NASS-37143 

MMC-NLS-SR.001 

-~ January 1992 

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llnAIRTIIV llnAIRIETTA 

MANNED &PACE &V.TEM& 

Book II - Part 2 

Propulsion 

National Aeronautics and Space Administration 

Marshall Space Flight Center 

Michaud Assembly Facility 

Cycle B'(CY1991) 

NLS Trade 

Studies and 

Analyses Report 

(N.\SA-CR-11':'14472) 

~LS T~~1f STUDIF~ ~~~l! O(CY!991) 

q.EP11F.T. -10~JK 7 P • NALYStS N93-23715 

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G3/31 0158787 

FOREWORD 

This document is Book ll, Pan 2 of the Cycle 0 Study Report containing trade 

studies and analyses perfonned by MMC in support of the Propulsion Working Group. 

The work was perfonncd under NASA Contract NASS-37143 between May 1991 and 

January 1992. This study was performed by Manned Space Systems, Martin Marlena 

Corporation, New Orleans, Louisiana for the NASA/Marshall Space Flight Center. 
INTRODUCTION 

This repon documents the propulsion system tasks petfonned in suppon of the 

NLS Cycle 0 preliminary design activities. The repon includes trades and analyses 

covering the following subjects: 1) Maximum Tank Stretch Study; 2) No LOX Bleed 

Petformance Analysis; 3) LOX Bleed Trade Srudy; 4) L02 Tank Pressure Limits; 5) LOX 

Tank Pressurization System Using Helium; 6) STME Heat Exchanger Performance; 7) 

LH2 Passive Recirculation Petformance Analysis; 8) LH2 Bleed/Recirculation Srudy; 

9) LH2 Tank Pressure Limits; 10) LH2 Pressurization System. For each trade study an 

executive summary and a detailed trade study are provided. For the convenience of the 

reader, a separate section containing a compilation of only the executive summaries is also 

provided. 
Table of Contents 

Section 1--Executive Summaries 

3-P-001 

3-P-018 

3-P-019 

3-P-025 

3-P-026 

3-P-027 

3-P-033 

3-P-034 

3-P-038 

3-P-039 

Maximum Tank Stretch Study 

No LOX Bleed Performance Analysis 

LOX Bleed Trade Study 

L02 Tank Pressure Limits 

LOX Tank Pressurization Using Helium 

STME Heat Exchanger Performance 

LH2 Passive Recirculation Performance Analysis 

LH2 Bleed/Recirculation Study 

LH2 Tank Pressure Limits 

LH2 Pressurization System 

Section 2-Complete Trade Studies and Analyses 

3-P-001 

3-P-018 

3-P-019 

3-P-025 

3-P-026 

3-P-027 

3-P-033 

3-P-034 

3-P-038 

3-P-039 

Maximum Tank Stretch Study 

No LOX Bleed Performance Analysis 

LOX Bleed Trade Study 

L02 Tank Pressure Limits 

LOX Tank Pressurization Using Helium 

STME Heat Exchanger Performance 

LH2 Passive Recirculation Performance Analysis 

LH2 Bleed/Recirculation Study 

LH2 Tank Pressure Limits 

LH2 Pressurization System 
Propulsion Study Reports 

Section1 

Executive Summaries 

-
Maximum Tank Stretch Study 

3-P-001 

Martin Marietta Manned Space Systems 

January, 1992 

Prepared By: 

R. Cronin 

Contributors: 

M. Werner 

S.Bonson 

R. Spring 

R. Houston 

Approved By: 

Z. Kirkland 
TABLE OF CONTENTS 

Page No. 

1.0 SUMMARY 1 

2.0 OBJECTIVE 1 

3.0 APPROACH 1 

4.0 RESULTS 2 

4.1 GROUND RULES AND ASSUMPTIONS 2 

4.2 MINIMUM LENGTH MPS 2 

4.3 TANK LENGTH VS FACILITY IMPACTS 2 

4..4 SENSITIVITY ANALYSES 3 

4.5 PAYLOAD PERFORMANCE 3 

5.0 TECHNICAL ISSUES 3 

6.0 CONCLUSIONS AND RECOMMENDATIONS 4 

APPENDIX A - DETAll..ED RESULTS 

II 
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1.0 SUMMARY 

The Maximum Tank Stretch Study, 3-P-001, was perfonned to investigate how much an 

LH2 tank can realistically be stretched to achieve more perfonnance for the 1 1/2 stage NLS 

vehicle. The areas examined were minimum length propulsion module (PM) concepts, 

manufacturing facilities impacts associated with Ul2 tank stretch and potential payload 

performance improvements associated with a stretched tank 1 1/2 stage vehicle. 

It was found that relaxation of some feedline geometry and routine constraints and 

utilization of different feedline flex concepts could save about 69 inches in PM length and 

allow a total of 11.9 ft. tank stretch (L02 and LH2). This includes a 10.8 ft LH2 tank 

stretch aft This can be accommodated by the MAF manufacturing facilities without major 

modifications. This can also provide a potential payload improvement of about 3000 lb for 

the NLS 1 1/2 stage vehicle. · 

Performance and configuration issues arising from this study addressed engine size and · mixture ratio, PM structural arrangement, packaging, staging feedline gimballing and PM 

length weight sensitivities. It was concluded and recommended that these issues should be 

addressed in Cycle 1 studies before the benefits of a stretched tank option could be fully 

evaluated. 

2.0 OBJECTIVE 

The objectives of the maximum tanks stretch study, 3-P-001, are twofold. 

One of the study objectives is to determine the realistic limits on how much the LH2 tank 

can be stretched to achieve more perfom1ance for the 1 1/2 stage NLS vehicle. It must be 

determined how much the Main Propulsion System (MPS) can be shortened. This 

translates into how much the LH2 tank can be stretched while retaining a propulsion 

module design concept similar to the NLS reference. The manufacturing and facilities 

impacts associated with stretching the LH2 tank must also be detennined to define realistic 

stretch limits. 

The second study objective is to detennine the 1 1/2 stage vehicle performance impacts 

associated with a stretched LH2 tank. These performance impacts should assume that the 

L02 tank is stretched slightly to hold engine mixture constant as the LH2 tank is stretched. 

3.0 APPROACH 

The approach taken in this study consisted of a three parallel path task flow as shown in 

Figure 1. One set of tasks consisted of development of a minimum length MPS concept 

and from that calculating parameDic vehicle performance and analyzing the tank stretch 

potential. A second set·of tasks were performed under another related contract study (3-S008A) and consisted of development of the MAF manufacturing and facilities impacts 

associated with LH2 tank stretch. A third set of tasks consisted of development of a list of 

technical issues associated with tank stretch and sensitivity analyses of parameters such as 

vehicle weight and payload performance affected by these issues. The results of all three 

1 
sets of tasks were coordinated to develop conclusions relative to tank stretch and a set of 

recommendations for Cycle 1 were developed 

4.0 RESULTS 

4.1 GROUND RULES AND ASSUMPTIONS 

Certain constraints imposed by the NLS reference configuration were ground ruled for this 

study. These included such items as engine location, a 4{1. PM, feedline geometry and 

routing, prevalves and feedline disconnects similar to those baselined in the NLS reference 

configuration. · 

Assumptions were developed to minimize the MPS length given the above constraints and 

consistent with a Propulsion Module (PM) design similar to the NLS reference. These 

assumptions included that the LH2 feedline to the boosters controls minimum length MPS, 

minimum length contoured feedline outlets are used, 00 slope is minimum for all lines, 1.5 

RID is minimum for pipe bends and lengthy scissors ducts would not be used in feedlines 

to accommodate engine gimballing. · . 

4. 2 MINIMUM LENGTH MPS 

All effort to shorten the MPS was concentrated in shonening the length (Z axis) of the LH2 

booster feedline. This length controls the minimum length routing of the MPS. The 

baseline configuration uses scissors ducts at the engine inlets with pipe bends of RID =2.5 

and minimum line slopes of 15°. By changing the line slopes to 0° and pipe bends to RID= 

1.5, the MPS was shonened by 37 inches relative to the baseline. This reduction translates 

into 37 inches of potential LH2 tank stretch. Replacement of the scissors ducts with 3 pipe 

gimbal joints plus the 1.5 RID bends and 0° slopes allows the MPS to be shortened 69 

inches. This is the preferred concept provided motion analysis shows that adequate 

clearance between. lines is maintained during engine gimballing. 

The use of Pressure Volume Compensated (PVC) ducts was also examined for potential to 

shorten the MPS. PVC length is controlled by engine gimbal requirements with longer 

PVC ducts required for larger gimbal angles. Use of PVC ducts can reduce the MPS 

length by 39 to 72 inches depending on length of the PVC. 

4.3 TANK LENGTH VS FACILITY IMPACTS 

An examination of MAF manufacturing processes and facilities in study 3-S-OOBA 

revealed several facility impacts relative to the ability to stretch the LH2 tank. It was found 

that modifications necessary to stretch the LH2 tank up to 5 feet (NSL baseline) are minor. 

Facility modifications necessary to stretch the LH2 tank from 5-1l.feet are considered 

significant but not major. To stretch a LH2 longer than 11 feet would require major 

modifications to existing production facilities and some new facilities. It was found that 

modification of certain one-of-a-kind facilities to accommodate LH2 tank stretch would be 

critical facility impacts. Cell A (core tank staCking) and Cell E (internal LH2 clean/iridite) are 

critical facilities. Cell A and Cell E have modification for tank stretch limits of 12 and 17 feet 

respectively. Tank stretch beyond these limits would require a new cell. 

The MAF cost impacts associated with these facility impacts were studied under a company 

funded project. This cost study developed a cost impact vs LH2 tank stretch length that 

2 

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increases in unique steps as various facilities are modified to accommodate increasing tank 

length. 

This cost trend reflects the facility modification break points at 11 ft and 17 feet of stretch 

discussed above. 

4. 4 SENSITIVITY ANALYSES 

Using the preferred concept to shonen the propulsion module, preliminary vehicle weight 

trends we~ developed to show the vehicle weight sensitivity to tank stretch. Tank weight 

increased with stretch while propulsion module weight decreased with an overall result of 

vehicle weight decreasing about 1134lb /foot of tank stretch up to a stretch slightly less than 

12 feet. 

The payload performance of the 1 1!2 stage vehicle was examined as a function of tank stretch 

and was found to increase in a non-linear fashion as the tanks are stretched. It was also found. 

that increasing the engine thrust from the NLS baseline (580 KSL) to 640 K (SL) improved 

performance and better utilized the stretch tank capabilities. 

4.5 PAYLOAD PERFORMANCE 

Payload performance of the 1 1/2 stage vehicle was calculated using the assumed vehicle 

weight trends for three LH2 tank lengths, SID ET, NLS refr ( +5 ft) and+ 10ft. The 

length of the L02 tank was adjusted to maintain an engine mixture ratio 6.0. Both the .l\i'LS 

refr ST.l\1E (580K) and a 640K engine thrust level were assumed. It was found that the 

NLS 1 1/2 stage vehicle payload requirement of 50 Klb could be met by either a 10ft 

stretched vehicle with 580K engines or a 5 ft stretched vehicle CNLS ref.) with 640K 

engines. Liftoff thrust/weight is marginal (1.2) for the 10 ft/580K vehicle. It appears that 

the I\TLS ref., length (5 ft stretch) with 640K engines is the better option. 

5. 0 TECHNICAL ISSUES 

Technical issues that evolved from the 3-P-001 configuration and sensitivity studies can 

be grouped into performance issues and configuration issues. 

Performance issues include: 1) Engine mixture ratio (can stretching only the LH2 tank and 

allowing engine mixture ratio to decrease improve stretched vehicle performance?); 2) 

Engine out capabilility (Can engine out requirements be lessened to eliminate the need for 

tank stretch?); 3) Increased engine thrust (should larger and more costly engines be used to 

eliminate the need for tank stretch?); 4) PM Weight vs Length is not well defined (should 

these analyses be refined?); and 5) 1 1/2 stage vehicle performance is extremely sensitive 

to PM vs length assumptions, ie, small changes in structure weight assumptions could 

negate an potential performance gains from increased propellant load (should structure 

weight assumptions be refined by more detailed design?) 

Configuration issues include: 1) Boattail structural design (more detail is needed) 2) 

How are feedlines structure, TVC and other systems packaged in a shonened PM?; 3) 

Should external routing ·of L02 feedlines be considered?; 4) Does the preferred 3 gimbal 

joint feedline concept exceed current gimbal joint technology limits?; and 5) Can the rail 

system used for the reference staging concept be used with a shone ned boattail? 

3 
6.0 CONCLUSIONS AND RECOMMENDATIONS 

The conclusions associated with tank stretch potential are: 1) The LH2 tank can be 

stretched 10-11 feet without major facility impacts; 2) The Ui2 tank can be stretched 10-11 

feet without a major change in the feedline concept; 3) An LH2 tank stretch of 10 feet can 

potentially provide a payload increase of about 3000 lb over the NASA 1 1/2 stage 

reference vehicle; and 4 ) Issues associated with shonened boattail sauctural design and 

packaging must be resolved to verify stretched tank performance improvements. 

These conclusions do not address the issue of, "Is tank stretch th_e best performance 

improvement option for the 1 1/2 stage vehicle or are other options such as increased 

engine thrust worthy of consideration?" 

The following recommendations relevant to stretched tanks were developed from the results 

of this study. Recommendations for cycle 1 study are: 

1) Analyze and develop a minimum length PM concept taking into account 

structural arrangement, packaging, staging, MPS arrangement. and feedline 

gimballing limits. 

2) Calculate minimum length PM mass properties and payload performance of a 

stretched tank/minimum length PM vehicle. 

4 
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· .. :: ... ::.: .. :=\)::~.: .. ·;:::. ::::=: .. ::·::·· .~·=::·= ... · ... ;.; = .. ·• .. ·:·····. 

Reference 

Conliguralion Constraints': 

"' 

Assumptions 

Required to 

Minimize PM Length 

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3-P-001 MAXIMUM T/ "JK STRETCH STUDY 

Develop Minimum 

Length MPS {I I 

·:::··.: ... : (" ·:·;( :'. :,:,:::::;;'.:,:_-:· .. :· 

. ·.···:·· 

. . Develop List of 

Technical 

Issues 

Tank Length 

v. Faclllly Impacts 

(3·S-008a) 

Perform 

Sensitivity Analysis on 

Payload, Weights 

· Parametric 

Performance: 

P/Ls v. Stretch 

Figure 1-Task Flow 

'' Analy~ls'~t Tank·'·''::·~::: 

Stretch Potential ·· 

& Recommendation 

'?,r Detail Analysis ·.·.· ·.:·· '··· .. ·: 

Coordinate 

Results with 

NASA/TIM 

Recommendations lor 

Cycle One 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
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Task Number 3-P-018 

No LOX Bleed Performance Analysis 

Prepared By: 

G. Platt 

Approved By: 

Z. Kirkland 

20 Dec, 1991 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
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Executive Summary 

Task 3-P-018, "No LOX Bleed Performance Analysis" of the National Launch System 

Phase B study done by MMMSS under the Shuttle C Contract reads as follows, "Analysis 

and testing if required to assess the feasibility of no engine and/or vehicle LOX bleeds 

considering probable engine start condition requirements, as well as antigeyser system 

design." This report is based upon the Marshall Space Flight Center study plan dated 

August 5, 1991, and presented at the Technical Interchange Meeting at Marshall Space 

Flight Center on August 28, 30, 1991, by Danny Davis, the cognizant Panel Chairman. 

The NASA Plan presented at the August 28, 1991, TIM does not require testing. 

• Upper loop performance is satisfactory -Temperature rise less than 5 F. for 

• 20 inch main feedlines. 

• 1 inch SOFI on downcomer. 

• Zero to 1/2 inch SOFI on riser. 

• 6 to 12 inch crossover duct diameter. 

• Zero to 35 lb/sec topping and replenish at 163 to 180 deg A at local pressure. 

• Engine feedlines likely to saturate at engine. 

• Geysering may occur. 

• Ambient helium bubbling will mitigate geysering effects, but will ·not cool LOX locally. 

• Most vapor will pass through screen unless screen is flat and horizontal. · 

• Local pressure above saturation for engine start must be established by prepressurization. 

3700 lb. tank weight impact estimated (25 psi higher tank pressure than with cold LOX). 

- -MARTIN MARIETTA 

MANNEDSPACESV~TEMS 

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Task Number 3-P-018 

No LOX Bleed Performance Analysis 

1.0 Summary 

The upper loop performance is satisfactory, the temperature rise for the natural convection 

loop is less than 5 F. The feedlines will be near saturation at the engine inlet, and tank 

prepressurization will be required to 20-25 psi higher than would be required with cold LOX. 

This will result in a LOX tank structural weight impact of 3700 lb. 

2.0 Problem 

Analysis and testing if required to assess the feasibility of no engine and/or vehicle LOX 

bleeds considering probable engine start condition requirements as _well as antigeyser system design. 

3.0 Objective 

General 

• To evaluate the NASA reference feedline design and determine its thermal performance. 

Specific 

• To evaluate the NASA reference feedline design from the standpoint of geysering. 

• To determine whether propellant conditions would be satisfactory for engine start. 

• To identify and evaluate thermal problems and technical costs arising from the NASA reference 

feedline design. 

4.0 Approach 

This study was accomplished by evaluating the reference system relative to geysering, heat up 

from ambient, natural circulation in the upper loop, and screen performance. 

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5.0 Results 

The results of this study are attached. The primary results are listed below. 

6.0 Conclusions and Recommendations 

• Upper loop performance is satisfactory. Temperature rise less than 5 F. 

• Engine feedlines likely to saturate at engine. Geysering may occur. Most vapor will pass 

through screen if screen is not flat and horizontal. 

• Local pressure above saturation for engine start must be established by prepressurization. 

A 3700 lb. tank weight impact is estimated. 

• It is recommended that a prechill system be incorporated. See Task 3-P-019. 

7.0 Supporting Data 

• NASA-CR-64-3, "Mechanics of Geysering of Cryogenics," Martin-~arietta Aerospace 

Corp., 1964. · 

8.0 Attachments 

Study 3-P-018 "No LOX Bleed Performance Analysis." 

MARTIN MARIETTA 

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Executive Summary 

Task 3-P-019 "LOX Bleed Trade Study" of the National Launch System Phase B study 

done by MMMSS under the Shuttle C Contract reads as follows, "Trade study to consider 

bleed vs. no bleed LOX system considering, at a minimum, operability, complexity, start 

sequence restrictions with no bleed, available propulsion module space, and tank stretch 

limits." This report is based upon the Marshall Space Flight Center study plan dated 

August 5, 1991, and presented at the Technical Interchange Meeting at Marshall Space 

Flight Center on August 28, 30, 1991, by Danny Davis, the cognizant Panel Chairman. 

Because of the difficulty in modeling the liquid heating, it was necessary to consider the 

total subcooling of the liquid necessary to start the engine to come from the subcooling 

accomplished by the prepressurization of the tank. This was estimated to be 50 psig in 

an analysis submitted to NASA by MMMSS and an analysis presented to the Chief Engineer 

by the Propulsion Team. Therefore, this value was used as a basis of comparison for 

subsequent subsystem concepts. 

Several subsystem concepts were considered for feedline and pump conditioning. Some 

of the concepts have characteristics that are obviously more desirable than others. 

• Effect on LOX tank design pressure 

• Predictability 

• Repeatability, engine test to vehicle 

• Precedence 

• Impact on engine design 

• Impact on engine test 

• Potential for required future change 

• Operational efficiency 

• Hazard introduced 

• Hardware complexity 

The evaluation of candidate subsystems is summarized as follows. MARTIN MARIETTA 

MANNED SPACE SYSTEMS ~. 

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• Reference No Bleed System cannot be expected to have subcooled propellant at engine 

inlets at start of prepressurization. 

• Reference No Bleed System causes tank prepressurization pressure increase. 

A 20 psi prepressurization increase to 50 psig results in a 3700 lb. tank weight impact. 

• Warm up (vapor pressure increase) after prepressurization is very slow - approximately 0.4 psi/min. 

• The Onboard Bleed looks viable and eliminates the penalty due to prepressurization requirements. 

( 

• The Overboard Bleed to the facility provides good performance but at the cost of increased complexity. 

• The Overboard Bleed Through the Engine to the atmosphere provides good conditions and is 

simple, only a bleed valve and a line to the nozzle exit are required. 

• The LOX dump appears to unduly burden the engine development program unless the engine 

is designed for LOX lead start. · 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
1.0 Summary 

Task Number 3-P-019 

LOX Bleed Trade Study 

The reference "No Bleed" system was compared with four alternative LOX Bleed Systems. 

All would require an engine bleed valve, and all would allow a reduction in LOX tank 

prepressurization pressure and the associated 3700 lb. payload improvement compared to 

the "No Bleed" case. 

2.0 Problem 

"Made study to consider bleed vs. no bleed LOX system considering, at minimum, operability, 

complexity, start sequence restrictions with no bleed, available propulsion module space, and 

tank stretch limits." 

3.0 Objective 

General 

To identify and evaluate alternate LOX bleed systems vs. the reference no bleed system. 

Specific 

To identify and evaluate alternate LOX bleed systems and determine their potential performance 

advantages as compared to a reference no bleed system considering the important attributes of each. 

( 

4.0 Approach 

The approach adopted in performing this study was to consider and analyze LOX bleed systems 

that had previously been used and that were suggested, identify a set of attributes by which the 

systems could be compared, and compare the systems with the reference no bleed system. 

MARTIN MARIETTA 

' MANNED SPACE SV'"',.EMS 

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5.0 Results 

The results of the study are attached. The primary results are listed below. 

6.0 Conclusions and Recommendations 

• Reference No Bleed System cannot be expected to have subcooled propellant at engine 

inlets at start of prepressurization. 

• Reference No Bleed System causes tank prepressurization pressure increase. A 20 psi 

prepressurization increase to 50 psig results in a 3700 lb. tank weight impact. 

• Warm up (vapor pressure increase) after prepressurization is very slowapproximately 0.4 psi/min. 

• The Onboard Bleed looks viable and eliminates the penalty due to prepressurization 

requirements. 

• The Overboard Bleed to the facility provides good performance but at the cost of 

increased complexity. 

• The Overboard Bleed Through the Engine to the atmosphere provides good conditions 

and is simple, only a bleed valve and a line to the nozzle exit are required. 

• The LOX dump appears to unduly burden the engine development program unless the 

engine is designed for LOX lead start. 

7.0 Supporting Data 

Task Number 3-P-018 "No LOX Bleed Peformance Analysis." 

8.0 Attachments 

Task Number 3-P-019 "LOX Bleed Trade Study." 

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MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
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Task Number 3-P-025 

L02 Tank Pressure Limits 

Prepared By: 

Tom Winstead 

20 Dec, 1991 

Approved By: 

Z. Kirkland 

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. Executive Summary 

NASA Statement of Work: 

"Establish L02 tank pressure limits vs. flight time considering engine start, 

shutdown and NPSP requirements, potential pressure stabilization of tank during. 

max airloads, structural weight considering proof test requirements ~nd performance." 

• With no-bleed L02 system, prepressurization will determine tank structural 

requirements. Current estimate of tank impact -4500 Ibm. For no impact on tank, 

prepress needs to be reduced to <30 psig. 

• Vent valve for baseline will be sized by prelaunch operations and will have no 

influence on flight. 

• Optimum NPSP at MECO is 30.8 psi at an ullage pressure of 20.0 psig. 

• Proposed system would have a prepressurization band of 30-32 psig with relief 

set at 34 psig. Structural impact of -500 Ibm is largely offset by a reduction in residuals 

MARTIN MARIETTA 

MANNEDSPACESVr ~MS 

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1.0 Summary 

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Task Number 3-P-025 

L02 Tank Pressure Limits 

The study has shown that the upper pressure limit will be determined by prelaunch 

operations. This results in a 4500 Ibm increase in structural weight. This impact can 

be eliminated by using a bleed conditioning system and reducing the pre-pressurization 

level to less than 30 psig. Lower limit is determined by the saturation pressure of the 

liquid up to terminal drain when the engine NPSP requirement of 30.8 psi becomes 

important. The optimum tank pressure is .... 20 psig which allows the autogenous 

pressurization flowrate to be reduced from the reference 3.0 Ibm/sec to 2.5 Ibm/sec. 

2.0 Problem 

Determine L02 tank pressure limits for the reference configuration. 

3.0 Objective 

Determine tank and system impacts for the reference configuration. 

4.0 Approach 

The approach to performing this study was: 

• Determine tank pressure vs. time for baseline trajectories. 

• Use inputs from 3-P-018, 3-P-019, 3-P-017 and 3-S-01 OA to determine system impacts. 

( 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
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5.0 Results 

The results of this study are attached. The primary results of the study are listed below. 

6.0 Conclusions and Recommendations 

The upper limit will be sized for pre-launch operations. 

Current autogenous flowrate can be reduced to lower tank pressure at MECO. 

Insulated LOX tank. 

Helium Inject. 

Recommend 30-32 prepress band with minimum relief at 34 psig. 

7.0 Supporting Data 

8.0 Attachments 

Study "Task Number 3-P-025, L02 Tank Pressure Limits," dated 12/20/91. 

MARTIN MARIETTA 

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Executive Summary 

NASA Statement of Work: 

"Select optimum L02 tank helium pressuri·zation system based on tank pressure 

limits and specified reference trajectories and considering safety, reliability, operability, 

simplicity, weight. including residuals and cost." 

1.5 Stage 

• Minimum pressurization system weight is achieved using cryogenic storage helium. 

• Ambient storage helium is the next best with fixed orifice autogenous being better at 

higher HEX temperatures. 

HLLV 

• Minimum pressurization system weight is achieved using cryogenic storage helium. 

• Fixed orifice autogenous weight performance is better at higher HEX temperatures. 

• Ambient helium system assumes no bottle staging and consequently will result in 

significant weight impact. 

MARTIN MARIETTA 

MANNEDSPACEsv-TEMS 

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Task Number 3-P-026 

LOX Tank Pressurization System Using Helium 

1.0 Summary 

A trade study was performed to evaluate LOX tank pressurization with ambient and 

cryogenic helium systems. A rough order of magnitude study for the autogenous 

pressurization system was done for comparison. Both ambient and cryogenic helium 

pressurization systems are lighter than an autogenous system, with the difference 

reducing as heat-exchanger outlet temperature is increased. The subsystem costs 

are significantly higher for the helium pressurization system. 

2.0 Problem 

A design concern for autogenous pressurization is that particulate ignition in the 

heat-exchanger discharge (pressurization supply) lines offers a catastrophic vehicle 

failure mode. The potential for failure increases in proportion to the selected heat 

exchanger discharge temperature. 

3.0 Objective 

The NASA statement of work is to "Select optimum L02 tank helium pressurization 

system based on tank pressure limits and specified reference trajectories and 

considering safety, reliability, operability, simplicity, weight, including residuals 

and cost." 

' 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
.( 

4.0 Approach 

The approach was to perform an analysis of the baseline (autogenous) system 

varying heat exchanger outlet temperature to obtain the residual weight sensitivity. 

A similar analysis was performed for helium pressurization system, and design 

features were selected for systems at ambient and cryogenic storage. Cost and 

weight estimates were performed for all three systems and comparisons made. 

5.0 Results 

The system weight impact was compared by summing component and ullage 

weight. The payload weight impact is identical to the weight carried to orbit 

insertion, and increases as pressurant residuals are reduced. The autogenous 

system impacts payload by 4250 lbs. at the baseline conditions. The cryogenic 

helium storage system has a payload impact of only 3180 lbs., but at a cost 

increase of $1.2M/flight when compared to autogenous pressurization. An 

ambient helium system impacts payload by 5500 lbs., but this value can be 

reduced to about 41 00 lbs. by staging bottles with booster engines. This system 

costs about $1.8M/flight more than autogenous. 

6.0 Conclusions and Recommendations 

Minimum weight is achieved using cryogenic stored helium. Ambient stored 

helium is the next lightest, with fixed orifice autogenous being better at higher 

HEX temperatures. There are more components and higher costs for the helium 

systems. The high-temperature GOX issue is traded with equally catastrophic 

bottle-failure issues. · 

MARTIN MARIETTA 

MANNEDSPACESV~TEMS 

( ( 
7.0 Supporting Data 

Task Number 3-P-017, "STME L02 NPSP Requirements" and Task Number 3-P-025 

"L02 Tank Pressure Limits." Nein, M. E. and J. F. Thompson, "Experimental and 

Analytical Studies of Cryogenic Propellant Tank Pressurant Requirements, 

"NASA TN 03177, February 1966. 

8.0 Attachments 

Task Number 3-P-026, "LOX Tank Pressurization System Using Helium." 

( 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
THIS PAGE INTENTIONALLY LEIT BLANK 

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Task Number 3-P-027 

STME Heat Exchanger Performance 

Prepared By: 

G.Piatt 

Approved By: 

Z. Kirkland 

20 Dec, 1991 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Executive Summmary 

NASA Statement of Work: 

Assess current STME heat exchanger performance and possible outlet temperature 

increase. Assuming LOX tank pressurization system uses the STME heat exchanger 

for an energy source, trade system performance (residuals) against engine impacts 

from increased heat exchanger outlet temperature. 

It was shown that for the Autogenous (GOX) pressurization system, the system 

using ambient helium storage, or the system using cryogenic (LH2 temperature) 

helium storage, the heat exchanger discharge temperature should be increased 

above the reference 700°R. It was found that the payload improvement due to 

pressurization system weight saved by increasing the temperature 200°R was 

600- 1000 lb. Tank wall temperatures were increased, as a result of the 200°A 

pressurant temperature increase, by 130°R at cutoff. This tank wall temperature 

increase is not expected to be detrimental. Greater tank wall temperature increases 

will require further evaluation. Heat exchanger cost per flight increases were small 

compared to the benefit in terms of payload improvement ($8400 for 600 - 1000 lb). 

The above work was based on a 1 1/2 stage vehicle. The overall effect is expected , 

to be similar for the HLL V. 

MARTIN MARIETTA --" MANNED SPACE sv-... t:MS 

( ( 
( ( 

Task Number 3-P-027 

STME Heat Exchanger Performance 

1.0 Summary 

It was shown that for the autogenous (GOX) pressurization system or the helium 

system the heat exchanger discharge temperature should be increased above the 

reference 700°R to at least 900° R. This would improve the payload capability of 

the vehicle by 600 - 1 000 lbs. · 

2.0 Problem 

To assess the value of increasing the STME heat exchanger outlet temperature. 

3.0 Objective 

General . 

To evaluate the desirability of increasing the STME heat exchanger outlet temperature. 

Specific . 

To assess, for three candidate LOX tank pressurization systems, the desirability of 

increasing the heat exchanger outlet temperature from the reference 700°A. 

4.0 Approach 

The approach to performing this study was: 

• To calculate a pressurization system weight, for 500, 700 and 900°A heat exchanger 

outlet temperatures, for each of three candidate LOX tank pressurization systems as follows: 

- Autogenous (GOX) pressurization system. 

- Helium system with helium stored in ambient temperature bottles. 

- Helium system with helium stored in bottles submerged in the liquid hydrogen tank. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Calculate the tank wall temperature for the 900°A tank inlet temperature to assure 

tank material integrity at the higher temperature. 

5.0 Results 

The results of this study are attached. The primary results of the study are listed below. 

6.0 Conclusions and Recommendations 

The payload capacity improvement due to pressurization weight saved by increasing 

the heat exchanger outlet temperature from 700 to 900°R was 600- 1000 lb. The· 

corresponding tank wall temperature increase was 130°A to .755°A. Greater tank 

wall temperature increases will require further evaluation. Heat exchanger cost per 

flight increases were small compared to the payload improvement ($8400 for 600- 1000 lb.). 

Further work is required to define the helium pressurization control system. 

7.0 Supporting Data 

STPT fax NM0-086-20 "STME Heat Exchanger Parametrics" dated 10/04/91. 

8.0 Attachments 

Study "Task Number 3-P-027, STME Heat Exchanger Performance" dated 12120/91. 

MARTIN MARIETTA 

MANNED SPACE s· ~Ms 

( ( 
( ( 

Task Number 3-P-033 

LH2 Passive Recirculation Performance Analysis 

Prepared By: 

G. Platt 

Approved By: 

Z. Kirkland 

20 Dec, 1991 

MARTIN MARIETTA 

· MANNED SPACE SYSTEMS 
( 

Executive Summary 

Task 3-P-033, "LH2 Passive Recirculation.Performance Analysis" of the National Launch 

System Phase B study done by MMMSS l:Jnder the Shuttle C contact reads as follows: 

"Analysis of LH2 feed system with passive recirculation system to assess feasibility, margins 

and performance including an assessment of engine prestart restrictions if any." This is a 

report of this study and is based upon the Marshall Space Flight Center study plan dated 

August 5, 1991, and presented at the Technical Interchange Meeting at Marshall Space Flight 

Center on August 28, 30, 1991, by Danny Davis, the cognizant Panel Chairman. 

Conclusions and recommendations were: 

• Simple System. 

• Screens make geysering correlation uncertain. 

• Non-horizontal screens will not become vapor-bound. 

• Saturated liquid hydrogen with 23 cubic inches/second of vapor being produced in pump 

expected after prepress. 

Rapid warmup after start of prepress reduces NPSP 5 psi/min. 

• Makes for short available hold time before depressurization-repressurization required. 

• May force tank design pressure to be increased. . 

• May complicate operations by forcing very short engine start window before recycle required. 

• Reevaluation required when engine start pressure requirement is established. 

MARTIN MARIETTA 

MANNED SPACE sv -w::MS 

( ( 
Task Number 3-P-033 

LH2 Passive Recirculation Performance Analysis 

1.0 Summary 

The LH2 "passive recirculation .. system appears to be capable of furnishing saturated, 

mostly liquid, hydrogen at the turbopump inlet at the start of prepressurization. The heat 

up rate after prepressurization of 5 psi/minute will limit hold time with tanks pressurized 

to 2 - 3 minutes. 

2.0 Problem 

To study and predict the performance of the reference hydrogen no bleed system. 

3.0 Objective 

General 

Determine the performance of the LH2 Passive Recirculation (no-bleed) system. 

Specific 

Gain an understanding of the performance characteristics of the LH2 passive recirculation 

system. Assess geysering situation, feedline screens, and hold time. 

4.0 Approach 

The approach to performing this study was: 

• Calculate engine inlet pressure with tank pressurized and unpressurized. 

• Convert the engine heat flux to vapor volume. 

• Calculate the rate of warmup of pump and propellant due to heating. 

• Research feedline stagnation and geysering; determine performance of system relative to 

geysering limits. 

• Calculate screen performance with regard to passing vapor. MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
"( 

5.0 Results 

The results of this study are attached. The primary results are listed below. 

6.0 Conclusions and Recommendations 

• Simple System. 

• Screen makes geysering correlation uncertain. 

• Non-horizontal screens will not become vapor-bound. 

• Saturated liquid hydrogen with 23 cubic inches/second of vapor being produced in pump 

expected after prepress. 

Rapid warmup after start of prepress reduces NPSP 5 psi/minute. 

• Makes for short available hold time before depressurization-repressurization required. 

• May force tank design pressure to be increased. 

• May complicate operations by forcing very short engine start window before recycle required. 

• Reevaluation required when engine start pressure requirement is established. 

7.0 Supporting Data 

NASA-CR-64-3, Contract NAS8-5418, Summary Report for the Period 1 July 1963 through 

30 June 1964, "Mechanics of Geysering of Cryogenics," dated June 1964. 

STPT CM No. NM0-089-17, "STME Start and Shutdown Requirements," dated 10/25/91. 

STPT CM No. NM0-076-05, "STME Turbopump Heat Leaks," dated 8/29/91. 

8.0 Attachments .. 

Task Number 3-P-033, "LH2 Passive Recirculation Performance Analysis." 

MARTIN MARIETTA 

MANNEDSPACESVrTEMS 

( { 
Task Number 3-P-034 

LH2 Bleed/Recirculation Study 

Prepared By: 

G. Platt 

Approved By: 

Z. Kirkland 

20 Dec, 1991 

MARTIN MARIETTA 

MA.NNED SPACE SYSTEMS 
( 

Executive Summary 

The NASA Statement of Work for this study reads as follows: 

Assess a no-bleed vs. bleed or recirculation system for the LH2 feed system considering at 

a minimum operability, reliability, propulsion module layout and tank stretch, weight, and cost. 

• Reference No-Bleed System will result in saturated LH2 in feed line and engine pump with 

vapor in engine pump and dry lines downstream of engine pump. 

Convection path complicated by screen. 

Analytical model and test program to anchor analytical model required. 

Warm-up after prepressurization increases saturation pressure 5 psi/minute. Would require 

depressurization of tank, repressurization for very short hold. (Engine start pressure not yet defined.) 

• On-Board Bleed has low flowrate, hydrogen quality in turbopump will be poor (80o/o vapor 

by volume). If this is satisfactory, would allow improvement in hold after prepress relative to 

no-bleed system. 

Test program required. 

Slight improvement in performance compared to no bleed. 

• Overboard bleed has adequate performance after prepressurization. Hold time limited due 

to loss of LH2 at 1.2 lb/sec per engine. 

Hardware complexity a disadvantage. 

SSME manufacturer uses this system, in principle, for single engine tests. 

Should be retained for further study. 

• Backward recirculation did not appear advantageous. 

Provides good engine/pump chill. 

Introduces large volume of vapor into feedlines. 

Hardware complexity a disadvantage. 

MARTIN MARIETTA 

MANNED SPACE SV"T'=MS 

( ( ,, 
( ( 

• Forward recirculation: 

Predictable, good experience with systems. 

Hardware complexity a disadvantage. 

Together with overboard bleed, provides best engine/pump chill. 

Provides best performance (best chill, no hold time limitation). 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
.... ,) .... 

Task Number 3-P-034 

LH2 Bleed/Recirculation Study 

1.0 Summary 

. Four alternates to the reference no-bleed system were studied to establish their 

performance characteristics and other attributes. Forward recirculation and the 

overboard bleed to the facility were both superior to the reference system. 

2.0 Problem 

Assess a no-bleed vs. bleed or recirculation system for the LH2 feed system considering 

at a minimum operability. reliability, propulsion module layout and tank stretch, weight, and cost. 

3.0 Objective 

General 

To compare candidate bleed systems with the no-bleed system for the LH2 feed system. 

Specific 

Compare performance, predictability, repeatability. precedence, engine impact, feed 

system impact, engine test impact, potential future change, operational efficiency, potential 

hazard, and hardware complexity of candidate bleed system concepts against the· reference· 

no-bleed system. 

4.0 Approach 

First, the candidate systems were identified and the performance of each was predicted. 

Then the systems were compared with each other and the reference with regard to the 

attributes listed in 2, above. 

MARTIN MARIETTA 

MANNEDSPACESYSTCMS 

( ( ( 
( 

5.0 Results 

The results of this study are attached. The primary results are listed below. 

6.0 Conclusions and Recommendations 

• Reference No-Bleed System will result in saturated LH2 in feedline and engine pump 

with vapor in engine pump and dry lines downstream of engine pump .. 

Convection path complicated by screen. 

Analytical model and test program to anchor analytical model required. · 

Warm-up after prepressurization increases saturation pressure 5 psi/minute. Would 

require depressurization of tank, repressurization for very short hold. 

• On-Board Bleed has low flowrate, hydrogen quality in turbopump will be poor (80°/o vapor 

by volume). If this is satisfactory, would allow improvement in hold after prepress relative 

to no-bleed system. 

Test program required. 

Slight improvement in performance compared to no bleed. 

• Overboard bleed has adequate performance after prepressurization. Hold time limited due 

to loss of LH2 at 1.21b/sec per engine. 

Hardware complexity a disadvantage. 

SSME manufacturer uses this system, in principle, for single engine tests. 

Should be retained for further study. 

• Backward recirculation did not appear advantageous. 

Provides good engine/pump chill. 

Introduces large volume of vapor into feedlines. 

Hardware complexity a disadvantage. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

• Forward recirculation: 

Predictable, good experience with systems. 

Hardware complexity a disadvantage 

Together with overboard bleed, provides best engine/pump chill. 

Provides best performance (best chill, no hold time limitation). 

7.0 Supporting Data 

3-P-033, "LH2 Passive Recirculation Performance Analysis." 

8.0 Attachments 

Task Number 3-P-034, LH2 Bleed/Recirculation Study "LH2 Passive Recirculation 

Performance Analysis." 

"MARTIN MARIETTA 

MANNED SPACE sv--rc:MS 

( ( 
( 

Prepared By: D. Vaughan 

20 Dec, 1991 

( 

Task Number 3-P-038 

LH2 Tank Pressure Limits 

( 

Approved By: 

Z. Kirkland -c 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Executive Summary 

NASA Statement of Work: 

"Establish LH2 tank pressure limits vs. flight time considering engine start 

and NPSP requirements, potential pressure stabilization of tank during 

max airloads, structural weight considering proof test requirements and 

performance. Also consider ascent venting criteria." 

• Current autogenous flowrate results in high tank pressures that will set 

the vent valve relief setting at -60 psig. 

• Structural impact of -7000 Ibm due to high tank pressures·. Pressure 

can be reduced with decreased autogenous flowrate. Proposed flowrates 

of .1.1 Ibm/sec/booster and 0.91bm/sec/sustainer still results in -1500 Ibm 

payload impact. 

• NPSP consideration indicate that the minimum ullage pressure to 

satisfy NPSP requirements will be -31 psig@ MECO. 

• To reduce further the structural impact an alternate pressurization 

system will be required, i.e., flow control valves, step orifice control. 

MARTIN MARIETTA 

MANNED SPACE SV~TEMS 

( ( 
1.0 Summary 

Task Number 3-P-038 

LH2 Tank Pressure Limits 

Baseline system results in very high tank pressure during ascent. Tank impact 

is .... 7000 Ibm. This can be reduced by selecting booster and sustainer pressurization 

flowrates of 1.0 Ibm/sec. The tank impact is then reduced to -1500 Ibm. 

NPSP requirements set the lower pressure requirement at -31 psia. 

2.0 Problem 

Assess LH2 tank pressure limits. 

3.0 Objective 

Determine tank and system impacts for the reference configuration. 

4.0 Approach 

The approach to performing this study was: · 

• To generate ullage pressure vs. time for the reference configuration and assess system impacts. 

• Develop system to minimize impacts to the tank and still maintain adequate NPSP margin. 

5.0 Results 

The results of this study are attached. The main results of the study are listed below. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

6.0 Conclusions and Recommendations 

The baseline autogenous flowrate of 1.41bm/sec results in high tank pressures that impact 

the tank structure by -7000 Ibm and maintain ample margin for NPSP requirements. Reduction 

of the autogenous flowrate to 1.0 Ibm/sec reduces the impact to the tank structure to -1500 Ibm 

and provides adequate NPSP. 

7.0 Supporting Data 

8.0 Attachments 

Study "Task Number 3-P-025, L02 Tank Pressure Limits," dated 12120/91. 

MARTIN MARIETTA 

MANNED SPACE SV~TEMS 

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Executive Summary 

NASA Statement of Work: 

"Select optimum LH2 tank pressurization system based on tank pressure limits 

and specified reference trajectories and considering safety, reliability,operabllity, 

simplicity, weight, Including residuals, and cost." 

Approach: 

Generate baseline pressure profiles for HLLV and 1.5 Stage 

Generate Issues and concerns to reference 

Evaluate reference with structural and NPSP requirements 

Generate alternate approaches 

Results: 

1) Baseline fixed orifice system results In high ullage pressures during flight that 

have a substantial structural weight impact. NPSP requirements are only -31 pslg 

which results In too much margin. 

2) Structural weight Impact can be reduced by reduction in fixed orifice flowrate to 

-1.0 Ibm/sec/engine. This still results in -1500 Ibm Impact due to the high ullage 

pressure that exists during the first portion of the flight. · 

3) Two approaches have been examined to reduce the Initial tank pressure without 

Impacting the NPSP requirement. These are a flow control system arid a step 

pressurization system. . 

4) Assisted customer In set-up and analysis of pressurization systems. . 

MARTIN MARIETTA 

MANNEDSPACESV~TEMS 

( ( 
( 

3-P-039 LH2 Pressurization System 

Background 

Martin Marietta Manned Space Systems was initially assigned performance of the 

subject contract task at the beginning of cycle 0. Early task planning was completed 

and preliminary analysis was done, after which the Propulsion Working Group made 

the decision to complete this task in-house. Martin Marietta continued to participate in 

the task in a review and advisory role. 

This report documents the planning and analysis work performed by Martin 

Marietta and Includes a summary of the results provided to MSFC for their completion 

of this task. · 

( 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
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Propulsion Study Reports 

Section2 

Complete Trade Studies and Analyses 

PREemtNG PAGE BLANK NOT FlLM£0 
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J " • 

Maximum Tank Stretch Study 

3-P-001 

Martin Marietta Manned Space Systems 

January, 1992 

Prepared By: 

R. Cronin 

Contributors: 

M. Werner 

S.Bonson 

R. Spring 

R. Houston 

Approved By: 

Z. Kirkland 

.. 
TABLE OF CONTENTS 

Page No. 

1.0 SUMMARY 1 

2.0 OBJECITVE 1 

3.0 APPROACH 1 

4.0 RESULTS 2 

4.1 GROUND RULES AND ASSUMPTIONS 2 

4.2 MINIMUM LENGTH MPS 2 

4.3 TANK LENG1H VS FAcn..ITY IMPACfS 2 

4.4 SENSITIVITY ANALYSES 3 

4.5 PAYLOAD PERFORMANCE 3 

5.0 TECHNICAL ISSUES 3 

6.0 CONCLUSIONS AND RECOMMENDATIONS 4 

APPENDIX A - DETAll..ED RESULTS 

-- ' II 
' I 

1.0 SUMMARY 

The Maximum Tank Stretch Study, 3-P-001, was performed to investigate how much an 

LH2 tank can realistically be stretched to achieve more performance for the 1 1/2 stage NLS 

vehicle. The areas examined were minimum length propulsion module (PM) concepts, 

manufacturing facilities impacts associated with Ul2 tank stretch and potential payload 

performance improvements associated with a stretched tank 1 112 stage vehicle. 

It was found that relaxation of some feedline geometry and routine constraints and 

utilization of different feedline flex concepts could save about 69 inches in PM length and 

allow a total of 11.9 ft. tank stretch (L02 and LH2). This includes a 10.8 ft LH2 tank 

stretch afL This can be accommodated by the MAF manufactUring facilities without major 

modifications. This can also provide a potential payload improvement of about 3000 lb for 

the NLS 1 1/2 stage vehicle. 

Performance and configuration issues arising from this study addressed engine size and 

mixture ratio, PM structural arrangement, packaging, staging feedline gimballing and PM 

lentrth weitrht sensitivities. It was concluded and recommended that these issues should be 

add-ressed in Cycle 1 studies before the benefits of a stretched tank option could be fully 

evaluated. 

2.0 OBJECTIVE 

The objectives of the maximum tanks stretch study, 3-P-001, are twofold. 

One of the study objectives is to determine the realistic limits on how much the LH2 tank 

can be stretched to achieve more performance for the 1 1/2 stage NLS vehicle. It must be 

determined how much the Main Propulsion System (MPS) can be shortened. This 

translates into how much the LH2 tank can be stretched while retaining a propulsion 

module design concept similar to the NLS reference. The manufacturing and facilities 

impacts associated with stretching the lJi2 tank must also be determined to define realistic 

stretch limits. 

The second study objective is to determine the 1 1n. stage vehicle performance impacts 

associated with a stretched LH2 tank. These perfonnance impacts should assume that the 

L02 tank is stretched slightly to hold engine mixture constant as the LH2 tank is stretched. 

3. 0 APPROACH 

The approach taken in this study consisted of a three parallel path task flow as shown in 

Figure 1. One set of tasks consisted of development of a minimum length MPS concept 

and from that calculating parametric vehicle performance and analyzing the tank stretch 

potential. A second set of tasks were performed under another related contract study (3-S008A) and consisted of development of the MAF manufacturing and facilities impacts 

associated with LH2 tank stretch. A third set of tasks consisted of development of a list of 

technical issues associated with tank stretch and sensitivity analyses of parameters such as 

vehicle weight and payload perfonnance affected by these issues. The results of all three 

1 
sets of taSks were coordinated to develop conclusions relative to tank stretch and a set of 

recommendations for Cycle 1 were developed 

4.0 RESULTS 

4.1 GROUND RULES AND ASSUMPTIONS 

Certain constraints imposed by the NLS reference configuration were ground ruled for this 

study. These included such items as engine location, a 4{2 PM, feedline geometry and 

routing, prevalves and feedline disconnects similar to those baselined in the NLS reference 

configuration. 

Assumptions were developed to minimize the MPS length given the above constraints and 

consistent with a Propulsion Module (PM) design similar to the NLS reference. These 

assumptions included that the LH2 feedline to the boosters controls minimum length MPS, 

minimum length contoured feedline outlets are used, 00 slope is minimum for all lines, 1.5 

RID is minimum for pipe bends and lengthy scissors ducts would not be used in feedlines 

to accommodate engine gimballing. 

4. 2 MINIMUM LENGTH MPS 

All effon to shonen the MPS was concentrated in shortening the length (Z axis) of the LH2 

booster feedline. This length controls the minimum length routing of the MPS. The 

baseline configuration uses scissors ducts at the engine inlets with pipe bends of RID =2.5 

and minimum line slopes of 15°. By changing the line slopes to 0° and pipe bends to RID= 

1.5, the MPS was shonened by 37 inches relative to the baseline. This reduction translates 

into 37 inches of potential LH2 tank stretch. Replacement of the scissors ducts with 3 pipe 

gimbal joints plus the 1.5 RIO bends and 0° slopes allows the MPS to be shortened 69 

inches. This is the preferred concept provided motion analysis shows that adequate 

clearance between lines is maintained during engine gimballing. 

The use of Pressure Volume Compensated (PVC) ducts was also examined for potential to 

shorten the MPS. PVC length is controlled by engine gimbal requirements with longer 

PVC ducts required for larger gimbal angles. Use of PVC ducts can reduce the MPS 

length by 39 to 72 inches depending on length of the PVC. 

4.3 TANK LENGTH VS FACILITY IMPACTS 

An examination of MAF manufacturing processes and facilities in study 3-S-OOSA 

revealed several facility impacts relative to the ability to stretch the LH2 tank. It was found 

that modifications necessary to stretch the LH2 tank up to 5 feet (NSL baseline) are minor. 

Facility modifications necessary to stretch the LH2 tank from 5-11 feet are considered 

significant but not major. To stretch a LH2 longer than 11 feet would require major 

modifications to existing production facilities and some new facilities. It was found that 

modification of certain one-of-a-kind facilities to accommodate LH2 tank stretch would be 

critical facility impacts. Cell A (core tank stacking) and Cell E (internal LH2 clean/iridite) are 

critical facilities. Cell A and Cell E have modification for tank stretch limits of 12 and 17 feet 

respectively. Tank stretch beyond these limits would require a new cell. 

The MAF cost impacts associated with these facility impacts were studied under a company 

funded project. This cost study developed a cost impact vs LH2 tank stretch length that 

2 
increases in unique steps as various facilities are modified to accommodate increasing tank 

length. 

This cost trend reflects the facility modification break points at 11 ft and 17 feet of stretch 

discussed above. 

4. 4 SENSITIVITY ANALYSES 

Using the prefem:d concept to shorten the propulsion module, preliminary vehicle weight 

trends were developed to show the vehicle weight sensitivity to tank stretch. Tank weight 

increased with stretch while propulsion module weight decreased with an overall result of 

vehicle weight decreasing about 1134lb /foot of tank stretch up to a stretch slightly less than 

12 feeL 

The payload performance of the 1 1/2 stage vehicle was examined as a function of tank stretch 

and was found to increase in a non-linear fashion as the tanks are stretched. It was also found . that increasing the engine thrust from the NLS baseline {580 KSL) to 640 K (SL) improved 

performance and better utilized the stretch tank capabilities. 

4. 5 PAYLOAD PERFORMANCE 

Payload performance of the 1 1/2 stage vehicle was calculated using the assumed vehicle 

weight trends for three LH2 tank lengths, SID ET, NLS refr (+5 ft) and+ 10ft. The 

length of the L02 tank was adjusted to maintain an engine mixture ratio 6.0. Both the NLS 

refr STME (580K) and a 640K engine thrust level were assumed. It was found that the 

NLS 1 1/2 stage vehicle payload requirement of SO Klb could be met by either a 10 ft 

stretched vehicle with 580K engines or a 5 ft stretched vehicle (NLS ref.) with 640K 

engines. Liftoffthrust/weight is marginal (1.2) for the 10 fr/580K vehicle. It appears that 

the NLS ref., length (5 ft stretch) with 640K engines is the better option. 

5. 0 TECHNICAL ISSUES 

Technical issues that evolved from the 3-P-001 configuration and sensitivity studies can 

be grouped into performance issues and configuration issues. 

Performance issues include: 1) Engine mixture ratio (can stretching only the LH2 tank and 

allowing engine mixture ratio to decrease improve stretched vehicle performance?); 2} 

Engine out capabilility (Can engine out requirements be lessened to eliminate the need for 

tank stretch?); 3) Increased engine thrust (should larger and more costly engines be used to 

eliminate the need for tank stretch?); 4) PM Weight vs Length is not well defmed (should 

these analyses be refined?); and 5) 1 1/2 stage vehicle performance is extremely sensitive 

to PM vs length assumptions, ie, small changes in structure weight assumptions could 

negate an potential performance gains from increased propellant load (should structure 

weight assumptions be refined by more detailed design?) 

Configuration issues include: 1) Boattail structural design (more detail is needed) 2) 

How are feedlines structure, TVC and other systems packaged in a shonened PM?; 3) 

Should external routing ofL02 feedlines be considered?; 4) Does the preferred 3 gimbal 

joint feedline concept exceed current gimbal joint technology limits?; and 5) Can the rail 

system used for the reference staging concept be used with a shonened boattail? 

3 

-
--

6.0 CONCLUSIONS AND RECOMMENDATIONS 

The conclusions associated with tank stretch potential are: 1) The LH2 tank can be 

stretched 10-11 feet without major facility impacts; 2) The LH2 tank can be stretched 10-11 

feet without a major change in the feedline concept; 3) An LH2 tank stretch of 10 feet can 

potentially provide a payload increase of about 3000 lb over the NASA 1 1/2 stage 

reference vehicle; and 4 ) Issues associated with shonened boattail structural design and 

packaging must be resolved to verify stretched tank perfonnance improvements. 

These conclusions do not address the issue of, "Is tank stretch the best perfonnance 

improvement option for the 1 1/2 stage vehicle or are other options such as increased 

engine thrust worthy of consideration?" 

The following recommendations relevant to stretched tanks were developed from the results 

of this study. Recommendations for cycle 1 study are: 

1) Analyze and develop a minimum length PM concept taking into account 

sttuctural arrangement, packaging, staging, MPS arrangement, and feedline 

gimballing limits. 

2) Calculate minimum length PM mass properties and payload performance of a 

stretched tank/minimum length PM vehicle. 

4 
Reference 

Configuration Constraints · 

Cl'l 

Assumptions 

Required to 

Minimize PM length 

( 

3-P-001 MAXIMUM TANK STRETCH STUDY 

·::>..) .. _ .. .;: ;.·: 

DeveloP Minimum 

lengt~ MPS 

Tank length 

Develop llsi ol 

Technical 

Issues 

v. FadHty Impacts 

(3-S-OOBa) 

Perform 

Sensitivity Analysts on 

Payload, Weights 

Parametric 

Perlormance: 

P/Ls v. Stretch 

Figure 1-Task Flow 

( 

Analysis ol Tank 

Stretch Potential 

& Reconmendatlon 

tor Detail Analysis 

Coordinate 

Results with 

··. NASAfTIM 

Recommendations lor 

Cycle One 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

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Reference Confi~furation Constraints - 3-P-001 

• 4/2 Engine Configuration 

• 1 Line Diameter Straight Length Before Engine Inlets 

• 1 Line Diameter Prevalve 

• Prevalve At Each Engine 

• Vertical Disconnects 

• Engine Mounting At Xt = 4383.28 

• PM Feedline Routing Internal 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

........ '"''H IQ1.flAI Conlio. 
Assumptions To Minimize MPS Le.ngth - 3-P-001 

• LH2 Feedline To Booster Controls Minimum Length Routing 

• Mimimum Length Contoured Feedline Outlets 

• oo Slope Minimum On All Lines 

• 1.5 RID Minimum For Pipe Bends 

• No Scissors Ducts 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

JMP.12111( lions ( ( 
·-

c 

·-:! 
Minimum Length MPS Using Scissors· Joints 

Minimum feedline routing length is controlled by the Z axis length of the booster LH2 feedlines. 

The baseline LH2 sustainer feedline uses pipe bends with an RID of 2.5, 15° minimum line slopes and scissors 

ducts at the engine Inlets to accomodate gimballing. 

Reduction of pipe bends to RID = 1.5 and line slopes to oo can shorten the LH2 feedline routing by 37 inches. This 

37 inch reduction translates to 37 inches of potential LH2 tank stretch. 

( Jf{ -:.12110/91·1 ( ( 
( 

LH2 F/L Comparison - Booster 

• Baseline with Scissors & Minimum Routing 

• 2.5 RID Bends 

•15° Slopes 

~ --Xn:4238 $=$ 

37" t 

Xn=4383.28 

• 1.5 RID Bends 

• oo Slopes 

• New Groundrules Allows 37" Additional Stretch 

MARTIN MARIETTA 

05.MPW.911003 MANNED SPACE SYSTEMS 
( 

Minimum Length MPS Using Gimbal Joints 

Replacement of the scissors ducts with three pipe gimbal joints plus 1.5 AID bends and oo slopes shortens the LH2 

feedline routing length by 69 inches. This translates to a potential tank stretch of 69 inches. 

Feedline movement during engine gimballing must be verified for this concept. 

~.12/10/91-2· ( ( 
( 

( 

LH2 F/L Comparison- Boo~ter 

• Baseline with Scissors compared to Minimum No Scissors 

• 2.5 R/D Bends 

• 15° Slopes 

04.MPW.911022 

$:=$ Xn=4238 1 \ 

. 69" • t.5 R/D Bends 

• oo Slopes 

Xn:4383.28 • ·No Scissors Duct 

• No Scissors & New Groundrules Allows 

69" Shorter PM 

• Engine Gimballing Dynamics to be Verified 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
Minimum Length MPS Using PVC Ducts 

Use of Pressure Volume Compensated (PVC) ducts can reduce the LH2 feedline routing length by 39 to 72 inches 

depending on length of the PVC. A 50 inch PVC and bends with 1.5 RID can shorten the feedlines and hence allow 

a LH2 tank stretch of 72 inches. 

PVC length is controlled by engine gimbal requirements. larger gimbal angles would mandate longer PVC ducts. 

( ,, :::.12/10/91-3 ( ( 
( ( 

LH2 F/L Comparison - Booster 

• Baseline with Scissors compared to Minimum PVC Ducts 

• 2.5 R/D Bends 

•15° Slopes 

39" 

Xn:4383.28 (}d \ 

cj:?f~ 

72" 

• 1.5 R/D Bends 

• 80" Long PVC • 50" Long PVC 

• PVC Ducts & New Ground rules Allows 39"to 72" Added Stretch 

• PVC Duct Lenath Varies with Gimbal Reauirements 

MARTIN MARIETTA 

MPW.911206 

MANNED SPACE SYSTEMS 
Minimum Length MPS Feedllne Clearance Critical 

The close proximity of lines in a shortened MPS dictates the need for a feedline motion during engine gimballing 

analysis. · 

( c. 12/1 0/91·4 

( ( 
( 

( 

MPS Arrangement 

• Minimum MPS Arrangement with No Scissors & oo Slope 

MARTIN MARIETTA 

MPW.911206 MANNED SPACE SYSTEMS 
.. 

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PREROING PAGE BLANK NOT F'ILM£0 
MAF Modifications For LH2 Tank Stretch 

The MAF modifications necessary to stretch the LH2 tank 5 feet are considered minor. 

Facility modification necessary to stretch the LH2 tank up to 11 ft. are considered to be significant but not major. 

To stretch a LH2 tank longer than 11 ft. would require major modifications to existing production facilities and some 

new facilities. · 

.·, 

( J C.12/10/91-5 ( ( 
• 1 .... ( 

Summary 3-S-OOSA 

• Reference Configuration LH2 Tank Stretch Feasibility Re-Confirmed 

- 5 ft Stretch Requires Minor or No Modifications 

• Tank Stretch up to 11 ft is Possible with Facility Modifications: 

- Cell E - Internal LH2 Clean & lridite 

- Cell A - Core Tankage Vertical Stack 

- Cell P - External Clean & Prime 

- LH2 Major Weld Assy 

- LH2 Proof Test(Bidg 451) 

• New Facilities/Major Mods are Required above 11 ft 

- New Proof Test Facility @ 11 ft 

-New VAB Cell A@ 12ft 

-New VAB Cell E@ 17ft 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Critical Facility Impacts, LH2 Tank Stretch 

Modification to accomodate LH2 tank stretch.of certain one-of-a-kind facilities are considered critical facility impacts 

The internal LH2 clean and iridite, Cell E., can be modified to accomodate tank stretch from 5 to 17 feet. A stretch of 

over 17 feet would require a new cell. 

For core tankage stacking, Cell A, stretch up to 12 feet can be accomodated with varying degrees of modification. A 

stretch of over 12 feet would require a new cell. 

~. 12/10/91-6 

( ( 
( 

Critical Facility Impacts - One-of-a-Kind Cells 3-S-OOSA 

• Cell E • Internal LH2 Clean and lridite 

-Stretch 5 ft - Minor Tool & Facility Modification 

- Stretch 5 • 11 ft - Raise Roof & Lengthen Door 

-Stretch 11 -17ft - Raise Roof, Lengthen Door and Lower Sill 

·Over 17ft -NEW CELL 

• Cell A - Core Tankage Stack 

- 8 ft 6 in LH2 Stretch Without Major Facility Modification 

- Over 8 ft 6 in - 12 ft - Modify TPS Closeout Room 

-Over 12ft. -NEW CELL 

MARTIN MARIETTA 

A ~OUST NLS 910916 . -- ...,. . 1\A!ANN&D APACE AVRT&UA 
( 

Groundrules And Assumptions 

• Delta Costs Are For Core Stage Vehicle 

-Tankage 

-Skirts 

- Propulsion Module 

-Avionics 

-IACO 

• 5' Stretch Common Core Tankage Is Baseline 

• Current ET Processes And Technology 

( 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( 
( 

NLS Core Tankage Stretch Summary 

20 ~ 

-

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- 15 1-

- Million :-

-

- 10 )-

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5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 

Feet 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

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Stretched NLS ROM Mass Properties - 1.5 Stage M:6.0 

Vehicle Element Refr. - 5.0' Stretch 11.9' Stretch 

Titan IV Shroud 13569 13569 

lnterstage/Transition 5563 5563 

Forward Skirt 2603 

L02Tank 14061 (1~9~31 lntertank 12071 

LH2Tank 36152 ~~~~~?I Aft Skirt 3746 3746 

Auxilliary Syst./Hdw. 13130 13130 

Avionics 3714 3714 

Propulsion Module 107261 . ~65711 Incl. Engines 

Total Dry Weight 211872 204049 

-

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

1.5 Stage Vehicle Impacts, LH2 Tank Stretch 

Payload capability of a 1 1/2 stage launch vehicle increases in a non-linear fashion as the tanks are stretched to 

accomodate more propellant. 

A LH2 tank stretch of 5 feet (NLS baseline) is insufficient to meet the NLS 50 K lb, 1 1/2 stage payload requirement 

for this vehicle when using the baseline 580 K (SL) STME and assuming one engine out. An 8 ft. stretch will meet 

the P/L requirement. 

Increasing STME thrust to 640 K (SL) will allow the NLS vehicle to meet the payload requirement (50 K lb) with the 

baselined tank size (5 ft. stretch). · 

"':.12/10/91·7 

( ( 
.. 

 

1.5 Stage VehJci~ __ PerforiJl(!_IJCe Impacts 

Objective : Impact to Payload Performance due to Engine Thrust 

and Tank Stretch for Sustainer Out at Liftoff 

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60000,-------------------~------------------~ 

58000 

56000 

54000 

P/l Rqmt 

52000 

~--- ..........__ 580 K 50000 .... - - - - - - - - - - - - - - - -

48000 

46000 

44000 

42000 

Ref (5 ft Stretch) 

40000 I • I I I I I • • I I I I I I I I I i • I I I I I I I I I I I I I I I I I 

74 75 76 77 78 79 80 81 82 83 84 85 86 

Tank Length (LH2) - ll 

Ground Rules 

• 5/28/91 Ref Wts 

• Mixture Ratio = 6.0 

• ISP = 430.25 

• 80 nm x 150 nm 

• MSFC TDDP Traj 

Assumptions 

• ~ LH2 Tank Wt Estimated 

• A L02 Tank Wt Estimated 

• A PM Wt Estimated 

-Staged Wt 

- Sustainer Wt 

• Residuals/Reserves Adj 

1.5 Stage Vehicle Sustainer Engine Out at Liftoff Performance 

Can exceed 50 Klb Rqmt with 640 K STME and Addnl 5 ft Stretch 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
> • 

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"" 

( 

1.5 Stage Vehicle-Stretch Tank Performance 

The propulsion module (boattail) was assumed to decrease in weight as its length decreased. This decrease 

outweighs the increase in propellant tank weight as the tanks are stretched. This overall decrease in inert weight 

coupled with more useable propellant produces a payload gain for a stretched vehicle. 

The magnitude of the propulsion module weight vs. length characteristic must be verified by detailed structural 

design and analysis to ensure the payload vs stretch gains shown in this preliminary performance statement. 

:.12/10/91-8 ( ( 
( 

STME vacuum Thrust 583,400 640,000 

LH2 Tank Length Std ET (75 ft) NLS ref (+5) (+10) Std ET (75 ft) NLS ref (+5) (+10) 

Gross Lift Off Weight 1,868,351 1,959,289 2,049,298 1,880,955 1,972,564 2,063,437 

Lift Off Thrust I weight 1.332 1.267 1.208 1.452 1.380 1.315 

Used Impulse Propellant 1,593,245 1,683,245 1,773,245 1,593,245 1,683,245 1,n3,245 

MECOWelght 193,737 200,218 205,772 200,818 208,665 215,776 

Tank weight 73,279 76,156 79,033 73,279 76,156 79,033 

Boattall Weight retalned(1) 53,523 48,378 43,233 60,666 54,428 48,190 

Boatlall weight staged(2) 69,746 64,203 58,658 75,185 68,947 62,709 

Total Inert Weight at MECO 125,269 124,534 123,819 133,945 130,584 128,948 

Reserves and residuals 16,093 17,166 20,433 16,470 17,568 20,835 

Transition Section 4,358 4,358 4,358 4,358 4,358 4,358 

Total non payload Weight 145,720 146,058 148,610 154,n3 152,510 154,141 

Gross Payload 48,017 54,160 57,162 46,045 56,155 61,635 

Performance Margin 4,802 5,416 5,716 4,605 5,616 6,164 

Net Payload 43,215 48,744 51,446 41,441 50,540 55,472 

Note: (1) with 2 STME's 

(2) with 4 STME's 

Mixture Ratio • 6.0 

Jsp • 430.5 seconds 

Direct InJect to 80 nm by 150 nm 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
tn 

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Maximum Tank Stretch - Issues 3-P-001 

Performance Issues 

• Engine Mixture Ratio 

• Engine Out Capability 

• Increased Engine Thrust 

• PM Weight vs. Length Not Well Defined 

• 1 1/2 Stage Vehicle Performance Extremely Sensitive To 

PM Weight vs. Length Assumptions, ie, Small Changes In 

Structure Weight Assumptions Could Negate Any 

Potential Performance Gains From Increased Propellant .Load 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS . 
Maximum Tank Stretch - Issues 3-P-001 

Configuration Issues 

• Boattail Structural Design 

• Packaging In PM - Feedlines, Structure, TVC, Other Subsystems 

• External Routing Of Feedlines 

• Feed line Gimbal Joint Technology Limits 

• Can Rail System Work With Short Boattail For Reference Staging 

Concept 

MARTIN MARIETrA 

MANNED SPACE SYSTEMS 

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Maximum Tank Stretch - Conclusions 3-P-001 

• LH2 Tank Can Be Stretched Aft 10-11 Ft Without Major Facility Impacts 

• LH2 Tank Can Be Stretched Aft 10-11 Ft. Without Major Change In 

Feedline Concept 

• A LH2 Tank Stretch Of 10Ft. Can Potentially Provide A Payload Increase 

Of About 3000 Lb. Over The NASA 1.5 Stage Reference Vehicle 

' 

• Issues Associated With Shortened Boattail Structural Design And 

Packaging Must Be Resolved To Verify Stretched Tank Performance 

Improvements 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( ( 
( 

Recommendations For Cycle 1 3-P-001 

• Using The Revised NLS Cycle 1 Propulsion Module Reference 

Configuration · 

• Cycle 1 Recommendations Relevent To Stretched Tanks 

- Re-Confirm Minimum MPS Length Using Cycle 0 Ground Rules - Develop Minimum Length PM Structural Arrangement - Develop Subsystem (TVC, Avionics, Pneumatic} Packaging For 

Minimum Length PM - Develop Staging Concept Details For Minimum Length PM And 

Analyze 

- Develop MPS Arrangement (Internal vs. External L02 Feedlines 

For Minimum Length PM - Analyze Feedline Movement During Engine Gimballing (Develop 

Feedline Gimbal Requirements} - Develop Minimum Length Propulsion Module Mass Properties ·. 

- Calculate Payload Performance Of Stretched Tank/Minimum 

Length PM Vehicle 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
.. 

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~ Task Number 3-P-018 

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2 " No LOX Bleed Performance ~nalysis 0 

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Prepared By: 

G. Platt 

Approved By: 

Z. Kirkland 

20 Dec, 1991 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( ,, 

Executive Summary 

Task 3-P-018, "No LOX Bleed Performance Analysis" of the National Launch System 

Phase B study done by MMMSS under the Shuttle C Contract reads as follows, "Analysis 

and testing if required to assess the feasibility of no engine and/or vehicle LOX bleeds 

considering probable engine start condition requirements, as well as antigeyser system 

design." This report is based upon the Marshall Space Flight Center study plan dated 

August 5, 1991, and presented at the Technical Interchange Meeting at Marshall Space 

Flight Center on August 28, 30, 1991, by Danny Davis, the cognizant Panel Chairman. 

The NASA Plan presented at the August 28, 1991, TIM does not require testing. 

• Upper loop performance is satisfactory - Temperature rise less than 5 F. for 

• 20 inch main feedlines. 

• 1 inch SOFI on downcomer. 

• Zero to 1/2 inch SOFI on riser. 

• 6 to 12 inch crossover duct diameter. 

• Zero to 35 lb/sec topping and replenish at 163 to 180 deg R at .local pressure. 

• Engine feedlines likely to saturate at engine. 

• Geysering may occur. 

• Ambient helium bubbling will mitigate geysering effects, but will·not cool LOX locally. 

• Most vapor will pass through screen unless screen is flat and horizontal. ~ 

• Local pressure above saturation for engine start must be established by prepressurization. 

3700 lb. tank weight impact estimated (25 psi higher tank pressure than with cold LOX). 

·MARTIN MARIETTA 

MANNED SPACE S' ..... T-=MS 

( ( 
( 

( 

/ 

Task Number 3-P-018 

No LOX Bleed Performance Analysis 

1.0 Summary 

The upper loop performance is satisfactory, the temperature rise for the natural convection 

loop is less than 5 F. The feedlines will be near saturation at the engine inlet, and tank 

prepressurization will be required to 20-25 psi higher than would be required with cold LOX. 

This will result in a LOX tank structural weight impact of 3700 lb. 

2.0 Problem 

Analysis and testing if required to assess the feasibility of no engine and/or vehicle LOX 

bleeds considering probable engine start condition requirements as well as antigeyser system design. 

3.0 Objective 

General 

• To evaluate the NASA reference feedline design and determine its thermal performance. 

Specn~ · 

• To evaluate the NASA reference feedline design from the standpoint of geysering. 

• To determine whether propellant conditions would be satisfactory for engine start. 

• To identify and evaluate thermal problems and technical costs arising from the NASA reference 

feedline design. 

4.0 Approach 

This study was accomplished by evaluating the reference system relative to geysering, heat up 

from ambient. natural circulation in the upper loop, and scr~en performance. 

( 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
.( 

5.0 Results 

The results of this study are attached. The primary results are listed below. 

6.0 Conclusions and Recommendations 

• Upper loop performance is satisfactory. Temperature rise less than 5 F. 

• Engine feedlines likely to saturate at engine. Geysering may occur. Most vapor will pass 

through screen if screen is not flat and horizontal. 

• Local pressure above saturation for engine start must be established by prepressurization. 

A 3700 lb. tank weight impact is estimated. 

• It is recommended that a prechill system be incorporated. See Task 3-P-019. 

7.0 Supporting Data 

• NASA-CR-64-3, "Mechanics of Geysering of Cryogenics," Martin-Marietta Aerospace 

Corp., 1964. · 

8.0 Attachments 

Study 3-P-018 "No LOX Bleed Performance Analysis." 

MARTIN MARIETTA 

MANNED SPACE ~''~TC:MS 

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The problem of liquid oxygen feedline and engine conditioning during prelaunch is composed 

of several parts, including loading, prevention of geysering when the feedline is full and when 

there is liquid oxygen in the propellant tank, predicting and/or promoting convection in the 

feedline, predicting the behavior of feedline screens (if used) , predicting the heat transfer to 

the LOX system within the engine, devising design solutions to provide adequate engine and 

feedline conditioning, and appraising and comparing the conceived solution for selection and adoption. 

NASA presented a reference configuration for the feedlines, shown on the facing page, which 

showed large, approximately 20 inch diameter feedlines which are interconnected at their 

lower ends by a line to permit circulation between the feedlines and to allow the LOX from 

both lines to be utilized in case of loss of one of the two sustainer engines. Each of these 

feedlines would feed three engines during boost for the 1-1/2 stage vehicle or two engines for 

the HLLV vehicle. For the 1-1/2 stage vehicle, four engines would be jettisoned at staging, two 

from each side. The HLL V vehicle would not jettison any of its four liquid propellant engines. 

facing 

3-P-018 

Page 1 

MARTIN MARIETTA 

( / 

MANNED SPACE SVC:TEMS 

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Upper loop 

It was logical to divide the feed system into two parts for analysis. The upper loop 

would comprise the tank, the two 20 inch feedlines and the crossover line. The other part 

would be each individual engine feedline. ·There were several analyses of the upper loop, 

one by NASA, and three by MMMSS. In Huntsville, the analysis was done by calculating the 

heat leak to each 20 inch feedline, assuming one inch of foam on one line and the other 

line bare. The foam insulated line would then act as a downcomer and the bare line as a riser. 

The loop flowrate was calculated to be 93 lb/sec for zero replenish, 100 lb/sec for the 

calculated replenish heat leak, and 145 lb/sec for 1 0 times the calculated replenish heat 

leak. The flow loop temperature rise was calculated to be 4.0, 4.6, and 9.3 deg F for the 

three cases. The calculation was also done by both Martin-Denver and Martin-Michoud. 

With different assumptions for the replenish flow and heat leaks for the tank and riser lines 

and engines using the Flow II model, all analyses show similar satisfactory system performance. 

facing 

3-P-018 

Page 2 

( 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

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3-P-018 

Page 2 

,. 

UPPER LOOP 

{Tank Bottom to Crossover) 

Good Performance. No Problem Found 

• Crossover Flowrate 85-1 00 lb/sec 

• Flow Loop Temperature Rise < 5 F 

• Variables: 

- Feedline Insulation 

1 inch on Downcomer 

None to 1 /2 inch on Riser 

- Riser and Downcomer Diameter 20 inches 

- Engine Heat Leak 2 to 1 0 btu/sec/engine 

- Topping/Replenish Flowrate 

0 to 35 lb/sec 

at 

163 to 180deg A 

- Crossover Diameter 6 - 12 inches 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

A plot from one of the Denver calculations is shown on the facing page. From these analyses 

it is concluded that the natural convection loop will condition the upper loop satisfactorily. 

facing 

3-P-018 

'Qe 3 

( 

MARTIN MARIETTA 

MANNEDSPACEsv~TEMS 

( 
Nominal L02 Feed System Temperatures 

Preliminary Temperature Profile in L02 Main Feedline 

·> ··~~9:··:·. ··.·· . . ..... · ' . . . 

· .)~~·~:::: t:'J.l~sif~iti~!19riif!ciw :~a!eM:. BSIJbi~fci:J•. ,1,~11~· (''c<•;;z: ''';< . . ··. . . .., . . .. · .. ·. • ... ··. 

' 166.5· .... : ... -.-::.:. =·:::·· 

Temp (Deg A .. I~~~ ' ... ·:· .. :·:·: 

165~5 

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Down Cotner Feed Line 

. Temperature 

Rise • A.J•f 

162.0 e-1 i ci_.U I I .4; ¢~--,.I I I I 4 I I I i o u; 1 1; & ;t I& ···I f f 0 . . • -. . ... ' • • ' a >. ' • • t.• ••" r • ' ' ' I 

10 lO · J~ 40 SO 60 70 00 90 100 110 1'20 

3-P-018 

Page 3 #'_ 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Feedline Conditioning 

The feedlines were analyzed separately. The length to diameter ratio of each feedline 

was, for the reference configuration, in the. range of 1 0 to 20. Since the reference 

configuration is not the final configuration. they may eventually have UD ratios even 

higher. The likelihood of stagnation, leading, possibly, to geysering may be evaluated 

by reference to the next page which shows the results of a large number of geysering 

tests done under contract to NASA by Martin Marietta. The indication from this figure is 

that geysering should not happen for a 12 inch line, however, for a smaller line or a longer 

line, it may well be in the geysering region. Calculations show that a 1 0 inch line 20 feet 

long would be in the geysering region. Also, it is considered mandatory to have a feedline 

screen similar to that of the Shuttle Orbiter. It is not known how this· screen will affect the 

convective pattern within the line. Attempts to model the convection within the line including 

screen effects have not yet been successful. 

facing 

3-P-018 

Page 4 

MARTIN MARIETTA 

MANNED SPACE SV~'TEMS 

( ( 
3·P·018 

Page4A 

( 

Feedlines Reference No Bleed 

•Rule of Thumb 

UD < 1 o Convection OK 

UD > 20 Convection Not Enough (Geyser Region) 

•Probable Design Between These or Even UD > 20 

•Screen In Line Will Inhibit Convection - Pores in 

Shuttle-Type Screen Will Be "Stable" 

•Most Vapor Will Pass Through and Rise Through 

Feedline if Screen is not Flat, Horizontal 

•Feedline may be Saturated or not have Repeatable 

Peformance 

•Analytical Models have not been Confirmed 

•Helium Bubbling, If Used, Will Not Result In 

Evaporative Cooling at Engine Inlet (Assumes 

Ambient Helium) 

•Possible Design Solutions 

On Board Bleed 

Overboard Bleed 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
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3-P-018 

Page 6 

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, 

Assumptions 

• MFL Flow Loop Temperature= 165 Deg R 

• EFL Heat Leak Same as Orbiter Feedlines With SOFI 

• No Flow in Engine Feedlines 

Discussion of Math Model Predictions 

• Engine Inlet Temperatures Strongly Influenced by Convection Coefficient 

• "Best" Convection Coefficient (Based on Best Match Between Predicted and Measured 

Temperatures for Orbiter Feedlines During 1978 Denver EPL Tests) 

• Convection Coefficient of 1 0 or 15 Gives Best Match When Flow and/or Helium Inject 

Flow Present · 

• Convection Coefficient of 5 Gives Best Match W~en No Flow and Helium Inject Present 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
... ,J ... 

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( 

3-P-018 

Page 7 

ENGttJE Hl:A 1 tEAK . 

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Convection Coefficient in Engine FDL 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
( 

No lOX Bleed 

Summary & Conclusions 

• Upper loop performance is satisfactory- Temperature rise less than 5 F. for 

• 20 inch main feedlines. 

• 1 inch SOFI on downcomer. 

• Zero to 1/2 inch SOFI on riser. 

• 6 to 12 inch crossover duct diameter. 

• Zero to 35 lb/sec topping and replenish at 163 to 180°A. at local pressure. 

• Engine feedlines likely to saturate at engine. 

• Geysering may occur. 

• Ambient helium bubbling will mitigate geysering effects, but will not cool LOX locally. 

• Most vapor will pass through screen unless screen is flat and horizontal. 

• Local pressure above saturation for engine start must be established by pre pressurization. 

3700 lb. tank weight impact estimated (25 psi higher tank pressure than with cold LOX). 

3-P-018 

Page 8 MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
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( 

Executive Summary 

Task 3-P-019 "LOX Bleed Trade Study" of the National Launch System Phase B study 

done by MMMSS under the Shuttle C Contract reads as follows, "Trade study to consider 

bleed vs. no bleed LOX system considering, at a minimum, operability, complexity, start 

sequence restrictions with no bleed, available propulsion module space, and tank stretch 

limits." This report is based upon the Marshall Space Flight Center study plan dated 

August 5, 1991, and presented at the Technical Interchange Meeting at Marshall Space 

Flight Center on August 28, 30, 1991, by Danny Davis, the cognizant Panel Chairman. 

Because of the difficulty in modeling the liquid heating, it was necessary to consider the 

total subcooling of the liquid necessary to start the engine to come from the subcooling 

accomplished by the prepressurization of the tank. This was estimated to be 50 psig in 

an analysis submitted to NASA by MMMSS and an analysis presented to the Chief Engineer 

by the Propulsion Team. Therefore, this value was used as a basis of comparison for 

subsequent subsystem concepts. 

Several subsystem concepts were cons.idered for feedline and pump conditioning. Some 

of the concepts have characteristics that are obviously more desirable than others. 

• Effect on LOX tank design pressure 

• Predictability 

• Repeatability, engine test to vehicle 

• Precedence 

• Impact on engine design 

• Impact on engine test 

• Potential for required future change 

• Operational efficiency 

• Hazard introduced 

• Hardware complexity 

The evaluation of candidate subsystems is summarized as follows. MARTIN MARIETTA 

MANNEDSPACEsv-TEMS 

( ( 
( -·· 

• Reference No Bleed System cannot be expected to have subcooled propellant at engine 

inlets at start of prepressurization. . 

• Reference No Bleed System causes tank prepressurization pressure increase. 

A 20 psi prepressurization increase to 50 psig results in a 3700 lb. tank weight impact. 

• Warm up (vapor pressure increase) after prepressurization is very slow- approximately 0.4 psi/min. 

• The Onboard Bleed looks viable and eliminates the penalty due to prepressurization requirements. 

• The Overboard Bleed to the facility provides good performance but at the cost of increased complexity. 

• The Overboard Bleed Through the Engine to the atmosphere provides good conditions and is 

simple, only a bleed valve and a line to the nozzle exit are required. 

• The LOX dump appears to unduly burden the engine development program unless the engine 

is designed for LOX lead start. 

f 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
1.0 Summary 

Task Number 3-P-019 

LOX Bleed Trade Study 

The reference "No Bleed" system was compared with four alternative LOX Bleed Systems. 

All would require an engine bleed valve, and all would allow a reduction in LOX tank 

prepressurization pressure and the associated 3700 lb. payload improvement compared to 

the "No Bleed" case. 

2.0 Problem 

"Made study to consider bleed vs. no bleed LOX system considering, at minimum, operability, 

complexity, start sequence restrictions with no bleed, available propulsion module space, and 

tank stretch limits." 

3.0 Objective 

General 

To identify and evaluate alternate LOX bleed systems vs. the reference no bleed system . 

. Specific 

To identify and evaluate alternate LOX bleed systems and determine their potential performance 

advantages as compared to a reference no bleed system considering the important attributes of each. 

( 

4.0 Approach 

The approach adopted in performi119 this study was to consider and analyze LOX bleed systems 

that had previously been used and that were suggested, identify a set of attributes by which the 

systems could be compared, and compare the systems with the reference no bleed system. 

MARTIN MARIETTA 

MANNED SPACE SVC:TEMS 

( ( 
( 

5.0 Results 

The results of the study are attached. The. primary results are listed below. 

6.0 Conclusions and Recommendations 

• Reference No Bleed System cannot be expected to have subcooled propellant at engine 

inlets at start of prepressurization. 

• Reference No Bleed System causes tank prepressurization pressure increase. A 20 psi 

prepressurization increase to 50 psig results in a 3700 lb. tank weight impact. 

• Warm up (vapor pressure increase) after prepressurization is very slowapproximately 0.4 psi/min. 

• The Onboard Bleed looks viable and eliminates the penalty due to prepressurization 

requirements. 

• The Overboard Bleed to the facility provides good performance but at the cost of 

increased complexity. 

• The Overboard Bleed Through the Engine to the atmosphere provides good conditions 

and is simple, only a bleed valve and a line to the nozzle exit are required. 

• The LOX dump appears to unduly burden the engine development program unless the 

engine is designed for LOX lead start. 

7.0 Supporting Data 

Task Number 3-P-018 "No LOX Bleed Peformance Analysis." 

8.0 Attachments 

Task Number 3-P-019 "LOX Bleed Trade Study." 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
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( 

The reference No-Bleed system is evaluated in Task Number 3-P-018. 

Because of difficulty in modeling the liquid heating in the feedline and because of the 

potential of having saturated liquid at the pump inlet at the time of prepressurization, it 

was necessary to consider the total subcooling of the liquid necessary to start the engine 

to come from the subcooling accomplished by the prepressurization of the tank. This 

was estimated to be 50 psig in an analysis submitted to NASA by MMMSS and an 

analysis presented to the Chief Engineer by the Propulsion Team. Therefore, this value 

was used as a basis of comparison for subsequent subsystem concepts which permit 

calculation of LOX temperature at the pump inlet and within the pump. A 50 psig 

prepressurization would result in a tank design pressure increase, compared to the 

reference tank design, of 20 psi. This would result in a 3700 lb. tank weight penalty. 

facing 

3-P-019 

Page 1 

MARTIN MARIETTA 

MANNEDSPACESV~TEMS 

( ( 
( 

lJ-. 

3-P-019 

Page 1 

Feedlines Reference No Bleed 

•Rule of Thumb 

UD < 1 0 Convection OK 

UD > 20 Convection Not Enough (Geyser Region) 

•Probable Design Between These or Even UD > 20 

•Screen In Line Will Inhibit Convection - Pores In 

Shuttle-Type Screen Will Be "Stable" 

•Most Vapor Will Pass Through and Rise Through 

F eedline if Screen is not Flat, Horizontal 

•Feedline may be Saturated or not have Repeatable 

Performance 

•Analytical Models have not been Confirmed 

•Helium Bubbling, If Used, Will Not Result In 

Evaporative Cooling at Engine Inlet (Assumes 

Ambient Helium) 

·Possible Design Solutions 

On Board Bleed 

Overboard Bleed 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

 

Several subsystem concepts were considered for feedline and pump conditioning. 

A specific recommendation will not be made at this time and in this phase of the design 

study because of the preliminary nature of the study, although some of the concepts have 

characteristics that are obviously more desirable than those of others. The attributes that 

were considered in evaluating the subsystem concepts are shown on the facing page. 

The evaluation of candidate subsystems against these attributes is summarized later. 

facing 

3-P-019 

Page2 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
3-P-019 

Page 2 4 / . 

Attributes Considered in Evaluating Systems 

• Effect on LOX Tank Design Pressure 

• Predictability 

• Repeatability, Engine Test to Vehicle 

• Precedence 

• Impact on Engine Design 

• Impact on Engine Test 

• Potential for Required Future Change 

• Operational Efficiency 

• Hazard Introduced 

• Hardware Complexity 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

The On-Board Bleed 

The first concept considered was the On-Board Bleed, which has no ground interface 

and does not vent to the atmosphere. A LOX flowrate of 0.851b/sec. in the return line 

was calculated for a 2 Btu/sec engine heat leak. This would allow the "hot" LOX to be 

carried up the return line and the LOX flowing down the feedline would be heated to 

171 deg A at the pump inlet and 180 deg R in the pump. This would provide a net positive 

pressure of 72 psi at the pump inlet and 50 psi in the pump for engine start. This concept 

would eliminate prepressurization as a LOX tank design factor, because the maximum 

tank bottom pressure in flight is only 14 psig lower than during prelaunch with a 50 psig 

prepressurization ullage pressure. The bleed was assumed to originate in the gas 

generator supply line downstream of the LOX pump. 

The predictability of the On-Board Bleed was considered "fair" because it has not been 

used recently, and it does not have a mechanical pump. The analysis is straightforward. 

The repeatability, engine test to vehicle was considered good, because the configuration 

would not be hard to duplicate. The Saturn IB precedent was virtually identical to this 

application. 

The impact on engine design and test is expected to be moderate. An engine bleed 

valve would be required. 

The potential that a future change would be required is considered low, however the 

available bleed flow rate is limited by the head available for natural convection. · 

Therefore, the only way to increase the flow rate is to enlarge the bleed line. Also, if the 

vehicle boattail is shortened to allow a longer hydrogen tank, the bleed line head would 

be reduced, reducing the bleed line flow rate. 

We consider the operational efficiency of this concept to be good because it has no 

ground interface, although it does add a bleed valve to each engine and this bleed valve might fail. 

facing 

3-P-019 

Page 3a 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS / 

( ( 
The hazard introduced would only be that one of the small lines might fail and create a 

LOX leak. The hardware required for this scheme is not complex. only a bleed valve and set of small 

lines and brackets for each engine is required. 

( 

facing 

3-P-019 

Page 3b 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
.. 

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3-P-019 

Page 3 (} 

Onboard LOX Bleed 

•Onboard Bleed Flows 0.91b/sec 

•Eliminates Stagnation in Feedline and STME 

•Provides Subcooled LOX at STME Inlet 

and in LOX Pump 

171°R at Inlet 72 psi above Vapor Pressure 

180°R in pump 59 psi above Vapor Pressure 

•Provides Repeatable Engine Operation Conditions 

from test Stand to Vehicle 

•Simple Design (Similar to S-1 Stage) 

Bleed Valve (Similar to SSME) Near GG Inlet 

One Inch ID Line with 12ft. Head 

•No Ground Interfaces · 

•No Separation Interfaces 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

The Overboard Bleed to the Facility 

The overboard bleed to the facility is quite similar to the on-board bleed in its performance 

except that the total head of the LOX above the engines can be used to drive the flow, 

rather than only the convective loop created by the feedline and the warmer return line. 

The flow was calculated for a system comprising 1/2 inch lines manifolded to a 1-1/2 inch 

line which carries the flow to a ground disconnect, thus overboard to the facility. This is 

very similar to the Shuttle LOX bleed and is predictable, repeatable, easy to duplicate on 

a single engine test stand, and robust from the standpoint that even with small lines a high 

flow rate (2.2 lb/sec) can be obtained. The predicted engine inlet temperatures are actually 

three degrees F lower than with the on-board bleed, even with the 1/2 inch lines which are 

half the diameter of the lines considered for the on-board bleed. For these reasons, the 

potential for a required future change was considered low. The operational efficiency was 

considered only fair because of the necessity of the ground interface and separate system 

to dispose of the oxygen bled to the ground. The hazard introduced was the same as the 

on-board bleed with the addition of the potential failure of the ground disconnect and potential 

leakage or failure of the on-board disconnect which is provided for booster separation. The 

hardware is more complex than the on-board bleed because of the in-flight disconnect, and 

because of the ground disconnect. Also, the 1-1/2 inch collector line and the ground 

disconnect add to the complexity. 

facing 

3-P-019 

Page 4 

MARTIN MARIETTA 

MANNED SPACE SY~TEMS 

( ( 
3-P-019 

Page 4 A' 

( 

Overboard LOX Bleed 

•Eliminates Stagnation In Feedline and STME 

•Provides Subcooled LOX At STME Inlet and 

In LOX Pump 

168°A at Inlet 

171°R In Pump 

•Provides Repeatable Engine Operating 

Conditions From Test Stand To Vehicle 

•Bleed Valve Required Similar To SSME 

•1/2 Inch Diameter Line Manifolded to 1 1/2 .. 

Disconnect 

•Adds Ground Interface (1 1/2 Inch Separation 

Disconnect) 

•Adds Two 1/2 Inch lnflight Disconnects 

( 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

The Overboard Bleed to the Atmosphere 

The overboard bleed to the atmosphere, is suggested by the performance of the on-board 

bleed and the fact that the SSME discharges 0.5 to 21b/sec through its LOX bleed line to 

the atmosphere during prelaunch. This indicates that a flowrate equal to that of the on-board 

bleed (0.85 lb/sec) could safely be discharged to the atmosphere at the engine exit plane 

which is the location at which the SSME LOX pump seal bleed is discharged. If the gas 

generator LOX supply line is not the critical location for engine chill. it is possible that the 

engine labyrinth seal leak (if a labyrinth seal is used) would be all the bleed that is necessary. 

The performance would be identical to the on-board bleed to the level of detail the analysis 

has been done, and an engine thermal model would be necessary to tell the difference. 

This would give an engine inlet steady state LOX temperature of 171 deg R. Like the other 

bleed concepts, this would remove the high prepressurization requirement that would be 

required if the feedline is saturated. The concept is predictable since the flow is steady. 

The repeatability, engine test to vehicle, would be excellent if the feedlines had similar 

heat leaks. If they did not, a means would have to be found to establish the same kind 

of a temperature gradient in the feedline, as expected for the flight vehicle. This method, 

except for the Shuttle experience cited above, has no known precedent. Again this method 

would require an engine bleed valve. There would be virtually no impact on engine test. 

This concept appears to be robust, and the possibility of requiring a future change appears 

to be low. The operational efficiency appears to be good. The hazard introduced by this 

concept is considered to be minimal. Hardware complexity is not a problem, since the 

concept requires only a bleed valve and a small drain line. If the engine pump seal leak 

can satisfy the need for the bleed. only the drain line would be needed. 

facing 

3-P-019 

Page 5 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
3-P-019 

Page 5 

Lox Bleed Valve 

{ 

....... 

Overboard Bleed to Atmosphere 

•LOX Bleed Overboard to Atmosphere at Engine: 

-Performance same as Onboard LOX Bleed for 

0.91b/sec Bleed from GG LOX Supply Line 

-SSME LOX Pump Intermediate Seal Currently 

Leaks 0.5 to 21b/sec of LOX with no known 

Ill Effects 

-Bleed Valve (similar to SSME) Required 

-May require adding Bleed Line to Engine Exit 

-No Interfaces, no Disconnects 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Start With LOX Lead 

Starting with a LOX lead did not appear to be a favorable concept. It does not appear 

realistic to dump several hundred pounds of LOX during start and follow it with partially 

burned combustion products and it does not appear favorable to require the engine to 

traverse the range of mixture ratios from LOX rich to its normal operating point during 

start. It has been suggested that a LOX lead could be followed by a purge, and then the 

engine started normally, and this might be satisfactory, but it would impose an additional 

development requirement on the engine. Neither method looks to be predictable. The 

RL-10 precedent does not look applicable because the RL-1 o cycle is completely different 

from the one selected for the STME. The engine test program would have to explore 

a range of inlet conditions to be expected in use, adding a number of tests to the 

development program. The method, because it does not look predictable, would have 

a tendency to have to be changed. The feedline temperature at engine start has not yet 

been predictable, and temperature interlocks may be required to assure a satisfactory 

start. There are so many unknowns about this method, that it appears to be impossible 

to evaluate the associated hazard. The advantage is that it does not add any. components. 

facing 

3-P-019 

Page 6 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
3-P-019 

Page 6 

Lox Bleed Valve 

( 

Start with lOX lead 

•Engine Start with LOX Lead: 

-Places Constraint on Engine Sequence Not 

Connected with Starting the Engine Itself 

-Forces Engine to Traverse Mixture Ratio 

Range from LOX Rich to Fuel Rich 

-Potential Increase in Engine Testing Required 

as Compared to Bleed 

~ . 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Summary 

The upper flow loop is considered to be a viable anti-geyser and propellant conditioning 

approach. 

The reference No-Bleed system is not predictable, so far, and may result in saturated 

LOX in the engine inlet and in the engine. Any attempt to make such a prediction is 

complicated by the assumed requirement for a screen in the feedline. To predict the 

No-Bleed system performance, an analytical model and test data to verify it are needed. 

Since this is the case, the tank must be designed to accommodate a high prepressurization 

requirement, which is expected to cause a higher tank bottom pressure than encountered 

in flight. Also, because of this unpredictability, it appears that the engine test program would 

be complicated by the need to evaluate a wide range of inlet conditions and hold time. Also, 

thermal constraints that might require real time evaluation may come out of such an engine 

test program. 

Of the bleed systems, the bleed to the atmosphere is the simplest, and the pump seal 

leakage may provide all or most of the necessary flowrate. The On-Board Bleed is viable, 

with an addition of complexity. The Overboard Bleed to the facility is low-risk, and may 

even be considered standard. since it is used on the Shuttle, but it does add operational 

and hardware complexity. - _ 

The idea was brought up to start with a LOX lead to expel any hot LOX. Since this would 

impose an additional development requirement on the engine and seemed questionable 

from a safety and engine operability standpoint, it is not recommended. 

3-P-019 

Page 7 

MARTIN MARIETTA 

MANNED SPACE SY~TEMS 

( ( 
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Evaluation Matrices 

The two evaluation matrices on the following pages summarize the above 

discussion of the candidate subsystems . 

facing 

3-P-019 

Page 8 

,, ( 

MARTIN MARIETTA 

MANNEDSPACESY~T~MS 

( 
( 

LOX FEEDLINE CONDITIONING 

Reference On-Board Overboard Overboard 

Attributes Start after LOX Bleed Bleed to No Bleed Bleed dump thru MOX To Facility Atmosohere 

•Effect on LOX Tank •Prepress to 50 psi In- •Deletes Prepress •Deletes Prepress •Deletes Prepress . •Prepress to 50 pslg. 

Design Pressure creases Tank Design Aeqt. Penalty Aeqt. Penalty Reqt. Penalty Increases UUage & 

Press. 3700 lb weight Tank Bottom Design 

Impact Pressure 

• Predictability •Poor •Fair •Excellent •Good •Excellent 

•Repeatability Eng. •Poor •Poor •Very Good •Good •Very Good 

Test to Vehicle 

•Precedence •None •Saturn 1B •Shuttle •Shuttle Shows •None 

S-1 Stage Safety, Not 

Performance 

•Impact on Engine •None •Adds LOX BV •Adds LOX BV •Adds LOX BV and ~auses MR Traverse 

Design small line from High to Normal 

•Impact on Eng Test •None •low •Very little •low •Potentially large 

•Potential for Req'd •large •low (limited Bleed •Very low (High ·low (limited Bleed •large 

Future Change Rate) Bleed Rate) Rate) 

•Operational •Good •Good (Pending •Fair •Good (Pending Test) •Unknown 

Efficiency Test) 

•Hazard Introduced •None •low •low •low •Potentially Severe 

•Hardware •low •LOX BV's & Small •LOX BV's Small •LOX BV's and •low 

Complexity lines Req'd lines & Onbd & Grd Drainlines Req'd 

Disconnects Req'd 

3-P-019 JJl~J:~I/11·J~T.!J:Il::tj ,~. 

Page 8 MANNED SPACE SYSTEMS 
( 

 

Attributes 

Effect on 

LOX Tank 

Precedence 

Impact on 

Eng. D•lgn 

Impact on 

Eng. Teat 

Potential tor 

3 • lmponenl 

2 • Somewtull lmportenl 

1 • Considered 

Weighting 

Factor 

3 

-

2 

-

3 

-

2 

-

1 

-

1 

-

Req'd Future 2 

~biDUI - Operational 

Efficiency 3 

- - Hazard 

Introduced 3 

-- - Hardware 1 

Gredes/Scores 

Reference 

No Bleed 

Complexity 

--

Total r:== 36 .. 

3-P-019 

Page 9 

LOX FEEOUNE 

CONDITIONING 

EVALUATION MATRIX Gredes • Aa4, 8:3, C•2, Da1 FaO 

Score • Grede • Welghl 

On Board 

Bleed 

Overboard Bleed I Overboard Bleed I Start after LOX 

to Facility to Atmosphere dump thru MOV 

v 

,~ r62 .. 

44 rss I 32 

MARTIN MARIETTA 

MANNED SPACE SVt'TEMS 

( ( 
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Cost Estimates 

The hardware cost estimates on the facing page are for the bleed systems described. The option alphanumerics are those assigned by NASA. 

( 

facing 

3-P-019 

Page 10 

( 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( 

'• 
( 

Item 

Engine Bleed Valve 

Line 

Elbow 

Line 

Elbow 

Line 

Elbow 

Inlet Fitting 

Total -

3-P-019 

Page 10 

I 

( 

L02 Options 

Option A-1 - On-Board L02 Bleed 

Description Qty/PM 

Similar to SSME (1 Per Engine) 6 

1.125" ID X 9.5' Line W/3' Flex 6 

1.125" ID 75°, R/d >= 2.5 6 

1.125" ID X 1 0.5' Line W/3'Fiex 6 

1.125" 10 75°, R/d >= 2.5 6 

1.125" ID X 2.5' Line W/3' Flex 6 

1.125" 10 75°, R/d >= 2.5 6 

Connection to Lower F/L Discon 6 

Cost/Unit Cost/PM 

81,613 489,678 

110 713 

23 138 

131 788 

23 138 

31 188 

23 138 ' 

150 900 I 

I 

I 

492,681 I 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
Item 

Engine Bleed Valve 

Line 

Tee 

Disconnect 

Elbow 

( 

Line 

Elbow 

Line 

Manifold 

line 

Elbow 

Line 

Disconnect 

Total 

3-P-019 

Page 11 

Option C-1 - Overboard L02 Bleed 

Description Oty/PM Cost/Uni Cost/PM 

Similar to SSME (1 Per Engine) 6 81,613 489,678 

0.5" ID X 9.5' Line W/3' Flex 6 119 713 

0.5" ID 6 50 300 

0.5" ID 1 26,000 26,000 

0.5" 10 90°, R/d >= 2.5 6 2 13 

0.5" ID X 1' 6 2 13 

0.5" ID 30°, R/d >= 2.5 6 1 4 

0.5" 10 X 12' 6 26 155 

0.5" To 1.5" ID 1 10,000 1·0,000 

1.5" 10 X 12' 1 234 234 

1.5" ID R/d >=2.5 1 57 57 

1.5" ID X 2' 1 39 39 

1.5" 10 1 26,000 26,000 

. 553,206 --- -~--

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
L.l -.. 

Engine Bleed Valve 

Line 

Total 

3-P-019 

Page 12 

( 

Option D- Overboard Bleed To Atmosphere 

(This is engine mounted hardware) 

Similar to SSME (1 Per Engine) 6 81,613 489,678 

0.5" 10 X 12.5' (Insulated) 6 27 161 

489,839 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

3-P-019 

Page 13 

Task 3-P-019 

Addendum 

Weight Estimates 

• On-Board LOX Bleed 

(Option A-1) 

• Overboard LOX Bleed 

(Option C-1) 

• Overboard Bleed to Atmosphere 

(Option D (Engine Mounted Hardware)) 

( 

lb. 

107 

68 

56 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( 
Summary & Conclusions 

• Reference No Bleed System cannot be expected to have subcooled propellant 

at engine inlets at start of prepressurization. 

• Reference No Bleed System causes tank prepressurization pressure increase. 

A 20 psi prepressurization increase to 50 psig results in a 3700 lb. tank weight impact. 

• Warm up (vapor pressure increase) after prepressurization is very slowapproximately 0.4 psi/min. 

• The Onboard Bleed looks viable and eliminates the penalty due to prepressurization 

requirements. 

• The Overboard Bleed to the facility provides good performance but at the cost 

of increased complexity. 

• The Overboard Bleed Through the Engine to the atmosphere provides good 

conditions and is simple, only a bleed valve and a line to the nozzle exit are required. 

• The LOX dump appears to unduly burden the engine development program unless 

the engine is designed for LOX lead start. 

3-P-019 

Page 14 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
THIS PAGE INTENTIONALLY LEFT BLANK 

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Task Number 3-P-025 

L02 Tank Pressure Limits 

Prepared By: 

Tom Winstead 

20 Dec, 1991 

( 

Approved By: 

Z. Kirkland 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

 

Executive Summary 

NASA Statement of Work: 

"Establish L02 tank pressure limits vs. flight time considering engine start, 

shutdown and NPSP requirements, potential pressure stabilization of tank during 

max airloads, structural weight considering proof test requirements ~nd performance." 

• With no-bleed L02 system, prepressurization will determine tank structural 

requirements. Current estimate of tank impact -4500 Ibm. For no impact on tank, 

prepress needs to be reduced to <30 psig. 

• Vent valve for baseline will be sized by prelaunch operations and will have no 

influence on flight. 

• Optimum NPSP at MECO is 30.8 psi at an ullage pressure of 20.0 psig. 

• Proposed system would have a prepressurization band of 30-32 psig with relief 

set at 34 psig. Structural impact of -500 Ibm is largely offset by a reduction in residuals 

MARTIN MARIETTA 

MANNED SPACE SY~TCMS 

( ( 
1.0 Summary 

( 

Task Number 3-P-025 

L02 Tank Pressure Limits 

The study has shown that the upper pressure limit will be determined by prelaunch 

operations. This results in a 4500 Ibm increase in structural weight. This impact can 

be eliminated by using a bleed conditioning system and reducing the pre-pressurization 

level to less than 30 psig. Lower limit is determined by the saturation pressure of the 

liquid up to terminal drain when the engine NPSP requirement ot30.8 psi becomes 

important. The optimum tank pressure is -20 psig which allows the autogenous 

pressurization flowrate to be reduced from the reference 3.0 Ibm/sec to 2.51bm/sec. 

2.0 Problem 

Determine L02 tank pressure limits for the reference configuration. 

3.0 Objective 

Determine tank and system impacts for the reference configuration. 

4.0 Approach 

The approach to performing this study was: 

• Determine tank pressure vs. time for baseline trajectories. 

• Use inputs from 3-P-018, 3-P-019, 3-P-017 and 3-S-01 OA to determine system impacts. 

.. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

5.0 Results 

The results of this study are attached. The primary results of the study are listed below. 

6.0 Conclusions and Recommendations 

The upper limit will be sized for pre-launch operations. 

Current autogenous flowrate can be reduced to lower tank pressure at MECO. 

Insulated LOX tank. 

Helium Inject. 

Recommend 30-32 prepress band with minimum relief at 34 psig. 

1.0 Supporting Data 

8.0 Attachments 

Study "Task Number 3-P-025, L02 Tank Pressure Limits," dated 12120/91. 

MARTIN MARIETTA 

MANNED SPACE SV~TEMS 

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Approach 

Generate baseline ullage pressure fqr HLLV and 1.5 Stage 

Generate issues with system and structural impact 

Trade residuals with engine NPSP requirements and engine cost sensitivities 

3-P-025 

Page 1 

( 

·MARTIN MARIETTA 

MANNED SPACE SY~T~MS 

( 
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The STME cost and weight impacts associated with NPSP variations are from 

the STPT consortium. The data are contained in NLS Data Book Log #35 in 

response to an action item from the propulsion TIM #1. 

facing 

3-P-025 

Page 2 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
3-P-025 

Page 2 

Assumptions 

• Analysis based on reference layout 

• Engine out trajectory used for MECO conditions 

• STME influences taken from NM0-090-20 

• L02 tank structural impact derived from 3-S-01 OA Trade . 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

The LOX Tank Trade Study 3-S-1 OA was done predominantly above 30 psig according 

to requirements for that study. The facing figure has been extended to 1 0 psig, where 

the stringer weights will start increasing. No analysis has been done below that pressure. 

The LOX tank skin will go into compression at about 25 psig, and some structural 

designers prefer to not operate the tank in compression. 

The raw material cost differential is estimated at about $350/lb. 

facing 

3-P-025 

Page 3 

MARTIN MARIETTA 

MANNED SPACE S"'-""'EMS 

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A typical LOX tank ullage pressure profile is shown for selected 1-1/2 stage and 

HLLV trajectories for the cycle-a baseline pressurization system (700 deg A 

autogenous, fixed orifice flow control, and no engine bleed). Note that the pressure near booster 

separation is about 8 psig higher than at cutoff. While the NPSP .requirement is 

approached only as the feedline is drained, the tank pressure at cutoff will generally 

have a minimal effect on structural design; the structure is more than likely designed 

by tank bottom pressure near booster cutoff when this pressure and other loads are 

converted to tank proof pressure requirements. 

If the no-engine-bleed option was retained, prepressurization and liftoff loads would 

probably design the tank. The weight impacts are unacceptable and retention of this 

option is not recommended. · 

facing 

3-P-025 

Page 4 

MARTIN MARIETTA 

MANNEDSPACESV~TEMS 

( ( 
( ( 

Baseline Ullage Pressure 

Minimum Relief Pressure 

ss. ~-----------------------------------------------+ 

..-.. t5-

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3-P-025 

Page 4 

0-0 

3.0 Ibm/sec/eng 

700°R Diffuser Inlet Temd>erature 

100. 200- ~00. 

Time for 11ftoff (sees) 

HLLV Nominal 

~ · / 1.sstage ·.....--

-

400. 

R~duce to Ullage 

Pressure Required 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

These curves show the payload sensitivity using the "baseline" NPSP of 

30.8 psia, evaluating residuals for cutoff from 70 and 1 00°/o power levels. 

The predominant factors differentiating these curves is the reduction in unusable 

propellant mass in the feedlines. The gain in payload is significant, and warrants 

a recommendation to plan to operate the engines at 70°/o for a predetermined period 

prior to cutoff. 

While the feed system layout and performance values are preliminary, this trend will apply 

to all configurations, although absolute values will vary. 

facing 

3-P-025 

Page 5 

MARTIN MARIETTA 

MANNED SPACE SY~T-:MS 

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Extending the analysis to parametric NPSP requirements shows that the baseline 

NPSP is probably near a minimum, but that there is a small influence on the system 

payload capability over a fairly wide range o.f N PS P and tank pressures. 

As tank pressure at MECO is increased, the structural weight increases, but residuals 

for an engine out are decreased so that the net payload impact is small. 

Elements considered here are residuals at a vapor pressure of 20.2 psig, ullage weight at 

MECO, and structural weights corresponding to conditions near booster cutoff, pump weights 

corresponding to the NPSP and cutoff from 70°/o power level. 

facing 

3-P-025 

Page6 

MARTIN MARIETTA 

MANNED SPACE SY~TEMS 

( ( 
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·-. 

Payload Impact at 70o/o Power Level 

14000 .,..-----------------, 

NPSP at 1 OOo/o pbwer level 

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~~"'~"' 4000~------~--~----~--~--~ 

3-P-025 

Page 6 

10 20 30 40 

Ullage Pressure @ MECO (psig) 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

The liquid vapor pressure at MECO has been assumed to be a constant 20.2 psia for 

all of the above trades. As minimum tank pressures are approached, there is an 

increasing gain for reduced vapor pressure. 

The 16.2 psia condition corresponds to Shuttle ET LOX vapor pressure, which 

results from a well insulated tank with a significant period of helium injection prior 

to launch which leaves the propellant in a well established recirculation mode 

throughout flight. 

The 20.2 psia condition corresponds to an uninsulated LOX tank without 

helium injection with the tank pressurized above 22 psia. 

facing 

3-P-025 

Page 7 

MARTIN MARIETTA 

MANNEDSPACESV~TEMS 

( ( 
3-P-025 

Page 7 

( 

Payload Sensitivity due to Liquid Vapor Pressure 

8000 -r-------------, 

7000 

- 6000 j Vapor Pressure 

E 

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C) 5000 ·-CD 

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4000 

3000 +---,.----,----.---r----.--i 

10 20 30 40 

Ullage Pressure @ MECO (psig) 

"r 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

 

The tank pressures and weight impacts associated with the reference design and the resulting 

requirement for a high prepressurization level corresponding to the no bleed engine start. The 

proposed minimum relief pressure and minimum tank pressure reduce the tank weight penalty, 

(compared with a tank operating at Shuttle ET pressure, from 4500 to 500 lbs.). 

facing 

3-P-025 

Page 8 ( 

MARTIN MARIETTA 

MANNED SPACE SV"''TEMS 

( 
3-P-025 

Page 8 

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12000 

10000 

6000 

6000 

4000 

2000 

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L02 Tank Structural Impact 

1--· 

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10 20 30 40 50 60 70 80 90 

PRESSURE (PSIG) 

' ... 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

'•. 
( 

The proposed LOX tank pressure ·band of relief valve set at 34 psig is a companion to fixed 

orifice autogenous pressurization, engine conditioning of some type, and system optimizing 

near 20 psig@ MECO. 

The nominal predictions on tank pressure are oversimplified by ground rules, and will experience 

some additional variations as components and STME operation are further defined. 

If a flow control system were incorporated, the maximum tank pressure could be reduced, and 

would provide performance improvements. 

facing 

3-P-025 

Page 9 

( 

MARTIN MARIETTA 

MANNED SPACE SY~TEMS 

( 
Proposed Ullage Pressure 

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3-P-025 

Page 9 

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Minimum Relief Pressure 

/ 

HLLV Engine Out 

100. 200. 380. 

T1me from 1 iftoff (sees) 

( 

1 .. 5 Stage, Nominal 

~ 1.5 gtage, Engine 

; ut 

400. 

Required 

tank 

pressure 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
" 

.( 

When system payload impacts are evaluated: 

• The tank pressure effect on weight is evaluated at BECO (~P = 8 psig 

higher than at MECO) 

• Liquid residuals are evaluated at the time when NPSP is no longer 

satisfied 

• Gas residuals are those required for tank pressurization 

Elements which would further increase payload are: 

• Vapor pressure less than 20.2 (page 7) 

• Tank pressurant flow-control (~P < 8 psig) 

facing 

3-P-025 

Page 10 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
3-P-025 

Page 10 

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E 

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Payload Impacts at Proposed Tank Pressure 

10000 I 

npsp=J0.8@11 OO%Power level 

9000 1 Psat=20 2psla 

8000 

7000 

6000 

5000 

4000 

3000 

2000 

1000 

structural 

impact 

0 -~ < I I 1->41 I I I 

10 20 30 40 so 

Ullage pressure@ MECO (pslg) 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Summary & Conclusions 

• With no-bleed L02 system, prepressurization will determine tank structural 

requirements. Current estimate of tank impact -4500 Ibm.: 

• Vent valve for baseline will be sized by prelaunch operations and will have no 

influence on flight · 

• Optimum NPSP at MECO is 30.8 psi at an ullage pressure of 20.0 psig 

• Proposed system would have a prepressurization band of 30-32 psig with relief 

set at 34 psig. The structural impact of -500 Ibm will be offset by lower residuals 

at MECO. 

• An intelligent tank pressure control system would result in a payload increase, 

but requires structural analysis. 

• Residuals are significantly lower for an insulated/helium injected tank than for an 

uninsulated tank. 

3-P-025 

Page 11 

MARTIN MARIETTA 

MANNEDSPACESY~TEMS 

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Executive Summary 

NASA Statement of Work: 

"Select optimum l02 tank helium pressurization system based on tank pressure 

limits and specified reference trajectories and considering safety, reliability, operability, 

simplicity, weight, including residuals and cost." 

1.5 Stage 

• Minimum pressurization system weight is achieved using cryogenic storage helium. 

• Ambient storage helium is the next best with fixed orifice autogenous being better at 

higher HEX temperatures. 

HLLV 

• Minimum pressurization system weight is achieved using cryogenic storage helium. 

• Fixed orifice autogenous weight performance is better at higher HEX temperatures. 

• Ambient helium system assumes no bottle staging and consequently will result in 

significant weight impact. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
Task Number 3-P-026 

LOX Tank Pressurization System Using Helium 

1.0 Summary 

A trade study was performed to evaluate LOX tank pressurization with ambient and 

cryogenic helium systems. A rough order of magnitude study for the autogenous 

pressurization system was done for comparison. Both ambient and cryogenic helium 

pressurization systems are lighter than an autogenous system, with the difference 

reducing as heat-exchanger outlet temperature is increased. The subsystem costs 

are significantly higher for the helium pressurization system. 

2.0 Problem 

A design concern for autogenous pressurization is that particulate ignition in the 

heat-exchanger discharge (pressurization supply) lines offers a catastrophic vehicle 

failure mode. The potential for failure increases in proportion to the selected heat 

exchanger discharge temperature. 

3.0 Objective 

The NASA statement of work is to "Select optimum L02 tank helium pressurization 

system based on tank pressure limits and specified reference trajectories and 

considering safety, reliability, operability, simplicity, weight, including residuals 

and cost." 

( 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

4.0 Approach 

The approach was to perform an analysis of the baseline (autogenous) system 

varying heat exchanger outlet temperature to obtain the residual weight sensitivity. 

A similar analysis was performed for helium pressurization system. and design 

features were selected for systems at ambient and cryogenic storage. Cost and 

weight estimates were performed for all three systems and comparisons made. 

5.0 Results 

The system weight impact was compared by summing component and ullage 

weight. The payload weight impact is identical to the weight carried to orbit 

insertion. and increases as pressurant residuals are reduced. The autogenous 

system impacts payload by 4250 lbs. at the baseline conditions. The cryogenic 

helium storage system has a payload impact of only 3180 lbs .• but at a cost 

increase of $1.2M/flight when compared to autogenous pressurization. An 

ambient helium system impacts payload by 5500 lbs., but this value can be 

reduced to about 41 00 lbs. by staging bottles with booster engines. This system 

costs about $1.8M/flight more than autogenous. 

6.0 Conclusions and Recommendations 

Minimum weight is achieved using cryogenic stored helium. Ambient stored 

helium is the next lightest, with fixed orifice autogenous being better at higher 

HEX temperatures. There are more components and higher costs for the helium 

systems. The high-temperature GOX issue is traded with equally catastrophic 

bottle-failure issues. 

MARTIN MARIETTA 

MANNEDSPACEsvr~~MS 

( ( 
7.0 Supporting Data 

Task Number 3-P-017, "STME L02 NPSP Requirements" and Task Number 3-P-025 

"L02 Tank Pressure Limits." Nein, M. E. and J. F. Thompson, "Experimental and 

Analytical Studies of Cryogenic Propellant Tank Pressurant Requirements, 

"NASA TN 03177, February 1966. 

8.0 Attachments 

Task Number 3-P-026, "LOX Tank Pressurization System Using Helium." 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
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Approach 

Perform baseline pressurization system analysis and vary heat exchanger outlet 

temperature to obtain residual weight sensitivity. 

Generate pro/con's of system. 

Perform helium pressurization system analysis to obtain similar pressure profiles. 

Generate residual weight sensitivity. 

Generate cost trade between baseline and helium alternatives. 

( 

\ 

·-... 

3-P-026 

Page 1 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

3-P-026 

Page 2 

Baseline Fixed Orifice Pressurization System 

• Pro's 

- Simple pressurization system 

- System weight reduced significantly with increase in HEX temperature 

• Con's 

- Requires engine heat exchanger 

- Potential combustion with particle impact 

MARTIN MARIETTA 

MANNED SPACE SV"' ... cMs 

( ( 
The components for the autogenous fixed-orifice configuration consist of a heat 

exchanger, orifices, and check valves for each engine to accommodate engine-out. 

The orifice at each engine provides tank pressurant flow corresponding to the 

propellant use for that engine. The check valve isolates an engine which is not 

firing. For the 1-1/2 stage, the booster engines are manifolded with a single 

in-flight disconnect. 

Tank prepressurization subsystems are identical, whether autogenous or helium 

pressurization systems are used, and are not shown as trade study discriminators. 

The STME cost and weight impacts associated with NPSP variations are from the 

STPT consortium. The data are contained in NLS Data Book Log #35 in response 

to an action item from the propulsion TIM #1. 

•..-... 

facing 

3-P-026 

Page 3 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
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This analysis of autogenous fixed orifice performance was done to establish a 

reference for the helium sys~em(s) studies presented later. 

Cycle-0 baseline parameters are shown for a nominal H LL V and 1-1/2 stage 

including engine-out (at liftoff) system response. 

The tank prepressurization does not correspond to the cycle-0 baseline; the 

No Bleed cycle-0 baseline resulted in severe system impacts as discussed in 

the appropriate trade study reports. 

facing 

3-P-026 

Page4 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

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a.. 25. 

ts. 

-u0. 

3-P-026 

Page 4 

Cycle Zero Baseline Ullage Pressure Profiles 

0-0 

Minimum Relief Pressure 

___, 

HLLV Eng Out 

100· zee. 

3.0 Ibm/sec/eng 

700°R Diffuser Inlet T emBerature 

HLLV Nominal 

~ / 1··~ .sstage 

-

Vapor Pressure 

300- 400. 

ng Out 

Ttme for lHtoff <sees> 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
' . .. 

THIS PAGE INTENTIONALLY LEFf BLANK 

; ;';; . '.t 
., 

:0 

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C> 

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The pressurization system weight is the sum of hardware and ullage residuals at MECO. 

A decrease in this weight for an otherwise fixed launch system allows a corresponding 

increase in payload capability. 

The hardware weight was assumed to be approximately that for the Shuttle ET, or 

about 450 lbs., and not a variable with heat exchanger outlet temperature. 

The hardware cost was estimated to be $254K/flight. 

facing 

3-P-026 

Page 5 

MARTIN MARIETTA 

MANNED SPACE SY~TEMS 

( ( 
Autogenous Pressurizat~on System Weight 

6000------------~------------~ 

~ 5000 

m 

.J 

-

~ CJ w 4000 

3: 

3 000 I . I I I I I I I I I I I I 

400 500 600 700 800 900 1000 

3-P-026 

Page 5 

HEAT EXCHANGER OUTLET TEMPERATURE (OEGR) 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

3-P-026 

PageS 

Cryo Helium Pressurization System 

• Pro's 

- Low pressurization system weight 

• Con's 

- Additional stretch of LH2 tank needed (approximately 1.5 inch) 

- Tank weight impact -110 Ibm. 

- Uncertaintly in LOX evaporation during mainstage 

- Additional LH2 tank penetrations for access and feedthrough 

- Requires engine heat exchanger 

- Additional components decrease reliability 

- System cost 

MARTIN MARIETTA 

MANNED SPACE SYC::TEMS 

( ( 
( 

.• 

Typical performance profiles are shown for constant flowrates at 700°A HEX supply temperature 

for the HLLV and 1-112 stage. Note that differential orificing was selected between Booster and 

Sustainer pressurant flowrates to minimize the difference in peak pressure (near BECO) and final 

pressure (at MECO). 

An equivalent to the autogenous fixed orifice system was used for comparison purposes. 

A companion analysis for engine-out performance is shown on the next page. 

( 

-· 

facing 

3-P-026 

Page 7 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

40. 

36. 

- a a. 

0) 

..-

Ul 

a. 

- 28. 

Cl> 

L. 

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Ul 

Ul 34. 

Cl> 

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--::::> 

18. 

12. 

&. 

3-P-026 

Page7 

00 

LOX Tank Ullage Pressure vs Time 

Helium Pressurization System 

Normal Performance 

700°A Heat Exchanger 

Outlet Temperature 

100. 200. 

, ---

300. 

Time from liftoff <T-0) <sees) 

HLLV J 

1.5 Stage 

400. 

MARTIN MARIETTA 

MANNEDSPACESY~T~MS 

( ( 
4.0. 

36. 

- 32. CJ) 

-U) 

a. 

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Cb 28. 

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16. 

12. 

&. 

3-P-026 

Page 8 

' ,, 

Cl.O 

LOX Tank Ullage Pressure vs Time 

Helium Pressurization System 

' -

100. 200 ~00. 

T 1me from 11ftoff (T-O)( sees) 

Engine Out 

700°A Heat ExhartJer 

Outlet Temperature 

1.5LStage 

----- _, HLLV 

4..00. bOO. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
The pressurization system weights are less than those for the autogenous (GOX) system, principally 

due to the reduced ullage mass. For this system, a decrease in weight for an otherwise fixed launch 

system allows a corresponding increase in payload capability. 

The LH2 tank stretch is to accommodate cryo-helium storage inside the tank (a volume of about 

75 ft3) and associated weight maintains a comparable LH2 usable volume/mass as a trade parameter 

between autogenous GOX pressurization and cryo-helium storage. 

An estimate of 450 lbs. LOX evaporation is based on data from Nein and Thompson, 

The cost for this system is estimated to be about $1 .43 M/flight. 

There are more components (bottles, orifices, check valves, etc.) in a helium pressurization system 

than in the autogenous GOX system as can be seen by comparing the sub-system schematics. 

These components decrease system reliability somewhat, but prior helium pressurization systems 

components have performed adequately. However, one applicable failure resulted in the loss of an 

S-IVB stage during ground test operations when a helium bottle ruptured. 

facing 

3-P-026 

Paae 9 , 

MARTIN MARIETTA 

(, MANNEDSPACESVrTEMS 

{ ( 
( 

Cryogenic Storage Helium. Pressurization System Weight 

4000 

3800 

3600 

-

E 

.a 

-- 3400 

.c 

-

0) 

·a; 3200 ~ 

3000 

2800 

3-P-026 

Page9 

300 400 500 600 700 800 900 . 1 000 

Heat Exchanger Outlet Temperature (DEGR) 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

3-P-026 

Page 10 

Ambient Helium Pressurization System 

• Pro's 

- Low system weight if helium bottles can be staged with engines 

- No additional LH2 tank penetrations 

• Con's 

- Uncertainty in LOX evaporation during mainstage 

- Packaging of system for staging 

- Propulsion module length 

- Requires engine heat exchanger 

- Additional components decrease reliability 

- Helium cost 

MARTIN MARIETTA 

MANNEDSPACESY~T~MS 

( ( 
THIS PAGE INTENTIONALLY LEFT BLANK 

·•: ' ~ 
~ 

i 

2 

G) 

,., 

8 

aJ 

~ 2 

~ 

z 

0 

-f 

l! 

r-

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( 

An ambient helium storage system was evaluated to determine cost and performance 

differences with cryo-helium storage. The system shows a cost advantage over 

operating with cryo helium stored in the LH2 tank, and system weights for the 1-1/2 

stage appear to be competitive if bottles are staged with booster engines. 

Other discussion items are identical to the cryo-helium pressurization system, 

except for the need to regulate the helium flow interactively to accommodate the 

continuously variable inlet temperature to the heat exchanger as the bottles blow 

down. A performance analysis is not appropriate at this stage, since the results 

could vary greatly depending on detailed design assumptions. 

Manifolding the Booster lines and using individual shut-off valves and check valves 

accommodates engine-out operations. 

The layout accommodates staging the bottles with the booster module. 

There are a large numuer of possible schematic layouts, each of which has a 

different component counVcosVetc. The layout selected should be adequate for 

. failure considerations. A detailed study of cost and failure considerations is 

required for this approach. ' I 

The motorized regulator is required to accommodate changing heat exchanger 

inlet temperatures as the bottles are blown down. However, a flow control scheme 

equivalent to the autogenous fixed-orif.ice system was used for comparison. 

facing 

3-P-026 

Page 11 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
STME 

HEX 

Disconnect 

Check 

Valve · 

3-P-026 

Page 11 

( 

' 

Helium Pressurization System 

I I I I I I ~JF&O 

-, 

1 

LOX Tank 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Helium Bottle Packaging 

Preliminary packaging of helium bottles appears feasible. Dependent on final thrust structure, 

feedline arrangement and propulsion .module length. 

3-P-026 

Page 12 

( 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( 
--

THIS PAGE INTENTIONALLY LEIT BLANK 
~ 

"" 

2 

~ C) 

., 

~ 

" 1"1 

C'J 

r· 

l> 

:z 

2 " 

0 

-t 

l! 

i 

0 ( 

The pressurization system weight is the sum of hardware and ullage residuals at 

MECO with an "effective" payload impact associated with staging bottles with the 

booster engines. A knock-down factor of 0.43 was used, considering booster 

module separation at 193 seconds of 465 second flight. · 

The pressurant line and associated hardware weight was assumed to be 

approximately the same as for the autogenous system, and not variable with 

HEX outlet temperature. 

The helium storage system weight is estimated to be 3940 lbs. and the hardware 

cost estimated to be $2.07 M/ flight. 

facing 

3-P-026 

Page 13 

MARTIN MARIETTA 

MANNED SPACE SYSTJ:MS 

( ( 
( 

Ambient Helium Pressurization System Equivalent Weight 

8000 -----------t 

,.... 

E 

.0 

-

-

~ 

I:. 

(1) 

-Q) 

~ 

3-P-026 

Page 13 

7000 

6ooo I "- No bottles staged 

5000 

4000 

I 

62.5% of bottles staged 

I 

3000 

300 400 500 600 700 800 900 1000 

Heat Exchanger Outlet Temperature <DEGR) 

( 

... 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Summary & Conclusions 

1.5 Stage 

• Minimum pressurization system weight is achieved using cryogenic 

storage helium. 

• Ambient storage helium is the next best with fixed orifice autogenous 

being better at higher HEX temperatures. · 

HLLV 

• Minimum pressurization system weight is achieved using cryogenic 

storage helium. 

• Fixed orifice autogenous weight performance is better at higher HEX 

temperatures. 

• Ambient helium system assumes no bottle staging and consequently 

will result in significant weight impact. 

Both 

• There is a potential for tank weight reduction by employing an intelligent 

flow control system; a time-consistent structural loads analysis is required 

to assess the potential payload benefits. 

3-P-026 

Page 14 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
... . ( 

Task Number 3-P-027 

STME Heat Exchanger Performance 

Prepared By: 

G.Piatt 

Approved By: 

Z. Kirkland 

20 Dec, 1991 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

NASA Statement of Work: 

Executive Summmary 

Assess.current STME heat exchanger performance and possible outlet temperature 

increase. Assuming LOX tank pressurization system uses the STME heat exchanger 

for an energy source, trade system performance (residuals) against engine impacts 

from increased heat exchanger outlet temperature. 

It was shown that for the Autogenous (GOX) pressurization system, the system 

using ambient helium storage, or the system using cryogenic (LH2 temperature) 

helium storage, the heat exchanger discharge temperature should be increased 

above the reference 700°R. It was found that the payload improvement due to 

pressurization system weight saved by increasing the temperature 200°R was 

600 - 1 000 lb. Tank wall temperatures were increased, as a result of the 200°R 

pressurant temperature increase, by 130°R at cutoff. This tank wall temperature 

increase is not expected to be detrimental. Greater tank wall temperature increases 

will require further evaluation. Heat exchanger cost per flight increases were small 

compared to the benefit in terms of payload improvement ($8400 for 600- 1000 lb). 

The above work was based on a 1 1/2 stage vehicle. The overall effect is expected 

to be similar for the HLL V. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
Task Number 3-P-027 

STME Heat Exchanger Performance 

1.0 Summary 

It was shown that for the autogenous (GOX) pressurization system or the helium 

system the heat exchanger discharge temperature should be increased above the 

reference 700°R to at least 900° R. This would improve the payload capability of 

the vehicle by 600 - 1 000 lbs. 

2.0 Problem 

To assess the value of increasing the STME heat exchanger outlet temperature. 

3.0 Objective 

General 

To evaluate the desirability of increasing the STME heat exchanger outlet temperature. 

Specific 

To assess, for three candidate LOX tank pressurization systems, the desirability of 

increasing the heat exchanger outlet temperature from the reference 700°A. 

4.0 Approach 

The approach to performing this study was: 

• To calculate a pressurization system weight, for 500, 700 and 900°A heat exchanger 

outlet temperatures, for each of three candidate LOX tank pressurization systems as follows: 

- Autogenous (GOX) pressurization system. 

- Helium system with helium stored in ambient temperature bottles. 

- Helium system with helium stored in bottles submerged in the liquid hydrogen tank. 

( 

l!o ... 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
f:' 

\ 

UJ 

( 

Calculate the tank wall temperature for the 900°R tank inlet temperature to assure 

tank material integrity at the higher temperature. 

5.0 Results 

The results of this study are attached. The primary results of the study are listed below. 

6.0 Conclusions and Recommendations 

The payload capacity improvement due to pressurization weight saved by increasing 

the heat exchanger outlet temperature from 700 to 900°A was 600 - 1 000 lb. The 

corresponding tank wall temperature increase was 130°R to 755°R. Greater tank 

wall temperature increases will require further evaluation. Heat exchanger cost per 

flight increases were small compared to the payload improvement ($8400 for 600- 1000 lb.). 

Further work is required to define the helium pressurization control system. 

7.0 Supporting Data 

STPT fax NM0-086-20 "STME Heat Exchanger Parametrics" dated 1 0/04/91. 

8.0 Attachments 

Study "Task Number 3-P-027, STME Heat Exchanger Performance" dated 12/20/91. 

MARTIN MARIETTA 

MANNED SPACE syr..--:MS 

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Guidelines and Assumptions 

• Did not differentiate between heat exchanger pressurant discharge temperature 

and tank inlet temperature. 

• Heat Exchanger weight and cost data taken from STPT fax NM0-086-20 "STME 

Heat Exchanger Parametrics" dated 1 0/04/91. 

• Practical upper limit for heat exchanger outleUtank inlet mean temperature taken 

as 900°R based on Shuttle experience. 

• Pressurization system hardware weight excluding high pressure bottles taken 

as 450 lb. 

• Pressurization gas bottles are assumed to be available in any size required. 

3-P-027 

Page 1 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

The facing page shows the heat exchanger location in the engine system. 

facing 

3-P-027 

Page 2 

( 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( 
) -·· 

•• 

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3-P-027 

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MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
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3-P-027 

Page 3 

LH2 

Tank 

Diffuser 

LOX TANK FIXED ORIFICE CONFIGURATION 

Disconnect 

Check 

Valve 

Orifice 

Heat 

Exchanger 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

From the data supplied by STPT, the graph on the facing page was derived. The heat 

exchanger weight is translated into payload weight penalty. The factor used for booster 

weight per pound of payload weight was 0.43. 

facing 

3-P-027 

Page 4 

MARTIN MARIETTA 

MANNED SPACE SY~,.EMS 

( ( 
Total Heat Exchanger Payload Weight Penalty 

60~----------------~ 

1.5 Stage Autogenous System 

40 

:E -

m 20 

..1 

-t-

::t 

~ 

-

w 

~ 

3-P-027 

Page 4 

0 Reference, 

3 lbm/sec@700 DEGR 

-20 

-40 I I I I I I I I I I I I I 

400 500 600 700 800 900 1000 

HEAT EXCHANGER OUTLET TEMPERATURE (DEGR) 

( 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

The total payload weight penalty is plotted on the facing page for the autogeneous 

(GOX pressurization system). Comparing this figure with the previous one, it is seen 

that the heat exchanger weight penalty is much smaller than the total system weight penalty. 

facing 

3-P-027. 

Page 5 

( 

MARTIN MARIETTA 

MANNED SPACE SY~1EMS 

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The facing page shows the results of an analysis to gain insight into the performance 

of the heat exchanger at two different power levels and a range of GOX flowrates. 

The curves were calculated for a particular heat exchanger configuration within the 

range supplied by the STPT. Several problems are evident from these calculations; 

first, the heat exchanger outlet temperature is a very strong function of power level. 

Therefore, the outlet temperature at 1 00°/o will have to be 25 -30 degrees higher than 

nominal to yield a mean as high as the nominal. Secondly, at the low gas temperatures 

characteristic of the 70°/o power level, icing of the duct may occur. 

facing 

3-P-027 

Page 6 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
.. 

....... 

a: 

C) 

G) 

c 

....... 

I! 

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..., 

1! 

800 

700 

600 

~ 500 E 

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t-

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Heat Exchanger Outlet Temperature vs. 

LOX-GOX Flowrate 

STPT Platelet HEX 

Provisional Power Balance 

B 100% 

• 70% 

400 I I I I I I I I I I I 

3-P-027 

Page 6 

2 3 4 

Flowrate (lb/sec) 

5 6 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Summary & Conclusions 

• The heat exchanger cost and weight are not significant compared to the ullage 

gas weight saved by increasing heat exchanger outlet temperatures in the range 

considered (500- 900°R). 

• Heat exchanger performance is a strong function of the engine power level. To 

achieve an average outlet temperature as shown will require a higher outlet 

temperature at full thrust. 

• Additional work is recommended to establish the maximum practical heat 

exchanger outlet temperature and resulting payload weight penalty and cost. 

3-P-027 

Page 7 

MARTIN MARIETTA 

MANNEDSPACESYrTEMS 

( ( 
.,, 

:::0 

"" 

~ ::z 

G) 

"'0 

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m 

r 

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2 

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The schematic for the helium pressurization system shown on the facing page is 

valid for either cryogenic or ambient temperature helium storage. In the case 

of ambient storage, 5/8 of the storage containers would be jettisoned with the 

booster. For cryo storage in the LH2 tank, the total bottle weight is charged 

against payload . 

facing 

3-P-027 

Page 8 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
--

THIS PAGE INTENTIONALLY LEFT BLANK 
., 

, 

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2 

~ C) 

"tt 

;;, ~ 

O:J 

r-

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STME 

HEX 

Disconnect 

3-P-027 

Page 8 

( 

Helium Pressurization System 

I I I I I I 1'1?1-IF&D 

-. 

I 

LOX Tank 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

3-P-027 

Page9 

Helium System 

Initial Pressure (psia) 

Final Pressure (psia) 

Initial Temper~ture (

0

R) 

Final Temperature (

0 R) 

Ambient Storage 

4500 

1000 

580 

400 (calculated) 

Tank Pressure (..,;;;ia) 24 at cutoff 

Tank Pressurant Inlet 

Temperature (

0 R) 500 to 900 

Resulting Heat Exchanger · 

Flowrates (Ibm/sec) 0.35 to 0.26 

( 

Cryo Storage 

4500 

1000 

40 

30 (calculated) 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( 
.. ( 

Helium Pressurization System Payload 

Weight Penalty 

5000 .....------------., 

~ 4000 

....... -

+J 

J:::. 

tJ) 

...... 

cv 3 3000 

2000 I I I I I I I I I I I I I 

D Ambient Storage 

•5/8 of Bottles Jettisoned 

whh Booster 

0 lH2 Temperature Storage 

•All Bottles Carried to Orbit 

400 500 600 700 800 900 1000 

Tank Inlet Temperature <Deg R) 

3-P-027 

Page 10 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Summary & Conclusions 

• The heat exchanger cost and weight are not significant compared to the ullage gas 

weight and storage bottle weight saved by increasing heat exchanger outlet 

temperatures in the range considered (500 - 900°A). 

• Improvements in payload capability appear possible by further increases in heat 

exchanger outlet temperature. 

• Additional work is required to establish the maximum practical heat exchanger 

outlet temperature and resulting payload weight penalty and cost. · 

• Further work is required to define the helium pressurization system control scheme. 

3-P-027 

Page 11 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
( ( 

Task Number 3-P-033 

LH2 Passive Recirculation Performance Analysis 

Prepared By: 

G. Platt 

Approved By: 

Z. Kirkland 

20 Dec, 1991 

MARTIN MARIETTA 

· MANNED SPACE SYSTEMS 
( 

Executive Summary 

Task 3-P-033, "LH2 Passive Recirculation Performance Analysis" of the National Launch 

System Phase B study done by MMMSS under the Shuttle C contact reads as follows: 

"Analysis of LH2 feed system with passive recirculation system to assess feasibility, margins 

and performance including an assessment of engine prestart restrictions if any." This is a 

report of this study and is based upon the Marshall Space Flight Center study plan dated 

August 5, 1991, and presented at the Technical Interchange Meeting at Marshall Space Flight 

Center on August 28, 30, 1991, by Danny Davis, the cognizant Panel Chairman. 

Conclusions and recommendations were: 

• Simple System. 

• Screens make geysering correlation uncertain. 

• Non-horizontal screens will not become vapor-bound. 

• Saturated liquid hydrogen with 23 cubic inches/second of vapor being produced in pump 

expected after prepress. 

Rapid warmup after start of prepress reduces NPSP 5 psi/min. 

• Makes for short available hold time before depressurization-repressurization required. 

• May force tank design pressure to be increased. 

• May complicate operations by forcing very short engine start window before recycle required. 

• Reevaluation required when engine start pressure requirement is established. 

MARTIN MARIETTA 

MANNED SPACE SV~TEMS 

( ( 
( 

Task Number 3-P-033 

LH2 Passive Recirculation Performance Analysis 

1.0 Summary 

The LH2 "passive recirculation" system appears to be capable of furnishing saturated, 

mostly liquid, hydrogen at the turbopump inlet at the start of prepressurization. The heat 

up rate after prepressurization of 5 psi/minute will limit hold time with tanks pressurized 

to 2 - 3 minutes. 

2.0 Problem 

To study and predict the performance of the reference hydrogen no bleed system. 

3.0 Objective 

General 

Determine the performance of the LH2 Passive Recirculation (no-bleed) system. 

Specific · 

Gain an understanding of the performance characteristics of the LH2 passive recirculation 

system. Assess geysering situation, feedline screens, and hold time. 

4.0 Approach 

The approach to performing this study was: 

• Calculate engine inlet pressure with tank pressurized and unpressurized. 

• Convert the engine heat flux to vapor volume. 

• Calculate the rate of warmup of pump and propellant due to heating. 

• Research feedline stagnation and geysering; determine performance of system relative to 

geysering limits. 

( 

• Calculate screen performance with regard to passing vapor. MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

5.0 Results 

The results of this study are attached. The primary results are listed below. 

6.0 Conclusions and Recommendations 

• Simple System. 

• Screen makes geysering correlation uncertain. 

• Non-horizontal screens will not become vapor-bound. 

• Saturated liquid hydrogen with 23 cubic inches/second of vapor being produced in pump 

expected after prepress. 

Rapid warmup after start of prepress reduces NPSP 5 psi/minute. 

• Makes for short available hold time before depressurization-repressurization required. 

• May force tank design pressure to be increased. 

• May complicate operations by forcing very short engine start window before recycle required. 

• Reevaluation required when engine start pressure requirement is established. 

7.0 Supporting Data 

NASA-CR -64-3, Contract NASB-5418, Summary Report for the Period 1 July 1963 through 

30 June 1964, "Mechanics of Geysering of Cryogenics," dated June 1964. 

STPT CM No. NM0-089-17, "STME Start and Shutdown Requirements," dated 1 0/25/91. 

STPT CM No. NM0-076-05, "STME Turbopump Heat Leaks," dated 8/29/91. 

8.0 Attachments 

' 

Task Number 3-P-033, "LH2 Passive Recirculation PerformanGe Analysis." 

MARTIN MARIETTA 

MANNED SPACE SY~TEMS 

( ( 
.. 

Task Number 3-P-033 

LH2 Passive Recirculation Performance Analysis 

Attachment-Detailed Data 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

The Reference Configuration 

Our analysis was based on a Martin-Marietta version of the NASA reference feedline 

configuration shown on the facing page. 

facing 

3-P-033 

Page 1 

( 

MARTIN MARIETTA 

MANNED SPACE SY~TEMS 

( 
{ 

( 

· NASA Sketch of Reference LH2 Passive Recirculation System 

--

--·---- -- -· ·--- ... - ·- -- ··-. -·---·-·-·- •·-·· -·---- -· ------ -·- -·-·····- .. - - .. -·--. - ·-·r--·--- ... \ 

3-P-033 

Page 1 

....... -·· .. -·- .. ·- ..... --- .......... --······ .. ·. --.;: -~ .. ·- ... ~ .. ---- .. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

The reference configuration to the level of detail known today would provide propellant saturated 

at the local pressure to the engine for start. The engine would be chilled to the extent possible, 

as limited by its configuration and the internal flow path and heat transfer situation within the engine. 

The local pressure would be about three psi above tank ullage pressure; this would mean that in 

the best case the tank design pressure could only be reduced by three psi by providing a positive 

chilldown system. The real benefit of a feedline/engine chilldown system is that it does a positive 

feedline/engine/chill and provides for a hold after prepressurization and before engine start. If 

there is no engine chill system, or if it is not a positive one, the state of engine chill must await 

a detailed engine analysis or engine test to be known at all, and a limitation of the duration of 

hold after prepressurization must be accepted. This duration, per several analyses, is about one 

to two minutes. This time can be extended by providing a chilldown system. 

facing 

3-P-033 

Page 2 

MARTIN MARIETTA 

MANNED SPACE SY~TEMS 

( ( 
k , _ 

3-P-033 

Page 2 

FEEDLINES 

REFERENCE NO BLEED 

• RULE OF THUMB. 

UD < 10 CONVECTION OK 

UD > 20 CONVECTION NOT ENOUGH 

(GEYSER REGION) 

• PROBABLE DESIGN BETWEEN 10 AND 

20 OR UD > 20. 

• SCREEN IN LINE WILL INHIBIT CONVECTION. 

• PORES IN SHUTTLE-TYPE SCREEN WILL 

BE "STABLE". 

• MOST VAPOR WILL PASS THROUGH AND 

RISE THROUGH SCREEN IF SCREEN IS NOT 

FLAT, HORIZONTAL. 

• FEEDLINE WILL BE SATURATED. BOILOFF RATE 

46 CU IN/SEC BEFORE PREPRESS 23 CU IN/SEC AFT 

PRE PRESS. 

• THERE IS NO VALID ANALYTICAL MODEL OF 

THE FEEDLINE WITH SCREEN. 

• POSSIBLE SOLUTIONS: 

• ON BOARD BLEED 

• OVERBOARD BLEED 

• RECIRC SYSTEM 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

The reference hydrogen no-bleed system will have a very rapid warmup rate 

after prepressurization. 

The graphs on the facing page show the calculated warmup rates for the insulated 

and the uninsulated turbopump cases. 

MARTIN MARIETTA facing 

3-P-033 

Page 3 MANNED SPACE SY<:TJ:MS 

( ( 
( 

3-P-033 

Page 3 

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Warmup of Propellants After Prepress. 

LIQUID HYDROGEN 

50--------------------------------

40 

8. 30 ftl ---e-1 uninsul. turbopump 

.> 

N 

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.,.__, insul turbopump 

20 

I 0 I I I I I I I I I I I 

0 100 200 300 

Time from Start of Prepress 

(sec) 

400 500 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
.. 

( 

The reference no-bleed system has an UD of 10 to 20. For such configurations, geysering 

is possible. This is shown by the facing figure which comes from the research at 

Martin-Marietta. The likelihood of stagnation leading to geysering for a 12 inch line is 

considered small, however, for a smaller line or a longer line, geysering would be expected. 

Also, a feedline screen similar to the one in the Shuttle Orbiter feedlines is considered 

mandatory. This screen will inhibit convection and may lead to geysering. In a hydrogen 

system, geysering is not damaging. As mentioned above, even a fully saturated feedline 

would be acceptable, but the engine thermal condition would not be clearly understood . 

MARTIN MARIETTA facing 

3-P-033 

Page 4 MANNED SPACE SYSTEMS 

( ( 
~- .. ( -· ( 

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.86 .7 .568 

L = 20 Z = 477600 184320 

3-P-033 

Page 4 

z = ( Q/R)·l 

12 oc (N,..}IIl 

log 10 Z = 5.6 7 5.26 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Summary & Conclusions 

• Simple System. 

• Screens make geysering correlation uncertain. 

• Non-horizontal screens will not become vapor-bound. 

• Saturated liquid hydrogen with 23 cubic inches/second of vapor being produced in 

pump expected after prepress. 

Rapid warmup after start of prepress reduces NPSP 5 psi/minute. 

• Makes for short available hold time before depressurization-repressurization required. 

• May force tank design pressure to be increased. 

• May complicate operations by forcing very short engine start window before recycle required. 

• Reevaluation required when engine start pressure requirement is established. 

3-P-033 

PageS 

( 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( 
IL ' ( 

( 

Task Number 3-P-034 

LH2 Bleed/Recirculation Study 

Prepared By: 

G. Platt 

Approved By: 

Z. Kirkland 

20 Dec, 1991 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
. ( 

Executive Summary 

The NASA Statement of Work for this study reads as follows: 

Assess a no-bleed vs. bleed or recirculation system for the LH2 feed system considering at 

a minimum operability, reliability, propulsion module layout and tank stretch, weight, and cost. 

• Reference No-Bleed System will result in saturated LH2 in feedline and engine pump with 

vapor in engine pump and dry lines downstream of engine pump. 

Convection path complicated by screen. 

Analytical model and test program to anchor analytical model required. 

Warm-up after prepressurization increases saturation pressure 5 psi/minute. Would require 

depressurization of tank, repressurization for very short hold. (Engine start pressure not yet defined.) 

• On-Board Bleed has low flowrate, hydrogen quality in turbopump will be poor (80°/o vapor 

by volume). If this is satisfactory, would allow improvement in hold after prepress relative to 

no-bleed system. 

Test program required. 

Slight improvement in performance compared to no bleed. 

• Overboard bleed has adequate performance after prepressurization. Hold time limited due 

to loss of LH2 at 1.2 lb/sec per engine. 

Hardware complexity a disadvantage. 

SSME manufacturer uses this system, in principle, for single engine tests. 

· Should be retained for further study. 

• Backward recirculation did not appear advantageous. 

Provides good engine/pump chill. 

Introduces large volume of vapor into feedlines. 

Hardware complexity a disadvantage. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
k>-.1 

_ 

• Forward recirculation: 

Predictable. good experience with systems. 

Hardware complexity a disadvantage. 

Together with overboard bleed, provides best engine/pump chill. 

Provides best performance (best chill, no hold time limitation). 

( 

) 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Task Number 3-P-034 

LH2 Bleed/Recirculation Study 

1.0 Summary 

Four alternates to the reference no-bleed system were studied to establish their 

performance characteristics and other attributes. Forward recirculation and the 

overboard bleed to the facility were both superior to the reference system. 

2.0 Problem 

Assess a no-bleed vs. bleed or recirculation system for the LH2 feed system considering 

at a minimum operability, reliability, propulsion module layout and tank stretch, weight, and cost. 

3.0 Objective 

General . To compare candidate bleed systems with the no-bleed system for the LH2 feed system. 

Specific 

Compare performance, predictability. repeatability, precedence, engine impact, feed 

system impact, engine test impact, potential future change, operational efficiency, potential 

hazard, and hardware complexity of candidate bleed system concepts against the reference 

no-bleed system. 

4.0 Approach 

First, the candidate systems were identified and the performance of each was predicted. 

Then the systems were compared with each other and the reference with regard to the 

attributes listed in 2, above. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
.. 

5.0 Results 

The results of this study are attached. The primary results are listed below. 

6.0 Conclusions and Recommendations 

• Reference No-Bleed System will result in saturated LH2 in feedline and engine pump 

with vapor in engine pump and dry lines downstream of engine pump. 

Convection path complicated by screen. 

Analytical model and test program to anchor analytical model required. 

Warm-up after prepressurization increases saturation pressure 5 psi/minute. Would 

require depressurization of tank, repressurization for very short hold. 

• On-Board Bleed has low flowrate, hydrogen quality in turbopump will be poor (80°/o vapor 

by volume). If this is satisfactory, would allow improvement in hold after prepress relative 

to no-bleed system. 

Test program required. 

Slight improvement in performance compared to no bleed. 

• Overboard bleed has adequate performance after prepressurization. Hold time limited due 

to loss of LH2 at 1.2 lb/sec per engine. 

Hardware complexity a disadvantage. 

SSME manufacturer uses this system, in principle, for single engine tests. 

Should be retained for further study. 

• Backward recirculation did not appear advantageous. 

Provides good engine/pump chill. 

Introduces large volume of vapor into feedlines. 

Hardware complexity a disadvantage. 

{ 

I / 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

• Forward recirculation: 

Predictable, good experience with systems. 

Hardware complexity a disadvantage 

Together with overboard bleed, provides best engine/pump chill. 

Provides best performance (best chill, no hold time limitation). 

7.0 Supporting Data 

3-P-033, "LH2 Passive Recirculation Performance Analysis." 

8.0 Attachments 

Task Number 3-P-034, LH2 Bleed/Recirculation Study "LH2 Passive Recirculation 

Performance Analysis." 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
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The problem of LH2 feedline and engine conditioning during prelaunch has been 

solved for different engines and feed systems in different ways, however, there has 

always been a chill of the LH2 turbomachinery and the thrust chamber. The STME 

has baselin.ed a zero chill system, and has no specific requirements to which a chill 

system should be designed. Whether the STME can be successful in developing a 

zero chill system is not known, therefore, it was considered necessary to consider 

several possible methods of chilling the STME prior to start. These methods have 

different degrees of potential effectiveness. The performance of each was predicted 

as applied to a reference configuration supplied by NASA, shown on the facing page, 

and is compared to the reference. 

MARTIN MARIETTA facing 

3-P-034 

Page 1 MANNED SPACE SYSTEMS 

( 

( 
--· ;· ··- . -#"' / _..7 - - - ---

~ 

~w\____ ···-··· . ··-···-·- . -

.- ...... --·~· -· 

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3-P-034 

Page 1 

..... 

Feedlines Reference No Bleed 

•Rule of Thumb 

UD < 1 0 Convection OK 

UD > 20 Convection Not Enough (Geyser Region) 

•Probable Design Between·1 0 and 20 or UD > 20 

•Screen In Line Will Inhibit Convection 

•Pores in Shuttle-Type Screen Will Be "Stable" 

•Most Vapor Will Pass Through and Rise Through 

Screen if Screen is Not Flat, Horizontal 

/ 

•Feedline Will Be Saturated. Boiloff Rate 46 ln3/Sec 

Before Prepress 23 In Cu/Sec after Prepress 

·There is No Valid Analytical Model of the Feedline 

With Screen 

•Possible Solutions 

•On Board Bleed 

•Overboard Bleed 

•Recirc System 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Ch illdown Systems 

The candidate engine/feedline chilldown systems were evaluated on the basis 

of the attributes shown on the facing page. 

All chilldown systems were assumed to have a 60 Btu/hr-ft2 heat leak if insulated 

and 600 Btu/hr-ft2 heat leak if uninsulated. 

facing 

3-P-034 

:1ge 2 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
( 

3-P-034 

Page 2 

( 

/ 

ATTRIBUTES CONSIDERED IN EVALUATING SYSTEMS 

• PREDICT ABILITY 

• REPEATABILITY, ENGINE TEST TO VEHICLE 

• PRECEDENCE 

• IMPACT ON ENGINE DESIGN 

• IMPACT ON FEED SYSTEM/ENGINE INSULATION 

• IMPACT ON ENGINE TEST 

• POTENTIAL FOR REQUIRED FUTURE CHANGE 

• OPERATIONAL EFFICIENCY 

• HAZARD INTRODUCED 

• HARDWARE COMPLEXITY 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

The On-Board Hydrogen Bleed 

The on-board hydrogen bleed was evaluated because of its simplicity. It has no 

ground interface and does not vent to the atmosphere. Based on the Martin-Marietta 

layout shown on the next figure, the available head would be approximately 12 feet 

or a maximum of 0.36 psi. The return line was assumed to be one inch in diameter 

and reenter the feedline just below the disconnect. The calculated flowrate would be 

less than 0.12 lb/sec per engine at the calculated heat load. The thermodynamic 

quality in the pump would be greater than 0.08 or 80°/o vapor by volume. For a doubled 

heat load, the flowrate would be less than 0.075 lb/sec with the quality in the turbopump 

and the return line greater than 0.13, or 85°/o vapor by volume. With a doubled return line 

flow capacity and the calculated heat leak, the volumetric quality would be approximately 

0.6. This large volume of vapor would enter the feedline at the elbow where the two 

feedlines split, and would flow up the 17 inch line into the tank. The vapor would tend to 

condense at the time the tan~ is prepressurized; the temperature difference to drive the 

condensation would be 8 deg A. With this system, the hold time after prepressurization 

would not be extended significantly as compared to the reference system. 

The on-board bleed scheme is considered poor from a predictability standpoint. This 

principle, natural circulation, has never been applied to the design of a hydrogen 

system. Similarly, the system performance is not expected to be repeatable from engine 

test to vehicle unless the feed system is virtually identical in the two cases including 

duct and line wall thicknesses and thermal response rates. 

The system adds an engine bleed valve. A very good insulation of the turbopump 

and feedline would be required. This might present a maintenance problem. There 

would be a high potential that a future .change would be required to make the system 

work. The system has no ground interface and introduces few joints which might 

leak, so it would not be expected to introduce much hazard. Similarly the system is 

simple, involving only hydrogen bleed valves and small lines. 

facing 

3-P-034 

.,age 3 

MARTIN MARIETTA 

MANNED SPACE SY~TEMS 

( \. 
3-P-034 

Page 3 

On Board Hydrogen Bleed 

•Requires No Disconnects 

•Requires Engine Hydrogen Bleed Valve 

•Available Head is Approx. 12 Feet or a Maximum 

of 0.36 psi 

•Fiowrate is Less than 0.12 lb/sec Per Engine at 

J 

Calculated Heat Load. Quality is Greater Than 

0.05 (70°/o Vapor by Volume). For Doubled Heat 

Load, Flowrate is Less Than 0.075 lb/sec With 

Exit Quality Greater than 0.13 (85°/o Vapor by 

Volume) 

•With Shortened Aft Compartment (Stretched 

Hydrogen Tank) Performance Would Be 

Reduced -

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

The Overboard Bleed to the Facility 

The scheme is shown in the next figure. It has more capacity than the on-board 

bleed because the full head of the hydrogen on board as well as the vent valve and 

vent line pressure drop are available to drive the flow, which must overcome only 

the vent system pressure drop. Because of this higher head, the system performance 

is more predictable than the on-board bleed, and more repeatable from engine test 

to flight vehicle configuration than the no-bleed or the on-board bleed. Also, the 

available hold time after prepressurization would be extended, limited by the amount 

of propellant lost overboard and the heating of the bulk of the hydrogen in the tank. 

The insulation requirements would not be. severe because of the relatively high flowrate 

that could be obtained. The precedence for the overboard bleed was the RL-10 engine 

which was used for the second stage application (the S-IV). It was bled in flight during 

first stage burn through long vent stacks which carried the hydrogen to the aft end of 

the first stage (the S-1). In this case the system had the full tank pressure plus the head 

of the fluid to drive the flow. The potential for a required future change was considered 

fairly large because of the lack of recent experience with similar systems for flight. The 

facility hydrogen disposal system would have to accommodate the larger hydrogen 

flowrate of the bleed plus tank vent rather than the tank vent alone. 

This system would require an engine bleed valve. It is felt that the impact on engine 

test would be small, since this kind of a bleed has been used in the past for engine 

development tests. 

The system would introduce a moderate hazard, since the added disconnects would 

add leak sources. The number of added disconnects would depend on the design, but 

at least two would be required. The hardware complexity added would be engine bleed 

valves, required for all concepts, the disconnects discussed above, and small lines on board. 

facing 

3-P-034 

age 4 

MARTIN MARIETTA 

MANNED SPACE SV~TEMS 

( ( 
( 

3-P-034 

Page 4 

{ 

' ( 

./ 

Hydrogen Overboard Bleed 

·Bleeds From Engines Are Manifolded and Directed 

to Ground Disconnect 

•Disconnects Must be Provided for Engine Lines (Inboard Engines if Disconnect on Same Plate as 

Fill, If F&D is Jettisoned) 

•Requires Engine Bleed Valve, Probably on Gas 

Generator Supply Line 

•Can be Manifolded to Minimize Number of In flight 

Disconnects 

•Approx. 2.8 psi Available to Drive Flow (Assumes 

1 psi Ground Disposal System Pressure Drop) 

•Provides Approx 0.1 lb/sec Flow to Chill Each 

Engine at Exit Quality of 0.1 (80o/o Vapor by 

Volume) with Calculated Heat Load 

•Provides High Flowrate (1.2 lb/sec) After Prepress. 

•Allows Extended Hold After Prepress. (Limited by 

Allowable Propellant Loss) 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Backward Recirculation 

Backward recirculation was considered for the Saturn S-11 and Saturn S-IVB stages. 

Performance calculations were run utilizing a computer program that was developed 

for Saturn and verified on Saturn and on the Shuttle program. 

The backward recirculation system is simpler than forward recirculation as used on 

Shuttle in that prevalves are not required. The system was calculated for one inch 

lines, since they would not require gimbals, but could utilize the simpler flex lines. 

The allowable hold after prepressurization would be increased compared to the baseline. 

The major disadvantage was that large amounts of vapor were introduced into the feedline. 

This is not necessarily a problem, but relies on the vapor recondensing during 

prepressurization. This recondensation is nearly unpredictable, and tests would be required 

to assess its acceptability. 

3-P-034 

. Page 5 MARTIN MARIETTA 

MANNED SPACE SVCTEMS 

( ( 
THIS PAGE INTEJit(fiONALLY LEFT BLANK 
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The backward recirculation system was judged to have fair to good predictability, 

because of the vapor in the feedline. The performance of the flow in the pump and 

lines was thought to be highly predictable and repeatable, the pump and small line 

would perform the same on engine tests and on the vehicle. An engine bleed valve 

would be required. Since the relatively large flowrate of 1 lb/sec would be assured, 

only moderate feedline and engine turbopump insulation performance would be 

required. The impact on engine test would be moderate, but to allow correct 

assessment of engine performance as it would be on the vehicle, the recirculation 

pump and recirculation line would have to be installed on the engine test facility. 

The potential for future change would be fairly large because of the large volume 

of vapor in the feedline. As to operational efficiency, the system is complicated by 

the addition of recirculation pumps. These pumps have been very reliable, except 

that a connector problem surfaced several times on the Shuttle. It is reasonable to 

assume that this problem has now been corrected. Also, as on any chilldown 

system, it is assumed that temperature interlocks will be required. The hazard 

introduced is considered moderate, in that leakage sources have been introduced, 

and four onboard disconnects have been added. 

MARTIN MARIETTA facing 

'l-P-034 

ge 6 MANNED SPACE SY~Tt:MS 

( ( 
( 

3-P-034 

Page 6 

( 

BACKWARD HYDROGEN RECIRCULATION 

•STANDARD SHUTTLE RECIRC. PUMP PUMPING 

BACKWARDS THROUGH THE ENGINE SYSTEM 

AND FEEDLINE. 

•REQUIRES ENGINE BLEED VALVE ON GAS 

GENERATOR SUPPLY LINE. 

•REQUIRES DISCONNECTS IN RECIRCULATION 

LINES. . . 

•OPERATES WITH PREVALVES OPEN, RETURN 

HYDROGEN FLOW IN MAIN FEEDLINES. 

•UTILIZES SAME RECIRCULATION PUMP AS 

SHUTTLE. 

•ACHIEVES HIGH FLOWRATES EVEN WITH ONE 

INCH RECIRCULATION SYSTEM LINES. 

•DISCONNECTS REQUIRED FOR OUTBOARD 

LINES. . 

·INSENSITIVE TO HEAT LOAD. 

•LARGE VOLUME OF GASEOUS HYDROGEN 

INTRODUCED INTO MAIN FEED LINES. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

The following figures show the backward recirculation performance characteristics. 

With saturated liquid at the pump discharge the flowrate at the point where the 

pressure drop versus flowrate curve intersects the pump performance curve is 

between 0.9 and 1.0 lb/sec. This is the point at which the system will operate. and 

at this point the vapor fraction of the volume flowing at the recirculation line exit is 

0.7. Similarly, with greater subcooling, the flowrate increases to approximately 

1.21b/sec and the vapor volume decreases to 30°/o at 3 deg F of subcooling. Since 

the head is only approximately 3 psi at the tank bottom, we can expect no more than 

1 deg of subcooling. The pump work will add less than 1 psi to the vapor pressure. 

or about 0.2 deg F. From this, the system performance will be between the saturated 

and the 1 deg subcooled curves, depending on the liquid level in the tank. When 

the system heat load is doubled and the turbopump is left uninsulated, the system 

performance changes only slightly. This shows that the system is not sensitive to 

the insulation. The turbopump would have 4.36 Btu/sec heat leak for the uninsulated 

case and 0. 7 4 Btu/sec for the insulated case. These data were supplied by Pratt and 

Whitney to DeWitt Westrope on August 29, 1991 (CM No. NM0-077-05). 

MARTIN MARIETTA facing 

1-P-034 

1ge 7 MANNED SPACE SY~TEMS 

( ( 
( 

.. 

Backward Reclrc 

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3-P-034 

Page 7 MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

\ 
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3-P-034 0 age8 

Backward Recirc Un1nsu1ated 

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Natural Conv. Head Press. (/) 

>-

(/) 0+---------~----~----~ 

0 2 

Flowrate (Jb/sec) 

( 

-a- Sat liquid 

A- 10 Subcooled 

IV 20 Sub coo Jed 

A- 30 Subcooled 

-a- Pump line 

Y= Volumetric Quality 

MARTIN MARIETTA 

MANNED SPACE SY~TEMS 

( 
( 

The Forward Recirculation System 

The forward recirculation system is the one used successfully on the Saturn V S-11 

and S-IVB Stages as well as the Shuttle. It is very predictable because of this 

experience, its performance is shown on the next figure for the predicted heat load 

and the following figure for the uninsulated turbopump, doubled heat load. The heat 

loads were calculated as described earlier. It is seen that this system is more 

sensitive to heat load than the backward recirculation system, however, this system 

does not put vapor into the feedline which must be condensed later during tank 

prepressurization. For the one degree subcooled case, with the insulated turbopump, 

the flowrate is one lb/sec. While the vapor volume in the exiting flow is 60°/o of the 

total volume, this is not critical because the vapor is reentering the main propellant 

tank. The flowrate is quite sensitive to the heat leak, unlike reverse recirculation. 

The system is considered very predictable. We have a great deal of experience with it. 

If the engine test facility is provided with a flight type system, and the feedlines have 

heat leak in the same range as flight type feedlines, repeatability from engine test to 

flight vehicle will be excellent. The potential for a requirement for a future change is 

considered small, because of the extensive experience with this system. The system 

in the past has had some difficulty in the operational efficiency category because of 

two things; connector shorts and the tendency to go into propellant loading without 

properly setting up the system. This should not be a problem. The hazard introduced 

is the potential for leaks. The hardware is relatively complex, involving small lines, 

pumps, valves to prevent pump spinning during powered flight, and on-board 

disconnects; the number of disconnects required is dependent on the design. 

For predictable performance without the potential for future change requirements, 

the forward recirculation system is superior. The trade is between the hardware and 

maintenance costs of this system and the unknowns involved with any other. 

facing 

3-P-034 

Page 9 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
3-P-034 

Page 9 

( / 

, 

( 

FORWARD RECIRCULATION 

•STANDARD SHUTTLE RECIRC. PUMP PUMPING 

AROUND CLOSED PREVALVE. 

-HYDROGEN BLEED VALVE, PROBABLY ON 

ENGINE GAS GENERATOR SUPPLY LINE, 

WILL BE REQUIRED. 

-DISCONNECTS MUST BE PROVIDED FOR 

OUTBOARD LINES. 

-RETURN CAN BE MANIFOLDED. 

-ATTACHED CURVES SHOW MINIMUM 

SPECIFIED PUMP PERFORMANCE. 

-ATTACHED CURVES SHOW SENSITIVITY 

TO ENGINE TURBOPUMP INSULATION AND 

TO HEAT LOAD OF REMAINDER OF SYSTEM. 

-PERFORMANCE CAN BE IMPROVED BY 

INCREASING SYSTEM LINE DIAMETER, 

BUT NEED FOR GIMBALS RATHER THAN 

FLEX LINES MAY BE PENAL TV. 

MARTIN MARIETTA 

MANNEDSPACESY~TEMS 

( 
( 

3-P-034 

Page 10 

12 

" ..... 

(I) 

~10 

c. 

0 

L 8 0 

Q) 

5 6 

Ul 

Ul 

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0. 

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( 

Forward Recirc 

Insulated Turbopump 

'',, __ !'!( .. ,_ 

'""·· ,,.,, 

' 

Flowrate (lb/sec) 

2 

( 

-a- Sat liQUid 

• 10 Subcooled 

11-- 20 Subcooled 

A- 30 Subcooled 

- m Pump line 

Y= Volumetric Quality 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

---------
( 

~-P-034 

'J9 11 

Forward Rec1rc Uninsulated 

'"'

12 Turbopump Double Heat Load 

...... 

(/) 

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a. 

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Natural Conv. Head Press 

Flowrate {lb/sec) 

( 

-a- Sat liquid 

t- 1 D Subcooled 

IF 20 Subcooled 

~ 30 Subcooled 

. ~- Pump line 

Y= Volume Quality 

2 

MARTIN MARIETTA 

MANNED SPACE SY~TEMS 

( 
-

.. 

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The following two figures summarize the comparison between the candidate systems. 

The numerical compa~ison reflects the verbal comparison. 

facing 

3-P-034 

e12 

( 

MARTIN MARIETTA 

MANNED SPACE SY~TEMS 

( 
( 

I 

HYDROGEN FEEDLINE/ENGINE CONDITIONING 

Attributes 

•Predictability 

•Repeatability Eng. 

Test to Vehicle 

•Precedence 

No Bleed 

•Excellent 

•Poor 

•None 

•Impact on Engine I •None 

Design 

On-Board Bleed 

•Poor 

•Fair 

•Used on LOX on S-1 

•AddsLH2 BV 

•Impact on Feed ~·Potentially Severe TP ~·Potentially High TP & 

System/Eng. lnsulla- & FL Insulation Per- FL Insulation Performlion formance Aeqt. ance Reqt. 

•Impact on Eng. Test I •Potentially Large 

•Potential for Aeq'd I •Large 

Future Change 

•Operational I-very Simple 

Efficiency 

•Harzard Introduced I •None 

•Hardware Com- I •Low 

plexity 

3-P-034 

Page 12 

•Potentially Large 

•large 

•Simple - May limit 

Hold After Prepress. 

•Low 

•H2 BV's & Small 

lines Req'd 

Overboard Bleed 

to Facility 

•Poor 

•Excellent 

Backward 

Recirculation· 

•Fair/Good 

•Fair/Good 

•Rl-10 On S-IV Single! •None 

Engine Tests 

•Adds LH2 BV Potential for Turbopump Windmilling 

•Adds LH2 BV 

•Insensitive to TP & ~·Moderate TP & Fl 

Fllnsulalion Perform- Insulation Performance ance Reqt. 

•None •Moderate 

•Fairly Small I •Fairly Large 

•May Impact Facility ~·Requires Recirc. 

H2 Disposal System Pumps 

Forward 

Recirculation 

•Excellent 

•Good 

•S-11, S-IVB, Shuttle 

•Adds LH2 BV 

•More Sensitive Than 

Backward Aecirc. 

•Small 

·Small 

•Requires Recirc. 

Pump 

•Adds most leak 

Sources 

•Moderate-adds •Moderate - adds 

•H2 BV's Small Lines 

On·Bd & Gnd Disconnects Req'd 

leak· Sources leak sources 

•H2 BV's, Small Lines, •H2 BV's, Small Lines, 

Recirc. Pumps, On-Bd Recirc Pumps, Onbd 

Disconnects Req'd Disconnects Req'd 

·r,~&fiiVII~&t;Uiill 

MANNED SPACE SYSTEMS 
( 

3 ·Important 

2 • Somewh•t lmport•nt 

1 • Con .. ctered 

Attributes I 

Precedence 

Impact on 

Eng. Design 

Impact on I Insulation 

Impact on 1 Eng. Test 

Potential for 

Future I 

:Cb•ogt 

Operational 1 Efficiency 

Hazard 

Introduced 1 

Hardware 

Complexity 

3-P-034 

19 13 

Total 

I 

I 

Weighting 

Factor 

2 

-

3 

2 

2 

2 

1 

2 

3 

3 

1 

I 

LH2FEEDLINE 

CONDITIONING 

EVALUATION MATRIX 

Grade• • Aa4, 8:3, C:2, 0..1 f..O 

GradeiiScor•• Score • Orede • Weight 

No Bleed On Board Overboard Bleed I Backward I Forward 

Bleed to Facility Recirculation Recirculation 

- __...... .. --- -

46 I 32 I 53 I 31 I 55 

MARTIN MARIETTA 

MANNED SPACE SV('TEMS 

( ( 
---

THIS PAGE INTENTIONALLY LEFT BLANK 
"tJ 

, 

:0 

! 

2 

" "'0 

~ OJ 

...... 

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::2 

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c 

-! 

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( 

The following tables give the hardware costs of the options, derived by a bottoms-up 

analysis using Martin-Marietta data. The Design, Development, Test, and Evaluation 

costs would be additive to these. 

facing 

~-P-034 

'ge 14 

MARTIN MARIETTA 

MANNEDSPACESV~TEMS 

( / \ 
( 

Item 

Bleed Valve 

Line 

Elbow 

Line 

Elbow 

Line 

Elbow 

Line 

Elbow 

Check Valve 

Insulation 

Insulation 

Total 

3-P-034 

Page 14 

LH2 Options 

Option B • LH2 Natural Convection Drive 

Description Qty/PM Cost/Unit 

Similar to SSME 6 78,057 

1" ID X 6' Line W/4' Flex 6 62 

1" ID 45°, R/d >= 2.5 6 10 

1" ID X 2' 6 21 

1" ID 30° I R/d >= 2.5 6 7 

1" ID X 6' 6 62 

1" ID 75°, R/d >= 2.5 6 17 

1" ID X 2' 6 21 

1" ID 90°, R/d >= 2.5 6 20 

1" Low pressure drop 6 27,300 

All flex with TPS equivalent 6 804 

All lines with 1" SOFI equiv. 6 2,813 

---

Cost/PM 

468,342 

372 

61 

124 

41 

372 

101 

124 

122 

163,800 

4,824 

16,877 

655,158 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
Bleed Valve 

Line 

Elbow 

Line 

Elbow 

Line 

Line 

Line 

Line 

Tee 

Line 

Disconnect 

Line 

Elbow 

Line 

Elbow 

Line 

Elbow 

Disconnect 

Line 

Tee 

Disconnect 

Tee 

Insulation 

Insulation 

Total 

( 

3-P-034 

Page 15 

Option C - LH2 Overboard Bleed 

Similar to SSME 

1" 10 X 6' Line W/4' Flex 

1" 10 45°, R/d >= 2.5 

1" 10 X 2' 

1" 10 30°, R/d >= 2.5 

1" 10 X 6' 

1" ID X 4' (Inboard) 

1" 10 X 6' (Inboard) 

1" 10 X 1 0". (Inboard) 

1" 10 (Inboard) 

1" ID X 2' (Outbd w/out fill & drn.) 

For 1" Line (Outbd. w/out f & d) 

1" 10 X 2' (Outbd w/out f & d) 

1" I 0 goo, R/d >= 2.5 (Outbd 

w/out fill & drain) 

1" 10 X 8' (Outbd w/fill & drain) 

2" ID goo, Rid>= 2.5 (Outbd 

w/fill & drain) 

2" 10 X 6' (Outbd w/fill & drain) 

2" ID X goo, R/d >= 2.5 (Outbd w/f & d) 

1.5" 10 Vehicle/Ground 

1.5" ID X 12' (Across tank bottom) 

12.5" X 1 

1.5" 

2.5" 

All flex with TPS equivalent 

( 

6 

6 

6 

6 

6 

6 

6 

6 

6 

6 

6 

1 

6 

6 

6 

6 

6 

6 

1 

6 

12 

1 

6 

6 

246 

120 

184 

120 

26,000 

234 

150 

26,000 

250 

804 

11,251 

1,473 

723 

1,105 

723 

26,000 

1,405 

1,800 

25,000 

1,500 

4,824 

67,507 

630,578 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( 
Item 

Recirculation Pump 

Shutoff Valve 

Elbow 

Line 

Elbow 

Bleed Valve 

Line 

Elbow 

Line 

Elbow 

Line 

Elbow 

Inlet Fitting 

Manifold 

Manifold 

Insulation 

Total 

3-P-034 

Page 16 

( 

Option D - LH2 Forward Recirculation 

Description Qty/PM 

Same as STS(MC281-0030-0002) 6 

1" dia equiv flow area (MC284- 6 

0395-0051) 

1" 10, Aid >= 2.5 6 

1" 10 X 3' Line Flex 6 

1" ID, R/d >= 2.5 (to dlstream end 6 

of prevalve) 

Similar to SSME 6 

1.125" ID X 6" Line W/3'Fiex 6 

(gimballing capable) 

1.125" ID 75°, Rid>= 2.5 6 

1. 125" ID X T Line 6 

1.125" 10 75°, r/d >= 2.5 6 

1.125" 10 X 5' Line 6 

1.125" 10 75°, R/d >= 2.5 6 

To Tank Bottom 1 

4 to 1 1 

2 to 1 1 

All lines with 1" SOFI or equiv. 6 

-

Cost/Unit Cost/PM 

137,500 825,000 

75,125 450,750 

20 122 

31 186 

20 122 

78,057 468,342 

78 450 

23 138 

88 525 

23 138 

63 375 

23 138 

8,000 8,000 

10,000 10,000 

10,000 10,000 

3,692 22,151 

1,796,437 I 

I 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Item 

Recirculation Pump 

Shutoff Valve 

Disconnect 

Line 

line 

Line 

Line 

Bleed Valve 

Insulation 

Total 

3-P-034 

... 1e 17 

--

Option F - LH2 Backward Recirculation 

Description Qty/PM 

Same as STS (MC281-0030- 6 

0002) 

1" dia equiv flow area (MC284- 6 

0395-0051) 

For 1" Lines 6 

1" ID X 6' 6 

1" 10 X 7' 6 

1" ID X 4' 6 

1 II ID X 3' 6 

Similar to SSME 6 

All lines with 1" SOFI or equiv. 6 

( 

Cost/Unit Cost/PM 

137,500 825,000 

75,125 450,750 

26,000 156,000 

62 372 

72 433 

41 248 

31 186 I 

78,057 466,342 I 

4,219 25,315 

1,924,646 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( 
( 

-·· 

3~P-034 

Page 18 

( 

Task 3-P-034 

Addendum 

Weight Estimates lb. 

• On-Board LH2 Bleed 93 

(Option B- LH2 Natural Conv. Driven) 

• LH2 Overboard Bleed 255 

(Option C) 

• LH2 Forward Recirculation 331 

(Option D) 

• LH2 Backward Recirculation 328 

(Option F) 

'' 

( 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Summary & Conclusions 

• Reference No-Bleed System will result in saturated LH2 in feedline and engine 

pump with vapor in engine pump and dry.lines downstream of engine pump. 

Convection path complicated by screen~ 

Analytical model and test program to anchor analytical model required. 

Warm-up after prepressurization increases saturation pressure 5 psi/minute. 

Would require depressurization of tank, repressurization for very short hold. 

(Engine start pressure not yet defined.) 

• On-Board Bleed has low flowrate, hydrogen quality in turbopump will be poor 

(80°/o vapor by volume). If this is satisfactory, would allow improvement in hold 

after prepress relative to no-bleed system. 

Test program required. 

Slight improvement in performance compared to no bleed. 

• Overboard bleed has adequate performance after prepressurization. Hold time 

limited due to loss of LH2 at 1.21b/sec per engine. 

Hardware complexity a disadvantage. 

SSME manufacturer uses this system, in principle, for single engine tests. 

Should be retained for further study. 

• Backward recirculation did not appear advantageous. 

Provides good engine/pump chill. 

Introduces large volume of vapor into feedlines. 

Hardware complexity a disadvantage. 

• Forward recirculation: 

Predictable, good experience with systems. 

Hardware complexity a disadvantage. 

Together with overboard bleed, provides best engine/pump chill. 

Provides best performance (best chill, no hold time limitation). 

3-P-034 

'1ge 19 MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( 

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( 

Executive Summary 

NASA Statement of Work: . 

"Establish LH2 tank pressure limits vs. flight time considering engine start 

and NPSP requirements, potential pressure stabilization of tank during 

max airloads, structural weight considering proof test requirements and 

performance. Also consider ascent venting criteria." 

• Current autogenous flowrate results in high tank pressures that will set 

the vent valve relief setting at -60 psig. 

• Structural impact of ~ 7000 Ibm due to high tank pressures. Pressure 

can be reduced with decreased autogenous flowrate. Proposed flowrates 

of 1.1 Ibm/sec/booster and 0.91bm/sec/sustainer still results in -1500 Ibm 

payload impact. 

• NPSP consideration indicate that the minimum ullage pressure to 

satisfy NPSP requirements will be -31 psig @ MECO. 

• To reduce further the structural impact an alternate pressurization 

system will be required, i.e., flow control valves, step orifice control. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
( 

1.0 Summary 

( 

Task Number 3-P-038 

LH2 Tank Pressure limits 

Baseline system results in very high tank pressure during ascent. Tank impact 

is -7000 Ibm. This can be reduced by selecting booster and sustainer pressurization 

flowrates of 1.0 Ibm/sec. The tank impact is then reduced to -1500 Ibm. 

NPSP requirements set the lower pressure requirement at -31 psia. 

2.0 Problem 

Assess LH2 tank pressure limits. 

3.0 Objective 

Determine tank and system impacts for the reference configuration. 

4.0 Approach 

The approach to performing this study was: 

• To generate ullage pressure vs. time for the reference configuration and assess system impacts. 

• Develop system to minimize impacts to the tank and still maintain adequate NPSP margin. 

5.0 Results 

The results of this study are attached. The main results of the study are listed below. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

6.0 Conclusions and Recommendations 

The baseline autogenous flowrate of 1.41bm/sec results in high tank pressures that impact 

the tank structure by -7000 Ibm and maintain ample margin for NPSP requirements. Reduction 

of the autogenous flowrate to 1.0 Ibm/sec reduces the impact to the tank structure to -1500 Ibm 

and provides adequate NPSP. 

7.0 Supporting Data 

8.0 Attachments 

Study "Task Number 3-P-025, L02 Tank Pressure Limits," dated 12120/91. 

MARTIN MARIETTA 

MANNED SPACE SY~TEMS 

( ( 
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( 

3-P-038 

Page 1 

Approach 

Generate baseline ullage pressure for HLL V and 1.5 Stage 

Generate issues with system and structural impact 

Trade residuals with engine NPSP requirements and engine cost sensitivities 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( ( 
( 

RESULTS 

• Baseline autogenous system results in high(60 psig) ullage pressures. 

Structural evaluation shows this to be - 7000 Ibm payload impact 

3-P-038 

Page 2 

5' 

~ 

~ 

~ ;..I 

~ 

LH2 Tank Weight Impact 

~:::: l I I I l lsalJ~vJJ~~L-Cl-0 I 6000 I I I I I I I 

::::: O;J:tyt~mlt~!il=t.-t.l.=tl--=~• ' : 

I 0000 I I I t I ~-1--+--H-· +-t I I f 

eooo IMtntmur N~sP --1-1--UI-t-~---f· I 

:::: t I I I I I t I I t 

20004-~~~~~_.,~~4-~~~+-~-4---~~ 

0 . 

20 30 40 50 60 70 80 90 

PRESSURE (pslg) 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

70 

65 

3 

en 

P.. 60 . ..., 

t! 

~ ~~ r.l 

8: 

r.l 

u 50. <I 

'""' ..J 

!:) 

~ 45. 

~ 

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3 40. 

I 

• 

Baseline Ullage Pressure 

~ __ --- HLLV Engin~t 

,-.1 ~--------- .... _ 

--- , 1.5 :ge \gine Out 

1.5 Stage 

-,-

HLLV 

1.416/sec/Engine Pressurant 

Flowrate 

Nan navnte- 1.4 Jbml.ealtllll 

3~. ~--~----~--~~--~----~--~----r---~----+---~--~ 

3-P-038 

Page3 

00 100. 200. :mo. 

TIM[ PROM LIFTOfF (T -0). (SP:C) 

( 

400. 500. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 

( 
( ( 

Results 

• NPSP results generated by AI indicated a minimum ullage pressure of -36 psig. 

This analysis did not take into consideration the benefit of throttling on NPSP requirement. 

This reduces the NPSP requirement from 13.1 to 7.6 psi. This effect lowers the minimum 

ullage pressure to -31 psig. 

• To reduce the structural impact the autogenous flowrate can be orificed to minimize this 

impact while still providing adequate NPSP margin. The proposed flowrates are 

1.1 Ibm/sec/booster engine and 0.9 lbm/sec/sustainer engine. This still results in a payload 

impact of -1500 Ibm. 

( 

3-P-038 

Page 4 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
( 

Proposed Ullage Pressure 

...... 

IV 

en 

55-

e .,'s- L-

:J 

en 

en 

., 

LQ. 

35-

25. 

3-P-038 

Page 5 

-189. 0.0 190. 

Flowrates (lb/sec) 

.9 Booster Engines 

1 . 1 Sustainer Engines 

Minimum Relief Pressure 

1.5 Nom 

" 

200. H9. 

TIme from I If torr (sees> 

( 

1.5 En 

/ 

ng Out 

498. 

MARTIN MARIETTA 

MANNED SPACE SV~TEMS 

( 
( ( 

Summary & Conclusions 

• Current autogenous flowrate results in high tank pressures that will set the vent valve 

relief setting at -so psig. 

• Structural impact of -7000 Ibm due to high tank pressures. Pressure can be reduced 

with decreased autogenous flowrate. Proposed flowrates of 1.1 Ibm/sec/booster and 

0.9 lbm/sec/sustainer still results in -1500 Ibm payload impact. 

• NPSP consideration indicate that the minimum ullage pressure to satisfy NPSP requirements 

will be -31 psig@ MECO. 

• To reduce further the structural impact an alternate pressurization system will be required, 

i.e., flow control valves, step orifice control. 

3-P-038 

Page 6 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
.. 

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PREtiEDING PAGE BLANK NOT FflM£o 
( 

Executive Summary 

NASA Statement of Work: 

"Select optimum LH2 tank pressurization system based on tank pressure limits 

and specified reference trajectories and considering safety, reliability,operabillty, 

simplicity, weight, Including residuals, and cost." 

Approach: 

Generate baseline pressure profiles for HLL V and 1.5 Stage 

Generate Issues and concerns to reference 

Evaluate reference with structural and NPSP requirements 

Generate alternate approaches 

Results: 

1) Baseline fixed orifice system results in high ullage pressures during flight that 

have a substantial structural weight Impact. NPSP requirements are only -31 pslg 

which results In too much margin. 

2) Structural weight impact can be reduced by reduction In fixed orifice flowrate to 

-1.0 Ibm/sec/engine. This still results In -1500 Ibm impact due to the high ullage 

pressure that exists during the first portion ~f the flight. 

3) Two approaches have been examined to reduce the initial tank pressure without 

Impacting the NPSP requirement. These are a flow control system and a step 

pressurization system. . 

4) Assisted customer In set-up and analysis of pressurization systems. . 

MARTIN MARIETTA 

MANNEDSPACESY~-~MS 

( ( 
•, 

3-P-039 LH2 Pressurization System 

Background 

Martin Marietta Manned Space Systems was initially assigned performance of the 

subject contract task at the beginning of cycle 0. Early task planning was completed 

and preliminary analysis was done, after which the Propuls~on Working Group made 

the decision to complete this task in-house. Martin Marietta continued to participate in 

the task in a review and advisory role. 

This report documents the planning and analysis work performed by Martin 

Marietta and includes a summary of the results provided to MSFC for their completion 

of this task. 

MARTIN MARIETTA 

MANNED SPACE SYSTEMS 
I 

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7/11/91 

3·P·039 

LH2 Tank Pressurization Syatam 

~select LH2 tank pressurization system (based on pressure limits from 3-P-038) 

and specified reference trajectories, and considering safety, reliability, operability, 

simplicity, possible integration with the core RCS system, weight including residuals, 

and cost." ·· 

Work Ststtmtnt 

The unquestionable advantages of using warm hydrogen gas to pressurize the 

LH2 tank and the potential use of hydrogen gas as a roll control propellant red~ce this 

study to ona of evaluating the advantages of warmer hydrogen as a presaurant, the 

possibility of obtaining warmer hydrogen from the STME, a control system evaluation, 

and evaluation of the pressurant diffuser for this system. The study should consider 

the flow within the pressurant lines and diffuser for both 1 112 stage and HLLV, and the 

manifolding configuration to allow use of hydrogen for roll control. 

Compute tank wall temperatures and pressurant weights for pressurant 

temperatures of o, 100, and 200 °F. In task 3-P-038, evaluate tank pressure capability 

for Increased tank wall temperature. With the STME project, evaluate the feasibility of 

obtaining higher temperature pressurant from the STME. Obtain control impulse 

requirements from vehicle dynamics studies 3·FM-02S "Generate FCS Requirements." 

Compare required impulse with Impulse available from hydrogen bled from 

pressurization system. Evaluate reliability effects of this increased control system 

complexity. Compute flow parameters of pressurant lines and pressurant diffuser and 

attitude control manifolds. Compute system cost and compare with independent 

reaction control system {RCS} cost. 

Joput pata 

Control impulse requirements from 3-FM·028 "Generate FCS Requirements." . , 

Engine characteristics regarding bleed f~ow temperature$ available. 

LH2 tank pressure limits data from task 3-P-038. 

Tank wall temperature model from MMC·Operatlons (0. Vaughn). 

Tank wall heat fluxes from task 3 .. FM..005. 

eroducts 

Study results showing advantages/disadvantages of pressurant temperatures 

·form 0·200 °F. 

1 
7/11/91 

Evaluation of use of H2 for attitude control ~. separate ACS with regard to 

advantages/disadvantages, reliability, operability, cost. · 

Une size and manifolding requirements for pressurf:ation/RCS system. 

.. 

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This figure (LH2 Tank Ullage Pressure vs. Time from Liftoff) is a preliminary 

analysis of LH2 Tank pressure profile for the reference trajectory. 

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