Report 2418-F 10 December 1982 NASA Composite … 2418-F 10 December 1982 NASA Composite Material Application To Liquid Rocket Engines (NASA-CK-170697) COMPOSITE flATEEIAL N83-15360

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Report 2418-F 10 December 1982

NASA

Composite Material Application ToLiquid Rocket Engines

(NASA-CK-170697) COMPOSITE flATEEIAL N83-15360APPLICATION TO LIQUID ROCKET ENGINES FinalEeport (Aerojet Liquid Socket Co.) 31 <4 pHC A 1 4 / M F A01 CSCL 11D Uaclas____ ________ _________________ G3/24 _ 02337 _

Final Report

by

D. C. JuddAerojet Liquid Rocket Company

Prepared For

National Aeronautics and Space AdministrationGeorge C. Marshall Space Flight Center-

Huntsville, Alabama 35812

Contract NAS 8-34623

https://ntrs.nasa.gov/search.jsp?R=19830007089 2018-06-29T01:28:27+00:00Z

ORIGINAL PAGE ISOF POOR QUALITY

IT "Report No.

2418-F2. Government Accession No. 3. Recipient's Catalog No.

4. Title and Subtitle

Composite Material Applicationto Liquid Rocket Engines

5. Report Date

December 19826. Performing Organization Code

7. Author(s)

D. C. Judd

8. Performing Organization Report No.

10. Work Unit No.9. Performing Organization Name and Address

Aerojet Liquid Rocket CompanyP.O. Box 13222Sacramento, CA 95813

11. Contract or Grant No.

NAS 8-34623

12. Sponsoring Agency Name and Address

National Aeronautics and Space AdministrationGeorge C. Marshall Space Flight CenterHuntsville. Alabama 35812

13. Type of Report and Period Covered

Contractor Report, Final14. Sponsoring Agency Code

15. Supplementary Notes

Project Manager - Dennis GosdinNASA-Marshall Space Flight CenterHuntsville, Alabama 35812

16. AbstractThe major objectives of this study were to (1) determine the extent to which

composite materials can be beneficially used in liquid rocket engines, (2) identifyadditional technology requirements, and (3) determine those areas which have thegreatest potential for return. The assessment was based primarily on weight savingsbut also considered material and fabrication costs, performance, life, and maintain-ability factors. Two baseline designs, representative of earth-to-orbit and orbit-to-orbit engine systems, were selected for analysis. All components of these base-line designs were evaluated to determine which could benefit most from fabricationwith composites. Weight savings from 50 to 80% were predicted for selected com-ponents with the substitution of reinforced plastic composite (RPC) materials formetal, and overall engine weight savings from 25 to 30% were found possible. Varioustechnology needs were identified before RPC material could be used in rocket engineapplications, and follow-on activities addressing these needs were proposed. Metallicor ceramic composites offered advantages in high-temperature or performance-drivenapplications but otherwise were not competitive to RPC on the basis of weight or cost.

17. Key Words (Suggested by Author(s)

Composite Materials, Weight ReductionAssessment, Lightweight Design, BarrierCoatings, Polymer Matrix Composites,Metal Matrix Composites, Liquid RocketEngine Components

16. Distribution Statement

Unclassified - Unlimited

19. Security Classif. (of this report)

Unclassified20. Security Classif. (of this page)

Unclassified21. No. of Pages

28022. Price"

FOREWORD

The work described herein was performed at the Aerojet Liquid Rocket

Company under NASA Contract MAS 8-34623, with Mr. Dennis Gosdin, NASA-

Marshall Spaceflight Center, as project manager. The ALRC program manager

was Mr. Roy Michel, and the project engineer was Mr. Craig Judd. ALRC

material engineering specialists for the study were Mr. George Janser and Mr.

Ed Carter.

The technical period of performance for the study was from 24 February

1982 to 30 November 1982.

The following individuals contributed significantly to this report:

Fred Fischietto (Structural Analysis)

Ralph Shultz (Drafting)

Craig Judd (Project Engineering)

George Janser (Materials Analysis)

Ed Carter (Materials Analysis)

Chuck O'Brien (Engine Systems)

PRECEDING PAGE BLANK NOT FElivlED

m

TABLE OF CONTENTS

I. Summary 1A. Study Objectives and Scope 1B. Results and Conclusions 3

II. Introduction 12A. Background 12B. Purpose and Scope 13C. Approach 13D. Program Schedule and Major Mileposts 15

III. Task I - Baseline Engine Configurations 17

A. Objectives 17

B. Engine Selections 17

C. Component Requirements Data Sheets 42

IV. Task II - Component Assessment and Identification 48

A. Objectives 48

B. Composite Properties and Fabrication Processes 48

C. Methodology and Evaluation Form 54

D. Component Selection 54

V. Task III - Conceptual Design Assessment 65

A. Objectives 65

B. Stress Analysis 65

C. Final Drawings and Weight Estimates 65

D. Outside Vendor Design Input Concerning Reinforced 73Plastic Composites

E. Outside Vendor Design Input Concerning Metal 81Matrix Composites

VI. Task IV - Cr1tical1ty Ranking of Technology Needs 84

A. Objectives 84

B. Technology Needs 84

C. Criticality Ranking of Technology Needs 92

TABLE OF CONTENTS (cont.)

Page

VII. Task V - Recommended Tasks 93

A. Objective 93

B. Recommendations 93C. Program Plans for the Selected Components 94

VIII. Conclusions and Recommendations 119A. Conclusions 119

B. Recommendations 120References 121

Appendices

A. Component Requirement Forms A-l

B. Reinforced Plastic Composite Properties B-l

C. Metal Matrix Composite Properties C-l

D. Task II Evaluation Forms D-l

E. Vendor Trip Memos E-lF. Technology Needs Definition Forms F-l

LIST OF TABLES

Table No. Page

I Technology Needs Cross-Reference Chart 5

II Advanced Expander Cycle Engine Weight Data 29III LOX/LCH4 Cycle C Engine Weight Data 43

IV Task II Evaluation Form 55

V Task II Selected Components 63VI Component Metallic Weight Versus Composite Weight 79

VII Outside Vendors Consulted 80VIII Plastic Matrix Versus Metal Matrix Comparison Chart 83

IX Technology Needs Listing 85

X Design and Analysis Steps for Recommended Parts 94

VI

LIST OF FIGURES

Figure No. Page

1 Overall Study Program Summary 2

2 OTV Valve Housing 7

3 LCH4 TPA Impeller Housing 8

4 OTV Nozzle Extension Shaft 9

5 OTV Skirt Support Ring 11

6 Major Milestone Schedule 16

7 OTV Engine Layout 18

8 Igniter/Injector Assembly (ALRC Drawing No. 1191990) 21

9 Chamber and Tube Bundle Nozzle {ALRC Drawing No. 1191991) 22

10 L02 Boost Pump (ALRC Drawing No. 1191994) 23

11 L02 Boost Pump (ALRC Drawing No. 1191996) 24

12 LH2 TPA (ALRC Drawing No. 1191997) 25

13 L02 TPA (ALRC Drawing No. 1191999) 26

14 Shutoff Valve (ALRC Drawing No. 1193176) 27

14A OTV Flow Schematic 28

15 LOX/LCH4 Engine Layout 33

16 High-Speed LCH4 TPA 35

17 High-Speed LOX TPA 36

18 Baseline Mode 1 Dual-Fuel and Alternate Mode 1 37Low-Speed RP-1 TPA

19 Baseline Mode 1 Dual-Fuel and Alternate Mode 1 38Low-Speed LOX TPA

20 Alternate Mode 1 Gas Generator Cycle Engine Thrust 39Chamber Injector

21 Alternate Mode 1 Coaxial Gas Generator 40

21A LOX/LCH4 Engine Flow Schematic 41

22 Sample "Component Requirements" Form 47

23 Task II Schematic 49

24 Fabrication Processes 50

25 Reinforcement Characteristics 51

vn

LIST OF FIGURES (cont.)

Figure No. Page

26 Matrix Characteristics 52

27 Properties of Various Composites 53

28 Detailed Description of Task II Evaluation Form 57

29 LCH4 High-Speed TPA (ALRC Drawing No. 1196001) 66

30 High-Speed LOX TPA (ALRC Drawing No. 1196002) 67

31 LOX Low-Speed TPA (ALRC Drawing No. 1196003) 68

32 Support Structure - Throat, Combustion Chamber 69(ALRC Drawing No. 1196004)

33 Seat - Gimbal Bearing (ALRC Drawing No. 1196005) 70

34 Shaft - Nozzle Extension (ALRC Drawing No. 1196006) 71

35 Injector Housing (ALRC Drawing No. 1196007) 72

36 Support Ring - Skirt Extension (ALRC Drawing No. 1196008) 73

37 Jacket - Tube Bundle, Nozzle (ALRC Drawing No. 1196009) 74

38 Manifold - Coolant, Thrust Chamber (ALRC Drawing No. 1196010) 75

39 Housing - Valve, Propellant (ALRC Drawing No. 1196011) 76

40 Gear - Actuator, Valve (ALRC Drawing No. 1196012) 77

41 Technology Risk Assessment Procedure 86

42 Technology Need Definition, Barrier Coating Process 88

43 OTV Valve Housing 100

44 OTV Valve Housing Fabrication Process Flowchart 101

45 1196011 Valve Housing Technology Program and 102Component Testing

46 Schedule and Budget for the OTV Valve Housing ' 103

47 LCH4 Impeller Housing 105

48 LCH4 TPA Impeller Housing Fabrication Process 106Fl owchart

49 1196001 LCH4 TPA Discharge Housing Technology ' 107Program and Component Testing

50 Schedule and Budget for the LCH4 TPA Impeller Housing ' 109

51 OTV Nozzle Extension Shaft Fabrication Process HOFlowchart

52 1196006 Shaft, Nozzle Extension Technology Program and mComponent Testing

vm

LIST OF FIGURES (cont.)

Figure No. Page

53 Schedule and Budget for the OTV Nozzle Extension 112

54 OTV Skirt Support Ring Fabrication Process Flowchart 114

55 1196008 Skirt Extension Support Ring Technology 115Program and Component Testing

56 Schedule and Budget for the OTV Skirt Support Ring 116

57 Schedule and Budget for Combining the Support Ring / 118and Extension Shaft

ix

I. SUMMARY

A. STUDY OBJECTIVES AND SCOPE

The major objectives of this study were to (1) determine the

extent to which composite materials can be beneficially used In liquid rocketengines, (2) Identify additional technology requirements, and (3) determinethose areas which have the greatest potential for return. The assessment isbased primarily on weight savings, but also considers materials and fabri-cation costs, performance, life, and maintainability factors as applicable.

The five-task study program summarized in Figure 1 was conductedto accomplish the stated objectives. Two engine systems were selected to bebaseline configurations for both orbit-to-orbit and earth-to-orb1t engines.All components from these baseline engines were assessed to Identify thosewhich could potentially benefit most from fabrication with composites.

Twelve components were ultimately selected for further study as aresult of this assessment. Preliminary drawings of the twelve selected com-ponents were reviewed by structural analysts to establish wall thicknessrequirements and to validate the design integrity. Subsequent to the struc-tural analysis, final cross-sectional drawings were prepared for the selectedcomponents, and accurate weights were determined by measuring the crosssections and determining the volumes.

Thought was then given to defining the technological barriersthat would need to be overcome in order to successfully build productionrocket components out of composite materials. A list of these technologyneeds was formulated, and a "Technology Needs Definition" form was filled out

for each of the individual technology needs. This form was used to defineeach technology need, assess its risk, suggest an approach to the problem,and propose a solution.

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Using the previously generated weight data and the information

from the "Technology Needs Definition" forms, a "Technology Needs Cross-

Reference Chart" was formulated. This chart displayed the number of compon-

ents common to each technology and showed the percentage of weight reduction

associated with the application of each technology need. This allowed the

ranking of technology needs as well as specific components in terms of weight

reduction payoffs. It also displayed an assessment of the risk associated

with overcoming each technology barrier.

The "Technology Needs Cross-Reference Chart" was used in

selecting four follow-on tasks recommended for further study and fabrica-

tion. The selected components not only showed promising weight savings

through composite substitution, but their construction also encompassed the

solution of a wide variety of technology needs.

Finally, a plan, schedule, and budget were formulated for the

design, fabrication, and testing of the four selected components.

B. RESULTS AND CONCLUSIONS

This study determined that weight savings between 50 to 80% are

possible on selected components when substituting reinforced plastic compos-

ite (RFC) materials for metal. This translates to an engine weight savings

of 31.4% for the OTV engine and 25.5% for the LCfy 600K engine when compos-

ites are used wherever possible. The lower weight savings percentage for the

LCH4 600K engine results from its greater number of hot-gas components

(>350°F) which cannot use RPC substitution. Metal or ceramic matrix compos-

ites (MMC or CMC) could be used if high temperatures or performance became

bigger "drivers" than weight savings. For the purposes of this study, how-

ever, RPC's were selected over MMC's and CMC's because of their lower cost,

greater ease of fabricability, and higher specific strength.

I, B, Results and Conclusions (cont.)

It was also determined that a wide variety of technology needs(thirteen major categories) remain to be explored in substituting RPC's formetallic parts. Many of these technology needs lend themselves to easysolutions and were only included out of a desire for thoroughness (I.e.,solar radiation effects, bearing surface lubricant, etc.). It is believedthat the remainder of the more difficult technology needs can be solvedthrough straightforward laboratory and fabrication test programs, asdiscussed in Section VI,B of this report.

Table I is an abridgement of the results obtained in Tasks Ithrough IV and was used to select the follow-on tasks recommended for fabri-cation in Task V. The vertical columns show which technology needs areapplicable to a given component and also give the percentage of engine weightsavings possible if the component used composite substitution. The horizon-tal rows show the number of components common to each technology need, aswell as the potential for weight reduction associated with the application ofeach technology need. Consequently, Table I allows the ranking of technologyneeds as well as specific components in terms of weight savings. It alsodisplays our assessment of the risk associated with overcoming each techno-logy barrier. A rating of "high risk" in no way signifies "next to impos-sible" in this chart; it merely indicates that the technology has only beenproposed or theorized and that a reasonable amount of technology testingremains to be done to completely solve anticipated problems. A rating of"medium risk" indicates that less research and testing will be required toimplement a given technology. A "low risk" technology need is one that isjust short of being operational.

The recommended follow-on tasks were selected, using Table I as aguide. This ensured that a combination of promising weight savings and tech-nology advancement features were incorporated into a minimum-cost, low-riskprogram. The components selected for fabrication in a follow-on program areas follows:

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1196011 OTV Valve Housing 8 14

1196001 LCH4 High-Speed 4 17TPA Impeller Housing

1196006 OTV Nozzle Extension 1 6Shafts

1196008 OTV Skirt Support Ring 5 4

The graphlte-epoxy OTV valve housing shown In Figure 2 was

selected both for its significant engine weight savings (1.3%) and because of

the fact that its fabrication would address a total of fourteen technology

needs. It is also a component which could be immediately useful on several

engines. This program could be completed 1n thirteen months for a cost of

$380K. It should be noted that this program could just as easily be split

into distinct segments (i.e., technology, design, fabrication, and test) and

performed over a 2- or 3-year time period with Incremental funding

(approximately $150K/year).

The graphite-epoxy Impeller housing shown 1n Figure 3 1s the

biggest engine weight saver on the 600K booster (2.5% engine weight savings)

and 1s also the most complicated in terms of advancing the state of the art.

Its complex shape and propellant exposure would result in the need to address

seventeen technology needs before a housing could be successfully fabricated

and tested. Its successful fabrication, however, would greatly facilitate

the construction of any of the other engine subcomponents. This program is

the most ambitious of the recommended tasks and could be completed in sixteen

months for a cost of $448K.

The graphite/fiberglass epoxy OTV nozzle extension shaft shown in

Figure 4 shows the greatest percentage of engine weight savings of any of the

components selected in this study (3.3%) and 1s also fairly simple to

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fabricate. This program could be completed 1n thirteen months for a cost of$231K. This 1s one of the simplest and least costly programs which could be

performed and still result In significant weight savings.

The honeycomb/graphlte-epoxy OTV skirt support ring shown 1nFigure 5 shows an engine weight savings of 2.2% and is also simple to fabri-

cate. Additionally, it could be mated with the aforementioned extensionshaft to form a subassembly. The support ring could be developed and tested

in thirteen months for a cost of $231K. Developing both the extension shaftand support ring together would result 1n certain economies because of the

commonality in the technology testing and subcomponent tests. They could

both be developed in the same 13-month time frame for a cost of $281K.

It is recommended that a follow-on program be funded to 1)

resolve technology needs, 2) design and fabricate a RPC subcomponent, and 3)test and evaluate the subcomponent. At the very least, a fabrication and

materials technology program should be initiated to find solutions for theidentified technology needs.

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II. INTRODUCTION

A. BACKGROUND

Many Improvements in liquid rocket propulsion are being evaluated

1n an effort to define an economical space transportation system. One such

Improvement Involves the use of composite materials in rocket engine design.

Composites are a family of high-performance materials consisting

of a matrix reinforced with a fiber. The matrix is usually a thermosetting

resin such as epoxy, a ceramic, or a metal such as aluminum. The reinforce-

ment can be Kevlar, fiberglass, graphite, or boron in the form of either con-

tinuous fibers, chopped fibers, or whiskers. The combination of a matrix and

a fiber results in a new composite material which 1s lighter, stlffer, and

stronger than either of Its constituents.

Recent studies indicate that weight reductions of 30 to 45% can

be achieved for specific components of existing engines by using current and

near-term composite technology. The Advanced Oxygen-Hydrocarbon Rocket

Engine Study (NAS 8-33452), conducted by ALRC for NASA-MSFC, suggested that

weight savings of 30 to 40% are possible for an entire LOX/hydrocarbon

booster engine employing near-term and future composite technology.

Previous applications of composites in military, NASA, and com-

mercial projects provide a broad base of experience. Rocket engine applica-

tion, however, will impose additional material, design, and fabrication

requirements due to such factors as hot gas and cryogenic temperature

extremes, maintainability requirements, and a dynamic environment. Further,

past experience has shown that the application of composites requires much

hands-on development of design procedures and fabrication techniques which

are unique to specific components.

12

II, Introduction (cont.)

B. PURPOSE AND SCOPE

The major objectives of this program were to 1) determine the

extent to which composite materials can beneficially be used 1n liquid rocket

engines, 2) Identify additional technology requirements, and 3) determine

those areas which have the greatest potential for return.

The scope included examining all major components from both

earth-to-orbit and orbit-to-orbit engines to determine which could benefit

most from composite substitution. The study guidelines dictated that the

major consideration be weight savings, although cost, life performance, and

maintainability were also considered. Drawings were made, technology needs

were assessed, and a program plan was presented for designing, fabricating

and testing four selected components.

C. APPROACH

To accomplish the program objectives, an effort Involving five

technical tasks was conducted. The tasks conducted were as follows:

1. Task I - Baseline Engine Configurations

Establish baseline representative engine configurations for

both orbit-to-orbit and earth-to-brb1t engines. This task was limited to

schematic descriptions of the two selected engine systems and to documenting

the general configurations and requirements of the engine components.

13

II, C, Approach (cont.)

2. Task II - Component Assessment and Identification

Establish a methodology for assessment of the component con-

cepts. Apply this methodology to all baseline components from Task I and

identify the components which can potentially benefit from fabrication withcomposites. A brief justification was also provided for each component wherecomposite materials do not show a benefit.

3. Task III - Conceptual Design Assessment

Prepare conceptual design drawings (cross sections) of each

component with potential composite application. Perform appropriate struc-

tural analyses as required to obtain realistic weight estimates. Define theminimum effort needed to illustrate the feasibility of technology development

for each component.

4. Task IV - Criticality Ranking of Technology Needs

Establish a critical 1ty ranking of identified technology

needs. The degree of risk for successfully developing the technology wascategorized between low (for components that are just short of being opera-

tional) to high (for components that have only been proposed or theorized).The degree of risk assessment Included such factors as commonality, cost,schedule, and performance. Life-cycle cost data, where available or readilyapproximated, were also utilized.

5. Recommended Tasks

Recommend a minimum of two follow-on tasks Involving compon-

ent fabrication and testing. The recommendations encompass the integration

.14

II, C, Approach (cont.)

of multiple technology needs and Include a detailed approach, schedule, andestimate of the resources required.

D. PROGRAM SCHEDULE AND MAJOR MILEPOSTS

Figure 6 presents a detailed program schedule of the events 1n

Tasks I through V. The mlleposts shown Include submlttal of data for NASA

approval, reviews, and task completions.

15


TASK DESCRIPTION

TASK I BASELINEENGINE CONFIGURATIONS

ORBIT-TO-ORBITEARTH-TO-ORBIT

TASK" 11 COMPONENTASSESSMENT ANDIDENTIFICATION

COSTPERFORMANCELIFE/MAINTAINABILITYWEIGHT BENEFITJUSTIFY EXCLUSIONS

TASK III CONCEPTUALDESIGN ASSESSMENT

LAYOUT DRAWINGSWEIGHT ESTIMATESTECHNOLOGY NEEDS

TAST TV~WTTCALTTYRANKING OF TECHNOLOGY NEEDS

FIGURES OF MERITCOMMONALITIES"COST AND RISK

TASK V RECOMMENDEDTASKS

COMPONENT FAB PLANCOMPONENT TEST PLANSCHEDULERESOURCE ESTIMATE

TASK VI REPORTINGPROGRAM PLANBIMONTHLYPROGRAM REVIEWSFINAL

Figure 6. Major Milestone Schedule

16

III. TASK I - BASELINE ENGINE CONFIGURATIONS

A. OBJECTIVES

The objective of Task I was to establish baseline representativeengine configurations for both orbit-to-orbit and earth-to-orbit engines.

The task was limited to schematic descriptions of the selected engine systemsand to general configurations and engine component requirements.

B. ENGINE SELECTIONS

After consideration of the three orbit-to-orbit point design

studies (Ref. 1, 2, and 3), the 15,000-1bF Advanced Expander Cycle Engine(OTV) was selected to represent the orbit-to-orbit configuration. A layout

of this engine is shown in Figure 7. Component layouts for the injector,chamber and nozzle, boost pump, main pump, and valves are given in Figures 8through 14. These layouts, plus the corresponding component dimensions thatwere utilized in preparing the layout, provide the means for estimating the

composites' impact on weight and structural integrity. A flow schematic forthe OTV engine is shown in Figure 14A. A detailed component weight breakdownfor the metallic OTV baseline engine is shown in Table II. This table alsoshows a preliminary estimate of what the engine would weigh if RPC's were

substituted wherever possible.

A study of the earth-to-orbit engines described in Contracts MAS8-33452 and MAS 8-32967 (Ref. 4 and 5, respectively) led to the selection of

the 600,000-1bF LOX/LCH4 (Cycle C - Gas Generator Cycle) engine to repre-sent the earth-to-orbit configuration. A layout of this engine is shown inFigure 15. Layouts of the major components (turbopumps, injector, and gasgenerator, etc.) are available from Contract MAS 3-19727 and are shown in

Figures 16 through 21. A flow schematic for the 600-K booster is shown inFigure 21A. A detailed component weight breakdown for the LOX/LCfy base-

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ORIGINAL PAGE ESOF POOR QUALITY

TABLE II

ADVANCED EXPANDER CYCLE ENGINE WEIGHT DATA

Page 1 of 4

Engine Components

Radiation-Cooled NozzleValves and Actuators

Two Valve BodiesFour Actuator BodiesActuator End ClosuresFour GatesShaftsGearsSprings

Nozzle Deployment SystemThree Extension ShaftsSupport RingGear BoxBall ScrewsFlex Shafts

Combustion ChamberLiner y

•* CloseoutManifoldsSupport Structure

Nozzle-Tube BundleTube AssemblyFour Reinforcing RingsManifold

CurrentWeight

(80)(72.8)10297.85.010101

(72)2427

21

(74.3)20.4

9.4

22.5

22.0

(38.4)

29.4

3

6

NewWeight

(80)(23.5)2.68.72.2

1.83.63.61

(39.9)4.914

21

(43.7)20.43.7145.6

(34.4)29.41.23.8

Weight*Reduction

(49.3)7.420.35.63.26.46.4-

32.119.1

13.0

-

30.6-5.78.516.4(4.0)-

1.82.2

NOTES:* Potential weight reduction with use

of composite materials

29

ADVANCED EXPANDER CYCLE ENGINE WEIGHT DATA (Cont'd) Page 2 of 4

Engine ComponentsCurrent New WeightWeight Weight Reduction

Propellent Lines

TubesFlangesFlex Joints

Controller

CaseAdd Other

Injector

Body

FaceV Coaxial Elements/ • —z*— 1

Manifold

Two Clevises

LOX TPA

Pump Housing

Seal HousingTurbine Housing

ImpellerShaftTurbineBearing

LH2 TPA

LHg Inlit Hey§ingPump Housing

Inducer

ImpellersTurbine Exit Housing

(37)9

1018

(35)11

24

(30.6)21

1.01.2

5

2.4

(25.1)5.4

7.0

4.0

.2

2.0

4.0

2.5

(21.5)

4.0

8.5

.1

3.0

3.7

(17.4)3.8

6

7.6(32.4)

8.4

24

(23.1)16

1.0.5

3.8

1.8(15.0)

2.4

5.3

1.8

.1

.9

2

2.5

(9.8)

L83.8

.05

1.51.4

(19.6)5.2

4

10.4(2.6)2.6-

(7.5)5-

.7

1.2

.6

(10.1)3

1.7

2.2

.1

1.12

-

(11.7)2,24.7.05

1.52.3

30


Engine Components

LH2 TPA (Cont'd)ShaftTurbinesBearings

Misc. Valves and Pneumatic PackageIgnition SystemLhL Boost Pump

Turbine-Impeller HousingExit HousingTurbine ImpellerImpeller BoltShaftBearings

LOX Boost PumpTurbine-Impeller HousingExit HousingImpeller ShaftBearings

Heat ExchangerOuter ShellInner Shell

GimbalThrust PadThrust MountCapShaftMonoballFasteners

CurrentWeight

1.4.6

.2[12.6][9.2](8.6)3.64.1

.3

.1

.4

.08(5.5)

2.4

2.7

.3

.08

(5)

3.5

1.5

(3.3)1.2

.7

.2

.8

.2

.2

NewWeight

.7

.3

.2[12.6][9.2](3.9)1.61.8.15.1.18.08

(2.5)1.051.2.15.08

(2.5)1.75.75

(2.1).54.49.15.58.15.2

WeightReduction

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.3---

(4.7)2.02.3.15-.22-

(3.0)1.351.5.15-

(2.5)1.75.75

(1.2).66.21

.05

.22

.05_

31


Engine ComponentsCurrentWeight

NewWeight

WeightReduction

Miscellaneous

Electrical Harness

Service Lines

TPA Protective Bulkhead

Attachment Hardware

Instrumentation

[37]

12.5

6.5

.4

15.02.6

[37]

12.5

6.5

.4

15.0

2.6

-

-

-

-

-

-

TOTAL WEIGHT 567.9 389 178.9

32

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Ill, B, Engine Selections (cont.)

line engine is shown in Table III. (The table includes estimates for engine

weight using composites wherever possible.) The majority of the component .weights shown in Tables II and III were based only on scaling equations or

preliminary drawing volumes. The ten components selected for further studyduring Task III (Section V of this report) were stress-analyzed and drawn in

detail to obtain more precise weights. (See Table VI.)

C. COMPONENT REQUIREMENTS DATA SHEETS

After selecting the two baseline engine configurations, physical

requirements for the major engine components were documented on "Component

Requirement Sheets." An example of a "Component Requirements Sheet" for theoxldizer TPA on the LOX/LCH4 engine is shown in Figure 22. The remainderof these data sheets is contained 1n Appendix A. This evaluation of eachcomponent's physical characteristics (i.e., thrust, Isp, temperature, pres-sure, weight, etc.) laid the groundwork for beginning Task II.

42

TABLE III

LOX/LCH4 CYCLE C ENGINE WEIGHT DATAPage 1 of 4

Engine Components

L02 TPAInducer HousingInducerImpellerImpeller HousingTurbine Inlet HousingTurbine VanesTurbine RotorsShaftBearings

CH4 TPAInducer HousingInducerImpellerImpeller HousingTurbine Inlet HousingTurbine VanesTurbine RotorsShaftsBearingsAdd Miscellaneous

r

InjectorBodyFaceLOX Manifold CoverFuel Manifold CoverHousingAdd MiscAcoustic Cavity

CurrentWeight

(895)122845114466

62

217

145

4

(661)

79

63

40

163

46

42

138

86

2.2

1.8

(611)

373

17

33

33

514

100

NewWeight

(624.2)90.741.725.4496662217

68.4

4

(396.8)

52.4

19.9

19.9

34

46

42

138

40.6

2.2

1.8

(477)

282

172525244100

WeightReduction

(270.8)31.342.325.695

-

-

-

76.6-

(264.2)

26.6

43.1

20.1

129

-

-

-

45.4

-

-

(134)

91

-

8

8

27

-_

43

LOX/LCH4 CYCLE C ENGINE WEIGHT DATA (Cont'd) Page 2 of 4

Engine Components

V High-Pressure LinesTubes (1)Tubes (2)Tubes (3)Flanges (1)Flanges (2)Flex Joints

Combustion ChamberLiner

V Qlqseoyt-ManifoldsSupport Structure

NozzleTube AssemblyManifoldManifoldReinforcing RingsJacket

L02 Boost PumpHousing HybridInducer BoltTurbineShaftBearingsAdd Miscellaneous

CurrentWeight

(384)3017840801640

(364)

154

61

50

99

(328)

109

93

49

22

55

(311)

193

7

12

60

1

38

NewWeight

(334.4)12.81784047.61640

(228.6)

154

24

11

39.6

(220)

109

58.5

31

10.5

11

(176)

97

7

6

27138

WeightReduction

(49.6)17.2

--

32.4

-

-

(135.4)-3739

59.4

(108)

-

34.5

18

11.5

44

(135)

96

-

633

-_

44

LOX/LCH. CYCLE C ENGINE WEIGHT DATA (Cont'd) Page 3 of 4

Current New WeightEngine Components Weight Weight Reduction

Low-Pressure Lines (253) (113) (140)Tubes (1) 12Tubes (2) 19Tubes (3) 54 72 113

Tubes (4) 100Flanges 68 41 27

Gimbal (204) (133.6) (70.4)Seat 29 13 16Body 166 114 52Block 7 5 2Shaft 2 1.6 .4

Miscellaneous [174] [174]LOX Valves (156) (117.1) (38.9)

Bodies 80 60.9 19.1Balls 38 25 13Shafts 13 6.2 6.8Miscellaneous 25 25

Pressurization System [138] [138]CH4 Valves (134) (110.9) (23.1)

Bodies 70 64 6Balls 33 21.7 11.3Shafts 11 5.2 5.8Miscellaneous 20 20

Controller (130) (120) (10)Housing 42 32 10Add Other 88 88

45

LOX/LCH4 CYCLE C ENGINE WEIGHT DATA (Cont'd) Page 4 of 4

Engine Components

Fuel Boost PumpHousingInducerTurbineInducer BoltShaftBearingsAdd Miscellaneous

Interpropellant SealGas Generator

Injector BodyManifoldDomeChamberElements

Ignition SystemHot-Gas Manifold

CurrentWeight

(103)62154.5

21

.3

.2

[90]

(76)

8

13

5

25

25

[40]

[23]

NewWeight

(62.5)427.52.5

10

.3

.2

[90]

(52.8)

6

9

3.3

25

9.5

[40]

[23]

WeightReduction

(40.5)207.52-

11---

(23.2)

2

4

1.7

-

15.5

-

-

TOTAL WEIGHT 5075 3632 1443

46

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IV. TASK II - COMPONENT ASSESSMENT AND IDENTIFICATION

A. OBJECTIVES

Task II involved the assessment of all baseline engine componentsfrom Task I and the identification of those components which could poten-

tially benefit the most from composite material substitution. Figure 23 dis-plays a schematic which outlines the steps taken in performing this task.

The basic contractual ground rules from NASA-MSFC mandated thatthe selections for composite substitution be based primarily on weightsavings, and secondarily on cost, life, and maintainability. It was also

desired that the selections address a variety of new technology needs and notmerely represent a direct application of existing airframe composite techno-logy to simple components such as frames and gimbal rings.

B. COMPOSITE PROPERTIES AND FABRICATION PROCESSES

The initial effort in Task II was to gather and organize infor-mation pertinent to the uses and limitations of composite materials.

Material properties for the most commonly used reinforced plas-

tic composites (RPC's) are contained in Appendix B. The material proper-ties data were obtained from one of the following sources: 1) vendor data,2) Air Force Composite Design Guide (Ref. 6), 3) technical seminars, 4) text-books (Ref. 7), and 5) literature search of technical papers. Appendix Ccontains similar material properties data for the most commonly used metalmatrix composites (MMC).

A short summary of composite fabrication processes, reinforcement

characteristics, matrix characteristics, and material properties is shown inFigure 24 through 27, respectively. These figures are included for the

48

ORIGINAL PAGE *OF POOR QUALITY

TASK ICOMPONENTS

APPLICATIONREQUIREMENTSAND ENVIRONMENT

• STATIC• DYNAMIC• THERMAL• COMPATIBILITY

COMPONENTCONFIGURATION

GENERALMATERIALSELECTION

FABRICATIONMETHODS

COMPOSITESTOCK LAY-UP

IN SITULAY-UP

SELECTED MTL/DES/FAB

JUSTIFICATION

FORM AND CURE_L

CURE

1— EXCLUDED MTL/DES/FAB

I WEIGHT SAVING 1'

[COST EFFECTIVENESS)

P E R F O R M A N C E |i

LIFE

MAINTAINABILITY |

Figure 23. Task II Schematic

49

ino

COMPRESSION MOLDING - RESIN IMPREGNATED FIBER MATERIAL (CALlfD PREPREG ORMOLDING COMPOUND) IS PLACED IN A MATCHED DIE MOLD AND HEAT AND PRESSUREARE USED TO FORM AND CURE THE PART. FINISHED DETAIL IS GOOD AND INSERTS AND iLINERS CAN BE INTEGRALLY ATTACHED

(AUTOMOTIVE AND ELECTRONIC STRUCTURAL PARTS) '

AUTOCLAVE MOLDING - PREPREG IS PLACED ON A FORMING TOOL ALONG WITH A BLEEDER ;SYSTEM. THE LAYED UP MATERIALS ARE COVERED WITH A PLASTIC FILM AND SUB-JECTED TO HEAT AND HIGH PRESSURE IN AN AUTOCLAVE. (EXAMPLE - ABLATIVE CHAMBER)

VACUUM BAG MOLDING - SIMILAR TO AUTOCLAVE MOLDING EXCEPT THAT AMBIENT jPRESSURE IS USED IN PLACE OF AUTOCLAVE PRESSURE (NON-STRUCTURAL AIRCRAFTPARTS) !

PULTRUSSION - PREPREG MOLDING MATERIAL IS DRAWN THROUGH A DIE IN WHICH RAPIDCURINGOCCURS. THIS PROCESS IS LIMITED TO MAKING STRAIGHT PIECES HAVING ACONSTANT CROSS SECTION (AUTOMOTIVE DRIVE SHAFTS) i

REACTION INJECTION MOLDING -PREFORMED FIBER MATERIAL IS PLACED IN A MOLD SCAVITY. RESIN IS INJECTED INTO THE MOLD WHERE IMPREGNATION AND MOLDING JOCCUR (AUTOMOTIVE AND ELECTRONIC STRUCTURAL PARTS)

Figure 24. Fabrication Processes

KEVLAR

GRAPHITE

FIBERGLASS-

BORON

A SYNTHETIC ORGANIC FIBER WITH A VERY HIGH TENSILESTRENGTH AND MODULUS. ON THE OTHER HAND IT DISPLAYSPOOR PROPERTIES IN COMPRESSION AND HAS PROBLEMSWITH MATRIX ADHESION TO THE KEVLAR FIBERS. GENERALLYLIMITED TO TENS ION APPLICATIONS

MADE FROM A SYNTHETIC ORGANIC FIBER AS A RESULT OF ASTRESS - GRAPHITIZATION PROCESS. IT POSSES HIGHTENSILE AND COMPRESSIVE PROPERTIES

MADE FROM SILICA AND WIDELY USED AS A REINFORCE-MENT IN COMPOSITE MATERIALS. NOT AS STRUCTURALLYEFFICIENT AS THE ORGANIC FIBERS, BUT VERY COST EFFECTIVE

FIBER IS MANUFACTURED BY VAPOR DEPOSITION OF BORONON A TUNGSTEN WIRE FILAMENT. CHARACTERIZED BY A HIGHMODULUS THAT SUITS IT FOR USE IN BOTH RESIN MATRIXAND METAL MATRIX COMPOSITES. VERY EXPENS IVE

o o

O j530 ¥

o -o

Figure 25. Reinforcement Characteristics

enro

°

EPOXY - WIDELY USED BECAUSE OF ITS CONVENIENT TRANSFORMATION FROM A LIQUIDTO A PLASTIC SOLID AND BECAUSE OF ITS ABILITY TO ADHERE TO A WIDE VARIETYOF REINFORCEMENTS. EPOXY MATRICES ARE ADVERSELY AFFECTED BY EXPOSURETO RADIATION, HIGH TEMPERATURE, HUMIDITY, AND CERTAIN CHEMICALS

POLYIMIDE RESINS - SIGNIFICANTLY LESS CONVENIENT TO USE IN COMPOSITEMATERIALS THAN EPOXIES. LESS AFFECTED BY HIGH TEMPERATURE AND CHEMICALEXPOSURE o|

:i£

ALUMINUM -DIFFICULT TO PRODUCE AND IS STRUCTURALLY LESS EFFICIENT THAN THE f|PLASTIC COMPOSITES. ON THE PLUS SIDE, ALUMINUM HAS A HIGHER TEMPERATURE ™

'

LIMIT AND IS LESS AFFECTED BY CHEMICAL FACTORS BECAUSE OF ITS IMPERME-ABILITY

CARBON -VERY DIFFICULT TO PRODUCE AND VERY EXPENSIVE. ITS STRUCTURALEFFICIENCY AT VERY HIGH TEMPERATURES IS UNMATCHED

Figure 26. Matrix Characteristics i

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IV, B, Composite Properties and Fabrication Processes (cont.)

reader's quick reference and information and are not intended to supplant theinformation in Appendices B and C.

C. METHODOLOGY AND EVALUATION FORM

The next effort in Task II was to establish a methodology for the

assessment of the component concepts. This was accomplished by preparing anevaluation form containing logic which led to the selection of components

most likely to benefit from composite substitution.

Table IV shows an example of the "Task II Evaluation Form" forthe oxidizer TPA on the LOX/LCfy engine. A detailed description of theparameters being evaluated in Table IV is found in Figure 28.

D. COMPONENT SELECTION

Each major component from the earth-to-orbit baseline engine(LOX/LCH4, Cycle C) and the orbit-to-orbit baseline engine (OTV) underwent

an evaluation in accordance with the methodology described in Table IV.Appendix D contains the evaluation forms for both engines. This evaluationnarrowed the field of 92 major components to 12 components which could poten-tially benefit most from composite material substitution. The component

selection rationale followed the steps shown below:

1. Ten parts were selected from each engine system solely onthe basis of weight savings to be gained through compositesubstitution. (A total of 20 parts.)

2. The twenty parts were categorized and numbered basedon percent of engine weight savings.

54

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ORIGINAL PAGE SSOF POOR QUALITY

Engine:Component:Weight:

% Engine Weight:Ranking:

Temperature and Pressure:Part:Material Density:Current Weight:

Length/Diameter:Proposed Fabrication Method:

Critical Failure Mode:

Proposed Material:

Volume Fraction:

New Weight:

AWeight:

Either LOX/LH2 OTV or LOX/LCH4 Cycle CName of major component (i.e., turbopump)Baseline metallic weight of componentComponent weight/total engine weightEach component is ranked according to percent ofengine weight. The heaviest is ranked No. 1.Design conditions that the component experiences.Subcomponent (i.e., shaft or rotor of the turbopump)Current metallic densitySubcomponent metallic weight (Estimated by obtainingthe approximate volume of the subcomponent andmultiplying it by the metallic density).Subcomponent envelopeA fabrication process will be selected for the com-posite component on the basis of cost, quality level,and functional properties.Most likely mode of failure, determined by pre-liminary structural analysis. This assessment aidsin the preliminary composite material selection.A composite material will be selected as a substitutefor the metallic part on the basis of the following:(1) propellent compatibility, (2) low temperaturetoughness, (3) elevated temperature and pressurestability and strength, and (4) storage deteriorationcharacteristics. An extensive material propertiesliterature search, in conjunction with consultationof material engineering specialists, will providethe information necessary to make these selections.Estimated percentage of the subcomponent which canbe replaced with a composite material.Obtained by multiplying the metallic volume by thecomposite density. The volume estimate will be refinedin Task III if the structural analysis indicatesany wall thickness changes.Difference in weight between the baseline metallicpart and its composite substitute.

Figure 28. Detailed Description of Task II Evaluation Form (Sheet 1 of 5)

57

ORIGINAL PAGE S3OF POOR QUALITY

Cost: The cost factor assessment methodology is illustrated onSheet 3 of this figure. It provides a relative cost com-parison for rocket engine components made from compositematerials. Manufacturing costs are broken down intofour categories: materials, labor, facilities andtooling. Material cost data are obtained from materialsuppliers. Labor manhours are estimated on the basisof the part's complexity and the required fabricationprocess. Tooling costs are similarly estimated.Facility costs are determined by the process selectedand are estimated only on a very general basis.The tooling, material, and labor costs are weightedequally because these manufacturing expenses are totallyrecovered in the price of the part. The facility costis assigned a weighted factor one fifth (1/5) that ofother component costs as these are prorated over otherproducts produced at the facility.Industrial engineering man-hour data available in theAir Force Composite Materials Design Guide (Manufacturing,Volume III) are being evaluated for use in estimatingmanufacturing costs (labor and tooling). Proposedmanufacturing flow sheets, vendors, and subcontractorsare also being used as sources of manufacturing costdata.

Maintainability: The maintainability assessment methodology is illustrated onSheet 4 of this figure. It provides a relative maintain-ability comparison for rocket engine components madefrom composite materials.The maintainability factor is broken down into threecategories: life, frequency of repair, and cost ofrepair. The life of a particular component is estimatedfrom the stability of the material in the operationalenvironments and/or its low cycle fatigue properties.The frequency of repair is estimated from the wear onthe component due to chemical or mechanical erosion(abrasion) and/or its high cycle fatigue properties.The cost of repair is estimated from the amount ofrefurbishment required to restore the part to an acceptablelevel of performance.A weighting factor of one fourth (1/4) was used to assessthe effects of the frequency of repair and the cost ofrepair maintainability factors. The life component ofmaintainability was weighted higher (1/2) than the othercomponents because it directly determines any need formaintenance.

Figure 28. Detailed Description of Task II Evaluation Form (Sheet 2 of 5)

58

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Selected/Excluded: All of the above-mentioned parameters will beevaluated for an acceptance-exclusion decision.AWeight will be given major consideration in thisdecision, with cost, maintainability, performance,and life being considered secondarily.

Justification for Exclusion: This will consist of a brief explanation why asubcomponent did not meet the requirements forcomposite substitution.

Figure 28. Detailed Description of Task II Evaluation Form (SheetJ5 of_5)

61

IV, D, Component Selection (cont.)

3. Any parts which duplicated the technology needs ofanother part were eliminated from consideration (retaining

the one with the highest percent of engine weight savings.)

4. Finally, the number of parts was narrowed based on cost,life, and maintainability.

Table V shows a list of the twelve selected components, along ;

with information on weight savings, proposed composite material, and proposedfabrication method. It must be stressed that these twelve components were

selected primarily on the basis of weight savings and secondarily on thebasis of cost, maintainability, performance, etc. It should also be under-stood that the numbers in the "Change in Weight" (AWT) column in Table V were

based solely on preliminary structural analysis at this point in time. These

numbers, however, were considered sufficiently accurate to select the best"weight savers." A more precise AWT number was obtained by measuring thecross-sectional area of the twelve finished drawings during Task III.

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V. TASK III - CONCEPTUAL DESIGN ASSESSMENT

A. OBJECTIVES

The major objectives of Task III were to prepare conceptualdesign drawings (cross sections) of each of the twelve components selected inTask II and to perform appropriate structural analyses as required to obtainrealistic weight estimates. A secondary objective was to visit outside ven-dors in order to supplement our knowledge of cost, life, fabrication, andmaintainability.

B. STRESS ANALYSIS

Preliminary drawings of the twelve selected concepts werereviewed by both structural and material engineering specialists, and thefollowing analysis steps were taken with respect to each drawing:

1. On- and off-axis material structural property calculations2. Von Mises failure criteria for biaxial mechanical stability3. Residual thermal stress calculations for interlaminar and

coating-interfacial stability at cryogenic temperatures4. Stress analysis calculations to determine minimum section

requirements

C. FINAL DRAWINGS AND WEIGHT ESTIMATES

Subsequent to the structural analysis, the final cross-sectionaldrawings were prepared for each of the twelve selected components. Thesedrawings, shown in Figures 29 through 40, contain fabrication notes andinformation on technology needs.

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V, C, Final Drawings and Weight Estimates (cont.)

The weights of the composite-substituted components were then

determined by using Simpson's rule to calculate part volume from the final

drawing dimensions. Table VI contains data on component metallic weight,

corresponding composite weight, and the percent weight savings achieved

through composite substitution.

D. OUTSIDE VENDOR DESIGN INPUT CONCERNING REINFORCED PLASTICCOMPOSITES

The design concepts for the twe>ve selectee' parts benefited from

consultation with a great variety of expert sources outside of ALRC during

the course of the program. Table VII contains a complete list of every

company contacted for consultation purposes.

A meeting was held with nonmetallic composite experts from the

Aerojet Strategic Propulsion Company (ASPC) to review the twelve cross-

sectional drawings from Task III. Their major commopts were as follows:

1) Some of the more complex shapes would be a challenge from a

fabrication standpoint, but are all considered feasible.

2) Extensive use of chopped molding compound (isotropic

material) in low stress areas would promote the produci-

bility of thick sections and complex surfaces where

maintaining fiber alignment is difficult.

3) The use of woven fiber in low stress areas instead of uni-

directional prepreg would aid in controlling fiber align-

ment.

78

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V, D, Outside Vendor Design Input Concerning Reinforced PlasticComposites (cont.)

These comments were incorporated into the design of the compon-

ents which were recommended for fabrication in Task V of this program.

In addition to the ASPC consultation, a trip was made to Los

Angeles to meet with six fabricators of reinforced plastic composite mater-

ials. Appendix E lists the persons contacted during the visits, along with

the product lines of the companies. Seven of the completed Task III drawings

were reviewed by each of the contractors, and they were asked to assess their

capabilities to make the parts together, listing the anticipated processing

difficulties. They made comments on minor part-geometry redesign which would

simplify fabrication and also indicated that they would like to be involved

in the design of any part which would ultimately be fabricated at their

facility. We also feel that if the bigger companies with composite design

experience were involved in the design from the beginning, a lot of unneces-

sary and expensive supporting technology testing could be avoided in any

follow-on fabrication program. All of the companies felt that the components

would be satisfactory for manufacturing, and some of the companies prepared

price quotes for inclusion in the Task V section of this report.

E. OUTSIDE VENDOR DESIGN INPUT CONCERNING METAL MATRIX COMPOSITES

Outside vendor consultations (Table VII) were made during the

course of the program with regard to the application of metal matrix compo-

sites to the selected components. The companies visted were Nevada Engi-

neering & Technology Corporation (NETCO), DWA Composite Specialities, and

Aerojet Solid Propulsion Company. A summary of their major comments is

presented in the following listing. (Also see Appendix E.)

81

V, D, Outside Vendor Design Input Concerning Reinforced PlasticComposites (cont.)

1) Selected parts too complex for fabrication by laying upMMC laminates and diffusion bonding.

2) TPA impeller housing is well beyond state of the art.

3) In complex shapes with multidirectional stress distri-butions, boron-aluminum and graphite aluminum becomeinefficient.

4) Valve housing could be fabricated from an aluminum-silicon carbide composite (powder metallurgy) such asDWA-AL. This material, however, possesses poor weldability.

5) MMC components would be lighter than those made from unrein-forced metal but heavier than those made from RPC's .

6) No problems with permeability or compatibility.

Based on the recommendations of both outside expert sources

and ALRC materials engineering experts, it was decided that reinforced

plastic composites were preferable for fabricating the selected designs (Task

V). Metal matrix composites are out of consideration at the present time

because of their higher cost, lower specific strength, and greater fabrica-

tion difficulties. Table VIII contrasts the weights and fabrication risks of

three individual components made from 1) RPC, 2) boron-aluminum, 3) silicon

carbide-aluminum, and 4) baseline metal, respectively. The weights in Table

VIII were determined analytically by performing a stress analysis to deter-

mine the appropriate wall thickness for each material. It can be seen that

the RPC components are lighter in every case and pose far fewer fabrication

difficulties.

If higher temperatures and greater performance become more

important "drivers" than weight savings in future programs, it is possible

that the use of certain MMC or ceramic materials would be indicated. For the

present study, however, it is clear that RPC components are lighter and more

cost-effective.

82

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VI. TASK IV - CRITICALITY RANKING OF TECHNOLOGY NEEDS

A. OBJECTIVES

The first objective of Task IV was to define the technology needsinvolved in developing the twelve selected concepts and to evaluate thoseneeds in terns of level of risk involved in bringing those technologies tooperational status. The second objective of Task IV was to provide a criti-cality ranking of the identified technology needs and to justify the rankingswith narrative and figures of merit.

i

B. TECHNOLOGY NEEDS

At the commencement of Task IV, a great deal of thought was givento defining the technological barriers that would need to be overcome inorder to successfully build production rocket components out of reinforcedplastic composites. A list of these technology needs is shown in Table IX.

Each of these technology needs was evaluated to establish theimpact of required technology. Figure 41 is a schematic which outlines thesteps taken during this evaluation. Each major component was examined toidentify any parts that had not been developed. Except for rocket nozzles,where composite nozzles have been used, essentially all composite-substitutedcomponents fall into the new-with-composite screening category.

Those parts requiring further development were then evaluated interms of the level of risk that would be involved in bringing them to opera-tional status. The degree of risk was categorized between low (for compon-ents that are just short of being operational) to high (for components thathave only been proposed or theorized). The sensitivity of each component tothese risks was assessed in terms of cost, schedule, performance, life,weight, and commonality of technology. The results of this evaluation were

84

TABLE IX

TECHNOLOGY NEEDS LISTING

1.0 H2 Compatibility

2.0 02 Compatibility

3.0 CH4 Compatibility

4.0 Low Temperature Toughness

5.0 Fabrication

5.1 Mechanical Fastening

5.2 Sealing Surface Finish

5.3 Vane Manufacturing Method

5.4 Vane Attachment Method

5.5 Fabrication Sequence

5.6 Barrier Coating Process

5.7 Plumbing Connections

5.8 Detail Joining Methods

5.9 Detail Fabrication Methods5.10 Mold Design

6.0 Cryogenic Properties

7.0 Interface Properties

8.0 Metal Coating Interface Properties

9.0 Differential Expansion Properties

10.0 Solar Radiation Effects

11.0 Low Cycle (Thermal) Fatigue

12.0 High Cycle Fatigue (HCF)

13.0 Bearing Surface Lubricant

85

ORIGINAL PAGE IS6F POOR QUAUTY

I SELECTED CONCEPT 2SELECTED CONCEPT 1

TECHNOLOGY SCREENING

•AVAILABLE• NEW• NEW WITH COMPOSITE

LOW RISK

• CURRENT TECHNOLOGY• FULL-SCALE• SIMILAR SYSTEM

I

MEDIUM RISK

• CURRENT TECHNOLOGY• SUBSCALE

EVALUATE IMPACT

WEIGHTCOSTSCHEDULEPERFORMANCELIFECOMMONALITY

1

HIGH RISK

• IN DEVELOPMENT• PROPOSED FORDEVELOPMENT

J

Figure 41. Technology Risk Assessment Procedure

86

VI, B, Technology Needs (cont.)

documented in a series of "Technology Need Definition" forms. One of theseforms was filled out for each of the categories shown in Table IX. This formdefined the technology need, assessed its risk, suggested an approach to theproblem, and proposed a solution. An example of one of these forms is shownas Figure 42 (Barrier Coating Process), and the balance is included asAppendix F in this report.

Many of the thirteen technology needs lend themselves to easysolutions and were only included out of a desire for thoroughness (i.e.,solar radiation, bearing surface lubricant, etc.). It is believed that thebalance of the technology needs can be solved through straightforwardlaboratory and test programs. This optimism about the ability to quicklysolve the thirteen technology needs through technology programs is based onthe following rationale:

1.0-3.0 Propel!ant Compatibility - Problems which can resultfrom chemical incompatibility or from freeze-thaw cycling (causing cracks)will be precluded by the application of internal barrier coatings. (See 8.0bel ow.)

4.0 Low Temperature Toughness - Thermoset polymer matrix com-posites are brittle materials at room temperature. Evidence of this is seenin the resin microcracking that results from the residual thermal stress fromcuring at elevated temperature. These composite materials, however, exhibita remarkable toughness capability because of the many individual fiber andresin interfaces, each of which is structurally redundant. This results in afracture process that is progressive rather than sudden, as is characteristicof conventional brittle materials. Such behavior is characteristic of out-standing fracture toughness. Because the strength of polymer matrix compos-ites is known to increase slightly at cryogenic temperatures, their fracture

87

ORIGINAL PAGE S3OF POOR QUALITY

ITEM: 5.6 - Barrier Coating Process

DESCRIPTION: High-performance plastic composites must be sealed in orderto contain liquids and vapors.

ASSESSMENT OF RISK:

* Medium Low

*Microcracking occurs in the resin matrix of high-performance com-posite materials. These cracks result from residual and appliedstresses. They render the resin matrix permeable to vapors and1iquids.

APPROACH: Several barrier coatings (metal foils, plastic films, etc.) havebeen used successfully to seal composite materials. Thesecoating materials and processes will be evaluated to select the

optimum system for a given rocket engine component.

PROPOSED SOLUTION: Ductile barrier coatings are required to seal compositematerials.

Figure 42. Technology Need Definition, Barrier Coating Process

88


toughness is not expected to decrease significantly at low temperaturesbecause of any embrittling effects.

Structural element tests (including impact) at cryogenictemperatures are planned to confirm that this is true for the candidate com-posite materials and processes being considered for this program.

5.0 Fabrication - Any potential fabrication problems (i.e.,fastening, barrier coating, sealing surface finish, etc.) will be solvedthrough the development of manufacturing techniques during the technologyprograms. None of the identified fabrication technology needs pose a seriousproblem in view of the current state of the art with RPC's.

6.0 Cryogenic Properties - This is a low-risk technology needdue to the existence of commercial composite materials which display highstrength and good fracture toughness at cryogenic temperatures. Before fina-lizing any design, the structural and physical properties of the candidatematerial will be verified by structural testing at cryogenic temperatures.

7.0 Interface Properties - Strong adhesive bonds between twocomposite parts or between a composite and a metal part have been demon-strated successfully in the aircraft industry. The stability of these adhe-sive bonds at cryogenic temperatures will be analyzed, and each proposed com-bination will be tested to validate the predictions. In the event that thepredicted structural performance is not attained because of poor adhesivebonding, other design alternatives will be investigated (i.e., bolting,riveting, sewing, etc.).

89


8.0 Metal Coating Interface Properties - The use of metallicbarrier coatings is anticipated in order to protect the composite materialfrom degradation caused by contact with the liquid rocket propellants. Thelarge temperature excursions caused by cryogenic conditions are expected toresult in significant bondline strain. This strain problem can be partiallymitigated by tailoring the thermal expansion properties of the compositethrough control reinforcement and fiber volume (Ref. 8). Another techniquewhich has been successfully used is to apply an elastomeric adhesive tie coatthat approximates the expansion characteristics of the metal barrier (Ref.9). A stable metal coating interface should result if the adhesive bond tothe composite is strong enough to prevent separation during thermal excur-sions. The use of more than one tie coat adhesive introduces additionalinterfaces and reduces the stress at the two key interfaces for greater bondstability.

Satisfactory metal-lined tanks and propellant (LOX) linesof polymer composite have been fabricated and tested by the Martin Companyunder contract to NASA (Ref. 10 and 11).

9.0 Differential Expansion Properties - Differential expansionbetween contacting dissimilar materials results in interfacial stress.Stress rupture and disbonding are two harmful effects of differential expan-sion. A preliminary structural analysis performed at ALRC indicated thatsatisfactory dissimilar material interfaces can be developed over the temper-ature ranges anticipated (see 8.0 above).

10.0 Solar Radiation Effects - The surface of epoxy matrix com-posite materials is degraded by solar radiation. This degradation results insurface crazing and decreases the composites' chemical resistance. Thisappears to be a low-risk technology need since protective coatings are

90


available to block solar radiation. The effectiveness of these coatings canbe demonstrated in ultraviolet (UV) weatherometer exposure tests.

11.0 Low Cycle Fatigue (LCF) - Low cycle (thermal) fatigueresults from structural damage in which thermal stress was a contributingfactor. In composite materials, these stresses can result because of flawsin the material, dissimilar material interfaces, anisotropy, or poor bonding.Past experience shows that once the cause of the structural damage has beenidentified (through structural element testing), one can design around it.Examples of composite components in use today which are designed to success-fully withstand low cycle thermal fatigue are 1) fiberglass LN£ bottles, 2)natural gas vessels, and 3) parts in cryogenic wind tunnels.

12.0 High Cycle Fatigue (HCF) - High cycle (thermal) fatigueresults from structural damage that is caused by stress-induced microstruc-tural changes in the material. These changes occur generally because of somediscrete mechanism that allows the stress to reach a destructive level. Thismechanism depends upon a crack, void, disbond, or dissimilar material inter-face and a dynamic change in the level of stress due to vibration or move-ment.

In the case of RPC's , designing for HCF is a low-risk tech-

nology need. Example of RPC's used in HCF applications abound (i.e., gears,machinery, compressor blades, etc.). This serves to underscore the fact thatcomposites are lighter, stiffer, and more structurally efficient than theirmetallic counterparts.

13.0 Bearing Surface Lubricant - Polymer composite materialsabrade more easily than metallic materials because of the resin micro-cracking. A lubricant can be helpful in sealing the surface and providing asmooth, low friction surface to reduce abrasion damage due to sliding and

91


rubbing. It should be a simple task to select a commercial lubricat ion sys-tem compatible with composite materials, the l iqu id rocket engine environ-ment, and the vacuum conditions of space.

C. CRITICALITY RANKING OF TECHNOLOGY NEEDS

Using the weight data from Table VI and the information from the"Technology Needs Definition" forms, a "Technology Needs Cross-Reference

Chart" was formulated (see Table I). This chart displays the number of com-ponents common to each technology and shows the percentage of weight reduc-

tion associated with the application of each technology need. This allowsthe ranking of technology needs as well as specific components in terms of

weight reduction payoffs. It also displays our assessment of the risk asso-ciated with overcoming each technology barrier.

Since Table I is basically a compendium of the studies conducted

in Tasks I through IV, it was used as a guide in selecting the follow-ontasks recommended for Task V of this study. Those selected for fabrication

in Task V represent the components which provide the greatest percentage ofweight savings through composite substitution and which also encompass thesolution to a wide variety of technology needs.

92

VII. TASK V - RECOMMENDED TASKS

A. OBJECTIVE

The purpose of Task V was to recommend a minimum of two follow-on

tasks involving component fabrication and testing. It was desired that the

selected components not only show promising weight savings with composite

substitution but that their construction also encompass the solution to a

wide variety of technology needs, thus paving the way for a large number of

composite substitutions in the future.

Another objective of Task V was to formulate a schedule and bud-

get for the analysis, design, fabrication, and testing of the selected com-

ponents.

B. RECOMMENDATIONS

The recommended follow-on tasks were selected by using the

critical technology need ranking developed in Task IV (see Table I). This

ensured that a combination of weight reduction and technology advancement

features were incorporated into a minimum-cost, low-risk program.

The components selected for fabrication and further study in a

follow-on program are as shown below:

% Weight Savings Number of TechnologyPN Component Rank Needs Addressed

1196011 OTV Valve Housing 8 14

1196001 LCH4 High-Speed TPA 4 17Impeller Housing

1196006 OTV Nozzle Extension 1 6Shafts

1196008 OTV Skirt Support Ring 5 4

93


VII, B, Recommendations (cont.)

The OTV valve housing shows significant weight savings and

requires the solution of fourteen technology needs. This component could be

immediately useful on several engines while solving problems connected with

many other components.

The LCH4 high-speed impeller housing is the biggest weight

saver on the 600K-lbF booster engine when made with composites. It also pre-

sents the most complications in terms of advancing the state of the art in

composite fabrication technology. If this component could be successfully

built, construction of any of the twelve components selected in Task III

would be facilitated greatly.

The OTV extension shaft yields the greatest percentage of engine

weight savings. At the same time, its fabrication process is fairly simple

and would address the solution of only six technology needs. The extension

shafts can be inserted directly into the OTV skirt support ring to form a

subassembly. This support ring was also selected because it is simple to

fabricate and shows a good weight savings.

Notice should be taken that the four recommended tasks are not

merely the four top weight savers based on the Task III weight analysis.

They were selected not only to show good weight savings but also to cover the

spectrum of technology needs from the simplest to the most complex. They may

be fabricated either together or singly, depending on NASA's schedule and

budgetary needs.

C. PROGRAM PLANS FOR THE SELECTED COMPONENTS

Table X shows a schematic representation of the general design

and analysis steps that would be followed for any of the four recommended

tasks.

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VII, C, Program Plans for the Selected Components (cont.)

The proposed program for each recommended task consists of fiveelements: (1) technology programs addressing the technology needs identified

for each selected part; (2) a preliminary design incorporating the results of

the material characterization portion of the techno1oqy programs and the

conceptual design from Task III; (3) a final design "corporating the resultsof the preliminary design and the fabrication development portion of the

technology programs; (4) fabrication; and (5) testing and evaluation of thetested part.

Each recommended task is costed separately, with a breakdown by

element to allow funding options for a follow-on activity.

Technology Programs

A general description of each technology need is given in Section

VI,B of this report. These programs are concerned with (1) the determina-

tion of mechanical and physical properties of candidate materials, (2) theeffect of propellant exposure on these properties, and (3) the development of

fabrication techniques for application to the final design. The purpose ofthe programs is to minimize risk in the design by providing the data requiredfor material selection, structural analysis, and the processing parametersrequired for fabrication of the part. Each program will be formulated to

identify the candidate materials, test specimens, test procedures, specialtest equipment, and test parameters. The results of cryogenic mechnical

tests will be evaluated for ductility and toughness values with respect to

those at room temperature. Propel lant compatibility tests results will be

evaluated as to chemical reactions and the effect of static and cyclicexposure on mechanical properties.

96

VII, C, Program Plans for the Selected Components (cent.)

Preliminary Design

The conceptual designs of Task III will be the basis for thedevelopment of a preliminary design that also incorporates the results of thematerial characterization portion of the technology programs (i.e., finalmaterial selection and design allowable properties). The Task III structuralanalysis hand calculations for the conceptual design will be refined todetermine new section sizes and detail design features such as laminate ori-entations and thicknesses, bond lines, and mechanical joints. The prelimi-nary design drawings will be reviewed with NASA for their approval.

Structural elements, based on the preliminary design, will befabricated and mechanically tested to verify design property selection andthe structural analyses. Failure to verify the conceptual design throughstructural element testing would require an iteration of the structuralelement tests.

Detail Design

The detail design will incorporate the results of the preliminarydesign and the structural element testing. A final computerized finite ele-ment structural analysis will be conducted to refine section sizes, bondareas, and mechanical joint details. Detail part and assembly drawings willbe presented at a final review with NASA.

The detailed designs will be submitted to ALRC manufacturing andto outside suppliers that provided quotes for the study to finalize costs.

97


Fabrication

The selected fabricators will be monitored to assure con-formance to drawing requirements and quality control procedures, includingraw material inspection, process controls, and nondestructive inspection offabricated parts.

Testing

The test plan for the completed part will be formulated to simu-late the duty cycle. If the part requires pressure testing, it will besealed by using the attachment method to the adjoining part as indicated inthe component design. The part will be proof-tested prior to cyclicpressure-testing with the propellant. In the event of failure during cyclictesting, a destructive post-test failure analysis will be conducted in aneffort to determine failure mode and to determine the degree of materialdegradation. If cyclic testing reaches full duration, the part will bevisually and nondestructively inspected to determine material degradation.At this time, the part will either be destructively analyzed or burst-testedand destructively analyzed.

As each of the four recommended subcomponents is entirelydifferent in function and fabrication complexity, it necessitates a separatebudget, schedule, and test plan for each one. The paragraphs which followdescribe a unique program plan for each of the recommended parts. Theseprogram plans consist of the following: 1) a conceptual drawing of thesubcomponent, 2) a fabrication process flowchart, 3) a summary of therequired technology and component testing, and 4) a detailed schedule andbudget for each program plan.

98

V I I , C, Program Plans for the Selected Components (cont.)

1. OTV Valve Housing

The OTV valve housing, shown in Figure 43, was selected bothfor its s ignif icant engine weight savings (1.3%) and because its fabricationwould address a total of fourteen technology needs. It is also a componentwhich could be immediately useful on existing engines.

Figure 44 shows that the valve body would be filament-woundfrom graphite epoxy and thereafter autoclave-molded. The valve bosses wouldthen be compression-molded separately from Kevlar-epoxy prepreg. Aftermachining openings in the valve body, the bosses would be adhesively attachedvia an autoclave bonding process. The internal barrier coating could beperformed on the mandrel before fi lament-winding the body, or they could bedeposited internally by a variety of methods (i.e., electro deposition,vacuum deposition, etc.) after the valve body is completed. The specificbarrier coating method selected would depend on the results of priorfabrication technology programs.

Figure 45 contains a summary of the material and fabricationtechnology tests which would be performed prior to fabricating the subcompon-ent. These tests would include 1) laboratory coupon tests to determine^,mechanical , thermal, and propel 1 ant compatibili ty properties, 2) fabricationand material process trials, and 3) structural element testing. After devel-oping the technology and fabricating a valve housing, the housing would beproof-tested, cyclic pressure-tested (in LOX, LH2, and LCH4), and des-tructively burst-tested in LOX.

The schedule and budget for performing the technology,design, fabricat ion, subcomponent testing, and evaluation of the valve bodyare shown in Figure 46. The entire program would take place in a 13-month

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time frame and cost a total of $380K. The budget was figured on the basis of

1982 company wage rates and includes all management, travel, and reporting

expenses. It should be noted that this program could just as easily be split

into distinct segments (i.e., technology, design, fabrication, and test) and

performed over a 2- or 3-year time period with incremental funding.

2. LCH4 TPA Impeller Housing

The impeller housing, shown in Figure 47, is the biggest

engine weight saver on the 600K booster (2.5% engine weight savings) and is

also the most complicated in terms of advancing the state of the art. Its

complex shape and propellant exposure would result in the need to address

seventeen technology needs before a housing could be successfully fabricated

and tested. Its successful fabrication, however, would greatly facilitate

the construction of any of the other subcomponents selected in Task III.

Figure 48 shows that the housing torus shell would be laid

up in a female mold and autoclave-molded. The openings for the vane roots

would be machined in the shell, and the vanes would be laid up against forms

inserted in the torus opening and then autoclave-molded. After this, the

balance of the pump housing and torus would be laid up and autoclave-molded.

The method of applying the barrier coating would be determined as in the case

of the OTV valve housing.

Figure 49 summarizes the material and fabrication technology

tests which would be performed prior to fabricating the impeller housing.

These tests would be similar to those performed for the OTV valve housing,

with the addition of vane processing and attachment method testing. The

final impeller housing would be proof-tested, cyclic pressure-tested in

LCH4, and destructively burst-tested in LN2-

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The schedule and budget for developing the LCfy TPAimpeller housing is shown in Figure 50. The program would have a duration of

16 months and cost a total of $448K. This is the most ambitious and costlyof the recommended tasks, but its successful completion would solve the

majority of identified technology needs connected with RPC substitution.^

3. OTV Nozzle Extension Shaft

The OTV nozzle extension shaft, depicted in Figure 34, showsthe greatest percentage of engine weight savings of any of the componentsselected in Task III (3.3%) and is also fairly simple to fabricate.

Figure 51 shows that the extension shaft would be fabricatedby co-pultruding graphite fiber overwrapped with fiberglass. The threadswould thereafter be machined in the fiberglass material overlay and lubri-cated. This simple process would result in a high-strength, lightweight,

hollow extension shaft which could be used as part of a nozzle extensionmechanism.

Figure 52 contains a summary of the material and fabricationtechnology testing which would be performed prior to fabricating the shaft.These tests would include 1) fabrication and material process trials,

2) structural properties testing at ambient and cryogenic temperatures (i.e.,torsion, bending, tension, shear, and impact testing), and 3) durability

tests at ambient temperature. The completed shaft would be subjected tovibration testing, cyclic testing (torsion and bending), and destructive tor-

sion testing.

The schedule and budget for developing the OTV extension

shaft are shown in Figure 53. The program would take place over 13 months

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and cost a total of $231K. This is one of the simplest and least costly

programs which could be performed and still result in significant weight

savings.

4. OTV Skirt Support Ring

The OTV skirt support ring, shown in Figure 36, shows an

engine weight savings of 2.2% and is also simple to fabricate. Additionally,

it could be mated with the aforementioned extension shaft to form a sub-

assembly.

Figure 54 shows that the support ring would be fabricated by

machining a honeycomb core and tape wrapping the honeycomb with graphite-

epoxy. The ring would then be cured in an autoclave mold and finish-

machined.

Figure 55 contains a summary of the material and fabrication

technology testing which would be performed prior to fabricating the ring.

These tests would include 1) laboratory coupon tests to explore low temper-

ature toughness and 2) structural element testing (i.e., fatigue, bending,

and pull-out testing). The completed support ring would then be subjected to

vibration testing, cyclic testing, and destructive bend testing.

The schedule and budget for developing the OTV skirt support

ring are shown in Figure 56. The program would take place over 13 months and

cost a total of $23IK.

It has been found that combining the development programs

for both the OTV nozzle extension shaft and the OTV skirt support ring

results in certain economies due to commonality in the technology testing and

113

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subcomponent tests. Figure 57 shows that a combined program could be per-

formed in the same 13-month time frame for a cost of $281K. An added benefit

would be the ability to test the two components assembled together as a sub-

assembly.

117

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VIII. CONCLUSIONS AND RECOMMENDATIONS

A. CONCLUSIONS

The following conclusions are drawn from this work:

1. Weight savings of up to 80% are possible on selected compon-nents when composite materials are substituted for metal.

2. Engine weight savings from 25 to 30% are possible with useof current composite technology. Future composite techno-may save 30 to 40% of the engine weight.

3. A variety of technology needs remain to be explored in sub-stituting composite materials for metallic parts. Thesetechnologies can be developed through straightforward lab-oratory and fabrication test programs.

4. Reinforced plastic composites were selected over metalmatrix composites because of their lower cost, greaterfabricability, and higher specific strength. If high-temperature applications become more important, or if

propellant compatibility becomes a major problem, the use ofMMC's would be more clearly indicated.

5. A variety of follow-on programs could be performed todesign, fabricate, test, and evaluate rocket engine sub-

components made from composite materials. The period of

performance would range from 13 to 16 months, with the

cost ranging from $231 to $448K for the simplest andthe most complex tasks, respectively.

119

VIII, A, Conclusions (cont.)

6. A follow-on program could just as easily be split into dis-

tinct segments (i.e., technology testing, design, fabrica-

tion, and test) and performed over a 2- or 3-year time

period with incremental funding (approximately $150Kper year).

B. RECOMMENDATIONS

The following recommendations are made on the basis of the pro-

gram results:

1. Conduct technology programs that address the fabrication,material properties, and propellant compatibility of rein-forced plastic composites.

2. Fabricate and test an engine subcomponent which shows prom-

ising weight savings with the use of composite materials and

which solves a wide variety of technology needs (togetherwith, or separate from, the technology programs, depending

on schedule and budget restraints).

3. Extend composite technology to additional rocket engine com-ponents as the technology is developed.

120

REFERENCES

1. Mellish, J.A., "Orbit Transfer Vehicle (OTV) Advanced Expander CycleEngine Point Design Study," Aerojet Liquid Rocket Company FinalReport 33574F, Contract MAS 8-33574, December 1980.

2. Brown, J.R., "Orbit Transfer Vehicle (OTV) Advanced Expander CycleEngine Point Design Study," United Technologies, Pratt & WhitneyAircraft Report FR 14615, Contract MAS 8-33567, March 1981.

3. Anon., "Orbit Transfer Vehicle Advanced Expander Cycle Engine PointDesign Study," Rocketdyne Report RI/R80-218-2 (Study Results),Contract NAS 8-33568, December 1980.

4. O'Brien, C.J. and Ewen, R.L., "Advanced Oxygen-Hydrocarbon RocketEngine Study," Aerojet Liquid Rocket Company Final Report 33452F,Contract NAS 8-33452, April 1981.

5. O'Brien, C.J., "Dual-Fuel, Dual-Throat Engine Preliminary Analysis,"Aerojet Liquid Rocket Company Final Report 32967F, Contract NAS8-32967, August 1979.

6. Advanced Composites Design Guide, Vol. I, 3rd ed., Air ForceMaterials Laboratory, Wright-Patterson AFB, March 1973.

7. Tsai, Steven W. and Hahn, H. Thomas, "Introduction to CompositeMaterials," 1980.

8. Wigley, D.A., "Properties of Materials: The Effect of Low Temperatureon the Strength and Toughness of Materials," AGARD Lecture Series No.Ill, Cryogenic Wind Tunnels, 1980.

9. Wigley, D.A., Properties of Materials: "The Physical Properties ofMetals and Non-Metals," AGARD Lecture Series No. Ill, Cryogenic WindTunnels, 1980.

10. Hall, C.A., Laintz, D.J., and Phillips, J.M., "Composite PropulsionFeedlines for Cryogenic Space Vehicles," NAS 3-14370, August 1973.

11. Caudill, C.L. and Kirlen, R.L., "Composite Overwrapped MetallicTanks," NAS 3-12023, March 1972.

121

APPENDIX A

COMPONENT REQUIREMENT FORMS

A-l

ADVANCED EXPANDER OTV

COMPONENT REQUIREMENTS

A-2

ORIGINAL PAGE ISOF POOR O.UALITY

TABLE III-I


Baseline Engine: Advanced Expander (OTV)

Chamber Pressure

Thrust (Vac)

I s p ( M i n i m u m ) ( v a c )

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15.000 1bf

475.4 sec.

LOX/LH?

1200 thermal cycles

10 hrs.

6.0

574.4 Ib.JD_

Components : Gimbal Assembly

Operating Pressure(s)

Operating Temperature(s")

Propellant Flowrate(s)

Start/Shutdown Conditions

Envelope (Length)

Weight

Material(s)

Subject to 15,000 lbf thrust

Ambient

2.4 in.

3.3 lb_

Tit/SS

A-3

ORIGINAL PAGE BOF POOR QUALITY

TABLE III-I



Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15,000 1bf

475.4 sec.

LOX/LH?

1200 thermal cycles

10 hrs. •

6.0

574.4 Ibm

Components: Chamber


Operating Temperature(s)



Envelope (Length)

Weight

Material(s)

PC = 1200 psia

Chamber coolant = 3.816 Ib/sec, Tube BundleFlowrate = .674 Ib/sec

18 in.

47.3 lbnimZirconium Copper - EF Nickel Closeout

A-4


TABLE 111-1


Basel ing Engine: Advanced Expander (OTV)

Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15,000 1bf

475.4 sec.

LOX/LH?

1200 thermal cycles

10 hrs.

6.0

574.4 Ib

Components: Copper Nozzle





Envelope (Length)

Weight

Material(s)

17.55 psi (forward), 0.52 psi (aft)

730°R

.674 Ib/sec

34.8 in

27 Ibm

copper

A-5

TABLE III-I



Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15,000 Jbf

475.4 sec.

LOX/LH?

1200 thermal cycles

10 hrs.

6.0

574.4 lb.m

Components: Injector





Envelope (Length)

Weight

Material(s)

POJ = 1434, PFJ = 1326, PC = 1200 psia

W =27.05 Ib/sec Wf = 4.51 Ib/sec

4.8 in.

30.6 lbm

304L CRES

A-6


TABLE III-I



Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15,000 1bf

475.4 sec.

LOX/LH?

1200 thermal cycles

10 hrs.

6.0

574.4 Ib

Components : Tube Bundle Nozzle





Envelope (Length)

Weight

Material(s)

2466 psia

730°R

0.674 Ib/sec

30 in.

38.4 1bm

347 ORES

A-7

OF POOR

TABLE III-I



Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15^000 lbf

475.4 sec.

LOX/LH2_

1200 thermal cycles

10 hrs.

6.0

574.4 Ib.m

Components : Radiation Nozzle





Envelope (Length)

Weight

Material(s)

Negligible

2450°F

31.56 Ibm

49.6 in.

80 Ibm

C-103 Columbium Alloy

A-8


IMkLJiblCOMPONENT REQUIREMENTS


Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15.000 Ibj

475.4 sec.

LOX/LH2.

1200 thermal cycles

10 hrs.

6.0

574.4 Ib.

Components : Nozzle Deployment System (3 shafts, DC Motor, Support Ring)





Envelope (Length)

Weight

Material(s)

Ambient

Ambient

60 in.

72 Ib.

A-9

OF P°OR

TABLE III-I



Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15,000 lbf

475.4 sec.

LOX/LH?

1200 thermal cycles

10 hrs.

6.0

574.4 Ib.m

Components : Propellent Flow, Control ValvesAModulating Poppet Valve





Envelope (Length)

Weight

Material(s) • Bodies

ShaftsPoppetsGears

SpringsSeal

8 Total

16 psi - 1450 psi

40°R - 600°R

VI = 27.05 Ib/sec Wf = 4.51 Ib/sec

Prop Flow Control Valve 10.12" x 13.75"McxLEoopet Valve 6.4" x 10.9"

1bm

(1) 6061-T6 Aluminum(2) A356-T6 Aluminum(3) 347 CRES(4) Nitronic 505) A-286(1) A-286(1) A-286(1) A-286(2) 15-5 PH H1150 M(1) 302 CRES(1) Filled Teflon

A-10

ORIGINAL PAGE g§OF POOR QUALITY

TABLE III-I



Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Rat io

Weight

1200 psia

15,000 1bf

475.4 sec.

LOX/LHp

1200 thermal cycles

10 hrs.

6.0

574.4 Ib.m

Components: LOX Boost Pump





Envelope (Length)

Weight

Material(s) Turbine Inlet & Housing

Pump Housing

Impeller & Shaft

Bearings

16-56 psia

-320°F

171 GPM

4.6" X 5.0"

A356

A356

15-5PH HI 150 M CRES

440 C CRES

A-ll

ORIGINAL PAGE SiOF POOR QUALfTY

TABLE III-I



Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15,OOOJbf

475.4 sec.

LOX/LH?

1200 thermal cycles

10 hrs.

6.0

574.4 Ibm

Components : LHp Boost Pump





Envelope (Length)

Weight

Material(s) Turbine Inlet & HousingPump HousingImpeller & Shaft

BearingsPreload Springs

18.5 - 50 psia

-420°F

456 GPM

5" X 2.87"

8.5 Ibm

A356 A1

A356 Al

Cast Ti 5A1 2.5 Sn ELI

440 C CRES

302 CRES

A-12

ORIGINALOF POOR QUALSTY

TABLE III-I


Baseline^ Engine: Advanced Expander (OTV)

Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15,000 1bf

475.4 sec.

LOX/LH?

1200 thermal cycles

10 hrs.

6.0

574.4 lb.m

Components : LOX TPA (Hi Speed)





Envelope (Length)

Weight

Material(s) • Turbine Housing

• Turbinet Pump Housing

• Seal Housing

• Pump Impeller & Shaft

• Bearings

A-13

Pump 48-1487 psiaTurbine 1512-1326 psia

Pump 170°RTurbine 489°R

194 GPM

11.85° X 8.8"

26.9 1bm

(1) A-356 Aluminum(2) Cast 316 CRES(3) Nitronic 50(1) A-286(1) A-356 Aluminum(2) Cast 316 CRES(1) 6061 Aluminum(2) 347 CRES(3) Nitronic 50(1) 15-5 PH H1150M(2) INCO 718(1) 440 C CRES

ORIGINALOF POOR QUALITY

TABLE III-I



Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15,000 1bf

475.4 sec.

LOX/LH2

1200 thermal cycles

10 hrs.

6.0

574.4 Ibm

Components: LHL TPA (Hi Speed)





Envelope (Length)

Weight

Pump 49 - 2531 psiaTurbine 2344-1522 psia

Pump 40°RTurbine 535°R547 GPM

11.05" x 7.52"

26.3 Ib

Material(s) • Turbine & Pump Housing (one piece) (1) Cast 316 CRES(2) Nitronic 50

• Pump Impellers (1) Cast Ti BA1 2.5Sn ELI• Stators/Diffusers (1) Cast Ti 5A1 2.5Sn ELI• Pump Inlet (1) Cast Ti 5A1 2.5Sn ELI9 Pump Inducer (1) Cast or Wrought Ti 5A1

2.5SN ELIe Pump Shaft (1) Wrought Ti 5A1 2.5Sn ELIt Turbine Exhaust Housing (1) Cast 316

(2) Nitronic 50o Turbine (1) A-286 CRES• Bearings (1) 440 CRES

A-14

OR1GSNAL PAGE £8OF POOR QUALITY

TABLE III-I



Chamber Pressure

Thrust (Vac)

I sp ( M i n i m u m ) ( v a c )

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15,000 lbf

475.4 sec.

LOX/LH?

1200 thermal cycles

10 hrs.

6.0

574.4 Ib.m

Components : Misc. Valves & Pneumatic Pack

Operating Pressure(s) 4000 psia

Operating Temperature(s) Ambient

Propellant Flowrate(s) None


Envelope (Length)

Weight 12.6 Ibm

Material(s) Titanium

A-15

ORIGINAL PAGE BPOOR



Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15,000 1bf

475.4 sec.

LOX/LH?

1200 thermal cycles

10 hrs.

6.0

574.4 Ib.m

Components : Lines





Envelope (Length)

Weight

Material(s)

18 - 1500 psia

Tox = 140°RTf = 40°R - 535°RWox = 27.05 Ib/secWf = 4.51 Ib/sec

27.0 Ibm

Ti t an ium

A-16

ORIGINAL PAGE ES-OF POOR QUALITY

TABLE III-I



Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propel lants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15,000 1bf

475.4 sec.

LOX/LH?

1200 thermal cycles

10 hrs.

6.0

574.4 lb_

Components : Ignition System





Envelope (Length)

Weight

Material(s)

1200 psiaTf = 550°RT = 140°R

te; -

9.2 Ibm

SS/Nickel

ecsec

A-17

TABLE III-I

COMIPONENT _RiE_QUJJRIEME_N_TS


Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15,000 1bf

475.4 sec.

LOX/LH?

1200 thermal cycles

10 hrs.

6.0

574.4 Ib.m

Components : Miscellaneous*





Envelope (Length)

Weight

Material(s)

*Electrical Harness - 12.5 IbService Lines - 6.5 Ib

TPA Protective Bulkhead - 0.4 Ib

Attachment Hardware - 15.0 Ib

Instrumentation - 2.6 Ib

37 Ib

A-18

ORIGINAL PAGE ESOF POOR -

TABLE III-I



Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propel 1 ants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15.000 1bf

475.4 sec.

LOX/LHp

1200 thermal cycles

10 hrs.

6.0

574.4 Ib,m

Components : Engine Controller





Envelope (Length)

Weight

Material(s)

Ambient

Ambient

16.8" x 10" x 8"

35 Ibm

Aluminum

A-19

ORIGINALOF POOR

ABLECOMPP.NENT_ REQl) IREMEN1 S


Chamber Pressure

Thrust (Vac)

Isp (Minimum) (vac)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

1200 psia

15,_OOCMbf

475.4 sec.

LOX/LH?

1200 thermal cycles

10 hrs.

6.0

574.4 Ib

Components,: Heat Exchanger





Envelope (Length)

Weight

Material(s)

1500 - 2300 osia

535°R

Very small amount

5.0 lb

A-20

ORIGINAL PAGE mOF POOR QUALITY

LOX/LCH4 ENGINE (CYCLE C)


A-21


TABLE III-I


Baseline Engine: LOX/LCH4 Engine (Cycle C)

Chamber Pressure

Thrust (S.-L.)

Isp (Minimum) (S.L.)

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

4300 psia

600.000 Ib1

309.1 sec.

LOX/LCH4

100

10 hrs.

2.82

5075 Ib-m-

Components : Gimbal System





Envelope (Length)

Weight

Materials)

Transmits 600K lbf thrust

Ambient

6.6" Long

207 Ibm

Tit/SS

A-22

POOR, QUALITY

TABLE III-I



Chamber Pressure 4300 psia

Thrust (S.L.) 600.000 lbf

Isp (Minimum) (S.L.) 309.1 sec.

Propel!ants LOX/LCH4

Duty Cycle (Burns) 100

Cumulative Life 10 hrs.

Mixture Ratio 2.82

Weight 5075 lbm

Components; Injector





Envelope (Length)

Weight

Material(s)

PC = 4300 psia

Wox = 1432.7 Ib/secWf = 508.43 Ib/sec

D = 15.4" L = 10.2"

611 lbm

Body - Inconel 625 or ARMCO Nitronic ~

Manifolds - CRES 347 or Nitronic'-50

Injector Face - Inconel 625

A-23


TABLE III-I



Chamber Pressure

Thrust (S.L.)

Isp (Minimum) (S.L. )

Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

4300 psia

600,000 Ib,

309.1 sec.

LOX/LCH4

100

10 hrs.

2.82

5075 Ib•m-

Components : Combustion Chamber





Envelope (Length)

Weight

Material(s)

PC = 4300 psia

WT = 1784.51 Ib/sec

L1 = 15.5" 0 = 15.44"

428 Ibm

Zirconium Copper

A-24

-ORIGINAL -PAGE 13OF POOR QUAUTY

TABLE III-I



Chamber Pressure

Thrust (S.L.)


Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

4300 psia

600.000 Ib1

309.1 sec.

LOX/LCH4-100

10 hrs.

2.82

5075 lb_

Components: Nozzle



Propel 1 ant Flowrate(s)


Envelope (Length)

Weight

Material(s)

6 psi

1580°R

WT = 1784.51 Ib/sec

L = 111.2" Dex = 76.4"

264 Ib

Nitronic 50 or A-286

A-25

OF POOR QUALITY

TABLE III-I




Thrust (S.L.) 600,000 1b


Propellants. LOX/LCH4

Duty Cycle (Burns) TOO


Mixture Ratio 2.82 '.

Weight 5075 lbm

Components : Gas Generator





Envelope (Length)

Weight

Material(s)

PC = 4200 psia

186p°RWox = 44.75 Ib/secflf = 111.87 Ib/sec

D = 10.5" x L = 14.3"

76 Ibm

Inj. body - Nitronic 50

Chamber - Inconel 625

A-26

ORIGINAL PAOE ISOF POOR QUALITY

TABLE 111-1


Baseline Engine: LOX/LCH. Engine (Cycle C)


Thrust (S.L.) 600,000 U


Propellants LOX/LCH4

Duty Cycle- (Burns) 100


Mixture Ratio 2.82

Weight 5075 1b

Component^ : Oxidizer Valves





Envelope (Length)

Weight

Material(s)

Up to 5300 psi

Cryo

Main Wox = 1432.7 Ib/secGG flox = 44.75 Ib/sec

Pump Valve D = 3.1"GG Valve D = 2"

,155 Ibm

Aluminum and A-286

A-27

ORIGINAL PAGE 13OF POOR QUALITY

TABLE III-I



Chamber Pressure

Thrust (S.L.)


Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

4300 psia

600.000 lbf

309.1 sec.

LOX/LCH4-

100

10 hrs.

2.82

5075 Ibm-

Components : Fuel Valves





Envelope (Length)

Weight

Material(s)

Up to 8000 psi

CryogenicMain Wf = 508.43 Ib/secGG ftf = 111.87 Ib/sec

Main Pump Valve DiamGG Valve Diam = 2.7"

= 3.4"

133 Ibm

Aluminum & A-286

A-28

ORIGINAL PAGE ISOF POOR QUAUTY

t

TABLE 111-1




Thrust (S.L.) 600,000 lbf


Propel 1 ants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

309.1 sec.

LOX/LCH 4-

100

10 hrs.

2.82

5075 Ib

Components : Oxidizer Boost Rump





Envelope (Length)

Weight

Materials)

= 26'7"' 27'2"

PD = 231 psia

Tox = 166°R PumpTox = 166°R TurbineWpump = 1 3 9 6 . 5 Ib/secUTurb = ?1R

Line Inlet D = 15.4" Outlet D =8 .2 "

311 Ibm~env

Shaft Inconel 718, 15-5 PH H1150MImpeller & Turbine 7075 T-73, Al AlloyHousing A356 T6, Al AlloyBolts A-286Housing Liner FEP Teflon Fused CoatingBearings CRES 440C; Haynes Star J Alloy PM

A-29

ORIGINAL PAGE KTOF POOR QUALITY

TABLE III-I




Thrust (S.L.) 600,000 lbf


Propel 1 ants ' LOX/LCH4




Weight 5075 1b

Components : Fuel Boost Pump





Envelope (Length)

Meight

Material(s)

Denvenv

20.9"21.9"

PD = 135 psia

207°R

W = 495.4 Ib/sec= 99 Ib/sec

Line Inlet D = 10.3" Outlet 0 = 7.3"

.103 Ibm

All materials.same as low speed LOX TPA

except teflon coating is not required.

A-30

ORIGINALOF POOR

TABLE III-I



Chamber Pressure

Thrust (S.L.)


Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

4300 psia

600.000 1bf

309.1 sec.

LOX/LCH4-

100

10 hrs.

2.82

5075 Ib-m-

Components : Oxidizer Main Pump



Propellant Flowrate<s)


Envelope (Length)

Weight

D = 26.9"envL = 36"env

Material(s) ShaftImpellerHigh-PressurePump & Turbine Hsg.Inducer HousingTurbinesBolts (pump)Bolts (turbineBearings

5262 psiaT = 166°Rpumpiurb =

wpump = 1396-5 lb/sec

•^ttrtr-^—1-56.61 lb/secDLin in = 8'2 1n'irt jt = 3.1 in

895 Ibm

A-286Inconel 718ARMCONitronic-50Inconel 718Inconel 718A-286WaspaloyCRES 440C

A-31

TABLE III-I


Baseline Engjng.: LOX/LCH4 Engine (Cycle C)

Chamber Pressure

Thrust (S.I.)


Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

4300 psia

600.000 lbf

309.1 sec.

LOX/LCH^

100

10 hrs.

2.82

5075 Ib.•m-

Components: Fuel Main Pump





Envelope (Length) Denv = 23.3 in.L = 34 in.

Weight env

Material(s)

7953 psia

W = 495.4 Ib/sec=156.61 Ib/sec

PL«n ' 7-3"

661 Ibm

3Lout = 3.4"

5 Al - 2.5 Sn ELI Titanium Alloy

All other materials the same as .high speed LOX TPA

A-32

. ORIGINAL PAGE'g&--- _OF POOR QUALITY

' ' 4

TABLE 111-1




Thrust (S.L.) 600.000 1


Propellents

Duty Cycle. (Burns)

Cumulative Life

Mixture Ratio

Weight

309.1 sec.

LOX/LCH4-100

10 hrs.

2.82

5075 Ibm-

Components : Hot Gas Manifold


Operating Temperature{s)



Envelope (Length)

Weight

Material(s)

363 psia

1780°R

156.61 Ib/sec

D = 6.6"

23 Ibm

Inconel 625

A-33

SSS

TABLE III-I


Baseline Engine: LOX/LCH, Engine (Cycle C)


Thrust (S.L.) 600.000 Ib



Duty Cycle (Burns) IQQ



Weight 5075 Ib

Components: Low Pressure Line





Envelope (Length)

Weight

Materials)

15 psia Ox24 psia CH4

Cryotax = 1396.5 Ib/secftf = 495.4 Ib/sec

Dox = 15.4"Df = 10.3"

253 Ib

A-34

ORIGINA^ PAGE-JS_OF POOR QUALITY

TABLE 111-1


Baseline Engine: LOX/LCH, Engine (Cycle C)

Chamber Pressure

Thrust (S.L.)


Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

4300 psia

600.000 lbf

309.1 sec.

LOX/LCH4

100

10 hrs.

2.82

5075 lb_

Components : High Pressure Lines





Envelope (Length)

Weight

Material(s)

4000-8000 psia

166°R - 1860°RWnY = 1396.5 Ib/secft° = 495.4 Ib/sec

Varied - See Calcs

383 Ibm

A-35


TABLE 111-1




Thrust (S.L.) . 600,000 lbf


Propel 1 ants LOX/LCH4

Duty Cycle (Burns) 100


Mixture Ratio 2.82

Weight 5075 lb[n

Components : Ignition Systems





Envelope (Length)

Weight 40 Ibm

Material(s) '

A-36

ORIGINAL PAGEJSOF POOR QUALITY

TABLE III-I




Thrust (S.L.) 600,000 1bf






Weight 5075 1bm

Components: Misc. (Frames, fastener,harness, instrumentation, etc),





Envelope (Length)

Weight . 17'4 Ibm

Material(s)

A-37

OF POOR QUALITY

TABLE III-I



Chamber Pressure

Thrust (S.L.)


Propellants

Duty Cycle (Burns)

Cumulative Life

Mixture Ratio

Weight

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600.000 Ib1

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2.82

5075 Ib

Components: Controller





Envelope (Length)

Weight

Material(s)

Ambient

Ambient

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Aluminum

A-38


TABLE III-!


Baseline Engine: LOX/LCH4 .Engine (Cycle "C)


Thrust (S.L.) 600.000 Ib.



Duty Cycle.(Burns) 100


Mixture Ratio 2.62 :

Weight 5075 Ib

Components : Pressurization System





Envelope (Length)

Weight

Material(s)

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A-39


TABLE III-I



Chamber Pressure

Thrust (S.L.)


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Mixture Ratio

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5075 Ibm-

Components : Interpropellant Seal





Envelope (Length)

Weight

Material(s)

5262 psia

Pump - 166°RTurbine - 1860°RPump - 1396.5 Ib/secTurhinp - 156.61 lb/S6C

90 Ibm

A-40

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APPENDIX E

VENDOR TRIP MEMOS

E-l

AerojetLiquid RocketCompany


REPORT NO MA-82-144iviciiericii£> Mnciiy&i£»'ORK REQUESTED

UBJ6CT

ROCRAM

ART NAME

BY

D. C. Judd

DATE 25 June

OEPT:

9733

Trip Report - Composite Material Fabricators(June 21-22)

Composite Material Applicationsto Liquid Rocket Engines

Conceptual Desiqn Assessment

W.O. NO.

2418-03-000

PART NO.

Task MlS/N

PAGE 1 OF

1982

4

TABLES: 2

FIGURES:

ENCLOSURES:

MATERIAL

Composites

HJRPOSE:

To meet with fabricators of composite materials to present ALRC

conceptual designs for composite parts, to obtain their comments on

producibility, and to assess their fabrication capabilities.

SUMMARY:

Six contractor facilities were visited during a two day trip to Los

Angeles. In these meetings it was explained that a significant weight savings

(between 30 and 40 percent) can be realized by substituting composite

materials for metals in liquid rocket engines. All six contractors indicated

that they would be interested in helping ALRC develop rocket engine parts.

ISTRIBUTION

DL KorsDE LemkcRW MichelCJ O'Brien

HO Davis (ASPC)REPORT BY

E. W. CarterREVIEWED BY

G. R. Janser ^.

V TRICK WANAGcH MATERIALS ANAI/SIS SECTION

E-2

ORIGINAL PAGE IS MA-82-144OF POOR QUALITY Page 2

INTRODUCTION:

The current NASA contract to study composite material applications in

liquid rocket engines has provisions for recommending designs for a minimum

of two parts for a follow-on development program. A description of this

follow-on program is required including detail fabrication plans. During the

visits to the six subcontractor facilities, subcontractor participation in the

follow-on program was discussed. The preliminary designs of seven rocket

engine parts to be made with composites were reviewed. Each contractor was

asked to assess his capability to make the parts together with anticipated

processing difficulties.

DISCUSSION:

The preliminary designs reviewed by the six subcontractors fall into

three categories of producibility:

(1) Parts whose fabrication is straightforward.

(2) Parts that require a process development program to establish

parameters for controlling fiber geometry.

(3) Parts that require redesign in order to successfully fabricate (Note

1A, Table I).

Table I summarizes the position of each subcontractor on these parts; it

shows the processes available at each facility for manufacturing the part and

the contractor's opinion of producibility. For five of the designs (1196001,

1196002, 1196003, 1196005, and 1196007), the complexity of the parts and the

lack of design detail make the assessment of producibility difficult. Two of

these designs (1196003 and 1196005) had enough obvious difficulties to

establish that a redesign would be necessary. It was assumed that the

redesigned part would be satisfactory for manufacturing if the subcontractor

consulted with the designer.

E-3

Swedlow Inc. indicated that they could assume responsibility for part

design. This suggestion was made during the review of the injector housing

(1196007). They feel that the high pressures and propellent flows dictate

extraordinary methods of structural analysis.

Four subcontractors (Swedlow, HITCO, Reynolds and Taylor, and M.C.

Gill) have excellent facilities for producing parts using composite materials.

Two contractors (Swedlow and HITCO) also have the capability to design with

composite materials.

The subcontractors have indicated that the response time to an RFQ for

participation in a development program would be three weeks or less. To

prepare a quote, they would require an SOW and a detail design of the part

to be developed. Additional technical meetings would be helpful to discuss

program requirements and familiarize the subcontractors with rocket engine

components. HITCO made the suggestion that one development part might be

a mock-up combining the design features and technology needs of several

parts. This approach would concentrate the objectives of the program in one

part and set of tooling to reduce costs.

Table II lists the persons contacted during the visits along with the

product lines of the companies.

CONCLUSIONS:

1. All six contractors have potential to help ALRC build rocket engine

parts out of composite materials.

2. HITCO, Swedlow, and M.C. Ciil are large diversified manufacturers

with significant manufacturing research capability. Reynolds and Taylor's

capabilities emphasize job shop production.

3. Poly-Trussions and Fibco are limited respectively to pultrusion or

vacuum bag and compression molding.

E-4

Page U

RECOMMENDATION:

1. Prepare an SOW defining subcontractor support requirements for

follow-on composite effort.

2. Coordinate SOW with candidate subcontractors and obtain quotations

for inclusion in follow-on proposal to be prepared in Task V (7/15-9/15).

E-5

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E-18

AerojetLiquid RocketCompany


Materials Analysis REPORT MQ. MA-82-156

DATE 21 September 1982

WORK REQUESTED

SUBJECT

PROGRAM

BY

D. C. Judd

Application of Metal Matrix Compositesto Selected Engine Parts

Composite Materials Applicationto Liquid Rocket Engines

P A R T NAME

OEPT:

9772

W.O. NO.

2418-05-000P A R T NO. S/N

PAGE 1 OFH

TABLES:

FIGURES:

ENCLOSURES:

MATERIAL

PURPOSE:

To consult with metal matrix composite fabricators to determine the

state-of-the-art for application to selected liquid rocket engine parts.

DISTRIBUTION

EW CarterD CulverDL KorsR MichelRO Schwantes

REPORT BY

G. R. JanserREVIEWED BY

APPROVED BY

E-19

MA-82-156Page 2

DISCUSSION:

The writer and Ed Carter visited Nevada Engineering and Technology

Corp. (NETCO) of Long Beach, California, and DWA Composite Specialities.

Inc. of Chatsworth, California, on 10 September, for consultation with regard

to the application of metal matrix composites to the subject program.

NETCO

The NETCO contacts were Leroy Davis, President, and Jack Williamson.

The purpose of visiting NETCO was twofold. In addition to the aforementioned

consultation, NETCO is expected to be a source for mechanical and physical

property data for proposed designs. NETCO is under contract to the

Department of Defense Metal Matrix Composites Information Analyses Center

(MMCIAC) to collect, evaluate, store and disseminate metal-matrix property and

processing data. Mr. Davis was asked to provide MMCIAC cyrogenic property

data for support for the design activity of the anticipated composites follow-on

contract. He stated that there was very little cryogenic information; and since

this activity is just getting underway, the material is not fully organized for

immediate retrieval. Mr. Davis is placing the writer on distribution for the

MMCIAC bulletin and will determine the availability of cryogenic data at a later

date.

When the drawings were presented; Messrs. Davis and Williamson stated

that they preferred not to comment on the viability of the design from the

standpoint of fabrication from metal matrix composites due to their complexity

as compared to their experience. They recommended consulting with either

AVCO, DWA or MCI for fabricability information. They believed that DWA

would be the best source for this information and did offer an opinion that

fabrication would be very difficult.

Mr. Williamson wanted to comment on the design with regard to reinforced

plastic (RPC) materials since he is also marketing director for Reynolds and

Taylor, a company which had declined to respond to an RFQ from ALRC on the

subject RPC designs. He said he would persuade his associates at Reynolds

and Taylor to reconsider their no-bid position.

E-20

^ MA-82-156ORIGIN*- P Page 3OF POOR

DWA

The DWA contacts were Joe Dolowy, President, Bill Harrigan, General

Manager-Operations and Roy Levin, Manager-Manufacturing Development.

After the writer and Ed Carter described the composites contract, Joe Dolowy

stated they were already somewhat familiar with the program since they had

consulted with Rocketdyne on their parallel effort. Their examination of the

drawings resulted in a unanimous opinion that the parts could not be

fabricated by laying-up metal matrix laminates and diffusion bonding. They

said that not only could the parts not be layed up due to the small and

varying radii, but that hot gas isostatic pressing (HIP) for diffusion bonding

would be a development program in itself. They specifically stated that the

most complex part, the pump housing, was well beyond the present state of

metal matrix composite fabrication capability. It was also their opinion that for

complex shapes with multi-directional stress distributions, boron-aluminum

becomes inefficient and that graphite-metal composites present the same

problem in addition to the problem of very low transverse and shear strength.

They proposed fabricating from their proprietary aluminum-silicon carbide

reinforced composite material, DWA-AI. This material is a powder metallugry

product, which is consolidated into semi-finished stock prior to being

processed by conventional metal shaping techniques such as forming,

extrusion, forging and machining. The material possesses poor weldability by

the CTAW process and is considered by DWA to be unweldable by the electron

beam process. Weld deposits are poro'us and have the same properties as

those deposited on unreinfoced aluminum alloys. DWA stated that they have

made good progress in reducing weld porosity, and presented

photomicrographs showing porosity limited to the weld deposit heat affected

zone interface. They expect to further improve weld soundess and are

currently developing composite weld rods for producing reinforced, higher

strength weld deposits. This development has led to their initiating a casting

program. DWA could not provide assurance that these technologies would be

sufficiently developed for application for our follow-on contract.

The writer inquired whether the aluminum graphite composite, which can

be tailored to match the expansion coefficient of the RFP material, could be

applied as liners for the interior of RFP parts. Their response was negative.

E-21

MA-82-156Page 4

Mr. Dolowy said that DWA would provide a written cost estimate for the

one part, the valve body. He would propose to machine the part from an

extrusion and gave a preliminary estimate of $50,000.

CONCLUSIONS AND RECOMMENDATIONS:

1. Fabrication of the candidate parts by diffusion bonding of metal

matrix composite laminates would be very difficult and is considered a high

risk item.

2. A material such as the DWA-A1 composite offers the lowest risk in

the application of metal matrix materials to the unwelded, candidate parts.

Welded designs must account for reduced strength and some weld porosity.

E-22

L.ORIGINAL mOF POOR QUALITY

composite specialties, Inc.

16 September 1982

Aerojet Liquid Rocket CompanyP.O. Box 13222Sacramento, CA 95813

Attention: Messrs. Ed Carter and George Janser

Subject: Meeting at DWA 10 September 1982

Gentlemen:

With regard to your visit of last week, we have reviewed thestructures discussed and have reached similar conclusions tothose reached during the meeting. Of the five structuresconsidered: 1) the LCH. High-speed TPA appears to be wellbeyond the present state of MMC fabrication capability; whentruly "castable" forms of discontinuous composites becomeavailable, large toroidal or volute type structures willbecome do-able. 2) & 3) The Support ring and nozzle extend-ing shafts were only considered in passing, since they didn'toffer any appreciable weight saving or technology challenge;both could be impacted with metal matrix. 4) Injector Housingrepresents a significant challenge to present discontinuous ^reinforced MMC fabrication technology. Utilizing forged DWAl 20for the bottom third and top third of the part while handlingthe oxidizer manifold as a separate problem seems most efficient(fabricate as a separate "pressure-vessel" type structure,utilizing conventional material, or possibly superplasticallyformed DWAl 20 ). Forged "cup"^"complex-base" cylinder shapeshave been produced with DWAl 20 . The most significant problemwith the injector housing would evolve from assembling the threesubcomponents; brazing and welding both seem reasonable, but adevelopment program would be, advised. If necessary, selectivecircumferential reinforcing could be applied, using graphiteepoxy; this could also aid in assembling the subcomponents.5) Housing-valve, propellant - represents the second structurediscussed 10 September, with reasonable options for MMC demon-stration. Extruded, heavy-wall cylinders represent a straight-forward approach, but necessitates significant machining tocreate bosses, attach-points and flanges. An alternative fabri-cation process could utilize ring-rolling to decrease the requiredmachining, but this would necessitate higher tooling costs.

In reviewing the Agrojet selected hardware, my associates andI felt the DWAl 20 (isotropic particulate-reinforced P.M. com-posite) would be most appropriate. Although our experience withboron-aluminum-type composites exceeds all other material systemswhen axi-symmetric, rather complex-shaped parts (with multi-directional stress distributions) are desired, B-A1 tends to

E-23SM33 scipeRfOR stRset, chzatscooRthx ca. 91311 (sisj 998-1504.

Aerojet Liquid Rocket Company -2- 9/16/82

• ORIGINAL. PAGE-IS-OF POOR QUALITY

become inefficient. Also, the complexities of fabricationwith continuous, stiff fibers limit available processes.The graphite-metal composites were not considered for similarreasons, with the added problem of very-low transverse andshear strength. The ability to utilize many conventionalmetal-working processes and to yield isotropic mechanicalproperties, has finalized the recommendation of DWA1 20materials.

In response to questions generated during the meeting, attachedare several tables showing mechanical properties and costs ofvarious MMC systems. A generalized development effort for thevalve housing might require five or six months, with initialextrusion experiments yielding 8-inch long x 2.2" ID x 4.8" ODsamples. Subsequent machining would create the final product.A $35K to $45K effort should be sufficient.

The injector housing structure, which would involve a muchgreater analysis effort, separate tools, and an assembly taskwould necessitate a ten to twelve-month effort. Initially, thedesign concept and part sizes would be finalized with the selec-tion of material for the oxidizer manifold. Forging, or hot-forming tools for the top and bottom sections will be produced;then axisymmetric tool-verification parts generated. Mechanicalproperty levels verified from the "tool try" parts followingNDT, would create mechanical-property data. The assembly tech-nique would be demonstrated (braze, or weld with option of Gr-Epoxy overwrap, or adhesive bond) with a final assembly, prooftest, and inspection sequence. This program would requiresupport to about $100K.

I hope your MMC study program with NASA is a success. We at DWAwould be happy to support Aerojet Liquid Rocket in any design,trade-off, or prototype hardware efforts using any metal-matrixcomposite materials. We look forward to your comments on thisletter.

Very truly yours,

DWA COMPOSITE SPECIALTIES, INC.

J.F. Dolowy, Jr.President

JFD:lsS0611

cc: T.D. Lynch

E-24

APPENDIX F

TECHNOLOGY NEEDS DEFINITION FORMS

F-l

TECHNOLOGY NEED DEFINITION

ITEM:

1.0 - H Compatibility

DESCRIPTION:

Select composite materials that are not degraded by exposure to

hydrogen.*

ASSESSMENT OF RISK:

l

High * Medium Low

* Freeze-thaw cycling of hydrogen rocket propellant (and othervapors) trapped in composite material has the potential tocause severe structural damage. The glass transitiontemperature of epoxy resin used in high performancestructural composites have glass transition temperatures aboveambient. At use temperatures below the glass transition,mechanical properties show improvement. Also at cryogenictemperatures, hydrogen is chemically unreactive and not acause of resin degradation.

APPROACH:

The effects of mechanical property degradation resulting from

freeze-thaw cycling can be demonstrated. The rate of degradation

should depend upon the permeability of the composite. It is believed

impractical to control mechanical property degradation by resin

selection.

PROPOSED SOLUTION:

Barrier coatings, liners, etc.

F-2


ITEM:

2.0 - O2 Compatibility

DESCRIPTION:


LOX.

ASSESSMENT OF RISK:

High * Medium Low

*1. If freeze-thaw (also condensation) cycling occurs.2. Most organic polymers used as matrices in high performance

composite materials are expected to fail LOX impact requirements(MSFC-SPEC 106).

APPROACH:

Poly (amide-imide) resin matrix may pass LOX impact requirements.

It is believed impractical to control degradation caused by

freeze-thaw cycling through material selection.

PROPOSED SOLUTION:


F-3


Compatibility

DESCRIPTION:


methane.

ASSESSMENT OF RISK:

High * Medium Low **

* If freeze-thaw (also condensation) cycling occurs.** 1. Temperature too low for chemical degradation.

2. Methane absorption will not make worse the effects of anypolymer transitions occurring because of low temperature.

APPROACH:'

It is believed impractical to control degradation of mechanical

properties due to freeze-thaw cycling through material selection.

PROPOSED SOLUTION:


F-4



ITEM; •

4.0 - ;Low Temperature Toughness

DESCRIPTION;

Select composite materials that are thermal shock and impact

resistant.

ASSESSMENT OF RISK:

High Medium Low *

* Composite materials have inherent toughness properties becauseof their energy absorbing fracture characteristics. Thesecharacteristics increase slightly at cryogenic temperatures to improvetoughness performance.

APPROACH:

Fiberglass reinforcement and graphite fiber reinforcement are used

successfully in cryogenic structural applications. Fiberglass is

preferred for applications requiring thermal insulation. Graphite is

preferred for applications requiring higher thermal conductivity or

greater stiffness. ,

PROPOSED SOLUTION:

Select composite materials on the basis of room temperature

structural properties and validate their structural performance at

cryogenic temperatures by conducting structural element tests.

F-5

TECHNOLOGY NEED DEFINITION j

ITEM: , [

5.1 -'Mechanical Fastening

DESCRIPTION: I

Bonded metallic inserts for composite materials may pull out because

of thermal contraction at cryogenic temperatures.

ASSESSMENT OF RISK:

High Medium Low *

* Mechanically anchored inserts are available that do not pull outeasier because of shrinkage.

APPROACH:

Review applications of mechanical fasteners to make sure that bonded

inserts are not used where insert shrinkage would result in an

attachment failure.

PROPOSED SOLUTION:

Avoid using bonded metallic inserts in composites at cryogenici

temperatures. '

F-6

POOR-•J.J-V

TECHNOLOGY NEED DEFINITION 4

"Ife.» v38p

ITEM;'W.,5.2 -'.si Finish of Fluid Sealing Surfaces

.

DESCRIPTION:. ' '-'-;#.

Smooth surfaces are required for O-rings and seals. Machining high

performance composite materials may not meet the finish requirementsfor seals.

ASSESSMENT OF RISK:

High Medium Low *

* Coatings can be used to improve the finish of machinedcomposite materials.

APPROACH:

Select coating compounds for use with lubrication seals.

PROPOSED SOLUTION:

Coatings to upgrade the finish of machined composite surfaces will be

used.i

F-7

ORIGINAL. PAGE ISOF POOR QUALITY


':..i$r.ITEM; "'I;-'

5.3 - Vane Manufacturing Methodi'

DESCRIPTION: 1

Select a manufacturing process to mold the TPA stator vanes.

ASSESSMENT OF RISK:

High Medium Low *

* Alternate methods for molding the stator vanes are available.

APPROACH:

Allow the manufacturer of the TPA housing to determine the best

manufacturing method for this part.

PROPOSED SOLUTION:

The stator vanes may be molded in-place (in situ) . or : molded

separately and bonded in place.

F-8



ITEM;

5.4 - Vane Attachment Method

DESCRIPTION;

An adhesive bonding process must be developed if the manufacturer

elects to mold these details separately.

ASSESSMENT OF'RISK:

High Medium Low

* Adhesives for bonding composite materials at cryogenictemperatures are available.

APPROACH:

The manufacturer of the TPA assembly will select an adhesive and

bonding process that will satisfy the structural requirements of the

design disclosure.

PROPOSED SOLUTION;

The shear stress between the stator vanes and the turbine manifoldi

shell may exceed the interlaminar s*hear capability of the composite.

In this case more composite material will be required to reduce this

stress.

F-9

ÔF POOR QUALITY

TECHNOLOGY NEED DEFINITION-' '

• . , . t

; '•• -,- '•^^f^lW'--- / 'ITEM: ;,rj , . .,. .,v .

5.5 - Residual Thermal Streses " " " : - •ti

DESCRIPTION: ' • ' • . , ' > : v %€ I:'" , . " • . . ' . ,;.V<2?-"~ ;

Determine the effect that the cure cycle has on the residual stress I

of each design.

ASSESSMENT OF RISK;

High Medium * Low

* Distortion and .ply failure due to thermal stress at cryogenictemperatures can be avoided by controlling the fiber orientation.

APPROACH:

Analyze the effects that the fabrication process has on residual

stress. Design the process and laminate to minimize thermal

distortion and failure.

PROPOSED SOLUTION:

Control residual thermal stress through design analysis and process

control. f

F-10

ORIGINAL PAGE S§OF POOR"'QUALITY ^

TECHNOLOGY NEED DEFINITION•

ITEM: •'?*' ' ' ' "!'" " ' -

5.6 - Barrier Coating Process

DESCRIPTION:

High performance plastic composites must be sealed in order to

contain liquids and vapors.

ASSESSMENT OF RISK;

High * Medium Low

* Microcracking occurs in the resin matrix of high performancecomposite materials. These cracks result from residual andapplied stresses. They render the resin matrix permeable tovapors and liquids.

APPROACH;

Several barrier coatings (metal foils, plastic films, etc.) have been

used successfully to seal composite materials. These coating

materials and processes will be evaluated to select the optimum

system for a given rocket engine component.

PROPOSED SOLUTION:

Ductile barrier coatings are required to seal composite materials.

F-ll

TECHNOLOGY NEED DEF.NIT.ON


ITEM: ;;; "'. - '

5.7 - Plumbing Connections.. • vlf:-

DESCRIPTION:.:-^!

Propellant line connections are a primary source of propellent leaks.

Composite materials are expected to cause more leak problems than

metals because they scratch and distort more readily.

ASSESSMENT OF RISK:

High Medium Low *

* Viton, Kryton, Teflon, Kel-F, and metal foils are materialsavailable to seal propellant line connections.

APPROACH:

Evaluate seals and O-rings to determine the most effective method to

produce leak-resistant plumbing connections.

PROPOSED SOLUTION;

More extensive use of metals in flanges will be necessary ifi

leak-resistant connections using corrtposites fail.

F-12

ORIGINAL PAGE


- • • • . " • > .' 5 .8 - Adhesive Bonding

DESCRIPTION^ , , /* i :^

v'Xdhesives and adhesive prepregs which join composite materials in

liquid rocket engine applications must satisfy the same compatibility

• requirements as the composite.

ASSESSMENT OF RISK:

High * Medium Low

* Melt processible PCTFE and poly (amide-imide) might providepropellant compatible adhesive systems (TBD). Otherwisebarrier coatings will be required.

APPROACH:

Freeze-thaw cycling damage is not as serious a problem for adhesive

bonding applications because resin microcracking does not occur.

Propellant exposure tests may indicate that a protective barrier will

only be required for LOX.i

4PROPOSED SOLUTION:

Barrier coating, liner, etc.

F-13

ORIGJJMAL PAGE ISPOOR QUALITY-


ITEM:5.9 - Fabrication Methods

DESCRIPTION;

Alternative methods are available to mold composite materials(compression, autoclave, resin injection, etc.).

ASSESSMENT OF RISK;

High Medium Low *

* An analysis of the structural effects of alternative processesmay indicate a change in the structural capability of the part.

APPROACH:

The effects of processing must be validated in structural tests. A

trade-off between structural performance and producibility will be

made.

PROPOSED SOLUTION;

Revised design allowables will be issued -if the choice of a processi

affects the structural properties of^he material.

F-14



. . . . v - ~. -y^ • I

ITEM; I

5.10 - Mold Design ''

DESCRIPTION;

A mold to produce composite parts should be capable of meeting all

the design objectives specified for the part.

ASSESSMENT OF RISK:

High Medium * Low

* Analysis of the mold performance before building the mold willreduce costly iterations in process and mold design to solveproblems involving producibility.

APPROACH;

An analysis of the mold performance should be made to determine if

any design changes are needed. The analysis should investigate the

effects of compaction, flow, temperature rise, cure uniformity, cool

down, and concentration of reinforcement and resin on residual

thermal stress, shrinkage, and voids.

PROPOSED SOLUTION;

Redesign the mold until analysis indicates that the part is

satisfactory or lower the part specification to reflect what is possible

before releasing the mold design for fabrication.

F-15



ITEM: .. ,. . /.

6.0 - Cryogenic Properties

DESCRIPTION: :: . .V6v-; '• • ''

Some commercial composite materials have useful structural < propertiesat cryogenic temperatures. Their outstanding fracture toughness isdue to microcracking of the resin matrix to relieve residual thermal

strain.

ASSESSMENT OF RISK:

High . Medium Low *

* Commercial composite materials have met the requirements ofapplication at cryogenic temperature because of good fracturetoughness and moderate increases in mechanical properties.

APPROACH;

Candidate composite materials will be tested at cryogenic

temperatures to determine their structural properties. Structural

tests following cryogenic temperature cycling will also be conducted

to evaluate the stability of the fibeif matrix bond.

PROPOSED SOLUTION:

Commercial composites will not be applied beyond their temperature

limits. Efforts will be made to obtain composite materials having

improved fracture toughness properties at cryogenic temperatures.

F-16

OF POOR

TECHNOLOCY NEED DEFINITION

ITEM: '- '

7.0 - Interface Properties

DESCRIPTION:

Many kinds of interfaces occur in composite materials. Interfacial

instability may occur because of corrosion, stress, debonding, etc.

ASSESSMENT OF RISK;"™--"L

High * Medium Low

* The low temperature stability of adhesive interfaces because ofthermal stress is a familiar problem, especially acute for metallicinterfaces.

APPROACH:

Tests will characterize the stability of composite material interfaces

in liquid rocket engine applications. These tests will determine

which kinds are acceptable in design and which are not.

PROPOSED SOLUTION:

Unstable interfaces can not be usjted because of the potential for

progressive structural failure. Metallic parts would replace

composite parts in any applications involving unstable interfaces.

F-17



ITEM:

8.0 - Metal Coating Interface Properties

DESCRIPTION;

Cycling temperature and pressure can fail the metallic barrier

coating protecting the composite. Two techniques may impart more

durability to the protective barrier: (1) Metallized plastic film, (2)

Elastomeric adhesive.

ASSESSMENT OF RISK:

High Medium Low

* Elastomeric adhesives have demonstrated low temperaturebonding capability to metal and metalized plastic film. Theirperformance at cryogenic temperatures needs to be determined

APPROACH:

Evaluate the bond stability by cycling the test specimens between

ambient and cryogenic temperature. Specimens consist of (1) metal

bonded to composite with an elastomeric adhesive and (2) metallized

film bonded to composite with an elastomeric adhesive. Failure of

the elastomeric adhesive bond will result in a more sophisticated

approach to stabilize the bond. A tie coat adhesive, approximating

the thermal expansion of the metal, will be evaluated with the

elastomeric adhesive.

PROPOSED SOLUTION:

Composite materials can not be used in applications requiring

protective coatings if the tests above and the supporting analysis

indicate that the barrier coatings fail to meet the strain cycling

requirements.

F-18

•

ORIGINAL PAGE*OF POOR QUALtTV

TECHNOLOGY NEED DEFINITION :

' ' ;:v^.-.v-::ife;.'!'''ITEM; • _ ,; :... .,"'.;-. ,' '

9.0 - Differential Expansion Properties

DESCRIPTION;

The thermal expansion'properties must be known in order to analyze

the thermal stresses affecting design.

ASSESSMENT OF RISK;

High Medium Low *

* The anisotropic thermal expansion coefficients (a and a )x ycan be used to calculate residual thermal stress.

APPROACH;

Thermal expansion data will be obtained from the material suppliers

or from ALRC lab tests.

PROPOSED SOLUTION:

Without this data the analysis of thermal effects to support design

will be less accurate. •*

F-19



ITEM:

- l ' • * • . ' • • ' >

4

. ' t

10.0 - Solar Radiation Effects

r i

DESCRIPTION: ,;

Protective coatings are needed to protect the surface of plastic

matrix composites from polymer degradation.

ASSESSMENT OF RISK:

High Medium Low

* The absorption of radiation by the reinforcement limits damageto the surface. In most applications the most serious effect is

cosmetic.

APPROACH:

Select protective coatings that are compatible with the rocket engineenvironment for use on surfaces exposed to solar radiation.

PROPOSED SOLUTION:

If coatings are not used to protect* against radiation, small losses in

strength will occur due to surface degradation.

F-20


ITEM::

11.0 - Low Cycle (Thermal) Fatigue

. DESCRIPTION:

Cycling temperatures can result in structural failure.

ASSESSMENT OF RISK:

'•a«KV /.-. .:,. .,;..-•,'. .-$&*?••• • ., •••;...•»

- . - . . - ' . H - ? » . , .

: - i v ; / f c - ; - ,

High Medium Low

* Distortion and ply failure due to thermal stress can be avoided

by controlling the reinforcement orientation.

APPROACH;

Conduct a thermal-stress analysis to identify any structural elements

that exceed allowable ply stresses.

PROPOSED SOLUTION:

Without this analysis the chances of LCF failures increase.

F-21

' ' ^JSSWS^^"OF POOR QUALITY ; ,

TECHNOLOGY^NEED DEFINITION

• V/. "

It:

• .*. -:-v:.ft.,i: ^»A; , ^-< • , •L-W:i ^-^r«^ ' ; ' '"V'-^»;/- -vtfV."^;.-- .- •

ITEM*

13.0 - Bearing Surface Lubricant3 •;..'- '.rv.i.' .

DESCRIPTION:

A lubricant applied to the nozzle extension shaft threads-will reduce^ ^: •-'.->,"'•

friction, stress, and wear.

ASSESSMENT OF RISK; , • - /

High Medium Low *

* Surfaces of plastic composite materials errode faster than metal

surfaces for a variety of reasons. Lubricants will reduce

friction and wear.

APPROACH;

Select a commercial lubrication system compatible with composite

materials, the liquid rocket engine environment, and the vacuum

conditions of space.

PROPOSED SOLUTION: <•'

Without a lubrication system, the useful life of the shaft will be

shortened due to excessive wear.

F-22

Documents

Report 2418-F 10 December 1982 NASA Composite … 2418-F 10 December 1982 NASA Composite Material Application To Liquid Rocket Engines (NASA-CK-170697) COMPOSITE flATEEIAL N83-15360