PROJECT TECHNICAL REPORT

APOLLO 10

LM-4

DESCENT PROPULSION SYSTEM

FINAL FLIGHT EVALUATION

NAS 9-8166

Prepared for

MANNED SPACECRAFT CENTER HOUSTON, TEXAS

NATIONAL AERONAUTICS AND SPACE ADMINI

Prepared by R. K. M. Set0

Propulsion Technology Section Power Systems Department

https://ntrs.nasa.gov/search.jsp?R=19690026326 2020-07-04T03:04:08+00:00Z

11 176-H314-R0-00 ~ ~ ~~~

PROJECT TECHNICAL REPORT

APOLLO 10

LM-4

DESCENT PROPULSION SYSTEM

FINAL FLIGHT EVALUATION

NAS 9-8166

Prepared f o r NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

MANNED SPACECRAFT CENTER HOUSTON, TEXAS

Prepared by R. K. M. Set0

Propul s Power

on Technology Sect on Systems Department

Approved by:

Approved by:

Approved by :

8 AUGUST 1969

Propulbi'on Technology Section

Pow& Systems Department

. 3 . Smith . Manaoer " Task E-19E

CONTENTS

1 . PURPOSE AND SCOPE . . . . . . . . . . . . 2. SWMARY . . . . . . . . . . . . 3. INTRODUCTION . . . . . . . . . . . . 4. PERFORMANCE ANALYSIS . . . . . . . . . . . .

Descent O r b i t I n s e r t i o n Burn . . . . . . . . P h a s i n g Burn . . . . . . . . . . . .

5. PRESSURIZATION SYSTEM EVALUATION . . . . . , . 6. PQGS EVALUATION AND PROPELLANT LOADING . . . 7. ENGINE TRANSIENT ANALYSIS . . . . . . . . . .

S t a r t and Shutdown T r a n s i e n t s . . . . . . . . T h r o t t l e Response . . . . . . . . . , . .

8. REFERENCES . . . . . . . . . . . . 0

ii

PAGE 1

2

4

6

7 8

9

11

14

14 16

17

TABLES

PAGE 1 . DPS M I S S I O N DUTY CYCLE . . . . . . . . . . 18

2. LM-4 DPS ENGINE AND FEED SYSTEM PHYSICAL CHARACTERISTICS 79

. . . . . . 3. DESCENT PROPULSION SYSTEM FLIGHT DATA 20

4. DESCENT PROPULSION SYSTEM STEADY STATE PERFORMANCE - . . . . . . . . . . PHASING BURN 2 1

5. DPS PROPELLANT QUANTITY GAGING SYSTEM PERFORMANCE - . . . . . . . . . . END OF PHASING BURN 22

. . . . . . 6. DPS START AND SHUTDOWN IMPULSE SUMMARY 23

ILLUSTRATIONS

FIGURE NO. PAGE

1 MEASURED SUPERCRITICAL HELIUM TANK PRESSURE - PHASING BURN . . . . . . . . . . . .

2 MEASURED HELIUM REGULATOR OUTLET PRESSURE (GQ3018P) PHASING BURN . . . . . . . . . . . .

3 MEASURED OXIDIZER INTERFACE PRESSURE - PHASING BURN

4 MEASURED FUEL INTERFACE PRESSURE - PHASING BURN

5 MEASURED CHAMBER PRESSURE - PHASING BURN . . . 6 MEASURED PROPELLANT QUANTITY, OXIDIZER TANK NO. 1

PHASING BURN . . . . . . . . . . . . 7 MEASURED PROPELLANT QUANTITY, OXIDIZER TANK NO. 2

PHASING BURN . . . . . . . . . . . . 8 MEASURED PROPELLANT QUANTITY, FUEL TANK NO. 1 -

PHASING BURN . . . . . . . . . . . . 9 MEASURED PROPELLANT QUANTITY, FUEL TANK NO. 2 -

PHASING BURN . . . . . . . . . . . .

24

25

26

27

28

29

30

31

32

i V

1. PURPOSE AND SCOPE

The purpose of this report i s t o present the resul ts o f the postfl ight

analysis of the Descent Propulsion System (DPS) performance dur ing the

Apollo 10 Mission. T h i s report docunents additional analysis of the DPS.

Preliminary findings were reported in Reference 1.

together information from other reports and memorandums analyzing specific

This report also brings

anomalies and performance i n order t o present a comprehensive description

of the DPS operation during Apollo 10.

The following items are the major additions t o the results as reported

i n Reference 1 :

1) The performance for the second (Phasing) DPS burn i s dis- cussed i n greater detail .

2 ) The Pressurization System performance is revised.

3) The Propellant Quantity Gaging System performance i s dis- cussed i n greater detail .

4) The transient performance analysis for DPS operation i s expanded.

2. SWIMARY e The performance of the LM-4 Descent Propulsion System dur ing the

Apollo 10 Mission was evaluated and found t o be satisfactory.

Because system data dur ing the DO1 maneuver was not recorded and due

t o the short length of b u r n s , no detailed performance study using the

Apol l o Propulsion Analysis Program was possible. However, the preflight

model was used w i t h f l igh t data t o approximate the performance a t repre-

sentative times d u r i n g the Phas ing Burn .

(13.1% of fu l l thrust) thrust, specific impulse, and mixture ratio were

calculated t o be 1371 lbf , 296.5 seconds and 1.605, respectively. For FTP,

the values were 9841 lbf , 304.0 seconds and 1.599.

considered as representative only.

For m i n i m u m th ro t t le operation,

These values can be

Instrumentation biases were determined on the regulator outlet pres-

sure measurement (GQ3018P) , the oxidizer interface pressure measurement

and the chamber pressure measurement w i t h values of t4.0, t7.5 and -0.8 t o

-1.6 psi , respectively.

e The supercri t i cal he1 i un t ank experienced an average pressure r ise

ra te of 5.84 psi/hr d u r i n g the coast period between launch and f i rs t DPS

engine f i r ing. This value was less t h a n anticipated from ground tests.

A l t h o u g h the fuel quantity gages (Fu 1 and Fu 2 ) never read off scale

(greater than the maximum 95 percent indication) as expected prior t o the

Phasing B u r n , they d i d respond with propellant consunption and were w i t h i n

b o t h the expected accuracy of 3.5% and the specification limits of 1.3%

a t the end of the burn . The oxidizer gages (Ox 1 and Ox 2 ) operated as

expected prior t o the phasing burn.

appeared that the Ox 2 gage was reading 1.6% higher than expected.

reading was s t i l l w i t h i n the expected accuracy of 2.7%.

However, a t the end of the burn, i t

T h i s

a 2

The engine s t a r t and shutdown transtents compared very well w i t h

predicted values. The shutdown transient time, however, was 0.09 seconds

greater than the specification limit o f 0.25 seconds.

from 13.1% t o FTP was acceptable.

The throt t le response

3

3. INTRODUCTION

The Apollo 10 Mission was the tenth i n a ser ies of f l ights using speci-

fication Apollo hardware.

manned f l igh t of the Lunar Module (LM).

f l i g h t o f Block I1 Command and Service Module (CSM) and the t h i r d manned

f l i g h t us ing a Saturn V launch vehicle.

I t was the t h i r d f l i g h t t e s t and the second

The mission was the fourth manned

1 The overall mission objective was to duplicate, as closely as possible,

a G type mission w i t h the exception of lunar landing and l i f toff . This

included the performance of the Descent O r b i t Insertion maneuver by the

Descent Propulsion System (DPS) and the rendezvous maneuvers by the Ascent

Propulsion System (APS). Also included as objectives were t o verify LM

operation i n a lunar environment, verify mission support of a l l spacecrafts

d u r i n g a l l mission phases a t lunar distances and t o obtain more information

about the lunar potential.

The space vehicle was launched from the Kennedy Space Center (KSC) a t

12:49:00 P.M. (EST) on May 18, 1969. Following a normal launch phase, the

S-IVB tage inserted the spacecraft into an orb i t of 102.6 by 99.6 nautical

miles. Two and a half hours af ter launch the S-IVB performed the trans-

lunar njection maneuver.

c ra f t s were ejected from the S-IVB approximately four hours a f t e r launch.

Dur ing the next 76 hours, four SPS burns were performed.

LM from the CSM i n l u n a r orbi t occurred 98.5 hours a f t e r launch. A t approxi-

mately 100 hours, the f i rs t DPS maneuver, the Descent O r b i t Insertion (DOI)

burn was performed. The burn duration was 27.4 seconds and included opera-

t i o n a t the m i n i m u n th ro t t le setting and throt t l ing to the 40% o f fu l l thrust

The CSM docked w i t h the LM and the docked space-

Undocking of the

'Reference 2.

4

level. This burn p u t the LM into a lunar orbi t of 61.2 by 8.4 nautical

miles. A t approximately 101 hours a f t e r launch, the DPS performed a Phasing

Maneuver burn , 39.9 seconds i n duration. The spacecraft was now i n a lunar

o r b i t of 190.1 by 11.0 nautical miles. The burn included operation a t the

minimum thro t t le sett ing and a short duration segment a t the Fixed Throttle

Posit ion (FTP). The Phasing Maneuver ended the DPS mission duty cycle. The

descent stage was separated from the ascent stage about two hours la te r .

The APS performed two f i r ings, the l a t t e r being t o propellant depletion and

the SPS performed one more burn during the subsequent portion of the

mission.

The actual ignition and shutdown times for the two DPS fir ings are

shown in Table 1.

The Apollo 10 Mission utilized LM-4 which was equipped with DPS engine

S/N 1039.

Table 2.

The engine and feed system characterist ics are presented in

Each DPS burn was prededed with a two j e t t X LM Reaction Control

System (RCS) ullage maneuver to se t t l e propellants.

There was one Apollo 10 Mission Detailed Test Objective (DTO) specifi-

cally related t o the DPS.

P13.14 LM Supercritical Helium.

The functional t e s t objective o f this DTO was:

1) Obta in data on DPS supercritical helium pressure profile dur ing standby and during DPS DO1 and phasing burns.

The detailed requirements of this objective are described i n Reference 3.

5

4. PERFORMANCE ANALYSIS

Due t o the insufficient duration of the two DPS maneuvers performed

dur ing the Apollo 10 Mission, a meaningful detailed analysis using the

Apollo Performance Analysis Program could not be made. Analysis was

further hampered by the loss o f the DO1 burn da ta . The burn was performed

behind the moon and the CSM failed t o record the LM data.

Upon activating the ambient helium start bottle i n preparation for

the DO1 burn , DPS pressures appeared nominal w i t h the exception of the

oxid izer interface pressure measurement (GQ 4111P) and the redundant

helium regulator outlet pressure measurement (GQ 3018P).

the regulator outlet pressure (GQ 3025P) and fuel interface pressure ' (GQ 3611P) were approximately equal a t 251 and 250 psia, respectively.

oxidizer interface pressure was 241 psia while the redundant regulator

out le t pressure (GQ 3018P) was 247 psia.

had a bias of from 0.8 t o 1.6 p s i a p r i o r t o the burn.

FTP time s l i ce during the phasing burn indicated t h a t the oxidizer pressure

transducer mus t have incurred a downward sh i f t . Had the interface pressure

A t this time,

The

The chamber pressure measurement

Simulation of an e

been as measured, the mean chamber pressure (with bias included) would have

been more than one psi lower than observed. I t was concluded tha t a t FTP,

the interface pressures were essentially equal and tha t there was a bias of

approximately 7.5 psi on the oxidizer interface transducer. Similar

reasoning,and the fact t h a t the regulator outlet pressure as measured by

GQ 3025P matched the predicted value during the burn, indicated t h a t the

regulator outlet pressure measurement (GQ 3018P) was biased low by approxi-

mately 4 psi. Table 3 presents the f l igh t measurements for the Descent

Propulsion Sys tem.

6

Descent Orbit Insertion Burn

Al though the da ta for the DO1 Burn was not recorded, indications are

that the DPS performed satisfactorily. Prior t o and a f te r the maneuver,

the system pressures appeared nominal. Astronaut reports of the burn i n d i -

cated noma1 operation. The burn was ini t ia ted a t the min imun throt t le

setting of 13.1% of ful l thrust. After approximately 14 seconds, the

engine was to be manually t h r o t t l e d t o the 40% level for the remainder o f

the maneuver. The length of the burn was reported t o have been approxi-

mately 27.4 seconds w i t h a measured velocity change of 70.66 ft/sec.

actual velocity ga in target was 71.25 ft/sec.

The

The p r e f l i g h t performance

predicted burn time was 28.0 seconds w i t h a simulated velocity change of

71.6 ft/sec. There are three primary reasons for the difference between

predicted and actual burn times: 1) differences i n velocity gain, 2 )

simulated m i n i m u n th ro t t le setting, and 3) simulation of the t h r o t t l i n g

transient from 13.1% to 40%. The p r e f l i g h t assuned minimum thro t t le posi-

t i on was 11.3% while the actual infl ight sett ing was 13.1%. In simulating

the bum, a step change between thrott le settings was assumed while the

actual maneuver requires approximately one second.

are accounted for, i t appears that the predicted and the actual burn time

would d i f fe r by less than 0.1 seconds. Other uncertainties about the b u r n

include actual s t a r t transient, time of t h r o t t l i n g t o 40%, actual throttle

position a f t e r t h r o t t l i n g (since the maneuver was performed manually) and

spacecraft weight errors.

performance was nominal .

a

I f these differences

In view of the above, i t was concluded that the

The attainment of the target velocity gain i s extremely c r i t i ca l t o the

descent trajectory. The smal 1 residual (difference between target and

actual) of 0.6 ft /sec was easily nulled by use of the LM-RCS. a 7

Phasing Burn

The Phasing Burn was performed sat isfactor i ly . The burn was i n i t i a t e d

a t the min imun th ro t t le sett ing. After 26 seconds the engine was automati-

cally thrott led t o the Fixed Throttle Position (FTP) for the remainder o f

the maneuver.

The actual burn time was 39.94 seconds w i t h velocity ga in of 175.8 ft/sec,

while the predicted burn time was 40.3 seconds for a velocity g a i n of 174.5

ft /sec. The actual target velocity ga in was 176.9 ft /sec. As w i t h the

DO1 Burn, the difference in predicted and actual velocity ga in and time

can be essentially accounted for by the difference i n the simulated and

inf l ight thrott l ing transients, minimun th ro t t le sett ing and start transient.

Table 4 presents the inf l ight measured da ta a t typical points dur ing each

System pressures appeared nominal d u r i n g and a f t e r the burn.

of the two throt t le positions experienced i n the Phasing Burn. The pre-

f l i gh t predicted values, obtained from Reference 5 , are also presented for

comparison.

dicted da ta .

are due t o the difference between f l igh t and predicted thro t t le sett ing.

Although detailed performance analysis could no t be made, the f l igh t data

was used i n the prediction model to give an indication of approximate in-

f l i gh t performance. Figure 1

through 9 present DPS inf l ight measured supercri t i ca l heliun t a n k pressure,

regulator out le t pressure, interface pressures , chamber pressure and gaging

system readings during the Phasing B u r n .

The inf l ight measured da ta compares well w i t h preflight pre-

Deviations a t the m i n i m u n th ro t t le sett ing (FS-1 -I. 10 seconds)

The results are also presented i n Table 4.

8

5. PRESSURIZATION SYSTEM EVALUATION

The performance of the pressurization system was considered satisfactory.

The ambient s t a r t bot t le was loaded w i t h approximately 1.1 lbm of

helium a t a pressure of 1619 p s i a a t approximately 72.5'F. A t launch, the

pressure was approximately 1612 psia. Five days prior to launch, the oxi -

dizer and fuel tank pressures were increased from the i r load pressures t o

186.2 and 193.3 psia, respectively. A t launch, the pvopellant tank pressures

had decreased t o approximately 168 and 188 psia, respectively. Approximately

30 hours prior t o launch, the supercritical helium (SHe) tank f i l l proced-

ures were completed w i t h approximately 48 lbm of helium loaded a t a pressure

of about 95 psia. A t launch, the pressure had risen to approximately 316

psia. The SHe tank pressure increase dur ing this period vas approximately

7.65 psi/hr due to normal heat leak into the system from the surrounding

environment.

sure rise rate was 7.31 ps i /h r .

During the 119 hour countdown demonstration t e s t , the pres-

A t 97.5 hours a f t e r launch, prior to pre-burn propellant tanks pressuri-

zation, the ambient helium bottle pressure was 1577 psia, the SHe t a n k

pressure was 885 psia, the oxidizer tank pressure was 97 psia and the fuel

tank pressure was 152 psia.

was attr ibuted t o heliun going into solution (Reference 6).

the ambient s t a r t bott le pressure was greater t h a n expected when only

temperature effects are considered.

s t a r t bott le pressure showed l i t t l e decay dur ing the four days prior t o

launch.

bo t t le i s located, pr ior t o the f i r s t DPS burn, were similar t o LM-3 (which

showed l i t t l e s t a r t bott le pressure decay from launch t o burn ) .

The pressure decay i n the propellant tanks

The decay i n

In the case of Apollo 9/LM-3, the

Indications were t h a t the temperature i n the bay where the s t a r t

I t i s ,

9

therefore, possible that there was a small heliun leak which could have

been caused by launch vibrations. An accurate analysis could n o t be made

due t o pressure measurement inaccuracies and the lack of system temperature

measurement.

increased t o 248.5'and 249 psia in the oxidizer and fuel tank, respectively.

Thus, although there may have been a heliun leak, the ambient s t a r t bottle

performed as expected and caused no anomalies i n propellant pressurization.

The average SHe t a n k pressure rise, from launch was approximately 5.84 psi/

hr. T h i s f l igh t pressure rise rate was somewhat less t h a n anticipated

based on ground tes ts . Similar reductions of infl ight pressure r i se rate

was experienced on LM-3. Because o f a known helium leak observed i n the

SHe system a f t e r the f i r s t DPS burn, however, i t was not clear whether the

reduced r i se rate was due t o zero-g coast conditions o r the existence of

the leak prior t o the f i r s t burn. Based on the similar pressure r i se rate

experienced d u r i n g the Apollo 10 Mission, it appears t h a t the LM-3 pressure

r ise was normal and that the leak occurred a f t e r system ac t iva t ion p r io r

t o the f i r s t burn.

t o be used for system predictions i s being revised.

Upon activation o f the ambient s t a r t bott le, the pressures

e

In view o f the above, the f l i g h t pressure r i se rate t o

From the available f l igh t da ta , i t appears t h a t the SHe system operated

normally during both DPS burns.

'Includes apparent 7.5 psi bias as discussed i n Performance Analysis section.

10

6. PQGS EVALUATION AND PROPELLANT LOADING

Propellant Quantity Gaging System

A t engine ignition for the second DPS b u r n , the oxidizer propellant

gages (OX 1 and Ox 2) were reading off scale, as expected (greater than

the maximum 95 percent indication). The fuel tank probes (Fu 1 and Fu 2)

had readings of 94.2 and 94.5 percent, respectively.

timate of consumed propellant dur ing the DO1 maneuver, the fuel tank meas-

urements should also have been reading off scale a t i g n i t i o n . T h i s devia-

t i on was a l so noted prior to launch. After ignition, the fuel quantities

remained relatively constant for approximately 31 and 27 seconds for Fu 7

and Fu 2, respectively, a t which time propellant consumption was indicated.

The oxidizer gages began to show consunption a t approximately 35 and 37

seconds fo r Ox 1 and Ox 2, respectively.

pellant gages were reading 92.4, 92.0, 93.8 and 94.5 percent for Fu 1 , F u 2,

Ox 1 and Ox 2, respectively. Table 5 presents a comparison of the measured

data and the best estimate of the actual values a t the end of the Phasing

B u r n . Although the Ox 2 gage is outside the specification limits of 1.3%,

i t should be noted t h a t the lack of data from the DO1 burn somewhat compro-

mises the calculated values. A1 t h o u g h i n i t i a l l y g i v i n g erroneous output,

the fuel gages appeared to be functioning w i t h i n specification limits a t

engine shutdown.

and 3.5% for oxidizer and fuel (Reference 7 ) . These accuracies were de-

veloped from recent tes t s conducted a t the White Sands Test Facil i ty (WSTF).

Based on the best es-

A t the end of the burn , the pro-

a

A l l values were w i t h i n the expected accuracties o f 2.7%

The fai lure of the fuel gages t o reach a maximum reading when greater

than t h a t amount of propellant was i n the tanks has been attributed to

either chemical reaction w i t h alodine or aluminum impurities w i t h the fuel , or contamination o f the fuel sensors due t o the referee propellant (used a

11

instead of l ive propellants i n probe manufacture and calibration) o r alodine

surface treatment (Reference 8). A chemical reaction between the fuel and

impurities, which are not clearly understood, could cause i n insulating

barrier t o be set up such t h a t the conductance w i t h i n the sensing portion of

the gaging system probe i s reduced, t h u s causing a reduction i n the fu l l

scale reading. This barrier could be in the form of bubbles forming on the

inner electrode when the sensor i s submersed in stagnant fuel. A small

quantity of residual from the referee propellant or from the alodine surface

treatment of the gage (prior t o installation) could combine with the pro-

pellant and form a conductive component in the fuel that s e t t l e s i n the

reference region a t the bottom o f the gaging probe causing the signal t o

be low a t gage activation.

would have t o be in a stagnant condition.

ence 8 t h a t under zero gravity conditions, these problems should n o t occur,

particularly due t o RCS and SPS act ivi t ies which would tend t o keep the

propellant reasonably active inside the tanks .

i t i s possible t h a t the propellant movement prior t o engine burn was not

great enough t o remove contamination from the reference region.

For either of these t o happen, the pmpellant

I t was t h u s concluded i n Refer-

In the case of this f l igh t ,

1 Propellant Loading

Pr io r t o propellant loading a density determination was made for the 3 oxidizer and fuel.

and a fuel density of 56.44 lbm/ft a t a pressure of 240 ps ia and a temper-

The analysis yfelded an oxidizer density of 90.22 lbm/ft 3

'Reference 9

12

ature o f 70' F. The oxidizer and fuel were loaded t o their planned overfill

quantit ies of 11400.4 lbm and 7136.7 lbm,yespectively. Off-loading was

planned such tha t the target loads o f 11209.4 lbm of oxidizer and 7054.8 lbm

of fuel would be obtained.

fuel was off-loaded than planned. The actual propellant loads a t launch

were 11209.2 lbm of oxidizer and 7009.5 lbm of fuel.

D u r i n g this procedure, however, 45.3 lbm more

13

7. ENGINE TRANSIENT ANALYSIS

The mission duty cycle of the DPS dur ing Apollo 10 included two s t a r t s

a t the minimum thro t t le sett ing, one shutdown a t approximately 40% th ro t t le

and one shutdown a t FTP. During the DO1 Burn the engine was manually

thrott led t o 40% th ro t t le and during the Phas ing Burn the engine was auto-

matically thrott led t o FTP.

Due t o data loss dur ing the DO1 Burn, only the transients for the

Phasing Burn were analyzed.

satisfactory since they compared well w i t h predicted values.

noted, however, t ha t the shutdown transient time was greater than the

specification limit by approximately 0.09 seconds.

The transients for this burn were considered

I t should be

Phasing Burn Star t and Shutdown Transients

In determining the time of engine f i re switch signals (FS-1 and FS-2),

the technique as developed i n Reference 10 was used. T h i s method, devel-

oped from White Sands Test Facility (WSTF) t e s t data, assumes tha t approxi-

mately 0.030 seconds a f t e r the engine start command (FS-1), an oscil lation

i n the fuel interface pressure occurs. Similarly, 0.092 seconds a f te r the

engine shutdown signal (FS-2) another oscil lation i n the fuel interface

pressure occurs. Thus, s t a r t and shutdown oscil lations of the fuel inter-

face pressure were noted and the appropriate time lead applied.

The ignition delay from FS-1 t o f irst r i s e i n chamber pressure was

approximately 0.85 seconds. I t has been shown from past f l ights that the

f i r s t start of a duty cycle i s generally longer t h a n subsequent starts by

a factor of approximately two. This difference appears t o be because of a

difference i n engine priming conditions, since prior t o the f irst s t a r t , 0

14

certhin engine ducts a re dry. Since this was the second s t a r t of the duty

cycle, the delay time appeared reasonable and compared favorably w i t h similar

s t a r t s experienced dur ing Apollo 5 and Apollo 9 f l igh ts .

0

The start transient from FS-1 t o 90% o f the steady-state t h ro t t l e

sett ing (13.1% of fu l l thrust) required 2.13 seconds w i t h a s t a r t impulse

of 728 lbf-sec. The transient time was well within the specification limit

of 4.0 seconds for a minimum throt t le start . The measured impulse compared

favorably w i t h the predicted (Reference 5) nominal value of 862 lbf-seconds

(a1 t h o u g h the nominal predicted time was approximately one second greater

t h a n measured) as well as similar s t a r t s performed dur ing Apollo 5. The

measured value was somewhat low when compared w i t h DPS s t a r t s on Apollo 9 .

One possible reason this deviation may be the coast time between burns .

Although there i s insufficient f l ight data t o fu l ly correlate the effects,

i t appears tha t the magnitude of the s t a r t impulse may be proportional to

the coast time between burns. This i s due t o residual propellants

freezing i n the injector a t engine shutdown before they can reach the

combustion chamber.

propellants t o sublime away. The resu l t can be par t ia l ly primed injector

a t engine res ta r t . The coast time between the burns performed on Apollo 10

was approximately 72 minutes which i s l ess than a l l coast periods w i t h the

exception of the coast between DPS 2 and DPS 3 on Apollo 5 (0.5 min) . The

magnitude of the s t a r t impulse for the Phasing Burn f a l l s between t h a t of the

Apollo 5 DPS 3 s t a r t and the other s t a r t s f rm Apollo 5 and Apollo 9.

0

An appreciable amount of time i s required for these

The shutdown transient required 0.34 seconds from FS-2 t o 10% of the

steady-state throttle sett ing (FTP) w i t h an impulse of 2041 lbf-sec.

the time and impulse for the transient are greater than observed d u r i n g

Apollo 5, where similar shutdowns were conducted, b u t compares favorably

Both

a 15

with the nominal predicted values o f 0.32 seconds and 2017 lbf-sec. The

transient time was, however, greater than the specification limit of 0.25

seconds for shutdowns performed from TTP. There i s no specification l imit

on impulse. The impulse from FS-2 t o zero thrust as determined by consi-

deration of spacecraft weight and vehicle velocity ga in was 2948 lbf-sec.

This agrees well with the predicted value of 3089 lbf-secs b u t i s somewhat

greater than the impulse experienced on Apollo 5 shutdowns. Table 6

presents a summary of the transients.

0

Th rot t 1 e Response

During the Phasing B u r n , the engine was automatically throttled from

the rninimun throt t le position t o FTP.

engine actuator, t o f ive psi less than steady-state chamber pressure a t FTP

was 0.94 seconds. This was within the specification limit of 1.0 seconds.

This value is 0.6 seconds greater than a similar throt t le change performed

dur ing Apollo 5 b u t was similar t o l ike thrott l ing performed d u r i n g Apollo

9 (40% t o FTP in 0.82 seconds).

The time from first movement of the

0

16

8. REFERENCES

1. TRW IOC 69.4354.2-54, "DPS Input t o Apollo 10 Mission Report," from R. L. Barrows to D. W . Vernon, dated 16 June 1969.

2. NASA/MSC Report MSC -00126, " A p o l l o 10 Mission Report," dated Aug., 1969.

3. SPD9-R-037, "Mission Requirements, SA-505/CSM-106/LM-4, F Type Mission, Lunar Orbit," dated 31 January 1969.

4. TRW Letter 69.7254.3-61, "Support to the Postflight Analysis of the GN&C Systems on Apollo 10," from D. L . Rue t o J . F. Hanaway, dated 1 July 1969.

5. TRW IOC 69.4354.2-12, "Apollo Mission F/LM-4/DPS Preflight Performance Report," from R. K. M . Seto, D. F. ROSOW, S. C. Wood and 3. 0. Ware t o D. W. Vernon, dated 21 February 1969.

6. TRW IOC 69.4354.1-70, "LM-4 DPS Fuel Ullage Pressure Prior to First Burn," from R. L. Barrows t o P. H. Janak, 6 June 1969.

7. GAEC Report LED-271-98, "PQGS Accuracy Study," dated 14 May 1969.

8. Trans Sonic, Inc. Report No. 4-04003C, "Final Report Analysis of Stagnation (Time Lag) Anomaly i n Propellant Quantity Gaging Section," dated April , 1969.

9. SNA-8-D-027 (111), Rev. 1 , "CSM/LM Spacecraft Operational Data Book," Vol. 111, Mass Properties, dated November 1968.

10. MSC Memorandum EP22-41-69, "Transient Analysis of Apollo 9 LMDE," from EP2/Systems Analysis Section t o EP2/Chief, Primary Propulsion Branch, dated 5 May 1969.

17

h

q. W

w -I u * u

w

* m CD N I- r o o 3

0- h h -I P

zt d d W II

z 0 U m m H r: v, CL n

w a u Q) E i=

18

TABLE 2

LM-4 DPS ENGINE AND FEED SYSTEM

PHYSICAL CHARACTERISTICS

ENGINE

Engine Number

Chamber Throat Area, In2

Nozzle E x i t Area, In2

Nozzle Expansion Rat io

1039

53. 7401

2569. 74

47. 64

Oxidizer I n t e r f a c e To Chamber

Resis tance a t FTP lbm-sec2 lbf-ft!’ 3904. 63

Fuel I n t e r f a c e To Chamber

Resis tance A t FTP lbm-sec2 lb f - f t ’ 6207.9

FEED SYSTEM

Oxidizer P r o p e l l a n t Tanks, T o t a l

Ambient Volume, F t 3 126.0

Fuel P r o p e l l a n t Tanks, T o t a l

Ambient Volume, F t 3 126. O 4

Oxidizer Tank To I n t e r f a c e

lbm-s e c Res i s t ance , l b f - f t 5

Fuel Tank To I n t e r f a c e

lbm- s ec l b f - f t 5 Resis tance,

496. 112

757. 682

TRW No. 01827-6125-T000, TRW LM Descent Engine Serial No. 1039 Acceptance Test Performance Report Paragraph 6.9, 8 December 1967.

GAEC Cold Flow Tests

TRW No. 4721.3.68-188, LM-4, Engine S e r i a l No. 1039 Descent Engine C h a r a c t e r i s t i c Equations, July 1968.

Approximate Values

19

MEASUREMENT NUMBER

GQ3435P

GQ3015P

GQ 30 18P

GQ3025P

GQ3611P

GQ4111P

GQ6510P

GQ3603Q

GQ3604Q

GQ4104Q

GQ4455X

GQ3 7 18T

GQ3719T

GQ4218T

GQ4219T

GQ6806H

GH1311V

G H 1 3 31V

GGOOOlX

TABLE 3

DESCENT PROPIRSION SYSTEM FLIGHT DATA

DES CRIPTION RANGE

Pressure , S u p e r c r i t i c a l H e l i u m Tank 0-2000 p s i a

Pressure , Ambient H e l i u m B o t t l e 0-1750 p s i a

P res su re , H e l i u m Regulator Out le t Mani- f o l d 0-300 p s i a

P res su re , H e l i u m U g u l a t o r Ou t l e t Mani- f o l d

0-300 p s i a

P res su re , Engine Fue l I n t e r f a c e 0-300 p s i a

P res su re , Engine Oxid izer I n t e r f a c e 0-300 p s i a

P res su re , Engine Thrust Chamber 0-200 p s i a

Quant i ty , Fue l Tank No. 1 0-95 pe rcen t

Quant i ty , Fue l Tank No. 2 0-95 pe rcen t

Quant i ty , Oxidizer Tank No. 1 0-95 pe rcen t

Quant i ty , Oxid izer Tank No. 2 0-95 pe rcen t

Low Po in t Sensor, P r o p e l l a n t Tanks Liquid Level O f f-On

Temperature, Fuel Bulk Tank No. 1 2 0- 12 0 F

Temperature, Fue l Bulk Tank No. 2 20-120°F

Temperature, OxidiBer Bulk Tank No. 1 20-120°F

Temperature, Oxidizer Bulk Tank No. 2 20-120'F

P o s i t i o n , Var i ab le I n j e c t o r Actuator 0-100 pe rcen t

Vol t s , Manual Thrust Command 0-14.6 VDC

Vo l t s , Auto Thrust Command 0-12 VDC

PGNS Downlink Data 40 B i t s

SAMPLE RATE SAMPLE/SEC

1

1

1

1

200

200

200

1

1

1

1

1

1

1

1

1

50

1

10

112

20

I v)

i, c, ln S H

In aJ L n

maJ S L Y 3

I Q

I- Y

E P SLL

L -

.I- 3 Uc, *I- m

m o

E E

2s-

P

n aJ c, lu L

P V L a J

aJ L 3 c, X

II c, c m 7 0

a J c ,

OCZ L

'I-

P .I-

n m

n

I E H

V

z

n

5 % 3- v 3 m a

rc n P

n c, v) 3 L c I-

5 3 V m > -

be

E: 0

v) 0 L W

c, m 0 L

n

.I-

L

c, X aJ c, T c,

3 aJ V c m Q L 0 u u a c

O aJ c, 0 aJ L L 0 u In aJ 3

lu > aJ ln a c I-

-r

.I-

P

P

21

TABLE 5 DPS GAGING SYSTEM PERFORMANCE

END OF PHASING BURN

PA R A M ETE R ~~

Oxidizer Tank 1 Measured Quantity, percent Calculated quantity, percent Difference, percent

Oxidizer Tank 2 Measured quantity, percent Calculated quantity, percent Difference, percent

Fuel Tank 1 Measured Quantity, percent Calculated quantity, percent Difference, percent

Fuel T a n k 2 Measured quantity, percent Calculated quantity, percent Difference, percent

Time, hr:min:sec 100: 59: 06

93.8 92.9 to. 9

94.5 92.9 t1.6

92.4 92.7 -0.3

92.0 92.7 -0.7

22

1 t 0 0

+ I 7

-

-

h + I

* m h P

-

h + I

b m h P

d In N

0 0 d

-

h

N P

0 * 0 c

h

N P

-

0 b

P 02 P

0 In ch

co d h

0, N 0 P

0 d a N

- c W S 3

J m

E L LL

-

In

0

0 m h 7

N I I I

- N

~

h h h - 7

In

cu

N I I I

0 e

LLI v, J 3 a 51 U

W Z 3

w o

@ E I v,

m I I I

m

N

P m h P - 0 m

0

G P S 0

a I- LL

5 L cc -0

E (eo L aJ

v) S

-0 C, 3 r v)

L 0 4-

aJ 3

m > S 0

n

2

P

b h ln E

W N

0

* N c

ID W

h N h ?

v,

0" b ch co E 7 m

7

_I

N h

N

co d ol N

m

(u

- d m 0

- d 0 N

co N h

a t

0 . c , aJ

P a J . m u 0 - 7 .I- aJ

n s

.I- C, m

a J a J m u

v, p. n

I

e 1

0811 1

O t L 1 -1 0211

1 080 1

l E- T

090 L 000 1

dS EPE bo WISd - Y N W l WCIIlIH lW3111kl3kl3dnS 3klnSS3tJd

24

~

0 m m m (0 m

I n d. In m 3

0 d m m W m

In m In m W m

0 m m 0 W m

In N In m W m

0 N Ln cr) W m

m In m W 0

c

0 m m W m

c

m 0 m m ID m

h

0 W v,

w

I-

v

E n w m n < J W

RPOL L 0 10 SC 106/L MU-DPS- [ RRW DFITFI 1 !

FIGURE 2

MEASURED HELIUM REGULATOR OUTLET PRESSURE

PHASING BURN

363490 363495 363500 363505 36351 0 26351 5 363520- 363525 363530 363’535 363’540 363’545 363’550 36i555 3613560 GROUND ELLAPSED TIME (SEC) 25

~ ~~

~~~~

~~

8

8 g.

0 0

0. cy d

8 i , H

36

HPOLLO 10 SClOG/LMU-OPS-[RFIW DFITFII

F I G U R E 3

M E A S U R E D O X I D I Z E R I N T E R F A C E P R E S S U R E

P H A S I N G BURN

S H U T W N I GN IT I O N

5 I I I I I I I I I I I I I

1 I

90 363495 363500 363505 36351 0 36351 5 363525 363530 363535 363540 363545 363555 363560 363550. 363520L GROUND ELAPSED T I M E (SEC) , 26

FlPOLLO 10 SClOG/LMU-DPS-~RAW DFITFII

Go361 LP

FIGURE 4 MEASURED FUEL

INTERFACE PRESSURE PHASING BURN

IGNIT ION SHUTDOWN

I I I I I I I I I I I I I 1 363490 363495 363500 363505 36351 0 36351 5 363520 363525 363530 363535 363540 363545 363550 363555 363560 s @

27 GROUND ELAPSED TIME, (SEC)

RPOLLO 10 SClOG/LMU-DPS-[RFIW DFITAI

I ~

I GN I T ION . SHU

FIGURE 5

MEASURED CHAMBER PRESSURE PHASING BURN

3WN

I . . I I I I I I I I I I I

363495 363500 363505 36351 0 36351 5 36 3 5 20- 363525- 363530 363535 363540 363545 363550 363555 363560 28 363490 GROUND ELAPSED TIME, (SEC)

RPOLLO 10 SClOG/LMU-DPS-[RAW DFITFI)

FIGURE 6 MEASURED PROPELLANT QUANTITY

1 OXIDIZER TANK NO. 1 ! PHASING BURN

! I

GQU030

I GN I T I ON SHUT DOWN

w

8

8 s!

I I I I I I I I I I I I 1

RPOLLO 10 SClOG/LM4-DPS-[RAW DFITFII

i

FIGURE 8

' MEASURED PROPELLANT QUANTITY FUEL TANK NO. 1

PHASING BURN

I G N I T I ON I SHU~~DOWN

RPOLLO 10 SClOG/LMU-DPS-(RAW DFITFI)

I GN I T ION 1

FIGURE 7 MEASURED PROPELLANT QUANTITY

OYIDIZER TANK NO. 2 PHASING BURN

SHUTDOWN

I I I I I I I I I I I I I 1 190 363495 363500 3635051 36351 0 36351 5 363520- 363525 363530 363535 363540 363545 363550 ?63555 363560

GROUND ELAPSED TIME, (SEC) 31

J w

5 D D

8 d

RPOLLO 10 SClO6/LM4-DPS-[RRW DATA1

I I I I I I I I I I I I I 1

IGNIT ION I

FIGURE 9 MEASURED PROPELLANT QUANTITY

I FUEL TANK NO. 2 PHASING BURN

SHUTDOWN

GWShUO