ASCENT PROPULSION SYSTEM FINAL FLIGHT EVALUATION

MANNED SPACECRAFT CENTERHOUSTON,TEXAS

JANUARY 1970

173-73555

Dnclas1504100/99

APOLLO 10 MISSION REPORT

SUPPLEMENT 6

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MSC-00126Supplement 6

APOLLO 10 MISSION REPORT

SUPPLEMENT 6

ASCENT PROPULSION SYSTEM FINAL FLIGHT EVALUATION

PREPARED BY

TRW Systems

APPROVED BY

James A. McDivittApollo Spacecraft Program

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

MANNED SPACECRAFT CENTER

HOUSTON, TEXAS

January 1970

11176-H347-RO-OO

PROJECT TECHNICAL REPORT

APOLLO 10

LM-4

ASCENT PROPULSION SYSTEMFINAL FLIGHT EVALUATION

NAS 9-8166 16 September 1969

Prepared forNATIONAL AERONAUTICS AND SPACE ADMINISTRATION

MANNED SPACECRAFT CENTERHOUSTON, TEXAS

Prepared byW. G. Griffin

Propulsion Technoloqy SectionPower Systems Department

NASA/MSC TR\~ SYSTEMS

Concurred

7}

// ~'/ ,~by: /Y;;!c ~ ~/;:&<dZ.»~irkland, HeadS1$tems Analysis Section

Annroved by: 0(.T&..dL.R. J. Smith, ManagerTask E-19E

Concurred/f /' ~"I '/' /Approved by: (>. /'/' -'C-'c~Jh___

P. H. Janak, HeadPronulsion Technoloqy Section,

\./

by:iN~D. W. Vernon, ManagerPower Systems Department

Aprroved. ~''­\

;'I . I \'

by: .(_..l. \_" ! .,l. ~-J7""-C. W. Yodzis, ChiefPrimary Propulsion Branch

Concurred

CONTENTS

PAGE

1. PURPOSE AND SCOPE

2. SUMMARY ...

3. INTRODUCTION

4. STEADY-STATE PERFORMANCE ANALYSIS

2

. 4

. 6

Analysis Technique. . . . . . . . . 6

Flight Data Analysis. . . . . . . .7

Comparison with Preflight Performance Prediction 12

Engine Performance at Standard Interface Conditions 13

5. PRESSURIZATION SYSTEM

Helium Utilization

Ullage Pressure Ouring Coast.Regulator Performance ....

15

15

15

16

6. APS PROPELLANT LOADHiG AND USAGE PRIOR TO BURrJ TO DEPLETTm: • 18

7. PROPELLANT SETTLING ANOMALY

REFERENCES

APPENDIX

i i

19

20

39

1.

Title

LM-4 APS DUTY CYCLE . . .

TABLES

. .Page21

2. LM-4 APS ENGINE AND FEED SYSTEM PHYSICAL CHARACTERISTICS 22

3. RCS PROPELLANT USAGE DURING APS BTD • . . . 23

4. FLIGHT DATA USED IN STEADY-STATE ANALYSIS. 24

5. PROPELLANT CONSUHPTION FROM ASCENT PROPULSION SYSTEMTANKS. . . . . . . . . . . . . . . . . . . . . . . 25

6. STEADY-STATE PERFORMANCE DURING BURN TO DEPLETION. . 26

iii

ILLUSTRATIONS

Title

1. Comparison of Throat Erosion, Predicted and Flight.

2. Acceleration Match During APS BTD

3. Chamber Pressure Match During APS BTD

4. Oxidizer Interface Pressure During APS BTD.

5. Fuel Interface Pressure During APS BTD.

6. Thrust During APS BTD .....

7. Specific Impulse During APS BTD

8. Oxidizer Flow Rate During APS GTD

9. Fuel Flow Rate During APS BTD

10. Oxidizer Interface Differential Pressure (Tank Bottomto Engine Interface) During APS BTD .

Page

27

28

29

30

31

32

33

34

35

36

11. Fuel Interface Differenti al Pressure (Tank Bottom toEngi ne Interface) Du ri nq APS BTD . 37

12. Comparison of Predicted and Flight Reconstructed Performance· 38

iv

1. PURPOSE AND SCOPE

The purpose of this report is to present the results of the postflight

analysis of the Ascent Propulsion System (APS) performance during the

Apollo 10 Mission. Determination of the APS steady-state performance under

actual flight environmental conditions was the primary objective of the

analysis. No formal analysis of APS transient performance was made.

Preliminary postflight analysis of APS performance ;s documented in

Reference 1. This report supersedes Reference 1, and includes such addi­

tional information as is required to provide a comprehensive description of

APS performance during the Apollo 10 Mission.

This report has been prepared as supplement 6 to the Apollo 10

Mission Report (MSC-00126).

Major additions and/or changes to results as presented in Reference 1

are listed below:

1) Revised performance values for the APS BTD.

2) Discussion of analysis techniques, problems and assumptions.

3) Comparison of postflight analysis and preflight prediction.

4) Discussion of the LM-4 propellant settling anomaly.

2. SUMMARY

The duty cycle for the lM-4 Ascent Propulsion System (APS) consisted

of two engine firings; a short manned burn and a longer unmanned burn to

propellant depletion (BTD). APS performance for these two firings was

evaluated and found to be satisfactory.

Engine ignition for the manned burn occurred at a ground elapsed time

(GET) of 102:55:02.1 (hours:minutes:seconds) with the engine being com­

manded off at 102:55:17.7 GET for a total burn duration of 15.6 seconds.

The unmanned BTD was initiated at a GET of 108:52:05.5 and was terminated,

as planned, by fuel-first propellant depletion. Activation of the fuel low

level sensor (llS) occurred at 108:55:24 GET and the chamber pressure began

to decay at 108:55:32.3 GET. Oxidizer llS activation time was 108:55:37

GET. The engine off command was given at 108:56:14.4 GET. One anomaly

was noted during the APS manned burn. The problem concerned the ascent

engine ~uantity caution light which came on approximately one second after

APS first burn ignition command; thus triggering a master caution and warn­

ing a1ann. The engine quantity light is controlled by the oxidizer and/or

fuel low level sensors. Since the low level sensors operated nominally

during the ren,ainder of the first burn and the entire second burn, it is

concluded that the low level sensor did not ~la1function but actually went

dry. Insufficient propellant settling prior to the burn due to the large

u11a£e volume that existed during the lM-4 flight is considered to be the

most probable cause of this difficulty.

r'leasurement data from the APS BTD indicated two abnormalities. The

first problem was observed in the oxidizer interface pressure measurement

data which was approximately 11 psia higher than expected. It was con­

cluded, after analysis, that the data was erroneously high. The second

2

involved oscillations of the helium regulator outlet pressure measurements.

These oscillations did not propogate to the interface or chamber pressures.

Both problems will be more fully discussed in the Pressurization System

section of this report. None of the above difficulties materially affect­

ed APS performance.

Steady-state engine performance parameters averaged over that portion

of the BTO analyzed are as follows:

Thrust - 3432.1 1bf

Isp - 309.49 sec

Mixture Ratio - 1.597

All performance parameters were well within their respective 3-sigma limits.

Calculated engine throat erosion for LM-4 APS was approximately one percent

higher than the predicted at the end of the BTO.

3

3. lNI'IDDocr1CN

Approximately two hours after canpletion of the DPS 001 and phasing

maneuver burn (about 103 hours GET) the ascent and descent stages were

separated and the APS engine was fired for 15.6 seconds (ignition to en­

gine cutoff). Upon canpletion of this insertion maneuver, the ascent

stage docked with the CSM and the crew and ~pnent transfer was effected.

Approximately six hours later, the ascent stage was separated and the en­

gine was ignited for the 248.9 seoond (ignition to cutoff canmand) burn

to propellant depletion (BTD). Exact data ooncerning ignition and cutoff

times and associated velocity changes are shewn in Table l.

The Apollo 10 IM-4 APS was aIUipPed with Rocketdyne Engine SIN 0002B.

APS engine performance characterization equations used in pre-flight pre­

diction and post-flight analysis are found in Reference 2. Engine accep­

tance test data used in the detennination of perfonnance are fran Refer­

ence 3. Physical characteristics of the engine and feed system are pre­

sented in Table 2.

Both firings of the APS engine were preceded by Reaction Control Sys­

tEm (ll:S) maneuvers to settle propellants. The OCS activity prior to the

BTD also included a separation maneuver. Total duration of the ullage and

separation maneuver prior to BID was approximately 90 seoonds.

Only one Apollo 10 Mission Detailed Test Objective (DIO), S13.13,

dealt specifically with APS perfonnance. The functional test objectives of

that DID are as follONS:

1) Confinn APS perfonnance characteristics

2) Confinn that APS propellant depletion shutdo.vn in a space environ­

rrent is not hazardous.

Specific re:JUirerrents of this objective are described in Reference 4.

4

(This page intentionally left blank)

5

4. STEADY-STATE PERFORMANCE ANALYSIS

Analysis Technique

Postflight performance analysis for the LM-4 APS was primarily con­

cerned ;~ith determining steady-state performance during the second ascent

engine firing, an unmanned burn to fuel-first propellant depletion of

248.9 seconds (engine on to engine off command) duration. The first ascent

engine burn was of insufficient length, 15.6 seconds, to properly determine

steady-state performance; however, an examination of the measurement data

(Appendix) was made and performance was found to be satisfactory.

The APS steady-state performance analysis was conducted using the

Apollo Propulsion Analysis Program as the primary tool. This program

utilizes a minimum variance technique to establish the best correlation

between an engine characterization model derived from ground test data and

selected flight test measurements. The minimum variance technique used in

the program consists of a series of error models using various ground and

flight data as inputs and a non-linear APS simulation model which is based

on empirically derived engine characterization equations. Successive itera­

tions through the program result in a "best" estimate, in a minimum variance

sense, of system performance history and weights.

Initial vehicle damp and propellant weights were obtained from

Reference 5, reaction control system (RCS) propellant usage was obtained

from analysis of the bi-level measurements and APS propellant usage for

the first burn was based on estimated steady-state usage. Vehicle damp

weight is considered to be constant, with the exception of the RCS propel­

lants consumed, throughout the run. Table 3 presents ReS consumption for

the propellant settling and separation maneuver just prior to the BTD and

6

during the BTD. All RCS consumption during the BTD was from RCS tanks.

Propellant densities used in the program were based on equations found in

Reference 6 adjusted by measured density data for the LM-4 flight given in

the Spacecraft Operational Data Book, Reference 7. Oxidizer and fuel temp­

eratures were taken from measurement data and were 69.8°F and 70.9°F, res­

pectively. These temperatures were found to be constant throughout the

segment of burn analyzed as steady-state.

The following flight measurement data were used in the analysis of the

LM-4 APS BTD: engine chamber pressure, engine interface pressures, vehicle

thrust acceleration, propellant tank bulk temperatures, helium regulator

outlet pressures, propellant LLS activation times, engine on-off commands

and RCS thruster solenoid bi1eve1 measurements. l~easurement numbers and

other data pertinent to the above measurements, with the exception of RCS

bi-levels, are given in Table 4. Plots of measurement data versus time are

presented in the Appendix to this report.

Flight Data Analysis

A 174 second segment of the APS BTD was selected to be analyzed for

the purpose of determining steady-state engine performance. APS BTD igni­

tion occurred at a GET of 108:52:05.5 and engine cutoff was commanded at

108:56:14.4 GET. The segment of the burn analyzed begins at 108:52:25.0

GET, 19.5 seconds after ignition, and ends at 108:55:19.0 GET. A longer

analysis segment was not feasible since filtered acceleration data was

distorted by the smoothing technique just prior to the start of chamber

pressure decay. Steady-state analysis of the APS BTD revealed no signifi­

cant anomalies. APS engine propellant consumption during the BTD is pre­

sented in Table 5. Consumption from the end of the steady-state analysis

segment to the beginning of chamber pressure decay was extrapolated from

7

strady-state analysis results.

The principal results determined by the APS steady-state analysis are

as follows:

1) Engine throat erosion was slightly higher than predicted, approxi­

mately one per cent at the end of the BTO.

2) Average specific impulse over the 174 second period analyzed was

309.5 seconds.

3) The average mixture ratio based on LLS actuation time was deter­

mined to be 1.597.

4) Oscillations in helium regulator outlet pressure were found to

have no adverse effects on engine performance.

LM-4 APS performance, was determined to be very close to predicted

with actual engine average specific impulse being approximately 0.8 second

higher than the predicted value. A portion of this increased performance

is attributable to a slightly increased throat erosion rate, however, an

increase in engine efficiency was also evidenced.

It should be noted that the number of APS flight measurements was sub­

stantially reduced for the LM-4 flight; thus automatically eliminating

some of the checks and balances that had been available on previous flights.

Of particular significance was the deletion of the tank bottom to engine

interface differential pressure (~P) measurements, since they provided an

excellent means of verifying throat erosion characteristics.

The general solution approach used in the LM-4 flight evaluation was

to calculate a vehicle weight (including propellant loads) for the beginning

of the segment of burn used to analyze steady-state performance and then

allow the Apollo Propulsion Analysis Program to vary this weight and other

selected performance parameters (state variables) in order to achieve an

8

acceptable data match. This technique led to excessive variations in

both propellant loads and vehicle inert weight. In order to resolve these

variations, the vehicle thrust and specific impulse were increased slightly

from predicted values until weight differences become more realistic. A

complicating factor in the analysis was the fact that the oxidizer inter­

face pressure measurement, GP1503P, was erroneously high by approximately

11 pounds per square inch (psi). This bias and an approximately 2 psi pos­

itive bias on fuel interface pressure, determined from ground support equip­

ment data, were applied to the flight interface pressure measurements input

to the program. Program results indicated that the above biases were essen­

tially correct. A further discussion of the oxidizer interface pressure

bias may be found in the Pressurization System section of this report. It

was also necessary to curve fit the interface pressure data since the

smoothing technique applied to the raw data resulted in extensive distor­

tion. The acceleration residual (measured data minus calculated) result­

ing from the program input data outlined above had a definite negative

slope indicating that a slight increase in calculated acceleration as

flight time increased was required to minimize the residual error. This

effect is gained by increasing engine flCMrates and/or increasing engine

thrust on a time basis. Since experience on the LM-3 flight indicated

that actual throat erosion rates could be considerably higher than pre­

dicted and since a slight increase in throat erosion would give the desired

result, a revised throat erosion curve was calculated using the partial

derivatives of throat area with respect to acceleration at ten second

intervals throughout the run. The inclusion of this calculated throat

area curve in the analysis program resulted in an excellent acceleration

match with essentially a zero mean and no significant slope. The derived

9

throat erosion curve was approximately one per cent higher than predicted at

the end of the BTD. This is considerably lower than the value of throat

erosion determined during the LM-3 postflight evaluation. It should be

noted that the segment of the burn analyzed as steady-state is somewhat

shorter for LM-4 than it was for LM-3, however, the LM-4 throat erosion

would not have been as large as that of LM-3 even if the burn times had

been equivalent. Figure 1 shows the calculated throat area curve in

comparison with the predicted curve for LM-4 and the maximum and minimum

curves for which the throat erosion characterization is valid.

Simulation of RCS activity was not as crucial to an accurate postflight

reconstruction of APS performance as it might have been since the thrusters

were theoretically close coupled, i.e., the net translational thrust for

the system is zero. In actuality, a small impulse, approximately equiva­

lent to 5 lbf thrust in the negative X direction, existed throughout the

BTD. The magnitude of this impulse was determined by analyzing accumulated

"on" times for all up and down engines and fllultirlying by a nominal 100 lbf thrust.

The RCS interconnect valves were closed during the BTD, therefore, RCS pro­

pellants came only from the RCS tanks. Weight overboard through the RCS

engines was accounted for by multiplying the nominal one engine flowrate

by tota 1 accumul ated sys tem "on" time and averagi ng thi s fi gure over the

period of the BTD. Usage during a 90 second ullaging and separation maneuver

prior to BTD ignition was accounted for separately but in a similar manner.

Figures 2 through 11 depict the principal performance parameters

associated with the LM-4 postflight analysis. Four flight measurements were

used as time varying input to the propulsion analysis program. Two of these

measurements, fuel and oxidizer interface pressure, were used as program

drivers. The other two, acceleration and chamber pressure, were compared

10

to calculated values by the program's minimum variance technique. The

acceleration and chamber pressure measurements along with their residuals

(measured minus calculated values) are presented in Figures 2 and 3.

respectively. Figures 4 and 5 contain oxidizer and fuel interface pressure

measurement data as it appeared after smoothing of the raw data, the curve

fit of this data that was ultimately input to the Apollo Propulsion Analysis

Program, and the residual between the two. Data shown in Figures 4 and 5

were adjusted by the fixed biases previously discussed prior to being input

to the program. Calculated steady-state values for the following parameters

are shown in Figures 6-11; thrust, specific impulse, oxidizer flow rate,

fuel rate, and oxidizer and fuel tank bottom to engine interface differen­

ti a1 pressures.

The principal indicator of the accuracy of the postflight reconstruc­

tion is the matching of calculated and measured acceleration data. A

measure of the quality of the match is given by the residual slope and

intercept data as shown in Figure 2. This data represents the intercept,

on the ordi nate, and slope of ali near fit to the res i dua1 data. It is

readily seen that the closer both these numbers are to zero, the more

accurate is the match. The acceleration match achieved with the LM-4

postflight reconstruction is excellent. A match of measured and calculated

engine chamber pressure is given in Figure 3 and is also considered to be

quite good. An additional indicator of the validity of the reconstruction

is the matching of the actual amount of propellant in the tanks at LLS

probe actuation times with the program calculated amount of propellant in

the tanks at corresponding times. Based on the densities of the propel­

lants, as determined from flight temperatures, and LLS probe heights, the

weights of oxidizer and fuel in the tanks at their respective LLS actua­

tion times were 48.2 lbm, oxidizer, and 38.4 lbm, fuel.

11

The weights of propellant in the oxidizer and fuel tanks at corresponding

times, as calculated by the propulsion analysis program, were 58.9 lbm,

oxidizer, and 45.0 lbm, fuel. The difference between actual and calculated

in both cases is well within the acceptable variance due to propellant

sloshing.

The LM-4 flight reconstruction is by all indications an accurate simu­

lation of actual flight performance. Residuals between calculated and

measured parameters are all within measurement accuracies.

Comparison with Preflight Performance Prediction

Predicted performance of the LM-4 APS is presented in Reference 8.

The intention of the preflight performance prediction was to simulate APS

performance under flight environmental conditions for the projected mission

duty cycle. No attempt was made in the preflight prediction to simulate

RCS operation.

Table 6 presents a summary of actual and predicted APS performance

during the BTO. Measurement data in Table 6 have been adjusted by known

biases as necessary and compare quite closely with the reconstructed

parameters. The most significant difference between the actual and pre­

dicted APS performance is that predicted regulator outlet and interface

pressures were higher than actual by about 3-4 psi throughout the BTO.

This difference is within the Class I primary regulator operating band.

The pre-flight prediction used a helium regulator outlet pressure based

on the Class I primary regulator operating band mid-point value of 184 psia.

Engine specific impulse determined by the postflight reconstruction is

somewhat higher than had been predicted but is still well within specified

3 sigma limits. A comparison of predicted and reconstructed values for

specific impuls~, thrust, and mixture ratio is presented in Figure 12

12

along with related three sigma disperstons. The variation of flight

specific impulse, thrust and mixture ratio were within their respective

three sigma dispersions.

Engine Performance at Standard Interface Conditions

Expected flight performance of the APS engine was based on a model

characterized with data obtained during engine and injector acceptance

tests. In order to allow actual engine performance variations to be

separated from variations induced by feed system, pressurization system and

propellant temperature variations, the acceptance test data is adjusted to

a set of standard interface conditions, thereby providing a common basis

for comparison. Standard interface conditions are as follows:

Oxidizer interface pre~sure, psidFuel interfacr pressure, psiaOxidiLer interface temp~rature, OFFuel interface temperature, OFOxidi zer density, 1bm/ft3Fuel density, lbm/ft3Thrust acceleration, lbf/lbmThroat area, in2

170.170.70.70.90.2150.391.

16.49

Analysis results (at 20 secrnds from igniticn) for tte 8TD correctRd to

standard inlE:t cOlcitiollS and compared to acceptance test values are

S:lown below:

Acceptance Test Flight Analysis %Data Results Difference

Thrust, 1bf 3508. 3513. . 1

Specific Impulse, 1bf-sec 308.7 309.3 .2----,-om _.

Propellant Mixture Ratio 1.594 1.594 O.

These differences are well within the engine combined repeatability and

acceptance test instrumentation uncertainties. Flight interface pressures

for that part of the BTD analyzed as steady-state were approximately 3 psia

13

below predicted levels. The engine thrust level for flight shows the

effect of the reduced pressure but is not as low as it might have been due

to a compensating increase in engine efficiency. Adjusting engine perfor­

mance to standard interface conditions and comparing with acceptance test

values shows good agreement with the small differences being attributable

to the previously mentioned engine efficiency increase. All differences

are within one standard deviation of acceptance test values. This indi­

cates that basic preflight prediction techniques are adequate, with the

possible exception of the throat erosion characterization. Based on the

results of the LM-3 flight analysis, a question as to the adequacy of the

APS throat erosion was raised. However, as previously discussed, throat

erosion for the LM-4 flight was only slightly higher than predicted. The

question could become somewhat academic in light of the fact that LM-5 and

subsequent vehicles will utilize the SWIP chamber which may exhibit differ­

ent throat erosion characteristics. It is noted that variations in throat

area do not enter into the determination of standard interface condition

performance since the acceptance test throat area becomes part of the stan­

dard interface conditions.

14

5. PRESSURIZATION SYSTEM

Helium Utilization

The helium load for the LM-4 flight was a nominal 13.2 lbm. Helium

tank temperatures and pressures prior to launch were 3067 psia, 71.8°F

and 3083 psia, 71.1oF respectively, for tanks 1 and 2. There was no indi­

cation of helium leakage during the mission. The calculated helium usage

during the burn agrees with analytical predictions.

Ullage Pressure Decay During Coast

The fuel and oxidizer interface pre-launch pressures were 167 and 178

psia, respectively. Telemetry data received during the initial lunar mo­

dule activation period indicated that the oxidizer interface pressure had

increased to 187 psia. The oxidizer temperature just prior to launch was

74°F and was 71°F at initial lunar module activation. The predicted oxi­

dizer interface pressure, based on temperature drop and helium solubility

in propellant, corresponding to the 71°F temperature was 172.5 psia. Fuel

interface pressure at initial lunar module activation was, as expected,

166 psia.

With temperatures holding near constant, pressure levels should either

remain constant or decrease due to helium solubility in the propellants.

In order for tank (or interface) pressures to rise, an increase in helium

mass in the ullage would be required. The data available, plus the system

configuration, indicated no mass could have been added to the ullage from

the only possible source, the helium supply tanks. The conclusion reached

from the above considerations is that the GP1503 measurement was erroneous­

ly high. This conclusion is further substantiated by the agreement of ob­

served data with predictions on the LM-4 fuel side and both fuel and oxi­

dizer measurements on LM-3.

15

Determination of the actual oxidizer interface pressure measurement

bias to be used during the BTD steady-state analysis was made by comparing

the level of the oxidizer interface pressure to that of the fuel interface

pressure after biasing the fuel interface pressure on the basis of ground

test data. It was assumed that the fuel and oxidizer interface pressures

would be within one psia of each other and that the level of the fuel inter­

face pressure, after biasing, was correct. This assumption was based pri­

marily on analysis of the helium regulator outlet pressure data which gave

no indication that the higher level of oxidizer interface pressure was

valid. This technique indicated an oxidizer interface pressure bias of

10-11 psia, with the higher value being chosen. As previously mentioned,

results from the propulsion analysis program substantiated this bias.

Regulator Performance

The regulator lockup pressure at initial APS pressurization \'Jas 184

psia (compared to 186 expected based on tests during checkout at KSC).

Regulation during the insertion burn and lockup after the burn were nominal.

At the start of the burn to depletion, the regulator outlet pressure dropped

to the expected (from KSC checkout) value of 181 psia. At 118 seconds into

the burn, pressure oscillations were observed in both helium regulator

outlet pressure transducers (GP0025P and GP0018P). These oscillations

were present for the remainder of the burn to depletion. GP0025P, located

4 inches downstream of the Class I primary regulator, indicated an ampli­

tude of 5 psi (peak to peak). GP0018P, located upstream of the oxidizer

check va 1ves (3 feet downs tream of GP0025P) i ndi ca ted a maxi mum amplitude

of 19 psi. Since both measurements are located in the helium manifold, the

differences in the magnitude of the oscillations may be explained by (1)

amplification of the pressure oscillations in the helium line or (2) varying

16

sensitivity to oscillations of the two transducers. Pre-installation tests

of the regulator module at Bethpage showed pressure oscillations of 3 psi

peak-to-peak at the regulator outlet (GP0025P) and 6 psi peak-to-peak at

the tank relief valve (downstream of the GP0018P location). The regulators

have a specification limit of 15 psi peak-to-peak.

Pressure oscillations of amplitudes up to 10 psi have been observed at

WSTF during PA-l testing. The oscillations did not propogate to the inter­

face or chamber pressures, due, probably, to the large propellant tank

ullage volumes present during the LM-4 BTD. It is also considered likely

that the pressure transducers amplified the oscillations to a certain extent.

No harmful effects to the system due to the pressure oscillations were noted.

17

6. APS PROPELLANT LOADING AND USAGE PRIOR TO BURN TO DEPLETION

The oxidizer tank was fully loaded at a pressure of 67 psia and an

oxidizer temperature of 72°F. The fuel tank was loaded at a pressure of

64 psia and a fuel temperature of 70°F. A density determination was made

for both oxidizer (1.4818 gm/cc at 4°C and 14.7 psia) and fuel (.8992 gm/cc

at 25°C and 14.7 psia) samples. Based on these density values, propellant

tank pressures, and propellant temperatures; a determination was made of

the quantity of propellant to off load. This off load (1072.0 lbm fuel and

1623.0 lbm oxidizer) was accomplished using the weigh tank five times. The

actual propellant load was determined to be 981.4 lbm fuel and 1650.1 lbm

oxidizer.

RCS consumption from the APS tanks during the Coe11iptic Sequence

Initiation (CSI) and Terminal Phase Initiation (TPI) RCS maneuvers was 28

lbm and 14 lbm, respectively. Therefore, the propellant available to the

APS was 1622 lbm oxidizer and 967 lbm fuel.

APS consumption during the manned burn was approximately 106 lbm of

oxidizer and 67 lbm of fuel. The ReS system interconnect valve was closed

during APS activity so that all RCS consumption was from the RCS tanks.

18

7. PROPELLANT SETTLING ANOMALY

Shortly after the ignition signal for the first APS burn, the ascent

engine quantity caution light came on triggering a master caution and

warning alarm. The quantity caution light is controlled by the oxidizer

and fuel LLS in each propellant tank. Approximately one second after the

ascent engine lI on li signal for the LM-4 APS manned burn, the oxidizer LLS

was activated for about one second, while the master alarm was on for two

seconds before being manually reset. Low level sensors in both oxidizer

and fuel tanks operated as expected during the remainder of the first burn

and the entire second burn. Based on this performance, it was concluded

that the signal received during the first burn was valid.

The cause of this anomaly is believed to be insufficient settling of

APS propellants prior to the burn. Based on Fiqure 4.8-10 in the Space­

craft Operational Data Book (SODB), Reference 9, an RCS propellant settling

maneuver of 3-4 seconds duration would be adequate for a vehicle with the

propellant load of LM-4. The actual RCS ullage propellant settling man­

euver prior to APS first burn ignition lasted for a period of 4.1 seconds.

This would seem to indicate that an adjustment to the SODB propellant set­

tling figures is required for missions utilizing vehicles with large ullage

volumes. Lunar landing missions do not fit into this category for two

reasons; first the positive gravitational field on the lunar surface

is sufficient to settle propellants and secondly the normal APS loading for

lunar landing missions is very near tank capacity.

19

REFERENCES

1. TRW IOC 69.4354.2-52, "Apollo Mission F/LM-4/APS 30-Day Postflight Pro­pulsion Analysis Report," from P. F. Thompson to D. W. Vernon, 10 June1969.

2. North American Rockwell Corporation Document No. PAR 8114-4102, "LunarModule Ascent Engi ne Perfonnance Characterization," prepared by T. A.Clemmer, 10 July 1968.

3. Rocketdyne Engine Log Book, "Acceptance Test Data Package for RocketEngine Assembly - Ascent LM - Part No. RS00039-001, Serial No. 0002B,"dated 1 May 1968.

4. NASA Document No. SPD9-R-037, "Mission Requirements, F Type Mission,SA-205/CSM-106/LM-4," dated 31 January 1969.

5. TRW IOC 69.6522.2.1-30. "Apollo 10 Post Flight Mass Properties," 8 July1969.

6. NASA Memorandum EP23-10-69, from EP/2 Head, Development Section toEP/2 Chief, Primary Propulsion Branch, "Propellant Densities (N204 andA-50);' 18 Februa ry 1969.

7. Spacecraft Operational Data Book, SNA-8-D-027 (III) Rev. 1, Amendment58 Vol. III, 2 May 1969

8. TRW IOC 69.4354.2-47, "Mission F/LM-4/APS Propulsion Performance Pre­diction," from C. G. Johnston to D. W. Vernon, 3 June 1969.

9. Spacecraft Operational Data Book, SNA-8-D-027 (II) Rev 1, Amendment2-180, 11 July 1969.

10. NASA Report MSC-00126, "Apollo 10 Mission Report," August 1969.

20

N ......

TABL

E1

-LM

-4AP

SDU

TYCY

CLES

Ign

itio

nE

ngin

eC

utof

fB

urn

Vel

oci

ty(1

)BU

RNFS

-1FS

-2D

urat

i on

Cha

nge

Hr:

min

:sec

GET

Hr:

min

:sec

GET

Sec

s.F

t/se

c

APS

1st

Bur

n

Lun

arO

rbit

Inse

rtio

nM

aneu

ver

102

:55

:02.

110

2:55

:17.

715

.622

1.

APS

2nd

Bur

n

Bur

nto

Dep

leti

on10

8:52

:05.

510

8:56

:14.

424

8.9

3837

.

(1)

Ref

eren

ce10

TABLE 2

LM-4 APS ENGINE AND FEED SYSTEM PHYSICAL CHARACTERISTICS

Engine(l)

Engine No.

Injector No.

Initial Chamber Throat Area (in. 2)

Nozzle Exit Area (in. 2)

Initial Expansion Ratio

Injector Resistance (lbf -sec2/1bm-ft5)@

time zero and 70°F

Oxidizer

Fuel

Feed System(2)

Total Volume (Pressurized), Check Valves

to Engine Interface (ft3)

Oxidizer

Fuel

Resistance, Tank Bottom to Engine Inter­

face (lb f -sec2/1bm-ft5) at 70°F

Oxidizer

Fuel

Rocketdyne SIN 0002B

Rocketdyne SIN 4094436

16.345

748.7

45.81

12630.

20148.

36.78

36.83

2723.

4402.

(1) Rocketdyne Report PAR 8114-4102, July 10, 1968.

(2) Per telecon with Tom Ervolina, GAEC, Janu~ry 28, 1969.

22

TABLE 3

RCS PROPELLANT USAGE DURING APS BTD

RCS Storage Tanks (lb~) APS Storage Tanks (lbm)

Oxidizer Fuel Total Oxidizer Fuel Total

Propellant Settling and 46.7 23.3 70. a a 0Separation Maneuver

APS (BTD) Ignition* 61. 5 30.7 92.2 0 0 0to Depletion

TOTALS 108.2 54.0 162.2 0 0 0

* Data presented in this table was derived from RCS thruster accumulated"on" time.

23

TABLE 4

FLIGHT DATA USED IN STEADY-STATE ANALYSIS

MeasurementNumber Description

GP2010P Pressure, Thrust Chamber 0-150 psia

Sample RateSample/sec

200.

Pressure, Regulator Outlet Manifold 0-300 psia

Pressure, Engine Oxidizer Interface 0-250 psia

Pressure, Engine Fuel Interface 0-250 psia

Oxidizer Tank Low Level Sensor Off-On

GP1503P

GP1501P

GP1408X

GP0025P

GP1218T

GP0718T

GH1260X

GG0001X*

Temperature, Oxidizer Tank Bulk

Temperature, Fuel Tank Bulk

Ascent, Engine On/Off

PGNS, Down Link Data

20-l20°F

Off-On

Di qi ta1 Code

50

50

* Acceleration determined from PGNS, Down Link Data

24

TABLE 5 - PROPELLANT CONSUMPTION FROMASCENT PROPULSION SYSTEM TANKS

GET APS Propellant on eoa rdEvent Tiiile,

hr T'; t;: sec Oxidizer Fuel

Launch 0:00:00 1650. 1 981.4Ignition for Lunar OrbitInsertion Maneuver 102: 55: 02 .1 1650.1 981.4

Ignition APS Burn to Depletion lOS: 52: 05.0 1516.3 900.3

Start of Chamber PressureDecav for BTD 108:55:32.3 107.0 17.7

25

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ons

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cted

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ulat

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18l.

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181

185

181

Pre

ssur

e(p

sia)

Oxi

dize

rB

ulk

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pera

ture

7070

7070

7070

6970

70o

f Fuel

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kT

empe

ratu

re70

7171

7171

7170

7171

of

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dize

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terf

ace

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ssu

re,

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116

816

717

116

716

716

816

616

6

Fuel

Inte

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116

716

717

116

616

616

816

616

6

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121

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120

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1.60

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Appendix

Flight Data

Figure

A-l APS Thrust Chamber Pressure (GP20l0P - FM) Insertion Burn

A-2 APS Oxidizer Isolation Valve Inlet Pressure (GP1503P-PCM) Insertion Burn

A-3 APS Fuel Isolation Valve Inlet Pressure (GP1501P-PCM) Insertion Burn

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. A-6 APS Helium Supply Tank No.1 Temperature (GP0201T-PCM) Insertion Burn

A-7 APS Helium Supply Tank No.2 Temperature (GPO?02T-PCM) Insertion Burn

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A-9 APS Regulator Out Manifold Pressure (GP0025P-PCM) Insertion Burn

A-10 APS Helium Supply Tank No.1 Pressure (GP0001P-PCM) Insertion Burn

A-ll APS Helium Supply Tank No.2 Pressure (GP0002P-PCM) Insertion Burn

A-12 APS Thrust Chamber Pressure (GP20l0P-FM) BTD

A-13 APS Oxidizer Isolation Valve Inlet Pressure (GP1503P-PCM) BTD

A-14 APS Fuel Isolation Valve Inlet Pressure (GP1501P-PCM) BTD

A-15 APS Oxidizer Tank Bulk Temperature (GP12l8T-PCM) BTD

A-16 APS Fuel Tank Bulk Temoerature (GP07l8T-PCM) BTD

A-17 APS Helium Supply Tank No.1 Temperature (GP0201T-PCM) BTD

A-18 APS Helium Supply Tank No.2 Temperature (GP0202T-PCM) BTD

A-19 APS Regulator Out Manifold Pressure (GP0018P-PCMj BTD

A-20 APS Regulator Out Manifold Pressure (GP0025P-PCM) BTD

A-2l APS Helium Supply Tank No. 1 Pressure (GP0001P-PCM) BTD

A-22 APS Helium Supply Tank No. 2 Pressure (GP0002P-PCM) BTD

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