Orbit Transfer Rocket Engine Technology

b

,' I"

.,

} -J

I"_ENEORP

y /

A_=ROJ_---T

Orbital Transfer Rocket Engine Technology Contract NAS 3-23772 Advanced Engine Study Task D.6 Final Report NASA CR-187216 June 1992

Prepared For. National AeronaulJcs and Space Administration Lewis Research Center 21000 Brookpark Road Cleveland, Ohio 44135

(NASA-CR-18?2|6} ROCKET ENGINE ENGIN_ SIUDY 1988 - Mar. 2_2

ORBITAL TECHNOLOGY: Final RepoFt, I990 (GenCorp

N92-32129

TRANSFER ADVANCED Oct. Aerojet)

Unclas

p

6]/20

Propulsion Division

0108999

NASA

ORBITAL

TRANSFER

ROCKET

ADVANCED TASK

ENGINE

D.6 FINAL

Prepared

Warren

STUDY

REPORT

By:

For:

Aeronautics and Space Administration NASA Lewis Research Center

Contract

No. NAS

June

RI'F/D041735a-C

TECHNOLOGY

R. Hayden

Prepared

National

ENGINE

1992

3-23772

NASA

CR 187216

Aerojet

2459-31-1

REPORT DOCUMENTATION

:J ,.,.,,d _¢mn.um,,..kl

PAGE

=_

i. Report No. NASA 4.

Title

CR And

187216

7 Recipienrs Catalog No. 5. Report Dale

Subtitle

Orbital

Transfer

Adwmced Task I).6 7

2. Government Accession No,

Rocket

Engine

Engine

Study

Final

June

Technology Report

Author(s) Warren

1_992

6. Performing Organization

Code

8. Performing Organization

Report No.

R. Hayden 10. Work Unit No.

9. Performing Organization Org.

9962,

Acrojet

Bldg.

Name and Address

2019

11. Contract or Grant No.

TechSystems

Sacramento,

NAS

3-23772

CA 95813 13. Type ot Report and Period Covered Final

12. Sponsoring Agency Name and Address

Oct NASA

I,cwis

('lcvcland.

15

Research

Ohio

Supplementary

Center

1988

to Mar

14. Sponsoring

1990

Agency Code

44135

Notes

16. Abstract This

report

concepts

documents

7.5 K Ibf to 50K versus The

the 6+1 major

Also,

,ind 485.2

scc

diameter

Continued

was

10%

engine

lack

Ibf to 305

with

up to a mixture

thrust

a mixture

ratio

operation

to include

in. exit more

engine

range

and operation

unit production,

and factors

contractors

ratio

diameter

The

basic

change. ratio

of

13.

Ibf.

Packaging

at a mixture in other

readily

platinum

An initial

envelopes will

requirements

engine

modeling

control

system.

from

joint

assessrnent

with

Key

Words

Rocket

(Suggested

Engine,

Throttling

by

sec

Lunar

Engine,

Mission,

Moon

Cryogenic

Mining,

Space

consideration. of the

Propellants, Unclassified

Transfer

- Unlimited

Vehicles

19.

Security

Classif.

(of

this

tlnclassified NASA

FORM

report)

20. Security Classif. (of this page)

21. No. of pages 27

Unclassified 1626

OCT

!

86

/J

7

rate of 4

at 20K

120 in. length/53

18. Distribution Statement

Author(s))

an Aerojet

A throttle

issues.

17

the 20:1

injector

Ibf, 487.3

be an important

contractor

12+1

costs.

accommodated

at 7.5K

from

of

in vehicle/engine

for bafllcd

varied

thrusts ratio

life cycle

all performance

design

in developing

engine

the issues

are 483. I sec

Engine

at five

to resolve

the baseline

1200.

contractors

obtained

of meeting

By using

with

contractor/engine

was

engine

predictions

at 50K

prime

capable

operation

Performance

vehicle

a 20:1

up to mixture

stable

of 6 and an area

in. length/136

is reconnnended

margins.

prime

data

shown

of 10 without

ratio

predicted.

prime was

and vehicle

design

over first

vehicle cycle

thermal

indicates

is also

throttling

expander

allowed

Parametric

for DDT&E,

of contact

propellant

for the use of NASA

engine

generated

MLETS)

to 100%

at 7.5K

also

study

and Phobos.

evaluated

margin

(named

Ibf with

Mars,

life due to the high

thermal

code

at 50K

work

was

dual

operational

increased

from

task

and opcration

sinu]lalion

to 5 seconds

data

Aerojet

long

the

a separate

of the study

rcquircn]enl

transient

LOX/LH2

to the Moon,

Cost

baseline

an expected

vtmslruction

exit

Ibf.

nominal.

The

throltling

missions

limitation

interfaces. with

an advanced

for manned

22. Price

interface

Ibf, in.

FOREWORD

This

report

planning

and

documents

design

for manned

of a hydrogen

data

base

cycle

within

mandated.

There

was

(OTV)

engine

expected

to generate

contractors

developing

The

baseline

range

lbf OTV

on the study technology

Mars.

the thrust

the 7.5K

on the

prime

and

was

constraints

Vehicle

study

engine

of 7.5K

engine

served

to focus

developed

technology

over

some

lbf to 50K lbf.

it within past

The

design,

or tradeoffs

the

program

concepts uses

preliminary

no comparison

useful

an

with

a limited

decade

sponsored

other

design

by the

by NASA

Lewis

Center.

Tile terms

Chemical

interchangeably engines for |he

study

dependent

Transfer

Research

Phobos,

cycle

was

These

highly

Orbital

the

parametric

for the vehicle

moon,

expander

engine

cycles.

range

to the

for starting

expander engine

information

missions

form

an engine

in this report

developed I,unar

Vehicle

Transfer

under

return

(LTV)

although

the CTP

mission

and

Propulsion

Lunar

(CTP)

the

OTV

program.

Excursion

engine

The

is designated

engine may

specific

(LEV)

OTV

be just

engine one

application

as a LTV/LEV

Vehicles

and

engine.

are expected

are

used

of several

of a CTP

engine

The

Lunar

Transfer

to use

the

same

basic

was

very

limited.

engine.

Interaction Interface

with

requirements

A Vehicle

Studies

This

study

the Design

and

of the study served

and

and

as senior

The

period

feedback

were plus

was

gleaned

direction

initially

Parametric

Subtask

by Jerry under

[/I)II417..55a

FM

prime

from

the

Pieper.

contractors NASA-MSFC

The

the direction report

throughout

for this study

°o° RI'

vehicle

sponsored

Phase

NASA-LeRC.

of the final

Engineer

of performance

the

primarily

from

directed

preparation Project

from

Ul

was

of Judy done

the period

was

work

October

was

continued

Schneider.

by Warren

Completion

ttayden

of performance.

1988

to May

through

1990.

who

TABLE

OF CONTENTS

Section

Title

1.0

Summary

2.0

Introduction

3.0

1 and

3

Background

2.1

Background

3

2.2

Scope

9

2.3

Relevance

2.4

Significance

to Current

Rocket

Engine

14

Technology

16

of the Program

18

Discussion 3.1

3.2

3.3

3.4

3.5

Design

and

Parametric

3.1.1

Engine

and

3.1.2

Power

3.1.3

Performance,

Engine

Cycle

Design

3.2.2

High

42 Mass

and

Envelope

Variation

71

Parameters

92

Studies

92

for 20:1 Throttling Mixture

Vehicle/Engine

Ratio

Study

3.3.1

Engine Design, Cost Estimates

3.3.2

Engine

Production

3.3.3

Mission

Related

3.3.4

In-Space

3.3.5

Engine

3.3.6

Issues

3.3.7

Weight

3.3.8

Parametric

Engine

20

Definition

Balance

Requirement

3.2.1

18

Analysis

Development,

Design

3.4.1

Dual

Propellant

3.4.2

Engine

Control

3.4.3

Engine

Components

and

3.5.2

Turbopump

Servicing

for the

135

Lunar

155 Mission

159 100

Basing

166

Requests

166

Update

171

Expander

Chamber

(DDT&E)

145

for Space

Data

Thrust

Engineering

145

Throttling

Penalties

3.5.1

and

Costs

in Engine

of Critical

Test

Cost

Requirements

Identification

135

Coordination

Maintenance

Baseline

108

Operation

Cycle

Baseline

171 171 176

Technologies Technology

Technology

179 179 186

TABLE

OF CONTENTS

Section

4.0

5.0

(cont.)

Title

3.5.3

Heat

3.5.4

Proportioner

3.5.5

Oxygen

3.5.6

Extendible/Retractable

3.5.7

Integrated

3.5.8

Engine/Vehicle

Conclusions

and

Exchanger

188

Technology Valve

Cooled

190

Technology

190

Nozzle Nozzle

Control

and

Health

191 Monitoring

System

191 191

Synergisms

Recommendations

192

4.1

Conclusions

192

4.2

Recommendations

193

References

195

Appendices A.

Detailed

B.

l-'ower

C.

Chug

RIH'/D0417.55a-l"M

Engine Balance Stability

Thermal at High of OTV

A-1

Analysis Mixture

B-1

Ratio

20K lbf Advanced

V

Engine

C-1

LIST OF TABLES

Table

No.

2.1-1

Technology

2.2-1

Engine

3.1-1

Advanced

3.1-2

Expected

3.1-3&4

CTP Engine Materials Selection, in a High Radiation Environment

3.1-5

Advanced Engine Pump Section

Study

Turbopump

Design

Point

-

46

3.1-6

Advanced Turbine

Study

Turbopump

Design

Point

-

47

3.1-7

CTP

Engine

Power

Balance,

Thrust

= 20K lbf, MR = 5

56

3.1-8

CTP

Engine

Power

Balance,

Thrust

= 20K lbf, MR = 6

57

3.1-9

CTP

Engine

Power

Balance,

Thrust

= 20K lbf, MR = 7

58

3.1-10

CTP

Engine

Power

Balance,

Thrust

= 50K lbf, MR = 5

59

3.1-11

CTP

Engine

Power

Balance,

Thrust

= 50K lbf, MR = 6

60

3.1-12

CTP

Engine

Power

Balance,

Thrust

= 50K lbf, MR = 7

61

3.1-13

Performance Loss Accounting for Various (7.5K lbf, 20K lbf, and 50K lbf)

3.1-14

LTV/LEV

3.1-15

7.5K

3.1-16

7.5Klbf Thrust Preliminary Complete Engine

Engine

3.1-17

Preliminary

Weight

3.1-18

Advanced

Flight

3.1-19

Advanced Normalized

Engine Design by the Throat

Study Engine Radius

3.1-20

Advanced

Engine

Design

Study

Normalized

3.1-21

Advanced

Engine

Design

Study

Basic

3.2-1

Study

3.2.1-1

Advanced

3.2.2-1

Engine

3.3.1-1

LTV/LEV

3.3.3-1

CTP

F,P I'/I_

17.._,5a

I:M

Goals

System

for the New

Requirements

Engine

Study

Results

Engine

Engine

Engine

Thrust

20 21

Chamber

Thrust

Assembly

Levels

Estimate Weight

Cost

Operations

vi

74

78 Estimate

80

Estimates,7.5K

lbf Engine

Estimates

at 20:1 Throttle

23 & 24

77

Turbopump

DDT&E

Removal

Rationale

Study

Requirements

Engine

Balance

10

Selection

Weight

- Engine

Expander

Goals

Weight

System Engine

5

Flowrates

Engine

Nozzle

Engine

Parametric

Propellant

lbf Preliminary

Power

and Thrust

of the

Engine Section

Baseline

OTV

80 82

Contour

88

Nozzle

Engine

Contour

Dimensions

Variation

91 93

Design Down

89

Specification

Condition

98 105 137 153

LIST OF TABLES

Table

(Cont)

No.

3.3.4-1

CTP 7.5K

3.3.4-2

Space-Based

3.3.4-3

CTP

3.3.8-1

Representative

3.4-1

CTP

3.4-2

Engine

3.4-3

Component

_¢_._x..._-._

Engine Space Maintainable lbf Thrust Engine CTP

Servicing

Engine

Engine

Maintenance

Operational

Functions

Engine Valves

Operation State

and

Electrical Sensors

Power

157

Functions

158 Requirements,

for Engine

Control

Watts

167 172 173

Sequence During

156

Components

Engine

vii

Operation

175

LIST

OF FIGURES

Ejgt_r_t:_N_9. 2.1-1

Dual

Expander

2.2-1

Start

Cycle-Space

2.2-2

Task

Order

3.1-1

OTV Engine Dual Expander 7.5K Ibf Thrust Engine

3.1-1A

Our Regeneratively As An Added Heat Stability

Cycle

Schematic

Transfer

Vehicle

D.6 - Advanced

Cooled Transfer

Damping

7.5K lbf OTV

3.0K

3.1-5

7.5K lbf Thrust

3.1-6

OTV Engine Turbopump

lbf Design:

Advanced

Activities

12

Cycle,

Series

Flow

19

Schematic,

Baffles Provide a Dual Function Surface as Well as a Combustion

Cooled

OTV

Preliminary Side Engine

Can

(Top

Study

3.1-13

Hydrogen

Power

3.1-14

Logic

in Computerizing

3.1-15

The

Oxidizer

3.1-16

The

Fuel

3.1-17

7.5K

3.1-18

Power

3.1-19

CTP

3.1-20

Dual Expander Schematic

3.1-21

Advanced

Engine

3.1-22

Predicted

Response

3.1-23

Predicted

Response

3.1-24

Advanced

Engine

3.1-25

Performance

3.1-26

Space

31

Concept

33

Layout

Design

Oxygen

29

Gimbal

Extension

- Regenerator

3.1-12

and

View)

Design

Preliminary

-

35

Details

37

Drawing

Channel Engine

44

Operating

45

Modeling Power

Path

Balance

Engine

Model

Transfer

Loss

from

51

Balance

52

OX Side

Preliminary

for 20K lbf Engine

for MLETS

Cycle

49

Efficiency

First

Engine

Results

48

Efficiency

TPA

a Power

Results

OTV

TPA

- Combined

is Balanced

Side Uses

Balance

- Combined

Balance

Side

Ibf Thrust

25A

28

Circuit

Nozzle

Engine

Component

Balance Design

53 Power

Balance

Engine

54 55

at MR = 6

64

Analysis

Advanced

Study

66

Alternate

Thrust

68

to 10% Throttle

Up Command

69

to 10% Throttle

Down

70

Study

Stability

at Rated

Command

Performance

72

Accounting

73

Vehicle

Propellant

.°° 55ad:M

Study

Showing

3.1-11

RI'T/1)0417

11

Engine

Flow

Sketch

Fuel

7.5K lbf Thrust Envelope

7

Version)

Device

TCA

3.1-4

Flow

Engine

Injector/Baffle

OTV Engine Attachments

(Parallel

Vlll

Flowrate

vs Engine

Thrust

76

LIST OF FIGURES

Figure

(Cont)

No.

3.1-27

Delta

3.1-28

Advanced

3.1-29

Engine

3.1-30

Advanced

3.1-31

Change

3.1-32

Engine

3.2.1-1

Predicted

Payload

vs Nozzle

Engine

Study

Thrust/Weight Engine in Engine Half

Percent

Bell

Parametric Ratio

81 Weight

Summary

vs Thrust

Study

Thrust

Length

with

85

Versus

Weight

86

Thrust

87

Section

OTV

90

Off-Design

Performance

Meets

10:1 Throttling

Operating

Requirements

3.2.1-2

Estimated

Head

3.2.1-3

Advanced

Engine

3.2.1-4

NARloy-Z,

3.2.1-5

Hydrogen

3.2.1-6

Enthalpy

3.2.2-1

NARIoy-Z

3.2.2-2

Blanching

and

3.2.2-3

Uncoated Reduction

NASA-Z Cylinder After Oxidation and Test Cycle - 26 Slot Area (1.8% Strain)

3.2.2-4

Uncoated NASA-Z Center Area (>2.7% Variation

84

Loss

Due

to Cavitation

Study

Wrought, Circuit

- Engine

Operating

Range

100 101

Pickup

- Oxygen

Exposed

97

Properties Enthalpy

Pickup

95

102

Circuit

104

to Oxidizing/Reducing

Environments

110

Cracking

111

Cylinder Strain)

of Combustion

After

Gas

Oxidation

Species

with

113

and

Test

MR and

-

114

Temperature

116

Variation in Specific Oxygen/Hydrogen

Impulse Propellants

with Mixture Ratio - 7.5K lbf Thrust

3.2.2-7

Variation in Specific Oxygen/Hydrogen

Impulse Propellants

with Mixture Ratio for - 20,000 lbf Thrust Engine

119

3.2.2-8

Variation in Specific Oxygen/Hydrogen

Impulse Propellants

with Mixture Ratio for - 50,000 lbf Thrust Engine

120

3.2.2-9

Advanced Flowrate

3.2.2-10

Engine

Maximum

Wall

3.2.2-11

Engine

H2 Bulk

Temperature

3.2.2-12

Engine

Maximum

Wall

3.2.2-13

Advanced Flowrate

RI'I'/I)O417.,_a-I:M

Engine Study and Thrust

Engine Study and Hydrogen

- Mixture

Temperature

Ratio

Versus

118

LOX

for Pc = 2000

Rise for Pc = 2000

Temperature

for Engine

psia,

psia,

for Pc = 2000

122

MR = 12

MR = 12

psia,

MR

- Mixture Ratio Versus Hydrogen Temperature at Maximum Thrust

ix

-- 12

123 124 126 127

LIST

OF FIGURES

(Cont)

l.'igure No. 3.2.2-14

Engine Hydrogen MR= 10

Bulk

Temperature

Rise - Pc = 1917

3.2.2-15

Engine Maximum MR = 10

Wall

Temperature

- Pc = 1917 psia,

Engine

Pressure

Drops

- Pc = 1917 psia,

129

130

MR = 10

Wall

3.2.2-18

Engine Maximum MR = 12

Hydrogen

3.2.2-19

Engine

3.3.1-1

LTV/LEV

Engine

Development

- Design

138

3.3.1-2

LTV/LEV

Engine

Development

- Fab

139

3.3.1-3

I,TV/LEV

Engine

Development

- Test

140

3.3.1-4

LTV/LEV

Engine

Development

- Total

3.3.1-5

LTV/LEV

Propulsion

3.3.2-1

LTV/LEV Nth Unit

3.3.2-2

Drop

for Pc = 1500

132

Engine Maximum MR = 12

Pressure

Temperature

128

psia,

Temperature

psia,

for Pc = 1500

for Pc = 1500 psia,

psia,

134

MR = 12

Program

133

Cure

Costs

141

Schedule

142

Engine at Complexity Production Cost

= 1.1 x RL -10 Engine

146

LTV/LEV Nth Unit

Engine at Complexity Production Cost

= 1.2 x RL -10 Engine

147

3.3.2-3

LTV/LEV Nth Unit

Engine at Complexity Production Cost

= 1.3 x RL -10 Engine

148

3.3.2-4

LTV/LEV Nth Unit

Engine at Complexity Production Cost

= 1.4 x RL -10 Engine

149

3.3.2-5

LTV/LEV Nth Unit

Engine at Complexity Production Cost

= 1.5 x RL -10 Engine

150

LTV/LEV

Reference

LTV Initial (26.2 MT)

Weight

in LEO

Lunar

3.3.6-1

CTP

3.3.8-1

0.030--in.

Columbium

3.3.8-2

0.050-in.

Carbon-Carbon

3.4.2-1

Control Engine

FM

Mission Engine

Program

161

Concepts

3.3.5-3

RI'I/I..N¼17.S5a

System

(1 Burn)

Versus

Isp

at Fixed

P/L

Profile Dual

162

163

Propellant

Expander

Nozzle Nozzle

Effectiveness-Parallel

Wall

Temperature

Wall Flow

X

Cycle

Temperature

Dual

Expander

165 vs Position

169

vs Position

170

Cycle

177

LIST OF FIGURES

l:igure

(Cont)

No.

3.5-1

Microchannel Thrust Load

3.5-2

Modified

I-Triplet

Injector

3.5-3

Modified

Injector

Element

3.5-4

Water

3.5-5

Structural

3.5-6

Dual

RI "1"/I )1.1417 _$,S,.i-I;M

Flow

Spool

Test Specimen Design)

Patterns Property Hydrogen

(Based

lbf

182

Elements

183

Pattern

184

for I-Triplet Variation

with

Turbopump

X

on 7.5K

i

Element Temperature

185 187 189

1.0

SUMMARY

The objective and

parametric

Parametric at five

data

design

engine

sure)

and

studies

The

formance

data

issues.

lack

and

propellant

NASA

expander

reduces

baseline.

This

lbf thrust

range,

baffled

cycle

needed

Simulation

(MLETS)

the basic

over use

code,

valve

throttling

range.

A throttle

predicted

using

the TUTSIM

Performance ratio

of 1200

50K lbf. 291.8

RlYl./iX_417.SSa

Ibm

at 7.5K

seconds engine

lbf, 486.3

free

psia

circuit

two

pres-

variation

to

parametric limitation

perof the

study

interface

for higher

and

pressure can operate

as a supplement

is predicted

Modified to be stable

is expected

of 4 to 5 seconds

from

chamber

accepted over

7.5K

high

to earth

the

design lbf to 50K

mixture

origin

Engine

at thermal

Transient

equilibrium, linear

to 100%

system

between

the

Liquid

gas

as the

at the

the

pressure

flow

to be nearly 10%

to drive purge

hydrogen and

of the dual

oxygen

for a helium

the

chamber

development

(400°F)

examined

the Aerojet

of thrust

over

the

was

code.

specific

at 7.5K

lbf, 484.3

weight

excluding

lbm

chamber

vehicle/engine

of the need

was

by delivered

dry

One

in the

heated

splitting

operation

dynamics

as measured

is 483.1

Predicted

rate

included

response

some

various

history

uses

oxygen

with

this cycle

control

year

a 20:1 range,

of lunar

engine

task.

variant,

a 2000

on an analysis

engine

to assess

assembly,

can maintain

for efficient Based

and

injector

can throttle

propellants.

costs.

cycle

a design

a cycle

study

to 100 psia

of the NASA

production

on the hydrogen

this study

the

unit

This

allows

the

the NASA.

cycle

6+ 1 to 12 + 1. The

included

an eight

engine.

expander

psia

input

primes

have

demands

During and

LeRC

This

pressure

operation.

ratios

vehicle

and

level.

Aerojet

first

descriptions

contractors

In addition,

(2000

in preparation

The

system

advanced

a nominal

as a follow-on

cycle

turbopump.

chamber

with

is recommended

Aerojet

from

lbf thrust

and

prime

a 20:1 range

to assist

DDT&E

vehicle

engine

7.5K lbf to 50K lbf.

ratios

initiative.

of contact

This

oxygen

space

transfer

over

20,000

advanced

for a LOX/LH2

from

expanded

plus

to develop

obtained

throttling

at the

was

Bush's

the

was

at mixture

done

study

was

ranging

operation

President

was

data

of engine

were

study

for use by space

thrusts

an evaluation

and

of the

at 20K lbf, and

I

impulse

seconds

at 20K lbf, and

gimbal 1362

at MR = 6 and

lbm

and

thrust

485.2

takeout

at 50K lbf using

an area seconds

at

structure

is

available

1.0, Summary,

technology.

Engine

retractable to 304.8

(cont)

section inches

extended.

envelopes varies

from

length/137

These

for a 1200:1 120 inches

inches

are large

area ratio length/58

diameter

engines

nozzle inch

using

exit

diameter

at the 50K lbf thrust

in terms

of envelope.

one extendible/

with

at 7.5K

lbf thrust

the nozzle

Packaging

will be an important

consideration.

The DDT&E well

accepted

used

assumptions

cost

typical

The total

first flight

in 1999. curve,

reference.

First

engine

range.

is about

dependent

Launch

of engine

DDT&E

cost

unit

and

return

on the mission

about

(ALS)

methodology

program.

numbers

$950M

costs

with

factor

Nth

of life cycle

RL-10 costs

life and maintenance

start

on production based

cost

was

scenarios

engine

which

as the

to be in the

to $4M

not feasible

and

numbers,

is expected

is $3M

MSFC

in FY91

on an RL-10

cost

to be as costed

in a NASA

a program

unit engine

the current

found

The program

and tests

are based

a complexity

mission,

Generation

a costing

System

fabrication

was

As references,

$6M.

using

production

thrust,

For the lunar

$6M to $12M engine

was generated

on the Advanced

program.

learning

data

and an OMS

as they

are

are still incompletely

defined.

The latest capable

version

of the dual expander

of meeting

all mission

MR = 12 operation.

All major

drive

are being

exchanger written,

technology and a vigorous

and health program. broadened interface

RPTI DO417-_Sa

evaluated

monitoring

technical

under

NASA

is in qualification program system

A continuation to include

performance

(ICHM)

vehicle

holds

promise

requirements

questions LeRC

for space

such

capability

shuttle made

prime/engine

oxygen

programs.

The

flight

issues.

2

platelet

heat

as this is

the integrated engine

but the scope

contractor

and

turbine

operations

the OTV

engine

20:1 throttling

as the 400°F

to develop

under

is recommended,

as a long-life

including

sponsored

start has been

of this work more

cycle

control

technology should

joint assessment

be of the

2.0

INTRODUCTION 2.1

AND

BACKGROUND

BACKGROUND

2.1.1

Orbit

Transfer

The NASA people

beyond

gram.

Over

been rected

main

contract

with

current

interest

the engine ments

The

integrated

health

gent

the

from

oxygen/liquid the basic middle

efforts

ponent, tem. tracts

this work

stage

RItI'/1XI417._5a

States

of tasks

Lewis

reported

beyond

was

pro-

LEO has

or Orbital

Flight

Research

herein

space

configurations,

as part of the Orbit Space

and

Center

Center

has

has di-

completed

under

a

of this contract on LEO-to-GEO

mission

and

system.

most

(500 starts, system.

technically

propellant

main

engine

engine

the reliability space

for both

basing

in

delivery

missions

to the

mission.

engine

advanced

and

and

effect

on

require-

multimission impulse

mandates resulting highest

and

from

these

performing

unmanned

engine

a sophisticated

in this century. and

The

redundancy

by specific

The

manned

an evolution

and

20 hours)

developed

been

Mars

as measured

control

has

payload

a manned

Also,

life goal

and

there

liquid

It could missions

strin-

serve until

as

the

century.

first

phase and

and

& Whitney,

of the NASA-LeRC

evaluate

engine

the most

innovative

system

Rocketdyne,

for this phase

of which

payloads

and many

for a variety

NASA

to emphasize

service

will be the

to generate

in 1982

been

monitoring

component, Pratt

of names

NASA-Marshall

on performance

long

The study

to move

of the United

developed

while

years

propulsion

very

of the next

vehicle

The work

Return

has

hydrogen

upper

studies

an emphasis

a premium

requirements

purpose

seven

Lunar

for a man-rated

throttling.

the inception

program.

development

place

for a vehicle

LeRC.

in the

capability

since

technology

Over model

Technology

concept

has had a number

development.

NASA

Engine

some

has been

for the vehicle

engine

the mission

(LEO)

(OTV)

this concept

(OTV)

responsible

orbit

of a general

1982,

Vehicle

has had

the vehicle

concept

Since

Transfer

earth

the years,

but the basic persisted.

low

Vehicle

of the notable

and work. was

levels

technology

program

Aerojet the dual

TechSystems initiated propellant

consisted

concepts

for an advanced

Aerojet

3

sponsored

O2/H2 were

several

at the

expander

subcom-

propulsion

each new

of

sys-

awarded

concepts cycle.

conduring

This

cycle

2.1.1,

Orbit

improves

Transfer

Vehicle

(OTV)

the conventional

oxygen

hydrogen

for use as working

and higher

chamber

fluids

I work

oxygen

an engine

(ICHM).

the existing

RL-10

engine

changes The

capable

impact

several

thrust.

This

thrust

from

that range. number

was

but three

straints

are considered. for the

LEO-to-GEO mission propulsion

RIrf/DOi17.55a

from

requirements.

The

to generate

may

mission

missions

may

very

set the engine

affecting

will

at least such

each.

one or two

thrust

engines

factors

will

as length

the current

will

in

engines.

is the be con-

baseline

An unmanned

of engines

vehicle The lunar

in the

set is established.

in the Spring

on the Advanced of 1990 with

Engine

this final

4

Study

report.

began

to

to fall within

engine

of these

the number

system

ranging

likely

two

mission

is engine

for engines

is very

when

solar

to change

baseline

vehicle

once

and gives

requirements

later in this report,

use only thrust

likely

thrust

System

defined.

data

use a set of four engines well

NASA-LeRC

in the inner

most

be optimum

As will be discussed

Lunar

requirements

mission factor

with

for the LEO-to-GEO

bodies

one

For a man-rated

engines

is better

parametric

selection

Monitor

goals

on the compat-

Health

1988

criti-

has been

oxygen/materials

the technology The

on the

the concepts

focus

Control

an OTV

or other

design

or four

Work concluded

and flexibility

builds

and testing

in accordance

mission

on the moon

per vehicle.

required,

vehicles

return

contract,

chamber,

as a comparison.

lbf to 50K lbf as the actual

of engines

hydrogen

in operating

technology

summarizes

in emphasis

planned

An important

fabrication,

has been

2.1-1

as the Lunar change

is the current

the thrust

work

technology

of the engine

7.5K

both

improvement

and an Integrated

Table

of landing

study

by heating

The Aerojet

tested), design,

goals.

which

analysis,

and development

technology

a vehicle

(cont)

cycle

consequent

engine.

(successfully

established

some

through

preliminary

Design

undergo

with

of the proposed

turbopump

ibility,

expander

II program,

by evaluating

cal to the success

Technology,

pressure.

The Phase Phase

Engine

in November

1988

and

for

2.1, Background,

(cont)

2.1.2

Aerojet

Dual

Propellant

In a conventional passages

in the combustion

ficient

thermal

energy

oxygen

flow

simple,

plumbing

is burned

with

cycles.

open

drive

fluid

oxygen

turbopump.

chamber

liner.

without

It is then routed forward,

With

cycle

is capable

expander

cycle

engine,

of only

the hydrogen eliminates

the need

requirement. element combustion of 400°F

by flowing

regeneratively

drop

across

high

a wide

used

results

hydrogen

the thermal

RP'r/D04,I

7.55a

associated

propellant

employed

to a for the

operating

the basic

as a

gas in the

be heated

life, long

as all

times

hydrogen

the current

production

by heated

energy

throttling

for the

release range.

oxygen.

The flow engine

study.

is the effluent

to the oxygen

from

gas-gas

then

injector

The

to a maximum

through

is shown

the

in Figure

hydrogen

the hydrogen

at an efficiency

It also purge

is heated

and

schematic

on

and excellent

The oxygen

heat exchanger

limitations

the demands

helium

I-triplet

efficiency

these

reduces

seal and the associated

extension.

energy

alleviates

This

is driven

is also needed

exchanger

cycle

2.1-1.

used TPA

cost of some

to heat turbine

pressure

exchanger.

of the 7.5K routed

alloys

and is fairly

and,

a purge

must

over

as hydrogen.

for the advanced

This cycle ering

one

and

requirement,

expander

turbopump

a LOX/GH2

nozzle

in the heat

the heat

cycle

improvement

as well

(-100%)

through

This is the schematic

and provides

fluid

oxygen

over

cooled

the cold oxygen

seals

based

by high

cycle

the losses

on only

suf-

the hydrogen This

not have

copper

acquires

potential,

the hydrogen

throttling

for an interpropellant

provides

stability

imposed

propellant

as the oxygen

The gasified

which

for the

a modest

dual

as a working

circuit

performance

and

through

the RL-10.

The Aerojet oxygen

limit

the wall

for combustion.

it does

power,

is routed

for both

interpropellant

and 10:1 or greater

expander

cools

to dependence

the needed

limits

hydrogen

of pumps

good

chamber,

to the design

maintenance,

by using

it offers

in turn, requires

the added

it both

drives

are related

To obtain near

where

Engine

engine,

to the injector

in the combustion

which,

very

wall

Cycle

cycle

the turbine

Its limitations

turbine

temperature

chamber

is straight

propellant

expander

to power

circuits.

Expander

from

has proven

lbf thrust the pump

engine outlet

more

efficient

preliminary

than originally design.

to the regenerator,

6

That

expected design

had

to the regeneratively

considcold cooled

,J

o o

w

_ >. x

W _

o

_

!

e_ L_

M.

Z > LULL (3U-

og

n_l._ clip

_m3::)Q > n D. i-- Z m II

>

7

I'1

II

II

m 0.. i-

II

_

2.1.2,

Aerojet

chamber

Dual

and then

drive.

This

limited

through

series

flow

flexibility

circuit

of wall

point

(1000

temperatures

the LOX/GH2

Shuttle 1990.

functions

Engine

(SSME)

heat exchanger

are very

compact

and thermally

nozzle

the turbine

baffles.

at or below

Baffles

Generally

they

the chamber. importantly, the total

mixed

The baffles hydrogen require,

however,

copper

efficient

structures

circuit.

The HEX

with

the balance

modulates

fabricated

to be operational

be used

provides

to trim

approximately

acquired

to keep

in

The regenerator

It will

in the oxygen

the oxygen

entering

400°F.

on an injector

face are commonly

used

to enhance

cooled

the baffle

surface

enthalpy

change

chamber

the thrust

a significant

fuel passing

for heat input comes

from

hydrogen

prior

percent

combustion

to enhance

the

stability

side

flow

of the baffle

to powering short

hydrogen

of the chamber

compared

barrel

40 to 60% of hydrogen

to be collected turbopump.

to a non-baffled

capability.

section

into

but, more

where

the hydrogen

heating

directly

From

circuits

cooled stability.

baffle

to the hydrogen.

to be relatively

of equivalent

are the hydrogen

through

the baffle

up the opposite

chamber

chamber

with

is still designed

area

side and back

thrust

and

on the Space

heat exchangers.

design

one

lbf thrust

wall

demonstrated

valve.

valve

This

regenerator

and expected

in the baffle

bypass

and baffle

the baffled

2.1-1.

20,000

are the hydrogen

program,

and

of thrust.

recently

circuit

at the

thermal

the regen

cooled

technology

in the oxygen

In this engine

allow

the range

also high.

between

in Figure

part of the engine

it provides

with

as shown range

were

the power

the flow

Chamber

over

of its bypass

The HEX

to split

Both are NASA-Z

the same

are transpiration

down

was

limited

turbine temperature

drops

An integral

hydrogen

is passed

(HEX).

for the hydrogen

extension.

drops

in the engine

using

gain

Pressure

chamber

limits

platelets

65% of the enthalpy cooled

design

by the setting

but was

lbf thrust).

exchanger

output

temperatures,

of a 21:1 throttle

key components

as a pre-heater

the engine

cooled

to 21,000

within

the hydrogen

and baffles.

remedy

heat

Main They

effective

(cont)

powering

and pressure

is capable

lbf thrust

are well

copper

hydrogen

temperature

version

Two

from

high

Engine,

before

of the chamber

A very

flow

Cycle

circuit

and the regeneratively

or parallel

design

the baffle

walls

of the cycle.

injector

Expander

generated

by the copper

The combination

split

Propellant

volume.

They The

and

2.1.2,

Aerojet

Dual

chamber

diameter

(chamber

injector

commonly

have

2.2

Propellant

Expander

is increased area

Cycle

to compensate

divided

contraction

by throat ratios

Engine, giving

area).

(cont) an unusually

Where

storable

of 2 to 4 this engine

has

high

contraction

propellant

engines

a ratio

ratio

of 15.3.

SCOPE

2.2.1

Objective The

descriptions

and

objective

parametric

of the data

study

for use

is to develop

by space

advanced

transfer

engine

vehicle

system

primes

and

NASA

planners. 2.2.2

Requirements

The propellant

engine

Specific

engine

engine

start

Figure

2.2-1

system

cycle

and

and 2.2.3.1

their

autogenous

data

for the

design

Contract

Subtask

Xk117

.SSa

the OTV

goals

are given

pressurization

Engine

oxygen/liquid

engine

hydrogen

technology

in Table

2.2-1.

requirements

Study

and

objective

engines

over

lbf thrust

3-23772.

This

task

•

Needed

program. The

baseline

are given

•

Identification

Parametric

a thrust engine

generates

in

advanced

and

the specific

range

of 7.5K

design

changes

assessment engine

9

cycle.

with

five

subtasks.

2.2-2.

Analysis

at a minimum,

cycle

activity

in Figure

is to develop

OTV

engine

is a 15-month

are presented

2 - Design

subtask

is the 7.5K

liquid

Description

for the

RI'I/I

under

interrelationships

on advanced

NAS

the

and tank

Advanced

The metric

continues

developed

requirements

Program The

subtasks

engine

technology

2.2.3

The

advanced

over

design

and

lbf to 50K lbf.

developed

under

The

NASA

parabaseline LeRC

the following: the

thrust

of advanced

range.

technologies

needed

TABLE Engine

System

2.2-1.

Requirements

and Goals

Propellants:

Liquid Hydrogen Liquid Oxygen

Vacuum Thrust:

7,500 Ibf to 50,000 Ibf (Study Range)

Vacuum Thrust Throttling Ratio:

10:1

Vacuum Specific Impulse:

#

Engine Mixture Ratio:

6.0 (Design Point at Full Thrust) 5.0 - 7.0 (Operating Range at Full Thrust)

Chamber Pressure:

*

Drlve Cycle:

Expander

Dlmenslonal Envelope: Length (Stowed/Extended) Dlameter (Maximum) Mass: Nozzle Type:

Bell With Not More Than One Extendible/ Retractable Section

Nozzle Expansion Ratio:

End of Regen Section to 1200 (Study Range)

Propellant Inlet Temperatures: Hydrogen Oxygen

37.8 R 162.7 R

Inlet Net Posltlve Suction Head: Hydrogen Oxygen

15 ft-lbf/Ibm at Full Thrust 2 ft-lbf/Ibm at Full Thrust

Deslgn Crlterla:

Human Rated Aeroassist Compatible Space Based

Servlce Life Between Overhauls:

500 Starts/20 Hours Operation (Goal)

Service Free Life:

100 Starts/4 Hours Operation (Goal)

Maximum Single Run Duration

##

Maximum Tlme Between Flrlngs:

_#

Mlnlmum Time Between Firings: #*

Maxlmum Storage Time In Space: Glmbal Requirement: Yaw Angle Acceleration (Maximum) Velocity (Maxlmum)

_tt It t#

Start Cycle

(Figure 2.2-1)

* Englne Parametric Study Result **Vehlcle/Mlsslon

Study Result

10

C om

B ,m

_=

>

_= _B x._

8 a w

!

2 w_

Zw

_

_z_ oO

_>Oo

_w

--

o wW_

Z_ O_

_

Oz

>D

<0

0

&

_0

zz_ o<

o_

_ _

°m

@

_w Oo

Z_

_0

Ow

IZ

_ _

DD

11

M=

m

,*,l

,,,-

W

w_

o_

0 9_'J _,Q w,.J O0

"o

uJ IIlI

.i

¢-

UJ

J

cO ml

I

[

T

1

__

.-._ [.-)

L_

0

_1 v

-" _[,_1

_

)-

r,,_

I-

ei

&

OR

m

u,.

E

@

12

2.2.3,

Program

Description, •

(cont)

Obtainable

design

point

chamber

pressure

for each

thrust

studied. •

Appropriate ance

data

mixture

ratios

thrust

performance, thrust.

of 5 and

are generated

turbine

bypass

Estimates

turbopump,

Power

balance

7 as well

of 0.1X nominal

• Plots

•

thermal, for each

versus

control

power

is obtained

as the design

(10 to I throttle for such

and

data

balat

MR of 6, and

at a

point). related

factors

as percent

thrust.

of delivered

specific

impulse,

engine

mass,

and

engine

envelope. •

2.2.3.2

oxygen

rich

thrust

point

A preliminary

definition

including

liquid

oxygen

and

line

thrust

takeout

structure,

sizes,

Subtask

3 - Engine

Defines

the effect

operation

(MR

selected

of the engine-to-vehicle

by the

liquid

Requirement of increased

hydrogen and

contractor

but

4 - Vehicle

Study/Engine

and

approved

inlet

gimbal

Variation throttling

= 12 + 1) on the design

interfaces location

and

system.

Studies range

engine

(up

to 20:1)

performance.

by NASA

LeRC,

is used

and

very

A single for this

subtask.

2.2.3.3

Subtask Discussion

shall

be used

to generate

and

data

supplied

following:

•

Engine

maximum

•

Maximum

time

between

firings.

•

Minimum

time

between

firings.

• Gimbal

single

storage

time

requirements.

13

run

Coordination

by the vehicle

the

• Maximum

KJ,r/l_,Tas,

with

Study

duration.

in space.

prime

contractors

2.2.3,Program Description, (cont) This information There any

was

insufficient

of these

items

2.3

during

propulsion Launch

System

Project

Pathfinder.

under

development

performance

(ALS),

This

in all three

by high

The

engine

expander

cycle

and

arguably

the most

generator gas generator expander

engine

under

the

engine

50%,

Rocketdyne and

the

engine

has

left even

capabilities.

The century

practise.

R_r,'DC_iT_.

The

most is the

technical

development

include

cycles.

have

The

dual

(CTP)

base

a need

chemical

program work.

under

The

to meet

in copper

engines program

alloy

for an integrated

100K

for a LOX/LH2

contracts

chamcontrol

work heat

and

to improve that

recent

improvement

This

recent

on rocket

propulsion

transfer

in expander

out

an

thrust

should

be con-

lbf thrust.

of enhancing Examples

to increase

obsolete

the

to

developed

cycle.

heat

are used

rule

to 500K

of ways

expander

chamber

exchangers

may

& Whitney

a number

the gas

compared

cycle up

is

among

extending

expander

Pratt

of the

even

thrust,

development

demonstrated

capability. texts

engine

cycle

efficiency

support

A modem

the effectiveness

chamber

abound

as engine

however,

of the hydrogen

expander

of increased

Rocketdyne,

have

propellant

such

lbf.

variations

variants

the merit

study,

beyond

platelet

but

considerations,

by Aerojet,

ribbed

(Ref.

for

the Advanced

propellants

under

to increase

fairly

new

(NASP),

the CTP

LOX/LH2

a common

of the current

LeRC

control

use

supports

and

other

study

Aerojet

requirements

developing

Propulsion

study

of the cycles,

cycle

trade

components

NOTE:

TECHNOLOGY

Plane

turbopumps,

cycles

innovations

and

20th

but

Results

NASA

currently

speed

gas generator

cycles,

in any

Also,

the

cycles

for an expander

sidered

the

level.

system.

Expander

cycle.

ENGINE

Transfer

programs

sophisticated

cycles.

thrust

to define

Aerospace

engine

All share

monitoring

areas

the Chemical

advanced

primes

ROCKET

the National

requirements.

health

range

and

for each

of the study.

program

technology:

pressurization

and

course

major

to be supplied

the vehicle

TO CURRENT

are three

system

with

the

RELEVANCE

There

bers,

contact

was

by about

are 40-

the throttle cycle

concerning

range

technology expander

cycle

1, p 156).

important reduction

development

rocket

engine

development

of the theoretical of the RL-10

advantages

engines

14

was

trend

in the

of a LOX/LH2

a major

achievement

last

third engine as was

of the to the

2.3, Relevance

to Current

Space

Shuttle

Main

States

Space

program

single

most

important

version

Engine

of this

evance

to current

1)

2)

4)

(SSME).

space

advanced rocket

Both

engine

of these The

be in use

engine

study

engine

technology:

engine

can be produced. engine

Oxygen

cooling

are now

practical.

Operation

of the next

seconds

or other

of nozzles

and

stoichiometric

several

statements

specific

(7.94:1)

copper

and

help

define

expander

requiring

engine

turbine

drive

its rel-

cycle

for the throttling.

for turbopumps

at mixture

chamber

a

The

can be developed

is practical

if the

that

United

will be the

century.

impulse)

vehicles

in the

decades,

of the 21st

hot oxygen

engine

see service engine

of 20:1 throttling

Vehicle

of a LOX/LH2

half

several

capable

will

engine/CTP

for the first

(>480

Excursion

engines

OTV

support

performance

A LOX/LH2

(cont)

development

A high

than

Technology,

decades.

will likely

Lunar

3)

Engine

for several

of this engine

results

Rocket

ratios

is given

greater

a protective

coating.

5)

Current

copper

an advanced

engine,

continued

6)

Chamber that

materials

but

some

design

and

a chamber/injector

system

given

chamber increased

technology capability

is acceptable

for

can be expected

with

manufacturing

will survive

a modern

techniques

500 starts

integrated

control

and and

give

100 hours health

confidence of

monitoring

(ICHM).

Current

state-of-the-art

control

8)

combustion

development.

thermal

operation

7)

alloy

and

health

operation

risks

Recently

proven

in electronics monitoring

and

extend

system engine

turbopump

static

bearings

and

start

100 hour

engine

and

sub-critical operation

that

controls will

makes

greatly

possible

reduce

a

engine

life.

technologies

such

operating

speed

requirement.

as self-aligning designs

help

hydromeet

a 500

2.3, Relevance

to Current

9)

10)

Platelet

heat

thermal

margins.

The

Chemical

can

now

cessfully

for Lunar

reported

herein

testing. repaid

major

the

in the

of the

Soviets

Japanese

panies

on licensing

their

engine

would

have

flight

qualified

does

Japanese

not continue, lost.

States

rocket

and

liquid NASA

has

form

of computer

toring

capability.

approaching

wr/_,Ts_

the leader

This

industry

area

theoretical

goals

chamber

Pathfinder for the

main

Vehicles.

and

This

discussions

has

great

at a time

of significance and

design limits

will be that

when

is the

as the OTV

16

successful com-

the United

States

presented

the position If the

of potential

by a

that

the

development

wasted

and of the

development

United

is vanishing

budget.

upgrading

sophistication assess

engine

advance

American

inducement

related

general

can better

with

for the survivability

defense

for a limited

a real

demonstrated

technology.

significance

The

to propulsion

this program

substantiates engine

the increased tools

now

to the economic

in LOX/LH2

be

program.

represents

already

program

should

ExcursionVehicle.

have

Without

fabrication

work

development

in this report

held

suc-

results

hardware

applicability

Lunar

study

technology

engine

have

of the engine

The

design,

the Japanese

counter

programs

Vehicle.

analysis,

is its direct

technology.

priorities

models The

operating

development

of this extended

its significance

engine

conflicting

A third

Project

risk

Excursion

presented

have

engine.

however,

opportunity

within

technology

for a low

Vehicle

and

technology

is still

to improve

Excursion

engine

of study,

then,

baseline

The

States

Lunar

Lunar

cost

Transfer

engines.

United

and

in risk for a full scale

program,

The

no high

and

years

moderate

design

state-of-the-art.

LOX/LH2

program

performance

and

on eight

relatively

engine

Propulsion

industry

Vehicle

for the Lunar

The

can be used

LOX/LH2

propulsion

by the reduction

requirements

(cont)

technology

Transfer

directed

are based

significance

Technology,

OF THE PROGRAM

Transfer

The

realistic

for Lunar

positioned

amply

Transfer

NASA-LeRC

needed

and

exchanger

SIGNIFICANCE

The

Engine

define

engine 2.4

Rocket

what

does.

of design of control

is practical This

gives

and

tools

in the

health

moni-

or possible

when

higher

confidence

in

2.4,Significance of designs can

prior

readily

the

Program,

to reducing deal

with

new

while

sensors

health

monitoring/management

operation

over

advances

this engine

to testable

the complexity

engine

I_]'l'/IX_llT_SSa

them

(cont.)

a service

and

and

a variety system

life previously

could

hardware.

The

interactions

will

algorithms greatly

unobtainable

not be developed

| "7

control

of a modern

of software that

new

beyond

increase

in a rocket drawings.

sophistication

expander

cycle

can be integrated the safety engine.

into

of engine

Without

these

a

3.0

DISCUSSION 3.1

DESIGN

AND

PARAMETRIC

This task began engine

preliminary

expander

series

the high

chamber

sure,

mixture

mixture

design flow

version

penalties

in system

the split

or parallel

that

temperatures

was

adopted

flow

proportioner

split

circuit.

to each

the wall

at full thrust

wall

psia

hydrogen

temperature.

When

chamber

pressure, to the barbut had

corrected

initial

penalty,

pres-

temperatures

This was

2.1-1.

dual was

chamber

to the regenerator,

high

no performance

is the

in the evaluation

(200 psia

the pump

in Figure

for that design

(2000

idle condition

produced

shown

directly

circuit

retain

by going

analysis

the parallel

to

proved

flow

version

concern

that a 20:1 throttle chamber

within

current

chamber

pressure thermal

development

requirement

pressure

over design

range had

for each would

This selection

all the thrusts

studied

decision

engine

below High

worked (see Table

lbf thrust

The task assignment

bounded

the engine

used

in the study

_C,a-3.0-3 .! .2

points

circuit.

now

is always

manifolding

was

funccooled

modified

In all other

100 psia.

out very 3.1-1)

made

psia.

pogo

chamber

for a 20,000

The intermediate

chamber

was

as 2000

lead to system

capabilities.

cooled

respects

a design

This was justified type

instabilities

Also,

pressure well and

to baseline

2000

psia

engines

by a

if the low was

well

do have

for a 10:1 throttling accommodated

a 20:1

engine.

|

as

cycle.

an arbitrary

to be extended

design risk.

in the engine

flow

the

as it can be used

The regenerator

The

the injector

of the chamber

function

life.

to the baffles. Also,

addition

to control

and the regeneratively

for chamber

the pump.

to the system

a worthwhile

the baffles

directed

of the study,

thrust

thrust

of both

independent

nominal

to be added

This has proven

stream

their

At the outset

had

implications

from

is now

the components

valve

temperature

hydrogen

RI_/D0417

and

with

A hydrogen

with

50K lbf.

A concern

lbf thrust

for the study.

on just the hydrogen

throttling

drop

lowered

3.1-1.

from

section

schematic

were

work

in Figure

of hydrogen

of the 7.5K

schematic

temperatures

turbine

pressure

tions

greater

wall

This has important

the baffle

2). The cycle

as given

The flow

then to the TPA

chamber.

of the results

ratio of 6) and at the tank head

ties, and

to optimize

a re-evaluation

(Ref.

and baffle

ratio of 5).

hydrogen

with

ANALYSIS

thrust

range

are given

between

in Table

7.5K

3.1-1.

lbf and

The 20,000

+

!+ m o_ _.r-

L_'_ ®2 t-ic,+,-

r,'J

P_

lg

3.1, Design

lbf thrust sultation

and

point and

schematic, Table

Parametric

was

Analysis,

also

direction

selected from

pressure,

could

be generated.

3.1-2)

3.1.1

for the engine

the

chamber

and

the

five

Cycle

program thrusts,

•

ENGINE

7.5K

lbf

STUDY

OTV

20K lbf

Lunar 25K lbf

35K lbf

50K lbf

3.1.1.1

machined line

alloy

lurgy

to final

a 100K

lbf total

thrust

gravity

losses.

is GLIDCOP

techniques.

throughout

the

the machining material.

This metal capability

An alternate

for three

thrust

stage,

and

results

Chamber Chamber

is pure matrix. over material

._

(see

Transfer

40 to 50,000-1b

Vehicle

and

of payload.

for a 4 engine

engine

LTV/LEV

LEO-to-moon

orbit

for

transfer

-- The

milled

copper

pure

This

is also

half

for the LEO-to-Mars

be extrapolated

chamber

coolant

manufactured

The

LTV.

the

transfer

to that

thrust.

Assembly Liner

with

engine

could

liner

channels

by the SCM with

small copper

0.15%

amount and

is the NASA-Z

TCA

RlVl-/DO417 _Sa. 3.0.3.!

results

engine

RATIONALE

Lunar

for two engine

of the baseline

ALl5

a baseline

con-

(LTV)

thrust

dimensions

after

available.

for a 4 engine with

Nominal

Thrust a.

results

thrust

vehicle •

baseline;

Vehicle

Nominal

With

study

SELECTION

Excursion

minimum •

the actual

tasks

3.1-1

thrust

Provides

monitor.

THRUST

Program

Minimum

variation

Definition

TABI.E ADVANCED

requirements

NASA-LeRC

and

Engine

(cont)

20

is a copper on the

enhances which

using

oxide

of aluminum

alloy

backside.

Company

aluminum

the

was

The

selected

basemetal-

dispersed

greatly

cycle

billet

powder

(A1203)

oxide low

alloy

fatigue

improves life of the

for the 3.0K

TABLE EXPECTED

•

Weight,

RESULTS

Envelope

and

intermediate

•

Power

Balance

•

Changes

•

Assessment

of 20:1 Throttling

•

Assessment

of high

thrust

OF THE

Performance

three

points Results

in Cycle

3.1-2

Predictions

plus

a 50K

at Each

Thrust

and/or

at 7.5K

and

on Scale-Up

at a Selected

ratio

for Engines

maximum.

Components

mixture

STUDY

Thrust

performance

Level

(MR

of 12) at a selected

level

•

Identification

•

Preliminary

•

Innovative

•

Interchange

with

•

Preliminary

DDT&E

Vehicle

of Critical

Propulsion

•

Final

Technologies

Engine/Vehicle

Design

(LTV)

•

report

Interface

Solutions vehicle

and

analysis

suitable

costs Lunar

Requirements

or Technologies

primes

at NASA-MSFC

for a common Excursion

for various

for use

engine

Vehicle

mission

by vehicle

requirements

R_"_/l_._7-sso-r

PARAMETRIC

21

and

primes

Conferences

for the

Lunar

Transfer

(LEV)

operational

in assessing

scenarios

propulsion

at

3.1, Design

design.

and Parametric

GLIDCOP

environment

than

for expected

hardened

NASA-Z

mission

tional the

through

channel

design

chamber

using

use

would

converging,

The

small

10 mil lands) quately

cool

in both

width

channel

and

depth

under

A nickel cooled

NASA

cobalt

Liner

(NiCo) for the

Alternatives

to the

Chromium.

Strength

the electroforming be an important

NiCo

process

LeRC

blockage.

widening

temperature to higher

funded

induced

(11 rail channels,

channels

Channel

lowest

in

variation

design

can ade-

geometry

pressure cooled

closeout

that

be about

give

given

and

but

varies

drop. throat

The

basic

program.

higher

a barrel

is 15.3. ratio.

The

hydrogen

A platelet

the microchannels

22

has

3 times cycle

could

the

strength

life. and Nickel-

be an improvement

in

of finished

liners.

This

machined

copper

liners.

of the

diameter

will be designed

simultaneously.

are Nickel-Manganese

E

devel-

on the hydrogen

improved

yields

value

technology

metals

alloy

there

and Manifolds

to protect manifold

bimetal

be used

the high

two

demonstrated

closeout

the same,

would

ratio

This

could

utilizes

with

was

program.

to a 28:1 area

The

liner

closeout

of 2.5 inches,

manifold

The

a thinner

that

contraction

of the

--

for electroforming

Dimensions

diameter The

Closeout

closeout

would

integral

17.._a-3.0-3,1.2

with

are

A conven-

channel

as adjacent

cooling

in the NASA

allowing

equivalent

RPT/DOt

life due

effective

3.0K lbf TCA

at a position

a flow

chamber

channel.

contract

consideration

a throat

cause

The

coolant

electroformed

c.

part

with

sections.

a blocked

performance

channels.

throat

reliability

are

3.1-4.

thermal

in the micro-channel

improves

for most

LeRC

electroform,

of 12 inches.

and

and

radiation

8)

throat

has

shortens

demonstrated

of a nickel

ber

barrel

changes

3.1-3

for the coolant in the

to the high

volume

life and chamber

design

variation

and

See Tables

30 mil channels

around

b. oped

doses.

and

less sensitive

strength

fatigue

appreciably

stresses

was

(See Reference

use

diverging,

the material

design

cycle

temperature

reduces

as yield

of a microchannel

the 30 mil channels

stresses.

material

radiation

The low improved

(cont)

is a dispersion

of space

negligible

Analysis,

The

20,000

lbf thrust

of 10 inches, inlet filter

and

manifold

TCA

for even

any

will

flow

debris

chain-

a length

(L')

be attached

will be designed from

could

as an that

distribution,

could and

Table CTP

•

THRUST

CHAMBER

Engine

ASSEMBLY

Swelling

3.1-3

Materials

Selection

IN A HIGH RADIATION

of Neutron

Irradiated Volume

Material

Copper

% Increase

ENVIRONMENT

Alloys After Irradiation

3 dpa 1.

15 dpa 2.

Copper: Marz**grade (99.999%) OF grade (99.95%)

1.8 2.1

6.8 6.6

DS Copper: C15720 C15760

0.8 1.1

0.9 0.6

Precipitation Hardened: Cu-Zr Cu-Mg-Zr-Cr

nil nil

3.6 nil

1.

3 dpa corresponds

to fluence

of 0.4 x 1026 n/m 2 (En > 0.1 MeV)

2.

15 dpa corresponds

to fluence

of 0.4 x 1026 n/m 2 (En > 0.1 MeV)

dpa = displacements Trade Name SCM Metal Products Cleveland, Ohio

17.44-7a/rV1

per atom

23

.c_ (30,--0

0_ _

U') ,q"

t_

0

(.O _'O O

o00,¢ (,O I._ u3

er

uJ o ',d"

"¢

"¢

.,,r..,r.

_OOLO 0_000

O

_10 T--

('_J

_ '_" O') C0 '_1" (O

.v',r--

oc01_

o3

_

¢..O (.O ',_

_,-o4 o000_

'_:t ',¢ o3

C

_..O .T- O')

I_0O

•,- O'_ '_f" o3 I'_ co

(3")o1",,. OOO0O3

f_ r_ O

CO _,-

8

CO ',,-

onOlo 000,-

"Oi/3 cO.,-

go.ocO

m -r i

m

c"

.o °.

c..

.o °°

OLO o

0

O o go a

2..4

I.O

_m

go 13_

N

6_ :b

rO

I--g0 rO

3.1, Design each

and

channel

coolant

Parametric

will be flow

hydrogen

hydrogen

collection

for the

two

for hydrogen

a structure

igniters

release

combustion

this in Task energy

release

have

accomplished chamber

a chamber

poses

engine

range

face

injector.

ciency.

This

The

the detail

design,

in the

series

metal

all approach

also

as well

flow

dome

pitch

as

circuits

of the injector

and

yaw

gimbal

(as

7.5K

exchanger

(HEX)

or potential

Bleed

- The

in such

has

performance walls.

and

The

injector

and

oxygen

change

phase

all oxygen

gas-gas

entering

elements

peris

the

is excellent

for a throttling

ratio

a

addressed

compatibility

is needed

injector

chamber

made

and

energy

release

Aerojet

element

The

100%

energy

or baffle

mixture

and

lbf thrust

the

with

This

injector

technology and

mixing.

face,

baffle

high

chamber

stability

nearly

problem.

for gas-gas

the

percentage

and

The

oxy-

range

fluctuations

of

as the oxygen

(See Ref. 3). Face

should

very

to chamber

for cooling

increases

but

inches.

of hydrogen not

exceed

plates

6%.

instead

for rocket

engine

it consistent

with

2

wall,

for face

A means

design.

etched and

cost

bleed

of nickel

flow

the baffle to energy

passages plates

release

face bleed and

copper

Recent

extensive

development

thrust

chambers

has

in

is to conas was of plat-

reduced

performance

near effi-

will be determined

of reducing

an ultra-high

_

precise

life at some

used

of platinum

engine

has

the chamber

Ref. 4)

RI'T/DI_417.SSa-3.0-3.12

are

performing

It delivers

pressures.

bands

region.

highest

for maximum

heat

of chamber

bleed

precise

the injector

baselined

(See

to the

inlet

3.1-1A)

The

for the

by tailoring

Combustion

thrust

- The

in relation

is designed

two-phase

for hydrogen

such

for routing

(Figure

source.

compatibility

contract

position

(2)

inum

The

it incorporates

There

actuators

of eight

of the element

excluded

the

to the baffles

laser ignition

I-triplet.

length

gas phase.

wide

20:1 without transits

is the

in the LOX/GH2

a very

in that

for combustion.

Design

length

element

is in the

struct

top of the chamber

is unusual

flow

their

in a chamber

versions

This

injector

the gimbal

element

for element

five

and

Element

C.4 of the OTV

formance.

the

uniformity.

3.1-3).

injector

efficiency

The

oxygen

for attaching

in Figure

gen/hydrogen

--

and

(1)

over

for acceptable

at the

for the hydrogen

GOX/GH2

illustrated

will

compared

in a manifold

Injector

manifolds

inlets

has

short

and

turbopump.

the usual

also

(cont)

checked

is collected

d. and

Analysis,

the

engine

risk

of

design.

_-I

j,

Figure 3.1-1A

_'::'°' /

log 189.035

Our Regeneratively Cooled Baffles Provide a Dual Function as an Added Heat Transfer Surface as Well as a Combustion Stability

Damping

Device

25A

3.1, Design

and

Parametric

Analysis, (3)

the injector

face

constructed inch

using

thick

body centered

the

the

structural oxygen

with

to accommodate

the oxygen

the

hydrogen

to accommodate

flow

the

two

(4) dent

of injector

gen.

A secondary

small

flow

circuits

compartments gas.

From

the baffles.

The

hydrogen

manifolds

the regen

that

cooled

injector

and

the

around

the tip,

regen and

cooled up

semi-circular

manifolds

basic

injector

design

flow

circuits.

drogen allow

the igniter

that has

are

three

The

ports

They

are built

collectors

a section

may

to be brazed

tors

have are

constructed

proportion

baffle

up

as there of nickel

The

The

and

and

segmented They

space

circuits

the heat

area

are

also

the

injector

hydrogen

of the

into

are

semi-cirfrom

between

the

in two

manifold.

to form

divided

to

the baffles,

inlet

are stacked

to

is due

terminating

baffle

into

flowing

connection down

hydro-

proximity

hydrogen

system

manifolds

to the

by two

mechanical flows

are indepen-

input

to the

to collect

which

outlet

flow

compartments

top side

manifolds

The

are

compart-

hydrogen

of the baffles.

collector

The

the hy-

in two

sections

below

the

to function.

baffle

plates

sections. surface

baffles

extend

They

is no convenient 200.

and

is a substantial

are partially

baffle

hydrogen

to the

sealing

This

transferred

to a separate

as discrete

at a time.

heat

line forms

actual

1/4

injector

of this

in close

the

the necessary

readily

of the injector

of wall

weld

inlet

dome

divide

used

separate

are

manifold

The baffles

to the baffle

welded

The tor face.

stability.

to the manifold

side

approximately

to the base

out

is to increase

chamber.

the back

and

function

is distributed

This

The

will be no internal

- The

40 to 60% of the total

chamber.

areas

hydrogen

compartments into

breaks

to the inconel

manifolds.

there

design

injector

The

welded

Construction

a high

are welded

the

baffle

ports.

is combustion

to maintain

combustion

cular

ignitor

as their

function

are

hydrogen

hub

brazed

element.

oxygen

passages

Baffle

small

pierce

the

to assure

and

These

and/or

posts

and

joints

and

technology.

posts

Oxygen

all welded

Both

welded

to complete

hydrogen

spoke

platelet

Oxygen

opening

the

sections.

beam

circuit.

divides

- The

developed

compartment.

mixing.

tailored

well

element

also part

independent

hydrogen

oxygen

which

several

Concept

will be electron

in each

covers

Injector

Aerojet's

sections

to form

ment

into

(cont)

welded

between means

are

are

above

and

to the inlet

the inlet of welding

NASA-Z

copper

and

outlet

the joint.

and

injecoutlet

collector The

in the baseline

collec-

to

3.1, Design and Parametric Analysis, (cont) design, but assembly

may

be constructed

techniques

the

nickel/platinum

are

formed

are

diffusion

correct

and

thermal

system

from

platelets

bonded

after

a unit.

The

bonding.

pierced ment

body

used

by the

oxygen

is formed

when

depending

the

baffle

system

is possible.

for the

end

on the

copper/nickel

so welding passages

rounded

See Figure

ignitor

ports.

A similar

the

engine

feeds

the oxygen

oxygen and

dome

yaw

is finish

3.1-2 for a cutaway

The

and

baffles

hydrogen

flow

machined

to the

of the 7.5K

be welded

good

for both

weld

in thermal

and

expected

has

cooled

and

injector joint

cycling.

baffle

circuit

fitted

The

body with

which

lbf TCA

the nickel

be kept

injector

from

into

the

same

bonded

will operate and

the

segments

that

a

can be used Differences

material

at fairly

up in the LOX/GH2

forms

alloy.

hydrogen

to

outlet

that

joint

of the

openings

to the baffle

will be of the

of

+6 ° pitch

injector

and

within

center

alloy

of

side

the conforming

or monel

face

top

oxy-

is welded

for the

plates,

hydrogen given

point

ele-

to one

dome

to be welded

assembly

heat

adjacent

At the

is

respective

on the opposite

requires

injector

Manifolds

is fed by liquid

considerably

assembly

will be an inconel

must

The

weld.

portion

complete

the

compartment

is slipped

dome

The

a passage port

to the

of this compartment

with

as an attachment

oxygen

service.

place

ignitor

oxygen

assembly

the

into

fusion

used

dome

from

to the

Final

of expansion

thermal

as the

adjacent

manifolds.

hydrogen

coefficients

are

or brazed

in the elements.

compartment

extension

plates

and/or

oxygen

injector

and

The

segments

injector

The

be centered

a circumferential

This

structure.

braze

atures

system.

to the baffle

manifold

must

compartment.

with

are welded

compartment.

passage

is a cylindrical

assembly

segments

is fed to the

dome

gimbal

in the baffle

that

Hydrogen

the hydrogen

injector

the injector

the

onto

inside.

posts

post

limits

low

stream heat

for

temperto the

exchanger

regenerator. e. of the chamber

complished connection the chamber side

etched

as a hydrogen

gen

area

Both

liquids

chemically

alloy

plate.

of the injector

and

required.

a common

The

the

or platinum

margins

have with

into

dimensions

baffle

of platinum

loads

RPT/D0417.55a-3.0-3.1.2

Stress

must

by using to the inlet

Figure

and

be protected

a "can" manifold

in the longitudinal (See

Relief

structure

Component from

and extends

is a sliding direction

damaging

Mounting stress

but acts as a rigid

27

or side

to the hydrogen

joint that allows

3.1-3).

Structure loads. inlet

for thermal stress

--

takeout

The throat This is ac-

manifold. expansion structure

The of for

Oxidizer Dome

I ,i

I

I

Oxygen Inlet

I

Injector Fuel Manifold

Baffle

Additional Injector Elements

Figure 3.1-2. 7.5K Ibf OTV Injector/Baffle Flow Circuit

28

- -

Vehicle Interface (Station 100)

_

Attachment Structure for + 6 ° Gimbal Actuators

Thrust

Takeout--re,-

Assembly

(1 of 2)

Component Mounting Can Assembly Rotated 90 ° from True Position Gimbal Ring

Gimbal

Attachment

to Hydrogen Manifold

Oxygen Inlet Manifold

en Cooled

Figure

3.1-3.

OTV

Engine

TCA

Sketch

Showing

29

Can and Gimbal

Inlet

Nozzle

Attachments

3.1, Design

and

Parametric

Analysis, This

ious

engine

Bracketry bilize

components. extending

the many

takeout

to the

gimbal

structure

is located

are

are

located

connected

actuator pitch

and

pivots

yaw

plished

where

actuator

rods

cooled

welded tubes

the

throat

throat

gimbal the

the

inertia

Oxygen

thrust

system

Cooled

baseline

concept

then

contoured

nozzle

where

ratio they

are

They

are

platelet

ical

structure

bility with

structure that

is a structure flow

passages

design

trade

three

concepts

RI "I'/DO4 |7 -r"5 a-3"0"3 "I"2

study that

that would

box The

actuators,

90 ° apart

and

of the

combination takeout

mounting

this engine

alloy

inlet and

of +6 °

is accomstructure.

The

mass.

ring

two

are

back

formed

by material to select some

that the

could

best

study.

30

for the oxy-

swaged

tube

that

bundle

is at area

ratio

bifurcated. for half

The

35.

is

All

bifurcations

the length

of the

manifold.

and brazed

contour

sheets

completed

then

in sections

copper

been

manifold

doubled

to the final

from

received

thrust

has

to the swaged

formed

have

The

be formed

be formed

is needed

design

at the outlet

would

circular

Movement

in any

to the engine

also serve

The

placed

dome.

Actual

the

gimbal.

are

point

must

construction.

throat They

gimbal

is a copper

then

An alternative per

robust

to moving

for forming

terminated

be of fairly

oxidizer

detailed

on a mandrel.

of 600:1.

and to sta-

Nozzle

The

limit

system

connecting

vector.

related

nozzle.

to the practical

on it.

valves

the flanges

attaches

loads

gimbal

the engine.

about

a centered

for var-

will be mounted

for a true

the engine

m No

to area

plane

above

Concept

extend

that

injector

from

structure

the various

The throat

manifold

a.

extend

to mount

top of the

see only

and

turbopumps

cylindrical

throat

3.1.1.2

gen

near

inlet

mounting

bundles.

This requires

hydrogen

change the

wire

on the structure

to the

rods

the two

Structure--

structure.

structure

as a convenient

the can will be used

Gimbal

gimbal

however,

serves

and electrical f.

as a thrust

"can"

In particular,

from

lines

(cont)

that

that

are

tube then

bundle welded

is a copinto

on a mandrel.

A third

are explosively

welded

be removed

configuration.

by melting Figure

3.1-4

a con-

possitogether

or solution. shows

the

A

a m i-

Q. 4) 0 C 0

C)

I

C 0

m

M_

xr-_

C X LLI

§

N N 0

z "0 0 0

|

m

i Q° a

I

,f.e4

A

IJ.

:o.!! J! .

,,i _n,

31

c

3.1, Design

and

Parametric

3.1.1.3

Analysis,

Radiation

Cooled

Material materials joint lighter

but

requires

does

not

a silicide

require

nozzle

a thicker

assembly would

have

the

have

be contoured

would

The by 28 volt cable and

DC electric

drive gives

shaft a doubly

ate all three millions

inside

a circular

jackscrews.

of garage

are attached

door

The ments. many ratio

The times

seal

sealing

problem.

movements

The

without 3.1.1.4

an over

potential.

The

takes

the oxygen

RPT/DO417.55a-3.0-3.1.2

The

first

at vapor boost

for the

thickness

assures

would oxidation. nozzle

The would

retraction/extension

to reduce

is to have

emphasis

three

weight.

of the

Both

as any

operation

motor

can oper-

is demonstrated

See Figure

seal

can readily

on leakage

one

engine.

leak about

a rugged

by a steel

motor

The

the

driven

connected

design.

and

is only

jackscrews

synchronized

a mechanism

or an increased

intent

uses

mechanism

that

motors

are

rate.

and With

0.3 psia. seal

and

design

break the

This

capable

cable

way

3.1-5.

critical

make

by the

ele-

the

seal

seal

at an area

simplifies

of many

the

nozzle

rates.

Pump

tank

determined head

rise,

pressure

and

boosts

turbine

pressures

that

is a low

pump

and

mechanically

are of a similar

the seal

propellant

mance

to the

across

The

has

shaft

of such

wear

Boost

Aerojet

This

leaf seal

Oxygen

low

the pads

are

to the nozzle

finger

design

problem.

motors

at the top

attachment

design

turbopump.

structure

pressure

columbium

carbon-carbon

carbon-carbon

mechanism

three

that

significant

of 600, the gas

of varying

reliability

is a double

without

sections

The will be

The

erosion

The

where

tube.

openers

to the gimbal

both

rings.

exit and

capability

The

columbium.

carbon-carbon

The

candidate

performance.

The

redundant

The

of the columbium.

to control

retraction/extension motors.

and

The

oxidation.

stiffening

have

configuration.

composite,

reliability

layer

two

for optimum

design

material.

to prevent

at the nozzle

Both

either

established

impregnated

section

attach.

with

coating

would

the

carbon-carbon

compatible

carbide

governs

wound

or other

a silicon

columbium have

are

Nozzle

selection

are a 2 directional

temperatures

(cont)

two

pose

separate

low speed

four

a very units

difficult give

stage

it to 55 psia

for feed

by a portion

of the

is driven

32

the

boost

best pump

to the first

turbopump

high

stage

perforthat speed TPA

pump

Throat Glmbal Clearance Zone

Throat

Glmbal

Brackets

Ox Tank Pressurization Valve

Nozzle Retractlon Motor (3)

t

Engine-out Glmbal

Centerllne

28 Volt DC Motor

_

/

)t

\\

/

/

j\\ Flexible Raceway

Fuel Tank Pressurization Valve

Figure

Pitch & Yaw Glmbal Attachment

3.1-5.

7.5K

ibf Thrust

33

OTV

Engine

(Top

View)

3.1, Design output.

and

Parametric

Turbine

ation

allows

pump

outlet

Analysis, flow

a conventional

is installed

line sections

is combined

with

ball bearing

design

just below

needed

the dual,

for gimbal

3.1.1.5

Oxygen

developed

program.

(See

admittance testing

Ref. 5 and 6).

turbine

limited

hydrostatic system.

to ambient

bearing

400 and

K-500

selected

for best

further

protected

diamond

film

diamond, pure

coating

With

of teflon,

in the

tank

3.1.1.6

speed

Hydrogen

first

stage

oxygen

by flexible

below line

boost

Component TPAs, removal

RI'I'/I

the

sections

pumps changeout

however,

are

for access.

)_ ¼ t 7 .c,.C,a- 3.I)-3.1.2

are

gaseous

OTV engine stages,

oxygen assembly

and nickel

on the non-moving elements.

(current

operates

with

areas,

a

and

were

rub or friction

areas

surfaces

conductivity

stress materials

and a newly

This diamond

a thermal

a full

seal or a purge

in low These

and (GOX)

for an interpropellant

but the potential

of materials

Boost speed

pump.

of the hydrogen

the line just

speed

boost

pump

oxy-

are

developed

film has the hardness several

times

and coatings

the GOX

unassisted

(i.e., rubbing)

greater

driven

of than

LOX TPA

bearing

is

starts

start operation.

The low low

two

is demanded.

life and all the required head

turbopump.

funded

rotating

are copper

oxygen,

and

this selection

of full service

required

plating

The

strength

on the moving

the slickness

copper.

capable

by silver

(maximum)

is no need

material with

Flexible

of the 3.0K lbf thrust

LeRC

section,

GOX).

There

compatibility

oper-

This boost

flowmeters.

speed

lineage

in the NASA

by 400°F

for the TPA

where

speed

elements.

generation

it to the high

is of the design

temperature

selected

vortex

It has an inducer

assembly.

The low

Assembly

tested

to be driven

Materials

monel

TPA

output.

for the rotating

connect

Turbopump

and partially

the pump

in-series

motion

The oxygen gen TPA

(cont)

hydrogen

boost

It is a four

TPA.

Rated

hydrogen that

Pump

pressure

flowmeters

and

accommodate

in an area

that

will be possible packaged (See Figure

stage

pump pump

is separated

from

suitable area

3.1-6).

34

that

motion

disconnects.

would

likely

to the

from

the

It is located

the hydrogen

requirements.

by an astronaut line

in design

by hydrogen is 55 psia.

can be reached

in a restricted

driven

similar

at full thrust

the gimbal

with

is very

Both

in a space The require

high engine

in

TPA low suit. speed

1 E 0 I-|

c-

a 0 N N

.J r-

.__ (n c_

¢-

E

\ Q. ec-

O

0

_> _3

> |

e4

U) U)

Q_ t)

I,i.

rn X

0

35

3.1, Design

and Parametric 3.1.1.7

Analysis,

Hydrogen

Turbopump

Aerojet that

needed

which and

for the

is a low two

thrust

turbines

housing.

design

between

190,000 The

materials

very

selection

so that

high

flow

expansion

operation

around

Hydrogen

drogen

gas exiting

the

through

the baffle

flow

for driving

the

exchanger

and

overall

injector

nology

originally

will

rotate

at

life

over

the wide

require temperature

assembly

or produce

also

susceptibility

lack

the

at rated

operatir/g system

must

temperature.

The

maintained

efficient

heat

copper

A proposed

Rl'IV_lT_5,-a.0-3.1.2

light

into

transfer the platelets weight

two

waste

injector.

shaped

for hydrogen passages

to

inlet

streams.

The baseline

either

alternative

36

a silver is to use

result

of the regen-

platelet

line

of platelet High

materials

or nickel beryllium

tech-

in various

outlet

3.1-7).

heat

is a lowered

Aerojet

and

hy-

energy

like a dogbone

characteristic (See Figure

usable

effect

application much

the

LOX/GH2

One

of the

from directed

into the

The

structure.

with

stream

from

wide

heat

heat

the injector.

finding

of metal

flow

was

is an example

but

used

proximity

by counter-flowing

are zirconium

brazing.

close

being

to the hydrogen

the engine

going

block

to transfer

is downstream

regenerator

it is a short

used

what

from

for injectors

at the ends fine,

converts

regenerator

stream

developed

exchanger

of the TPA

This

the hydrogen

structures

an exceptionally

circuit.

upstream

rounded

very

section

immediately

In appearance,

The

turbine

is a heat

The

devices.

the

Engine,

speed

bearing

the

does

Regenerator

turbopump.

is to cool

erator

Materials

which

shaft

maximum

to meet

bind

in the same

CTP

Each

contraction

do not either

the bearing.

A regenerator

tions.

procedure.

a hydrostatic and

For the

stages

embrittlement.

3.1.1.8

erator

with

six pump

assembly

range.

to

engine

contained

rotating

are hydrostatic

clearances

features

but

in concept

XLR-134

shafts

requirements,

All bearings

tight

balance

similar

Force

design

operating

and

for the output

for cryogenic

unacceptably

assembly

rpm.

engine

for a stout

the normal

very

on the Air

contra-rotating

over

is needed

TPA

separate

speed

for an easier

needed

That

critical

requirements.

hydrogen

engine.

provides

is about

range

LOX/LH2

TPA is used

system

speed

careful

This

shaft

provides

thrust

engine.

a hydrogen

double

the

whatever

Assembly

has developed

divided

This

not exceed

CTP

(cont)

flow with

connecdesign

give

AT's are for the

bonding platelets.

aid

regenused

in

C

C m ,m

m

a m

fJ

0 e= _=

fJ fJ fJ

II.

0

C fJ |

0

*0

fJ fJ

e= om

l, IJ "10

fi

(.1 rm > "10

i

< r,: I

I1 0

0

0

0

37

0

0

3.1, Design

and Parametric

3.1.1.9

Analysis,

Liquid The

engine. gas

The

for the

enthalpy gen

engine

change

out

turbine

the poor

appearance

to minimize

passages

will side

prevent

unmixed

oxygen

changes

At rated

place

exiting

stream

to have

predictable

for propellant

See Figure

2.1-1

functions

the cold

cooled

chamber

can be commanded existing DC motor reliability.

design driven The

RPr/l_iT._s_-3 _3.la

is only

HEX

The

platelets

with

the

in more

The

flow

oxygen

used

film

flow

on the

pressure

and

the phase

must

as the

operation. change change

as the

be homogeneous

cooling

task.

monel

and/or

with

to boiling

at low

oxygen

oxy-

transfer

site of the phase

for a critical

or NASA-Z

from

is similar

by rapid

a problem

nozzle.

The

devices/geometries

is caused

is the preferred

cooled

heat

design

is supercritical

characteristics

copper

passage

of the

nozzle.

of oxygen.

generating

flow

the HEX

The

transfer

and

HEX

drive

The

materials

inconel

inlet/outlet. Engine

Valves

and

The

flight

engine

for the

valve

position

are described

a. divide

phase

to a gas

the oxygen

zirconium

turbulence

Two

traversing

enters heat

3.1.1.10

their

a liquid

uneventfully.

next

are

include

flow.

the oxygen

take

tubing

will

flow

expander

coming

It is larger,

flowrate

rectangular

cooled

hydrogen

The

regenerator.

the high

thirds

AT for efficient

of oxygen.

with

the straight

they

from

drop

high

dual

for turbine

to two

with

(HEX)

of the

in the oxygen

a very

to the

Exchanger

of oxygen

one-half

flowed

characteristics

two phase

should

for the HEX

from

in that

thrust

gives

construction

heating

gained

is counter

Heat

in the operation

provides

balance

This

pressure

depart

HEX

pump

transfer

general

passages

the

outlet.

heat

and

hydrogen

with

element

on the efficient The

pressure

Hydrogen

is a critical

depends

needed

TPA

despite

HEX

turbopump.

of the high

hydrogen

Oxygen/Gaseous

cycle

oxygen

(cont)

and

stream

the engine to adjust

for this valve valve dominant

Engine

requires in the

Control

a set of 12 valves schematic.

The

for normal major

operation.

control

valves

valve

is used

and

below:

Hydrogen

hydrogen

Basic

with

Flow from

Proportioner the pump

baffle

plates.

the flow

+25%

and

several

separate

failure

mode

Valve into

Its neutral of total

coils

with

is to fail safe

38

This

streams position

flow

mechanisms

drive

two

--

are

to either possible.

independent to a centered

feeding

the regen

is at a 50-50 circuit.

split.

There

It will power flow

to

split.

It

is no

be a 28 volt sources The

for valve

3.1, Design

and Parametric

is commanded controller TPA

has

which

to the the

to a particular

turbine.

tion

the flow Any

the bulk

bypass.

At full

energy

thrust

flow

to the hydrogen

When

the valve

flow

always

and

through

flow

open

through

it is pressure

this valve

by changing

the

ure

of this

is fail-in-place

valve

drop

and

closed

becomes

effective. With

only

possible

or hydrogen has

more

open a tank

RPT/DOll

the

main

mixture

components

head

7.55a- 3._3.1.2

amount

pressure start

This

posi-

or directly

will gives

of

loop.

At

the chamber

valve

with

head

to

will go to full

be directed

to the baffle

the greatest

thermal

ratio The

control oxygen

adding drop

condition,

Valve

pressure

path the

ratio

for some relatively

valve

valve

until

drops.

39

open

idle

and

drop

idle With

is capable

the valve

supplies. mix-

a low

pres-

full open bypass

oxygen

valve flow

operation,

in either

but

hydrogen

line

turbine

or closed

pressure

of the hydrogen. small

between

fail-

valve

to provide

modulates

the pressure

The

power

of

The

The

to provide

hydrogen

some

missettings

open

total

as a

is

the turbine.

is designed

hydrogen

all

to be orificed

operation minor

the

for simple

tank

so that have

independent

it is fully

from

lowest

may

around

direct

operation.

as an alternative.

with

off the

can

25% is reserved

regenerator

accommodates

This

valve

abnormal

pressure

n

is to change the

the

bypassed

designed

has

The

chamber

On start

light

This

line

valve

is computed

valves

with

DC motors

the mixture

--

or other

fail-full-open

start. On

Valve

regenerator

for low

Idle

ratio

shutoff

The

bypass

28 volt

in setting

Bypass

of hydrogen

injector.

Mixture

circuit.

a low

the turbine

tank

to the

to assist

regenerator

through

flow

valve

feedback

bypass

balanced

needed

Hydrogen

during

passage

rates.

flow

by redundant c.

sure

flow

overthrust

the regenerator.

above,

control

more

to the

by an adjustment

the regenerator.

during

As noted

ratio

the

independent

hydrogen

through

can be reduced

attained.

ture

its own

Regenerator

baffle

the minimum

be powered

going bypass

is compensated

the regenerator.

that

will

of the

temperature

through

the regenerator

portion

to assure

mode

has

the engine

the regenerator

going

will direct

Hydrogen

is fully

goes

with

valves

which

while

stream

that

TPA.

25% of the hydrogen margin

rise

a larger

maximum

control

valve

two

the proportioner

b. but

in conjunction

of these

temperature

a flow-split

hydrogen

for the hydrogen

bypass

operation

control

split

missetting

turbine

with

the

be done

baffles.

circuit

optimize

This sets

thrust

representing

will

must

(cont)

position

computed

hydrogen

low

Analysis,

the

the oxygen

hydrogen

circuit

is positioned

the circuits

balanced

of the control

to for

needed

3.1, Design and Parametric Analysis, (cont) using a modulating valve. The valve is designed to that

normal

system

hydrogen

TPA.

hydrogen

turbine

Unlike

A "close"

should

valve

modulating

Its failure

can continue

signal

bypass

the other

closing.

pressurization

can

valves,

position

at line

Electrical

from

make

redundancy

ratio

a tight, is also

so

the

when

for mixture

must

of 200 psia

output

by the controller

responsibility valve

pressures

increased

be given

the idle

is closed.

under

also

assume

close

the control.

leak

free

needed

seal

on

for this

valve.

d. leling

the circuit

hydrogen

through

through

through

the

capable

Hydrogen

hydrogen

the turbine.

turbine

be at the

at any

10% bypass

thrust

or mixture

change

new

position.

movement

to the

the change

is synchronized.

hydrogen

turbine

tuning. erator ature

The

bypass

bypass and

and

pressure

bypass

minimum.

circuits

to derive

thrust

calculated

flowrate useful

The

controller

chamber

calculated at low

from

thrust

the

turbine

it was

judged

completed

RI'T/DO417

55a-3.0-3.1.2

is controlled better prior

to route

to entering

final

Bypass

for any

initial

as reflected

flowmeter

Valve

HEX bypass

output

The valve.

all oxygen

through

the oxygen

cooled

have

temperature This

nozzle.

40

the

hot

the ratio

of the regen-

up

temperto the

10%

and

oxygen

with

compared

with

is particularly

greatest

errors.

entering

is on the hydrogen

to assure

The

so that

is correlated

This

the

hydrogen

a

the

stabilized,

of the oxygen

valve

the HEX

start

for mixture

are also

pressure.

readings

--

has

for the hydrogen

readings

and

will follow

missetting

This

will

When

by the hydrogen

information.

and

margin

will

will be automatic

Flowmeter

speed

valve

it

thrust

bypass

change

readings

ratio

thrust

operation.

valve

bypass

position

flowmeter

mixture

turbine bypass

compensation

pressure.

by the

the

valves

uses

where

oxygen

turbine

compensate

This

as to actual

turbine

the oxygen

turbopump

HEX

the

At nominal

to the commanded

on feedback

hydrogen

will make

and

operation

e.

corresponding based

The

to full open.

valve

The

controller

from

It is a modulating

operation.

turbine.

thrust

flow

engine

proportioner

both

still be some

abnormal

will also

to the

balanced.

closed

is commanded

hydrogen

will

all

control

valve

valve

there

directs

This

After

bypass

be pressure

the valve

authority.

modulate

ratio

full open

is in a line paral-

control

position

then

Closing

valve

for normal

to a particular It will

valve

This

position

direct

ratio.

turbine.

--

from

or other

mixture

Valve

position

for overthrust

valve

the

will

can be invaded the

Bypass

TPA

With

as the circuits

of stabilizing

should

the

Turbine

phase from

side

change the

as is

3.1, Design

TPA

and Parametric

turbine

oxygen

discharge

TPA

power feedback 400°F

from and

commanded closely

the oxygen TPA

the

engine

instance,

the 25% bypass point,

the

bypass hydrogen

position bypass

hold

a stable

temperature

regenerator

bypass

temperature bypass

control.

responding

The to the

propellant responding

the oxygen

time

the hydrogen

cent

ratio.

bypass

The is very

engine

bypass

with

requires

various

found ceramic

RPT/,X_tT_,-3.0-3 1_

compatible materials

valve

--

well

reliability oxygen

It is the

established

turbine valve

valve

position

position

on total

to a setting

valve

valve

operating

cor-

information

moves

corin step

is reached.

At that

commanded

engine

for thrust

control

valve.

for engine

position

bypass

selected

bypass

is used

to establish

was

100%

range.

position

to a specific

control

as per-

for the engine. in a "look-up

characteristics

The table"

and

a

accuracy. bypass

hot oxygen.

are good

turbine

master

valve

turbine

with

the

as the

to

to about

feedback

the designated

valve

the HEX

related

to a designated

will modulate

and

This

is made,

hydrogen

bypass

thrust.

both

of high

Valve

to trim

and

range.

bypass

go to

At that

thrust

operating

Bypass

until

a modulating

The materials

valve

turbine

linear

The

400°F.

For

will

The

hydrogen

is used

changes.

a narrow

the engine

the move

of

circuit

25%

on

will be

valve

of movement

from

based

to the hydrogen

to the

pressure

oxygen

thrust indicator

After

thrust.

within

it to move

bypass

the selected

similar

commands

bypass

for setting

to cover

Turbine

turbine

of commanding

runs)

thrust

approaches

amount

will operate

chamber

turbine

a small

valve

thrust.

and

only

HEX

stability

will stay

will be very

Oxygen

technique

position

balance

desired

to actual

with

mixture

The

controller

flowrate

has

adds

bypass

the

to

maximum

valve

turbine to hold

This

flow

turbine

during

from

needed

to the

the oxygen

will modulate

change

the hydrogen

"hunting"

to the

discharged

corresponding

the oxygen

is no controls

the oxygen

then,

for this valve f.

thrust

and

just

for a not-to-exceed

positioning

to the regenerator valve,

adjusts

temperature

temperature.

oxygen

of the enthalpy

valve

up command

valve

on power

Specifications

there

until

compensation.

(based

inlet

valve

a throttle

turbine

will

bypass

This

so that

on receiving

The

the liquid

thirds

for an initial

thrust.

synchronized

two

turbine

controller

with

about

turbine.

the oxygen

from

(cont)

is counterflowed

to give

the oxygen

Analysis,

valve

Copper,

candidates.

41

must

nickel,

be fabricated several

from

of the monels,

and

3.1, Design

and

on the one brated

path

Each

Ignitor two

quarter

leak

cooling.

flow

lasers

must

Parametric

close

Ignitor

Valves

inch

ignitor

propellant

around

valve

the

will

is routed

to two

psia)

line

pressures.

Engine

Controller

Aerojet

has

(ALS)

engine

Health

monitor

change

portioner

continued to increase setting

to protect lead

making intelligence

hardware

supports

RI'I'/D¢_417.5S

some

a calicavities

coils

for reliability.

laser

ignition

with

The

valves

engine

for

requirements.

more

selected

chamber

hydrogen

extend

thrust

More

to a catastrophic software

for later

the

load

will

require

failure

The

using

circuit.

fact

expert

that

Should

the

engine

a mixture

or a change

pressing

lower

effectiveness. signal

failure.

development

or stored

will

to the chamber

its life.

computed

system

cooling)

is evalu-

and

operation cooling

have

data

streams

controller,

control

will

Sensor

data

reduces monitor

an integrated

components

throttle-down

the health

(and

with

requirements.

by the engine

chamber

flow

prior

to CTP

Selected

during

at the

and

have

in the

This

pro-

is decision-

indications this is a decision-

systems

and

artificial

techniques. 3.1.2

puted

800°F

to direct

for use

for the Advanced

designed

system

action

hydrogen

will

for reliability.

All engine

in a computer.

operation

to a shutdown

system

overall

the regen

approach

valve

making

through

use

controller

was

baselined.

As an example,

of hydrogen

that

to meet

for immediate

decisions.

ignitors

suitable

the ignitor

actuation

modified

Engine

system

algorithms

are used

temperature

would

selected

an engine

valve

through

supplies

are

System

Advanced

(ICHM)

sensors

maintenance

Monitoring CTP

These

can be readily

valves

hydrogen flow

power

developed that

type

electrical

ignitors.

(>300

information

ratio

separate

at high

monitoring

controller

redundant

separate

by health

throat

dual

The

hydrogen

two

designed-in

flowrate

to allow

and

This

ated

have

poppet

lines.

ignitor

System

health

seat

m Simple

in each

3.1.1.12

and

(cont)

g.

3.1.1.11

Launch

Analysis,

at a particular

a-3.0-3.1.2

Power

Balance

A rocket

engine

power

balance

point

within

the engine

is an energy

operating

42

envelope.

and

mass

balance

If a computation

comdoes

3.1, Design

balance

and

Parametric

(i.e., there

point

is a solution

is, by definition,

(generally tions), ance define

engine

operating

for chamber

Figure

3.1-10

the 7.5K

is outside

pressure

lbf thrust

engine

3.1.2.1

gram

Power

achieved

data

input

at a heavy

Balance

to a parallel

flow

dual

expander

components.

For this study

turbopumps

3.1-12

curves

rated

thrust/design

studies study

were

nominal

full thrust

MR = 5 was chamber

RPT/DO417.55a-3.0-3

selected

design

.1.2

was

preliminary

chamber

could

to be sub-

As an example,

expander

Tables

cycle

used

in

equations

power

time

in developing

balance

savings

pro-

was

the code.

to change

from

for a look-up

of the algorithms

is then

This was design

It is now

ill

a series

flow

developed

geometry was

A worst

maximum

for each

Also,

case

so that hot gas side wall

wall

for each

design

point

operating temperature

within

for various at each

of the

true

for each curves

of the

turcalculated

curve,

or points

table.

The

Very

on

thrust

temperatures

preliminary

and length.

selected

expected

43

point

for use in a lookup

the maximum.

table

particularly

and performance

so that hot gas side

pressure

used

had to be defined

The initial

and 3.1-6)

baffle

thrust).

for the thermal

be modeled

of numbers

below

chamber

adjusted

the basic

engine

This

modeling.

3.1-5

modeled

were

(100%

from

OTV

designs

injectors.

to determine

point

calculated

required

must

component

thermally

2000 psia

hours

examples

An equation

thrust

a common

dual

15 minutes.

modeling

some

to an array

completed

about

recent

actual

(See

are converted was

takes

gives

and 3.1-13).

chamber/injector

flowrate

envelopes.

flow

The Aerojet

component

for accurate

was determined

(Figures the

3.1-11

and the thrust

bopump

by Aerojet

cycle.

active

Figure

thrusts

4 hours.

the most

the program.

selected

can be hand

of programming

Each

five

balance

to printout

with

for the series

of bal-

Development

in about

input

its fifth generation

operating

of itera-

An array

is used

propellant

the

design.

3 or 4 iterations from

or total

number

envelope. ratio

then

is no balance

after a defined

mixture

similar

components),

If there

operating

versus

Thrust

envelope

A power through

pressure

to generate

is the operating

envelope.

on a solution

of the engine

envelope.

of the engine

operating

to close

on a plot of chamber

the

stituted

point

the capability

the engine

by a failure

the selected

(cont)

within

within

indicated

points

Analysis,

at

configuration early

thrust

as the rated

of 2000 psia point would

in the

and

(MEOP). not

The

exceed

or

0

z

i

o__

o

8

i

_

.

,

Ox P

,,=,

P,ressure > 540b psia ILl

I

el

I

I_.

:'_ I

:

_k.

:

___

_

_'--'-

0J

I_

o,.

............ =. / / /

_o i\_ _ "-

i \ _i_

ol_e_l

_

I °_

i

.I

_

I

i _

I l,

_ z

I Io

a_n_Xll_l

44

o _"

m

O CD

-Im

"O

2 O
II

r4) e0 Q.

E 0

0

,.t e4

O

z

2

__

U.

II

"_ o X iii

o0 o0 O.

CXl

45

,,,o.

0 _

o,,go

•_

__

0

•_

__

O_ _

__ ..... _0_

Z_

__

__

Q.

E Q. ,m

0 Q. C O)

_0_ __

Z_ _.,

_0_

a S

0__ 0 II I..i

(1) tt

_

_0_0 9

0 ..Q

9 9

X_ a_

4

L._

=1

00000

0__

I-. 9

>.,, "0 =;

00000

00000

U) G) r..

9 9

__0 __0

O) r. ILl

Z_ _

"0 4) (,1 C a > "0

_

__0 o00_ o0000

Z_ _ _

__ 000_ 00000

_00o0

000oo

M W -M

0

u.; II

r_ 134-_

Cq D-'- t-- t-- t-.uO {0 _'0 _0 O_

0

_0_ _00_

r_

C_ r,,,Oo_ gO

_,_

_1C_ 0_ 0 b'_ 09 cO "_; -'_

0

_0_ _0_ _0_

M W W

W

.,0 m

I-

\

_

00000 00000

00000 00000

00000

00000

00000 00000 _0000

00000 00000 _0000 _0_0

W W W 96

_ _

W

46

.w,,0

O_

io,°o

__

O_

OO090_O •

0

_---

od_do

[-_'-_

(DQ0(D(D(D

0

,:::.-.

_0_

£-4,-.

_0_ 00_

0(00040

_0_0_

._

_d.dd

"

..-.

__0 _0_ _0_0

_) 0,1 0,1 0,1 ¢,,1 ,'-0 0

0

0

0

00000

_

00000

O_ • _ _

_OOOO __ 0__ .... 00000

"_.o_

_._

__ _0000

_

_g

•

00000 _0_

0_1 00000

_0_

0

.,--I

0_0 _0_ __0

0__ .,-4

00000 00000 _0000 _0_0

00000 00000 _0000 _0_0

D>m

m>m

47

0

0 0 0

0 0 _

0 0 0

0 0 _

0 0 0

0 0 _0

0 0 0

0 0 _

(e!sd) eJnsseJd e6Jeqos!(] dLund

48

0 0 0

C) 0

0 C)

C) 0

C) C)

(e!sd) eJnsseJd

0 C)

0 0

0 0

e6Jeqos!O 49

0 0

dLund

C) ¢)

3.1, Design the

and Parametric

design

maximum

at a chamber culated

at MEOP.

pressure

within

vative

the actual

design

feature

would

point.

followed

Figure

between

3.1-15).

3.1-16).

ance.

tables

for accurate

capability

architecture Power

ability

lbf thrust

example results sures Tables

are much

lower

through

the power

what mance. over

and

balance

over

the

entire

the maximum For the 2500

psia

Rl'r/I)0417 5,_,,I-3.0-3.1.2

chamber

20,000 giving

was

thrust

wheel was

parallel

the

range

circuit

up

to assist

if the

in NBS

in the bal-

VAX

in

iterations

or NASA

property

spread

(see

the iterations

This

or the

coupling

circuit

adjustments

cycle.

of the

diameter used.

flow

data

derived lookup

computers.

sheet

The

program.

ability

a 25% overthrust

of the

on TPA

performance.

Figure

3.1-18

gives

operating

version

but

the chamber

temperatures ratios

of the cycle The

would

be, given

capability.

50

program

the

In this power

pressures 7.

balance

envelopes

For the 50,000

pres-

is the same. for engines

psia

were

the baseline chamber

and

A gratifying

a 2000

for

is the

pressure

and

to sustain

results

temperature

of 5, 6, and

engine

the maximum

balance

features

The

of interest.

engine

the power

engine.

at mixture

pressure

lbf thrust

time

summarizes

the important

50K lbf thrust

work

the oxygen thermal

be adjusted

is looked

One

20K lbf thrust

give

is a useful devel-

hydrogen

can watch

in the

3.1-17

thrust.

diameter

3.1-12

can

on a commercial

of turbine

for the

It is a conser-

balance

significant

the

on the PRIME

in Figure

flow

not neces-

pressure

the power

cal-

Results

at rated

wheel

for the parallel

of 20K lbf thrust

sure

chart

the effect

a 2.35 inch

3.1-7

Balance

engine

to iterate

be run

is based

data

point

chamber

most

the operator

property at each

that

real

was

levels.

by balancing

to balancing

make

to be

and temperature

thrust

psia

expected

3.1-14.

is the

inputs

and

Propellant

engine

begins

rate

in that

terminal

The the 7.5K

transfer

several

the program

3.1.2.2

program

was

This point

modeled,

in Figure

as an entry

calculations

requires program

heat

point

the pressure

at 2300

the components

program

closing.

Also,

at all five

as shown

has

operating

calculations.

capability

balance

operator

on the computer it is not

actual

path

It is used

The

show

envelope

With

HEX

It is an interactive

progress

for stress

power

The

the circuits.

Figure

be used

design.

The (see

and MR = 7.

operating

the logic

limiting

psia

The overthrust

of the engine

opment

(cont)

The power

of 2300

for this point

sarily

Analysis,

chamber explored

component

pressure lbf thrust

result

presto see perfor-

at MR = 7 was engine

of

the

r

e.,

(o

0.--_

u

_ 0 elU

I11

0 n m O) c N Im im

L.

E 0

Q

_._

C

,'_>

eO. U

l

l

0

ffl

0 -J

4 |

e4

,m

11

>

if)

51

0

,4,,o i.. Im

IJ,.

eB

m O0 It

In

.m

3<

@

e-

|

L_

im

IJ.

E ::3 E .m .m

E II

O.

52

i !-

oo

•,ii_

,l_

iii

X

Om

w _-

E ®

_® _

=: o':,-u.

-,

" ;

_" R

o'._

4)

Ig

,'_

.,

ai-

4..,

(U

g.

•

_

v

" E --

.--

i.,

__

(0 g

Ii_

q,,_ "l"!

=

_

_:

•

•

•

tU

_

•_l II

I--

i_.

II

,ii..IIIL

_ o..-c-_

e.. (_.l

"

I

II

I-- _." II

II

P,°

)

[ o. _._ _ 2'-_ _ @r-

k'_'/-\_ ) _ _ l__h_ o_

_

_O

53

-I _I _-a-I _.L_.

II

nIll i'rI,LI m

v 0 ¢"i1 L

0 ul -I ill I'1" ¢)

m

0 13.

|

e4

.i

U.

55

TABLE

3.1-7

CTP ENGINE POWER THRUST = 20K lbf

BALANCE MR = 5

Oxidizer Tank

Conditions

Pump

Conditions

Ox:

Cool

Side

Heat

Fuel:

Exchanger

Cool

Side

P= T=

15.0 162.7

(psia) (deg R)

20.0 37.8

P= T=

3629.8 182.4

(psia) (deg R)

4543.4 93.5

(psia) (deg R)

P= T=

3319.7 537.0

(psia) (deg R) 4363.5 478.4

(psia) (deg R)

4122.8 601.5

(psia) (deg R)

4319.1 994.1

(psia)

2595.7

(psia)

p_

T=

Regenerator

Regen

Fuel

Jacket

p_-

T= Baffles

p_

T= Ox Nozzle

Turbine

Hot

Side

Cooling

Conditions

HEX

P= T=

3292.2 860.0

(psia) (deg R)

P= T=

2307.6 794.4

(psia) (deg R)

p_-

T= Gas Side

Regenerator

p_

T= Injector

Combustion

Chamber,

Pc

P=

2009.6

T=

800.4

(psia) (deg 2000

R) psia

723.2

(psia) (deg

(deg

(deg

R)

R)

R)

2409.8 462.0

(psia) (deg R)

2224.1 373.9

(psia) (deg R)

2003.3 373.9

(psia) (deg R)

TABLE

3.1-8

CTP ENGINE POWER THRUST = 20K lbf

BALANCE MR = 6

Oxidizer Tank

Conditions

Pump

Conditions

Ox:

Cool

Side

Heat

Fuel: Side

Exchanger

Cool Regenerator

Regen

Jacket

Baffles

Ox Nozzle

Turbine

Hot

Cooling

Conditions

Side HEX

Gas Side

Regenerator

Injector

Combustion

_r/_7.ss,.r

P= T=

15.0 162.7

(psia) (deg R)

20.0 37.8

P= T=

3764.2 183.1

(psia) (deg R)

3958.7 87.1

(psia) (deg R)

P= T=

3441.0 555.0

(psia) (deg R)

P= T=

3822.5 466.3

(psia) (deg R)

P= T=

3635.0 646.1

(psia) (deg R)

P= T=

3782.5 1012.5

(psia) (deg R)

2397.5 759.5

(psia) (deg R)

P= T=

2300.5 435.6

(psia) (deg R)

P= T=

2169.8 365.7

(psia) (deg R)

2002.9 365.7

(psia) (deg R)

P= T--

3411.3 855.0

(psia) (deg R)

P= T=

2343.3 786.1

(psia) (deg R)

P= T= Chamber,

Fuel

2015.4 792.5

Pc

(psia) (deg R) 2000

57

psia

(psia) (deg

R)

TABLE

3.1-9

CTP ENGINE POWER THRUST = 20K lbf

BALANCE MR = 7

Fuel

Oxidizer Tank

Conditions

P= T=

15.0 162.7

(psia) (deg R)

20.0 37.8

(psia) (deg R)

Pump

Conditions

P= T-

3875.3 183.8

(psia) (deg R)

3705.1 84.9

(psia) (deg R)

P= T=

3547.1 586.8

(psia) (deg R)

P= T=

3600.7 403.7

(psia) (deg R)

P= T=

3393.0 675.9

(psia) (deg R)

P=

3573.1 1009.7

(psia) (deg R)

2305.8 776.7

(psia) (deg R)

2231.8 377.0

(psia) (deg R)

2133.4

(psia)

Ox:

Cool

Side

Heat

Fuel: Side

Exchanger

Cool Regenerator

Regen

Jacket

Baffles

T= Ox Nozzle

Turbine

Hot

Cooling

Conditions

Side

P= T=

3512.8 856.8

(psia) (deg R)

P= T=

2378.2 785.2

(psia) (deg R)

P=

HEX

T= Gas

Side

P= T=

Regenerator

P= T=

Injector

Combustion

Chamber,

Pc

349.4 2011.6 792.9

(psia) (deg R) 2000

psia

2002.9 349.4

(deg

R)

(psia) (deg R)

TABLE

3.1-10

CTP ENGINE POWER THRUST = 50K lbf

BALANCE MR = 5

Oxidizer Tank

Conditions

Pump

Conditions

Ox:

Cool

Side

Heat

Fuel: Side

P= T=

Exchanger

Cool Regenerator

Regen

Jacket

Baffles

Ox Nozzle

Turbine

Hot

Cooling

Conditions

Side

Gas Side

37.8

(psia) (deg R)

5555.5 104.7

(psia)

P= T=

5479.5 606.9

(psia) (deg R)

P= T=

4044.3 554.7

(psia)

P= T=

5376.1 1354.0

(psia) (deg R)

2504.7 862.7

(psia) (deg R)

P=

3670.7

T=

182.9

P= T=

3404.6 560.0

(psia) (deg

R)

(deg

R)

(psia) (deg R)

P= T=

3317.4 860.0

(psia) (deg R)

P= T=

2311.9 793.2

(psia) (deg R)

(deg

R)

P= T=

2351.7 585.6

(psia) (deg R)

Regenerator

P= T=

2224.1 448.5

(psia) (deg R)

2003.6 448.5

(psia)

P-T= Chamber,

2006.2 799.3

Pc

(psia) (deg R) 2000

59 RPTI DO417-SSa-T

20.0

(psia) (deg R)

HEX

Injector

Combustion

15.0 162.7

Fuel

psia

(deg

R)

TABLE

3.1-11

CTP ENGINE POWER THRUST = 50K lbf

BALANCE MR = 6

Oxidizer Tank

Conditions

Pump

Conditions

Ox:

Cool

Side

Heat

Fuel:

Exchanger

Cool

Side

R)

20.0 37.8

(psia) (deg R)

3800.3 183.6

(psia) (deg R)

4722.9 95.9

(psia) (deg R)

3523.0 582.0

(psia) (deg R) 4665.6 589.1

(psia) (deg R)

3588.6 589.6

(psia) (deg R)

4577.5 1368.6

(psia) (deg R)

2357.1 896.7

(psia) (deg R)

2253.1 550.2

(psia) (deg R)

2169.7 453.1

(psia) (deg R)

2004.6 435.1

(psia) (deg R)

P=

15.0

T=

162.7

P= T= P= T=

(psia) (deg

p_

T=

Regenerator

Regen

Fuel

Jacket

p_

T= Baffles

p_

T= Ox Nozzle

Turbine

Hot

Side

Cooling

Conditions

HEX

P= T=

3425.8 860.0

(psia) (deg R)

P= T=

2346.0 790.1

(psia) (deg R)

p_-_.

T= Gas

Side

Regenerator

p_

T= Injector

Combustion

Rl'l/i;_o417-_,-'r

P= T= Chamber,

2006.1 796.9

Pc

(psia) (deg R) 2000

60

psia

TABLE

3.1-12

CTP ENGINE POWER THRUST =50K lbf

BALANCE MR =7

Oxidizer

Fuel

Tank

Conditions

P= T=

15.0 162.7

(psia) (deg R)

Pump

Conditions

P= T=

3943.0 184.4

(psia) (deg R)

P= T=

3658.1 607.8

(psia) (deg R)

Ox:

Cool

Side

Heat

Fuel: Side

Exchanger

Cool Regenerator

Regen

Jacket

Baffles

Turbine

Hot

(psia) (deg R)

P= T=

4149.4 562.0

(psia) (deg R)

P= T=

3315.1 611.3

(psia) (deg R)

P= T=

4058.4 1399.9

(psia) (deg R)

2271.3 929.5

(psia) (deg R)

P=

2194.8

(psia)

T=

506.4

P= T=

2135.1 419.6

(psia) (deg R)

2004.8 419.6

(psia) (deg R)

P= T=

3553.1 857.8

(psia) (deg R)

Conditions

P= T--

2387.2 784.6

(psia) (deg R)

Side

Gas Side

HEX

Regenerator

P= T=

Injector

Combustion

Chamber,

2011.6 792.2

Pc

(psia) (deg R) 2000

61 RITr/DOIIT-_a-T

4193.9

(psia) (deg R)

90.3

Cooling

Ox Nozzle

20.0 37.8

psia

(deg

R)

3.1, Design

and Parametric

maximum

chamber

increasing

thrust,

completed.

With

was

kept

Analysis,

pressure

was

these

2100

was

encouraging

psia

at MR = 7.

expected

results

before

There

is a dropoff

the study

the baseline

of 2000

with

calculations

psia

were

chamber

pressure

the study. Additional

ratio

about

but it is less than

throughout

mixture

(cont)

power

balances

are given

in Appendix

B for the high

operation.

3.1.2.3

Modified

Liquid

Engine

Transient

Simulation

Program

(MLETS)

Analysis The power assumes

steady

the engine but adds MLETS

state

operation. the

time

for several

years

the OTV power

recent

presented

balance

to make below

show

with

program

features version

discussed

in the previous

no consideration

of the time

performs

mass

similar

of the components

and

of the LETS program

which

Its complexity

program.

is about

an order

It was

also developed

to the OTV engine.

In its present

it easily

the OTV

usable

with

it to be a valuable

engine

engine,

tool for predicting

dependency energy

system.

of

balances The

has been

in development

of magnitude

greater

for other form

and

paragraphs

engine

systems

it still requires

and

some

but the preliminary engine

than

results

operation

and

requirements. a.

able insight lines.

MLETS

at Aerojet.

to be adapted

development

control

The

program

operation

dependent

is the most

has had

engine

balance

into many

Concerns

Uses

of the MLETS

design

of current

•

concerns

interest

Power

Program

as it includes

--

The MLETS

models

can provide

for all components

valuand

are:

balance

(backup

to the independent

power

balance

program).

•

Engine

•

Control operating

Rl-r/ixnlT.SSa-3.0-a.l

._

sensitivities

to component

sensitivities

to component

scenarios.

62

design

changes.

operation

and engine

3.1, Design

and Parametric

Analysis,

(cont)

•

Engine

stability

•

Engine

transient

start, •

more

than

the concerns

given

above b.

engine

schematic

numbers

refer

to the TPA turbines.

to the hydraulic

The HEX

the MLETS

performance pressure

components.

change

operating

is limited

is suspect

performance

while

until

are measured.

design

tool.

was

can be fed back

found.

to the component

insensitivity

very to control

scenario

calls

MLETS

analysis

changes

for thrust confirmed

over

Capability. the

throttle

by use

this as the preferred

_SSa-3.0-.3.1 2

for the TPA

to model

that these

of component rates

of temperature

efficiencies, until

point

has been

for details

range.

This

circuit.

The

side

reflects current turbine

and

and other

a relatively makes

can be adjusted

The oxidizer

scenario.

of

The LI

be evident

used

lags,

of the oxygen

63 RPT/DO417

It should

and

to be used

in the hydrogen

to be adjusted

design

None

The simplified

3.1-19.

operator

design

team

--

to do this in real time

that a stable

to do

PU 101 and PU 201 refer

dependency

by the program

design

insufficient

to the models

the component

the component

Bootstrap

effectively

line.

are built

The ability

shows

1. "bootstrapped"

adjusted

In effect,

this program

in Figure

of the algorithms

components

In practise,

were

configuration

When

actual

structure.

operation.

Analysis

is HE 102.

the time

without

desired.

TU 101 and TU 201 refer

by the accuracy

were

transient

for that particular

In particular,

parameters

a valuable

is given

configurations to its modular

to this task

of the MLETS

is HE 101 and the regenerator

analysis

and the other

circuit.

sections

due

in the depth

analysis

model

engine

of the engine

Results

This includes

shutdown.

assigned

addressed

for the MLETS

and linearity.

reprogramming

assessment

were

time

of different

hours

Preliminary

used

pump

code

program

a preliminary

and

assessment

extensive

much

response

throttling,

Ready

The

in operation.

stable

the MLETS

to match

the

determined of the actual of the

this design.

circuit

its relative engine bypass

operating valve.

The

Fuel Circuit

OX Circuit

BC 101

I BC201

LI209

15

LI101 LI201

12

38

_1. 23

16'

E101 17

I 3

_,_,,,_ 15_

LI102

37 18

TUI01 11,

LI202 44

122 LI103

19

21

2O

f

LI203

29

LI211 45 24

LI208

LI204

36 LI105

35

LI205

38

27

10

LI207 r46 26

32

25

LI206

LI 104 Iog16.1.9

Figure 3.1-19.

CTP Engine Model for MLETS 64

Analysis

3.1, Design

and Parametric

Analysis,

(cont)

The HEX bypass interaction

with

the

hydrogen

particular.

This

interaction

circuit

valve

and hydrogen

can be reduced Allow

a small

points

limiter

Ox TPA

This

bypass

Reverse the

engine

given

There efficiency

was

but

the required

may

be an error

"bootstrapped," indicate. energy

There should

be more

was

adjusted

in a larger

than

bypass

There and

the proposed

has not been energy mum range, valve oxygen thrust

developed

extraction value.

approaches temperature

HEX bypass

_-,r/_7_,-3.0-312

operations for the HEX

in the

As long

the oxygen

control

control

HEX

as the

to the turbine

valve

was

bypass

turbine

be limited

With

in

until

oper-

it must

going

act

into

"hunting"

HEX

and

the

of the

HEX

the regenerator decouples

circuit.

schematic

in

the There

may

to the lower

delta

for this option

is

due

problem

bypass

far lower that

the

has

throttle

bypass

range

valve

must

be

marks

the present

As the

response. algorithms

An algorithm sufficient

This is a maxi-

the normal

oxygen

the bypass

schematic,

to reduce

Available

assumes

the end of the

engine

to a few set points

65

is within

would

detected.

coupling.

cir-

circuit

requirements

is at 400°F.

reduce

the

the MLETS

program

turbine

until

for quick

between

than 400°F.

That

not been

and the

to the

bypass

energy

the circuit

valve

in the hydrogen

decreased than

one mismatch

is reached.

side.

oxygen

sized

that increased

the HEX

until 400°F

should

over

can be less

10% bypass,

on the oxidizer

was

setting

3.1-20.

turbine

so that oxygen oxygen

for the

the oxygen

A circuit

algorithms

adequate

valve

a "bootstrapping"

and

in the

bypass

effectively

from

in Figure

valve

dependent

will prevent

This

circuit

temperatures.

of thrust

of the

circuit.

a penalty

bypass

valve.

the position

hydrogen

Turbine

number

as a temperature turbine.

considerable

of ways:

for the HEX

regenerator

causes

regenerator

in a number

only

ating

cuit.

placement

control

turbine

bypass

and

increase

usable

range

movement

the circuit

the for of the

coupling.

w f_jffWO_ff_JUa_ff_Wf_ffOU_ad

w

W

0

66

3.1, Design

and

Parametric

Analysis,

2.

Operating

shows

the engine

variation

engine

at thermal

equilibrium.

without very

any

control

minor

drift.

attributed The

to minor

engine

analysis

The

does,

design

is so slight

however,

using

oxygen thrust

integral

bypass

change

ber of runs

but

and

from

Engine

dynamic

made

better

used

There

a solution

cycle

completed

RI_/DOt

engine

control

the

shown

as easily

a

be

instability. close

to it. More

condition.

be readily

should modeling

called

in about time.

in the turbopumps 10% thrust This

is an important

to be better

17 .$5a-3.0-3.1.2

to 100%

defined

Thrust

hydrogen

What

controllable. rates

was

thrust

not enough

come

from

7.5K

circuit

controller

bypass

should

be

time

available

and

flow

lbf thrust

lags

engine

(See Ref. 7). Throttle

in Figures

ratio

3.1-22 The

and

flow

at low thrust

speed.

The

engine

in 4 to 5 seconds.

performance early

time)

number

in the engine

6 7

case

the engine

even-

this control

prob-

through

I IF.X

modeled

should Throttling for the

development.

a different

for this series

expander

decreases

the

using

A 10% change

dual

A num-

system.

response

3.1-23.

parallel

(A thrust/unit

was

plus

commands.

coupling

in the

to follow

Proportional

to solve

the circuit

was the

attempted

ratio

In each

throttling

by changing

valve.

mixture

gains.

reducing

of the thermal

controlled

and

controller

of engine

was

turbine

the

different

0.3 seconds.

Throttle

the

for both

TUTSIM.

is shown

A simulation

the hydrogen

using

unstable.

code

expander

needs

with

have

or very

slow

3.1-21

to operate

the actual

should

Relatively

circuit.

while

The

faster.

point

to establish

Throttling.

for each

setting

were

were

went

lem,

from

it could

as to actual thrust

point

variables

that

the engine

in Figure

is allowed

the operating

so regular

control.

plot

operating

engine

algorithms

that

loop

by adjusting

controllers

to lags

the

will be needed

is for closed

thrust

at the rated

is confirm

a controller

turbine

response

in the

composite

for stability.

attempted

tually

the rated

and

stable

testing

3.

the

from

The

By t = 148 seconds

dynamically

baseline

adequate

time

off errors

experimental

the analysis

Stability.

At t = 126 seconds

change

round

is either

and

with

changes. The

(cont)

flow

in thrust

should

dual

could

have

a similar

in the low

thrust

be capable

of accelerating

down vehicle

be

range

due

will be somewhat prime

contractors

and

0

k.

r-

4me

I

00

n-

m

i Im

O) 1

LLI •-

g) ¢3 eCO

,(

O)

O i

•

elsd 'Od 'emsse:d _eq,.,eq:_

,

r

d/O 'olleH eJnlXlW

1

!

r

tud: 'peeds dmnd -oqJnl ueOo, pAH

G8

|

r

tudJ 'peed S d-,ndo(pnj. ue6_xo

106.4

I

I

I

I

104.4 _102.4

I

I

_ j

Mixture Ratio

7.0

.

6.0

_

___

'i

1

I

,----,

_ 100.4

5.0

98.4

4.0

96.4

3.0

_

94.4

o.

92.4 90.4 88.4 86.4 1.0000

1.5000 Time, sec

Figure 3.1-22.

Predicted

Response 69

to 10% Throttle

Up Command

_

104.4

Mixture

lO_4_[,,,

,R.t,o, ' '

102.4

7.0 .__,-

6.0

100.4 98.4 96.4

I X

_,---'_z-----'-_

Thrust (Actua)

| 3.0

n" 94.4 o.

92.4 90.4

86.4

I

I

I

I

I

!

I

0.500000

I

I 1.0000

Time, sec

Figure 3.1-23.

Predicted

Response

to 10% Throttle

7O

Down Command

°"

3.1, Design

and Parametric

3.1.3

Performance,

3.1.3.1

Performance

formance

measurement

formance

comparison

area ratio

(Ae/At)

is shown

the One

losses

using

engine

Layer

the 7.5K

lbf thrust

layer

Table over

ered

specific

Also,

impulse.

the maximum

with

only

minor

tage

of the vehicle

can program with

changes

is no significant

specific variations

impulse from

designers

mixture

no delivered

selection

ratio

system

average

specific

engine

specific

impulse.

impulse.

The dropoff composition weight

of the exhaust

increases

causing

thrust

in changing

is delivered

over

at MR = 6.3. an active

this range

to use

The result

This is usually

gas species. a reduction

and with

first

then corrected

for

further

corrections

and boundarv

program)

is plotted

variation

was

for divergence Kinetics

and

specific

(ODE)

and

in Figure

corrections

impulse At higher

thrust

on other a mixture

all available

BLM

3.1-25

for

for 7.5K,

ratio

important

above

the range

to improve

design

range

ratios

the

Vj, according

system

that

in the tanks

of the propulsion the nominal

MR = 6.3 reflects

mixture

of 5 to 7

to the advan-

propellant

than

deliv-

considerations.

management

is a maximization

more

over

This can be used

propellant

in the jet velocity,

71

minor

engine

1/2

Rr'r/DOi17_SS.-3.0-3 aa

psia

thrust

performance

are very

be based

in specific

and

performance

should

penalty.

of 2000

per-

of 5 to 13.

by baselining

performance

An initial

engine

(ODK)

Dimensional

with

the peak

within

program

loss accounting

advantage

Thrust

model

summarizes

an MR range

Performance There

3.1-13

for per-

This plot is for delivered

The corrections

(Two

pressure

on the X-axis

3.1-24.

losses.

standard

specified.

a chamber

plotted

Kinetic

The performance

engine.

area ratio

Equilibrium

TDK

is the best

The theoretical

a One Dimensional

20K, and 50K lbf thrust

studied.

on Figure

Dimensional

Model).

used

was

operation.

using

and

range ratio

Parameters

impulse

propellants

Mixture

are calculated

_Boundary

and Envelope

N Specific

both

and boundary

lossess

Mass

of 3 curves

from

(cont)

the thrust

of 1200.

state

for divergence layer

over

at steady

calculated kinetic

with

as the family

impulse

Analysis,

average

the changing molecular

to the equation

,,e.._

V

"_l') ('_1 r,.. 03

•r ="

S

.#=-

0

°_

t_

rr

o e" 0 r,, W 'O Q o c al > 'O m

x

--

CO

,I{ ,4 |

Q L.

OR

ti. --f43

O O)

I

I

o 03

O I_

I

I

i

I

I

,I

O (.10

O I.o

O ,,_

O c")

O C_l

O v-

I O O

uJql 1 dSl peJe^!loa oes-tqu

72

I O O'J

I O 0O

!

o

6

w_ _a OO

D

a

_

O,I

C m ,4.,,*

t.._ 0 U U

O

.< O .D

r_

(n o _1 Q

E 0 -- aO Q

ft.

|

--

.o in

I,L

-

--

I O O 14)

O O_

I

O 00

I O

I O r,D

I O

I

I

I

I

t

O

I O

O'_

_

"--

O q,)

mq I ] oes-lqll

73

dSl _e^lle

G

I O co cO

O o0

(.O

TABLE PERFORMANCE

LOSS ACCOUNTING

THRUST

LEVELS

7.5K

Pc=2000

MR ODE ODK TDK ABLM P.I. & Del

3.1-13

(7.5K,

lbf Thrust 6

7

491.1 490.0 488.0 7.4 480.6

494.4 492.5 490.5 7.4 483.1

494.5 490.5 488.5 7.4 481.1

Radius,

20K lbf Thrust 5

11 409.7 424.9 423.2 7.4 415.8

12 413.9 409.8 408.2 7.4 400.8

13 399.7 396.2 394.6 7.4 387.2

rt = 0.765"

Level

MR ODE

491.1

494.4

6 494.6

11 429.7

12 413.9

13 399.7

ODK TDK ABLM PT & Del

490.3 488.3 6.6 481.7

492.9 490.9 6.6 484.3

491.4 489.4 6.6 482.8

425.7 424.0 6.6 417.4

410.5 408.9 6.6 402.2

396.8 395.2 6.6 388.6

Throat

7

Radius,

50K lbf Thrust

Pc=2000

50K lbf)

Level

5

Throat

Pc=2000

FOR VARIOUS

20K, AND

5

rt = 1.25"

Level

MR ODE ODK TDK

7

491.1 490.5 488.5

494.4 493.3 491.3

6 494.5 492.1 490.2

11 429.7 426.3 424.6

12 413.9 411.1 409.4

13 399.7 397.3 395.7

ABLM PT & Del

6.1 482.4

6.1 485.2

6.1 484.0

6.1 418.5

6.1 403.3

6.1 389.6

Throat

Radius,

rt = 1.97"

3.1, Design and Parametric Analysis, (cont) where M

__

Ru

The

=

average

exhaust

universal

gas

molecular

ratio

N j=

efficiency losses

T 0

=

stagnation

chamber

temperature

Po =

stagnation

chamber

pressure

P j=

jet pressure

To/M

ratio

of unreacted engine

of gas specific factor

at high

formance

steadily

mixture

relating

consistently point

oxygen

heat

at constant

pressure

theoretical

and

(MR

as mixture = 7.94).

increases

ratios

total

At any

(>7) must

propellant

other

flowrate.

engine

operating

Figure

3.1-26

gives

the propellant

per

engine

basis.

For a vehicle

with

a four

can be readily thrust

used

for a given

the flowrate used

288,000

to calculate

propellant is about

for the Lunar

system

propellant

weight

to determine

3.1.3.2

engine

presented

RFr/Do_7_,-3.0-3.1a

was

would

there

of

the amount

operation

completed

in Table

3.1-15.

derived

flowrates

of this

is no per-

propulsion

(LTV).

flowrate

detail.

This

(burn

20,000

Ibm

load

For a total and

For a 4 hour

propulsion

a correction from

total

is on a This

time)

at

lbf thrust is one

of the

operating

time

for attitude

the loaded

engine

of the

by four.

time

propellant

is

for each

set multiply operating

to be subtracted

about

performance

for a set of four

residuals

have

from

in more

propulsion

A 288,000

propellant.

Mass

A detailed

are

The

presents

For instance,

Vehicle

this represents

Engine

total

Propellant

the usable

maintenance,

engine

174 lbm/sec.

Transfer

use

the

load.

+ 174 = 1655 seconds.

control

thrust

engine

the maxima

stoichiometric

decision;

graphically

3.1-14

without

to the

despite

stream.

parameter

Table

of:

increases

be an economic

levels.

figures

volume

jet velocity

MR above

in the exhaust

thrust

engines,

delivered

ratio

five

rated

constant

justification.

One

figure

and

at the exit plane

decreases

stoichiometric

weight

constant

7=

To at the

the

gas

propellant

operating

time

8 missions.

Computations

weight during The

and

center

the preliminary interface

of mass

design

for separating

75

computation task engine

for the 7.5K

(Task weight

D.5). from

The

lbf results

vehicle

1

UJ

E m

w,--

o i== a. Q i

.Q

> ._J

.1

2 >

o

Q Q.

G_ @d |

o L--

°l

LL.

I

c)

o o

I

o cO

0 m"

1,--

76

o

o

TABLE LTV/LEV

ENGINE

Thrust

3.1-14

PROPELLANT

7.5Klbf

Propellant

Flow

•

Total

•

Oxygen

•

Hydrogen

(lbm/sec) (lbm/sec)

Tank

•

Total

•

Oxygen

•

Hydrogen

Propellant

Pressurization

Flow

(lbm/sec) (lbm/sec) (lbm/sec)

Flow

•

Total

•

Oxygen

•

Hydrogen

20Klbf

25Klbf

35Klbf

50Klbf

Rate to Engine

(lbm/sec)

Autogenous

FLOWRATES

16.38

43.49

54.36

75.96

108.29

14.04

37.28

46.60

65.12

92.82

2.34

6.21

7.77

10.85

15.47

0.82

2.17

2.72

3.79

5.41

0.70

1.86

2.33

3.26

4.64

0.11

0.31

0.39

0.54

0.77

15.56

41.32

51.65

72.17

102.88

13.34

35.42

44.27

61.86

88.18

2.22

5.90

7.38

10.31

14.70

Engine

for Combustion

(lbm/sec) (lbm/sec) (lbm/sec)

Notes: 1.

Autogenous

propellant

2.

Table

on a mixture

°

based

Propellant extrapolation down.

flow ratio

assumed

to average

to engine.

of 6.

flow at thrusts below nominal are higher than would predict as specific impulse decreases

77 RPT/D0417-SSa-T

5% of flow

a straight line as the engine

is throttled

U.

113

78

3.1, Design

and Parametric

weight

the mounting

was

sidered

above

interface ation

define

exclusive

and engine

additions

be defined

to the vehicle

(cont)

The gimbal

weight

some

structure

contractor

of all items

in Table

For example,

is a reasonable

system

summary.

lines

3.1-16.

definition

of engine

a figure

for engine

valid oper-

the mounting

other

interfaces

dry weight weight

con-

for engine

above

Several

the engine

were

An equally

necessary

and propellant

are given

validity.

and ICHM

in the weight

as the sum

takeout

and a total

with

actuators

are not included

of the thrust

These

could

plane.

this plane

would

plane.

Analysis,

as delivered

based

on a contract

requirement. Table could

be applied

3.1-16

after engine

this is commonly

a function

snubbing

to secure

system

There

is no weight

Those

engines

vehicle

weight,

required

weight

that a helium and would

A lightweight

is unlikely

to be used.

favor

one material

Table

3.1-17

for the final

material

over

contrasts design,

ture, and

part on a trade-off

Advanced components

Engine

Study

as would

P,PT/rx_17 s5,-30-3.1.2

3.1-27

normally

versus

same

basis

several

in Table

be delivered

before

hundred

is the radiation

cooled

any one of which lightest

requirements, delivered

3.1-18.

is material

even

when

though

resolved,

nozzle could

payload.

Part

of the

temperaof pay-

3.1-17. used

that this is the contractor

will

be selected

An example

of Table

it

material.

part on operating

by the engine

79

engine.

to heaviest.

for the engines Note

to the

estimates

that,

it

but it is

cycle

uncertainties

results

pounds

can be proposed

ratio from

without

comparisons

weight,

weight

for the four nozzles

computation

one.

data are cautioned

in engine

nozzles,

on structural

not require

weight

expander

of variability

is retracted.

of engine

of a conventional

of four

are given

does

as

for any nozzle

section

as engine

of weight

The weight

nozzle

be considered

An example

in Figure

is no weight

structure

weight

can add

is a 2:1 weight

will be based

is given

system

takeout

this

engine

on the

may be design

the weights

criteria

sensitivity

there

yet there

Users

although

compute

still in development

another.

selection

load

source

Also,

often

computations

assembly

Another selections.

system

not normally

for the turbopump

There

as this engine

system.

purge

insulation thrust

the extendible

system

a helium

that it is a vehicle

not include

design.

when

purge

that do require

Note

It does

of the vehicle the nozzle

manufacturers

are made.

delivery.

for a helium

on the assumption to put each

includes

in the weight

for the

to the vehicle

TABLE 7.5K

LBF THRUST

PRELIMINARY COMPLETE

3.1-16 ENGINE ENGINE

Component

Current

Propellant

Flowmeters

Hydrogen

Main

Oxygen

Main

Primary

Gimbal

Engine

Out

ICHM

System

Shutoff

Estimate

3.0 Ibm 8.5

Valve

8.0

Valve

Actuators

ESTIMATE

Weight

(4)

Shutoff

Gimbal

WEIGHT

17.0

(2)

14.0

Actuator

12.0

Electronics

Insulation

10.0 Sub-Total

Nominal (From

72.5

Gimballed Component Table 3.1-15)

Weight

298._____!

TOTAL

370.6

TABLE PRELIMINARY

NOZZLE 7.5K

3.1-17

SYSTEM WEIGHT LBF ENGINE

Columbium 0.020-in.

ESTIMATES,

Carbon-Carbon 0.030-in.

0.050-in.

0.060-in.

Nozzle

Skin

48.3

72.4

20.6

24.7

Nozzle

Attach

18.1

18.7

7.0

7.0

Nozzle

Stiffener

4.7

4.7

1.5

1.5

(3)

8.4

8.4

8.4

8.4

8.1

8.1

8.1

8.1

5.4

5.4

5.4

5.4

9.0

9.0

9.0

9.0

4.2

4.2

4.2

4.2

5.0

5.0

5.0

5.0

69.2

73.3

Ballscrew Gearbox Ball

(3)

Nut

28VDC

(3) Motor

Flex Cable Support

(3)

(6) Strut

TOTAL

(6) (per

I
TCA)

111.2

135.9

80

O

O

O

6

6

.Q

.D

C.)

l .N¢.=

.=

¢.=

0_

in o

co o

o

o

0_

&

6

O O

O

0

I11 4-U

.m

o'_ o II1

II

D.

J_

0 Z

¢=. (P

_o

>

N N O Z

"0 N 0 Z

u

i

0

>,,

D.

O I-

m

|

o

(P

s0 •

._o.> O

."

0

ql 'POl,_ed

81

etleO

e_

TABLE

ADVANCED

FLIGHT

3.1-18

ENGINE

WEIGHT

ESTIMATES

Enqin_ Thrust, PoundsForce Material

Component

7.5K

20K

2_K

35K

50K

7.5K

W_i_t in Povnds

20K

25K

35K

50K

Percent of Total Weiqht

GlidCop & NiCo

Thrust Chamber

36.66

64.51

84.95

131.34

197.63

14.79

13.47

13.93

14.68

14.51

Ni Base

Injector

24.08

45.97

56.24

79.07

121.15

9.71

9.60

9.22

8.84

8.89

ZrCu

Baffles

8.50

9.79

14.03

22.08

35.17

3.43

2.04

2.30

2.47

2.58

69.24

120.27

155.22

232.48

353.95

27.93

25.10

25.45

25.98

25.99

TCA SubTotal Be

Ox Cooled Nozzle

14.20

38.80

49.50

71.80

109.10

5.73

8.10

8.11

8.02

8.01

C-C

Rad. Cooled

73.34

143.95

175.11

248.50

378.37

29.58

30.05

28.71

27.77

27.78

Ox TPA

10.00

16.67

23.33

36.67

56.67

4.03

3.48

3.83

4.10

4.16

Ox Boost

18.60

31.00

43.40

68.20

105.40

7.50

6.47

7.11

7.62

7.74

10.00

16.67

23.33

36.67

56.67

4.03

3.48

3.83

4.10

4.16

6.40

10.67

14.93

23.47

36.27

2.58

2.23

2.45

2.62

2.66

45.00

75.00

105.00

165.00

255.00

18.15

15.66

17.21

18.44

16.72

0.84

7.12

7.10

7.06

8.58

0.34

1.49

1.16

0.79

0.63

3.98

14.10

16.40

20.90

30.00

1.61

2.94

2.69

2.34

2.20

4.82

21.22

23.50

27.96

38.58

1.94

4.43

3.85

3.12

2.83

41.32

79.85

101.67

149.15

227.00

16.67

16.67

16.67

16.67

16.67

SubTotal

206.60

399.23

508.33

745.74

1134.99

83.33

83.33

93.33

83.33

83.33

Total

247.92

479.08

610.00

894.89

1361.99

100.00

100.00

100.00

100.00

100.00

Orig. Total

291.80

486.33

680.87

1069.93

1653.53

Nozzle

Pump Fuel TPA Fuel Boost Pump TPA SubTotal Be

H2/H2 Regenerator

Be

H2/O2 HEX

HEX & Reg SubTotal Valves, Lines, & Misc.

3.1, Design

prime.

and Parametric

A graphical

based

on materials

choice

could

add

Analysis,

presentation choices

10% to the totals

of interest.

28 lbf thrust

3.1-29

for each

pound

is given

This plot

can be used

3.1.3.3

thrust,

formance very

large

lbf thrust tion

engine

cooled

lbf thrust

psia

it is 106 inches.

brake

or of using

design

should lengths

using be used in terms

the

weight

at intermediate

for the nozzle

is plotted

thrust

with

directly

over

a ratio

of the

thrust

is an area

performance

programs

with

of

same

cal-

against

thrust.

points.

in conjunction

length

engines

engine

behind

they

and nozzle in Tables

Figures half

3.1-32

used

area

ratio.

3.1-31.

and

Note

enough

do not

interact

nozzle

such

3.1-19

and

with

which

defines

engine

dimensions

each per-

with

the radiaAt the

to warrant with as a plug

the

the symbols

83

study

is given

aero-

cluster.

thrust These

2(IK

an

can be tables and

section.

of the basic

is a

that the 50K

doors.

for any 3.1-20.

a con-

The result

13.1 ft long

contour

with

maximum

the aerobrake

to a conventional

given

were

is serious

where

configuration

with

extended

of 1200:1

to deliver

in Figure

on the vehicle

an alternative

data

found

ratio

the selected

are shown

is the

the

was

and

the nozzle

mounting

parametric

nozzle

pressure

impact

of a rocket

length

3.1-21.

_r/_Tss.._.o-3.1_

This is

materials

engine

presentation

engine

The

A summary Table

graphical weight

This

The engine calculated

versus

20K lbf thrust

engine

half sections

retracted.

of either

3.1-28.

conservative

ratio

where

chamber

Engine

investigation

A more

at about

Another

110% RAO

is 25.4 ft long

nozzle

optimizes

Aerojet

and a common

engine.

3.I-18.

in Figure

Envelope

performance.

for the 2000

is given

the thrust/weight

cycle

to estimate

The criteria

engine

plots

3.1-30

Engine

for optimum

in Table

of engine.

in Figure

weights

as given.

This engine

culations

tour

of the engine

as indicated

Figure the range

(cont)

in

iiiiiiiii 0

E E O0 or)

.Q

(I)

E m

Od

..C m-

"0

iEi:i:i:i:i:i:i: C Qm

C mmm "0 0 od

(.1 m > "0

<[ (_1 |

r_

0 0 •_t

0 0 Od

C) 0 C)

0 0 CO

0 0 CO

C c- .._, OOZE

o_'_ .o 0

o

0 C) "d"

0 C) Od

0

oo oo

D B

o o o

8 D

°ooO

0

n oO_

D

c0 0 O O

D

a. ttG m

0 0 0

C "'

v-

C_ |

o 0

o 0 'I 0 0

D_ I 0

0

0

0

_A

_s.CD 2 _ c-

85

0 0 0 0

F1

0 0 0 0 m

eO'J Q

0 0 0 _0 CO

[1

,m

>

o o I-

o_

n

c_ -

2

o_ o

D

Od

erIAI "O e-

O O O LO

>

c_ |

O O O O

e_

8 IA

O O

n

I_.

I O O LO OJ

I

I

I

IIII O O O O,I

I O O CO

I

I

I

I

I

O O O

I

I

I I I I 0 0

-_

86

O

Engine

Half Sections

Engine

Mounting

Plane

I I

1 I

i

i

_

,

i,

J

59"

7.5K--_-

10'

106"

I 167

I

115'

18.3'

20K

._3_

i

Overall Engine Length

I m

I

21 .7

°

25K--

r

136"

Aerobrake

L

_

!i

35K-

L

50K,

25.4'

I

20K

I I

;

I

157"

25K

I!

35K

._t_

50K IbI Thrust

I

Figure 3.1-31.

Change

in Engine Length With Thrust 87

Engine Length Wlth Nozzle Stowed For

--

TABLE ADVANCED ENGINE NORMALIZED

Engine

Thrust,

lbs

Throat

Radius,

inches,

Parameter

DESIGN STUDY ENGINE CONTOUR BY THE THROAT RADIUS

7.5K

20K

25K

35K

50K

0.765

1.25

1.395

1.65

1.97

Normalized

(non-dimensional)

Chamber Barrel

rt

Length, Section

Chamber

Inner

Radiused

9.60

9.32

9.09

8.12

Lc

6.81

6.58

6.34

6.14

5.30

at

6.86

4.80

4.66

4.55

4.06

Barrel

2.0

2.0

2.0

2.0

2.0

Con-

2.0

2.0

2.0

2.0

2.0

2.0

2.0

2.0

2.0

2.0

141.318

141.318

141.318

141.318

141.318

34.641

34.641

34.641

34.641

34.641

Radius rc

Transition,

to Converging, Radiused

ri

Transition,

verging Radiused Throat

to Throat,

ru

Transition, to Nozzle,

Nozzle

Length,

Nozzle

Exit Radius,

rd

Ln re

Angular Chamber

Barrel

•Converging Nozzle

Angle, Nozzle

Dimensions

13.73

L'

Length,

the Injector,

Iintial

3.1-19

to Section,

Divergence

Relations

30 °

30 °

30 °

30 °

30 °

30

30 °

30 °

30 °

30 °

4.9 °

4.9 °

4.9 °

4.9 °

4.9 °

0i

On

Exit,

0e

TABLE ADVANCED NORMALIZED

Nozzle

Length/Throat

Station

L/rt

3.1-20

ENGINE DESIGN STUDY NOZZLE CONTOUR*

Radius

Nozzle

Radius/Throat r/rt

1

.75

1.2842

2

2.50

2.3357

3

5.00

3.8379

4

7.50

5.340

5

8.459

5.916

6

10.344

7.006

7

12.100

7.949

8

15.727

9.727

9

19.152

11.243

10

23.758

13.089

11

30.096

15.349

12

35.370

17.036

13

41.893

18.927

14

56.310

22.513

15

65.471

24.458

16

78.448

26.865

17

103.304

30.589

18

123.367

32.936

19

141.318

Engines contour.

of any

thrust

in the 7.5K

L=Ln to 50K lbf thrust

89 RPT/D0417.SSa-T

Radius

34.641 range

have

the

same

r=r e

nozzle

a_ (D

q,llm m

"!-

I-C_ | ,i,E

!---

C_ 1,1. em

90

TABLE ADVANCED BASIC

3.1-21

ENGINE DESIGN STUDY ENGINE DIMENSIONS

Thrust, Parameter Throat

Radius,

Throat

Area,

Chamber

rt

inch 2

Diameter,

Contraction Baffle

inches,

Ratio

inches Geom.

Cross-Section

Area,

inch 2

lbf

7.5K

20K

25K

35K

50K

1.53

2.50

2.79

3.30

3.94

1.84

4.89

6.12

8.54

12.18

7.64

9.80

10.92

12.89

15.40

24.9

15.3

15.3

15.3

15.3

14

26

32

45

64

17.3

10

10

10

10

Contraction Ratio Baffle Area

Less

Chamber

L', inches

10.5

12

13

15

16

inch 2

2208

5868

7344

10,248

14,616

53.

86.44

96.70

114.73

136.42

1395

1200"

1200"

1200"

1200"

Length,

Nozzle

Exit Area,

Nozzle

Exit Diameter,

Nozzle

Area

*Assumes

ILPTID0117.E6a.T

inches

Ratio

e = 1200

is fixed

91

3.2

ENGINE

REQUIREMENT

The first of the two extended

throttling

missions land

within

a vehicle

a function

range.

Project

selected,

although

Pathfinder,

there

Phobos,

range

was

of the

effects

reason

for this investigation

as a propellant.

of high

For

20,000

for a throttling

main

mixture

ratio

of the

No

study,

were

throttle

a high

engine needed

to is

and

the thrust

a four engine rate

was

set.

given

planning.

requirements

was

an investi-

on the engine

oxygen

for

of the

used

with

for mission

of using

range

requirements

(MR > 7) operation

is the possibility

engine

g-load

engine

at an

capability

lbf/engine

number

the baseline

of the

engine

3.2-1.

basic

mechanical

evaluated

platinum

ratio

The

study

line NASA-Z

mined

design.

from

MR operating

point

Lunar

The rocks

of MR = 12 + 1

Design

3.2.1.1

using

ratio

were

Component

ization.

worst

case

7 (oxygen

calculated

TPA)

based

See Section

a single

tll-r/I)O4iT.tS_/31-3.8

design

used

for the

of the engine

alloys

for the baffle

variation

is unchanged plates

study but

as well

as the

the

is high

base-

Design

design

design

on flow 3.1.1 point

This

design

psia

at these

for the

and

mixture

conditions

thrust

performance curve

for the

at a chamber

10:1 throttling

at a nominal

performance

points

conditions

or 2000

A normalized 3.2.1-1.

parameters

for 20:1 Throttling

The turbopump selected

design

material.

3.2.1

Figure

mission

to look

on engine

throttle

10:1 on the baseline.

A summary Table

mixture

case

was

was

investigated.

given

ture

The exact

the Apollo

from

this portion

emphasis

is a requirement

as an important

variation

baseline

in the set and the g-load

20:1 versus

gation

program

coordination,

this is recognized

STUDIES

the engine

or Mars.

of engines

considerable

The second

was

the new

For this subtask

after

from

With

of the number

The throttle

variations

on the moon,

vehicle/maneuver.

VARIATION

of 20,000

lbf was

and

tank

pressur-

calculated.

is given

stage

mix-

Flowrates

For the 20:1 throttling

for this TPA to a single

were

psia

TPA).

5% for autogenous baseline.

is applicable

engine

of 2300

5 (hydrogen

engine

chart

92

pressure

ratio plus

10:1 throttling

as

for either

TPA.

TABLE STUDY

Parameter, Engine

BASELINE

Thrust,

Mixture

Ratio

Chamber

20:1 Throttling

lbf (Nominal)

Impulse,

lbf-sec/lbm

Range

Pressure,

Pc, psia, Wall

(Nominal)

Temp.,

Thrust

Plate Maximum Wall Temp., °F Copper (NASA-Z Alloy) Platinum (pure or 10% Rhodium)

Oxygen TPA (Maximum)

Turbine

Chamber/Baffle Turbine

Bypass

Regenerator

Hydrogen

How

Minimum

Flow,

Bypass

LOX/GH2

HEX

Idle Valve

Inlet Temp.,

Minimum

Bypass Range,

Split,

Minimum

Oxygen Cooled (for Property

Nozzle Data)

Material

Regen

Inlet

%

Area

Ratio,

Varies

TBD

1050

1050

1050 2000

1050 2000

400

400

30 to 70

30 to 70

10

10

25

25

10

10

NASA-Z

NASA-Z

Nozzle

Inlet Area

Oxygen

Cooled

Nozzle

Exit Area

Diameter,

Ratio, Ratio,

¢

inches

93

Ratio

7 to 13

2000

0tol0

¢

Mixture

484

Copper

Cooled

Rr,r/Do_lT.ss.-T

Flow,

High

TBD

Copper

Oxygen

Internal

%

% of H2 Flow Material

Chamber

%

% Flow,

Regen Cooled Chamber (for Property Data)

Chamber

°F

VARIATION

20,000

5to7

Maximum Gas Side Chamber, °F

Fuel

REQUIREMENTS

dimension

Specific

Baffle • •

- ENGINE

3.2-1

0 to 10 NASA-Z Gold

Copper, Plated

NASA-Z Gold

Copper, Plated

28:1

28:1

35:1

35:1

600:1

600:1

9.8

9.8

TABLE STUDY

Parameter,

Contraction

Radiation

Nozzle • •

Regen

Channel

- ENGINE REQUIREMENTS (CONTINUED)

dimension

Chamber

AE/At

BASELINE

Cooled

VARIATION

20:1 Throttling Ratio,

Nozzle

Ainj/At

Area

Ratio,

High

Mixture

15.3:1

15.3:1

1200:1

1200:1

Ratio

(_) Contour 27 ° Angle

To E = 35 To _ = 1200

Channel

Land

Geometry

Width

in Throat,

27 ° Angle

Rao, Optimum

Rao, Optimum

Bell

Bell

.011" x .083" Depth

in Throat

.010

inches

94 RPT/D0417-55a-T

3.2-1

.010

i

il

.................................................... _ ........L......... r.......+,.......... _ I _ I " I " : 'z " ... _

__

! ........

JJJ

_

i,,

_ ........

•

......

t

r

Itl_L:,_ I<-

#"

•

J

| I...- _

-i, .... _f .... _I

I "I"

q

" i i _

i

: !

i ......i :, I

I

I

°

11

I "

O

4

i

.......

#

I_'_ o _1

I

l_i

li_i

I i_<>

I

! !-'1"----.-I I P_

!

-"_I

I

I

1 I _I !.=--._L_ _

i-tli

r_._/'1.,,.. / %1

/,

_ii

I O

O

O

.....

i_

: :

-.

o

ii ,

~o "-

'r"

-_ 4i

: .,

1_o

i

I: f

g

!

_.

i

I

_P',,,! _' I<_

If _i_ _l--i--i_ __ _,' _ ' I I!

o

',4,--

:

!/_ o:

e.--.-../_.,._ t

ua

_

\.'_i,,, %11

_> o_

" _

O

_PH/H

I

d O3 =3 iI

u.

_95

3.2, Engine

Requirement

Variation

Note from

Q/Q

design

considered

that

the speed

= 0 throughout

necessary

engine

throttle

design

falls on the

where

head

is also

low

plotted

Y-Axis.

pump

against

chamber

performance

the top end

selected

design

capability

suction

and stability.

The engine

operating

With frequency

combustion

stability

over

discussion

stability

a range

and

a wide

liner

and baffle

and baffles,

treated ature

at 1700°F

plot for the Narloy-Z

0.2% yield reversion desired

strength

curve

to annealed normal

to exceed

900°F.

hydrogen

flow

per baffles, flow

path

is needed required

the and

With

to the chamber recourse transient

at nominal

for the nominal

added

With

at 900°F

capability.

speed

controllability

3.2.1-3.

to the Rockwell

at 1050°F

that will

limit

and

plates

letting

baffle

reduces

the

3.2.1-5

shows

Figure

condition

C includes

There

a

on the chamber

on material versus

The design

set a life limit

down

this is readily

time.

It also

based

A stress

chart.

shown

is dependent

is 1050°F

low

heat

temper-

allowable

will be a gradual for the chamber.

The

is 900°F.

at the 20:1 throttle baffle

limit

injector

have

Appendix

3.2.1-4.

The

a 5% overthrust

in Figure

for 4 hours.

stable

pressure.

NASA-Z/Narloy-Z

as Figure

the

thrust

give

for low

throttle

temperature

is to reduce

thrust.,

only

of the 20:1 in this study.

temperature

platinum

the overthrust

injector

at equilibrium.

The chamber

chamber

of the I-Triplet

is given

3.2.1-2

to assure

Tests

material

operating

selected

or "chug".

aging

TPA

the TPA

to consider

with

for this

Figure

it is important

at the low

properties

was

was

is depicted

operation

was

from

range

operating

predicted

line

throttling

temperatures

for 2 hours

stall

to a 100 psia

is traded

stability.

the maximum

point

will

for "chug"

Extended chamber

3.2.1-1

envelope

was The

The

slope

speed.

in particular,

representative

test reference

instabilities.

can be inferred

corresponding

performance

a negative

characteristic

for 20:1 throttling

in Table

band

This

have

for reference.

The design

speeds

design)

other

specific

in designing

given

at MR = 7. Broad

and

performance

performance,

points

stall

pressures.

at the low

This changed TPA

speed

(N/N

range.

on this figure

The concern at low

ratio lines

Cavitation

loss is plotted

(cont)

the operating

to avoid

line

performance

Studies,

chamber

for each

point

enthalpy

be designed

accomplished

temperatures length.

should

reach This

the hydrogen of the three

by biasing

2000°F.

shortens

pickup

With

the cop-

the hydrogen

of the hydrogen circuit

active

not

enthalpy

components:

that pickup

,,Q.-

I

I

]

Ee_

,

!

i

.

;

:_ o

....

t

;

:

_

_ _

_.-,',- N--

_'-

_

•r-

L

,

• r_

_h. 13 _,

"0

'-_

, _

OO

E_ _, X

:_ ,-O

"_ En E:

O Z

_:_

- -_

>

(1,1 I:_. 0 or),-

o .......

J ___.

_--

0

i

.....

0

C_

Z

"'-" (..)

,.

¢-

.E

_D E:

o

.... o

[.___L._I.... '

.............

' "b'-:.'

Z

=

:'-_._.-."_,.

i

O0 "," 0

........

(D 0

_o

_.

o

_P

.....

{J') t_J

___

rm

....

(J o

- _

t

-!

o .._

--

f T

•

........ _ -_!_

7 """_'--q: .... ..............

t ............ 1----'---_, ......

:

r - "--.. t"-'1'......._

.....

"o

_o .... _ .......

c_.

"r" "0

E ffl

I

o

N ----|

.................... t .............................. -t-4 ---

N

U

I

-t

_

-J

t

..........

P.

,

zV1 0 I--

0 ,

......

•

,

i

_

_

_L ........................

L

;

,

'

i

,

:

o

I

.....

---,-----

J e,j ,

i

,

,

:

.

,

t

{ i

!

•

_

._

0

_

.......

_

M')

1

.........

:

1N33_13d

........ -

NOIIVIIAV9

i

i

i

i

i

'

__[

_

o

,

i

'

i,-_ ......... 01

1o

.[

3NG S507

-+t

,

,

aV3H i

97

ORIGINAL

P_.QE

OF POOR QUALI_

TABLE ADVANCED Engine

EXPANDER

ENGINE

3.2.1-1

TURBOPUMP Value

Conditions

Rated

Thrust,

Throttle

F, lbf

Propellant

Inlet

Temp.,

Propellant

Inlet

Pressure,

Speed

°R

Boost

Conditions

Pumps**

38

163

20

15

5O

5O

Density

Dimensions

Fuel

lb/ft

4.42

Shaft

Speed

Total

Discharge

Total

Suction

Pressure

psia

Total

Pressure

Rise

psi

Total

Head

Weight

LOX

psia

To TPA's

Propellant

LH2

2,300*

Propellant

Pump

Oxidizer

2,000

Pc, psia

Pc, psia

To Low

Fuel

20:1

Pressure,

Overthrust

SPECIFICATION

20,000

Range

Chamber

DESIGN

Pressure

Rise

(cavitating)

Flow

Capacity Specific Speed Cavitating

(Based Head)

on

3

Oxidizer 71.2

rpm

150,000

55,230

psia

4,650

4,650

50

5O

4,600

4,600

138,231

9,313

lb/sec

5.903

35.42

gpm

599.4

223.3

1,448

1,464

ft

rpm

x _zDm 1/2 ftf72

Efficiency

%

Fluid

Horsepower

h.p.

1,483

600

Shaft

Horsepower

h.p.

2,279

881

ft

1,792

111

Net

Positive

Suction

INLET

Specific

DIA.

DISCHARGE Q/N AH/N

Suction

2

DIA.

Speed

Head

rpm

65

x _zc)m I/2 ft 172

68

13,334

24,133

in.

TBA

TBA

in.

2.44

2.44

3.993

x 10 -3

4.043

x 10 -3

6.144

x 10 -6

3.053

x 10 -6

TABLE ADVANCED

Turbine

EXPANDER

3.2.1-1

ENGINE TURBOPUMP (CONTINUED)

Conditions

DESIGN

Dimension

Fuel

Gas Power

Gas

Mass

Gas

Inlet

Pressure

h.p.

Flow Total

lbm/sec

Back

Shaft

Speed

Pressure

Inlet

31.9

770

400

1.67

2.1

% Total Area

Specific

Heat

Specific

Heat

Pressure (effective)

55,230 80

psia

4,109

4,038

in.2

0.324

0.414

Ratio ft/°R

(1)

(2)

(1)

(2)

(1)

(2)

in.

Mean

1,920

80

BTU/lb°R

Constant

REFERENCE

5.31

150,000

rpm

Nozzle

Diameter,

908

2,460

psia

Efficiency

GO 2

2,349

oF

Temperature

Ratio

Static

Gas

Oxidizer

GH 2

Shaft

Gas

SPECIFICATION

2.77

(1) Hydrogen

--

NBS Technical

(2)Oxygen

--

NBS Technical

Note Note

2.42

617

April

1972

384

July1971

The design point was based on a 15% overthrust rating. The power balance confirm this as a viable operating point. The TPAs are slightly overcapacity this design point.

Refer boost

Rr, r]_;7-ss.-+

to the engine schematic in Figure pumps to the high speed TPAs.

3.1-1

99

for the

relationship

of the low

did not using

speed

!_

i

i

_i_®_

: rj

:

:

i

i

i

........... ]

?

i

i,._,-._

'

x---__'

:

:

:

'

:

-- _

_

_

i

Z

i

........... i.°, _, i............. !............. i............. i-7o--y ....... _ ............. _............ "... ............ _............ "... ............ o

............ !o,._.i ............. i............. _.......... ..----I..-i- -.'-..d .................... ............ .......... ....... 8,

.= ............ _............ _............ _......... _.......... ,, ............ i ........ .....I............ I............ ............ l..x......i ............. ! ................

............. i.......................... !............. i.......................... i....... ......L. ......... 1........................ l........................................................ _

i 0

_=

,m

!i

i

eni-iii

o_1

i

_-

............. i.......... I i.......E i............. i .......

!

"_ ;

"-

_, ..........._ ......_

_o

............. '..........._........ ' ............ •....... , ............ I ............ 1 .....,®...... ..... -

i _.

i

i ._1 i

I

m_

"

e= I,i,I

I _,,

"

I __ _

- _

o° JE

............ i............. i.....x...,...i ............. i....... .'....I ............ !............ ._ ............ l...... _o...... __

r= i,= UJ

i i \i i : _ z _,, ._............ :: ............ i............ _.......... _......................... _............................................ _ o..., m i=*

I T,=,

............ i............ i............ i............ i............. i.-X.....i ............................................... _........... i............ i............. i _.

i._

.................................................... !............. i............. i........ '_........... ,............... X....... t; ...... !............ i............ i............ i............ " ............ +........... i............ !............ i............ !......... !.........x.. ...., ............ '_t _! ............ i............. i............. i............. i iii \,. _,'_i il !

!

i

":

!

i

i

i.

i.

.i, ,i,

i

i

:

i

=

!

:

i

i

i

i

i

i

_

i

\

_]=

,__

aHNV Z-,OM'_eEl oJn|x!_ Ii

•

tO0

i

" :,

i

i

i

i

i

_---_._ ....... I.=..-.._ ............ i............ i.............

_ _'_

:

E

• _

i _

_ _ _

o

.__ U=

NARIoy-Z

1700UF-2

h-WQ,

WROUGHT"

900°F-4

h

7002.26.

ROCKWELL ¢_b

INTERNAT

IONAL DiVISlCN

MATERIALS PROPERTIES MANUAL

I O. 70-01

t'V_qLOYZ HT

17001:

2

I_S

9001"

4 _

_ _D61LE S_ni

,2B

WO.o_

.IP

SIA Q_-

9-t-_ PAOE I'_.kIBER [DITION P_L'¢I- "/.B.2. !._,]_

tO0 T_n"T"_Tl"n'nrm'_'rrrrI"

n lTn'Wl'e'n_n'rl TEMPERATURE, '; .... I""1 .... I'n'_'_FrmPnYIrn'le"lw'rr'lnWI__'_r'1"_nr'l./O0] C

........ I............................. _o ........................ I....I....................... '.............J _

' "dl

I

\

I

.......

tO

l EMPERATURE,

Figure 3.2.1-4. 101

F

7

7OOO E I-

'_

6000 ,50(_

d 09

4OOO 0 c-

3000 t-

'"

2£E_ 1000 0 7.5

20

25

35

50

Thrust Level, Ibf

Regenerator

Jacket -,:_.-_ Baffles

Figure 3.2.1-5.

Hydrogen

102

Circuit Enthalpy

Pickup

3.2, Engine

Requirement

hydrogen

regenerator,

the 20K lbf thrust input

Variation

regeneratively

level

more

energy

require

increase

in its size

and

weight.

with

pound

3.2.1-6

The

the balance

important.

extraction

cooled

Reducing

baffles.

At

the enthalpy

at the regenerator

This can be one of the penalties

pounds

per pound

and regenerator extracted

lower

than

the incoming

has lower

operation,

all components

with

a con-

for a wide

the engine

actual

throttle

Power

sec/lbm. 20,000

at this point

This represents lbf thrust.

boundary

throttle

down

come

losses.

cooled

Maximum

temperature

by the power

the baffle.

_r,t_lT_s.,3_.3_

interplay

in size

if more

energy

"oversized".

During

The flow

by the

around

in HEX is

HEX.

throttle

regenerator

The

down

and HEX

these

per

is six

is some

transfer.

was

run at a chamber

is 20.67:1

and predicted

Results

a small

The thrust

decrease

chamber

pressure

are given

of 30.6 lbf-sec/lbm

from

chamber

but they

with

balance)

The proportioner

temperatures,

for heat

increase

flowrate

turbine

in the

bypass

components.

specific from

of 100 psia

in Table impulse

the 484.3

in mixing

is no longer

to eval-

3.2.2-1.

The

is 453.7

lbf

lbf-sec/lbm

efficiency optimum

at

but mainly size

at the

conditions.

out of the regen

(assumed

enthalpy

There

the turbopump

20:1.

Temperatures

wall

increase

throttled

a decline

Losses layer

from

The

is

at 20:1 Throttling

balance

when

nozzle.

circuit

increase

but the oxygen

of the hot hydrogen

Balance

operation ratio

must

are effectively most

cooled

of its enthalpy

conditions.

temperatures

A power uate

at nominal

plot for the oxygen

2/3

for the hydrogen,

hydrogen

delta

will be bypassing 3.2.1.2

up about

The regenerator

regenerator

valves

picks

pickup

in the oxygen

of hydrogen

sizing.

from

the enthalpy

oxygen

acquired

is substantially

from

is very

and hydrogen

range.

in Figure

HEX

chamber,

input

For reference, given

cooled jacket

would

throttle

(cont)

the regen

in the jacket

sequent

Studies,

valve

are within

the design

is at 587°F a 50/50

split

are still within

and out of the baffles hydrogen

is 1030°F

limits.

split

for the regen

should design

103

from cooled

be changed limits.

Note

that the hydrogen

it is at 653.7°F. the proportioner chamber

to better

valve

and 810°F

optimize

these

for

Oxygen

E

Circuit

2OO

D m

d

oo

150

o e-

¢L

IO0

e" W

5O

0 75

20

25

35

Thrust Level, Ibf

Hex

Ox Nozzle

Figure 3.2.1-6.

Enthalpy

104

Pickup

5O

Table 3.2.2-1 Engine Power Balance At 20:1 Throttle Down Condition ÷ .............

÷ ...............................

÷ ...............................

|

OTV

+ .............

÷ ...............................

(

I Off-Design

ENGINE

POWER

÷ ............................... I

4. .............

4. ...............................

[

+

I

R

Fuel

4. ............................... -

670.70

(In/deC

R)

I

Pout

-

15.00

Tank

I

Tout

-

182.70

(dec

I

Condltlona

I

Hour

-

-67.17

(BTU/t)

I

_

Rho

4. .............

4. ...............................

8hut-off Valve

out

-

(pule)

71.17

R)

Heat

Fuel: Side

Regenerator

20.00

8

37.80

Hour

out

(pule) (dec

-117.36

•

R)

(BTUt#)

4.34

(t/©utt)

(pile)

Pin

-

20.00

(pals)

(p$1a)

Pout

-

20.00

(pale)

-

0.00

Delta

(pml)

T

-

162.70

(dec

H

-

-87.17

(BTUIt)

R)

P

-

H

-

In

-

71.17

(t/cult)

Rho

in

Rho

out

-

71.17

(t/cult)

Rho

out

-

8.03000

(In^2)

•

T

Rho

m

CdA

0.00 37.80 -117.35

(psi) (dec

R)

(BTU/I)

4.34

(tloutt)

4.34

(#/Cult)

•

10.00000

(in^2)

Pin

,,

21

(psia)

158.66

4" ............................... (pale)

+

Pin

,.

le.00

Pout

-

107.0S

Pout

-

Tin

-

le2.70

(dec

R)

Tin

•

37.83

(dec

R)

Tout

-

164.37

(dec

R)

Tout

=

43.44

(deg

R)

(pall|

Rho

in

-

71.1705

(Ib/ft3)

Rho

out

-

70.9094

-

1.919

Elf

(hyd)

-

0.360

Eft

(meh)

-

1.000

N

-

7361.48

HP

-

2.34

-

P

in

-

4.3340

(Ib/fl3)

(Ib/ft3)

Rho

out

•

4.1748

(Ib/tl3)

(ibluac)

Wdol

=

0.320 O.301

(rpm)

(gpm/rpm)

-

137.85

(pule)

-

123.20

|pale)

-

14.45

(hyd)

-

Eft

|mch)

(Ib/sec)

•

1.000

N

•

17529.54

HP

= -

(rl_n)

8,07 .00leeS0

(HP) (gp_Irpm)

Hour

-

-Sd.SO .5e.30

Rho

in

-

71.00

Rho

out

•

71.00

: -

158.27

(pale)

-

0.42

(psi) R)

(dec

R)

Tout

-

67.30

(dec

R)

(dog

R)

Delta

-

23.85

(dec

R)

-

-97.38

-

122.O2

184.87

-

-

(dec

-

HIn

(pale)

43.44

Tout

0.00

P

16e.88

-

(do

-

Pout Delta

-

T|n

104.87

T

4.

Pin

R)

-

Wdot

Eft

O/N

(psi)

Tin

Odor

(pale)

4. ...............................

Pout

Delta

.00

Rho

(HP)

.O010438

Pin

Delta

Cool

-

Tout

15.00

P

9

(BTU/t)

HIn

(gTU/t)

Hour

(#l©uft) (tlcuft)

(BTUI#;

-

4.17

(#/cult)

Rho

out

•

0.88

(t/cult)

Qdot

1.92

(e/sac)

Wdot %

bypaes

4. ..............................

105

(BTU/t)

In

(STU/I)

4. ...............................

T

Rho

0.00

] 4.

15.00

4. ...............................

Exchanger

R)

-

QIN

Cool

(In/dog

-

Wdot

4" .............

0202.00

Pout

Rho

(e/cult)

÷ ...............................

Pump

4. -

Pout

CdA

Condltlona

R

Pin

Delta

4" .............

|

4. ...............................

(

..............

1

Run

I

Side

Page

÷ ............................... Oxidizer

+ .............

OX:

÷

BALANCE

• -

28.41

(BTU/I)

0.12

(#leec)

25.oo ÷

Table 3.2.2-1 Engine Power Balance At 20:1 Throttle Down Condition ÷ .............

÷ ...............................

÷ ...............................

I

OTV

÷ .............

ENGINE

POWER

BALANCE

÷ ............................... I

..............

(Cont.)

Peg#

2

÷ ............................... Oxidizer

Fuel

÷ ................................................................ Pin

-

166.68

(pile)

Regen

Pout

-

143.92

(pile)

Jackal

Delta

P

12.76

Tin

-

Tout

= -

Delta

T

HIn

-

Hout

-

Rho

In

-

Rho

out

-

Wdot ÷ .............

÷ ...............................

R)

1000.00

(deg

R)

-97.30 3302.64 4.17

(BTU/#) (BTUI#) (#/cult)

0.03

(#/ouft) 0.16

(#/seo)

-

166.27

(pile)

Pout

=

148.74

(pile)

P

-

Tin

=

Tout

-

T

HIn

-

Hout

-

Rho

in

-

Rho

out

-

Wdot + ...............................

+ ...............................

Pin Pout P

-

123.20

(pile)

-

123.12

(pale)

Nozzle

Della

Cooling

Tin

-

164.37

(deg

R)

Tout

-

540.70

(deg

-

R)

-

376.33

(deg

R)

-

-56.30

(BTU/@)

T

Hlfl Hour

-

Rho

in

-

Rho

out

=

Wdot

(deg

Pin

Delta

Delta

R)

1043.44

-

Delta

Ox

(p|l) (deg

÷ ...............................

Baffles

+ .............

43.44

0.08

116.87 71.00

(BTU/#) (e/cult)

0.06 .

(psi)

(#;¢uf_) 1.92

(#/eec)

÷ .............

÷ ...............................

÷ ................................

4. .............

4. ...............................

÷ ...............................

106

7.53 57.31

R)

1113.31

(deg

R)

1056.00

(deg

R)

67.74 3805.88 0.99

(BTU/#) (BTUI#) (#/tuft)

0.03 -

(pll) (deg

(#/¢uft) 0,16

(#/see) +

Table 3.2.2-1 Engine Power Balance At 20:1 Throttle Down Condition 4. .............

4. ...............................

I 4. .............

4. ................................ OTV

ENGINE

POWER

4. ...............................

I

I

4. .............

4. .................................................................

Page

3

Fuel

Pin

=

123.11

(psle)

Pin

-

143.92

Pout

-

111.82

(pale)

Pout

-

120.31

Wdot

=

0.885

Wdot

=

0.129

Tin

-

840.10

(dog

R)

Tin

-

1078.33

(dog

R)

Tout

-

532.ge

(dog

R)

Tout

-

1084.34

(dog

R)

Hin

-

110.67

(BTU/e)

H|n

-

3684.19

(BTU/@)

•

115.00

(BTU/#)

Hour

-

3835.02

Hour Rho

in

Rho

out

= -

(e/cult)

Rho

in

0.83

(#/ouft)

Rho

out

=

HP

=

PR

-

1.100

-

0.206

-

2,542

-

46.64

U/Co dla

%bypass

(Ib/seo)

0,08

Eft.

Wheel 4. .............

BALANCE 4. ................................

Oxidizer

Turbine Conditions

(Cont.)

0.032 2.33

(HP) (Pin/Pout) (fps/fpm)

=

(BTU/#)

0.02

(t/cult) (#/cufl)

-

0.02

=

0.400

HP

-

PR

-

1.122

(Pin/Pout)

-

0.090

(fps/fps)

-

3.090

:

59.74

Wheel

die

% bypaIc 4. ................................

4. ...............................

(pale) (Ib/sec)

Eft.

U/Co (in)

(pale)

8.95

(HP)

(in)

Pin

-

128

31

(pale)

Hot

Pout

-

127

31

(pale)

Side

Delta

Heel

Tin

-

1072

73

(dog

R)

Exchanger

Tout

=

1072

79

(dog

R)

O0

(dig

R)

45

(BTUI#)

P

Delta

-

T

=

HIn Houl Rho

In

Rho

out

4. ...............................

-

3684

-

3364.45

=

-

bypass

-

O0

(pit)

(BTU/Ol (#/cult)

0.02

-

Wdot .............

0

0.02

Qdot

%

I

(#/cuft) 0.00

(BTU/s)

0.23

(#/sac)

25.00

...............................

4.

Pin

-

127.31

(p$ie)

125.51

(pale)

Gee

Pout

Side

Delta

Regenerator

Tin

-

1072.79

(dog

R)

Tout

-

1047.38

(dog

R)

(dog

R)

P

Delta

-

T

-

HIR Hout

Injector

-

3664.45

-

3876.42

(BTUI# (BTU/@

tn

-

0.02

(_/cuft

Rho

out

-

0,02

(e/cult

-

-20.41

Wdot 4. ............................... Pin

24.81

(p_l]

Rho

Odor + .............

1.30

-

0.30

(BTU/s (e/sac

+ ...............................

+

-

111.82

(pale)

Pin

-

Pout

-

100.07

{pale)

Pout

-

Tin

-

338.74

Tin

-

(dog

R)

125.51 09.37 1047.83

(psla)

(dog

in

=

0.02

(#/cuft)

Rho

in

-

0.02

(e/cult)

Rho

out

-

0.53

(#/gull)

Rho

out

-

0.02

(#/cuft_

(Ib/sec)

Wdct

-

0,305

Drop

-

25.64

(psle)

CdA

-

0.679_3

(In^2)

•

1.827

Drop

-

11.88

CdA

-

1.08918

4. .............

4. ...............................

I

Combustion

_

PC

-

]

Chamber

I

DPcc

-

I

I

ERE

-

I

I

F

-

(pile) (In^2)

I

(Ib/sec)

;

4. ............................... 100.00 0.18

MR

-

3.00

(pale)

Wdol

-

2.13

Dthroel

-

Is0

-

(Ibf)

107

t 4.

(pale)

1.000 987.20

I R)

Rho

Wdol

I

(psla)

2.600 453.66

(O/F) (Ib/sec)

I I

(in)

I

(sac)

I

3.2, Engine

Requirement 3.2.1.3

Variation

only

a shift

hydrogen

and

performance the come

oxygen

effect

High

for the high

capable

The

3)

Maximum

4)

A thrust/MR

Oxygen

6) plates

limits

(1050°F

that

wall,

at high

engine.

speed

operation

at

10:1 throttling is reduced

There

MR need

seems

to 21,000

to be no

(oxygen

to keep

not be as great

side

through the gas

are outside

the nominal the

thermal

the

were: variation

is to be

as at nominal

MR.

changes.

is within

above

subtask

20:1 throttling

be below

violating

flowrates

ratio

component

could

flowrates

throat)

versus

overthrust

for the

combination

envelope

800°F

low

for unrestricted

mixture

defined

without

insufficient

high

major

thrust

Hydrogen

were

for the

ratio

operation

as usable

stable

Operation

without

Throttle

adequate

lbf thrust

design

2)

in the operating

baffle

rules

engine

to assure

of the

or life. Ratio

MR variation

5) considered

20,000

Mixture

3.1-1)

20:1 throttling

the control

for 20:1 throttling

capability

weight

1)

of continuous

penalties

nominal

ground

are

accommodates through

(Figure

margins

overthrust

The

readily points

valves

The

on engine

design

operating

Thermal

lbf for the

3.2.2

used

engine

pressure

condition. engine

23,000

significant

back

control.

down

in top end

lbf from

baseline

in the turbopump

and

throttle

(cont)

Conclusions

The with

Studies,

oxygen

for MR

operating

design TPA

= 6.

envelope

if it is

limits. design

point

were

not

limitation). either side

the

regen

cooled

wall

temperature

of the operating

envelope.

chamber below

(Hydrogen

or

design side

limitation).

7) outside gram

of the operating to evaluate

k_-r/,_,,7__s_/32-3 _

high

Any

high

envelope. MR control

MR region (There

where were

stability).

108

control insufficient

instability hours

is encountered available

is

in the pro-

3.2, Engine

Requirement

Variation

In essence, MR operation special that

must

the engine

plates,

and

envelope

these

be used

modification.

One

Thrust

platinum

Chamber

small

amounts

approaching oxygen 3000

stoichiometric

is present psia,

gases

at chemical

it to diffuse

copper

oxide.

be 2.7 volume

thermal

formed

precent

the copper

cycling

pressures

subsequent atomic oxide

and water

of 200,000

psia.

in the surface

Grain

where

or

the confirmation copper

The

oxidize a major

For instance,

baffle

operating

diameter

it reacts

with

hydrogen

that diffuses copper

issue

can generate

= 10, atomic pressure

of the oxygen

atom

a copper

atom

water

water

are enlarged,

with

and cracks

9 through

psia,

surface

vapor.

bubbles

=

to form

Pc = 3000

into the copper form

of

at any MR

(chamber

the MR = 7 and

and

(See References

in the presence

at MR

percent

where

boundaries

of the metal.

life becomes

operation

coalescence

was

3.2.1-3).

will rapidly

The small

to elemental

change

Ratios

at 1.6 volume

equilibrium). lattice

Figure

Mixture

or higher.

gases

design

to MR = 10 with

(See

(>600°F)

= 7.94)

into the metal

During

it will reduce

(MR

for nominal

engine.

Chamber

in the combustion

allows

will

of free oxygen.

up

plates.

at High

surface

any

used

of this analysis

operation

baffle

Life

the engine

without

results

for a rocket

A hot copper even

MR operation

includes

expanded

3.2.2.1

say that

gratifying

envelope

to MR = 13 with

(cont)

groundrules

for high of the

operating

is unusually

Studies,

there where

Later

internal

and blisters

11, and Figure

are

3.2.2-1

for

photomicrographs). Chamber

Blanching

The progressive leaves

visual

effects

to the whitish,

new-copper

interconnected running (Center

on a copper

to the

Line Average)

on the order properties degraded

of all common at such

due to crack See Figure

of 1988°F

coolant

flow

substrate

copper

of coolant

severe

areas

(NASA-Z,

Blanching channels

termed

surfaces

roughness

have

indicated

greater

than

OFI-IC,

ZrCu,

can progress and

described

subsequent

until

above

"blanching"

due

are commonly

interconnected

Surface

temperatures

alloys

been

Blanched

passages. Blanched

reaction

that have

appearance.

is common. with

wall

"wormholing,"

temperatures.

penetration 3.2.2-2.

chamber

penny

by subsurface

parallel

oxidation/reduction

porosity,

and

cracks

of 300 microinch surface 1700°F. etc.)

chamber

temperatures Structural

are

the thrust

CLA

severely chamber

burn

fails

through.

ij,,_Gtm_l_ BLACK

AND

WHITE

FAGIPHOTOGRAPh

Mag 2000X

%1_i ili_¸

i

Mag 2000X The TGA Specimen Exposed to a Reducing Environment Shows a Reasonably Smooth, Undulating Surface (Top), While an Oxidized Specimen Shows a Granular Surface (Bottom) Figure

3.2.2-1.

NARIoy-Z

Exposed

to Oxidizing/Reducing

110

Environments

, ....

o ,.-,w

pHOInGRAP_

I

_m

v

_u e,

c

IB

i

u c

I i

c

c c ii

i

|

I"1 u t,m

in

_k

Jr" 0 L_ I-

111

3.2, Engine

Requirement

Variation

The surface

roughness

contact

at many

Figures

3.2.2-3

nation

high

which

3.2.2-4

where temperature,

properties

over

several

relates

Refer

to rocket

engine

gen/hydrogen

objectives

ment/chamber impingement tailoring stream

pattern.

will

have

compatibility

Careful

injector

and

next

ratio

more

wall

concern

at the

due

Strain

of

in

The

and

increased

to the loss

are given

integrity

discussion

NAS

combi-

degradation

is an effect

of

of blanching,

of the phenomenon

Task

by tailoring

contact

as it

of the injector

hydrogen

or so of the chamber

is to prevent

means

bleed

of assuring

hydrogen of this element

the

oxygen

stream

openings walls,

occurring

rich

increases.

protect but

This

the

face

of the

element compatibility.

despite

some

element

asymmetry.

wall

temperature

tective

surface

becomes and

as material manhour nel tical,

Preventive

measures

coatings.

This

intensive. that may

Rr,r,_lTss,/3a-3_

The

reduces

SSME

will keep be ineffective

time

second the wall

to date

oxygen

Narloy

it is done.

and It also

wall

temperature

as coolant

channel

1) surface

concentration

porosity, requires

but

and

after

blanching

form

in application and

depends This

or injector

4) pro-

of peening

disassembly

reduction, 600°F.

2)

wall,

is limited

engine

at or below deposits

smoothing,

by a combination

temperature

112

at the

is refurbished

the surface

roughness

method,

include

Z chamber

smoothing

surface

each The

tried

3) reducing

by mechanically

is removed

cooling and

Prevention

reduction,

evident

grinding.

Blanching

is

of the injector

injector/chamber

from

increase. ele-

and

2 or 3 inches

baffle

oxy-

element/wall

of the gas

and

blanching

MRs

In short,

within

face

the

the effectiveness

MRs.

chamber

in the

ratios

ratio

At high

potential

design

the best

restricts

at lower

of the

as mixture

C.4 is to improve

ratio.

the oxidizing

imbalance

the momentum

flow

than

progressive

element

3-23772

in the portion

inch

is the

"I'-triplet

concern

solution

and

is apparent.

to cracking.

the momentum

while

is considered best

surface

structural

leads

hydrogen

increasing

the first

optimization The

chamber

Reduced

degrades

of particular

conductivity

in roughness

reduced

compatibility

by greatly

increase

11 for an in depth

of Contract

wall

thermal

to the porosity

chambers.

momentum

of the

due

of a blanched

cycles

to Reference

Another

face.

the

are

material

Photomicrographs

wall

not a cause.

(cont)

temperatures the

of increased

material

One

surface

decreases

points. and

Studies,

may

is

on channot

element

be prac-

%e,-_ = ,. _ _ _. I

BLACK

_

_-_ _ _;_

_

i -- I'r_, t_r'_ At'%") ?,,';-'J-f _: _-_,..3t ,.,,.-,_{APH

b

a

Tested at 11000F, MR = 10, 300 psi, O2/H 2 Cycle

a) b) c)

Figure

3.2.2-3.

Uncoated Cycle-26

NASA-Z Slot

Area

Cylinder (1.8%

113

After Strain)

Cu-Skin Indication of Blanching on Cu.Skin Oxide Layer Beneath Cu Skin

Oxidation

and Reduction

Test

'_ F DHQi.'-',7,-_:i, :--'. /_r',i,..m. - "",-',.r:,T.

St.ACK

a #

•

Tested at 1100OF, MR = 10, 300 psi, 02/H 2 Cycle

a) b) c)

Figure

3.2.2-4.

Uncoated Center

NASA-Z

Area

Cylinder

(>2.7%

114

Strain.)

Cu Skin Indication of Blanching on Cu Skin Cu Skin and Oxide Layer Underneath

After

Oxidation

and Test -

3.2, Engine

Requirement

asymmetries

lead

hydrogen

film

for all phases operating per

Variation

to higher

cooling

flexibility. was

mixture fully

ratio

range

mixed

molecular

oxygen

expected

to operate

nonoxidizing results

gas

and

by a gas tance The

nickel

and,

and

was

gold

on a 30 millionths

thick

formation,

typically

called

the

surface.

After

copper

Aerojet choice

team

has

operational

to baseline

the

that

to the

cop-

three

were

and

brush

but

coatings

nickel

nickel

engine

use

of a

The

aluminide,

on copper

checked

test articles

for oxidation

subjected

plated

but

cracks

were

had

tests.

suffered

some test

articles

observed

formed

technique

resis-

to bend

on the copper

pin holes

plating

The

An

are practical.

blanching

many

in a

be considered.

then

subsurface

an effective

present.

coatings:

from

oxygen

present.

must

tested,

surface was

oxygen

deposited

bomb

the engine

molecular

species

such

were

coating

was

nickel

strike.

reviewing

blanching

excellent Without

voids, the

due

discussion

in recommending of the

it for the engine

copper design

requirement.

RI'T/ I)O417 .5.S a/ 3.23.8

engine

MR = 6. At MR = 13

liner

tested

nickel

Kirkendall

no reservations

for preventing

protection

chamber

transverse

that

7

in the

in some

would

require

time. The

done

alloy

or spalling

many

concluded

development

the hot

the copper

no cracking

on sectioning, It was

with

concentration,

The

below

compromises

over

and

above

cope

aluminide

oxygen

test.

atomic

among

coatings

protected

in the bond

There

specimens.

The

ratios

coating

distribution

that

11) show

of nickel

This

non-oxidizing

vapor

investigators

types

process.

aluminide

tested.

some

Two

temperature

cracking

material

MCA

species

rapidly

to water

(Reference

require

by Aerojet.

increase

or a noncopper

would

at mixture

or tailoff).

a thin

shows

MR must

program

diffusion

at high

tensile and

gold.

only

at high

The nickel,

stream

is second

of the MCA

selected

method

or operation

is to apply

3.2.2-5)

third

lead

of the gas

(see Figure

coating

chamber

last option

evaluation

combustion

The

(no oxidizer

the method

An

(cont)

temperatures.

operation The

This

wall

for the complete

of engine

surfaces.

Studies,

115

in all respects the nickel to the gold

chamber. even

strike

if high

This

plating was

(Ref.

MR operation

into

11) the

as the method

is cheap

was

a void

to diffuse

report

plating

gold there

propensity

in the MCA the gold

when

enough is not

an

of

G) ::2 L_

Q.

E dUm'L IleM @

i

q)

I-

"0 C

g

n,"

ql'

Q.

_

o

o.

qJ

,-

q,,,

m

(e_d) _nmm_OI,m_ed

e,,0

Wo 'wm_t,mdt_.t.

J I

|

I

Qi

I /

/

/

I

|

_3 JQ

E 0 0 C 0

im

0

c_

.m L_

>.

&

0') U,.

116

3.2, Engine

Requirement 3.2.2.2

Variation

High

Mixture

For specific

Studies, Ratio

a constant

(cont)

Performance

propellant

mass

flow

rate

engine

thrust

will vary

with

impulse:

F = Isp m

Where

F = Thrust

Isp and:

= Specific

Impulse

M = Propellant

Flowrate

FI2 _ Isp @ 12 Isp @ 6

F6

Subscripts

but:

Refer

T = Absolute

to Mixture

Ratio

Temperature

of Combustion

Isp @ 12 o_ _/T12/Ag. T6/Ag.

Isp@6

The gen

flowrate

rigorous

3.2.2-6,

under

the

chamber

an engine

at any impulse which ular

engine

theoretical

that

given

there

MR and

is at or slightly confirms

weight

Rl'r/Do417._Sa,'3.2-3.8

the

contract.

simpler

charted and

400 seconds.

analysis

Figure

based

with

a family

This

represents

3.2.2-8 improvement

At MR = 12 and This

1 17

more

is given engine

in baselined

of curves

for

a throttling not

3.2.2-7

temperature

be assumed

that

for a 50K lbf thrust

for the Pc = 2000

range

is the performance

is the chart

is 83% of specific

on combustion

changes.

thrust

oxy-

and

ratio

it should

Figure

maximum

extensive

mixture

have

and

range.

performance

pressure.

above

psia.

predictions

engine,

is a slight

2000

A more

is for the 7.5K

All 3 charts and

the entire

chamber

3.2.2-6

_-__ 0.81

at MR = 12 and

thrust.

change

Figure

performance

over

operated

performance

100 psia

for a 20K lbf thrust Note

Engine

81% of rated

3.2.2-8.

between

can operate

prediction

about

and

OTV

pressures are

OTV

of engine

3.2.2-7,

current

These

engine.

produce

presentation

Figures

20:1.

should

baseline

Gas @ 612 Gas Mol Mol Wt. Wt @

higher psia

the

impulse and

thrust

cases

specific at MR -- 6 gas

molec-

of

o o _N

o OC o

._.c

E_ m

m

|

C ,D

_ C

r- 0s o_._ •£ o

(D

& 4) O)

,I I

I

!

I

O

I

O

LL.

I

O

Wql/Oes-Jql 'd$1

118

o o

0

2

o

t-- c 0 u rr

2

(¢) _B r-

m

I

c iI

_ ik_

e- w o=

|

i,J

:3 i

I

I

I

Im

14=

WqllOes-lql

119

'dsl

co

mqllOaS-lql

120

'dsl

3.2, Engine

Requirement

Variation

These release

efficiency

triplet

(ERE)

element

inches.

Pc = 2000

E = 1200)

and 485.2

2000

to 100 psia

high

mixture

by using

and

ratio

oxygen

from

maximum point

oxygen

on the assumption the

design

from

thrust

engine

includes

reaches

LOX TPA

MR and higher

shows

however. does

become

oxygen

capacity

rated

capacity.

thrust

an overthrust

2000

pressures:

psia

about

1550

Results hydrogen NASA-Z

was

The

2000

regenerator or Narloy

x,_/,_lT_,/32-3.s

psia

analysis preheating

baffle

curve

wall

used

overthrust).

The design The

curve

in

thrust

versus

stoichiometric).

does

balance

thrust

maximum

flowrate

plots

MR = 8 (near

is that engine

lbf

at MR = 9 as it

3.2.2-9

cycle

psia

This was 20,000

lower

flattens

a

The design

For a nominal

The engine

could

loss

engine

At

at MR = 13,

at high

thrust

MR

is attained

at

lbf is reached.

cooled

1500

also

thrust

for MR = 7, Pc = 2300

in Figure

above

chamber

performance. psia.

Later

for MR = 12 operation

analysis

to

realized

the full MR range

(15%

This curve

at slightly

thermal

and

point

engine

tank pressurization.

of 21,200

psia

not attainable

psia.

of the

2000

from

loss in Isp by going

parameters.

design.

MR > 10. Actual

regen

to be

down

by the economies

of oxygen.

of this analysis

until

is expected

in throttling

condition

preliminary

falls off rapidly.

result

runs

operating

Ibm/second

peak

(Isp)

The

be justified

set at the flow

The upper

thrust

at MR = 12 for their

chamber

impulse

The Isp loss

to the injector.

The baseline extensively

specific

as one of the limiting

a realistic

flow

a concern

MR = 8.3 where

= 6,

over

An interesting

not

(MR

can only

5% for autogenous

ratios

engines

balance

an engine

mixture

20K, and 50K lbf thrust

material.

in the 7.5K lbf engine

plots

10

at MR 6 + 1. Power

of 43.66

on I-

than

MR study,

was

based

greater

for the high

this is a flow

3.2.2-9

assumption

the guidelines

turbopump

an additional

Figure

and lunar

that this was

point

and

a 100% energy

length

is 30 to 45 lbf-sec/lbm.

greater,

turbopump

for the oxygen

assume

by a chamber

respectively.

pressure

recovered

the nominal

afforded

in delivered

lbf-sec/lbm,

Under vary

the mixing for the 7.5K

is much

predictions

This is not an unreasonable

measured

chamber

(cont)

performance

is attained.

performance

psia,

483:1,487.3,

delivered

efficiency

Baseline

Studies,

evaluated

are given

both in Figures

hydrogen temperature

121

and

baffle

The

evaluation

power

balance

while

the

NASA-Z

and

plates

is above

thermal

was

evaluated

done

that

maximum

3.2.2-11.

(Tin. limit

at two

showed

platinum

and

to the baffle

were

work

actual

3.2.2-10

the

plates

was

baffle With

= 499°F)

plates. the

the

for all settings

of

Engine Thrust, F, Ibf O O

0 0 0

0 0

l-

p.

!

| -

_

"0 c-

--1-

I

O

!

i1

I

X

0 .J

> :

:

i

!

a

0

Im

0 ...... _......._............. . 4...... X

om

I$ |

0 ....... im

C ........I..... t ..............._........ i ......... 4-......... I I i

"O G) 0 m >

'lO

<

O) I

¢_

It.

122

Cq

d

II

.m

m

_0

L_

0 0 0 II

d_

i__

0 tl,.-

-

2

I--

_ _

0

m

0 c_

E

E

.m

-'t._ -0

-

_:

- 2

ee-

& t_ IIIIIIIII

0 0 cO

IIIIIIIII

IIIIIIIII'

0 0 _D

0 0 _

(,4)

IIII1111!

IIIIIIIII

0 0 C_

0 0 0

e_n_.teduloj_

123

IIIIIIIII

0 0 oD

0 0 _D

0

N e4

I,L

00 0

I-1

ot I!

I:I: o J .... I_ I _" I

L_

,_

CII

_i.

cff

L

oCD

I

el 0 0 0 II

I

¢)

L___

J-

_

_l_

._. IX; C:)

I

I_

° ,.-.I

L

_

I

'

V\\

i

', /I,

n" Q)

1:1

__.L___

I

o

L_

I

I

m_

ffl

"k I

_o

"_ 0 C:; .,,.,t

0

I.

Q)

E Q) I'm m

.',_l-

- -,-

_-

I I I

Z

_._

"r" Q) r't_ rU.I

I_ I--

T-_111111111

0 0 ',-

Ilillllllllllllllll

0 0 0

IIIIIIIII

0 0 (:_

0 O _

IIIIIIIII

O 0 I_

Illllllll

0 0 (D

0 0 u_

O

& Q) i._

(_)

_Jrnl_.Iadma,I,

124

t_ U.

3.2, Engine the

flow

Requirement proportioner

capability design and

The

hydrogen

alloy

indicated

3.2.2-10

has

gave

the

peratures,

but

the

the hydrogen various

in the

settings

continued gains

3.2.2-13

channels.

This

chamber

rise

at all

optimum and

margin

temperature

valve

the

as

where

and

and

MR = 10.

as well

side

wall

temperature

3.2.2-14

be to optimize

geometries

hot gas

of the bulk

in Figures

would

baffle

the

chamber,

more

set by the

is a function

proportioner

and

bulk

case,

to the

provide

thermal

is not used

For this

going

would

life are

is presented

of this engine

below

regenerator

flow

design

temperature.

engine

capability

of the hydrogen

by adjusting

wall

and

the

the actual

within

is also

at 90°R.

baffle

maximum

power

development

baffles

gives

margins

engine

assumes

and

is well

chamber

54% of the hydrogen

actual

Design

the

3.2.2-12,

A platinum

Figure

however,

where

chamber

is adequate.

on the plot.

baffle,

settings Figure

the

setting

baffle

(cont)

A platinum

chart,

both

valve

Studies,

valve

next

enters

proportioner

Figure

valve.

at proportioner limits.

copper

Variation

rise

3.2.2-15

A design

the sum

temof

for

goal

for

of the enthalpy

as the proportioner

valve

setting. One of pressure

drops

pressure are

drops

over

size,

chamber

and

outlet

downstream equalize desired

2) orificing

the baffle

A high study.

(which

includes

the

condition

condition errors smooth

This

that

in effect,

Baffle

mixture

ratio

in Figure

5% autogenous of maximum

cause

The

with

and

is shown

is determined

drop.

by an iteration as plotted

possible

the

solutions

by restricting

it mixes

with

is to equalize

flow

shows

the

the

regen

the pressure

Otherwise,

they

distribution

from

drops will

that

valve. Wall

Temperature

implies

process to have

12 5

Limits

a reduced

hydrogen

where

hydrogen

flow)

is plotted

for that

curve.

R1,r/l:x_,7._/3.2-3._

the

3.2.2-16

circuit

elements.

change

3.2.2-13

thrust

before

is the equalization

Two

baffle

criteria

pressurization

engine

the curve

in the

circuit

Figure valve.

circuit

at the proportioner

Chamber

circuits.

drop

outlet

valve and,

commanded

3.2.2.3

pressure

pressure

proportioner

hydrogen

in the design

of the proportioner

and

to increase

to be addressed

parallel

settings

hydrodynamically

design

has

the velocity

of the

and

that

the two

for various

to 1) increase

channel

issue

mixture during

some

ratio.

flowrate against The

the power

minor

flow

inflections.

rate

for this

to the TPA mixture

maximum

balance,

and

It should

ratio

at

thrust has

small

be a

(:E)

6

L

"

-P--

II

L I

C)

I I

.p-I

Im

E :3 E

I L I I I I I I

X m Q C

llJ

_q

& III

0 0 00

III

cq

IIII

0 0 ¢D

Z--

.m

U=

126

Hydrogen Temperature From Cooled Chamber and From Baffles, °F 0 0 r

I

!

:

aas/wql

'au!Gu3 o], eleJMol-I uaGoJP_H

127

o

_-.__

0 II

!

0 eu

o.a

Q. f,.

-0

II

t

-0

p I

I

0

I

I

I

I

,@.I

I

I

I

I

0

4-

I....

I

I

I

I

I

--I

2L E

-I.

I.-

lm

no r-

I o L-

I

I

I

I

-r-

I

I

C

I

I

I I

I

I

I

o

r-c") -o

o o o

I

I I I

I|

I l I I

o

I I I I

I

I I I

o

I I I

I I

I I

I I II

II

I

I I

IIII

o o

III

IIIII

0 0

LU '1""

I o

I I"I

"0 >,

I IIIII

& N e4

I o o

o

tl

.,,,-.4

(,4)

oan eaod oj.

128

0

o

'=;_

,-

II

"

_

-©rj - c;

0: m "+ Q.

-

l:::l

t,-

-

(_

T-

-_rJ

•

'1"

-_. -o

II

o.u

I I I

I

Q

L

=

L

_

E X

I I

<:_

I

0,1

r-

I,u

I -iI

llllllllllllllllll 0

I

I I

I

I

IIIIIIIIII 0

O_

IIII!!1111 0 O

II

0 0

129

, _

II

III

I I

IIIIIIIII 0 0 (_

0 ¢) p,.

& N (6

0

_ L

_

.__ I1

0

Ir,.+

_ _._._

O II

|

m

+m

Q. T-T--

II

I

I

a, o

I

a

I

O

I-

L_

I_

--I

L_

G. O r= .i

e= 0

W

_5 & --e'3 -0

Q

°_

IIIIIIIII

Q

r ' i Q o P3

1,1= i

i

i

i

i

i

i

i

0

lllllllll o o

IIiiiiiii o o

(p!sd) do J(] eJnsseJcl

130

0

0

3.2, Engine

Requirement

Variation

As hydrogen chamber

slows

The result

13. The thermal throat

ature

near

changed sates

baffles.

if platinum

compromise

used

wall

independent performance

ratios

above

for the NASA-Z hydrogen shown

in Figure

of 0.52 flow inum

fraction

baffles

are

used,

This

limit.

mixture

mixture

ratios

chamber

from

life at lower

chamber

there

psia

chamber is some

hydrogen

the

in Figure flow

3.2.2-13

and

temperature allowable

wall

are for copper

temperature

would

be little

of platinum

This indicates selected

temper-

that actual

compenenthalpy

the use of platinum

does

that

the

chamber

setting the

to 10, copper

within

design

mixture

ratios

baffle limits.

increases.

This

chamber

and

circuits give

would The

adequate

have

as the operating

relationship

is

baffles

point

valve

baffle.

could

chamber

design

to keep also

be biased

wall design

margin.

would

setting

If plat-

an important

to be used

platinum

of 1100°F more

reducing

establishes

At

to route

is at a proportioner

copper

baffles

limit

is adjusted

can be increased,

two

temperature.

the design

valve

point

for the

wall

to exceed

temperature

crossover

between up

maximum

At

the regen improve

towards

lower

temperature.

1500

regen

increase

conductivity

begins

the baffle

10 to 13 a platinum

Similar the

would

proportioner

the proportioner

relationship

temperature

engine

Note

ratio

wall

3.2.2-

to a maxiwall

gas side

is the chamber

temperature

to the regen

temperature. At any

concern

circuit,

3.2.2-15.

thermal

temperature.

If the hydrogen

to the chamber

correspond

a baffle

outlet

increases.

in Figure

a chamber

the

capability.

10 the wall

material.

baffle

of the material

The design mixture

and

as shown

as the lower

hot gas side

program

through

channels

This is shown

(1259.6°R),

baffle

plots

but hydrogen

were

is relatively

(1509.6°R),

The

rates.

of 800°F

the velocity

in the heated

balance

A platinum

(2459.6°R).

and baffles,

for the higher

pickup not

alloy

to 2000°F

chamber

temperature

face of 1050°F

for copper

flow

set in the power

wall

reduced,

of the hydrogen

rise at lower

limits

hot gas side

temperature alloy

control

(cont)

is progressively

time

temperature

the injector

of 1050°F

flow

and the residence

is a bulk

mum

Studies,

bulk

Rx'r/_7_,,/3_-3.8

chamber

plots

pressure

maximum separation temperature

that

are given condition.

in Figures Again,

should

improve

of the curves

for the

rise

is evaluated.

131

3.2.2-17,

3.2.2-18,

the platinum

engine

life.

copper

versus

This

discrepancy

Note

baffle

and allows

in Figure

the platinum

3.2.2-19

for

for a lower 3.2.2-18

baffles

can be readily

that where

evened

__PI

I

I

I

I ,

II)_@ ,_ _

,

,zm

t

_

iX

, m i r ....

\

' -',.....

,

,I

I

,

,

_

..-'-"

_

Xt

,

/

8

t,,,cf

'V

-

"

-o .,-,

E

'

¥' '7_ ' "_r .... ,7- _--r-

\,

y

\ ,

,

i i'#i I

I

'

'

,

,

,

,

,

,

I

I

I

I

I

I I

I I

'

II

.

*lml

_,

-

E =I

............

I, oo

IIII

,

,

:°I_

,

,

. _

I

I

"_

I

I

I

LI,I

I I

I I

I I

_v_

'

,,,,i,,,,,,,,,l,,,,,,,,,I,,,,,,,,,i,,,,,,,,,I,,,,, o

o

o

o

(_[) az[n_,_aadcua,l,

132

II

°o

I

==

, -_

'_

°°

(6

E

.. I I

_

I

_

I

I_

I

I

I¢0

i

I

IE"

I

___ _ __ +- ...... i i I I ''

I_

zm

||

+_---_--i_ /

I

I

:

I

i_

I

,/

,

, \

7,

\

g

..

_ B

X

-_. _

.... r---

-0

_ -_--'....

-r-I

._. _ 0

_ _

m

@ _

,.. °_

c-

II1111111

0 C) ---

I11111111'

0 0 0

I11t11111

0 0 _

(d)

'111111111

0 0 O0

II1111111

0 0 r--

o-Inl_-t_dm°±

II11111

0 0 co

I

o

_

u.

133

p..

(p!sd) do_o eJnsseJd

134

3.2, Engine

out

Requirement

by a slight

chamber

adjustment

and

baffle

proportioner drops

3.3

engine

study

vehicle

study

the NASA

centers

on data

problem

was

Advanced

President

engine/vehicle

new

The

start

of the

study

material

summary

comments

section

covering

in specific

used

to cost

priate

improved

for the

m,r/D0417_,/32.3.a

areas

DDT&E following

had

very

vehicle vehicle

little

on later

work

done

1989, Aerojet during

reviewers

were

used

highly

of the system

The

reasons:

1 35

and

methodology

the material

as submitted.

Cost

major

Advanced

to

Launch ALS engine

to the ALS

with

same

Estimates

changes

Reference

was

that

In some

favorable.

The

gen-

study.

in estimating

test.

was

to

coordination

MSFC.

made

on the

let.

a response

The

Engineering

work was

fabrication,

engines.

and

initial

were

in the material in the

to

The

In particular,

TechSystems

contract

An

contractor

and

to

NASA

in this section

to LeRC found

the

to prepare

material

were

contract contract

contracts

or discussion.

submission

Test

under

contract.

study

need

work

contractor

attendance.

study

accomplished.

cost methodology

for the CTP

meeting

prime/engine

Development,

as design,

between

errors

the application

such

coordination

all of the

never

based

from

are under

before

was

where

techniques

initial

costs

made

1988 and

The

pressure

Martin-Marietta

primes

and

of vehicle

and

vehicle

on the

Nearly

of the original

During

(ALS).

the

of the three

are

by the NASA's

kind

Design,

System

primes

the work

Rocketdyne

required

underway

Engine

cost estimating

the

hydrogen

and

distribution,

definition

is updated

system

gives

to equalize

of the Boeing

two

This

initiative.

effort.

were

the

affected

that

corrections

and

well also

space

version

3.3.1

costs,

was

interface

some

3.2.2-19

of the

to be orificed

Whitney,

while

of the two

is an edited

instances

the range

to coordinate work

and

Center.

was

effort

at the

Figure

COORDINATION

format,

start

Study

to support

envisioned

Center

Flight

late

Bush's

valve.

have

was

Pratt

requirements,

Engine

would

the simultaneous

Aerojet,

Space

coordination

Also,

circuit

objectives

Research

the

bypass

at MR = 12 over

STUDY

with

teams.

Marshall

follows

baffle

study

teams

Lewis

(cont)

mixing.

of the

NASA

erated

drops

VEHICLE/ENGINE

One

The

The

to stream

Studies,

of the regenerator

pressure

valve.

prior

Variation

14 has

a breakout

methodology

considered

appro-

a of was

3.3, Vehicle/Engine

Study

Coordination,

Customer

test and

as MSFC version

of the

Both engines

•

Production

•

As both

endar

assumed

year

size

start

in 1992 with

would

also start

summarized $950M

assumptions number table.

dollars. shown

cost

It should in Table

of development The development

gories

and numbers

lent engines

bility

demonstration,

engines and

have

major

= 960 tests.

is given

in Figure

3.3.1-3.

testing

time

the

a minor

Note

be emphasized Costs

is given

effect

flight

program

testing and

is about

are directly

related

to the

for instance,

as the

the 960 given

in the

3.3.1-5. costs

a clarification

The 960 tests

assumed

This

15 equivalent

readiness

would

are totaled

total cost

from

by cal-

Fabrication

or decrease,

are changed

of the DDT&E

program.

A breakout

charts

that the

increase

3.3.1-1.

Development

three

that the costs

will

requested.

in Table

3.3.1-1.

3.3.1-2.

These

3.3.1-4.

life of 30 full duration The

has only

are given

in Figure

in Figure

includes

and a total of 43 engines

an average

overhauls.

in Figure

submission

life testing,

at the same

start on 2 Jan 1991.

in Figure

development

the flight

of magnitude.

engines

cost study

schedule

for the development

component

for the

costs

3.3.1-1.

was

and

similar

are the same.

of the two

tests

of tests

order

capabilities

and qualification

Following

same

be in development

in 1998 as shown

in 1992 as shown

the

full funding

and engineering

completion

the ALS engine

be very

cost.

assumptions

by cumulative

in 1990

would

difference

an on-contract,

of design

should

propellants.

are within

and industry

on DDT&E

The study

for both

use cryogenic

engines

Physical

requirements

CTP engine.

runs

technology

The basic

acceptance

is the customer

•

•

(cont)

a total

for development, firing,

tests

with

is equivalent

and

of the cate-

cluster

appropriate to 36 engines

of 58 equiva-

engines

for

qualification, testing.

The

relia43

refurbishments x 30 tests/engine

Table 3.3.1-1. LTV/LEV

Engine

DDT&E Cost

ASSUMPTIONS: e1990

FIXED $

eCOMMON eENGINE eFIRST

ENGINE THRUST

MISSION LBF

COST=$9-13M

LIFE = 30 MISSIONS

e.85 LEARNING eNUMBER eSINGLE

RANGE 20-30K

UNIT PRODUCTION

eENGINE

eTOTAL

FOR LTV/LEV

CURVE USED

OF DEVEL/QUAL # OF DEVEL/QUAL ENGINE

ENGINES TESTS

RELIABILITY

@90% CONFIDENCE

= 960

= 0.990

LEVEL

eCONTROLLER/HEALTH SYSTEM DERIVATIVE PROGRAM

137

MONITORING OF ALS STME

= 32

f-. (/)

O 0

i (3

Z m

Z Iii Z m

W 0

(_$) 1800

7VnNNV

LL

Z

0 I-,

0

GO

!11

a z

z LU

OJ

n

0 .-I

UJ LU 0 0

0

0

0

oO

_

_

0_1

(IAI$) ZSO0 -iVnNNV

139

0

m

GO

0

|

0 _D

0 _t"

|

0 0_1

i

0 0

i

,

0 O0

0 _D

--

i

0 _

0 O4

(_$) 1SO0 qVnNNV 140

0

0

O0 0')

0

_O) LLI _._l

,4 I

0

0 O)

i2.

:}

C_I CI)

I--

0 I-

,-I

0 0 0

0 CO

0 0 (,0

0 0

0 0

(_$) iSO0 qViOi IAINO 141

0

.u

I,I.

,, _

"-

,-

I1=_

°o°

<...

<:

I I_

-_

--

II_ IJ_

°

o. ee

ffl

u


-,"_ _

II

!

a<

O.

F,

0

-

2

O.

-I

< |

];Z IJ.l a >-

Z

< Z

LU a

i-. Or) I,LI I--

I-Z_:

Oa

.-I LU

rr n

LU C_

Lu< a>

142

Z

Z

I-

I-

::D

I:I:

0

ILl I-Z

o o

n

o

3.3, Vehicle/Engine

the

oxygen

and

Study

testing.

zone

room,

turbopump to obtain and

tests the

performed

tests

Component

Tests

throttling,

use

allow

system

testing

date

controller

and

existing

control

controller/engine

stands mentation

and

hydrogen

tests

existing

facility

Oxygen

and

heat

inlet

at the component

stability

capabilities

flowrates

will

mixture

level.

to the

fed

be

(bomb),

at this due

(HX's)

be a pressure

level

can be performed

diffuser

hydrogen

exchangers

will

for

"J" area

temperatures,

GG's

Combustion

tests

altitude

for cham-

and

exit conditions.

fed propellants.

and

prototype

workhorse

vacuum

with

These

cooled

level

"E" area

These

thrust

tests

also

test

expansion

ratio

demonstrate

instrumentation.

and

Testing

provisions

stands to vali-

uses

for facility

development.

systems

- also

and

2 additional

at sea level

supplemental and

testing:

chamber

systems

interface

Engine

engine

diffuser.

monitoring

capabilities

in the

at required

chamber

require

controller

required

(GG's)

performance

without

c)

generators

each

stands

using

facilities.

2 stands

ratio.

- may

health

or existing

farm

gas

Breadboard

room

tank

as follows:

acquisition:

at Aerojet

facility

of J area

expansion

in "E" area

2 new

chamber/injector

and

are planned

database

and

and

engine

thermal

facilities

development

oxygen

pressure

b) required

new

facility

the

and

will be performed

use

gas

require

chamber/nozzle

test type,

instrumentation, will

(cont)

turbopumps,

All thrust

with

may

a)

to simulate

design.

ratio,

program,

warm

pressures

O2/H2

Test

hydrogen

ber/injector control

Coordination,

testing:

requires

will

utilize

supplemental

the

2 additional

control

room

and

test instru-

capabilities. d)

in J area

using

existing

testing

will

be done

facility

limitation

burns.

Also

Reliability control with

for thrust

requires

demonstration room

and

the exception level

I additional

testing:

existing

identical

high

facility

143

1 additional

capability.

cooled

(>1000)

capabilities.

Rr-rjD0417 ssa/32-3.s

altitude

of radiation

at very

requires

area

at (TBD),

nozzle ratios

Full extension

and

possibly

test up

stand

engine (possible

long

duration

with

altitude

3.3, Vehicle/Engine

cells

Study

e)

Engine

f)

Preliminary

at SSC - verifies

environments. ration

Coordination,

qualification

system

May

Flight

level

require

testing:

same

Readiness

transients

and

a diffuser

adapter

Readiness

Testing:

as reliability

Testing

testing

(PFRT):

engine

interactions,

for unique

engine

(d) above.

Adaption

to test

thermal

radiation

cluster

nozzle

configu-

envelope.

tested

in J area

g)

Flight

h)

Acceptance

at altitude

One failures

assumed

question

for the test

well

the test engine.

is no longer history

usable.

of the RL-10

the ICHM per

failures

system engine,

would

From and

asked

SSME

to predict however,

and

will result

answer

violence failures

One

acceptance

number

is "none"

goals

of the

failures.

engines

based test

test

of failure

a test stand

where on

total

design

as

a test article the

program

The

exceeding

type

damage

can be expected

of the

of engine

if the

can be expected

catastrophic

in some

the

to seriously

failures

preclude

will be individually

concerned

short

serious

engines.

(f) above.

as d).

The

6 to 10 of these

as PFRT

engine

by reviewers

be of sufficient Less

same

each

(same

program.

This

as destroy

testing:

conditions

is catastrophic.

tests

(cont)

test

is to use number

life, and

of some

can be expected. To summarize:

•

6 test

stands

•

2 test stands prototype

•

1 engine

required

in J area

and an additional and engine

test stand

level

control

1 test stand at altitude

•

Special

required

required

room

required

in E area

in J area for reliatesting

at (TBD)

at altitude

conditions.

for reliability/qualification/testing

conditions.

diffuser

may

be required

development.

_',riDo_17a.13_-3J

development

testing.

bility/qualification/acceptance

•

for component

144

for cluster

PFRT

and

FRT

for

3.3, Vehicle/Engine

Study

3.3.2

Coordination,

Engine

Production

Estimates comparison,

2) the

comparison in the

is the

portion

(nominal)

of the

thrust

of 20K to 30K lbf. technology series

will

be more

with

thrust,

percent

a different

curves

a good

through

were

the

factor

on each

chart.

The

charts

are used

by selecting

cost

cost

then

was

The

study

due

to

bounds

are

range of used

for this

parameterized

as the entering

results

in the level

impact

now

lbf

higher

to build.

likely

quantity

a 15,000

and

for

product

a thrust

considered

Production

production

mature

to have

costly

baseline

given

using

agruments

for Figures

with

3.3.2-1

3.3.2-5.

a thrust,

say

appropriate

20K lbf; a learning

curve,

complexity

For the example

factor.

Complexity

These

are

1990

spread

ment

dollars.

over

and

support.

This

expect

Cost

A basic

production

production was

to get

($M)

One

not estimated level

cost. likely but

and

then

given,

quantity,

entering N th unit

the

say chart

cost

($M)

1.2

1.3

1.4

1.5

5.5

6.0

6.6

7.2

7.7

is that This

area

DDT&E

makes

costs

a clean

of additional

is highly

of supporting

a production

1.1

assumption

engine

costs.

any

say 85%;

Factor

N th Unit

can

figure.

more

for

The

It is also

capability

to gauge

90%

curve.

is a fully

flat.

its increased

10% to 50%

and

This

is expected

therefore,

of 80 and

curve,

The

not

and,

median

learning

complexity

with

from

engine.

engine

1) a baseline

learning

is essentially

the CTP

complex

factors

85%

that

engine,

cost require

3) an expected RL-10

curve

CTP

Learning

engine

and

where

The

production

Whitney

learning

of complexity

complexity.

thrust, and

engine

Cost

of engine

engine Pratt

(cont)

engineering

break

cost

discretionary

are

with

costs

between

case;

subject

the

is:

and

develop-

is continuing

in any

desired

sunk

48 engines;

engineering the customer

to funding

constraints.

3.3.3

DDT&E some

costs

unplanned

_/_17_,,,3.2-3.s

Mission

Related

One

of the

concerned

the

engine

Costs

questions impact

asked of having

servicing.

145

after

receipt

planned

of the

engine

Aerojet

servicing

input

on

only

versus

are

a

C) cO

@ (N

.i

Tm

II

m

G

oO 0 c m_

_

C

g

(MS)iso3 ilNnHiN

146

_

"°

$ @

ILl

II

_z

"_

o

,,>,_ -I

F-

14 e.i

& cO

-_-

("4

Cb

03

ca

_f

(mS) ISOO fiNn HiN

147

("4

00

(3"}

03

o l,i

P

iiii 0 |

-I II: x

1!. >FZ
II

c)

o_ _c: la._o

I--

0 _t_

Z

/

w_

l i

._I


0

-IZ

I

_

Cb

2 U.

O0

u3

_I-

O4

C3

03

u3

(IA_) ZSOS ±INnHIN

148

O4

4) r"

im

$

tUJ 0 1-,. |

I

.-I

rrx

II >.,

J_ z <_

q) Q.u)

G

oO (j e-

Z

4-

0

C

=

EO

Q 0) 0 C ,uJn

t

I_1 r-..,..

14 0

i

Y

¢9 q) _m

ol U.

I

(_)

ISOO

llNn

149

,m

HIN

©

@

$ @ e"

=m

e. U.I o T,= |

-I I¢ X

>tl--

m. Ip=,.

II

Z ,,<

0

= _.= Eo

-I-

oO 0 ,''_

0

@ 0 C =. .1

•1" I !

J

wa.

I-0 I--

u,l'E

_z

j,

Q

|

@

1,1. Q

i.n

c]

(mS) lso9

J_lNnH..LN

150

3.3, Vehicle/Engine

Study

The and

whether

maintenance assist

the

Coordination,

answer

is dependent

the integrated decisions

planners

(cont)

control

and

A maintenance

on what health

monitoring

scenario

architects

system

is used

Some

guidelines

is needed.

plan

to do,

for real that

time may

are:

An engine

should

capability

is not going

"Don't

pay

Series

production

nance

capability

basic

first

is engine

The

newer

engines

maintenance

removal

should

vehicle

you

to space-based

in-space

include

design

is needed

to determine

out

versus

engine

Maintenance

costs

in a space

of in-space

change-out. include:

- Training - Facility - Downtime - Spares

and

Spares

- Tools/Diagnostic - Personnel

are

provide

the

Storage Devices

Costs

- Inspection/Certification - Administration

15 1

to expand

maintenance

life.

as the

the facility

mainteis

built.

capability.

a lot of maneuvering

assessment

can be used

this

suits

its operational

if the

need."

replacement.

to a person

A realistic

don't

and

must

maintenance

during

capability

components require

for in-space

to be available

of an engine

and

The

vehicle

not be designed

for something

developed

k_/D_7_/3.2-3.S

the mission

access suit.

that Even

to the This

must a block

space

be developed I design

maintainable

is non-trivial,

the

room.

maintenance

cost effectiveness

costs

and

of component

timelines change-

3.3, Vehicle/Engine

Study

Coordination,

In-space

(cont)

maintenance

mal shops

or flightlines:

- Vacuum

Environment

- Specialized

Tool

- Lack of Gravity - Limited

not encountered

in nor-

Assistance Working

Problems

Reliability

problems

Kit

Effective

- Access •

presents

Time

for Mechanics

in reconnecting

in Space

flow

lines

after

Suits maintenance

is a

concern.

•

Insulation

and connections

complicate

•

Actual, more

component

The Aerojet careful

look

develop ations

at designing

a rapid, except

face plane. arm reach

consuming

7.5K

for in-space

uncomplicated

All operations once

access

removal

of some

steps.

are done

to the vehicle/engine

sort is expected

Design

far

a

concern

All engine

was

removal

interface.

Table

within

physical

internormal

3.3.3-1

of this sequence.

to

oper-

at the engine/vehicle suit,

for the actual

been

included

The first

in a space

is the reverse

to be available

have

expected.

Preliminary

by one person,

Replacement

than

procedure.

nozzle

sensors

operations

(see Ref. 2).

changeout

can be done

is gained

Engine

maintenance

engine

repair

and fatiguing

lbf Thrust

management

out.

in-space

for the extendible/retractable

the required mover

change

demonstrated time

for health

is a list of

A prime

handling

of the

engine. Engine for instance, requires should

removal

has a design

only

two

not require

can be done

for the engine

ball lock devices more

than

with

attachment

for keeping

an hour

a tool kit of very

once

to the thrust

the engine

access

is gained

few

takeout

in place.

items.

Aerojet,

structure

Engine

that

removal

to the engine/vehicle

inter-

face area.

Some various

design

RP_/_17_,/32-3J

of the questions

requirements.

In general,

asked

related

there

to the effect

152

is no impact

unless

on DDT&E requirements

cost

of exceed

TABLE CTP

•

Engine

•

Propellant

•

Electrical

•

Manual

Centered,

ENGINE

Nozzle

Isolation Power

Valves

OPERATIONS

Closed

Removed

from

the Engine

Operations:

Electrical

--

Extendible Nozzle Removed, Edge Protector Installed

--

Main

Hydrogen

--

Main

Oxygen

--

Hydrogen Tank Valve, Capped

--

Oxygen Tank Valve, Capped

--

Engine

Handling

Fixture

--

Control

Gimbal

Actuators

--

Flex Lines

--

Prime

--

Engine

kmvi_,_-_,-'r

REMOVAL

Extended

--

Engine

3.3.3-1

harness

Line Line

Thrust Out

Connectors Screw

Disconnected

Autogenous

Below

Pressurization

Harness

Secured,

Shutoff

Shutoff

Pressurization

Autogenous

Connected

Capped,

Assemblies

Below

Disconnected

Restrained,

Mover

Moved

Disconnnected,

Line

Line

Regen

Valve, Valve,

Stowed Cooled

Capped Capped

Disconnected

Disconnected

Below

Below

Connected Disconnected

Upper

Engine

Covered

with

to Engine

Structure

Locking

Devices

of Engine

Compartment

153

Removed

Nozzle

Protective

Material

Shutoff

Shutoff

3.3, Vehicle/Engine those listed

in Table 2.1-1.

requirements. between

Study Coordination,

This is true of such things

Life cycle costs are implied

missions

and whether

removal

operational

and refurbishment.

history.

may be planned operation would

During

in space. service

(by remote

change.

video

is a possible

of things

hours

operation

the engine

to earth engine

for refurbishment

to earth

deployed system, later date.

for refurbishment

orbit burn,

The LTV duty

cycle (typical)

and earth

starts each mission.

about

1655 sec operation

would

,m'z_17__,,z32_.s

Total LTV propellant

on 4 engines

this is about

a translunar

operation

350 starts.

With 5 reuses

35 starts

7 times

and about and about

the operating

154

versus space

engine

design

to operation

injection

correction burn

at a

burn,

burns,

a

a

for a total of

A fifty mission

for

life would

This is close to the life

prior to scheduled 2.4 hours

is

of a space

(288,000 Ibm) is sufficient

of 80K lbf total thrust.

program.

accumulate

it was returned

have

of 20

If the engine

in terms

circularization

engine

goal set for the OTV engine

it could,

and a post airbraking

seven

of engine

would

following

(equivalent

years of operation.

though

has

that will be the most

two midcourse

correction

24 hours

burn,

maintenance

burn,

pre-airbrake

requirement

Our baseline

assuming

even

return

to the end of

The present

facility.

a

maintained

are not defined.

after five missions expended

a de-orbit

Our design

for an

building

routinely

after 40 missions

the early

be considered

when

in space

capability

during

a cause

of refurbishment

for a maintenance

feature

system

(ICHM)

of the engine

detect

to predict

with periodic

be discarded

change-out

maintainability

in perspective,

some

and inspection

of uses prior

scenario.

The economics

has no provision a quick

each engine

with

inspection could

that are hard

to 40 missions would

and 500 starts).

astronaut)

The number

on the maintenance

The engine

monitoring

with a complex

exception).

a removal

Also, an optical

unit and/or

are a number

fifth mission.

total about

control

for an engine

every

lunar

system

to

system

life equivalent

returned

be for a mature

prior

and health

by the integrated

been an operational

important

is for five missions

or out of limits

We do not have experience

emphasizes

requirement

In all cases anomalous

There

design

after five missions.

of uses

missions.

life is dependent

station

is expended

This would

the number

after two or three

(Skylab

returning

regarding

of the vehicles

after each mission new system.

and reliability

operation

be justification change.

range

early

recorded

engine

design

as throttle

by questions

the engine

The OTV engine engine

(cont)

operating

time accumulated

maintenance time.

To put this

by the space

shuttle

3.3, Vehicle/Engine

Study

main

engine

(SSME)

gone

seven

such

as the OMS

on an orbiter

missions

maintenance

Coordination,

without

engines

mission.

extensive

with

free goal.

(cont)

maintenance.

a fairly

Meeting

We are not aware Some

high-margin,

it with

a high

of any SSME storable

conservative

performance

that has

propellant

design,

cryogenic

engines,

can meet engine

this

will

be a

challenge. One and

storage.

of the capabilities

At some

considered

time

necessary

establishing

such

in its evolution,

if the Lunar

a capability

3.3.4

In-Space

of factors

the Aerojet ability

version

of the engine.

received

major

because

design

that could

The list is both vehicle

engine

structure

and/or

engines the

are given

Phase

engines

A vehicle health

to the mission system

indicates

the scheduled

studies.

monitoring operation there time."

adhere

to a timephased

engine

condition

wr/t_7_./3_.3.8

in-space

3.3.4-2

and safety.

With

as noted)

and 3.3.4-3

is used. Despite problem

a sophisticated

components

not time

This system a scheduled the decision system

Preventive

determined.

155

(see less

Table

of

3.3.3-1)

consideration in place.

the engine

Without

stand-

features

the engine

is given

in Table access

compo3.3.4-1.

through

the

will be inaccessible.

and servicing

ICHM

plan.

with

the

the maintain-

given

capability

sensitive.

respectively.

assessing

was

changeout

of the six servicing

maintenance

determined,

of these

design

features

be replaced

design

some

In three system

when

maintenance

of maintenance

is an engine

be

The cost of

from

addresses

replacement

in component

aerobrake

in Tables

may

there.

maintenance

removal/replacement

and vehicle

The full range

capability

is based

in-space

will be of concern

to limitations design

(LEV)

This section

Component

a competent

(subject

cost.

engine

for reliability

With nents

that

emphasis.

of the concern

Vehicle

addressed

the life cycle

The rapid

maintenance

changeout

and Servicing

section

of the engine

base is engine

be substantial.

Maintenance

comprising

at the Lunar

a greater

Excursion

will

The previous point

needed

functions These

were

operational would

from the

to be very

maintenance

plan,

critical

if this

to be "fix it now,

it will be very maintenance

tile

adapted

functions,

appear

is likely

involving

may

difficult very

not to

well

be

at

TABLE CTP ENGINE

3.3.4-1

SPACE MAINTAINABLE COMPONENTS 7.5K LBF THRUST ENGINE

Comment

Component Hydrogen

Main

Oxygen

Main

Hydrogen

Shutoff

Boost

Oxygen

Boost

Hydrogen Valve

Shutoff

Valve

Engine/Vehicle

Interface

Requires

Access

Near

Engine/Vehicle

Interface

Requires

Access

Near

Engine/Vehicle

Interface

Press)

Requires

Access

Near

Engine/Vehicle

Interface

Pressurization

Requires

Access

Near

Engine/Vehicle

Interface

Requires

Access

Near

Engine/Vehicle

Interface

(Low

Autogenous

Hydrogen

Near

(Low

Autogenous

Oxygen Valve

Access

Valve

Pump

Pump

Requires

Press)

Pressurization

Regenerator

Bypass

Bypass

Valve

May

be a Bolt-On

to a

Design Dependent; Manifold

May

be a Bolt-On

to a

Oxygen

Regenerator

Gimbal

Motors

Requires

Access

Near

Engine/Vehicle

Interface

Gimbal

Actuators

Requires

Access

Near

Engine/Vehicle

Interface

Extendible

Nozzle

Extendible

Nozzle

Deployment

Extendible

Nozzle

Dep.

Valve

Design Dependent; Manifold

Ready Motors

Mechanism

Access

Requires

Access

Near

Engine/Vehicle

Interface

Requires

Access

Near

Engine/Vehicle

Interface

Fuel Flowmeters

Design Pressure

Dependent; May Boost Pump

be Bolt-On

to Low

Oxygen

Design Pressure

Dependent; May Boost Pump

be Bolt-On

to Low

Flowmeters

Controller

One

or Two

Removable

Boxes

with

Cannon

Plugs Sensor

Signal

Conditioning

Units

Miscellaneous Hardware, Brackets, Wires, External Sensor Elements

RPT/r_lT-SS,-T

Designed to Allow Sensors only Electronics Changed Recalibration) Dependent

156

on Access

to Remain in Place, (Requires System

TABLE SPACE-BASED

Perform

scheduled

propellant

Perform

visual

Determine

inspection

ACS

fault

FUNCTIONS

storables engine

Perform

system

Service

batteries

Replenish

(includes

engines)

Control

System)

modules

(after

each

mission

if

are used)

module*

(after

operational and

stored

unscheduled

LTV

status

(Altitude

Replace

Perform

to and from

LTV

packaged

Perform

MAINTENANCE

maintenance

Transfer

Replace

CTP ENGINE

3.3.4-2

fuel

helium

TBD

mission

time)

testing cells (if used)

maintenance

damage

Verify

any

Isolate

fault

electrical

assessment

damage

Perform

required

inspection)

unit

repair "remove

The vehicle can be developed engine replacement. In-space design solution. Table 3.3.3-1 design concept.

RI'I'/I)0417-._a-T

scheduled

failure

to replaceable

Perform

(beyond

and

replace"

due

to failure

with replaceable propulsion modules or for individual handling requirements may dictate one or the other lists the steps in removing an engine in Aerojet's

157

TABL_ CTP

BERTH

SERVICING

Rendezvous

•

Capture

•

Berth

LTV with LTV

FUNCTIONS

Station

at Station

LTV at Station

TRANSFER

PROPELLANT

•

Verify

•

Perform

Interface

•

Transfer

INSPECT

Integrity

Propellant

Leak Check

Residual

Propellant

from

LTV To Station

Tank

Farm

LTV

•

Perform

•

Determine

•

When

Visual

Inspection

LTV Fault

Fault

Status*

or Damage

Detected*

Perform

Damage

Assessment

Initiate

Electrical

Initiate

Fault

Formulate

(TV/EVA)

Test Routine

Isolation

Integrated

PERFORM

to Verify

Fault

Routine

Maintenance

Plan*

LTV MAINTENANCE

•

Perform

Scheduled/Unscheduled

•

Mission

Reconfigure

•

Perform

System

•

Deactivate

MATE

OPERATIONAL

LTV

•

•

ENGINE

3.3.4-3

•

Transfer

•

Mate

•

Verify

•

Perform

Operational

and

LTV AND

Stow

to LTV Interface

LTV/Payload

Integration

Test

LTV/PAYLOAD

Perform

Prelaunch

•

Transfer

Propellant

•

Launch

a_T/_IT-SS,-T

for mission

to LTV

•

* Operations

Testing*

LTV (if not required

LTV/Payload

LAUNCH

Tasks*

PAYLOAD

Payload

Payload

Maintenance

Operations* from

Station

to LTV

LTV/Payload*

where

the LTV engine

Health

Monitor

System

is used.

158

at that

time)

3.3, Vehicle/Engine

Study

3.3.5

Engine

3.3.5.1

Mission.

With

Vehicle

the

One

Freedom.

provided

actual

studies

vehicle NASA

lunar

MSFC

transfer

is a modular

four

pletion

TLI burn

moon

or, possibly,

to earth Crew

Module/Cargo

cargo

This

four

engines

extendible/retractable reduces

the length

The

along

second

crew

with

from

nozzle

LTV and

identical

to those

extension for the

used

landing

two

core

propul-

lunar

vehicle.

of these

tanksets

are retained

for com-

to the lunar for impact

base.

The

on

LTV

the

returns

tankset.

A payload

of 27 metric

tons

ton crew

module

logistics

support

can be transferred

is plus

to the

LEV

in lunar

(LEV)

which

has

module.

is the lunar

the

the

released

for an unmanned

cargo

to and

core

basic

for attaching

propellant

of a 4.8 metric

module

the

operation

LEV are

required

Module. can consist

vehicle

for actual

propellant

on the

in the

the

main

provisions

two

then

at the

with

stage.

burn

other

are

landed

or can be all cargo

mission.

with

soft

two

Station

the

around

to transfer

They

on the propellant

proposed.

is serviced

(LEV).

at Space

has

(TLI) The

and

proposes

comprise

pattern

emptied.

support

of the core

injection

change.

structure

engines

shaped

Return

modules:

The core vehicle

when

vehicle

base

tankset

is part

in a cross

may

is based

three

Four

aerobrake

the trans-lunar

excursion

which

system.

of the

lunar

is a single

are jettisoned

orbit

of the

This

for the Lunar

the concept

up form

Tanksets.

concept

started

(LTV)

built

An

tanksets

During

legs

vehicle

propulsion

Separable

The

just

Stage.

array.

a vehicle

concept

vehicle

stage sion

Mission

Concept

LTV Core

set of landing

for the Lunar

NASA-MSFC

is the

This

(cont)

Requirements

The vehicles.

Coordination,

excursion

from

the lunar

is actually on the on LTV

legs.

RlYr/I)O417.55a/3.2-3.8

159

vehicle surface.

based

at the lunar

LTV except engines

During

As noted

has

operation,

a

above,

station.

The

the been only

removed. two

This

engines

it

3.3,Vehicle/Engine Study Coordination, are

operated

than

those

above

pumped

for the

LTV.

and of a controlled with

weight baseline The

are given

in Figure

This vehicle

concept

One of the study orbit

inflections

that were

the vehicle

weights

440 seconds

specific

curve

be prepared

a direct

shows

necessary.

lunar

issues

•

Control

•

Hot

•

Oxygen

impulse

program

mission

in Engine

The

into

towards

on payload.

is given

on high in Figure

system

stability

any range,

are:

and change

rate.

temperature

circuit

Combustion

rise on throttling

liquid-to-gas

stability

over

160 RPT/DO417.K$a

/ 3.2-3.8

specific

performance 3.3.5-3.

Throttling over

phase

change

the

throttling

down. at low

range.

3.3.5-2. has

to generate is why,

at a

A companion

As it stands,

pressures. •

shown

versus

for the

as plotted used

infinity.

payload

of the LTV

in Figure

curve

is not dearly

20,000

impulse

the program

emphasis

profile

in throttling,

section

goes

specific

are presented

What

maximum

of specific

the OTV engine

Issues The

have

vehicles

set of four

the sensitivity

is evident. built

weight

point

of both

an engine

to the engine

engine

concept.

the initial

of the effect

The proposed 3.3.6

on the MSFC

that would

why

using

of this trade

impulse

sketches

to determine

of the assumptions

impulse,

evaluation

this curve

The results

specific

a result based

was

the touchdown

3.3.5-1.

was evaluated items

for the LEV are lower

over

Conceptual

at the start of the mission

ton payload. of a high

requirements

of hovering

and weights

27 metric

could

is capable

of a "g".

earth

importance

as the thrust

at a fraction

engines.

in low

thrust

The vehicle

landing

dimensions

lbf thrust

idle

(cont)

system

impulse however, is

for

-

I

.....

I L_

161

== i

"

:

.

_Z

i

.1

i

'-'

i

/

S i= i

,.Y

,,,,,,,,,,,,"""'

_==. "E"

i

!

!

!fl

!

i

i

,:

!

i

i

,_

_==

® ,,o, .J e¢3)

N 0 0

0

0

0

¢o

0 0

¢4 4) =._

=l

(mql_) lqO!eM AZ'I le!llUl

162

IL

;o_ .o I i_._1

_=_.

_

_

_

_z_.l

m im

UJ ..4 a LL. 0

0 C 0 Im

m

om

Z 0

m c

al

_I

N

e4 m Iz:

'_.

0 n,, Z _4

_°

_

d_

,_

!iI' _1

' !_ _ _ "'1

_1

i_ .!

163

L.

I1

3.3, Vehicle/Engine

Study

Coordination,

•

Turbopump

operating

•

Engine

operating

envelope

•

Specific

impulse

degradation

As the throttle with the

hot section lower

may

(primarily

mass

flow

be set by the

alloy

chamber).

control

and

This

velocity level

is also

speed

system

instability

the oxygen

all times. diagram

At lower the

erratic

due

show

erratic

with

to film

the range

where

phase

where

to gas

temperature

reads

due

throttle

1050°F

operating

curves This

change.

At high

to

range

or lower

change.

phase

uneventfully

with

enters

(copper

may

show

lead

to some

system

pres-

can

flow.

and

energy

release

properties

"dome"

and

flow

A gas-liquid

mass

combustion

changes.

Aerojet

in a LOX/GH2 the problem

predictable

the 2-phase

are unpredictable

removes

concern

increases

A maximum

pressure

the oxygen

performed

is increased

rate.

2-phase

This

wall

turbopump

and

change

10:1, there

coolant.

is the oxygen

properties

of the injector.

side

decreases.

temperature

response

the transition

ratio

beyond

for the output

problem

boiling

mixture

the oxygen

upstream

the wall

thermodynamic

as thrust

gas

where

pressures

restrictions.

increases

chamber)

or a lower

makes

range.

of the hydrogen

changes

A special sures

range

thrust

thrust

dis-proportionate

(cont)

heat

from

of the rates

element

uses

T-S

are could

a gas-gas

exchanger

at

element

(HEX)

the combustor

to a less critical

component. The chamber

channel

resulting

chamber

full thrust. in the cold

length

gas

hydrogen

erator tributor

and

exiting

to transfer

by 40%.

the pump

Figure

for CTP

engines

of various

to the

thermal

energy

Refer

l_'r/D_,7_,/3.2-3.8

3.3.6-1

variations

circuit TPA

shows

thrusts.

energy regenerator

turbine

section.

for the cycle

to Section

establishes

the thermal

the hydrogen

from

rise

For wide

is to use a hydrogen

entering

thrust

temperature

geometry.

is too short

A solution

hydrogen

the engine

hot section

The the

Note

that

section resulting

164

in throttle needed

discussion

to run

cooled the the cycle

the excess

is counterflowed heat

transfer

of the

the regenerator above

ratio

where

effectiveness

at all thrusts

3.2 for additional

the regeneratively

7.5K

hydrogen

is a significant lbf.

of engine

energy with

can

throttling.

at

the

increase regencon-

Hydrogen Circuit Enthalpy Change

o_ o;

0.8

........

¢0 ¢3

_c

0.5

Q. t-UJ

0.4-

O2

0 7.5K

20K

25K

Thrust

Level,

35K

50K

klbf

Regenerator

Jacket

Baffles

Figure

3.3.6-1.

CTP

Engine

Dual

165

Propellant

Expander

Cycle

1

3.3, Vehicle/Engine 3.3.7

Study

Weight

Penal_es

Space purge

system

basing

is not needed

propellants

are self-purging

turbopump

baselined

nal conditioning

units

controlled

must

under

will

need

included,

for the most

be answered

as generation

study.

that specific

Now

preliminary

(worst

case)

Rvr/_17._,/32-3_

electronics,

gimbal

weight

not add

to request

A few items

weight

be in engine

directly

and

LOX

purge

sensor with

gas.

sigthermo-

to the assembly. spares

and

to LTV or LEV weight

maintenance

for various in Section was

are

but

capability.

not covered

Requirements

Electrical

power

requirements

(see Ref 2).

for 13 different

Table

engine

were

operation

extension/retraction

0

Engine

Out Gimbal

(2 engine

1

Configure

engine

for chilldown

2

Chilldown

02 TPA

3

Chilldown

H2 TPA

4

Engine

5

Tank

head)

idle

166

of parametric

Some

state

vehicle prior

could scope

to this study 3.1 are given

determined gives

data

requests

of the contract

a follow-on

3.3.8-1

Nozzle

(tank

3.1.

in Section

00

Start

types

outside

known

Power

head

seal

conditioned

pounds will

driven

Request8

of interest

task

actuators

thermally

penalty

Electrical

in watts

and

add a few

of the information

design

the hot oxygen

control

in the information

areas

and

as an inert gas

The cryogenic

an interpropellant

for a space-based

responses

that information.

engine

conditions

This does

part,

weight

of the CTP engine.

This will

Data

engine

not require

significant

Parametric

some

does

circuits.

into orbit

Aerojet

save

to be insulated

equipment/facility.

3.3.8

actually

vacuum

actuators,

heater

still be carried

generate

may

Basing

for the AT version

The most maintenance

(cont)

for Space

for the engine

Valve

statically

Coordination,

for the the power

points.

The

configuration) to start

are not

for this should

below.

7.5K

lbf thrust

required state

points

are:

i-

E 4) L1

-i O" 4) rr L_

0 a. C

I1

(2) C Q >

C

Q. IIC

I

m

,D

167

3.3, Vehicle/Engine

Study

Coordination,

6

Pumped

7

Normal

8

Overthrust

9

Normal

Shutdown

10

Normal

Gimbal

11

Operational

External

operation

the throat.

With

and from the

portion

nozzle

The two

engine

(near

curve).

curves

side

Figure

3.3.8-2

Cluster

engines

on side loads, flexibility, cussion

and gimbal with

design

Attach

The

work

tolerance.

point

above

the lines

to

The

higher

area ratio

outside

for a columbium

of heating

radiation

for

of the oxygen

600).

much

versus

by an adjacent

cooling

but for a carbon-carbon

to date The

(far nozzle

Points

and Gimbal

of gimbal

side extension.

attached

brackets.

Aerojet

with

ball lock devices

nuts,

bolts,

fastening

of this clearance

movement,

engine

to the thrust has a design

that can be connected

devices.

clearance is dependent

nozzle

length

that requires

and

some

dis-

Method

is actually

made

a 12 inch

This is one of the items

to determine.

engine

has assumed

adequacy

on engine

primes

sets

or other

variables

constraints

pair of the two two

have

for direct

and

at full

will be negligible

ratio

for the effect

to space

3.1-5)

temperature

(area

even

28) to some

radiation

external

insulated,

exit plane.

system

the vehicle

ratio

touch

Constraints

at the nozzle

mechanical

external

to account

to the

(see Figure

the temperature

the same

All Aerojet between

gives

unless

or exposure

gives

(area

near the exit manifold will,

Engine)

be cold

manifold

insulated,

are needed

curve)

would

The maximum

nozzle 3.3.8-1

(Buried

a can structure

circuits

be 400°F

Figure

inlet

inside

of the engine.

extendible/retractable

nozzle.

chamber

the hydrogen

would

temperatures.

Environment

thrust

the chamber

range

Storage

Radiation

the injector/baffle

upper

cooled

from

Idle Mode Operating

The engine thrust

(cont)

The pitch

and

takeout where

by a person yaw

structure

by one

this attachment in a space

axis gimbals

is

suit; no

are attached

to a

0 _00 O0

-

g

ioo O0

e

[ o _130

I°_

-

0 rY

-

o

o

0

Z

<

-OO -0']

d q o

I

-_)

I,..

-I .I

_L

-0

169

O _00 u_

-

-00 0"_

i

L

¢:

,--

O

0 C 0

_ -

£

00 "_

u_i-5

-O

q

-r"-

e4

o

|

_,,_-

m O'1

OH

u. I

I

I

O00g,

I

I

I

l

1

I

I

t

I

I

t

OOgL .J ';gJr'i°JSdWg.]_

I

I

_

I

1

I

i

l

-_0 r_

000[ _0

170

-C -,'r3j ,F)

3.3, Vehicle/Engine

structure

at the

be for a true done

Study

top of the

throat

by slip-on

for assembly made

with

without

3.4

Dual

The

of the study

based cycle

when

engine

preliminary

version

in Figure

regenerator oxygen

circuit.

surface

area

same

engine

capability with

excellent

cycle

state-of-the-art.

for the

engine

operation

sequence

results thermal

basic

3.1.1.10. Liquid

indicate

that

for very

were

be

required can

be

of an astronaut

the engine

be an excellent

developed

design

latter

resulting life.

has

weight

due

based

on engine

from

the

to increased would

promises

the

hydrogen

circuit

version

lbf thrust

although

the

Either

engine

propellant 7.5K

baseline

hydrogen

in HEX

be made

dual

in the

version

of the

at the beginning

flow

developed

engine

isolation

long

give

the

control

to be highly

It represents

the current

loop

system,

a dosed

and

sensors

in Table

corresponding Transient

control

to the integrated

valves

is outlined

Engine

cycle

The

3.4-2.

control are listed

flexible

expander

Simulation

is controllable

17 1

and

Analysis

states

health

in Table

are given

stable

control

monitoring

3.4-1.

of valve

(MLETS)

is dynamically

a dozen

and

A discussion

component

equilibrium.

RI'r/tX_7._S,/3_-3.8

could

Baseline

engine

The

current

The

should

for the parallel

flow

would

for input

The

engine

connections

the task

was

temperatures.

of sensors

(ICHM).

Modified

delta choice

Control

2.1-1

by an increase

will require

system

in Section

series

engine

and

The

lower

margins

valves,

given

an array

results

for better

minimization.

This

no tools

of this cycle

be evaluated.

The

Engine

so that

simplify

in Figure

It is the

be penalized

thermal

3.4.2

version

to the

of the HEX

weight

to the

would

Cycle

design

task.

performance. and

connection

point

UPDATE

given

should

It may

needed

flow

thermal

design

upstream

gimbal

all engine-to-vehicle

greatly

Expander

compared

3.1-20

would

DESIGN

schematic

on the

that

of the

in space.

parallel

engine

arm

by thumbscrews

This

Propellant

The performer.

tools.

placement

actuator

It is possible

BASELINE

3.4.1

Actual

gimbal

securing

changeout

ENGINE

(cont)

injector.

The

or disassembly.

an engine

expander

engine

gimbal.

clamps

or broken

making

Coordination,

The

basic

function

in Table

is 3.4-3.

preliminary once

it is at

I/I 0 m

m

>

Q C m

I,LI G. I-0

172

Table Engine

3.4-2

Operation

Sequence

Table Entry Number

Actions/State Nozzle Extension Retraction

28V DC Motor Driven Ball Screws

0

Engine Gimbal

28V DC Driven Actuators

1

Configure Engine for Chilldown

Close Open Close Open Open

2

Chilldown

02 TPA

Open Ox Main Valve (Gaseous 02 Flows Through the TPA Pump, Hex, Ox Cooled Nozzle, TPA Turbine, Injector, and Out the Engine Nozzle). Close Ox Main Valve on TPA Reaching Operating Temperature, Close Ox Igniter Valve

3

Chilldown

H2 TPA

Open H2 Igniter Valve, Open Fuel Main Valve (Gaseous Hydrogen Flows Through the LH2 Pump, Chamber and Baffle Circuits, H2 TPA Turbine, Regenerator, Injector, and Out the Engine Nozzle). Close Fuel Main Valve on Reaching the TPA Operating Temperature, Close H2 Igniter Valve.

4

Lightoff

Open Ox Main Valve, Open Fuel Main Valve, Open Igniter Valves, Actuate Igniters

5

Tank Head Idle

Modulate Fuel Idle Valve for Mixture Ratio Control, Hydrogen Proportioner Valve for Chamber/Baffle Temperature Control, Engine Temperatures Stabilized, Combustion Smooth, Igniters Off

6

Pumped

Modulate Turbine Bypass Valves Towards Closed, Close Fuel Idle Valve, Close Regen Bypass Valve, Close Hex Bypass Valve until Ox Turbine is 400°F. Accelerate TPA's to

O0

Terminated

by Limit Switches/Torque

Idle Mode

(2)

Fuel and Ox Turbine Bypass, Fuel Regen Bypass Valve, Fuel Idle Valve, Hex Bypass Valve, Ox Igniter Valve (Ignition Off)

Pumped Idle Speed, Hold by Modulating Turbine Bypass Valve, Stabilize Engine Temperature. Evaluate Health Monitor System Readings. Begin Main Tank Autogenous Pressurization.

17,44-Ta/rV2

173

Table

3.4-2

(cont.)

Table Entry 7

9

Normal Operating Range

Command Engine Thrust. Ox Turbine Bypass Valve Moves to Thrust Schedule Setting With Fuel Turbine Bypass Valve Following. Regen Bypass Valve Moves Towards Closed Position to Meet Thrust Requirement Hex Bypass Valve Modulates to Keep Ox Turbine Inlet Temp at 400°F. Hydrogen Proportioner Valve Adjusts to Keep Throat Temperature Within Limits. Mixture Ratio Trimmed by H2 Turbine Bypass Valve.

Overthrust

Command Engine Thrust With Override on Turbine Bypass Control Lower Range. Increase Mixture Ratio to 7. Ox Turbine Bypass Moves to Thrust Schedule Setting With Other Valves Following as in Normal Operation. Health Monitor System Will Reduce Thrust on a Trend Towards Unsafe Temperature. Thrust May Fluctuate as Controller Maintains Mixture Ratio With a Turbine Bypass Valve at Zero Bypass.

Normal Shutdown

Shutdown Command

Initiates Throttle

Down

to Pumped Idle Range. At Idle TPA Speeds the Fuel and Oxygen Main Valves are Closed, Turbine Bypass Valves Commanded Full Bypass, Regen Bypass and Hex Bypass Valves to Full Bypass. Idle Valve Full Open. Igniter Valves Open, Ignition On. Residual Propellant is Vented to Space Through the Nozzle. Engine Centered. 10

Normal Gimbal

28 V DC Gimbal Actuators are Activated Per Controller Instructions at Any Time During Engine Operation.

11

Operational Storage

Thermostatically Controlled Heater Power for Valves and Sensor Electronics and DC Motors, Thermal Control Power to ICHM System, Engine Centered, Nozzle Extension Retracted.

174 17.44-7a/rt/3

r if)

_ _

g

g

o

o

_._. _ _z_z_

,-.-

_

O00z

O) IJ.l

o

o

0

W

W

_

o

o

g

o

_

o

o

_.

_

_

lib oo

0

o8

8_

¢,Z) ILl

t.-

a_ 0

_0

_o_

ILl

W _J

o "

o"

ggg

_,-

ILl

000

I/J

==

111 IJJ

I/J

0

o

I/,,I

I.U

I.U

gg8ood

ILl

¢_

0

S

0

0

oB

0

_ o

C 0

im

o _ o,o_ oo (t) ILl

l_

a) o.

0 0

z lla n 0

z uJ o. 0

z

z

I_.

Q.

0

,-

£:) tu o)

z

no.

g

_.

®"

_ _-,-

oO W

oO I/J

o_ _

o _ o _o

I_I

o ooZZ_= _ _ o• oo....o g g _00_ o o

o o ._

_,:_ 0

,I

O)

o_tU I 0.

o=

Ii

0

joe_

C q) (.. 0 Q.

o

o

n

w ,-

®

_ bbbz

o

o

_z_z a

a _

_8

_

o o

OoOz

o

_

o_

_ _ a oo

a d

z

E

@ 0

0 n

z b

8 d

mzz

_mo_

-oodz

_

_

_

o

o

o

o_

z

o

_cA I--

o ir

o

o o9

o. o. o_

o _o o _

z_ 8 gg

ill

Ill

g

g

e°

tu 0. 0.. 0 I-I.-

z

8 z nO 0 z O. m

z I-

z i,,u o

0 O.

0 o

_II

175

o_

a. 0

3.4, Engine

Baseline

Design

Basic valves. pass

The valve

oxygen

The

stream

at higher

baffle

circuit

shutdown lines

and

a control

regenerator

control bypass

then

thrust

levels.

The

operating, power

balance

curves

for both

turbine

plotted

against

engine

chamber

are indicated balance

have

for controls

pressure

points

was

head

idle.

margin

power

available

adjustment chamber

jected

turbine,

pressure

(and

thrust)

3.4.3

Engine

settings

to give there

points

even

more

2000

five

are

margin

of the flow

and settings the

power

acceptable chamber

from

pumped

balance Note

available

is reached.

is available

for

helium

bypass

very

at low

the

valves.

that

at this point.

psia

and

bypass

considering

flow

and

These

the power

bypass

until

chamber

assessment

transitions

a 10% bypass

is not critical

hydrogen

interaction.

because

is positive

in the

HEX

inflection

valve

section.

turbine

of the

as the engine

bypass

turbine

just

These

by-

or required

shows

good

hydrogen

HEX

for an initial

settings.

crossed

system

3.4.2-1

bypass

are for start

through

used

will be an upward

the curves

variables

design

idle

program also

that

to increase With

some

for additional

at the top end.

Components

engine

component

for the engine

to a complete

questions

psia

purge

is very

in the valve

The

valves

the control

linearity

error

at which

to the

Discrete

(Pc < 100 psia)

TPA

thinking

The

Other

Figure

to reduce

There

curve

condition,

in the

used

percent

the other

The design

were

At 2000

set to adjust

power

pressure.

modeling.

on the

valves.

the

the cooled

is done was

with

energy

between

valve.

program

bypass

as well.

a few

curves

to tank

valves

on the plot

will

split

control

turbine

ratio.

available

is no helium

engine

thrust

two

to the OX TPA

to supplement

There

the

mixture

temperature

by the proportioner

Once

with

the engine

to hold

hydrogen

pressurization.

bypass

sets

oxygen is used

control

regenerator

valve

valve

The

is accomplished

modulating

to regulate

is controlled

valves.

(cont)

bypass

or tank

and

engine

turbine

following

is primarily

Update,

baseline.

reevaluation.

discussion During

To remain

in Section the study,

in the design

3.1.1

each it had

represents

component

the latest was

to survive

sub-

such

as: •

Can

the engine

•

Can

its function

weight

,u,r/,_7_,/3.2-3._

and

operate

if it is removed?

be assigned

complexity?

176

to another

component

at a savings

in

G) e=

°_

"_

I,IJ t,)

0 L_

rX Iii

D

m m

¢B n I

tO

0

& I

l

l

i

I

I

|

l

i

l

I

I

I

*

i

I

I

I

a

i

I

i

i

a

G) L_

% 'ssedX8

eu!qJnj.

.i

!1

177

3.4, Engine

Baseline

Design

Update,

(cont)

•

Can the design

•

Is it producible?

•

Does

it have

be changed

safety

to save

implications?

weight?

If so, what

are the

failure

modes

and effects?

•

Is it an engine improve

Some ation

If so, what

changes

are needed

to

its life?

of the potential

design

changes

that came

out of the re-evalu-

included:

1) tioned

on grounds

Combine

patible

with

either

Eliminate

tation

with

ability

Use

and

over

would

tional

a continuously

to each

preclude

the latter

stages

be difficult

Provide

valve

engine

with

valves.

The

start.

of chilldown

valve

only

massive

This was

ques-

unit.

rejected

as it was

3 or 4 positions

This may

that there

variable

dumped

one,

This was

incom-

merit

the hydro-

but requires

consul-

will be any improvement

loop

for chiUdown

in chilldown would

A preferred

have

for both

in reli-

valve.

a recirculation

circuit.

with

unit.

engine.

to be sure

use the propellant

valve

would

a stepped

designers

4)

and HEX into one

the regenerator.

the HEX bypass

the valve

or cost

may

a 10:1 or 20:1 throttling 3)

gen regenerator

the regenerator

that packaging 2)

This

life limiter?

but would

also have

option

if that can be done

add

a line and

to be positioned

is engine without

of the turbopumps.

operation severe

where

an addiits failure

at tank head

popping

idle in

or pressure

fluctuations.

5) This

was

tenance lant lines,

considered scenario. several

Rr-rlr_17_.i32_J

Add

line quick

premature The design electrical

until baseline

cable

connects/disconnects the NASA

decides

calls for simplified

connectors,

two

178

gimbal

for each on a specific engine actuators,

major engine

removal. and

component. mainFour

locking

propel-

3.4, Engine devices

Baseline

at two

through

Design

points

on the

an opening

large 6)

as adding

coils

on a ground open

fault

against

lost to effect manually

closed

erational

or fail

3.5

shutdown.

goal

progress

has

discussion technology

failure

of the OTV

engine

issues

to practice.

been

made

over

and

by a solenoid

The

will

are

open

powered

if power

valves

valving

dual

so they

is either

is can be

fail op-

TECHNOLOGIES

technology develop

program

design,

program

report

is to identify

materials,

personnel

the life of the contract a status

circuit

releases

these

engine

have

valves

that

rejected

mode.

Aerojet

be considered

shut-off

to build

was

All valves Each

processed suit.

This

circuits.

propellant

station.

plane

in a space

reliability.

possible

OF CRITICAL

to fulfill

as well

and

critical

systems

believe

a great

this goal.

as a call

the

The

to pursue

deal

of

following additional

efforts. Thrust

Chamber

Performance

is usually

stated

of this engine

as the using

mixture

ratio our

using

of 6).

reasonable

In general,

established

combustion difficult

performance

a/ 3.2_3.8

to pick may

figure

come

With

Performance specific

performance

computation

pressure

Aerojet

has

injector

energy

element chamber

--

for delivered

(at a chamber

methodology. well

Technology

Improvement

Aerojet's

484 lbf-sec/lbm

Rlrl-/D(3417.55

main

crew

be in one

applications.

electrical

in place

on the

them

3.5.1.1

will be

the

to put

3.5.1

using

The

It is also

from

technology

should

for critical

improving

latched

safe depending

engine

approaches

deliver

and

would

an astronaut

separate

system.

IDENTIFICATION

LOX/LH2

95%

the

or opened

A major

valves

necessarily

force

an engine

structure

to accommodate

by physically

within

a spring

takeout

redundant

without

powered

(cont)

thrust

enough

Add

complexity

actuating

Update,

designs

length

up additional

impulse

a good

such

psia,

record

release

179

The

Aerojet will,

of 1200:1,

of test stand

verification

coaxial nearly

(ERE)

engine version

theoretically,

ratio

Isp in the future,

from:

area

efficiencies

as the very

(Isp).

rocket

methodology

of 2000

will yield engine

for a LOX/LH2

greater

and

of lsp than

and

swirl

coaxial,

100%

ERE.

Although

real

gains

a

in vehicle

a it

3.5, Identification

of Critical •

Technologies,

Reduction

(cont)

of initial

chilldown

propellant

dumping

prior

to engine

start. Maintenance

•

of injector

throttle

range

injector

elements).

Active

Propellant

residuals

vehicle

specific

engines Reduction ating

cooled

their

envelope

strong,

engine

Long

attitude

and

are

as a black circulating

control

system

up by the main

weight.

mixture

down.

ratio over

Weight

engine

nominal)

and

in developing a high

oper-

(MR _>7) overthrust

thrust,

premium with

The

that

is a function

selection,

materials

(2000

psia

can be

of the

num-

chamber and

pres-

using

or greater)

engine.

engine

service

transportation

cost

maintenance,

using

cost $100,000

per hour.

lant expanded

in

to the tanks.

pumped

in thrust

materials

pressure

lent of higher

RI'r/_17.._./3.2-3.8

size

a high

is a continuing

life for a space

tank

the engines

hydrogen

and back

gas supply

size

lightweight

chamber

most

reduction

the engine

the liquid

chamber

(5 to 25% increase

There

with

to minimize

when

utilize

with

physical

ber of components, sure.

periods

of a LOX/LH2

includes

to keep

systems

the entire

operation.

in engine

capability used

space

pressure

during

over

range

and consequent

would

the efficiency a high

(APM)

during

to deep

the regen

Improving

off at the lower

boiloff

One approach

through

efficiency

mission.

impulse

radiator

by using

is a drop

in hydrogen

shutdown.

release

Management

on each

Reduction

body

(there

energy

life.

Every

that usually

based specific

people

engine. impulse

Long

into space

that of the item.

suits

formidable

has been costs

service

as it reduces

the actual

180

carried

exceeds

in space

These

to support

item

mission.

has

In space

estimated

mandate

a large

to

a long

life can be the equivathe amount

of propel-

3.5, Identification

of Critical

Technologies,

Engine 10:1, 20:1, or whatever Aerojet

design)

weight

reduction

consideration

without

an added

transient

state-of-the-art

engine

for mission

indicates

operating

detail

unit

time).

must

One

range

The

in Section

3.1.1,

in 0.3 seconds

should

be adequate

be verified

engine

performance

change

This

range:

(for the

is a small other

As noted

is the

requires

There

20:1.

a 10% thrust

but

Thrust

in Section

Chamber

3.1.1.

Design

It embodies

Microchannel better

by the vehicle

by a factor

NiCr)

of three

over

NiCo

program.

Optimized

I-triplet

best

element

This

a

is possible for present

prime

contractors

needed

of the element

based and water The sured

/ 3.2-3.8

The

hydrogen flow success

and

the

The

mass

of the element

by an increase

nickel

chamber, high

various

distributions modification

in chamber

OTV

has

selected

injector

life.

are

ratio patterns

better

in Figure between

were

plotted program

is not

prompted

to assure

shown

It

face.

chamber

release

(See

a wall 3.5-3

the oxygen

verified

by

Figure

3.5-4).

will

I-

the

elements.

of the injector

C.4

the

as having

as a long

Task

strength

electroform in the

energy

patterns

combinations

to increase

Aerojet

to the momentum

streams.

and

Three

for construction)

under

resulting

on adjustments

closeout.

4 to 8 inches

The

and

specimen.

of all LOX/LH2

a short

in

strain

A for dimensions

3.5-2

ERE.

thermal

demonstrated

ERE within

high

compatibility.

for reduced

elements.

potential

to get

modification

been

is defined

features:

a conventional

has

baselining

design

are recommended

(see Figure

N100%

allows

test

injector

performance

can attain

throat

electroformed

and

The

TCA

See Appendix

technology

triplet

in the

bimetal

NiCo,

Aerojet

state-of-the-art

for a machined

Codeposited

closeout.

The

several

transfer.

3.5.1

(NiMn,

i

design

heat

Figure

Rlrl'/DO417.&%

from

parameters.

controllers,

throttle

a regenerator.

is decreased per

aspects.

suitability. 3.5.1.2

some

A high

and

(A thrust

analysis

engine

valve

range

rate

performance

requires.

control

as throttle

violating

has two

the mission

is throttle

preliminary

throttling

(cont)

be mea-

"t'cT i

Channel Depth CO

J

Scale 40X

LAND

tLT

tLB

0.01094

0.00875

®

0.01109

0.01156

0.01125

0.01125

®

0.0109

0.01078

NOMINAL

Figure

3.5-1.

CHANNEL

0.011

Microchannel

Test

tor

NOMINAL

+_.002

Specimen

182

Based

tc_. B

0.010

on 7.5K Thrust

Level

+ .002

Design

! Z

0

0

U

_j_

D

WIJ

e_

u.

n-

Z

e-

E

<s20

m

0

O. e"

m

0 e_

tu Z

_

om

I-m

WU

|

¢_" mm DI

0

e_

im

&l.

183

Predicted

Mass

Flow Patterns

for Fuel and Oxidizer

Elements

Left Baffle Element

Right Baffle Element

Center Element Wall Element

3K Element

OX

Figure

3.5-3.

Modified

Fuel

Injector

184

Element

Patterns

SCMMLt

•

1

...,>: DISTRIBUTION

I I

1

7OO

5 O0

I I

0 O0

g

i

I

g

g

g

g

!

i

8

iA\

j t6" ')

I

-4

%

SCI']P1LL1

OX

HASS

DISTRIBUTION

Figure 3.5-4. Water Flow Patterns For I-Triplet Element 185

0 0 o

3.5, Identification

of Critical

Technologies,

Hydrogen

(cont)

cooled

the regen

cooled

baffles.

By dividing

chamber

and baffle

face area is increased dance

of energy

Chamber

possible in throttle

The major the use of improved copper sion

alloys

of finely

dispersed

lent machining copper,

the SCM

for a comparison). show

some

promise.

other

noble

metals

of choice work

should

should work required

concentrate on low

friction

at elevated

done

very

size

life

psia

may

be

and an increase

will come

of a new hardened

GLIDCOP

have

compared

temperatures

(See

successful Platinum

may

cost.

reducing

Figure

3.5-5 that

platinum

well

be the

The emphasis

engine

excelto pure

composites

in adapting

of

by the inclu-

alloys

matrix

in

family

conductivity

metal

material

life while

and

material

in future

weight.

Technology an oxygen

including

turbopump

seal leakage treatment

bearing

system

with

bearing

operation

with

400°F

would

wear

50°F oxygen

system.

and improving

to minimize

186 RPT / DOll 17,&'.'Sa / 3.2 -3.8

in copper

the raw

engine

on reducing

in a hydrostatic

loss in thermal

5 and 6) and a hydrostatic

surface

pressure.

technology

dispersion These

applications.

despite

has tested

work

in engine

in TCA

copper,

properties

Turbopump

turbopump

little

chamber

be on increasing

sur-

is an abun-

to 3000

an investigation

particles.

has also been

plates

gas (See Reference

using

is also being

to thrust

has begun

oxide

very

Aerojet

Aerojet

the oxygen

company

mechanical

for the baffle

3.5.2

drive

Aerojet

aluminum

Work

reduction

metal

by hot section

of capability

area for improvements

characteristics,

and excellent

consequent

only

between

range.

materials.

from

extension

result

into chamber

limited

flow

the heated

The practical

is effectively

A further with

plates

that can be turned

pressure

capabilities.

by 80%.

the hydrogen

oxygen overall

during

be worthwhile.

the

turbine

A continuation turbine efficiency. rubbing

drive

of gas

Minor starts

¢'I

:I

¢_

0

0

0

0 C')

© O_

>.

z=

_.

_

_

0 ell

_=_ _;

z_ _ 0

0

A U. o

.=

x.

o O_

E )-

0 0

L-

iD

ii

C_

!s_l 'ql6ueJls ple!A 187

3.5, Identification

of Critical

Technologies,

The hydrogen the-art

uses

hydrostatic

Figure

3.5-6

for a representative

six stage

hydrogen

1989 and

early

1990.

uses

a conventional

with

a hydrostatic

200,000 could with

rpm,

the

same

problems

at the

chemical

transfer

The CTP

current program

3.5.3

system.

system

greater

using

limits

reached

exchanger advanced with

engine

the dual

for NASP.

Heat

Exchanger

twofold:

A bonus ponents.

is the ability Nearly

_'r/L_7_./3.2.3.8

any variant

spool

program

in late

but the XLR-134

of this critical

TPA

technology

psid.

cycle

could

turbopump

in the

7.5K

speeds

static

capability

TPA

design

assembly

Various

seal

hydrogen

for engines

to verify

rotating

so that

expander

heat

exchanger

stage

with

Engine.

cycle

seal

of

systems

could

be assessed

identify

the design

needed

by the

to 50K lbf thrust

for any

projected

technology

using

provides

(HEX)

3.1.1

about

range.

engine

for the

The value range

the stream

of these greatly

expander

188

leaving

with

heat

of the

regenerator

used

components increased

needed

around

engine

alloys

of a platelet

is by using

to operate

the hydrogen

bypass cycle

copper

has a discussion

65% of the energy

operation

of a hydrogen

testing

and the hydrogen

and throttle

heat from

the engine

current

Section

engine.

can be reduced

to trim

engine

practical,

ft/sec,

be demonstrated

Main

excess

A dual

Technology

expander

with

XLR-134

state-of-

(see

operation.

also be designed

to 8000

heat exchanger

and 2) the HEX

turbopump

hardware

platelet

Shuttle

size

TPA

qualification

LOX/GH2

1) chamber

regenerator, oxygen

the flight

propellant

of 7500

for any

also

on the Space

2000

rise

could

The Aerojet has now

than

program

technology

speed

proven

should

This demonstration

propulsion

configuration

A demonstration

pump

pressure

TPA.

was

considered

is needed.

a hydrogen

operating

and

for subcritical

concept

for test

critical

design)

currently spool

spool

pump

basic

a dual

The dual

tip speeds

be designed

with

for the Air Force

bearing

and

systems

tested

ball bearing

at turbine

technology

was

Such capability

turbopump

bearing

pump

(cont)

the

turbopump.

these

can benefit

comfrom

one

a

o or)

i i

0

I 0

Or)

I

¢/_ b--

E ::3

!

I,

0 s_

:3

IC 0 t-

"0 >,

.Ira "1-

0 0

m

_6

tJ)

c-

,m i.-

CO

e,m

1,1.

0

U3 ._o

u._--

o >, -To

_E

._o .,,.,, t-

LLO.

O

0

189

3.5, Identification

of these

heat

exchangers.

technology

the weight

alloy.

oxygen

side

gas phase

under

mechanical as a two

be greatly

conditions

production

engine

design

operation.

Figure

2-1.1

reliability.

would

and

leave

tested

LeRC

(Ref:

Oxygen

to operate

use of oxygen Cooling

fabrication

weight

over

concerning

oxygen

70%

beryllium

geometry

as it reverts

transfer

rates.

generators coefficient

fabricating,

R_r/r_17s5,/32-38

oxygen

for

to the

Without

the oxygen

would

for the HEX

would

main

valve

have

of the proposed development

design

dual

at pressures

actuating

has a are pre-

system

be adequate

coils

system.

expected

based

valves issues

would

with

separate

by the controller. turbine

bypass

in normal

engine

on current

for

A loss-of-power

commanded

an oxygen

propor-

task for the 7.5K lbf thrust

for a man-rated

have

on several

engine

control

valve

models.

Nozzle

the oxygen as a coolant Pressure

and testing

The

be verified

propellant

This is a critical

None

in the last position would

Cooled

is dependent

that 28 volt dc motors

needed

would

of High

balance

preliminary

that each

400°F

response

NASA-TM-81503). designing,

of their

of a channel

heat

transfer

3.1.1.10).

2) confirmed

the valves

with

thermal

configuration.

demonstration

Control

needed The

A critical

Technology

The valve

proposes

The dual

sion.

of liquid

turbulence

heat

and Section

for the redundancy

3.5.5

energy

questions

and high

channel

and

in the desired

Aerojet

technology

fabricated

control

and

supplies

critical

the component

is the selection

velocity

Valve

(see Reference

valves.

failure

reduce

mixing

and the overall

Engine

history

metering

type

weight.

be a demonstration

are also serious

turbulent

Proportioner

(see

cision

power

and

reduced.

valves

these

would

for the HEX

designed-in

stream,

complexity

to be resolved.

of high

from

phase

3.5.4

tioner

issue

that will assure

mixing

added

could

There

that need

A design the oxygen

is some

substitution

in copper

with

(cont)

heat exchanger

This material

compatibility

exit

Technologies,

The price

for the platelet

in beryllium. from

of Critical

expander

cycle

turbopump has been Rocket technology an oxygen

190

extracts

from

about

the oxygen

demonstrated Thrust

cooled

nozzle

cooled

in chamber

Chambers

that should

35% of the thermal

With

be reduced

nozzle tests

Liquid

exten-

at NASAOxygen,

to practise

on a test bed engine.

by

A

3.5, Identification

of Critical

3.5.6

Technologies,

Extendible/Retractable

Nozzle

The NASA-MSFC materials

and joint

the contract

had

successful,

for an OTV

not been

a follow-on

long

seal

for leak

free

3.5.7

should

nozzle

is required

Integrated

work

(Task

E.7 to Contract

engine.

Additional

E.7 task

work.

3.5.8

under

This cycles.

to operate

and

Health

recognized

the OTV

NAS

3-23772)

development

engine addresses

is likely

based

Synergisms

The

for vehicle

list of areas

of vehicle

several

of the items.

The

collaboration

of engine/vehicle

gimbal

design,

information. carried number place.

The

out

the main

proposed

to the extent

of critical Also,

thrust

of propellant

with

the efficiency

tank

engine vehicle

intended

technology

items of both

and

take-out

system,

and

flight

station

of the study.

191

for the

engine

when design

aerobrake

such

display not that

a collaboration

(synergy)

be door

has

believes

for

attitude

and

interfacing

the

in

also

of the

Aerojet

CTP

designers

should

control

task

from

given

design,

with

recent

coming

designers

contractor

and

most

system

integration

will be identified vehicle

The

technology

improvement

structure

prime/engine at the start

as a critical

propulsion

pressurization,

improved.

Rr,r,'_lTa_,,'32-3j

so

with

the requirements

performance

collaboration

system

technology

(ICHM)

on recommendations

the

control

an assessment

is a critical

program.

in depth

requires

optimization

include

System

LeRC

technology

3.5.1

design,

the

an aerobrake

Monitoring

by NASA

Engine/Vehicle

to include

This

As of this writing

the development

should

through

Section

extended

and

the

required.

Control

has been

numerous

to demonstrate

nozzle.

to assess

for reliability.

retraction

a program

it is funded

be considered

after

and

This task

Assuming

operation

engine

extension

continued

awarded.

to fund

extendible/retractable

mechanism(s)

as the CTP

alternate

is expected

engine

program

deployment/retraction of the

(cont)

will be

been a takes

4.0

CONLUSIONS 4.1.

AND

CONCLUSIONS

Three goal

RECOMMENDATIONS

conclusions

of producing

a new

stand

high

out due

to their

performance

importance

LOX/LH2

in meeting

rocket

engine

the ultimate

for space

transfer

applications: 1) the-art ment

The NASA

in LOX/LH2 program

LeRC-sponsored

rocket

engine

can be started

work

technology

at any time

has materially

to the

with

point

relatively

advanced

the state-of-

an engine

develop-

where

low

technical

and schedule

risks.

expander RL-10

2)

The Aerojet-developed

cycle

engine

engine

margins,

in terms

prime

The

define

engine

primes

are

efforts

is rapidly

working

this engine

on study

capability.

in response

to President

such

moving

The

and

program

as thrust,

a real

there

vehicle

to focus

envelope,

Bush's

space

showed

on the

initiative

that

vehicle

basis.

application; interest

by the

thermal

on a regular

is national

needs

propellant

represented

and

are needed

towards

dual

technology

monitoring

interchanges

contracts;

is needed.

the 1960

of the

operational

parameters

contractor

version

impulse,

and health

design

contractor/engine development

where

control

flow

over

specific

collaborative

vital

engine

advance

of delivered

and modern 3)

helped

is a major

parallel

This

vehicle

in specific

real

near-term

results

include:

missions

application. Conclusions

expander

4)

The

cycle

improved

ments,

and

expanded

where

long

engine

5) purge

specific

system.

to engine

development

life and

A major

of the parallel

thermal

the

margins,

operating

benefit

flow

reduced

envelope.

operating

This is a major

technology

of the benefit

and

version

of the original

pump

This may

flexibility

study

output well

Aerojet

pressure

require-

cycle

of choice

be the

are emphasized.

Aerojet

cycle

in the context

is the

elimination

of the

overall

the

operating

of any vehicle

helium

design

requirements. 6) baseline

design

RZr/_175S,/4.0-5.0

A throttle with

only

range minor

of 20:1 is well adjustment

within to the

192

turbopump

capability

design

points.

of the

4.1,Conclusions, (cont)

baseline gold

7)

High

engine

cycle

plating

platinum

New

for better

8)

Engine

should

individual

chamber

specific

impulse

improvements scenario

copper

plates

chamber

should

is very

close

to propulsion

needs

versus

limit

Materials

by the

will

require

be constructed

to realistic

system

to be developed

complete

component

use,

and

engine

changeout

answered

The

sensor

of places

TPAs

12)

Integrated

a fruitful

area

industry

a of

upper

specific

to assess

removal.

despite

technology

hold 4.2

great

exchange

impulse.

the

need

Access

any

limits.

quick

and

for engine

connect/dis-

and

a decade

with

promise

alloys

versus

for continued and

research

platinum

lightweight

life as a limiter bearing

health

work behind

area

alloys

and

for thrust

composite

materials

service

life may

metals.

hydrostatic

control

to rocket

Work

copper

of turbopump with

a major

to replace

for continued

is at least

13)

for heat

question

by using

remains

the GLIDCOP

beryllium

in a number

11)

remain

technology

In particular

can be used

to engine

systems.

monitoring

system

for the forseeable other

be

industries

development

future. in adapting

The

will

rocket

new

propul-

electronic

and

engines.

the MLETS in reducing

code

showed

development

that

analytical

risk,

but

they

tools

of this sophisti-

are time

intensive.

RECOMMENDATIONS

The needed

address

of the

accommodated

margins.

changeout will

Protection

is readily

features.

development.

was

emphasize

constraints

> 7) operation

MR = 10, the baffle

thermal

A maintenance

10)

cation

(MR

Above

delivered

component

packaging

ratio

components.

mils.

alloy

9)

sion

and

of I 1/2

work

connect

mixture

study

between

work engine

was and

left incomplete vehicle

contractors.

this need:

_'r/_;_,,/4.o-s.o

in those

193

areas The

first

where three

a cooperative recommendations

effort

4.2, Recommendations,

1) coordination head

The study among

start,

included

(cont)

should

vehicle

gimballing

be continued

contractors

requirements,

with

and and

more

the engine

thrust

specific

tasks

contractors.

takeout

structure

requiring

Such

topics

design

should

as tank be

for resolution.

2) as possible

as it has engine

cooperative

tasks 3)

included

A maintenance

scenario design

for the vehicle

A focused

for these

implications. contractors

task to improve

in any follow-up

engines

study.

This should working

with

propulsion

Again,

needs

be included engine

system

this requires

to be developed

vehicle

as soon

as one

of the

contractors.

specific

impulse

and engine

should

contractor

collaboration.

4) include

Continued

a number

5) bearings,

the industry

technology

program

should

tasks.

program

design,

for a hydrogen

and a variety

ICHM

In particular, to handle

RP'r/D0417.55./4.0_.0

of an engine

should

Test Bed Engine

7)

be adapted

rotor

the OTV engine

development

A demonstration

Development

throughout

opment.

under

TPA

of seal designs

using

hydrostatic

would

resolve

code

adaptable

concerns

component.

6)

Expander

of materials

subcritical

on this key

work

be funded.

(This

state

analysis

may

be a product

for use

of the Advanced

program).

capability some

steady

should

techniques

this propulsion

be improved of artificial system.

] 94

by continued intelligence

research

decision

and devel-

making

should

be

5.0

REFERENCES

Sutton,

°

George

P., Rocket

Propulsion

Wiley

Interscience,

U.S.A.,

80027-9,

Hayden,

.

Final

Warren

Report",

ATC

NASA

Lewis

23772,

Schneider,

°

Judy,

Aerojet Lewis

Final

Brannam, Technology Aerojet

Final

Report,

TechSystems, Lewis

Holtzman,

Research

W.A.,

Technology

1988.

Verification", C.4, CR 4387,

NASA

TechSystems,

Final

Program,

NAS

Engine and

NAS

Ohio,

"Integrated

3-23772,

Orbit

Task

Research

Hydrostatic

Task

Ohio,

B.4 and

August

Turbopump

Oxygen

Testing",

B.7, CR 185262

October

1989.

Control

and

Transfer

Bearing

Task

Oxygen

E.3, CR 182122, Center,

Turbopump

3-23772,

Ambient

3-23772,

Report",

Oxygen

Cleveland,

Vehicle

Cleveland,

II Task

Lewis

3-

Oxidation/Reduction/

and

NAS

Center,

Contract

Phase

for NASA

Fabrication,

Vol. II: Nitrogen

Hayden,

TechSystems

NAS

1990.

Engine

Contract

Transfer

W.R.

Management,

Task

Aerojet

Vehicle

Research

Center,

and

Design

1987.

NASA

NASA

October

Gas-Side

87-12-NLRFC1,

Vol I: Design,

R.J., et al, "Orbital

0-471-

Contract

and

January

Transfer

Lewis

Ohio,

3-23772,

Chamber

TechSystems, NASA

ISBN

Preliminary

CR XXXXX,

NAS Ohio,

August

Report,

Aerojet

B.6, CR 185175,

°

ATC

Engine

Cleveland,

Contract

P.S., et al, "Orbital

Testing",

NASA

Report

edition,

Design,Fabrication,

Cleveland,

California,

Technology,

°

NASA

"OTV NASA

Center,

E., "Combustion

IR&D

Buckmann,

Ralph,

Injector

Center,

Jerrold,

Sacramento,

°

Research

fifth

1986.

2459-34-1,

"Baffled

Research

Corrosion",

Sabiers,

Report

TechSystems,

Franklin,

.

R., and

Elements,

Health

Rocket

Engine

Aerojet

Cleveland,

Ohio,

October

1988.

.

Schneider,

J. and

Chamber

Assembly

NASA Research

RPT/DO4175Sa/4.0-5.0

Contract Center,

Hayden, Hot

NAS

W., "Orbital Fire

Test

3-23772-C.2,

Cleveland,

Ohio,

Transfer

Program", Aerojet September

195

Vehicle

Interim TechSystems 1988.

3000

Report,

lbf Thrust CR-182145,

for NASA

Lewis

1989.

5.0, References,

.

(cont)

Anderson,

R.E.; M. Murphy;

Oxygen/Hydrogen Aerojet

May 10.

H.G.,

Kobayashi,

A.C.

Study",

12.

Klueh,

R.L., "Bubbles

13.

Klueh,

R.L. and

Brandt,

Bair,

Center,

Alabama,

NASA

Science

Combustion Contract

August

Liquid

Observations

and Cooling

NAS

8-36167,

1989.

and Technology,

"Some

With

1980.

Chamber

Alabama,

Chambers

pp.

5-13,

1969.

on Hydrogen

of the Metallurgical

Society

of AIME,

1968.

in Crystals",

Scientific

American,

Vol.

218, No.

1968.

Configuration

I, Contract Marshall

9984:0055G,

"Main

Transactions

February

W., "Channeling

Ostrander,

MSFC,

W.W.,

E.V., and Schindler,

Engine

_,r/_lT__s,/40,_ 0

Alloys",

Earth-to-Orbit

Flight

Thrust

June

TechSystems,

in Solids",

of Silver",

March

D.B.,

NASA

Mullins,

242, pp. 237-243,

90-98,

16.

Space

Rocket

NASA-LeRC,

Aerojet

1370-F-1,

Embrittlement

NASA

Marshal

of

on Copper

for the Advanced

Pressure

and Morgan,

No.

Vol.

of High

NASA-TM-81503,

Report

15.

Conference,

"Cooling

Technology

14.

paper

Environment

"Effects

1986.

Price,

Vol.

J.A. van Kleeck,

Chamber

Technical

Technology

Oxygen," 11.

Rousar;

Combustion

TechSystems,

Propulsion

D.C.

N.,

NAS

C.M.,

Studies", 8-36876

Spaceflight "Chug

5 March

Stability

"Space

Transportation

Final Study and NAS

Center,

Report,

8-36855,

Alabama,

of 20K CPT

1990.

196

Main Executive

Aerojet August

Engine,"

and

Booster

Summary,

TechSystems

for

1989. Engineering

- EAR

3, pp.

APPENDIX DETAILED

ENGINE

A

THERMAL

CONTENTS

Introduction Design

Methodology

A.

Regeneratively

B.

LOX/LH2

C.

Oxygen

Boundary

Cooled HEX

Cooled

Chamber

and Hydrogen

and

Regenerator

Nozzle

Conditions

Discussion A.

Geometry

B.

Regeneratively

C.

Baffles

D.

Heat

E.

Oxygen

A.5

Conclusions

A.6

References

A.7

Nomenclature

I_t417.55a-App

A

Allowables

Exchanger Cooled

Cooled

- All Components Chamber

and Regenerator Nozzle

A" 1

Baffles

ANALYSIS

A.1

INTRODUCTION

The material Engineering 1989.

contained

Analysis

Report:

It documents

dual

propellant

lead

analyst The

A.2

expander

engine

was

cooled

nozzle,

and independent

and

hydraulic

regeneratively

then

flow

10 May

version was

of the

also

the

design. 50K lbf thrust.

The 25K

point.

defines

of the

when

required

temperature

hot hydrogen

entering

the regenerator

inlet

temperature

turbine

defines

to operate

HEX

defines can attain.

the regenerator

is greater

is sized

to give

the assumed

baffle

A-2

an outlet

turbine

inlet temperature.

chamber inlet.

The

condition

to

minus

the

transfer inlet

in the

condition.

the maximum If the resulting than

the solution

temperature

were

to the turbine.

the H2 inlet

exit/regenerator

Finally,

and

cooled

pressure

the regenerator

is evaluated.

the

predictions

of energy

then

to the

temperature

the oxygen

the amount

were over

condition

at the turbine

inlet

Estimates thrust

components, an inlet

thermal

constant

the regenerative

to the regenerator

the regenerator

equals

defines

the hot hydrogen

the cold H2 entering

side which

the

considered

and hydraulic and

their

for each

the hydrogen

Next,

as sepa-

approach.

conditions

first.

chamber,

not be analyzed

analytical

the hydrogen

across

(thrust

to consider

on the other

of the baffle

in the nozzle

This defines

had

and were

Because

defines

energy

available

Group

the hydrogen

The total

components could

discharge

conducted

temperature

drop

inlet

pump

and its thermal

and pressure

inlet temperature,

A

assumed

the chamber

temperature

hydrogen

l)041735a-App

was

was

major

led to an iterative

is not dependent

chamber

through

The hot hydrogen

converged

9985:0234,

to the 20K design

that the analysis

range.

drop

heat exchanger.

baffle

thrust

This

of energy

temperature

thermal

and baffles)

Engineering

chamber

mean

heat exchanger.

possible

This

determined.

The temperature

amount

similar

but rather

and

of the

The mixed

pressure

the

ratio

to the baffle

made.

TCA

TechSystems

author,

for 20K, 35K, and

HEX,

and hydrogen

cooled

characterization

was

entities,

the AT Systems

mixture

pressure

Study",

the report's

that the five

H2 regenerator,

oxygen

from

the entire

indicated

interdependence.

assumed

obtained

Aerojet

of the parallel

K. Dommer,

to be very

analysis

rate

of the

Feasibility

analysis

designs

from

METHODOLOGY

A preliminary oxygen

derived

Thrust

OTV engine

evaluated

assumed

was

thermal

engine.

lbf thrust

analysis

High

in-depth

cycle

for the 7.5K

DESIGN

"OTV

the first

thermal

lbf thrust

in this appendix

the assumed is considered

on the cold

A.2, Design

Methodology,

Table gives

A-I gives

chamber

conditions

the analytical

design

cooled

to predict

of the analysis.

and A-V

CHAMBER

chamber

their hydrogen

acteristics.

The parameters

ysis.

A-II defines

Table

A-III, A-W,

COOLED

The regeneratively code

Tables

for the start

Table

summarize

the

A-II

inlet

components.

REGENERATIVELY

computer

assumptions

parameters.

for all the

A.

(cont)

defined

and baffles

pressure

in Table

the parameters

AND

BAFFLES

are modeled

drop

and bulk

temperature

constant

throughout

A-I are held

and their values

using

which

are varied

the SCALE rise

char-

the anal-

for each

thrust

level.

A preliminary ture

and pressure

the regenerator

the H2 inlet

O2/H2

heat exchanger

ponents

is interelated.

estimated

initially.

baffle

evaluated

was A-1

thrust cooled

geometry

is maintained assumed

Figures

of H2 inlet

evaluated,

chamber

baffle

drop

temperathrough

pressure

is assumed

analysis

preceded

the

but the operation

of these

com-

temperature

to the baffle

and hydraulic

temperature.

an optimum

is determined

is then held

cooling

channel

and hydraulic

channel

thickness,

has to be

characterization

The

of the

trends

are shown

in

channel

geometry

profile

through

for the MR = 6 operating

constant design

and the thermal is determined

predictions

configuration

channel

for this study.

back-side

The

baffle

proportionally

at the

three

A-9 and

A-10.

_17.._,-^pp^

thermal

discharge

The

work,

of the H2 inlet

the initial

The pressure

the H2 pump as well.

inlet

condition.

and

hydraulic

at mixture

of the chamber

are

ratios

shown

charequal

in

and A-8.

to increase

evaluated

level

profile

The baffle wall

chamber.

of hydrogen

A-6.

to 5 and 7. The thermal

gas-side

estimates

regenerator

An estimate

of the chamber

A-7

cooled

to the baffle

as a function

the regeneratively

Figures

provided

therefore,

pressure

As a result,

At each

acteristics

small;

and hydrogen

through

This channel

balance

for the regeneratively is typically

to represent

Figure

power

mixture

ratios

(channel wall

and land

thickness)

cross-sectional

with

thrust,

however.

and

thrust

levels.

A-3

widths,

of the 7.5K area

channel lbf design

and number The baffle

The predictions

depth, (Ref.

of channels

is

characteristics are shown

1)

are in

TABLE

A-I

OTV RELATED

I.

ANALYTICAL REGENERATIVELY

TO

Operating

Conditions

and

Pressure / Baffle Ratio

(psia) Flow

Chamber Chamber Mixture

II.

ASSUMPTIONS COOLED CHAMBER

Regeneratively

Cooled

Flow

2000 5,

of throat / throat radius Convergence angle (degree) H2 inlet pressure (psla) Maximum channel width in nozzle (in) Channel width in throat (in) Land width in throat (in) Land width in Barrel (in) "" Gas-Side Wall Thickness at Throat (in) Gas-Side Wall Thickness in Nozzle (in) Gas-Slde Wall Thickness in Barrel (in) Back-Side Wall Thickness (in) Channel roughness (in)

Gas-Side Close-Out Channel Channel Channel Channel Channel Channel

Cooled

6,

7

Chamberz

of Curvature upstream throat / throat radius of Curvature downstream

Regeneratlvely

50/50

Split

Maximum Aspect Ratio Radius of Curvature upstream of convergent section / throat radius

III.

BAFFLES

Splits:

Actual Contraction Ratio Geometric Contraction Ratio Inlet Area Ratio Gas-Slde Wall and Land Material Close-Out Material

Radius of Radius

AND

I0 15.3 28 Narloy Ni-Co i0

2.0 2.0 2.0 40.0 5500 0.030 0.011 0.010 0.025 0.020 0.060 0.060 0.020 60.E-06

Baffles:

Wall and Land Materlal Material width (inch) land width (inch) wall thickness (inch) backside thickness (inch) depth (inch) roughness (in) A-4

Pt-ZGS Pt-ZGS O. 020 0.020 0.025 0. 020 O. I00 60.E-06

Z

TABLE

A-II

REGENERATIVELY COOLED CHAMBER AND BAFFLE GEOMETRY AND PROPELLANT FLOW RATE ASSUMPTIONS VERSUS THRUST

THRUST

Throat

area

Barrel Barrel L'

(In**2)

Dlameter Length

(in) (inch)

(inch)

Baffle

Length

Baffle

Cross-Sectlonal

Area

(inch)

(In**2)

Total Propellant Flow Rate (Ib/s)

(Ibf)

2 OK

35K

5 OK

4.89

8.54

12.18

9.76

12.89

15.40

6.31

7.49

7.01

12

15

16

4.86

6.45

7.01

26

45

64

41.32

72.17

A-5

102.88

TABLE

REGENERATIVELY

A-Ill

COOLED CHAMBER, BAFFLE, COOLANT INLET CONDITIONS THRUST 20

A.

CHAMBER Inlet

Inlet

- coolant

Temperature MR" 5,6,7

B.

BAFFLE Inlet

Inlet

Flow MR MR MR -

NOZZLE Inlet

Inlet

Coolant

-

-

50

K

5500

5500

5500

90

90

90

3.44 2.95 2.58

6.01 5.16 4.51

8.57 7.35 6.43

5500

5500

5500

H2

Rate 5 6 7

503 499 497

459 459 459

430 430 430

3.44 2.95 2.58

6.01 5.16 4.51

8.57 7.35 6.43

4862 4849 4844

4908 4897 4888

4906 4895 4887

610

610

610

34.43 35.42 36.16

60.14 61.86 63.15

85.73 88.18 90.02

(1b/s)

-

02

(psia)

Temperature MR" 5,6,7 Flow MR-5 MR-6 MR-7

K

(°R)

coolant

Pressure MR-5 MR-6 MR-7

35

(Ib/s)

Temperature MR - 5 MR6 MR - 7 Flow MRMR MR-

(Ibf)

(°R)

Pressure (psia) MR m 5,6,7

Coolant

Co

Rate 5 6 7

- coolant

NOZZLE

- H2

Pressure (psia) MR - 5,6,7

Coolant

K

AND

Rate

(°R)

(ib/s)

A-6

TABLE A-IV

HEX

INLET

CONDITION8

THRUST "

.....

20K IUmililI

_

02

_ _

L

INLtTTn(P_TURZ

INLET

PRESSURE

GAS

INLET

GAS

INLET

TEMPERATURE

PRESSURE

188.

188.

188.

5168.

5168.

5168.

805. 841. 874.

869. 905. 940.

868. 903. 940.

3229. 3287. 3295.

2971. 3096. 3180.

2601. 2829. 2980.

41.32

72.17

102.9

5.17 4.43 3.87

9 • 02 7.73 6.77

12.86 11.02 9.65

36.15 37.19 37.96

63.15 64.95 66.31

(°R),

(PSIA)

MR-5 MR'6 MR'7

TOTAL _LOW _TZ

50K

(PSIA)

MR'5 MR'6 MR-7 H2

I

II

(°R)

MR'S,6,&7 H2

I

--

MRm5,6,&7 02

35K

_

(LSS/SZC)

H2 FLOW RATZ WITH 25t sypASs (LBS/SZC) MR-5 lqR-6 MR-7

02 rLOW RATX (LBS/SZC) (lO5% OF rLOW_Tm) MR'5 MR-6 MR-7

A-7

90.02 92.59 94.52

TABLE A-V

REGENERATOR

INLET

CONDITIONS

THRUST 20K

H2

HOT

INLET

TEMPERATURE

HOT

INLET

PRESSURE

COOL

INLET

TEMPERATURE

COOL

INLET

PRESSURE

HOT

FLOW

RATE

FLOW FLOW

642. 633. 624.

641. 631. 624.

3096. 3190. 3221.

2838. 3001. 3109.

2448. 2725. 2904.

90.

90.

90.

5500.

5500.

5500.

6.89 5.90 5.17

12.03 10.31 9.02

17.15 14.70 12.86

3.58 3.10 2.69

6.32 5.41 4.74

9.00 7.72 6.75

(LBS/SEC) MR-5 MRI6 MRs7

H2 COOL (50% OF

579. 572. 564.

(PSIA)

MR-5,6,7

H2

50K

(°R)

MR_5,6,7 H2

_

(PSIA)

MR-5 MR-6 MR-7 H2

35K

(oR)

MR-5 MR-6 MR-7 H2

_

RATE RATE)

(LBS/SEC)

MR-5 MR-6 MRs7

A-8

b,.

Lr)

II r_

0 r-O _.00

r_ rY

m

I I I I I

r_

"0

_.

2 .o o S

.-

'

E

E

Y

_

(1.) _

• I-"

0 O_

-

_

-

II

(1)

12) f" -O_

u_

i X •

Em lid

.E

F--

l

-

,_

-

_

-

_

_

__ ii

jlllllllllIillllllllllllillllliiiiiilllliiiiiiiii

-O O

I

O

0

0

0

C)

0 0 p,.

c;

o

(:5

o

oO LO

LO LO

_LO

oJ (.0

(aaJ6ap)

aJn],0JadLua]

A-9

D}la(]

O

d O tD

II

II

II

i i I I I

I I I I I

I I

0 -0 o3

2 -0 -0 I_,

rc

E

0

i-

V

Q) t,-.

-

Z3

ra C

'I-

-o 2 .cO(]') O_

E

0 ..I Q L_

Q I._

_

.._

_

_

-o° _c if-)

c_ O4

2: Q

m

rn

Q I.-

01 a n

-0 '0 I

I

!

I

I

I

!

1

1

I

i

I

I

I

I

I

I

I

I

i

i

I

0

0

Ln

L._ rO

L_ C',,I

A-IO

I

I

I

I

I

I

I

0

0

(p!sd) _Jnss_Jd

I

011_C]

urb

LI.

II

II II

0 -0 -GO

I I I I I

I I I I I I I i I I -0 -

I,--

-g 4:3

r_ 0 C" "0--

_ --• m I1)

U} el'-

_

• m

lJ i" I

Q 0 If) C_

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

1

i

I

0 0

i

0

0

0

0

d

d

d

d

0 O_

_ O0

0 O0

LID t'--

(ggJ6gp)

gjnloJgdLug I o_iga

A-11

L_

cD

r'-.

II

II

0 -0

ry

I I I I I

I I I I I

I I I I I

o o

rr 0

E Q

v

I-

(D

"

t,--

£L

E

0 ...I Q

-

I--

Q i_ D.

_

©

-r

-o

(-"

m

--O

-u3

Q m

m

,f ,II ..C H,m

IJ.

I

0 u'3

I

I

I

I

I

l

I

I

|

I

I

I

I

I

0 _

I

I

I

I

I

I

I

I

I

I

I

0 uO

(p!sd)

I

I

I

I

I

I

0 IF)

aJnssaJd

A-12

0 iaC]

|

I

I

I

I

I

I

I

-0 0 0 If) Cq

II

II II

I

II

I I I I

II II II II

0

n

..D

0 u')

o-E -0--

:D 5,,,.

¢I-'-

_: e in

[

I

I | I I I I I |

[

i

I I

I i , , '

i I I I

' ! I , , '

, I | '

o o

I I , I | , I

o

o

0

o

o

o 0 0

o ,.f') O_

0 0 O_

0 if) 00

o 0 cx_

(eeJBep)

eJnl0aedwel

A-13

011eC!

_L

i.r)

_

r-,,

It

It

II

o -o

r_

r'r

rY"

-00

t I t I i

t I I I I

I I I I I

2 _.=

-o °

i

"E" 0

v

m

L

:CO

E P

3::

Q)

,..10

Q-

m

E

"

-

(1)

_

-

__.

2

"

_

"1-

-

_

®

Y 0 Lr)

-0

C

. L.f')

m

5") o L_

(.}--

I&.

0 0 I

I

t

I

I

I

I

I

I

i

i

I

I

I

I

I

I

I

I

i

t

i

I

t

I

I

t

t

I

i

I

I

i

i

I

I

i

i

i

i

t

I

I

1

I

o

0

0

0

0

0 r'-,,.

0 _0

0 L.O

0 _

0

(p!sd)

@Jnss@Jd o_1@@

A-14

0 0 _D

0

0 0

0 tO

_3

(3_03G)

3WDIVW3d_N31 A-15

V1930

0 0

LO

0 I-

_2 c_

II.

I

I

I

I

' I 0 0 0,4

0 0 _t-

0 0 O0

(p!sd)

3_JnSS3_Jd

A-16

VIq3@

0

II II r_

I I I I I

I I I I I

! o fOfOm

_>=

OE ol..

|

o lit

(aeJBep)

3_I"I.LWl3dI,_3J.

A-17

VI-130

L.r')

(4D

P"

II

II

II

[E

rY

rY

I I I I I

I I I I I

I I I I I

_Lr) _r3

_0 -Lo

m

r_

0

_>>

ur3

D I OG.

| m

Q L-m

U.

E

IIIIIIIIIII1

IIII

IIIIIII111111

lill

Illll

II

O

O

O

O

O

Lr)

_"

tO

C,J

,--

(p!sd) 3EINSS3EIcI V173C]

A-18

A.2,

Design

Methodology,

(cont)

OXYGEN/HYDROGEN

B°

HYDROGEN

the

H2/O2 uses

pressure

heat

exchanger

and

and

Design

AND

fluid

and

exchange

and the H2/H2

to solve

heated

heat exchanger

the energy

(HEX)

inlet conditions

program.

counterflow

technique

for the cooled fluid

state

drops

an iterative

EXCHANGER

HYDROGEN/

REGENERATOR

The steady mine

HEAT

the desired

parameters

between

hydrogen

the steady

streams.

model,

state

outlet

for the HEX

and

is used

the working

fluids

regenerator.

energy

The channel fluid

HEXSS,

conditions

hydrogen

in the

The HEXSS

and momentum

geometry,

to deter-

equations

the core

length,

are provided regenerator

code

and

to run

the

are given

in Table

A-VI.

The the same

as those

of the two similar

heat exchanger

of the 7.5K lbf/thrust

components

mass

The total

and the regenerator

are assumed

flux per channel

number

temperature

are determined

level.

The thermal

and hydraulic

thrust

level

are shown

in Figures

in Figures

A-15 The

02

a preliminary estimate in the drop

through

power

oxygen

inlet

through loss

balance

using

Because

The

iterated

until

erator

is greater

cooled

of the

the

chamber

points

value

assumed

provided

of the baffle

downstream

the corresponding

pressure

on the turbine

design

the initial

than

the required for each

MR range

fluid thrust

at each

predictions

inlet

from turbine

to the HEX

for the

are shown

loss used

the HEX and

inlet

determined is based

the energy

is based

temperature the baffle

the

on an available

turbine.

group. and

temperature

to the hot side of the hydrogen This

assures

This

power

is unknown inlet

in

on the pressure

in the preliminary

by the turbomachinery

inlet temperature.

A - 19

were

and a 37% loss across

components,

temperature

baffle

to the HEX

temperature

to run the oxygen

H2 inlet

is based

the performance

_, 7.ss,-^n, ^

required

the regeneratively

turbopump

a

component.

The predictions

The regenerator

and pressure

The 02 outlet

nozzle.

percentage

A-14.

temperature

balance.

cooled

through

to maintain

condition

for the complete

area

A-18.

of the total energy

pressure

effects

liquid

A-11

to obtain

geometry.

to be

The cross-sectional

7.5K lbf engine

needed

that baseline

are assumed

however,

corresponding

characteristics using

2).

thrust,

at the MR = 6 operating

determined

geometry

(Ref.

with

and the core length

outlet

HEX

design

to increase

as that of their

of channels

are then

engine

channel

that there

is regenis

TABLE

A-VI

DESIGN PARAMETERS AND ASSUMPTIONS FOR THE H2/O2 HEX AND H2/H2 REGENERATOR

HEX

WALL

ZrCu

MATERIAL

CHANNEL

DEPTH

02

.03

IN.

CHANNEL

DEPTH

H2

•04

IN.

CHANNEL

WIDTH

.06

IN.

WALL LIKE

THICKNESS CHANNELS

BETWEEN

WALL THICKNESS HOT AND COLD

BETWEEN CHANNELS

OXYGEN

CRITICAL

REGENERATOR

PRESSURE

WALL

MATERIAL

.043

IN.

.036

IN.

730.

PSIA

ZrCu

CHANNEL

DEPTH

H2

COOL

.02

IN.

CHANNEL

DEPTH

H2

HOT

.04

IN.

CHANNEL

WIDTH

WALL LIKE

THICKNESS CHANNELS

BETWEEN

WALL THICKNESS HOT AND COLD

BETWEEN CHANNELS

HYDOGEN

ASSUME

CRITICAL

PRESSURE

INCOMPRESSIBLE

FLOW

A-20

.056

IN.

.043

IN.

.034

IN.

188.

PSIA

_ ¢0 P,-

II II II n- n,, n,-

o o

o o

o

d

NIVO

A-21

_r_ nf,

/,D

illillll

0

0

0

0

0

c5

c_

c5

0

0

(_o)

SS07

3_n.LV_3dF_3J.

A-22

_0

|

0 O

(5

O

O

O

O

0

c_

d

d

d

O

F_

F3

_'N

(015d)

SSOq 3anss3ad

A-23

t'N

¢Ii II:

I_ I_

_o m

.oA : 2 -

D

•

: _

_,

un

_ N

nl

123 iiiiiiiiiiiiiiiiiii

Jllllllllllll

iliiiiiiii

IIIIIIIIIIIIIIII

_

0

0

0

0

0

0

0

c_

c$

6

o

o

o

o

to)

_I"

CN

0

00

¢0

_I"

(alSd)

SSOq 3anss3ad

A-24

t.v,o::rv. 0

.o / /

/ 0

(l:Jo)8807 31_nlvl_3dlAl31

A-25

I_ i'I,,.I_

/

ii

/

I

I

m

I,I,.

(19o)NIV9 3wn1Yw3d_31

A-26

Ix :E _0 u*)

/

/ / I

0 0

(5 0

0

0

0

0

_

o

c_

LrJ

(OISd) SS07 3WnSS3Wd

A-27

0

0

0

0

d

c_

(5

o

0 e,,l

(OlSd)

£SOq

3_lnSS3;_a

A-28

0 0

A.2,

Design

sufficient

Methodology, energy

available

the pump

to result

evaluated

and sized

charge

(cont) in the hot H2 from

in the assumed

conditions

to achieve

baffle

to transfer

inlet temperature.

the assumed

are assumed

the HEX

baffle

to be the inlet

inlet

to the cold H2 from

The regenerator temperature.

conditions

is then

The H2 pump

to the cold

side

dis-

of the

regenerator.

The are given

HEX

and hydrogen

in Table

C.

OXYGEN

COOLED

Table

the bulk

A-VIII

using

are held

inlet

estimate

of the

in an assumed Since

power

are fairly

energy

oxygen

the pressure

neglected

The

02 outlet

ation

with

for all components.

(assuming

design

option

mum

channel

nel geometry teristics shown

geometry

assuming

profile

depth

of 0.500

profile

constant

in Figures

analysis

ratios A-20

inches

A-19

through

equal and

thrust

pressure

low

thrust

A-29

the

modified

determined.

drop

nozzle

and The

this

predictions

is

the

a maxi-

cooling

hydraulic

vari-

depth

using

to reflect

Holding

to be

channel

of 10) is determined

the thermal

are

pressure

cooled

results

enthalpies

is assumed

the

an

levels.

effects

maximum

the nozzle. level,

oxygen

pressure

The

is then

7 are then

A-21.

and

tur-

with

of 610 R for all thrust

summarizes

ratio

oxygen

of the nozzle

to the nozzle

level.

aspect

throughout

to 5 and

ratio

in

requirement

the regeneratively

The profile

at a given

defined

portion

of this study,

Figure

program.

(area

the coolant

of the required

cooled

is typically

pressure

allowable

oxygen

parameters

temperature

nozzle

inlet

an MR = 6 for each

of the SCALE

the

range

profile

a maximum

at mixture

t_lT._,-AppA

condition.

with

to predict

estimates

to the

oxygen

to the HEX

determined

code

in the oxygen

in the

equal

A channel

Coupling

the nozzle

to pressures

cooled

The

provided

termperature

through

in the estimate.

thrust

in the

the analysis.

balance

available

inlet

drop

insensitive

throughout

and pressure.

total

computer

rise characteristics.

constant

temperature

that is regeneratively

the SCALE

temperature

A preliminary bine

as determined

NOZZLE

of the nozzle

of 28 to 635) is modeled and

geometries

A-VII.

The region

drop

regenerator

chancharacare

Table

A-VII

H2/O2 Hex Geometry

Thrust

Core Length

20K Ibf

35K Ibf

50K Ibf

17.6

15.0

15.1

58.8

87.1

125.

(in.)

Core Weight (Ibm) Total No. of Channels

02

966

1679

2392

Total No. of Channels

H2

966

1679

2392

H2/H2

Regenerator

Geometry

Thrust 20K Ibf

35K Ibf

50K Ibf

15.4

8.6

7.4

41.08

39.73

48.2

Core Length (in.) Core Weight

(Ibm)

Total No. of Channels

H2 - Hot

861

1491

2100

Total No. of Channels

H2 - Cold

861

1491

2100

17.44-7a/rl/4

A-

30

_kr)

_kr) ,@

:

v rO__ Or) D n" I F-

U.

e n

IL

u') illllllllllllllllllllllll

%.--,

0

0

0

0 0 O0

0 0 _l"

0

(p!sd)

3_Jnss3_ld A-31

V173C]

Lr]

(D

r_-

II

II

II

rY

rY

rY

:E

:E

:E _0

I I I I I

I I I I I

I I I I I

m

E

o -0

d I

m

i

_6 0 I

0 0 LQ r_

I

i

1

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

I

I

I

I

1

I

I

i

I

I

1

I

I

0

0

0

0

0

0

0

0

0

_0

rO

04

C_

0 u_ CN

(aaJ6ap)

3wn±vw3d ]± A-32

v±]]0

I"',- _0

L_

II

II

il

I I I I I

I I I I I

I I I i I

_0

2o

z_ 0

.:--° n,.

:

-1--

-__o1_

_c_

: o

J"'l

0 0 09

I

!

!

I

I

I

I

I

|

0 0 tO

I

i

I

I

I

I

I

I

I

J

|

I

i

I

!

0 0 _t-

I

I

I

I

i

I

I

0 0 t_l

(p!sd) 3_nss3ad

A-33

vJ_fl30

I

I

I

t

I

I

' 0

I

0 0

$ o

A.3

BOUNDARY

CONDITIONS

The Bartz

equation

regeneratively

is used

cooled

to evaluate

chamber

the gas-side

08

equal

are defined

in Section

to the average

property

and

is smaller

heat fluxes

near

that gaseous

The injector actual

the 7.5K same should

1.0.

states

than

barrel

The higher

of 1.0 throughout

in the 7.5K

two

fore end heat

from

the

fluxes

from

This

created

the 7.5K

for the

for all thrust

with

due

element

The complete

as that of

remain

the

value

of 1.0

barrel

contraction

of

Cg value

ratio,

the

to overpower

characteristics. cooled

ratio. value

a constant

large

the

Because

contraction

reflects

profile

when

same

barrel

to the generic

sufficiently

3 asserts

injector.

to the smaller

profile

the

gas velocity.

flux should

that due to the smaller

design.

fluxes

to be the

to the end of the regeneratively

lbf engine

heat

the generic

and compared

by the injector

however, Reference

engine

of the fore end heat in Cg over

in the present

chamber

the OTV

All

program.

levels

to be similar.

is assumed

has become

temperature.

lbf design,

than nominal

The resulting

infers

wall

isentropic

study

at the film temperature

72 computer

in the 7.5K

higher

lbf engine

gas velocity

aft end of the barrel

maintained

assumed

increase

is assumed.

isentropic

high

the TRAN

in this study

are determined

the barrel.

one-dimensional

from

are assumed

yield

used

the relative

of the

the adiabatic

used

the magnitude

Cg's

and

that this is the case

design,

are evaluated

the one dimensional

velocity

However,

be lower

often

exceeds

injection

lbf engine

The corrected

ratio

17 to 1 value

injectors

designer

as well.

temperature

the fore end of the barrel

velocity

GH2

All properties

contraction

the near

hydrogen

GH2 injection

the

than

coefficient

04 Prf"

Cg(z)Ref"

data are obtained

The 15.3 to I chamber study

0.026

A.7.

of the wall

temperature

transfer

and baffles: Nuf=

Symbols

heat

the

The Cg profile nozzle

is shown

is in Figure

22. The Hess heat the

transfer heat

and

Kunz

coefficients

exchanger

and

correlation for hydrogen the regenerator.

Nuf=

0.0208

(Ref. 4) is used in the baffles, This

to describe

the regeneratively

relationship

Ref 0.8 Prf0.4. (1.+

the forced

0.01457

is:

g_°b)

gbOw

r_lT.ss,-^p_^

A-34

convection

cooled

chamber,

A-

U'1 -r,-) _D

o rl--

o

& mi -0

|

o u. 0 Jllll

0 0

I

l

!

I

l

I

I

I

I

I

I

I

I

I

I

C) 0 0

0 0 0")

0 0 aO

C> 0 r---

_-

c_

c_

6

83

A-35

A.3,

Boundary

Conditions,

The supercritical transfer

coefficients

NUb=

(cont)

LOX correlation of oxygen

Nuref*

(pb/[3

(Ref. 5) used

w)"'5. (kb/kw)5*

chamber

and

calculated

pressure

baffle

drop

relates

the convective

heat

is:

(Cp/Cp_

the

for a cooling exit

static

(P/Pc_

2/3.

,

NUref=.0025*Re

The coolant

to evaluate

"'2. [1 + 2/(1/d)]

where

04

b Pr b"

channel

pressure

within

to the

the regeneratively

inlet

stagnation

cooled

pressure.

It is

as:

A Ps = A Pinlet + A Pfric + A Pdyn

One half of the dynamic flow

loss is assumed

at the channel

inlet

to account

for

contraction.

The flow two

head

through

components.

compared

the HEX

and regenerator

This method

to a compressible

The friction

factor

results pressure

calculation

is assumed

in a conservative drop

to be incompressible

pressure

for the

loss estimate

when

assumption.

(Ref. 6) used

in all friction

pressure

loss evaluations

is:

f:00[15 A roughness fles, and

of 60.0

nozzle

and

E-6 inches

is assumed

a roughness

of 125.0

Re;'l for the regeneratively E-6 inches

is assumed

cooled for the

chamber, HEX

baf-

and

regenerator. A.4

DISCUSSION A.

GEOMETRY

The maximum bending

_17._,.^pp ^

loads

applied

ALLOWABLES

channel

- ALL COMPONENTS

width

to fully-elastic

to wall

hot walls

A-36

thickness

for all components

and is described

as:

considers

A.4,

Discussion,

(cont)

The maximum considers

tensile

B.

channel

loads

applied

in the regeneratively III describes The straddle

mill

rather

tion in increasing sizes

chamber

in Table

barrel

channel

decreased

pressure

drop.

to center

mum

used

width

at the throat values

width

is maintained to 0.010

with

allowable

in Tables

A-I and

channels

A-II.

Table

A-

to the chamber. to be fabricated

This method to minimize used

allows

pressure

in the 7.5K

at the throat

used

inches

flow

channel

in the throat

area

from

with

a smooth loss.

a transi-

The

lbf engine

and the

The increase

in the barrel

was

decreased

increased

channel

design

land width

increase

area

from

0.040

Because

the earlier

design,

the 0.011

with

in the

ratio

the total

used

(channel

velocity

thrust

geometrically

saves

pressure

inches

used

A-37

drop.

in the

bar-

to decrease

inches

channel widths

rather

than

width

center

are decreased

previously.

Coupling

depth/channel

decrease

levels

widths

design

to 0.011

the land

inches

and flow

are, in general,

lbf engine

is increased

design.

For the three

in flow

and the land

in the 7.5K

10 to I aspect

depth. region

was

at the throat

from

the maximum

overcooled.

r_lr.ss.-^pl, ^

cutter.

widths

in the 7.5K lbf engine

in a channel

channels

as:

of the coolant

are assumed

on those

and land

those

The channel

correspondingly

results

width

channel

width

from

spacing

assumption

and is calculated

descriptions

conditions

in the chamber

A-I are based

for all components

section.

rel were

inches

hot walls

used

CHAMBER

inlet

a constant

of the channel

The

0.010

than

criteria

are summarized

hydrogen

channels

width

and general

chamber

or decreasing

indicated

the exception

COOLED

cooled

cooling

elastic

assumptions

the assumed

to land

to fully

REGENERATIVELY

The geometric

width

this

width)

limit

of 21% at the maxi-

evaluated,

the design

limited

and

The

maximum

in the 7.5K lbf design

of the

are therefore land to 0.025

width inches.

A.4, Discussion, (cont) Becausethe total number the decrease

in barrel

approximately

27%.

through

the barrel,

engine

however

must

the

only

is determined

in an increase control

total

flow

constraint

applied

structural

support.

provide

in channel

dictates

coolant

by the throat

the

area

width

channel

is still

of

depth

larger

to the minimum

profile

with land

Fabrication

geometry,

and

the 7.5K

width

weight

lbf

definition

issues

were

considered.

split

balance

system

occuring was

assumed.

schematic upstream

evaluated

This

pressure

flow

path

compared

to the series The bulk

each

the

level

the

thrust

level.

the

square

root

temperature

flow

C.

hydrogen

reduced

and the

occurs

of thrust. as thrust

rise

channel

amount

because

in Figures

throat

flowrate

and

chamber

in hydrogen

increases

is the

result.

assumption

and

general

the

thrust

A-8,

as

and

trends

respectively.

remain

channel

drop

chamber

versus

constant

increases

diameters

pressure

was

cycle.

throat

per

psia

of 90 R related

A-7 and

at the

power

of 5500

through

drop

a 50/50

preliminary

expander

geometry

with

temperature

pressure

of coolant the

The

pressure

and

version

pressure

inlet

propellant

are shown

the

A-23).

hydrogen

dual

flow

discharge

the delta

of the

An increase

geometric

the

regeneratively

inlet

conditions

The

NASA-Z maximum

cooled

initial current

baffle

are

less

material

effort

started in the

wall

with earlier

1050°F

was

description in Tables

for

with

increase

with

a decrease

in bulk

of the coolant A-I and

A-II.

channels

The

in

hydrogen

A-III.

assumption

temperature.

to the turbine than

summarized

in Table

baffle

assumed

gas-side

temperature

temperature

baffle

are summarized

The

_,7._,-^p_^

ratio

(Figure

a low

chamber

parallel

a pump

temperature

cooled

This

to the

BAFFLES

The

design.

with

version

evaluated,

rise

and

significantly

contraction

thrust

changed

of the regenerator

coupled

for the regeneratively Because

was

parametrically

to the new

inlet

results

temperature

The

the lands

The flow

width

Wall

assumption.

is that not

land

of channels

differed

a platinum

alloy

work.

The

In the

previous

high

at the coolant

plane

A-38

that

(Pt-ZGS)

NASA-Z

(1000°R/540°F) exit

from

baffle

study, and

the

of the 7.5K baffle

is limited required

maintaining

of the baffle

rather

was

lbf than

the

to a 1050°F hydrogen a wall

difficult.

_.II .I" /

•,_ e" C

Jill iD • II O) •H - _-. r,. c II

3

\ _-'_;

:,.,i

i- I_. _!:_-,OD II U

_ o) _

_o \

I t_ I_

II li 0.1--

°_

4) "£ ILl

_

i,i _iIiI

,,>i_-I _i

_.<:> i,,,:_ i,i_

P->=J

C-

i

>'0

e_4

_ A-39

A.4,

Discussion,

Platinum

(cont)

was

chosen

conductivity A later

(similar

evaluation

operations

up

showed

cooling

lbf engine

tip and

to Nickel)

the corner

chamber

diameter,

whichever

desirable

in order

to maintain

higher 1/2

heat the barrel At the

20 and

baffle

length.

levels

evaluated.

The affects

for both

the

Section

A.4B,

to the

relationship

baffle

has

the

baffle

are

assumed constant

which

differentiates

baffle

length

The

turbine

The

sum

(963 and

the

increases

area

levels. heated

as thrust

increases

shows

the resulting

exposure

to

lengths

for the

MR = 6 condition.

the

is very

68 degrees

A-40

inlet

bulk

which,

temperature

with

The

flux

rise

thrust

the

in the

parameter levels.

rise in both

Because

of the fluid. components.

50K lbf thrust

the resulting

to the limited

due

channel

the

thrust

for the 35 and

in

channels

becomes

in

H2 through

per

temperature

temperature

thrust

area

down

temperature

R due

the

As discussed

different

condition

than

defines three

of coolant

then

the bulk

close

the

H2 mass

for the

of the delta

considera-

to thrust.

length

area

stability

surface

goes

the number

baffle

turbine

shows

H2 flowrate

longer

for the

heated

the

A baffle

criteria

A-25

At the 20K lbf thrust

by approximately

is not

Figure

up, so does

temperature

section

to prevent

the chamber

The

the barrel

stability

total

the

or 1/2

of combustion

and

evaluated; study.

section.

baffle

of the

section

the

surface

goes

region

affects

thrust

The

of these

A-26

convergent

combustion

and

with

is a function

Figure

the

Because

temperature

area.

and

through

inlet

is lower

pp A

directly

total

960 R, respectively).

shape

resulting

however.

for all thrust

temperature

190417.$5a-^

trend,

to scale

to 2000°F).

to those

parametric

baffle

length

surface

only

than

the

the

rise

is the

longer

the baffle

temperature

to be identical

A baffle

temperature.

of heated

(up

thermal

for all engine

of the barrel

because

levels,

and

baffle

length

considered

illustrates

inlet

opposite

remains

not

chamber

the bulk

the

with

of the baffle

turbine

trends

associated

A-24

selection

the

as either

35K lbf thrust

Figure

be used

for the present

a rectangular

was

of its high

capability

could

are assumed

of the

is smaller.

conditions

temperature

baffles

not analyzed

is taken

diameter

tions.

turn,

length

flux

the core

are

baffle

alloy

in the baffle

however,

maximum

wall

because

_< 10.0.

channels

regions

for this effort

its high

copper

ratios

design;

material

and

that

to mixture

The 7.5K

as the baseline

turbine heated

as a function

levels inlet

surface of thrust.

r"

('_)

_u") uF)

"0_ • ) (:L)

•_

..E)

_u')

:3 L_ r_Q

,y,, 'v

LO _0_O0 D " _

Q I

I _--

|

:3 m

IA.

-aO

_O I

0

i

I

I

i

I

I

I

I

I

I

0

I

I

I

I

I

i

I

f

I

!

I

I

I

I

0

(u!) HION37

I

i

i

i

i

i

0

37JJ'v'8

A-41

i

i

i

i

i

i

i

i

I

0

1,1 ._1

bJV

i,

W<

<

133

(eeJBep)

V173Ci

3_IFI/V_I3d_31

A-42

i

_uO U'D

_

4.--

JO

o

i

u') (/3 ID Of I I--

_

Q I

C m

Q e" m L_

LO C_ |

Q L_

,m

U.

_0 rl

ii

ii

iiilll

il

ii

iiii

ii

ii

Ii

iii

i ii

iii

ii

ii

ii

ii

ii

ii

ii

i i iiii

ii

iii

0

0

0

0

0

0

0

6

6

c5

c5

c5

6

c_

03

_D

_-

C,_

0

O0

<0 o3

(Wo) 3WFIIVWHd_31

137NI

A-43

3NI8WI71

A.4,

Discussion,

(cont)

The

pressure

to those

of the chamber,

assumed effect

drop

to achieve

also

however.

the same

is negligible

regeneratively

exit pressure

temperature

cooled

with

An orifice

for the pressure

The bulk

increases

baffle

thrust.

directly

The losses

are

downstream

compared

of the baffle

as that of the chamber.

range

small

is

The Joule-Thompson

of this study.

rise and pressure

drop

for the MR = 5-7 range

versus

thrust

are shown

trends

in Figures

for the

A-9

and A-10,

respectively. D.

HEAT

The and

cold

EXCHANGER

divergences

H2 in the

because

the total

maintain

number

drop

conditions

the

range

50/50

based by the

the

fluid

on that

assumed

turbopump

throughout

at the three

thrust

of the

the

=

HEX

hydrogen

regenerator. bypass

pp

is sized

a turbine exit

are

evaluated. directly range.

with

The

attributable

the hot This

is

thrust

to

differences

to the

from

insensitive

to pressure

to the

is approximated

the baffle

The

percent

power

HEX and

fluid

balance

for the

inlet

using

temperature design

pres-

as a

the regeneratively

of turbine

.94"(.5"T

curves

+.5*T

to vaporize inlet

the oxygen

temperature

conditions

cooled loss was supplied

lbf analysis.

prior

as the inlet

the HEX The

by:

was

inlet

) Baffle

to entering

of 860 R (max)

are used

of H2 around

in the 7.5K

is determined

A

thrust

and

or regenerator.

A -44 L,X_I 7.SSa-A

levels

to increase

levels

temperature

turbine.

in the HEX

Regeneratively Cooled chamber

A 25% bypass

determined

regenerator

thrust

the

is fairly

temperatures

02

group.

to achieve

resulting

HEX

enthalpy

Mixture

nozzle

three

is assumed

in the preliminary

T

The

at the

channels

H2 inlet

a 6% loss across

for the H2 and

channel

the hydrogen

of the

and

per

length

of this study,

mixture

chamber

flux

overall

Because sure

of fluid

mass

drop

are small

for the fluids

and

REGENERATOR

in pressure

regenerator

a similar

in pressure

AND

the

02

cooled

at a MR = 6 condition.

conditions assumed

temperature

to the based

hot

side

The of the

on the optimum

to the hot

side

of the

A.4, Discussion,

(cont) TMi×ture =

The inlet and A-V,

conditions

respectively.

.75

* TH2outHEX

+

for the HEX and

These

conditions

.25

* TH2inHE

X

the regenerator

are given

are listed

as a function

in Tables

of thrust

and

A-IV

mixture

ratio.

The geometry in Tables heat

A-VI

The delta

temperature

are shown

in Figures

A-11

loss

across

pressures

assumed

to be equal

to the pump

throughout

the thrust

perature

and temperatures

at the HEX thrust

temperature

is due to the low

ratio.

flatten

levels

A higher

out beyond

pressure a thrust

stream

conditions.

drop

and baffle

at the lower

of 35K lbf due

thrust

ratio.

condition,

levels

thrust

case

of the shows

HEX

are

the H2 temis longer

potential.

results.

than

The low

area

to hydrogen

The curves

in H2 inlet

level.

temperature

Because

surface

to the similarity

A-13

outlet

the

driving

heated

H2 sides

of the thrust

The oxygen

for the lower

are listed

Figure

and

mixture

at the 20K lbf thrust

chamber

and

as a function

ratios

for a given

to compensate

for the 02

respectively.

for all mixture

range

inlet is low

at the higher

rate

and A-12,

discharge

and the regenerator

trends

the HEX for the oxygen

The inlet

is constant

for the HEX

and A-VII.

exchanger

the pressure

and the dimensions

flow

tend

temperature

to to the

HEX. Figure ture

ratios.

stant,

at a thrust

levels

temperature) lengths head

because the low

pressure

compensates

are similar and total

H2 side

for thrusts

of 50K lbf losses

While

H2 dynamic

the HEX

pressure

are similar

inlet conditions. lower

illustrates

The H2 delta

while

thrust

A-14

from

increase.

thrust

head

(due

HEX

for the 35 and

The pressure of the HEX

is longer

than

to the higher

drops axial

for the three relatively

for the

length

two

and

mixcon-

lower

the H2

that of the 35K lbf design,

The H2 inlet

engines,

loss for the 50K lbf design

trends

H2 inlet pressure

length.

50K lbf thrust

loss

20 to 35K lbf remains

of the interaction

for the additional

pressure

pressure

due

inlet

temperature

and HEX

the dynamic

pressure

however,

is higher

and lower

the

to the lower

H2 inlet

pressure. The temperature and cool

hydrogen

respectively. perature cold

streams

losses

H2 to compensate

Lx_17_,-App A

as a function

of the regenerator

At the 20K lbf level, of approximately

and gains

500°R,

in order more

for the reduced

of the

are depicted to maintain

energy heated

A-45

in Figures

the specified

is required surface

thrust

area

to transfer

level A-15

for the and

H2 baffle from

per hydrogen

hot

A-16,

inlet

tem-

the hot to the flow

rate

in

A.4,

Discussion,

(cont)

the chamber

and baffle.

temperature

loss

As a result,

across

the temperature

gain

the hot side of the regenerator

across

the cool

are the highest

side

and

the

at the 20K lbf

condition.

The H2 pressure

drops

across

function

of the thrust

level

the inlet

temperature

and pressure

thrust

levels

the highest

and

loss of the

cool

also becomes

thrust

the driving

the velocities

attributable

to the different OXYGEN

Table

down

deviations and

Since for all

to the baffle

and

length

the

pressure

of the regenerator

loss on the

as thrust

goes

up.

are relatively in dynamic

hot H2 side

low

pressure

temperatures

Since

head

are small.

NOZZLE

summaries cooled

length

channels

pressures

respectively.

the regenerator

the pressure goes

as a

is constant

The axial

the regenerator range,

stream

regenerator

in determining

hot H2 inlet

and A-18,

from

required

and thrust

side of the regenerator

hydrogen

the inlet pressure

ratio

A-VIII

A-17

at the 20K lbf condition.

COOLED

in the regeneratively

the geometric

nozzle.

assumptions

The assumed

for the coolant

inlet conditions

channels

are listed

in

A-HI.

The cooling fabricated

with

pressure

drop.

optimized

width,

of both

channels

a straddle The

values.

channel

They

inches.

As a comparison

channel

width than

region

channel

maximum wall

allow

allowable thickness.

_17_,-^_p^

based

and gas-side

and the land aspect

ratio

widths

wall width

channel Due

depth

to the low

channel

thickness

aspect

with ratio

ratios

The channel less

than

allowable 02

inches

to represent

in the region

Strength

depth

or equal

of the chamber,

as 0.200

inches

between

the nozzle.

of 10 for the entire

velocities

do not reflect

reasonable

widths

to be

to minimize

configuration

a maximum

is set at 0.500

A-46

cutter

throughout

portion

to be as large

coolant

are assumed

width

are considered.

cooled

inches

are assumed

nozzle

than a constant

of 10 for channel

is 0.056

a maximum

cooled

on maintaining

to the hydrogen

in that region

10. Rather

rather

for the cooling

are, rather,

wall

for a maximum

(the

mill cutter

land width,

the gas-side

in the regeneratively

assumptions

allowed

total

the

of the hot H2 through the mixture

Table

level,

though

throughout

cool

H2 temperature

parameter

even

in Figures

for the cool

outlet

H2 is largest

of the regenerator

E.

are depicted

the required

at the lowest

the hot and

profile to 0.050

the maximum aspect

cooled

ratio nozzle

near the exit), a more

of

the

reasonable

of the nozzle

is

Table

A.Vill

Oxygen Cooled Nozzle Assumptions

20K

Thrust (Ibf) 35K

50K

Coolant

Inlet Area Ratio

28

28

28

Coolant

Exit Area Ratio

635

635

635

4.89

8.54

12.18

41.32

72.17

102.88

.200

.200

.200

2.0

2.0

2.0

0.5

0.5

0.5

Yes

Yes

Yes

.100

.100

.100

Throat Area (in. 2) Total Flow Rate (Ibs) Maximum

Channel

Ratio of Channel Maximum

Channel

Width (in.) Width to Land Width, max Depth (in.)

Single Bifurcation Channel

17.44-Ta/r

Width at Bifurcation

t/4

(in.)

A-47

A.4, Discussion,

affected

(cont)

by this prescription,

mum

0.500

inch

drop

related

the

seen

manner A.5

assumption

coolant

delta

in Figures

decrease

the

is negligible ratio

A-20

in bulk

surface

and

and

penalty

when

maintained and A-21,

temperature cooled

area

drop

temperature

for the regeneratively

in which

the

respectively.

The

with

increasing

and

is also

flow

to the

throughout across

chamber

rate

to the

compared

pressure

rise

propellant

related

pressure

nozzle.

the oxygen increase thrust

in is similar

attributable

scale

maxi-

with

to

to the

thrust.

CONCLUSIONS The

tions

thermal

and

determined

values.

values

(discharge

for the

for either pressure

variations.

pressure

same

predictions

pump

drop

for each

of 02 and

and

02

(90°R

should altered

5500

psia,

exit

be made

if the pump

the

reference

respectively)

to account

insensitive

can be used

for the H2 pump

to pressure

directly

and

condi-

as reference

from

are fairly

trends

H2 pump

be viewed

are

enthalpies

thermal

temperatures

and

and

02

should

or hydrogen

the H2 and the

on

component

H2 are 5168

of this study,

discharge

based

balance

the oxygen

Since

range

were

power

in pressure

pressures

for density

hydraulic

in a preliminary

Adjustments

discharge

the

and

pressure

10 to I aspect

are shown

drop

trend

depth

resulting

nozzle

pressure

channel

to a maximum The

cooled

the additional

assuming

188°R

for the

02

pump) Of the are the

five

limiting

components

evaluated,

the

components

for the system

regeneratively

cooled

chamber

delta

pressure

on the H2 and

occur

through

the regeneratively

and 02

the

HEX

sides,

respectively. Prohibitively chamber is used

when and

upstream the inlet

flow

the 50/50

flow

drop

_lT.ss.-,,pp ^

flow

(50%

rise

of the H2 to the chamber

regenerator.

through

to the

the regeneratively

approximately

A-48

2%.

and

the coolant comes

is substantially

attributable

is small,

When

the H2 to the chamber

the chamber

is predominantly

series

split

and

through

temperature

drops

by the

of the regenerator

hydrogen

versus

H2 pressure

H2 is preheated

pressure

in H2 bulk

high

flow

directly

decreased. lower

The

temperature.

cooled

chamber

cooled

50%

to the baffle)

split from higher The

is moved the

pump, density

difference

for the parallel

A.5, Conclusions,

(cont)

The H2 temperature engine exit

thrust

to the

of 35K lbf thrust.

turbine

inlet

inlet

temperature

H2 turbine but

system

power

an increase limited and

energy

in the outer currently

chamber's limited

additional The

additional of the

hydrogen

pressure

higher

the hydrogen

compared

side

to that

section.

of the cylindrical

drop

drops heat

regeneratively

effect

on the side

attempt

was The 02

out

the

hydrogen

side

of the

heat

exchanger

made

to weight-optimize

cooled

nozzle

for a single

accommodate flow

design

cooled

nozzle

the

the

is adequately

only

penalty

Mechanical

section

the collection

The

at the inlet

of the oxygen.

manifold is a small

of the

increase

L', and

the

however. by allowing drop

penalty

regenerator

is small

at thrusts

in the

would

have

a small

pressure

drop

on the

to reduce

weight.

No

regenerator. area

ratio

design

of the nozzle

can be located

only

length

section

pressure

pressure

cooled

area

the baffle

be reduced

(particularly

or hydrogen

than

for the 20K lbf

surface

case

The

HEX

a

By increasing

Additional

case,

is already

to be considered,

also be investigated

extendible/retractable

so that

nozzle.

could

delta

length

heated

sides

the

loss.

D,barrel)

could

chamber

balance.

condition,

to obtain

cylindrical

have

on both

additional

of the system

pass

and

cooled

of 35K lbf to 50K lbf); therefore,

< 1/2

the components.

exchanger

range

oxygen

would

the pump

35K lbf thrust

the baffle

the outer

at an

it is recommended

section.

and regenerator

through

level,

more

barrel

penalty

of the

50K lbf thrust

in both

exchanger

of the

provide

For the

from

is not sufficient

(L,baff

U would

be obtained

heat

of the

the hydrogen

considerations

drop

(963°R)

the 35K lbf thrust

Because

pressure

The weights

twice

value

50K lbf thrust

that

(L') be considered.

additional

could

than

or thrust

cylindrical

area

lower

pressure

by the length

surface

baffles.

to heat

stability

engines,

For the

is nearly

chamber

length

to combustion

slightly

a peak

H2 pressure

590 psid.

pressure

for any

reaches

corresponding

is only

available

balance

35K lbf thrust

The

delta

in chamber

due

to the turbine

is approximately

the pump-to-turbine If the total

on

at the inlet

of 28 and

of the

requires

away

in pressure

from drop

through-

nozzle

to

a pass-and-a-half the due

end

of the

to loss

in the

180 ° turn.

In summary, margins

the

for the dual

for the required

_17.ss,-^pp ^

parallel expander

operating

flow cycle

schematic

has

engine.

There

envelope.

A-49

significantly are no major

improved thermal

the

thermal

design

limits

is

A.6

REFERENCES

1.

Dommer, ATC

2.

3.

K.T.,

TAR

"OTV

9985:008,

Hayden,

W.R.,

Contract

NAS

Regen-Cooling

4.

Hess,

and

Sabiers

3-23772,

and

Supercritical

5.

°

Kunz,

R.G.,

and

Contract

NAS

3-20384,

Fuel

_7-_.-^pp ^

H.R.,

NAS

and 3-21030,

Engine

Design

"A Study

November, Chamber

ALRC,m

Design

Dependent

Design,"

Final

Report,

for Estimating

1987.

Convection

63-WA-205,

"Supercritical

Approach

21 Aug.

of Forced

Paper

D.C.,

Cooling

Preliminary

TAR 9980:2024,

ASME

Rousar,

Calorimeter

Combustion

Contract

ATC

Hydrogen,"

R.L.,

Preliminary

1988.

Mechanistic

Spencer,

Ewen,

tL, OTV

October

Compatibility," H.L.,

and Baffle

14 Jan. 1988.

Ito, J.I., "A Physically Thermal

Jacket

Heat

Transfer

to

1963.

Oxygen

Heat

Transfer,"

1977. and

Cooled

Investigation, 31 May

A-50

Resonator Report

1978.

No.

Design,

High

TFD 9752:0185,

Density

A.7

NOMENCLATURE Cg

Turbulent

Cp

Specific

C F

Integrated

d

inside

D

diameter

f

friction

Fry

Yield

k

Thermal

L

Length

1

Length

L'

Axial

Land

land

Nu

Nusselt

Nuref

Reference

P

pressure

Pr

Prandtl

Re

Reynolds

T

Temperature

tw

Wall

V

Velocity

W

Channel

Greek

flow

Correlation

Coefficient

Heat Average

tube

Specific

Heat

from

Tw to Tb

diameter

factor Strength conductivity

from length

Start of Heated from

injector

Tube

to Temperature

to throat

width Number Nusselt

Number

--- .0025 * Reb * Pr_

Number Number

Thickness

Width

Letters: dynamic

e

roughness

p

density

E_ZSS,-Ap_A

Pipe

viscosity

A-51

Measurement

A.7, Nomenclature,

(cont)

Subscripts: b

Evaluated

cr

Critical

dyn

Dynamic

f

Evaluated

fric

friction

H

Hydraulic

inlet

inlet

max

maximum

min

minimum

S

static

W

Evaluated

at bulk

temperature

property

at Film Temperature

at Wall

Temperature

A-52 D0417.55a-App

A

APPENDIX ADVANCED

ENGINE

CONTENTS B.1

Introduction

B.2

Power

Balance

At Mixture

Ratio

= 8

B.5

Power

Balance

At Mixture

Ratio

= 10

B.8

Power

Balance

At Mixture

Ratio

= 12

B POWER

BALANCE

B.1

INTRODUCTION

This appendix normal

operating

Variations 10, and line

range

subtask 12 were

connecting

envelope.

investigation

made such

pumps.

cooled

In general

hydrogen

TPA

points

balances

thrust point

high

temperatures

all other

copper

surfaces)

limit

before

mixture

balance

work

ratio

operation.

thrust

thrust

boundary

is set by either limit

limit

balance

flowrate

thermal design points.

B-- 2

Balances

at MR = 8,

limits

operating

cooled

chamber

for the turbolimited

is effective.

(see Appendix

(800°F

A

maximum

temperature

limitation

design

for the engine.

TPA

energy

wall

of the

the Engine

point

the oxygen

chamber

outside

supports

for the regeneratively

the oxygen

will be within

ratios

for the engine

is set by available

is both

by a detailed

at these

at three

to the maximum

temperature

MR operation

are supported

wall

A

results

mixture

the high

Another

limited

confirms

DO417.55a-App

near

marks

baffles.

energy

balance

of high

at or very

or by a design

or hydrogen

power

of MR = 6 + 1. This power

This maximum

flowrate

power

contains

and These

A) that

for the throat,

1050°F

for

J

a.

II

0

--4.

_

4"--4-

....

@

9-

• .,,:.-..+

IT D m I L

m •

_

_

'_

1::1 a

[+[

;;;.....--..

+1:1 rs

v

G OQ_O_

• ,i

0

0 I

I

l

I

I

l

I

•

II

D II

I

m --

_ :I 0

0

0

II

U

I

|

0..

10,-

I

41

c 0 tM

0

;=;o 0

--+

•

C_

_+;o

00(

_

_

oo--

z

II

I--

/

l--

I

U

_o _?. o_'io

O_

....

• IL 0

OJI

11 S| +,

+

I

i

!1

II

--

m'

I[Ir

qk qk

_Z

_

_

_v_ 00_o_

4"4 I _ _-

i + UI

0

q"

_

0

-

(_

I_

...... qO.O

• N

Q q" Ik. qN 4i _L_ 11_ aID _4 t_i * • IIl_l_l_"

II_ _1 * U)

slr_4

114_

*1)

• I,,.

lID OlD *

.

.lllq' v-.

Ill

II I

I

I

|I|O

I

t

IIIIIIII l

I

I

I

!

I

I

I

I

I

+

I

l

I

I

Ill I

|

c 0 =8mO 000_

41 C

:1--

_

0

0

O(

:

II _

+_+ F. _

c+_c:oo7

+ -

z

0

= a, -a: m i; 0

0 ,ill I: +.. I-I--'U I; 0 (..1

0

--lie 0 II

0 *,

S"

41

O. ,G In

m

o

--g •

• I

OI-J

0 4.

L

I

B-3

18 4)'-

..-;

ee_ • .1I

•

• 0

_z:

U d= ,0

0

D I O.

_

II

L

D

II

|

II

I0

I

II I

|

_

0

O000*

iilil

II II

.

¢:

:

* II

I

i

L W

II

--

0

U_ OO4( Z

+

_+

+ m cr

v_

UJ Z

Z UJ

o N

||l|#l|i

0

| •

|

--

0

Iilll

* •

L •

m

•

_.

+

_--

00_

÷

_+

Im C •

• ,V

• n,

I .-_

N

,1_

iow

:::

I m

Z(J

m O,

B-4

|

I

--+--÷

--

.........

+

I

A_A il. 41

C

D LL'_

0

Q_i_

0

Q

_

_

qk

q,

¢

O*

Jl

L

V

g

I

O

_

0

• * qr ...... v.- eV • t Q qllO Ir QO +- 0i II Q IQ qe m

I

|

I

Q

_

_Ig

IQ _o i_

g f.h. I_

0

_

0

oo_owmo_

e

Im

|

mm II

0

|

. O 0 a I' • I_ _ _ 0 • o O • • • A _ i.- O _ _. sII

|

I

|

II

+" Illillil

I

I

Illiliil

III

I

--0

_

O

I)

I

•

OOO OmO • •

lid

•4r

IX

I

I

'+"" O

• ill

•

|

I

c --

O

am 0

0

l

l

I

II

41

# W G

O

C

311:

:l

O

Oq-

•

• (._

• II

a

t

•

._>,+

--O

oo

c_

18 O 0

04[

OAm

||_,

;. o | ?. o i o +..1 .+.._ = _ | ,t •I.

+i

4"

4.

+m+

_skq)

--

A g

_ I

o

m

v

_

W v

i .

41 O *

N

O P* , * • S'L* m O

ID _r • lO I".

,o. 0

_ _) ID • "l)

_t O O

oleo ooo • o A o

o • +-

+.+

--

0

illllliiilllNNi

I

|

m

+

_

_

I

l

c --

_ •

O

O

I

II

O

Oq[

|

II

I

I

u

U,

I

+

.I.

T,

b I::

O

J_

,,-

I-

I: 0

C II

41 G 01_

O T'

+ I 1O "T'

X _ +

B-5

UI

0 w o q I:

h o

:l

,0

,..--.,

,4. ....

+

_.

÷

+

¢11 ii G.

--

0oo0_ 00_

......

_

0_00_ mO_ON__O

,pt"

lllll Io

n

io

U

|l

_!

II

II

II

0

II|||||l

I

l

|1|

|

II -_, .c

0

•

'-

--0

0

tU 0

©=--

_--_

O0

Z

c_--_--¢_0000_

_z_:_ E

[_ m ri-

i

¢LCLO vv_

0

_

"-.

,,

_..

.o

_

vv_ v

c

vv

v_

Z 0

N

_o;_

0

•

|

II

II

I||l

II

II

0 II

II

II

IlllIIII

I |

lilil_ |

I

0 c

=--

0

+

÷

•_

0

_

•

--

0, _ :l

> m ill _

II _ I--'U

Q

¢=--©_--¢_0000

O<

................

+

0 o'r (.1 ¢: o

_+

_

O

• *'0

Omw

u

o +

_

O

OI c el

_

0

4"

B-6

0

O 'g_ ..--

•

"_

_ 0 ¢ Ok

¢

I

I

÷

m

,i*

_

.I.

4"

÷

4*

m 8. vv

_OO_ _

_.-_

m O L

IIIIllI|

I |

IIINIIll

U

I

IIIm|

t

O+ I

O

in

@

O_

|

L G

I

C_ -Q

I

I

--

@ ©_

4'

÷

I,l*

÷

oo_O_Sm_

_v , Ill

....

vv m

....

b • N

OO

B|I|IIOB

B

N O

•

_'._

II|ll

• l

I

Q

4"

4" --4"

m c

•

llo •

all

N X N OoO ZO

m

÷

--4"

--+

B-7

O

4Jr @ 8 n0

p • 0 al 0_' -. 0.

--O

I

l

+

_+

÷

i+ _AA_AAA_

l0

+

°

_-

A_

¢D II L V

VV

_b

v

vv N

omo_ o_o_

_m

om,_ 0 =m U.

_ _0

1600

_0

|m|||HNIIM|tU

C_ -0 UJ

t

C_OC_C=OO

IIIIIIII #

ill

A

III

__. • o

I

I

I

_ _D

•

--

°

IIN )_

0

IIIIII

l

_

0

_=--¢_000

IIII

l

0

0 0

_0000_

+

+ A

;;._:-_

..:_)_

°" II.

,

.....

; II,

0_0_ OmO_ .,0 _._

_,_

m

Jill

llllUllJ

illilililillill

-¢

:_

_

0000

0 _[

+

b _

II b

0 m c

.Q

b 0

)-

¢ 0

o_ II -C) O[

0_

o _+

I •

G

_+

I_

ll_-"

,lh

+

J=

+

.

B-8

m •

m

ii o

o

÷

÷

n. m 18

....

m

¢

Q

Q

q* v

qh v

. vvv

m

v_

v 00_,

.....

,

'q" 5r

llmll

||m|

n

IN

l

iNIH|IINNK|N

II A "O .r.

tu

o--rw +

4[ _0

4.

_

4-

4- _ C

8.

"0

IU Z

c

ILl =) I0

(j

.......

4-

÷

g,.*o

vv_

0l

|l|l| lull

IN

0 :1 :_ 000

::1

Oo

q, v

I

0 J:

n,,

0

Q

_A_

OO_ _vqk

v

• N 'O --

............. _AA_

alll--_'q,--.

al -I- "0

b 0

8-9

v_

|

II

I

÷

--

--

+

@ AAAAAAA

iI r: _v

II

II

II

li

II

m

_

It

vw

H

N U

|

c_ --

0

mlmlNmml

_EUEI

i

• I

_

_

I

0

I

--0

c=

0 Z

I

00_

J

_ vv

......... Z W

_ • N

_,-

"0

0

--

li

il

_

_

il

il

ii

li

II

......

IO |

_ _

_._

U

_v_

l

._:

l

am II

II II

O

--0 ¢=--

4-

00(

m

°_ C

•

•

I

N _< N 000 zO

nl m

0

n

B-IO

ii

|

+

+

+

_---

Q 1 v_w

m

•

•

a

_-_00._0,

k

||llll|limmlmll

II

I

II

lii

_

lillllgl

|

G_ --0

I

I

o i i

C_ --0

W

||; _

+

+_+

+

IL

A_ II

II

¢L

Q.

v_-_v

W

i Ul

• N

0

-

• ° 0 _'_ _D *ID

l|l||lllDlllNlB

II

I|1|1111

0

|

0 •

0

e

_

_J C3C0000_ O_ n G

_

+

+

c

a

0

J_ b ::= *o l© 0 U +

_

_*

_+

_

0

¢: II

b

_E J_ t:

0 -zm-rw

•

_,_-

_(

÷

B-11

(_

APPENDIX

CHUG

STABILITY

OF OTV

20K LBF ADVANCED

CONTENTS C.1

Introduction

C.2

Technical

C.3

References

Discussion

TABLES C-1

Hot

Fire Data

(Ref.

C-2

Hot

Fire

(Ref. 2)

Data

1)

FIGURES C-1

Injection

C-2

Data

Plot

- Ox Pressure

C-3

Data

Plot

- Fuel

Velocity/Effect

Pressure

of Throttling Drop Drop

C-1 D0417.SSa-App

C

C

ENGINE

C.1

INTRODUCTION

A study instability, OTV

performed

commonly

to assess

known

the potential

as "chug"

arising

for low

from

deep

frequency

combustion

throttling

of the advanced

engine. The

engine

lbf thrust, tling

parameters

a nominal

range.

sources (2).

was

The

mixture

injector

of stability

This

ratios

was

ratio

was

over

to chug.

a GH2/GO2

pressure

of Premixed testing

tested

observed

include,

of 6, chamber

for hot-fire

type never

in the study

is composed

data

element

and

used

of 2000

I-Triplet

of this injector

a wide

range

The

evidence

system psia

elements.

from

and

20,000

a 20:1 throt-

Some

are given

of chamber

with

excellent

as references

pressures

reference

(1) and

and

mixture

(2) is especially

compelling. In the OTV

series,

only

The

injector

characterization. injector

by simply

losophy

using

for the 20,000

the injector

has

(1) program. then

the

been

verified.

stability

high.

C.2

The

understood

The P/Pc

two

ratio

The

injection

and

the

droplet

and using

velocity,

vaporization

at full Pc, but

IX'_17,SSa-App C

orifice

must level

as the engine

be inferred

frequency

the The

geometry

the previous assumes

of the

phi-

that

called

the initiation are

until

predicts

instability

the process

then

reference

rationale

rationale

and

lbf

chambers

by analytical

system

same

elements),

increased-thrust

combustion

3000

that

injector

by virtue

analytical

characterizing

size,

size

stable

variables time

If one

for the

in the

detailed

from

of the same

feed

of comthe injection

lag.

propellants,

rate.

elements.

be used

the low

combustion

liquid

derived

the propellant

important

the

undergone

to be chug

margin

between

most

size

must

that

has was

more

confidence

It is generally

In a system

stable

element

DISCUSSION

bustion.

same

(using

stability

is a coupling

chamber

demonstrated

TECHNICAL

chugging

Delta

size

chug

experimentally is very

of the

lbf injector

already

lbf chamber

for that

more

If a larger

the series,

the 7500

the type

time

lag is dependent

of injector

typical

is throttled

C- 2

element,

situation down

is that

the injection

on many

the

factors:

atomization

an injector Delta

P/Pc

length,

will

be chug

decreases,

in

C.2, Technical

Discussion,

the injection injector

velocity

becomes

(con0

decreases,

chug

and the combustion

unstable.

This is why

time

throttling

lag increases

is a great

and

concern

thus

the

for liquid

propellants.

That

is not the case

propellant and

Delta

propellant

P/Pc

for a system

ratio

temperature.

Since

to Pc, the injection

independent

of Pc.

with

the gas being

being

the actual

for the case Pc. with

range

ization

types

doublets with

Pc.

propellants,

or shear delta

mix I-triplet)

slow

is such

of fuel

the injector

face.

the injection

would

may

time

oxidizer,

under

and hence

So the propellants

first

is throttled

velocity

velocity

falls off with

increases

due

With

to atom-

some

ele-

like-on-like chugging,

if designed

for OTV

to an absolute combustion

burning

over

fiat.

to cause

incipient

slightly

by 3 percent

consideration

are already

the

that as the engine

to mixing.

lag due to mixing

be

the second

as impinging

be enough

type

cases:

lag has no component

such

linearly

a plot of injection for two

been

due

ratio

and

gas

have

time

characteristics,

the element

and

therefore

the injection

can be a component

the

lags will

on the plot),

This shows

curve

this component

as to bring

are both

to the fact that C* decreases

the

mixing

P. However,

lines

the

of mixture

rates

injector

(dashed

the combustion

but there

coaxes,

impingement

within

solely

flow

shows

OTV-type

flow,

a function

time

This

propellant,

of compressible

Otherwise,

and mass

C-1.

lines).

(liquid)

This is due

that have

a low

The

case

or vaporization,

ment

(solid

of an incompressible

gaseous

in Figure

In that situation,

lag are only

and combustion

as incompressible changes

of pressures.

With

the

density

decreasing

that

velocity

of Pc for a hypothetical

treated

For the actual

time

propellants.

the gas density

This is illustrated

as a function

gaseous

and combustion

proportional

velocity

using

when

(the

pre-

minimum. actually

they

occurs

emerge

from

injector.

There element

exist

under

data from

several hot-fire

reference

C-2.

Of particular

these

have

mixture

_lTs,-^ppc

of data on low

conditions.

(1) is given

Table

respectively).

sources

ratios

note

These as Table

in Table

are slightly

The difference

are listed

less than

the tests

with

of this type

of injector

as references

(1) and

reference

(2) is given

from

numbered

that considered

pressure

C- 3

stability

above

C-1, and data

C-2 are

in chamber

frequency

the two

122 and for OTV tests

115.

(2).

The

as both

(5.4 and 5.6,

is 106 versus

of

HOT-FIRE

For

DATA (REFERENCE

Stability Single-Element

1)

Data Injectors,

DelP Test No. 135 138 139 140 141 143 144

151 152 153 154

DelP

Pc Chamb. -3

-2

145 146 147 148 149 150

1973

-i

-3

MR

(psia)

Pc

ox

Pc

f

Stability

6.01 7.63 2.05 3.97

291 258 325 522

.182 .181 .410 .227

.208 .194 .581 .291

Stable Stable Stable Stable

4.02 0.99 3.93

103 302 303

.217 .707 .226

.278 1.159 .291

Stable Stable Stable

2.03 6.01 4.12

304 291 98

.438 .166 .221

.625 .191 .279

3.94 2.06 3.90

317 329 524

.221 .398 .218

.283 .569 .283

Stable Stable Stable Stable

5.99 4.14 6.08

290 103 91

.176 .213 .185

.203 .270 .208

1.96

105

.420

.600

C-4

Stable Stable Stable Stable Stable Stable