 stage.

26



RP-I LOX TIP-1

i

I 1I I

1,500.000 I bs. Thrust

fiaure 14



The S-E Stage propellant system is composed o f one LOX tank, one RF-1tank, prope11anc lines, control valves, uents, and pressurization subsystems.

Loading of LOX and W-1 tanks ;s controlled by ground computers. W-1 loadingis compieted at a considerable time prior to start of LOX loading. LOX bubblingbegins and continues throu@, the LOX tank loading to prevent possible LOXgeysering. Approximately 90 seconds prior to ignition command rhe RF-: tankis pressurized from a ground source. Approximately 60 secon ds prior to ignition

command the LOX tank is pressurized from a ground source. Prior to start ofautomatic sequence and up to 72 seconds before li ftof f ground source heliumis bubbled through the LOX lines and tank to prevent stratification in engimLOX suction lines. After l if to ff fhe LCX tank pressurization is maintained byGOX converted from LOX in the heat exchanger. The W-1 tank is pressurized withHe stored in bottles in the LOX tank and heated by passing the He through theheat exchanger.

100

80

L 60

51D 4a

w-1Sy5tet.7

Ijtart.oadiNq RD.1

Low Re-

Pressurize TscuK-

--+I ’I ‘.I I”

IqNitiON

commaNc

reTmuk

30



17 iwdiau\eter -A-

33rierqewy DraNI

Ved &Reiiet Valve f-

Ye& - elief

30.5 psia 35.0 wia

I12 iw. diw

31 cu. +t. 3000 p&9. He Bottle5 (4)for W-1 Tamk Pressuriz

Veut C Relief ValveOprlv -25.5 prig

Close -24.0 psiq

Lox Fill L Drain

-RP-1 ?Gll CDrairLiNe

er etupe)

iN hNeS 01 Liftoff - b‘tS0 Ibs.

Total propellant at liiioif 4,3%OOO

bs. Total pmpel\ard co~suwwd dterl&t& -4,320,2OO IbS.



1 ! /f ‘vehicle ThrustS*ructure.A+tuatorAitachueM Poinrt

Tw’L,~~~~,,fuel dlscharqeduct

I I

Hydraulic Actustor(2 per .eN9iW)

\ Servo vcllve C‘JNtt’OiScktuator ~~OV.SMPH+

Actuators III

CaNted go NOMINAL thrust



14

~---. --Signal SOUtXCs---

fhcrnocoupls

Measuring Rack

2, in S-K stage \

r- Sign *ls (liF:off,c.o., OtCJ F



r-i raruring Lnstrum a~tati~n d

Diatri butor TRtcmctry systems

Battery Battery

No. 1 No.2

640 amp. min. 1250 amp. min.

Main Power Distributor6reund P.w.r

2.9 vdc

In~ttrumrnt.tion

.$ -iolomdry

systems

*E”ginr system(Thrust)

. ED5 system

Nate:

All Components Show”

Arr Located I. The Thrust

Frame Area Except as N&cd.

1

~tl~chsniral Systems

. DDAS sy5tms

. Rang* safety system s

- staging System

- Squcnciny System

36



37



The stage sfrucfure includes: an aft interstage, an aft sk.ir t andthrust structure, a heat shield, a LOX tank, an LH2 tank, and a fc .skirt.

The stage has fi ve J-2 engines which generate a total thrust of1,000,000 pouuds.

Two recoverable film cameras will be flown an the S-II stage of rhefirst two Saturn V launch vehicles. 'These cameras are mounted an thethrust structue as shown on opposite page. The primary objective ofthese cameras is to view second-plane S-X/S-II separation. The secand-ary objectives are to view S-IC/S-II first-plane separation and J-2rngine ignition,

The cameras are turned on by the switch selector shortly beforefirst-plane separation and operate for approximately 40 seconds. Thefilm is 'marked" one-tenth of a second before first-.plane separation,one-tenth of a second after engine start and one-tenth of a secclndafter second-plane separation. The camera capsules are ejected ap-proximately 8 seconds after second-plane separation.

Imediately following ejection, the camera capsule stabilizationflaps are deployed. After the camera capsule descends to an altitudeOf 4,3?0 meters, a paraballoan is inflated, which CRUSTS the stabili-zation flaps TO fall away. Six seconds after the paraballoan is in-flated, a recovery radio tranmit.te+ and flashing light beacon located

an the paraballaon are turned on.

After touchdown, the camera capsule effuses a dye marker to aidvisual sighting of the capsule, and a shark-xepellant co protect thecamera capsule, paraballoon, and the recovery team.

38



-2 Fill # Drain

LH2 F..dl;.c

Ullrg.- Aockot. (8)

Lights (28)

For Cameras

5tag. wcigy-. Dry :- 88,200 Lbs.. fi t s-IC Ignition:-- 1,035,300Lbs.

. A+ S-II Cutof+:- 103,500 Lbs.



The operating cycle of the 5-2 Engine ccmsists of prestart, starf,

steady-state operation and cutoff sequeoces. During prestart, Lox andLIZ2 flow through the engine to temperature-condition the engine components,and LO assure the presence of propellane in the turbopumps for starting.Following a timed cooldown period, the start signal is received by rhesequence controller which energizes various control solenoid valves toopen the propellant valves in the proper sequence. The sequence controlleralso energizes spark plugs in the gas generator and thrmt chamber foignite the propellant. In addition, chr sequence concro11er releasesP,Wz from the start tank. Abe GHz provides the initial drive for the turbo-pumps that deliver propellant fo the gas generator and the engine. 'i&opropellant ignites, gas generator output accelerates the turbopumps, andengine Chrust increases to main stage operation. At this rime, the sparkplugs are de-energized and the engine is in steady-state operation.

Steady-state operation is mintained until a cutoff signal is receivedby the sequence controller. The sequence controller de-energizes thesolenoid valves which in turn close the engine propellant valves in theproper sequence. *s a result, engine Chrust decays and the cutoff sequenceis compleLe.



Lsr Propel!ant utilization valve

I-!-

1J

II eE:*cll-

GHe for LHztank pressurizatiob

4

200.00d Ibs. Thrust

tic&Ire 21



The S-II Stage prope:lant system is composed of integral LOX/LH~ tanks,

propellant lines, control valves, vents, and prepressurizaCion subsystems.Loading of propellant tanks and flow of propellants is controlled by thepropellant utilization systems. The LOX/LH2 tanks are prepressurized byground source gaseous helium. During powered flighr of the S-II Stage,the LOX tank is pressurized by GOX bleed from the LOX heat exchanger.The LH2 tank is pressurized by OH2 bleed from the chrusc chamber hydrogeninjector manifold: pressurization is maintained by the LH2 PressureRegulator.

S-II PKOPELL4NT LOAD AND OPERATIONAL SEQUENCE

1--------- --*mouqal/P~ \ 1000 qal/he4

96% , \ 98% Start

L!!z-jAutowotic

I

I

II I

7.hr. -6.5hr. -S-T&. -4.shr. -T.t.w toLi:toff -1e7 SK. -6

‘-

-ie

A IStart

LoadiNq

LHZ

I

+c.



LH2 Tat&-153.000 Ibs.

dt IqNitiQN

Lam. TaNk-‘?92.6~ lb5.

dt IqHltiQN

DraiN

HeatErchahlqer--.d...,,. F-+-Y I\coverts Lox tGox far Lox TaNk pressurizu&iv~duriNq S-11 powered fliqht .

Total ~ropellcwt dt liftoFf-9%. .M !L-.Total propeik~t cousu~eddfter lif*f -936x00 It6

FiquW 22 I[



The propellant utiiization (PU) system controls loading and enginemixture ratios (LOX to LH2) to ensure balanced consumption of WX and

LH2.

Capacitance probes mounted in the LOX and LH2 containers mm&for t-bemass of the propellants during powered flight. At PU activaeiou (5.5 secondsafter J-2 ignition) the capacitance probes sense the LOX to LA2 imbalance andcommands the engine to bum at the high rate engine mixture ratio of 5.5:1.When the high mixture ratio is removed, the PU system will then command fheengine to burn the reference mixture ratio of 4.7:1, striving for siw.ultaoeousdepletion of LOX and LHz for maximum stage performance. Engine cutof f isinitiated when any two of the five capacitance probes in either Caskindicate dry.



ITelemetered

Xrror Signal

to GroundLoadingcomputer

4 Telemetry

Tiqure 23

4

Telemetetsd V4lvrCommand Siqnal

CONtlOlalveTeietieteredValve P&t iou SiqWal

TO other4 enqiqes

F’

E



Total6 +

t+ All ES E Display F; I;

Microphone _(acoustic 1

,’4

79 *To Blockhouse thru pwflighf

data acqu isitioe System

rI



DistribUtDr

-- -------- ---_--__ .&& 5kirt- - --_

Thrust Framr

A,%&

Proprllmt Monitoring

FMahagrmenf Systems

POW.%VRange Safety

Recirculation system* IBatterv No. 1 4

Figure t27--



Intentionally Left Blank

57



--++t

lkTop View Looking Aft

AT&mr+qntcnnrs (4)

Forwwdkirt

S-D?3 sta3* Weight *

.Dry:-~6,500 Lb..

*At Ignition:~Zbz,ZoO Lbs.

‘Al- I& Cutoff:-- lE7,3OOLbs.

*At 2Ud&Cutoff:--29,300Lbo.

SEx*cludrs Wt. of aft..

Auxiliary Tunnel

comm*n BUlKhDId

interstage



Thz operating cycle of the J-2 Engine consists of prestart, start,steady-state operation and cutoff sequences. During prestar?, LOX and

LH2 flow through the engine to temperature-canditian the engine compon ents,

and to assure the presence of propellant in the turbopumps for starting.

Following a timed cooldo-m period, the start signa l is received by the

sequen ce controller which energizes various control soleno id valves to

open the propellant valves in the proper sequen ce. The sequen ce controller

also energizes spark plugs in the gas generator and thrust chamber to

ignite ctle propellant. In addition, the sequer.ce controller releases

ai* from the start tank. The GE2 provides the initia l drive for the turbo-

pumps that deliver propellant fo rhe gas generator and the engine. The

propellant ignites, gas generator output accelera tes the turbopumps, and

engine thrust increase s to main stage operation. At this time, the spark

plugs are de-energized and the engine is in steady-state operation.

Steady-state operation is maintained until a cutoff signa l is received

by the sequen ce controller. The sequen ce controller de-energizes the

soleno id valves which in turn close the engine propellant valves in the

proper sequen ce. As a result, en.ci.ine thrust decays and the cutoff sequen ce

is complete.

I

.

COMMAND

N TlME FROM IGLIITIOH



GCJS forLOX ton*pressurization

A

GHz for Ltiztank p~essUrizC,tiOC,

i !

1200,000 Ibs.Thrust

fiqure 29

II”.’

--



The s-IW stage propellant system is composed of integral LOX/L~ tanks,propellant lines, control valves, YeDts, and pressurization subsystems.Loading of the propellant tanks and flaw of propellants is controlled by theprope1Lanr utilization system. Both propellant tanks are initially pressurizedby ground solnrce cold helium. LOX tank pressurization during S-WE *tagebum is maintained by helium supplied from spheres in the I,Ez tank, whLch isexpanded by passing through the helium heater, to maintain positive pressureacross the common Lank bulkhead and to satisfy engine net positive suctionhead. The LHz pressurization strengthens the stage in addition to satisfyingnet positive suction head requirements. After engine ignition the presst,reis maintained by GH2 tapped from the engine supply

L------- ----

T-ltart

AUtOttlstiC

Sequence

t

liftoff

A

Ignitiontom owlI 9

III

15 ht. -7.0 hr. -4.i hr. -319 hr.

I -f;me to liitotf



; r-'

I/ i

I ‘MlI 1 r1 [I-

’ I//I ,clL-I-_

I I’ ILI-

,r;Kj

i

l

LHITciHk-41.200 lb5 al

lqrri+ioN

i3 5 cu.+t. 31W sia rjHs

~sbhores(8). IN. Iiqht Losrqqrrk preswruut&w ,

/f--Los Fill 4 Drsiu

Hpot ExchaNqef u

Total propeitnb&t at i; ftof f-234.100 4trs.TO-M propellaN+ com%medafter liftoFf-232.7m lbs.

Technical Information Summary AS-501 Apollo Saturn V Flight Vehicle

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Transcript of Technical Information Summary AS-501 Apollo Saturn V Flight Vehicle