space shuttle thermal scale modeling application …

SPACE SHUTTLE

THERMAL SCALE MODELING

APPLICATION STUDY

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Contract No.: NAS 9-12991

DHL No: 6

DRD No: MA-129T

LMSC NO: D313907

SPACE SHUTTLE

THERMAL SCALE MODELING

APPLICATION STUDY

FINAL REPORT

CONTRACT NAS 9-12991

FEBRUARY 1973

BY

K. N. MARSHALL

W. G. FOSTER

LOCKHEED PALO ALTO RESEARCH LABORATORYI O C K H E E D M I S S I L E S & S P A C E C O M P A N YA G I O U P O I V I S I O N O F l O C K H f f D A I I C I A F T C O I P O I A T I O N

COMPEMPS

Section

Foreword i

Illustrations ii

Nomenclature ±v

1 SUMMARY •„ ' 1-1

2 INTRODUCTION ~ 2-1

2.2 Technical Content 2-2

2.3 End Product 2-3

3 SHUTTLE PHASE B ORBITER CONFIGURATION REVIEW 3-1

3.1 Shuttle Orbiter Configuration 3-1

3.2 Critical Thermal Control Areas 3-4

3.3 Level of Detail for Modeling Systems 3-4

4 THERMAL ANALYSIS 4-1

4.1 Thermal Analysis Requirements 4-1

4.2 Thermal Math Model Development 4-2

4.2.1 Payload Configuration 4-18

4.3 Orbital Conditions 4-22

4.4 TMM Results 4-22

4.5 Internal Power Sources 4-304.5.1 RCS and CMS 4-30

4.5.2 Fuel Cells : 4-31

4.5.4 Hydraulics 4-32

5 TSM DESIGN 5-1

5»1 Design Requirements 5-2

5»2 Conduction-Radiation Modeling 5-2

5.3 Convective Modeling 5-6

6 TSM TEST REQUIREMENTS 6-1

6.1 Solar Simulation 6-1

6.2 Test Facility 6-2

6.3 Test Requirements 6-2

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CONTENDS(Cont'd)

Section

7 POTENTIAL APPLICATIONS OF THE TSM 7-1

8 CONCLUSIONS ' 8-1

9 REFERENCES 9-1

APPENDIX A ANALYSIS AND STUDY PLAN A-l

1.0 Introduction and Summary A-2

2.0 Task of Work Breakdown and Schedule . A-3

3.0 Thermal Math Model Refinement and Calculations A-3

14-.0 Important Thermal Control Areas and TemperatureRequirements A-6

5.0 Trade Studies A-7

6.0 TSM Design Criteria Development A-8

7.0 TSM Design Requirements A-9

8.0 TSM Detailed Design Development A-9

8.1 Conduction-Radiation Modeling A-10

8.2 Convective Modeling A-ll

8.3 Level of Detail for Modeling Systems A-12

9.0 POTENTIAL APPLICATIONS OF TSM FOR USE IN SHUTTLETHERMAL DESIGN A-3A-

10.0 TEST FACILITY DEFINITION AND REQUIREMENTS A-15

APPENDIX B CONCEPT REVIEW HANDOUT ' B-l

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A G R O U P D I V I S I O N O F t O C K H E E D A I I C I - A F T ' C O R P O R A T I O N

FOREWORD

This report is submitted to the MSA Manned Spacecraft Center by the

Lockheed Palo Alto Research Laboratory, Lockheed Missiles & Space Company,

Inc. The work was performed under Contract HAS 9-12991 and was administered

"by the Manned Spacecraft Center, with Mr. R. A. Vogt as technical monitor.

The final report provides a technical summary of the work performed

from June 30, 1972 to October 10, 1972.

The authors wish to acknowledge K. W. McGee for his efforts in developing

the Thermal Mathematical Model used in the program and R. K. Wedel for his

part in establishing portions of the modeling criteria.

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A C R O U P D I V I S I O N O F L O C K H E E D A I R C R A F T C O R P O R A T I O N

ILLUSTRATIONS

Figure

3-1 Air Transportable Orbiter Vehicle Arrangement 3-2

3-2 IMSC Space Shuttle 040A Configuration 3-3

3-3 Shuttle Overall Dimensions 3-5

3-^ Original Equivalent t Values 3-6

3-5 Revised Equivalent t Values 3-7

I*-1 TMM External Surface Nodes (Preliminary) 4- 3

4-2 TMM External Surface Nodes (Preliminary) 4-4

14--3 TMM Internal LI-1500 Nodes (Preliminary) 4-5

4-4 TMM Internal LI-1500 Nodes (Preliminary) 4-6

4-5 TMM Structural Nodes (Preliminary) 4-7

4-6 TMM Structural Nodes (Preliminary) 4-8

14—7 TMM Payload Bay Nodes (Preliminary) 4-9

4-8 TMM Crew Cabin Section Nodes (Preliminary) 14-10

4-9 TMM Nodal Description 4-11

4-10 TMM Nodal Description 4-12

4-11 Space Shuttle Structural Detail 4-l4

4-12 Orbiter Flight Stations 4-19

4-13 Station Numbers for Typical Payload Configuration 4-20

4-l4 Steady State Temperatures, Hot Case Orbit, Upper Surface Sun Side 4-23

4-15 Steady State Temperatures, Hot Case Orbit,Upper Surface .Shade Side 4-24

4-16 Steady State Temperatures, Hot Case Orbit, Lower Surface 4-25

4-17 Steady State Temperatures, Hot Case Orbit,-Payload Bay Structure 4-26

4-18 Steady State Temperatures, Cold Case Orbit, Surface, LI-1500Midpoint and Structure Nodes 4-27

4-19 Steady State Temperatures, Cold Case Orbit Lower Surface 4-28

4-20 Steady State Temperatures, Cold Case Orbit Payload Bay Structure 4-29

5-1 Non-Dimensional Temperature Difference as a Function of ReynoldsNo. for One-G Field 5-8

5-2 Non-Dimensional Temperature Difference as a Function ofReynolds No. for Zero G. Simulation 5-9

ii

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ILLUSTRATIONS(Cont'd.)

Figure Page

5-3 Velocity Scale Ratio Versus Dimensional Scale Ratio forZero G. Simulation 5-H

APPEHDIX

A-l Task 2 Work Breakdown Structure A-k

A-2 Program Schedule A-5

111'

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NOMENCLATURE

A_ — Surface area ratio for incident radiant heat rates

A. — Surface area ratio of element i

A — Area ratio normal to conductive heat flown

C — Ratio of model specific heat to prototype specific heat

C — Specific heat at constant pressure

G — Ratio of Grashof Numbers in the model and prototypeT*g — Ratio of gravitational fields

h — Ratio of convective heat transfer coefficient in model to that inprototype

•x-I — Ratio of incident radiant heat flux

K — Ratio of model normal thermal conductivity to prototype conductivity

K * — Ratio of conductivity of gas in model to the gas in the prototypeo

L — Length ratio of model to prototype

M — Ratio of molecular weights*P — Ratio of pressures#

p.r — Ratio of Prandtl Numbers

Q — Ratio of energy transfer*

Re — Ratio of Reynolds Numbers

t — Ratio of insulation thickness

t" — Equivalent skin thickness

T — Model temperaturem

T — Prototype temperature

T . — Gas temperature6V — Volume ratio, Velocity ratio

iv


A C R O U P D I V I S I O N OF L O C K H E E D A I II C II A F T C O R P O R A T I O N

Greek Letters

3 — Orbit orientation

e — Emittance

cv_ - Solar absorptance8

p - Density*

p - Ratio of densities*9 - Time scale ratio£.

y, - Ratio of fluid viscosities

Subscripts

L/T — Laminar flow in model and turbulent flow in prototype

c — Convective coefficient

sub — Substrate

ins — Insulation

n — Normal direction

Superscript

*— . Indicates the ratio of properties on the scale model to properties

on the prototype.


A G R O U P D I V I S I O N O F L O C K H E E D A I R C R A F T C O R P O R A T I O N

Section 1

Summary

The objective of this contractual effort was to assess the capabilities

and sensitivies of a Space Shuttle Thermal Scale Model (TSM) to support the

Shuttle Orbiter integrated thermal control design, define a test article and

preliminary test plan, and provide cost and schedule information for final

design and fabrication of the TSM. To accomplish this objective the study was

divided into five major tasks as follows:

Task 1: Shuttle Vehicle Design Review and Study Plan

Task 2: Thermal Scale Model Design

Task 3: Preliminary Test Plan Development

Task : Shuttle Thermal Scale Model Cost and Schedule

Task 5: Program Reports and Documentation

The program was originally scheduled to be a 7-m°ath effort beginning on

June 30, 1972, and ending on January 30, 1973. On October 10, 1972, approxi-

mately 3 months and 10 days after receipt of contract, IMSC was notified that

the program was being cancelled at the convenience of the Government. During

the period of performance prior to termination, the activity on Task 1 had

been completed and work has started on Task 2 which involved detailed design

of the thermal scale model. Technical progress had included development of a

coarse thermal mathematical model for the entire Shuttle Orbiter, temperature

calculations for the extreme hot, cold, and worst transient orbital cases,

and development of scale model design and test requirements. In addition,

the level of detail for modeling various subsystems was established, several

potential applications for the scale model were identified, and a list of

trade studies to aid the TSM detailed design was developed. The project was

on schedule and within budget at the time of termination.

The purpose of this report is to document the progress on the project and

to present the results of the completed studies. Since the project was not

1-1


A G R O U P D I V I S I O N OF L O C K H E E 0. A I « C R A F I C O R P O R A T I O N

completed/ many of the questions regarding the application and sensitivity of

a TSM for use in Shuttle thermal design remain unanswered. However, several

significant items, especially those involving TSM design criteria for model-

ing convection in the crew cabin and .modeling of the conduction-radiation

heat flow paths throughout the vehicle, were established.

Results of the initial studies indicate that a thermal scale model with

an overall scale ratio of 1/3 would be feasible to construct and would provide

valuable information during preliminary thermal design of the Shuttle and

would serve as a useful tool throughout the entire thermal design and verifica-

tion program. The studies also show that a 1/3 TSM could be tested in the

NASA/MSC Chamber A test facility under solar simulation with rearrangement of

the present side and top solar lamp bank systems. For a TSM to be of maximum

benefit to the Shuttle design effort, final design and fabrication of the test

article must be completed early in the program such that test results can be

obtained prior to the FDR.

Included in this report as Appendix A is the Analysis and Study Plan which

was generated at the conclusion of Task I, but was not released due to contract

termination. Appendix B contains an outline of a design review meeting held

with NASA/MSC to review the results of Task I and to present the approach to

be followed in accomplishing future work on the contract. The outline contains,

in summary form, the major considerations in modeling the conduction-radiation-

convection fields, the level of detail for modeling various systems, prelimin-

ary test requirements, and a list of potential applications for the TSM.

1-2

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Section 2

lOTRODUCTIOW

2.1 Background and Program Objectives

The Space Shuttle concept definition of Phase B defined a number of

critical thermal control problems. These problems include: crew compartment

temperature control and thermal loading, external insulation bond line tem-

peratures and rate of temperature change, payload bay heat balance and tem-

perature control, QMS and RCS propellant tanks and RCS thruster temperature

control, thermal radiator performance, landing gear minimum temperatures,

elevon actuator temperature control, minimum temperature of hydraulic fluids

throughout the vehicle, and payload/orbiter thermal interface effects. These

and other potential thermal problems can be investigated by analytical means

if a thermal mathematical model (TMM) can be developed that provides exact

definition of the internal and external heat flow fields. Verification of

the analytical design.on a total systems basis is not possible, however,

due to the massive size of the Shuttle vehicle, as compared to existing ther-

mal vacuum test facilities. Thus, a different approach to verification of

the thermal design, as compared with that employed for previous spacecraft,

is necessary. This combination of circumstances leads to consideration of a

small-scale thermal model. :

The potential for use of small-scale thermal models during spacecraft

preliminary design and in design verification studies has been under investi-

gation for a number of years. A variety of analytical and experimental

programs has demonstrated the advantages and limitations of using such models,

These programs involved a wide range of complexity — from studies of simple

geometries that lend themselves to rigorous analytical verification of test

results, to studies of complete spacecraft where every effort was made to

include all design details of the prototype vehicle. A summary of findings

from all studies completed to date leads to the conclusion that a thermal

scale modeling program can be a valuable tool for assessment of the strengths

and weaknesses of a thermal design approach.

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A G R O U P D I V I S I O N OF L O C K H E E D A I R C R A F T C O II P O R A T I O N

With this background in mind, the present program was undertaken to

determine the utility of a thermal scale model in the overall Space Shuttle

thermal design and thermal verification program. Specifically, the applica-

tion of a TSM to assist in the development of a thermal math model and to

aid in design of critical thermal control areas for the Orbiter were to be

studied. In addition, the study had the objectives of determining the

utility of a TSM for thermal verification testing of the final Shuttle design

and the application of a TSM for thermal testing of the Shuttle in combination

with typical payloads.

2.2 Technical Content

Reference 1 presents the plan by which the Shuttle thermal scale modeling

application study was to be accomplished. The program plan document (Ref. l)

describes the specific tasks and subtasks and the approach for accomplishing

the program objectives. In terms of the major tasks, the technical work

planned for this study is outlined as follows:

Task 1; Shuttle Vehicle Design Review and Study Plan

o Review of Phase B Shuttle Orbiter configuration

o Establish Thermal Analysis Requirements

o Development of prototype preliminary TMM

o Establish preliminary TSM design requirements

o Establish preliminary TSM test requirements

o Define role of TSM in the Shuttle thermal design plan

o Preparation .of Study plan for use in Task 2

2-2



Task 2; Thermal Scale Model Design

o Finalize prototype TMM

o Thermal analysis using prototype TMM

. o Trade studies to define TSM design criteria

o Finalize TSM design requirements

o Establish test facility requirements

o Develop TMM for scale model

o Detailed design of TSM

o Documentation of TSM design and specifications

Task 3; Preliminary Test Plans

. b Develop test requirements document

o Establish post test analysis requirements

o Develop requirements for correlation of test results

Task k: Shuttle TSM Cost and Schedule

: o Definition of tasks and preliminary WBS for final design

and fabrication

o Develop preliminary task schedule

o Establish cost estimates for final design and fabrication of

Shuttle TSM •

2.3 End Product

The end product planned for this contractual effort was to be a complete,

documentation of the application of a TSM to Space Shuttle Orbiter thermal

design. Specifically, a complete set of preliminary engineering design draw-

ings for a Shuttle Orbiter TSM was planned. Scaling factors, model materials,

assessment and identification of the level of detail in the area of subsystems,

2-3

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assessment of applicability to the Space Shuttle Orbiter integrated thermal

control design, and identification of sensitivities in terms of the model's

ability for predicting prototype thermal behavior were to be specifically

addressed. Details of test plans, test facility requirements, and schedules

and total cost of a TSM final design and fabrication program were to be es-

tablished.

Because of project termination before completion, the final objectives

of the study were not realized. Work completed to date of cancellation in-

cluded those items listed under Task 1. This effort included preparation of

an analysis and study plan for use in accomplishing Task 2. The contract

was terminated before the plan could be formally issued; therefore, in order

to document all work accomplished, the study plan is included in Appendix A

of this report. In addition, work had begun on Task 2 with refinement of

the TMM and initiation of TSM detail design. Results of this effort are re-

ported herein. No work had been done on Tasks 3 or to date of contract

termination.

2-k

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Section 3

SHUTTLE PHASE B OKBITER CONFIGURATION REVIEW

3«1 Shuttle Orbiter Configuration

The. proposed Shuttle configuration was reviewed in detail to familiarize

personnel assigned to the program with the vehicle design and to establish

the depth of thermal analysis required to support the scale model design

activity. The vehicle's thermal design requirements were reviewed to identify

critical thermal control areas, natural heat-flow boundaries, and areas

with unique thermal response characteristics.

Per NASA and IMSC agreement, the Shuttle Orbiter concept as proposed by

LMSC in the Space Shuttle Technical Proposal, Volume III,. IMSC/D15736 , May 12,'

1972, was used as a basis for the thermal modeling application study.

Use of the IMSC Shuttle concept was selected because of the immediate

availability of design information to IMSC personnel and due to the compressed

schedule required for the TSM .program. Taking this approach would not impede

achieving the primary objective of the study which was to prove feasibility

and utility of a TSM to support the integrated Shuttle thermal design/verifica-

tion program. Major differences in critical thermal control areas between

the IMSC concept and the final Shuttle design were to be considered during this

study to assess their impact on the model's final configuration and test per-

formance. NASA agreed to keep IMSC informed of major differences which should

be considered during the modeling program.

The Shuttle concept, shown in Pigs. 3-1 and 3-2, shows the major sections

and systems of the vehicle. As shown in Fig. 3«1> the vehicle structure can

be divided into six major sections. These include the forward fuselage,

center fuselage, aft fuselage, wings, payload bay doors, and the fin and rudder

3-1


A G R O U P D I V I S I O N OF L O C K H E E D A I I ) C B A f T C O R P O R A T I O N

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3-2


A G R O U P D I V I S I O N O F L O C K H E E D A I R C R A F T C O R F - O R A T I O N

SIDE THRUSTERS

HEAT SHIELD

Fwb THRUSTERS

s:s PROPELLENT S » S T £ M

VIEW A-A

CREW CABIN A AVIONICS BAY

-MANIPULATOR iRM SUPPORI

RCS MOSUL E

RCS MODULE(»LTEB«»IE POS'llON)

Fig. 3.2

BASE HEAT SHIELD

LMSC Space Shuttle 040AConfiguration

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sections. Pig. 3«2 shows the location of major equipment and systems within

the -vehicle. The overall prototype Orbiter dimensions as given in Fig. 3«3

are 119. ft. long by 78.3 ft. wide by 52 ft. high.

The IMSC concept proposed the use of a honeycomb skin structure. How-

ever, for purposes of the TSM study a solid aluminum skin structure was

assumed per NASA request, with equivalent t (average thicknesses) for this

structure provided by NASA. The equivalent t values used in the analysis

are shown in Fig. 3-^ with updated values given in Fig. 3-5« The t values

include the skin and skin stiffeners. Material for the prototype skin struc-

ture was assumed to be 202*4—T . Other major structural materials included

6A1-14-V Titanium for the landing gear and k kO steel for the elevon and fin

actuators.

LI-15QO with an infrared emittance of 0.8k- and a solar absorptance of

0.80 was assumed for the external thermal protection system (TPS). Further

details on this and other thermal control components are described in Section

k.

3.2 Critical Thermal Control Areas

Various critical thermal control areas for the Shuttle Orbiter were

identified during the Task 1 studies. These areas are listed along with their

temperature requirements in Section k- of Appendix A.

3.3 Level of Detail for Modeling Systems

During Task 1 an assessment was made regarding the level of detail required

for modeling various systems and subsystems within the Shuttle Orbiter. This

level of detail is presented in Section 8.3 of Appendix A.

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Section k

THERMAL ANALYSIS

l)-.l Thermal Analysis Requirements

During the initial phase of the program, a plan was developed for the

analytical studies to determine the required complexity of the TMM in terms

of the support it would provide to: (l) assessment of the feasibility and

usefulness of a thermal scale model, (2) performance of trade studies, (3)

design of the thermal scale model, and (U) definition of test facilities and

source simulation requirements. The requirements for developing the TMM in-

cluded definition of the nodal network for all internal and external sections

of the vehicle and definition of the orbital cases to study for determination

of external heat rates. Requirements were also defined for including a

typical payload configuration in the TMM.

A thermal analyzer model is considered as one of the basic tools in

design of a TSM. Development of the preliminary TMM was initiated as soon

as possible after contract go-ahead and was essentially completed during

Task 1. Early development of the TMM was considered necessary in order to

complete a satisfactory design for the scale model within the specified time

schedule.

In developing the TMM, the approach followed was to construct a math

model.only in sufficient detail to support this study contract. Thus, a

coarse nodal network was constructed for purposes of defining important heat

flow paths and temperature fields throughout the vehicle. A detailed TMM,

typical of.that required for vehicle thermal design, was not required for

this study. The TMM was set up to include the three dimensional conduction,

radiation, and convective networks for each major area of the Orbiter. Pro-

visions were made for incorporating a typical payload configuration in the_ . .

TMM. Temperature performance of the vehicle was obtained for the extreme hot,

cold, and worst transient orbital cases as described in Section .4.


A G B O U P D I V I S I O N O F L O C K H E E D A I R C R A F T C O R P O R A T I O N

To aid in TSM design and to provide a means of comparing scale model re-

sults with prototype results, a requirement for developing a TMM of the

thermal scale model was established. Computer analysis were planned as the

TSM design progressed to support decisions regarding any necessary deviations

from exact, thermal scale modeling.

k.2 Thermal Math Model Development

Development of the TMM was accomplished in two stages. The first stage

consisted of a preliminary TMM developed during Task 1., and included k O internal

and external thermal nodes connected by approximately 700 conduction resistors

and 1500 radiation resistors. The second stage incorporated a typical payload

configuration and various design refinements that were defined as a result of

Task 1 studies. This math model consisted of 550 thermal nodes connected by

approximately 800 conduction resistors and 1100 radiation resistors. The re-

vised TMM was about 90$ complete when the program was terminated.

Figures k-1 through k-Q show the nodal breakdown for various sections of

the vehicle. The TMM is symmetrical about a plane passing vertically through

the logitudinal centerline of the vehicle. All nodes on the pilot's left are

odd numbered, while corresponding nodes on the right are even numbered. The

right side value of each corresponding node is one unit greater than those on

the left side. Figures U-9 and -10 give the surface and. underlying structure

node numbers, the node area, and the. average aluminum skin thickness assumed.

U.2.1 Basic Assumptions

Various assumptions were made during construction of the TMM to simplify

the model and speed its development. These assumptions were divided into two

general categories:


A G R O U P D I V I S I O N O F L O C K H E l i D A I R C R A F T C O R P O R A T I O N

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1) structural, in which the physical shape of the vehicle was altered or

distorted, and

2) thermodyhamic, in which the overall vehicle heat transfer was altered

without affecting the orbital temperatures.

Structural Assumptions

The general shape and physical relationship between surfaces and components

of the Shuttle Orbiter were maintained. Some of the exterior surface contours

were altered as follows:

1. The region in front of the crew cabin was changed to a conical sur-

face.

2. The vertical, stabilizer was assumed to be a trapezoid rather than in

its true shape.

3- The body cross section aft of the payload bay was changed to a

rectangular shape.

• k. The wing upper and lower surfaces were assumed to be flat, with the

upper surface canted down U to account for the taper of the wing.

5. The base area where the orbiter main engines are located was modeled

as a flat plate and modified to account for the engine cutouts. The

modifications included a thinner aluminum structure and reduced sur-

face infrared emittance and absorbtance to account for shading by

the engines.

The interior structure was assumed to be one of two types, either shear

web or trusswork construction. Inside the wing' and the fuselage lower surface,

0.032 inch thick aluminum shear webs spaced on 30 inch centers were assumed.

Between the fuselage outer surface and payload bay inner surface, trusswork

structure was assumed. Figure 4—Ushows the assumed construction. 'The truss-

work was assumed to be at the same stations as the shear webs in the body lower

surface and consisted of 1.5 inch O.D. tubes with 0.030 inch wall thickness.

4-13

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L O C K H E E D : M I S S I L E S 4 S P A C E C O M P A N YA G R O U P D I V I S I O N - OF L O C K H E E D A I R C R A F T C O 8 P O » A T I O N

PAYLGAD BAY STRUCTURAL CONCEPT^ WING STRUCTURAL CONCEPT

VUL.T1LA YER INSULATION

PAYLOAD BAYSTRUCTURE

TRUSSWORK

1.50 IN. OH TUBES

0.030 /A:. \'.ALL

SHEAR ViE3

•PAYLOAO BAY DOOR

BODY UPPER SURFACE

STRUCTURE

LH500

BODY LOWER SURFACE

STRUCTURE

LHSOO •

'IGUREW-SPACE SHU

WING UPPER SURFACE, L1-1500 —

STRUCTURE

SHEAR WEB1-0.032 IN.

Wlf/G LOWER SURFACESTRUCTURE

LH500

TTLE STRUCTURAL DETAIL

All structural material used in the TMM was assumed to be 2024-T3 aluminum

throughout, except for the landing gear "bogeys where 6 A1-4V Titanium was used.

By NASA directive, a solid aluminum skin structure was assumed throughout

the vehicle in place of the honeycomb structure proposed by IMSCI in Ref. 2'.

Equivalent skin and structural thicknesses (t) as furnished by NASA (Figure 3-l4)

were used in the TMM development.

The landing gear thermal model was based on the IMSCI Orbiter details. Main'

gear weight was 35 1 Ibs. per side, including three wheels per side at 210 Ibs.

each. For the nose gear, 183 Ibs. was allowed for the two wheels and 875 Ibs.

for the bogey. Titanium was assumed for the landing gear bogeys and rubber

(C = 0.48 BTU/LB°F, p = 75 lbs/ft3) for the tires.

Thermodynamic Assumptions

Thermodynamic assumptions used in developing the TMM have been directed

toward simplifying the computer model while maintaining the correct thermodynamic

relationships and orbital temperature response of the vehicle. The following

assumptions were made:

l) Heat conduction through the HI-1500 parallel to the surface is negligible

compared to conduction through the aluminum skin structure. This was

verified by hand calculations and computer analysis. Hand calculations

showed that for the thickest LI-1500 (1.74 in.) on top of the thinnest

aluminum skin (0.060"), the ratio of the heat conducted through the

LI-1500 to that conducted through the skin was 1 to 30. Therefore, for

the same temperature difference between two points, 30 times more

energy would be conducted through the skin than through the LI-1500 in

a direction parallel to the vehicle surface. Computer analysis showed

that for two adjacent LI-1500 nodes on the Orbiter, the temperatures

were +37.3°F and -67.6°F with or without the lateral LI-1500 conduct-

ance.

LOCKHEED PALO ALTO RESEARCH LABORATORYL O C K H E E D . M I S S I L E S & S P A C E C O M P A N Y

A G H O U P D I V I S I O N O F ' I O ' C K H E E D A I U C R A F T C O B P O K A t l O N

2) The payload "bay doors and landing gear doors were assumed not to be

connected to the structure by any heat conduction paths. Although

these doors are secured by hinges and latches, these items were

assumed to have high thermal resistance. Before any other assumption

regarding hinge thermal resistance can be made for the thermal model,

it will be necessary to obtain data on the resistance across the

actual hinges proposed for the prototype.

3) Wo ascent aerodynamic heating effects are included since these effects

are assumed to be non-critical in determining the orbital temperature

response of the Space Shuttle for the modeling program.

h) The crew cabin with its interior equipment was input to the TMM as an

insulated sheil with an interior convective environment. The cabin

insulation was assumed to be 0.75 inch thick polyurethane foam with an• " p

internal convective heat transfer coefficient of 0.25 BTU/Ft Hr. No

windows were added to the cabin since they have their own cooling

system and movable shades. The structural concept consisted of a

single pressure shell on the upper surface where the inside cabin

contour, matches the body contour and a pressure shell and aerodynamic

shell on the sides and lower surface where the two contours do not

coincide. The internal atmosphere temperature was set at 70 F.

5) Optical properties of the various materials used on the Space Shuttle

are shown in Table .1. On the interior structural surfaces, such as

the wing upper and lower surfaces, the emittance (e) was assumed to be

0.8, which is characteristic of a surface painted for corrosion con-

trol. The payload bay interior surfaces were assumed to have an e .

.of 0.9, as directed by NASA. Radiation coupling between various nodes

in the TMM was assumed, not only between exterior surfaces, but also

between interior structural nodes such as wing upper and lower surfaces,

vertical stabilizer sides, and body structure and payload bay sides.

This radiation heat transfer was found to be the primary mode of heat

transfer in these areas.

4-16


A G « O U r D I V I S I O N OF L O C K H E E D A I R C R A F T C O R P O R A T I O N

TABLE 1

OPTICAL PROPERTIES OP MATERIALS USED IN THE SPACE SHUTTLE

Material Location

TABLE 2

LI-1500 THICKNESSES

or/.

LI -1500 • 80

Payload Bay Interior Surfaces 0.9

Structure Interior Surfaces

Carbon/Carbon Nose Cap 0.9

.8k .952

0.9 i.o

0.8

0.9 l.O

Body Lower Surface

STA 250 to 575

375 to 10251025 to 1500:

Wing Lower Surface

Eleven Lower Surface

- 1.56"

1.66"

1.62"

1.7V

2.00"

Body Upper Surface

STA 250 - 575 0.63"

575 to 1500 0.40"

Vertical Stabilizer O.V

Wing Upper Surface 0.4"

Eleven Upper Surface O.V

Base 1.6"


A G R O U P D I V I S I O N OF L O C K H E E D A I I C IT A F T C O I P O I ' A T I O N

6) The only source of power dissipation included in the model to date of

program termination were the radiators which dissipate a total of

67000 BTU/Br to space. Other, heat sources which were to "be included

at a later date include the Reaction Control System propellant, stor-

age areas, fuel cell units, and avionics packages not actively cooled.

The 67000 BTU/Hr heat dissipation for the radiators is the maximum

value, and any additional heat load would be handled by waste water

sublimators. The minimum radiator heat load was assumed to be 0 HPU/

Hr for a docked condition with no personnel aboard.

7) LI-1500 thicknesses used for the TMM are shown in Table 2, with the

Orbiter stations shown in Figure -12. The thickness ranges from O.k

, -inch on the upper vehicle surface to 2.0 in. on the elevon lower sur-

•" "•.face.; ''.:.;'•' '. ' ' ' :. .-. - '• ' '' '-: .' •'"' ; .-.-.'

14-.2.1 Payload Configuration

Per agreement between NASA and IMSC, a typical payload was defined for

inclusion in the thermal analytical studies. The configuration and properties

of the payload are as follows: .

o Cylindrical shaped payload 60 ft. long by 15 ft. O.D.

o 202U aluminum skin with a thickness of 0.03 inch.

o Emittance of payload external surface of 0.9.

o Emittance of internal payload bay surfaces of 0.9.

o Passive payload with ho internal energy generator. .

o Assume only radiation coupling between payload and orbiter

o Assume multilayer insulation on the interior of the payload bay asfollows (Refer to Fig. lt-13 for stations):

U-18

LOCKHEED PALO ALTO RESEARCH LABORATORYL O C K H E E D M I S S I L E S i S P A C E C O M P A N Y


LOCKHEED MISSILES & SPACE COMPANY SPACE SHUTTLE PP.OJECT

PREPARED BY

EM HO.

DATE

PAGE OF

o

CO

93Q:O

PreparedLOCKHEED MISSILES & SPACE COMPANY

Page

Checked

Approved

Model

Report No.

J$OT70S*1 /*Jr££/U*. jrv/lffAcig 6* fAYtoA/ S / / / SS S S S /S S / S S ' SSS'S/

i£

PORM LMSC 362B-1 LIETZ I 118 JTOPFLIGHT

Surface Ko. MU Thickness (in.j

1 0.852 0.58.3 0.58.4 0.585 0.586 0.587 1-988 1.989 2. in10 2.ia11 1.9812 1.98

o Assume properties for the multilayer insulation

Cp @ 70°F = .315 BTU/LB °F

p - 1.58 LBS/FT3 (?0 layers/in.)

Temp (°F) Cp(T)/Cp(70°F)

-360 0.16-310 0.23-210 0.3570 1.0

K @ 0 atm. @ 70°F = 0.7 x 10~3 BTU/FT-HR-°F

Temp (°F) K(T,P)/K(70°F, 0 ATM)

-300 .338-200 .375-125 . 38- 50 .588o .72570 1.0150 1.312225 1-75350 2.1 75

1+-21LOCKHEED PALO ALTO RESEARCH LABORATORYL O C K H E E D M I S S I L E S & S P A C E C O M P A N Y

A G R O U P D I V I S I O N O F t O C K H E E D A I R C R A F T C O R P O R A T I O N

U.3 Orbital Conditions

Heat rates and vehicle temperatures were calculated using the TMM for

the extreme hot, cold, and worst case transient orbital conditions. The

three orbital cases are as follows:

Hot Casej P = 90°, solar oriented with the vehicle rolled 5° to

provide the maximum projected area of vehicle top side

toward the sun, pay load doors closed, radiators open.

Cold Case; p = 90 , bottom of vehicle earth oriented, tail facing sun,

payload bay doors open.

Worst Transient; Equitorial orbit, payload bay doors open and facing

sun in solar orientation; vehicle to pitch or roll 90 >

depending on initial orientation, as it enters earth's

earth while in shadow, then pitch or roll 90 to repeat

shadow and to remain earth oriented with bottom facing. the

earth while in shadow, then pitch or r<

solar exposure of open payload region.

TMM Results

Steady state temperature results for the hot and cold orbital cases are

presented J.n Pigs._..; -_l!iLlL.thraugh h-j.... For the hot case, the aluminum skin

temperatures ranged from approximately 190 F on the wing upper surface to - ihO F

on the fuselage- lower surface. For the cold case, all structural temperatures

are in the range oJ

and radiator area.

are in the range of -200°F to 0 F, except for the sun illuminated aft section

Results were obtained for a short transient run which demonstrated that the

computer program was operating correctly. The data was not reduced prior to

contract termination, and therefore, will not be presented in this report.


co

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b.5 Internal Power Sources

As discussed in Section k.2, the initial TMM assumed only power dissipa-

tion due to the crew cabin and its associated equipment. Prior to contract

termination, various other sources of internal power generation that may have

affected internal temperatures were identified. These would have been included

in the revised TMM had the program continued. A discussion of these are in

the following sections.

4.5.1 ECS and QMS

The Reaction Control System (RCS) and Orbital Maneuvering System (QMS)

propellants have an allowable temperature range during storage of +50 P to

+90 F. For the TMM, a thermal node insulated from its surroundings with 2.5

in. of TG-15000 was assumed. A heat source (or sink) was provided such that

when the temperature exceeded the set limits, enough heat would be provided

or allowed to dissipate to keep it within the stated limits. Although the

RCS thrusters have a catalyst bed which is maintained above 150 F, the heat

leak effect was neglected since the thrusters are well insulated and any heat

leak from this source can be assumed to keep the propellants warm. A schematic

of the thermal model would look like this

RCS.

2.5 in TG-15000

/// /

£_//-7"

/ / S / /

50° <T

7 s / / s

/

<

/

/ // /

90

/

°F

/ /

//>

. •'// /

Q to surroundings

- radiation

- conduction

Two of these systems will be placed in the aft structural compartment and one

forward of the crew cabin. The two located aft would encompass both the RCS

and QMS on one side of the vehicle into one node.

LOCKHEED PALO ALTO RESEARCH LABORATORYL O C K H E E D M I S S I L E S 4 S P A C E C O M P A N Y

A C R O U P D I V I S I O N O F I O C K H F . E D A I R C R A F T C O R P O R A T I O N

It-.5-2 Fuel Cells

The fuel cell plants, operated continuously in orbit, are actively cooled

with most of their excess heat being rejected by the space radiators. Some

excess heat will escape in the fuel cell area, however, since they operate at

temperatures around 180 P. For the TMM, the recommended approach was to

select a thermal node that is maintained at the fuel cell operating temperature

of l8o F and insulated from its.surroundings as shown below.

2.0 in TG-15000

FUEL C£LL

2.0 -*

/ S / / / / / / / / / /

180°F

Vs s / /////.

V-* Q to surroundings

- Radiation

- Conduction

One of these units would be placed between the crew cabin and payload bay.

Avionics

Avionics in the space shuttle concept studied during this program are

located in three areas, l) the crew cabin, 2) clustered around the orbiter CG,

and 3) in the area behind the payload bay. The crew cabin avionics are actively

cooled by the cabin air and .are not included as a separate node, in the TMM.

Passive cooling is assumed for the other two avionics packages, with all heat

produced being absorbed by the .surroundings. Power dissipation for the aft

electronics is 137 watts. For the CG avionics, the power dissipation is.

estimated as 200 watts. The recommended setup for the TMM is shown below:

w Q ^ Q

^out = 200 watts •

///////£ ^out

/ / / / / / / / /

= 137 watts

>

/,/ / / s s / / /

CG avionics aft avionics


A G R O U P D I V I S I O N O f t O C K H E E D A I K C B A F T C O « P O « A I I O N

h.J.k Hydraulics

The hydraulic system fluid has a lower temperature limit of -UO F. When

the fluid temperature reaches this point, pumps are started to circulate the

fluid until the temperature goes above -20 F, at which time the pumps are

shut down. Due to the small size of actuators and lines inside the vehicle

extremities, these items were not planned for inclusion in the TMM. The pumps

were assumed to be located in the region behind the payload bay where other

heat producing items are located. This would provide the pumps with enough

heating such that they would need no active thermal control of their own. Due

to the fact that their primary duty cycle is during ascent and entry only, the

heat dissipated during their orbital operation was assumed to be small enough

such that it can be neglected.

lf-32



Section 5

TSM DESIGN

Detailed design of the scale model was to be accomplished during the

second phase of the program. This effort had just started when notification

was received of contract termination. However, enough progress had been made

to select a design scale ratio of 1/3 and some of the materials for modeling

the conduction heat paths. Also, relatively firm requirements were established

for modeling the crew cabin atmosphere and candidate gases were selected for

eonvective modeling.

TSM design was to be accomplished in accordance with the standard modeling

laws developed for the "temperature preservation" approach. The general

criteria for thermal scale modeling in the space environment has been esta-

blished and reported by numerous authors, and thus, will not be repeated in

great detail in this report. The basic approach to be followed in design of

the TSM was to select a scale ratio that appeared to satisfy the majority of

the model's requirements and then proceed to model all areas of the vehicle to

meet that scale ratio. It was anticipated that various compromises in model-

ing would be required for some areas of the vehicle that could not be scaled

exactly. It is also a possibility that the scale ratio of 1/3 would have to

be adjusted + 10$ as the process of selecting model materials progressed.

5-1

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A G «OU P 0 I VIS I ON OF L O C K H E E D A l B C R A f f C O R P O R A T I O N

Had the program "been completed, the end result of the model design

activity would have been the specification of materials, size, shape, weight,

internal power sources, external heat flux requirements, details of critical

thermal conduction joints, details for modeling the crew cabin atmosphere,

and details of modeled components and their location relative to each other.

A complete set of specifications and engineering drawings was planned to

document details of the TSM design.

5«1 Design Requirements

During Task 1, preliminary TSM design requirements were established.

This effort involved consideration of which modeling approach to follow (i.e.

temperature preservation, materials preservation, etc.), scale ratio, mat-

erials, modeling of convective heat flow, TSM size in relation to that of

available test facilities, important conduction and radiation heat flow paths,

level of detail for modeling various systems and components, and the potential

applications of the TSM. The design criteria established during this initial

phase are listed in Appendix A, Section A. 6.

5.2 Conduction - Radiation Modeling^

From Ref . 3, the criteria for modeling the conduction and radiation tem-

perature fields, are as follows :

5-2


*For temperature preservation, T = 1, and Eg. 1 becomes

•*• ** * * K Ap V C * * n n * /r»£—r- - AX =Q = — r- =A± (2)

For the model to represent prototype behavior, Eq. 2 must be satisfied for all

conditions and. for all heat flow paths. This would be an easy task if an

infinite number of materials were available with thermophysical .properties that

spanned the entire range. However, in reality this is not the case, therefore,

compromises must be made in the scale criteria. These compromises generally

are involved with distortion of the conduction heat flow path by adjusting the

cross sectional areas or the lengths of heat flow paths to properly scale the

amount of energy conducted. Such distortions must be done in such a manner so

that the important radiative heat transfer areas are not adversely affected.

Approximately 9&jb of the Shuttle Orbiter's external surface is covered

with insulation termed the thermal protection system (TPS) that protects the

vehicle structure and internal components from the severe heating during the

ascent and entry stages of flight. For purposes of' this study, the LMSC

material LI-1500 was assumed for the TPS. The thermal conductivity for

this material is 0.01 2 BTU/Hr Ft °R at 0°F and is essentially isotropic.Q

The density is approximately 15 Ibs/ft , and the thickness ranges from a

minimum of Q.k inch to 2 inches over the Shuttle surface. .

Because of the wide usage of this material on the Orbiter, particular

attention was paid to modeling the LI-1500 in combination with the underlying

5-3


A O ROU t ,DI V I S IO N OF L O C K H E E D A I R C R A F T C O R P O R A T I O N

aluminum substrate. For purposes of this study, 202*1—T3 aluminum was assumed

for the skin, and much of the other structural components. This aluminum

alloy has a thermal conductivity of 70 BTU/Hr Ft °R at 75°F. In modeling

conduction through the LI-lpOO and aluminum skin, certain assumptions have

to be made due to this combination of materials and due to the materials avail-

able for use on a thermal model. The required assumptions result in distor-

tion of the thickness scale ratio. Based on results from the TMM analysis,

the assumption can be made that heat flow in the LI-1500 is one-dimensional

(i.e. perpendicular to the skin surface) and that heat flow in the aluminum

skin is two dimensional (i.e. circumferential and longitudinal in the aluminum

skin).

Designating LI-1J500 as material No. 1 and the aluminum skin as material

Wo. 2, and equating the various dimensionless groups in Eq. 2, the following

equations are obtained:

•X- j(f-For the case of heat flow perpendicular to the vehicle surface A = L..,

•3f #and L. in Eq. 2 becomes the thickness ratio t.. .

Thus,* *

K . = tnV L

* * * #e - - K

x*-i= L1


A G H O U P D I V I S I O N O F L O C K H E E D A I R C R A F T C O R P O R A T I O N

For the case of heat flow parallel to the vehicle surface, the aluminum

skin is the predominate mode of heat transfer and the following identities

result:

* ••**.*e - P2c2-t2

*-i =1

> - L2

Q represents the heat flux by conduction, internal

radiation, or any internally generated heat flux within

the vehicle

Therefore, to be more general

Q* =L*

where L is the overall length scale ratio which for

the Shuttle TSM is equal to 1/3.

In attempting to model the LI-1500, it was found that very few materials

are available which have a conductivity lower than LI-1500. Multilayer in-

sulation does, however the conductivity of this material is not isotropic.

One candidate material for modeling LI-1500 is a Johns-Manville insulation

called MIN-K . This material has a thermal conductivity under vacuum con-

ditions that is approximately 1/2 that of LI-1500 at room temperature. There-

fore, from the equation K = t,, we find that the thickness ratio t., is 0.5.

5-5

LOCKHEED PALO ALTO RESEARCH LABORATORYL O C K H E E D M I S S I L E S & S P A C E C O M P A N YA G R O U P D I V I S I O N O F L O C K H E E D A I I C B A F T C O R P O R A T I O N

This means that the thickness of insulation on the thermal model would be

1/2 that of the prototype.

For modeling the aluminum skin structure it was found that some of the

stainless steels can be used; however, in using stainless steel for this

application, distortion of the skin thickness ratio is required to model the

heat flow because the conductivity of stainless stell is not exactly 1/3 that

of aluminum.

Additional problems occur when trying to model the transient response.

The full impact of compromises that may have to be made to predict transient

behavior of the spacecraft were not determined prior to contract termination.

5.3 Convective Modeling

The important criteria for convective modeling are derived in Ref. k. There

it is shown that the criteria can be determined from the general equation:

* * » • * • » ^ o*V*C*T* * * K*A*T*h A T =Q = A. T = P . Z. = A_ I = —

• . • n x e* ^ L*

From this it is found that the similitude criteria are

"5f 3v * / vT.

h = K (1/L )

In order to meet the above criteria the Reynolds, Grashof and Prandtl

numbers must remain invariant. If this is done then the exact temperature

preservation criteria becomes

# *Kg =.L.

However, in order to meet these criteria a suitable gas must be found which

will preserve the Prandtl number while at the same time have a conductivity•£

ratio equal to L . The Reynolds number can usually be preserved by adjusting

the flow velocity while the Grasholf number can be preserved by adjusting the

gas pressure according to

5-6

LOCKHEED PALO ALTO RESEARCH LABORATORYL O C K H E E D M I S S I L E S 4 S P A C E C O M P A N Y


P* = ( y,*/k*) (1/g V)

Generally it is difficult to find a suitable gas which meets the above

criteria. To overcome this problem a system using "Scaling Compromises" was

derived by Shannon (Ref. 5). These compromises involve the so-called "heat

transfer coeffocient preservation" method and the "mass flux preservation"

method.'

These two methods use the same gas in the prototype and model and adjust

to flow conditions in order to attempt to maintain similitude. It was shown

in Ref. 5 that if either mass flux or heat transfer coefficient were preserved

then reasonable similitude could be obtained.

Both temperature preservation techniques and scaling compromises techniques

were investigated during Task I of the program. The results of these studies

are given in Appendix B and will be reviewed briefly here.

The preliminary analysis of the prototype heat transfer coefficients showed

that in a 1-g 1-atm test the free convection heat transfer coefficient would be

hc =0.14-3 BTU/Hr Ft °F while the forced convection coefficient would only be

h P = 0.25 BTU/Hr Ft °F. That is free convection would be dominant in bothP prototype and model.

.The "Scaling Compromise" method of heat transfer coefficient preservation

could be used with reasonable results in a 1-g test for scale ratios of L = 1/3

(Figure 5«l) and Reynolds greater than 10 .' Likewise the mass flux preserva-

tion would also give good results at any Reynolds number. Therefore, either

method could be used to model the prototype in a 1-g test.

In order to simulate zero-g conditions the pressure of the gas in the

model would have to be lowered. If the pressure is lowered so that free con-

vection is eliminated then very good results could be obtained using scaling

compromises (Fig. 5*2).

5-7

LOCKHEED PALO ALTO RESEARCH LABORATORYL O C K H E E D ' M I S S I L E S & S P A C E C O M P A N YA G H O U P D I V I S I O N OF L O C K H E E D A I * C B * F T C O R P O R A T I O N

W)•H

5-8

cf

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5-9 '

However, if the pressure is lowered then flow velocities must be in-

creased to maintain the correct Reynold's ratio for thermal similitude. This

creates a problem as shown on Figure 5-3-

.£For a scale ratio of L = 0.3 it is found that the velocity of the gas

jt

in the model must be increased so that V W 18. Since the anticipated proto-

type flow rates are around 30 fps this would give a model flow rate approaching

Mach 0.5. At these velocities compressibility comes into play, and the similitude

criteria do not hold

Therefore, we find that Scaling Compromises can give good results for

demonstrating 1-g performance, however, it is impractical to use them to

demonstrate anticipated 0-g performance in the Space Shuttle.

A different situation arises when strict "temperature preservation"

method is used. With this method the main problem is to find a suitable gas

which will meet the criteria

* * *Kg = L and Pr = 1

Once this is done the Reynolds and Grashof numbers can be adjusted by

varying the flow velocities and pressures.

During the study it was found that sulfur dioxide (SO ) could be used if•x*

. the scale ratio was L = 1/2.5. With this scale ratio and sulfur dioxide

strict temperature preservation could be maintained in a one-g test if the%.

velocity ratio was reduced to V = 0.575 and the pressure ratio increased to* ,-P =1.65, both obtainable values.

In addition zero-gravity performance could be simulated by reducing the« #

model pressure so that P =0.21 and increasing the velocity ratio to V = 2.8.

* iSimilar conditions can be found using Freon 11 and a scale ratio of L = 1/3.

However, Freon 11 has some problems associated with its use.

5-10


A G R O U P D I V I S I O N O F L O C K H E E D A I B C 5 A F T C O R P O R A T I O N

co«o

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5-11

In conclusion the convective studies shoved that all three methods

"Mass Flux", "Heat Transfer Coefficient" and "Temperature" preservation could

be used with good results to simulate a 1-g combined forced and free convect-

ion prototype. However, only the Temperature Preservation method would be

practicable for 0-g simulation.

5-12


A G R O U P D I V I S I O N O F ' L O C K H E E D A I R C R A F T C O R P O R A T I O N

Section 6

TSM TEST REQUIREMENTS

An important objective of the study was to define the requirements for

testing the thermal scale model once the design had been completed. Pre-

liminary test requirements were established and are presented in Appendix B.

6.1 Solar Simulation

At the beginning of this program, the thermal control surfaces proposed

for a major portion of the Shuttle Orbiter's external surfaces had values of

solar absorptance and infrared emittance that gave an a /e of near unity.- . • 8

However, the crew compartment windows and the deployable radiators have a /esvalues that are different from the other surfaces and are not unity, and in

addition, it is desirable to test the model under conditions where the payload

bay doors are both open and closed. Considering these factors and the pos-

sibility that a change to a low a /e thermal control surface for the TPS mayS

be made as the Shuttle design progresses, it is highly desirable that the

incident solar heat flux on the model duplicate the solar energy spectrum.

The use of infrared lamps for simulating the external incident heat flux

was considered but not thoroughly investigated. Applying this technique re-

quires either duplication of adsorbed energy into the surfaces using analyti-

cal methods as a .basis for determining the heat flux required, or it requires

that the external surface temperatures be controlled to analytically pre-

dicted levels during the test. Both of these methods present limitations in

testing of the thermal model. The use of infrared lamps for this application

should not be entirely ruled out, however, further study is required to com-

pletely evaluate the problems associated with their use.

6-1


A G R O U P D I V I S I O N OF L O C K H E E D A I B C B A F I C O B P O B A T I O N

6.2 Test Facility

Due to the size of the Shuttle vehicle, the desirability to design a

thermal model with a scale ratio of near 1/3, and the potential necessity

for testing the model using exact solar simulation, NASA ISC Chamber A test

facility would be required for the test program. The present solar simulator

in Chamber A is too small for testing a 1/3 scale model, and thus, would have

to be modified to meet the requirements of this program. There are two pos-

sibilities that could be considered for modification of the solar simulator.

The first would be to increase the total number of lamps to provide a side

sun of 14-0 ft. by 26 ft. and a top sun of 2.6 ft. in diameter. The second

approach would be to re-arrange the present lamps bank system in a modified

triangular shape as shown on pages B-32 and B-33 in Appendix B. This shape

corresponds to the general shape of the shuttle orbiter.

6.3 Test Requirements

A list of preliminary test requirements for the TSM program is presented

on page B-3U of Appendix B.

6-2


A G R O U P D I V I S I O N O F L O C K H E E D ' A I R C R A F T C O R P O R A T I O N

Section 7

POTENTIAL APPLICATIONS OF THE TSM

The primary advantage of using, a physical scale model as described

herein is that it would provide test data on both a system and subsystem

level during the early design stages in order to guide the thermal design and

to assure that an optimum design was accomplished for the entire shuttle

orbiter. In addition, final verification of the thermal design could be

achieved on a complete system level. A list of potential applications for

the TSM is presented on page B-36 of Appendix B.

7-1

LOCKHEED PALO ALTO RESEARCH LABORATORYL O C K H E E D M I S S I L E S & S P A C E C O M P A ' N Y

A - G R O U P D I V I S I O N O f I O C K H E E D A I B C H A F T C O R P O R A T I O N

Section 8

CONCLUSIONS

Due to termination of the contract before its completion, all the ob-

jectives could not be achieved. However, from the results obtained during

the initial studies that were completed, the following can be concluded:

1. A 1/3 scale model can be designed that models the conduction,

radiation, and convective heat flow fields for steady state con-

ditions.

2. The basic modeling approach to be followed would be that of tem-

perature preservation. Some distortions in the scale ratio would

be required due to limitations in available materials, however,

the major heat flow paths and important temperature fields would

be preserved.

3. The thermal math model proved to be a useful tool in establishing

the criteria for modeling the conduction heat flow paths. It

was concluded that for steady state conditions, one dimensional

heat flow perpendicular to the vehicle surface can be assumed in

modeling the low conductivity TPS while neglecting the heat flow

in the aluminum substrate in this direction. Results from the

TMM indicated that the circumferential and longitudinal heat flow

is predominately due to the aluminum substrate, and the heat flow in

the TPS in these directions can be neglected.

k. Materials are available for modeling the conduction heat paths.

5. The radiation paths would be modeled by preserving geometrical

identity of the major surface areas and by using identical thermal

control surfaces.

6. Gases are available for modeling the convective fields in the

crew compartment area.

8-1


A G R O U P D I V I S I O N O F L O C K H E E D A I R C H A F T C O R P O R A T I O N

7. There would be some problems associated with modeling the

transient conditions, however, the extent and ramifications

of these problems -were not fully explored due to the con-

tract termination.

8. The preliminary studies indicated that it would be desirable

to test the model using a solar simulator which closely

approximates the solar spectrum. For a 1/3 scale model, the

solar simulator in NASA/MSC Chamber A test racility would

have to be modified to accommodate a model of this size. The

trade off in using other means of applying the external incid-

ent heat flux, such as tungsten lamps, was not determined.

9. Several potential applications of the TSM, beneficial to

achieving an optimum thermal design for the shuttle, were

identified.

8-2


A C R O U P D I V I S I O N OF L O C K H E E D A I R C I I A F T C O R P O R A T I O N

Section 9

REFERENCES

1. K. N. Marshall ; w. G. Foster; "Space Shuttle Thermal Scale Modeling

Application Study Program Plan" IMSC $)309067, Lockheed Palo Alto Research

Laboratory, Palo Alto, California, Aug. 11, 1972.

2. "Space Shuttle Technical Proposal, Volume III," LMSC/D157361»-, May 12, 1972.

3. R. E. Rolling, "Thermal Scale Modeling in a Simulated Space Environment,"

N-05-66-1, Final Report for NASA/MSFC Contract NAS 8-1152, Lockheed Palo

Alto Research Laboratory, Palo Alto, June 1966.

k. "Space Shuttle Thermal Scale Modeling Application Study Proposal, Volume I"

LMSC/D082269, May 3, 1972.

5« R. L. Shannon; "A Thermal Scale Modeling Study for Apollo and Apollo Ap-

plications," D180-15C48-1, Boeing Co., Seattle, Washington, June 1972.

9-1


* G R O U P D I V I S I O N OF I O C K H E : : D A I R C R A F T C O R P O R A T I O N

APPENDIX A

ANALYSIS AND STUDY PLAN

The Analysis and Study Plan was prepared to describe the approach to

be used in accomplishing Task 2. It was in the process of being prepared

when the Space Shuttle TSM contract was terminated. Since it was not sub-

mitted, it is included here as Appendix A.

A-l



ANALYSIS AND STUDY PLAIT

NAS 9-12991

1.0 INTRODUCTION AND SUMMARY

The overall objective of the scale modeling application study is to

assess the capabilities and sensitivities of a Space Shuttle thermal scale

model (TSM) to support the integrated thermal control design, define a

test article and a preliminary test plan, and to provide cost and schedul-

ing information for final design and fabrication of the test article. The

study consists of five major tasks as follows:

Task 1: Shuttle Vehicle Design Review and Study Plan

Task 2: Thermal Scale Model Design

Task 3: Preliminary Test Plans

Task k: Shuttle TSM Cost and Schedule

Task 5: Study Reports

The primary activity on the program to date has been on Task 1, which

is now complete with the issuance of this document. Task 1 was involved

with preliminary studies to define the role of a TSM in the overall Shuttle

integrated thermal design and verification plan. The activities under Task

1 consisted of: l) a complete review of the Phase B Shuttle Orbiter con-

figuration and systems; 2) development of a. thermal mathematical model (TMM)

for the prototype; 3) view factor and heat rate calculations; ). development

of preliminary TSM design and test requirements; and 5) preparation of an

analysis and study plan for use in Task 2.

The analysis and study plan presented herein describes the approach to

be used in accomplishing Task 2 which is the detailed design of the TSM0

Reference 1, which presented the program plan for the entire contractual effort,

A-2

LOCKHEED PALO ALTO RESEARCH LABORATORYL O C K H E E D M I S S I I E S & S P A C E C O M P A N Y


and provides an overall schedule and WBS to be used for all Tasks including

Task 2. The present document deals only witti Task 2 and describes in detail

each of the subtasks to be accomplished for Task 2. In addition, the major

results of Task 1 studies are presented. These include l) a listing of the

important thermal control areas and their temperature requirements, 2) TSM

design criteria ' 3) preliminary TSM test requirements, ij-) Level

of detail for modeling the various subsystems, 5) a tentative list of trade

studies to be performed, and 6) a list of potential applications for the TSM.

2.0 TASK 2 WORK BREAKDOWN STRUCTURE AND SCHEDULE

The work breakdown structure for Task 2, showing the major subtasks and

their interrelationships, is presented in Fig. A-l. A schedule for Task 2

activity is presented in Fig. A-2. Each subtask is described in detail in

the following section of this document.

3.0 THERMAL MATH MODEL (TMM) REFINEMENT AND GALCULATIOBB

A TMM of the prototype Shuttle Orbiter was developed during Task 1.

In developing the TMM, the approach followed was to construct a TMM in suf-

ficient detail to support this study. Thus, a course nodel network was

constructed for purposes of defining important heat flow paths and tempera-

ture fields throughout the vehicle, A detailed TMM, typical of that required

for vehicle thermal design, is not required for this study and is certainly

beyond the scope of this contractual effort. Background information and

details of the TMM development for this program may be found in Reference 1.

A-3



TSM DESIGN

TMM REFINEMENTAND

CALCULATIONS

TSM DESIGN CRITERIATRADE STUDIES

TEST FACILITYDEFINITION ANDREQUIREMENTS

TSM DESIGNCRITERIA DEVELOB4ENT

1TSM DETAILED

DESIGN REQUIREMENTS

TRADE STUDIESREPORT

ITSM DETAILED

DESIGN DEVELOIMENT

TSM DESIGNDOCUMENTATION

Figure A-l Task 2 Work Breakdown Structure


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During the initial stages of Task 2, the TMM will be refined to take into

consideration further details of the Space Shuttle thermal design and changes

resulting from the Task 1 studies. Computer calculations will be performed

using the up-dated TMM to study steady-state and transient vehicle response

for the extreme hot, cold, and worst transient orbital cases as defined in

Ref. 1. These analyses will form the basis for detailed TSM design. Speci-

fic refinements in the TMM will include,but not be limited to the following:

a) The revised set of t for the aluminum skin structure received from

NASA/MSG on August 25, 1972, will be incorporated into the TMM provded that

hand calculations show these changes may have significant affects.

b) Additions to the TMM in the crew cabin region, with specific considera-

tions given to modeling window effects, will be added as required to accomplish

TSM design.

c) Heat generated from internal sources will be incorporated in the TMM.

d) A cylindrical shaped payload with properties as agreed upon by NASA

and LMSC will be incorporated into the TMM.

e) Changes will be made to include multilayer insulation on the interior

walls of the payload bay region.

To aid in TSM design and to provide a means of comparing scale model re-

sults with prototype results, a TMM of the TSM will be developed. Computer

analysis will be conducted as the TSM design progresses to support decisions

regarding any deviations from exact scale modeling that may become necessary.

iKO IMPORTANT THERMAL CONTROL AREAS AND TEMPERATURE REQUIREMENTS

Various important thermal control areas for the Shuttle Orbiter were

identified during Task 1. These areas are listed along with their temperature

requirements as follows:

A-6


A C R O U P D I V I S I O N O F L O C K H E E D A I U C B A F T C O H P O B A T I O N

1) Crew compartment (TL™, M to 113°F)

2) Payload bay (-ICXPF to 150°F)

3) Landing gear and tires (-65°F to 350°F)

M Avionics (-65°F to l60°F)

5) Elevon and rudder actuator system hydraulics(-65°F to 250°F)

6) Propellants (50°F to 90°F)

7) Payload/Orbiter thermal exchange

8) Radiators

9) External insulation bond line (-250°F to 350°F)

10) Structure temperature prior to entry (100 F max)

11) Auxiliary power units (APU)

12) Fuel cells

13) Transients for various subsystems

1*)-) Cryogenic tanks

15) RCS thrusters

5.0 TRADE STUDIES

A trade study as defined for this program is any study in which one ap-

proach is evaluated against another for purposes of establishing an optimum

thermal scale model for utilization in Shuttle thermal design. Trade studies

will be performed to establish the degree of complexity required for the model

design, test facility requirements, and a testing program that best meets the

needs of the overall shuttle thermal design requirements. The objectives of

the trade studies are threefold: (l) to provide a guide to the model designer

in terms of final selection of scale ratio, materials, scaling sensitivities

for various thermodynamic features, level of detail for scaling systems and

subsystems, and defining the degree to which distortion in temperature fields

and compromises in eact modeling can be tolerated; (2) to determine whether

A-7


A G R O U P D I V I S I O N Of L O 'C K H E E D A I R C R A F T C O R P O R A T I O N

the model can be used as a thermal testbed for evaluating certain materials

and thermal-control systems proposed for use on the prototype; and(3) to

determine the proposed complexity of the model design and test program in

terms of the effect on manufacturing difficulties and costs involved, as

well as the impact on the overall program schedule.

The TMM will be utilized during the trade studies largely to determine

the impact of distortion of various heat flow fields. The TMM will show

which thermal paths are critical and which can be distorted without affect-

ing vehicle performance.

6.0 TSM DESIGN CRITERIA DEVELOBffiKE

Before starting final TSM design, definitive design criteria will be

established to provide firm guidelines for the model designers. The criteria

will be developed from the analyses and studies performed during Task 1 and

from Initial results of the trade studies. The criteria will include defini-

tion of overall modeling approach (e.g., temperature preservation, materials .

preservation, etc.) overall scale ratio, level of detail for modeling sys-

tems and subsystems,and the approach to use in modeling the various radiation,

conduction, and convection heat flow paths. Design criteria established during

Task 1 are as follows:

•£1) Use "temperature preservation" approach, T =1, for conduction-radiation

modeling.

2) Study both pure "temperature preservation" and "scaling compromises"

approaches for convection modeling.

3) Overall length scale ratio to be between 1/2.5 and l/ .

4) Assume one-dimensional modeling for external insulation and internal MLI

wherever possible.

5) Assume two-dimensional modeling for the aluminum skin structure.

6) Use identical optical properties between model and prototype for both in-

ternal and external surfaces»

A-8



7) Geometrical identity "between model and prototype for all major external

surfaces.

8) Heat exchangers and fluid loops will not be modeled.

9) Radiators to be modeled with simulated heat load applied.

10) Wings and vertical stabilizer extremities will "be deleted or heated with

infrared energy as required so that model will fit within solar simulation

capability.

11) Cylindrically shaped payload will be included in design studies per NASA/

IMSC agreement.

12) With the exception of the crew cabin, active thermal control systems will

not be modeled, however, their effects on various temperature fields will be

considered in TSM design.

7.0 TSM DETAILED DESIGN REQUIREMENTS

Throughout the initial stages of Tank 2, specific requirements will be

established for design of the TSM. These will include such items as require-

ments for structural integrity of the pressure shell and other structural areas

and requirements for modeling crew cabin windows and other discrete sections

of the vehicle. .

3.0 TSM DETAILED DESIGN DEVELOPMENT

Detailed design of the TSM will be accomplished in accordance with the

design criteria outlined in Section 6 and in accordance with the standard

modling laws developed for the temperature preservation approach. During this

design effort, modeling of the three basic modes of heat transfer—conduction,

radiation, and convection — will be considered. Deviation from strict adher-

ence to the exact modeling laws for convection and for certain conduction paths

may be necessary and is discussed .in 8.1 and 8.2.

The basic approach to be followed in design of the TSM is to select

a scale ratio which appears to satisfy the majority of the model's requirements

and then proceed to model all areas of the vehicle to that scale ratio. Scal-

ing compromises will be made for those areas of the vehicle that cannot be modeled

exactly.

. A-9



Results from computer runs of the prototype TMM will be used to guide the

model design. In addition, a TMM of the scale model will be developed for pur-

poses of evaluating the affects of scaling compromises and for comparing TSM

and prototype temperatures. Comparison of results from the two math models will

provide the means by which the model design is judged on its ability to predict

thermal behavior of the prototype.

The end result of the model design activity will be the specification of

materials, size, shape, weight, internal power sources, external heat flux re-

quirements, details of critical thermal conduction joints, details for model-

ing the crew cabin atmosphere, and details of modeled components and their

location relative to each other. A complete set of specifications and engineer-

ing drawings will be produced to document details of the TSM design* These

drawings will conform to Category A, Form 3 of Specification MIL D-1000 and

MIL STD-100.

8.1 Conduction-Radiation Modeling

In modeling the conduction and radiation heat flow paths, adherence to the* * # .* * #

modeling criteria of K = L , I = 1, and 6 = (£ C will be complied with wher-

ever possible. Due to the limitations in available materials, it is anticipated

that in modeling certain conduction paths, distortion of the thickness ratio may

be required to model the overall heat flux. Where distortion of the conductive

path is required, it will be done in such a manner that the radiant exchange be-

tween components is not affected.

In modeling conduction through the external insulation (i.e. LI-1500) and

aluminum skin, certain assumptions have to be made due to the combination of

materials involved which distort the thickness scale ratio. Based on results

derived from the TMM during Task 1, the assumption can be made that heat flow

in the LI-1500 is one-dimensional (i.e., perpendicular to the skin surface) and

that heat flow in the aluminum substrate is two dimensional (i.e. circumferential

and longitudinal in the aluminum skin). During the detailed design phase,

additional analysis will be made to determine the validity of this assumption

for all external surface areas of the vehicle.

. A-10



In modeling the interior portions of the Orbiter, it is important for the

TSM to have the ability to demonstrate the temperature environment in critical

thermal control areas and around temperature critical systems such as fuel

lines, fuel tanks, and hydraulic lines.

8.2 Convective Modeling

During Task 1 it was shown that thermal modeling of the pressurized sec-

tions of the space shuttle is feasible from a theoretical point of view. It

will be possible to obtain good correlation between the one-g test of the mock-

up and the one-g test of the model. Three methods can be used in the model

test, each of which has certain advantages and disadvantages. The three methods

are: l) scaling compromise using heat transfer coefficient preservation, 2)

scaling compromise using mass flux preservation, 3) strict temperature preserva-

tion.

The first two of these methods permits the use of any scale ratio for the

model. In addition, both heat transfer coefficient preservation and mass flux

preservation can give accurate results for a combined forced and free convec-

tion test such as will be done on the full size mockup. The third method re-

quires selection of a suitable gas and scale ratio combination. However, if

such a combination can be found then not only can the one-g performance be

compared with the full size mockup tests, but in addition zero-g simulation test

will be possible. This will give accurate data as to the thermal performance

which can be expected in orbit. Two possible gas-scale-ratio combinationsjt o

are sulfer dioxide with a L = 1/2.5 and Freon 11 with a L =1/3.

The conclusions reached during Task 1 are that accurate scale modeling can

be accomplished on the crew section of the space shuttle. If the right scale

ratio is chosen three different approaches can be tested making the model an

extremely useful tool as a TMM verifier and as a zero-g performance demonstrat-

or. The one question which remains is whether or not the modeling of the pres-

surized areas can be accomplished with enough accuracy and with reasonable costs.

A-ll

LOCKHEED PALO ALTO RESEARCH LABORATORY•St. - ' L O C K H E E D ' M I S S I L E S & s P A c E C O M P A N Y> ' A G R O U P D I V I S I O N O F I O C K H E E O A I R C R A F T C O R P O R A T I O N

During Task 2 two approaches formodeling the convective systems will be

evaluated. The first will be to evaluate the feasibility of modeling the

pressure shell. This feasibility study will include scale ratios and mater-

ials and compromises where needed. It may well be that the scale ratio chosen

for the overall model may not suit the modeling of the pressure shell. Such

compromises must be evaluated in order to determine the feasibility of the

model.

The second approach will be to determine how the model can be constructed

to act as a thermal systems tool. This will include access to the interior

and the level of detail of the interior components such as hard mounts and

equipment racks.

During Task 2'the convective system will be modeled but the detail and

effort will depend on the trade studies outlined above which will be conducted

s imultaneously.

8.3 Level of Detail for Modeling Systems

During Task 1 an assessment was made regarding the level of detail re-

quired for modeling the various systems and subsystems within the Shuttle

Orbiter. As design of the model progresses and an analysis provides better

definition of thermal behavior, it may be necessary to change the level of

detail for some of the components. The following list presents the major

systems and a brief statement about the level of detail to which they will

be modeled as presently defined:

1) Crew compartment; The cabin atmosphere and internal equipment will

be modeled to the extent discussed in Section 8.2.

2) Avionics; The avionics in the forward section are convectively

cooled by their own closed cycle air supply and as such, the details

of this section will not be modeled. The heat that the avionics

may supply to surrounding structure and to other areas of the vehicle

will be scaled in the model.

A-12



3) Landing gear and tires; This system will be modeled in .detail toprovide in the model a means of duplicating the thermal environment

of the prototype.

k) RCS and QMS: These systems will not be modeled in detail, however,

their mass specific heat and the heat that they provide to their sur-

roundings will be scaled in the model.

5) Payload bay; This section will be modeled in detail.

6) External and internal insulation; One-dimensional effects will be

modeled very closely. Two-dimensional effects will be studied and

scaled where important.

7) Skin structure and supporting structure; These members will be

modeled in detail, however geometric similarity may not be possible

to achieve due to material limitations.

8) External and internal thermal control surfaces; The optical proper-

ties between model and prototype will be preserved.

9) Wings and vertical stabilizer; The shape and major thermal control

characteristics of these sections will be modeled as closely as pos-

sible.

10) Hydraulic pumps and lines; These will not be modeled, however the

model will be designed such that the temperature environment

in which these items are located will be duplicated from prototype to

model.

11) Miscellaneous internal jpower sources; The thermal energy dissipated

from these sources will be scaled in the model.

12) Fluid loops,.heat exchangers, and radiators; The fluid loops and heat

exchangers will not be modeled. The radiators will be modeled in de-

tail, and their heat input will be supplied by attached heating ele-

ments.

13) Connecting joints and hinges; Due to lack of information about

actual thermal conductance across joints and hinges for the Shuttle,

the model will be designed assuming infinite conductance for solid

connections and zero conductance for hinged connections.

A-13


A C R O U P D I V I S I O N O f I O C K M E E O A I R C R A F T C O R P O R A T I O N

Cryogenic tanks; The extent to which these will be modeled has not

yet been decided.

9.0 POTENTIAL APPLICATIONS OF TSM FOR USE IN SHUTTLE THERMAL DESIGN

The most important application of the TSM is that of providing a tool

for preliminary and final thermal design studies, however, there are other

potential applications that must be considered for optimum utilization of a

thermal scale model. During Task 1 studies, a list was compiled of potential

applications for the TSM. These and other applications will be studied

throughout the remainder of the program, with a major goal of defining the

role of the TSM in the overall Shuttle integrated thermal design program.

Potential applications of the scale model are as follows:

1) To provide experimental data on both a system and subsystem level

during the Shuttle's initial design stages for use in optimizing

the orbiters thermal design.

2) Provide information for refinement of the prototype TMM.

3) To aid in design of the crew cabin convective cooling system.

.'**•) To study the performance of the convective cooling in a simulated

zero g environment.

5) To determine the effects of external thermal control surface specular-

ity on vehicle temperatures.

6) To study Shuttle/Payload thermal interaction, and to evaluate the

adequacy of payload thermal control.

7) To study radiator performance in a total system environment.

8) Test vehicle for external insulation t° study performance of the

thermal/mechanical design.

9) Multipurpose test object serving as a test bed for subsystem testing,

including those systems with active thermal control.

10) Thermal test vehicle for evaluating performance of thermal design

changes and for confirmation of overall system final thermal design.

11) Test vehicle for verification of launch pad thermal protection system.


A . G R O U P D I V I S I O N O F L O C K H E E D A I R C R A F T C O R P O R A T I O N

10.0 TEST FACILITY DEFINITION AND REQUIREMENTS

During Task 2, studies will be undertaken to define the test facility

requirements for TSM testing. In defining these requirements, the role that

the model will play in the overall Shuttle program must be kept in mind.

Such items as the use of solar simulation or infrared lamps for external heat

flux irradiation will be studied.

-A-15

LOCKHEED PALO ALTO RESEARCH LABORATORYL O C K H E E D M I S S I I E S & S P A C E C O M P A N YA G R O U P D I V I S I O N O F L O C K H E E D A I 8 C R A F T C O R P O R A T I O N

APPENDIX B

CONCEPT REVIEW HANDOUT

At the conclusion of Task I a program review of the work accomplished

to date was held at NASA/MSC Houston. A handout was prepared and presented

at the review. This handout which is included as an appendix gives an

outline of the work accomplished, potential model applications, test require-

ments, and program approach.

B-l


A C R O U P D I V I S I O N OF I O C K H E C 0 A I R C R A F T C O R P O R A T I O N

IMSC' D30961820 September 1972

APPENDIX B

CONCEIT EEVIEW

HANDOUT

Presented Under

NASA/M3C CONTRACT HAS 9-12991

SPACE SHUTTLE THERMAL SCALE MODELING

APPLICATION STUDY

By

LOCKHEED PALO ALTO RESEARCH LABORATORYLOCKHEED MISSILES & SPACE COMPANY, INC.

PALO ALTO, CALIFORNIA

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