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HREC-6370-2LMSC-HREC TR D306346

LOCKHEED MISSILES & SPACE COMPANY, INC.

HUNTSVILLE RESEARCH & ENGINEERING CENTER

HUNTSVILLE RESEARCH PARK

4800 BRADFORD DRIVE, HUNTSVILLE, ALABAMA

SUMMARY OF RESULTSOF PARAMETRIC STUDIES OF SPACE

SHUTTLE BOOSTER, ORBITER,AND LAUNCH VEHICLE

CONCEPTS

31 December 1972

Contract NAS8-26370

Prepared for National Aeronautics and Space AdministrationMarshall Space Flight Center, Alabama 35812

by

Dale BradleyRobert E. Buchholz

APPROVED:B. Hobson Shirley, Super or

Aerophysics Section

$'1/ IS. Farrior

fesident Director

LMSC-HREC TR D306346

FOREWORD

This report presents the results of analytical and experimental space

shuttle aerodynamic studies conducted by the Lockheed Missiles & Space

Company, Inc., Huntsville Research & Engineering Center (Lockheed-Huntsville),

while under contract from the Aero-Astrodynamics Laboratory of the George C.

Marshall Space Flight Center (Contract NAS8-26370). The experimental results

presented in this report were obtained in tests conducted between 15 June 1971

to 31 December 1972. These studies were conducted by the Aerodynamic De-

sign Group of the Aerophysics Section of Lockheed-Huntsville. This report

has been prepared in response to the requirement of the subject contract for

a final summary report. The NASA Technical Monitor of this contract is

Mr. Paul E. Ramsey, S&E-AERO-AAE.

ACKNOWLEDGEMENT

The authors are grateful to Roger R. Ellis, James A. Moore, Robert

A. Lott, Douglas J. Elder and Mickey D. Gamble, all of Lockheed-Huntsville,

who made substantial contributions to this analytical/experimental study.

ii

LOCKHEED- HUNTSVILLE RESEARCH & ENGINEERING CENTER


SUMMARY

The results of analytical and experimental parametric studies of space

shuttle booster, orbiter and launch vehicle aerodynamics are described for

the period 15 June 1971 through 31 December 1972. Results of studies con-

ducted from 1 July 1970 to 15 June 1971 have been previously reported in an

interim report. These studies were conducted by the Aerodynamic Design

Group, Aerophysics Section of Lockheed-Huntsville while under contract to

NASA-George C. Marshall Space Flight Center, Contract NAS8-26370. During

this study over 1700 hours of experimental wind tunnel tests were conducted

by Lockheed on several versions of the shuttle booster, orbiter and launch

vehicle. Fifteen separate tests were conducted in three different test facilities.

Test data were published by NASA-MSFC through a separate contract with

Chrysler. The test data documents are referenced later in this report. Due

to the number of test programs conducted and the time required for test prep-

aration, analysis of the test data has been limited to that required to drive the

experimental program. No documentation of the data analysis has been pub-

lished nor will it, be included in this report. A brief description of each of the

experimental tests conducted including the test purpose and approach is in-

cluded in this report. Several test models were designed and fabricated by

Lockheed for NASA-MSFC in support of the experimental program. These

models are described in this report in the sections which relate to a specific

test program.

oio. .



CONTENTS

Section Page

FOREWORD ii

SUMMARY.. iii

NOMENCLATURE ix

1 INTRODUCTION 1

2 DESCRIPTION OF EXPERIMENTAL STUDIES. 5

2.1 Parametric Study of a 0.00325-Scale Orbiter Model(TWT 498 and TWT 499) 7

2.2 Static Stability and Control Study of a 0.004-ScaleModel Orbiter (TWT 542) 24

2.3 Cruise Engine Placement and Lateral-DirectionalStability Study for a 0.015-Scale Flyback Booster(NSRDC 3310) 34

2.4 Base Drag Reduction Study on a 0.015-Scale FlybackBooster (CAL-T-18- 0 63) 50

2.5 Static Stability and Trim Characteristics of a 0.00227-Scale Parametric Pressure-Fed Booster (TWT 526) 68

2.6 Reentry Study of a 0.00513-Scale Solid RocketBooster (TWT 541) 84

2.7 Static Stability and Control Study of a 0.004-ScaleParametric Launch Vehicle (TWT 544 and TWT 544X) 97

2.8 Orbiter Pressure Distribution of a 0.004-Scale OrbiterWhile Mounted in the Launch Configuration (TWT 550) 121

2.9 Pressure Distribution on 0.004-Scale HO Tank and SRMsWhile Mounted in the Launch Configuration (TWT 543) 129

2.10 Mutual Interference Study of the Orbiter, HO Tank andSolid Rocket Booster of a 0.004-Scale Space ShuttleLaunch Configuration 140

3 REFERENCES 155

iv



LIST OF ILLUSTRATIONS

Chart Pag e

1 Evolution of Experimental Tests for Contract NAS8-26370Showing Relative Sizes of Vehicles 2

Tables

1 Experimental Test Programs Conducted Under ContractNAS8-26370 6

2 0.00325-Scale HCR Orbiter Reference Dimensions 10

3a Run Schedule for TWT 498 11

3b Run Schedule for TWT 499 16

4 0.004-Scale Space Shuttle Orbiter Reference Dimensions 26

5 Run Schedule for TWT 54Z 27

6 0.015-Scale Booster Reference Dimensions (NSRDC 3310) 38

7 Run Schedule for NSRDC 3310 39

8 0.015-Scale Booster Reference Dimensions (CAL-T-18-063) 53

9 Run Schedule for CAL-T-18-063 ' 54

10 0.00227-Scale Parametric Pressure-Fed BoosterReference Dimensions 71

11 0.00227-Scale Parametric Pressure-Fed Booster Base Areas 72

12 Run Schedule for TWT 526 73

13 0.00513-Scale 156-Inch SRM Reference Dimensions 87


15 0.004-Scale Space Shuttle Launch Configuration ReferenceDimensions 100


17 Run Schedule for TWT 544X 106

18 Orbiter Surface Pressure Tap Locations 123



21 North American Launch Configuration ReferenceDimensions 143


Figur e s

1 Cross Section of the MSFC 14 x 14-Inch Trisonic WindTunnel 18

v



Figure Page

2a General Arrangement-Orbiter Model 2A 19

2b General Arrangement - Orbiter Model 2B 20

3 HCR Orbiter and Associated Model Parts 21

4a Installation Photograph, Model 2A (TWT 498) 22

4b Installation Photograph, Model 2B (TWT 498) 23

5 General Arrangement, Space Shuttle Double Delta Orbiter 30

6a Installation Photograph, Double Delta Orbiter (TWT 542) 31

6b Installation Photograph, Double Delta Orbiter (TWT 542) 32

7 Double Delta Orbiter with Associated Parts 33

8 General Arrangement of the Baseline Flyback Booster Model 47

9 Installation Photographs, Flyback Booster (NSRDC 3310) 48

10 Flyback Booster Engine Configuration and Location ofDummy Engine Pods, G6 49

lla Flyback Booster Base Plenum Orifice Location (End View) 58

llb Flyback Booster Base Plenum Orifice Location (Side View) 59

12a Flyback Booster Body Flaps (End View) 60

12b Flyback Booster Body Flaps (Side View) 61

12c Flyback Booster Body Flaps (Top View)- 62

13a Flyback Booster Base Venting (Side View) 63

13b Flyback Booster Base Venting (Top View) 64

14a Installation Photograph, Flyback Booster Base Plenum(CAL-T-18-063) 65

14b Installation Photograph, Flyback Booster Base Flaps(CAL-T- 18-063) 66

14c Installation Photograph, Flyback Booster Base Vent(CAL-T-..18-063) 67

15a 0.00227-Scale Parametric Pressure-Fed Booster ModelGeometry (Cones, Cylinders and Flares) 79

15b 0.00227-Scale Parametric Pressure-Fed Booster ModelGeometry (Fins) 80

16 0.00227-Scale Parametric Pressure-Fed Booster Model Parts 81

17a Installation Photograph, Pressure-Fed Booster (TWT 526) 82

17b Installation Photograph, Pressure-Fed Booster (TWT 526) 83

18a General Arrangement, 0.00513-Scale 156-Inch Solid RocketMotor Geometry 91

18b Strake Geometry 92

vi



Figure Page

19a Mounting Arrangements for 156-Inch SRM, Angle of Attack-:10 to 90 Degrees 93

19b Mounting Arrangements for 156-Inch SRM, Angle of Attack90 to 190 Degrees 94

20a Installation Photograph, Solid Rocket Motor (TWT 541) 95

20b Installation Photograph, Solid Rocket Motor (TWT 541) 96

21a 346-Inch HO Tank With Three Alternate Noses and One-BodyDiameter Extension 110

21b 400-Inch HO Tank With Two Alternate Noses 111

21c T8 346-Inch HO Tank Nose Cone 112

21d T9 312-Inch Diameter HO Tank 11321e T10 346-Inch Diameter HO Tank 114

22a 156-Inch Solid Rocket Motor with Standard and SkewedNoses and One-Body Diameter Extension 115

22b 178-Inch Solid Rocket Motor 116

23 HO Tank Ventral Fin 117

24 Baseline Double Delta Wing. Orbiter Launch Vehicle 118

25a Installation Photograph, Double Delta Wing OrbiterLaunch Vehicle (TWT 544) 119

25b Installation Photograph, Double Delta Wing OrbiterLaunch Vehicle (TWT 544) 120

26 Static Pressure Tap Positions for Double Delta Orbiter 126

27a Installation Photograph, Orbiter Pressure Test (TWT 550) 127

27b Installation Photograph, Orbiter Pressure Test (TWT 550) 128

28 General Arrangement for HO Tank-SRM Pressure Test 133

29 SRM Orientations with Respect to the HO Tank for s = 750 134

30a HO Tank Pressure Orifice Location 135

30b SRM Pressure Orifice Location 136

31 Sting Support for HO Tank - SRM Pressure Test 137

32a Installation Photograph, HO Tank - SRM Pressure Test(TWT 543) 138

32b Installation Photograph, HO Tank - SRM Pressure Test(TWT 543) 139

33 General Arrangement North American Orbiter 01 147

34 General Arrangement North American External Tank T3 148

vii



Figure Page

35 General Arrangement North American SRB S1 149

36 Space Shuttle Parallel Staging System for the MSFC14 x 14-Inch Trisonic Wind Tunnel 150

37 SRB Nonmetric Mounting Sting 151

38 General Arrangement North American Launch ConfigurationT301 S1 152

39a Installation Photograph, North American LaunchConfiguration on Dual Sting (TWT 545) (Metric SRB) 153

39b Installation Photograph, North American LaunchConfiguration on Dual Sting (TWT 545) (Nonmetric SRB) 154

viii



Symbol

CND POS

i

m

S

STK LOC

WNG POS

X

y

Zz

Greek

a

F

6

8

NOMENCLATURE

D e finition

canard position for flyback booster; 1 is high position,2 is low position

incidence angle, deg

length, in.

mass flow rate, lb-mass per second

surface area, in

strake location for solid rocket motor, deg, left sideof body is zero position

wing position for flyback booster; 1 is high position,2 is low position

longitudinal position, in.

lateral position, in.

vertical position, in.

angle of attack, deg

angle of sideslip, deg

dihedral angle, deg

control surface deflection, deg

semivertex angle, deg

roll angle or in the case of solid rocket motors in thelaunch configuration, their position relative to the HOtank, 0 being at the top of the tank, deg

ix



Subscripts

a

b

C

CY

e

F

L

NOZ

ORB

R

r

SRM

STK

T

w

aileron

base

canard

cylinder

elevon

flap or flare

left

nozzle

orbiter

right

rudder

solid rocket motor

strake

tail

wing

x

LOCKHEED -HUNTSVILLE RESEARCH & ENGINEERING CENTER

HUNTSVILLE RESEARCH & ENGINEERING CENTER P. O. BOX 1103 WEST STATION HUNTSVILLE, ALABAMA 35807

& SPACE In reply refer to:LMSC-HREC D306412

COMIPANY 18 January 1973

National Aeronautics and Space AdministrationGeorge C. Marshall Space Flight CenterMarshall Space Flight Center, Alabama 35812

Attention: A& TS-PR-M

Subject: Contract NAS8-26370, Submittal of Summary Report

Gentlemen:

Enclosed, is the Summary Report as required under the subject contract.

Very truly yours,

LOCKHEED MISSILES & SPACE COMPANY, Inc.,

N,?

7HuntsVile Research & Engineering Center

JEE: pgp

Encl: LMSC-HREC TR D306346, HREC-6370-2, "Summary of Results ofParametric Studies of Space Shuttle Booster, Orbiter, and LaunchVehicle Concepts," Summary Report, Contract NAS8-26370.

cc: DCASO, DCRA-DBGHC (C. L. Weber)

Distribution:A&TS-MS-IL, one copyA&TS-MS-TU, one copy

~A&TS-MS-IP, two copiesS&E-AERO-R, seven copies plus one reproducible

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


Section 1

INTRODUCTION

The NASA-George C. Marshall Space Flight Center (MSFC) is currently

conducting extensive technology studies to support the aerodynamic develop-

ment of a space shuttle system. The Aerodynamic Design Group, Aerophysics

Section, of Lockheed-Huntsville is under contract to provide assistance to

MSFC in these studies. Lockheed-Huntsville overall objectives are: (1) plan

and conduct analytical studies of booster and orbiter geometry as an aid in

specifying the parameters to be studied experimentally; (2) provide liaison

between MSFC and contractors during the design and manufacture of test

models; (3) plan, conduct and coordinate experimental wind tunnel tests of the

booster, orbiter and launch vehicles in the test facilities which will be pro-

vided by MSFC; and (4) analyze and document results of all studies. The

Lockheed-Huntsville support began in July 1970 when the original contract

was awarded and was continued by an extension in April 1971 and another in

January of 1971. The initial study was based on completely reusable flyback

booster and orbiter configurations (MSFC delta-wing booster and MDAC model

256-14 orbiter). Lockheed provided assistance inthe design of the models,

in planning the experimental test programs, and conducting the tests. Results

of studies conducted between 1 July 1970 and 15 June 1971 have been reported

previously in Ref. 1.

* Booster Studies

The progression of the test program and overall summary of this study

is presented in Chart 1. Tests of the booster model were conducted in the

7 x 10-foot transonic wind tunnel at the Naval Ship Research & Development

Center (NSRDC) and in the 8 x 8-foot transonic wind tunnel at Cornell Aero-

nautical Laboratory (CAL) (Refs. 2 through 5). The flyback booster static

stability and control characteristics were measured for variations in geometry



Orbiter

TWT 494,498,499

Solid I

TWT 544, 544X,550, 543

545

Chart I - Evolution of Experimental TestsSizes of Vehicles

for Contract NAS8-26370 Showing Relative

2


such as high and low wing and canard positions, wing and canard exposed area,

vertical surface size and location of cruise engines. A cold flow non-metric

engine simulation test was conducted to measure the interference effects of the

cruise engine exhaust on the booster aerodynamics including control effective-

ness. A short test was also conducted to determine possible techniques for

reducing the large base drag of such a blunt based configuration (Ref. 5).

Shortly after the completion of these booster tests (November 1971) the de-

cision was made by NASA to drop the flyback concept in favor of water re-

coverable boosters. A parametric water recoverable booster model was then

tested (Ref. 6, February 1972) to determine the high angle of attack (50 deg <

a < 90 deg) aerodynamic characteristics of such a concept. Included in the

geometrical parameters of the cone-cylinder-frustum-fin concept were: cone

angle, cylinder length, frustum angle and fin exposed area. Results from this

test were to be used by MSFC in preliminary design efforts. The last booster

alone test conducted under this study (Ref. 7, May 1972) was to obtain data on

a solid rocket booster design at very large angles of attack (-10 deg < a < 190

deg), to simulate all possible reentry attitudes.

* Orbiter Studies

Tests of the MDAC model 256-14 orbiter began in June 1971 (Ref. 8)

but this configuration was abandoned by MSFC in favor of a more parametric

symmetrical airfoil model. Lockheed aided MSFC in the preliminary design

of the new model with results of analytical studies. Tests of the new orbiter

configuration were conducted in August 1971 to determine the static stability

and control effectiveness and the effects of wing tip fins, a center dorsal fin,

wing exposed area and aspect ratio, and elevon exposed area (Ref. 8). Soon

after these tests were completed NASA abandoned the large reusable orbiter

in favor of an external fuel tank configuration which required a much smaller

orbiter. In support of the development of the new orbiter Lockheed designed

and fabricated a 0.004-scale model and tested it to determine the static sta-

bility and control characteristics (Ref. 9). This test was conpleted in May 1972.

In late August 1972 Lockheed was directed by MSFC to design and fabricate a

3


LMSC-HREC TR:D306346

0.004-scale model of the North American Rockwell (NAR) orbiter. This task

was completed by Lockheed and the model was tested by the MSFC COR (Mr.

Ramsey) during September 1972. An additional test was planned by Lockheed

to determine orbiter wing panel loads (Ref. 10) but was cancelled by MSFC.

* Launch Vehicle

Launch vehicle tests (Ref. 11) began in May 1972 to determine the aero-

dynamic characteristics of the launch configuration and to study the effect of:

orbiter incidence relative to the external fuel tank (HO tank) and solid rocket

motors (SRMs); SRM longitudinal and radial location on the HO tank; and

orbiter-HO tank separation distance. Orbiter wing panel and body centerline

pressures were measured while mounted on the launch vehicle in a test con-

ducted in July 1972. Orbiter iniidence angle, SRM radial location on the HO

tank and HO tank nose shape were varied during the test. The pressure dis-

tributions over the HO tank and SRMs when mounted with the orbiter as the

launch configuration were obtained in a test conducted during August and

September 1972 (Ref. 12). Models for this test program were designed and

fabricated by Lockheed from Emerson and Cumings Stycast material. The

final test conducted under this contract was a launch vehicle test to determine

the mutual interference effects of the orbiter, the HO tank and the SRMs (Refs.

13 and 14). The orbiter, HO tank and SRMs were all mounted on separate stings

in order that all mutual interference effects could be determined.

* Analysis of Test Data

Analysis of the test data obtained during this study was limited to that

required to drive the experimental program. It should also be mentioned that

several of the tests conducted were of a very general nature and are worthy

of detailed analysis. Such analysis was not possible under this contract be-

cause of the rapid response required to keep pace with the shuttle development.

Data analysis reports were published by Lockheed concerning the analytical

trade studies of a LOX-RP booster concept (Ref. 15) and when it was deemed

necessary to document possible sources of data anomalies (Refs. 16 and 17).

4



Section 2

DESCRIPTION OF EXPERIMENTAL STUDIES

A chronological listing of the experimental tests conducted during this

study is included as Table 1. These tests, the test models and test facilities

are described in the following sub-sections. For each test the test

purpose, data reduction and a brief discussion of the test operation are pre-

sented.

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2.1 PARAMETRIC STUDY OF A 0.00325-SCALE ORBITER MODEL(TWT 498 and TWT 499)

This test was an extension of the Space Shuttle orbiter tests performed

during the first phase of this contract. The orbiter body was identical to the

orbiter model previously tested (Model 1) but the wing was of higher aspect

ratio. The model had provision for either tip fins or a center dorsal fin. Data

from this test were to be compared with Model 1 data to provide a basis for

optimization of orbiter performance and stability and control. The test was

conducted at Mach numbers ranging from 0.6 to 4.96, angles of attack from

-4 to +50 degrees and angles of side slip from -9 to +9 degrees at angles of

attack of 0,10 and 15 degrees. A pretest report for this test was published

in April 1971 (Ref. 8).

2.1.2 Test Facility

A complete description of the MSFC 14 x 14-Inch Trisonic Wind Tunnel

is presented in Ref. 18 from which the following excerpts were taken:

"The tunnel is an intermittent trisonic blowdown tunnel operatedfrom pressure storage to vacuum or atmospheric exhaust. The testsection measures 14 x 14 inches in two of the interchangeable testsections. The transonic section provides for Mach numbers of 0.20through 2.50 and the supersonic section provides for Mach 2.75 through5.00.

"Air is supplied to a 6000 cubic foot storage tank at -40°F dewpoint and 500 psia. The compressor is a three-stage reciprocatingunit driven by a 1500 hp motor.

"The tunnel flow is established with a servo-controlled gate valve.Air from the control valve flows through the valve diffuser into thestilling chamber where the air can be heated up to 200°F. Air thenflows into the test section which contains the nozzle blocks and testarea.

"Speeds are varied in the subsonic range by a controllable diffuser,in the transonic range by perforated tunnel walls, in the low (1.5 to 2.5)supersonic range by tilting fixed contour nozzle blocks.

"The transonic section has variable porosity walls that allow foroptimum wave cancellation in the transonic flow region.

7



"Downstream of the test section is a hydraulically controlledsector that provides for angles of attack of +10 degrees with variousoffsets extending the pitch limits to 90 degrees.

"The variable diffuser, with its movable floor and ceiling panels,is the primary means for controlling the subsonic speeds; it also allowsfor more efficient supersonic runs. The sector assembly and diffusertelescope to allow easy access to the model and test section.

"The tunnel flow is then exhausted through an acoustically dampedtower to atmosphere or into the vacuum field of 42,000 cubic feet. Thetanks are evacuated by five vacuum pumps driven by a total of 500 hp.

"Data are recorded by a solid state digital data acquisition system.The digital data are transferred to punched cards during the run to bereduced later to proper coefficient form by a computer."

Fig. 1 shows a cross section of the tunnel.

2.1.3 Model Description

The model tested was a 0.00325-scale modified version of a McDonnell-

Douglas Space Shuttle orbiter. The body was identical to the unmodified con-

figuration, but the wing was of higher aspect ratio (AR = 2.0 for the no tip fin

case, 2A) and contained no twist or camber. The model was tested in two

forms, with straight wing tips and centerline dorsal fin (Model 2A) and with

wingtip fins (Model 2B). The elevons and rudders are capable of being de-

flected to provide control effectiveness data. Sketches of the models are

shown in Figs. 2a and 2b.

Model nomenclature for the various model components is listed below:

Symbol Definition

B2 model 2A and 2B body

D1 dorsal fin (Se = 900.3 ft2 )

D2 dorsal fin (Se = 722.1 ft )

e3 model ZA elevon

e4 model 2B elevon

r l rudder-dorsal D 1

r2 rudder-dorsal D2

8



r4 rudder-tip fin V22

V2 tip fin-wing W5 (S e = 632.4 ft 2 )

V4 tip extension-wing WZ (S e = 240 ft ea)

W2 model 2A wing (Se = 3963 ft )

W5 model ZB wing-w/o tip fins (S e = 3468 ft 2 )

All model parts are shown in Fig. 3.

Figure 4 shows the model installed in the tunnel. The model is fabricated

of 17-4 PH stainless steel. Model design and construction was performed by

the Naval Ship Research and Development Center, Carderock, Maryland.

2.1.4 Data Reduction

Model forces and moments were measured by MSFC balances 201 and

231. All forces and moments were reduced to nondimensional coefficients.

Reference dimensions used for data reduction are listed in Table 2. Data

were corrected for weight tares and sting deflections. The data for TWT 498

were entered into the Chrysler Corporation's System for Analysis and De-

velopment of Static Aerothermodynamic Criteria (SADSAC) and published as

a data report (Ref. 19). TWT 499 test data were not published because of

possible data discrepancies due to blockage and sting interference.

2.1.5 Discussion

An outline of the configurations tested is shown in Table 3a for TWT 498

and Table 3b for TWT 499. The parametric portions of this test were run first.

During the parametric investigation, anomalies in the data were discovered.

To investigate these anomalies the test was extended and various stings were

tested with the basic model to determine the effects of sting interference and

tunnel blockage. The results of these studies are published in Refs. 16 and 17.

The combined tests required 227 wind tunnel occupancy hours to complete.

9



Table 2

0.00325-SCALE HCR ORBITER REFERENCE DIMENSIONS

10

LOCKHEED - HUNTSVILLE RESEARCH & ENGINEERING CENTER

Parameter Full Scale Model Scale

22Reference Area (S ) 5935 ft 0.063 ft 2

~~~~ref ~~~(9.025 in )

Reference Length (eref' b)

Models 1,2A and 2B 157.487 ft 0.512 ft(6.142 in.)

Balance Location (BMC)Models 1, 2A and 2B (from nose) - 0.352 ft

(4.222 in.)

Moment Reference CenterModels 1, 2A and 2B (from nose) 102.367 ft 0.333 ft

(3.996 in.)(above model ~) 0 .0 0.0

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Model 1 (baseline) 362.50 ft 2 0.004 ft 2

(0.551 in )

Models 2A and 2B 303.20 ft 2 0.003 ft2(0.461 in )

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2.2 STATIC STABILITY AND CONTROL STUDY OF A 0.004-SCALEMODEL ORBITER (TWT 542)

2.2.1 Test Purpose

The NASA decision to adapt a parallel burn launch concept with an orbiter

external Hydrogen-Oxygen (HO) fuel tank resulted in a redesign of the orbiter

airframe. The purpose of this test was to determine performance and stability

and control for this new orbiter. The test was conducted at Mach numbers

from .6 to 4.96, at angles of attack from -4 to 30 degrees and angles of side

slip from -6 to 6 degrees at angles of attack of 0 and 10 degrees. A pretest

report for this study was published in May 1972 (Ref. 9).

2.2.2 Test Facility

The test was conducted in the MSFC 14 x 14-Inch Trisonic Wind Tunnel.

See Section 2.1.2 for a description of this facility.


A sketch of the model is shown in Fig. 5 and installation photographs

are included as Fig. 6. The model nomenclature is as follows:

Symbol De s c ription

Al abort SRM pods

B1 orbiter body, including canopy andhousing along top centerline

B2 B1 with off-block for body alone

P1 ACPS pods (wing upper surface)

Vil twin vertical tails (5-degree LEsemi vertex angle)

W4 wing (including glove)

R2 nozzle shroud

F2 body flap

24



The model was designed and constructed by Lockheed-Huntsville. The body

and wings are made of aluminum and the vertical tail and control surfaces

were fabricated from 17-4 PH stainless steel. Figure 7 is a photograph of

the completed model showing all component parts.


Model forces and moments were measured by MSFC balance Z01. All

forces and moments were reduced to coefficient form using the reference

dimensions shown in Table 4. The data were corrected for sting deflections

and weight tares.

The data were entered in the Chrysler SADSAC program and published

as a data report (Ref. 20).

2.2.5 Discussion

An outline of the configurations tested is shown in Table 5. This was the

first test of the new NASA orbiter configuration. As such, the data were of

extreme interest to MSFC and the Manned Spacecraft Center (MSC). Unfor-

tunately this test was severely curtailed by failure of the roll gage of the

balance, requiring that all runs after the failure be pitch plane runs only.

Wind tunnel occupancy time for the shortened test was 58 hours.

25



Table 4

0.004-SCALE SPACE SHUTTLE ORBITERREFERENCE DIMENSIONS

26



Reference Area (Sref) 3420.0 ft 2 7.880 in 2

(Wing Theoretical Area)

Reference Length ( ref) 5070 in. 2.028 in.(M.A.C.)

Reference Span (bref) 1115.0 in. 4.460 in.(Wing Span)

Balance Location (BMC) 3.883 in.(Balance Moment Center)

(from Nose)

Moment Reference Center 92D.5 in. 3.682 in.(MRP) (0.70 AB)

(from nose)

2~~~~Base Area (Ab) 317.7 ft z 0.732 in 2

Cavity Area (Ac) 0.314 inc~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~

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2.3 CRUISE ENGINE PLACEMENT AND LATERAL-DIRECTIONALSTABILITY STUDY FOR A 0.015-SCALE FLYBACK BOOSTER(NSRDC 3310)

2.3.1 Test Purpose

Earlier tests (Ref. 1) of the Lockheed-Huntsville modified McDonnell-

Douglas flyback booster had shown unfavorable lateral-directional character-

istics (NSRDC 3110) and suggested favorable cruise engine placement location

(NSRDC 3210). Further modifications to improve these areas were investi-

gated in tests conducted during August 1971. The model was tested at Mach

numbers of 0.4 to 1.2 at angles of attack from -4 to 20 degrees and at angles

of sideslip from -6 to 6 degrees while at angles of attack of 0, 6 and 15 degrees.

A pretest report was published in August 1971 for this test (Ref. 4).

2.3.2 Test Facility

The wind tunnel used for this test was the 7 x 10-foot transonic facility

at the Naval Ship Research & Development Center. This facility is described

in detail in Ref. 21, from which the following excerpts were taken.

"The drive system is composed of two 12,000 hp synchronouselectric motors which operate at 720 revolutions per minute. Eachmotor is connected to a fan by a variable-speed dynamatic coupling.There are two stages of axial flow fans, each 18 feet in diameter.

"Two air dryers maintain condensation-free flow in the tunnelcircuit at all Mach numbers.

"The tunnel cooling system consists of a heat exchanger throughwhich water is circulated to maintain a stagnation temperature of lessthan 140OF in the tunnel circuit.

"Evacuation pumps provide a range of test stagnation pressuresfrom atmospheric to about 700 lb/ft2. By venting the test sectionarea to atmospheric pressure, a stagnation pressure of approxi-mately 1.75 atm can be reached at a maximum Mach number of 0.9.The wind tunnel operates continuously over a range of Mach numbersfrom 0.4 to 1.17 at the reduced pressure condition.

34

LOCKHEED- HUNTSVILLE RESEARCH & ENGINEERING CENTER. .?.


"The test section is rectangular and measures 7 ft in height and10 ft in width. The floor and ceiling are slotted and have diffuser flapsat the downstream end of each slot to control flow reentry and Machnumber.

"The main support system, a cantilevered boom and sting mountedon a vertical strut, has three degrees of freedom, all of which areresolved about a fixed point in the test section (tunnel station 100.787).The range of angular displacements are:

Angle of Attack -4 to 29 deg

Sideslip -25 to 25 deg

Roll -180 to 180 deg

"As many as 12 forces and/or moments may be read out withoutadditional equipment. A Fischer-Porter pressure recording unit willmeasure and read out as many as 100 pressures.

"The raw data are stored on a punched paper tape, and are tabulatedby a Flexowriter."


The model tested was a 0.015-scale modified McDonnel-Douglas flyback

booster. A general arrangement of this model is shown in Fig. 8. The basic

model can be tested with high or low wing positions, high or low canard posi-

tions, variable wing and canard incidence, variable wing dihedral, variable

body length, wing tip or dorsal vertical fins, and canard flap, elevon and rudder

control deflections.

Model nomenclature for the various model components is listed below.

B1 modified MDAC 256-14 body(. = 3.453 ft)

C2 aerodynamic canard (Se = 0.169 ft2

D1 dorsal fin (Se = 0.442 ft2 )

Fl canard trailing edge flap (cf = 20%c)

F2 canard trailing edge flap (cf = 40%c)

G5 dummy sidebody mounted engine pods,high position

35



G6 dummy sidebody mounted engine pods,low position

T1 dorsal fin end plate. Trailing edgeflush with dorsal tip trailing edge(S = 0.087 ft2 )

T2 dorsal fin end plate. Trailing edgeflush with dorsal tip trailing edge(S = 0.070 ft2 )

V1 wing tip vertical fin (Se = 0.248 ft 2 each)

V2 wing tip extension (S = 0.15 ft 2 each)

V3 identical to V1 except rolled out 40°2 '

V4 identical to V1 except mounted inverted

V5 twin ventral fins constructed of sheetaluminum. Trailing edge of fin mountedflush with booster base. Rolled out 10 degfrom vertical (S = 0.113 ft 2 each)

W1 baseline wing (Se = 0.891 ft 2 )

The model was designed and constructed at the Naval Ship Research and

Development Center, Carderock, Maryland. Figure 9 shows several installed

configurations.


Model forces and moments were measured by NSRDC balance TSB-24.

All forces and moments were reduced to nondimensional coefficients. Refer-

ence dimensions used for data are presented in Table 6. The data were cor-

rected for tunnel flow angularities and sting deflections. Axial force was

corrected for weight tares.

The data were entered in the Chrysler SADSAC program, and published


2.3.5 Discussion

Several combinations of vertical fins and wing dihedral angles were in-

vestigated during the test to gather information on the vehicle lateral-directional

36


LMiSC-HREC TR D306346

characteristics. It is extremely difficult to achieve a proper ratio of lateral-

to-directional stability for vehicles of this type because of the extreme aft (70%

of body length) location of the center of gravity. Data from this test were used

to select a configuration that would provide a reasonable ratio of lateral-to-

directional stability while still maintaining a workable configuration with regard

to performance and control, and having no geometrical restraints upon landing.

Previous tests had suggested that the best engine location would be forward

on the body. This position exhibited less drag and was favorable from a weight

and balance standpoint. However, there was a disadvantage as the original loca-

tion of the engines (forward under the body) caused a loss in directional stability.

The two engine locations investigated during this test did not exhibit the de-

creased directional stability of the former under the body location tests. Figure

10 shows the dummy engine pod.

Control deflections were also obtained during this test for the wing in the

low position, and for the wing with negative dihedral in the high position.

Table 7 lists the configurations tested. The test required 56 hours of

wind tunnel occupancy time to complete.

37



Table t

0.015-SCALE BOOSTER REFERENCE DIMENSIONS (NSRDC 3310)

38



Reference Area(S ef)6020 ft2 1.355 ft

2

Reference Length( fb)229 ft 3.453 ft

Balance Locationbody station 2. 694 ft

Moment Reference Centerbody station (c.g.) 172.92 ft 2.594 ftwater line (above i) 1.25 ft .01875 ft

Base Area (Ab)

high wing 1076 ft 0.242 ft 2

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49


2.4 BASE DRAG REDUCTION STUDY ON A 0.015-SCALE FLYBACKBOOSTER (CAL-T-18-063)

2.4.1 Test Purpose

The base drag of the flyback booster is as much as 50% of the total ve-

hicle drag at zero lift angle of attack. Any reduction of base drag would

significantly improve cruise characteristics of the booster. Three methods

of reducing base drag were evaluated during this test. Test conditions in-

cluded a Mach number range of 0.4 to 1.1, an angle of attack range of -4 to

20 degrees and angles of sideslip from -6 to +6 degrees while at angles of

attack of 0, 6 and 15 degrees. A pretest report for this study was published

in September 1971 (Ref. 5).

2.4.2 Test Facility

The test was conducted in the 8 x 8-Foot Transonic Wind Tunnel of the

Cornell Aeronautical Laboratory, Inc., Buffalo, N.Y. The tunnel is described

in Ref. 23, from which the following excerpts were taken.

"The tunnel has a perforated throat and an auxiliary pumping systemfor plenum pumping. The continuous circuit tunnel is capable of operatingfrom 1/6 to 2-1/2 atm total pressure thereby providing a wide range oftest Reynolds numbers as well as Mach numbers. The range of operatingpressures is necessarily limited by the total power available at the higherMach numbers.

"Angles of roll and attack as established by the roll and vertical strutpitch mechanisms are accurate to within +0.10 degree.... A considerableimprovement in model accuracy can be realized by installing electrolyticbubbles inside the model thereby eliminating sting-balance deflectionsand sting joint hysteresis. Accuracies of +0.02 degrees are possible withthis system.

"The storage capabilities for either auxiliary air or nitrogen consistsof three farms.... The total storage capacity is approximately 280,000cubic feet of gas at 3000 psig. The air is supplied from a 1000 hp Clarkreciprocating air compressor which is rated at 2.365 pounds per secondat 3000 psig and -70OF dew point for continuous duty.

50



"Thirty-three readouts are available for simultaneous reading of data.

"Digital data from the readouts are recorded in punch cards which areintroduced into the system through an IBM 1442 card punch and reader."


The basic model has been described previously in Section 2.3.3. For

this test modified base regions were fitted to the basic model. Drawings of

the modified regions are shown in Figs. 11 through 13. Figure 14 shows the

actual base regions as installed on the model. The model nomenclature for

this study is listed below.

Symbol Description

B4 base plenum configuration

B5 base flap configuration

B6 base venting configuration

B7 same as B4 except has rocket nozzlesremoved

C2 aerodynamic canard (Se = 0.169 ft 2 )

F2 canard trailing edge flap (cf = 40%c)2V1 wing tip vertical fin (S e = 0.248 ft each)

W3 baseline wing modified for new baseshape


Cornell balance number CAL-TASK-Mk-XIX was used to measure balance

forces. Data were corrected for tunnel flow angularities. A bubble pack mounted

in the model was used to measure angle of attack so no sting deflection cor-

rections were necessary. Axial force was corrected for weight tares. All

model forces and moments were reduced to coefficient form using the dimen-

sions shown in Table 8.

The base flow plenum total pressure taps and instrumentation to compute

mass flow rate from the base nozzles were fitted to monitor base flow conditions

during that portion of the test.

51




as a data report (Ref. 24). The tabulated data were also published as a Cornell

data report (Ref. 25).

2.4.5 Discussion

The methods used in an attempt to reduce base drag were:

1. Base flaps were used to induce freestream air into thebase region,

2. A vented base was employed to open the base region tofreestream air, and

3. Air was blown into the base region.

All three of these methods depend upon the use of air with relatively high total

energy to relieve the low pressure, stagnated air present in the base region.

The fitting of these devices required that the base be reworked, so a short

transonic study and a control effectiveness study was done to determine the

aerodynamic effects of the modified base. A summary of the configurations

tested is shown in Table 9.

None of the devices tested were particularly successful in reducing

base axial force. There was very little or no change for the base flap or base

vent methods and a significant increase for the blown base method. It is felt

that the blowing base was unsuccessful because the nozzles were too localized

and the flow velocity too high.

Twenty-eight wind tunnel occupancy hours were required to complete

the test.

52



Table 8

0.015-SCALE BOOSTER REFERENCE DIMENSIONS (CAL-T-18-063)

53

LOCKHEED-HUNTSVILLE RESEARCH & ENGINEERING CENTER


Reference Area(Sref) 60Z0 ftZ 1.355 ft 2

refReference Length( refb) 229 ft 3.4'53 ft

Balance Locationbody station 2 694 ft

Moment Reference Centerbody station (c.g.) 172.92 ft 2.594 ftwater line (above i,) 1.25 ft .01875 ft

Base Area (Ab)

high wing 1076 ft 2 0.24Z ft

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65

LOCKHEED-HUNTSVILLE RESEARCH & ENGINEERING CENTER


Fig . 14b - Instal lat ion Photograph, Flyback Booster Base F laps (CAL-T-18-063)

66



Fig . 14c - Instal lat ion Photograph, Flyback Booster Base Vent (CAL-T-18-063)

67



2.5 STATIC STABILITY AND TRIM CHARACTERISTICS OF A 0.00227-SCALE PARAMETRIC PRESSURE-FED BOOSTER (TWT 526)

2.5.1 Test Purpose

Studies for a space shuttle launch configuration suggested that a largepressure-fed, water recoverable booster offered possible advantages. Todetermine the static aerodynamic characteristics of such a configuration aparametric test was conducted. Model parameters included nosecone angle,body length, body flare angle and fin size. Tunnel parameters included Machnumbers of 1.96, 2.74 and 4.96, angles of attack from 50 to 90 degrees, andangles of sideslip from -10 to 10 degrees while at an angle of attack of 60 de-grees. A pretest report for this study was published in January 1972 (Ref. 6).

2.5.2 Test Facility

The test was conducted in the MSFC 14 x 14-Inch Trisonic Wind Tunnel.For a description of this facility see Section 2.1.2.


The model included three nose cones, three body lengths, three baseflare angles and three sets of fins. A maximum of 136 combinations can becreated using these model parts. Figure 15 shows the dimensions of thevarious model parts. All model parts were made of aluminum. Figure 16

is a photograph of the associated model parts. Model nomenclature for the

test was:

Symbol Definition

C1 short cylindrical body (I = 2.457 in)

C2Z medium cylindrical body (A = 3.357 in)

C3 long cylindrical body (Q = 4.257 in)

68



F0 straight flare extension (Q = 1.35 in)

F1 7.5-degree flare extension (I = 1.35 in)

F2 15-degree flare extension ( = 1.35 in)

F3 20-degree flare extension (I = 1.35 in)

N1 25-degree nose cone (Q = 0.965 in)

N2 33-degree nose cone ( = 0.693 in)

N3 40-degree nose cone (I = 0.536 in)2T1 small tail fin (S = 0.848 in )

T2 medium tail fin (Se = 1.696 in2 )2

T3 large tail fin (Se = 2.536 in )

Figure 17 shows the model as installed in the tunnel.


Model forces and moments were measured by MSFC balances 201 and

2Z7. Two balances were necessary because of the center of pressure shifts

caused by running with fins on or off and varying normal force loads. Balance

201 was used for configurations and Mach numbers when loads were low,

balance 227 was used for conditions when loads were high. In this manner

maximum accuracy was obtained.

After correcting for sting deflections and weight tares, data were reduced

to coefficient form by the use of the reference dimensions shown in Table 10.

The moment reference point varied with configuration but was always at 60%

of the body length. Actual distance of the moment reference point from the

nose for each configuration is shown in Table 10.

Two methods were used for computing base axial force. The first

method used a base pressure and a cavity pressure, while the second used

only a base pressure. Areas used for base axial force computation are shown

in Table 11.

69



The final coefficient data were entered into the SADSAC program and

published as a data report (Ref. 26).

2.5.5 Discussion

Data that were obtainedduring this test represent a comprehensive study

of high angle-of-attack supersonic aerodynamics for cone-cylinder-flare-fin

configurations. Analysis of these data would provide design data useful for

future studies of any vehicle of this type.

Table 12 shows the configurations run during the test. The study re-

quired 154 hours of wind tunnel test time to complete.

70



Table 10

0.00227-SCALE PARAMETRIC PRESSURE-FEDREFERENCE DIMENSIONS

BOOSTER

71


Full Scale Model Scale

Reference Area (Sref) 857 ft 2 0.636 in 2

(Cylinder Cross-Sectional Area)

Reference Length (Qref) and Span (bref) 396 ft 0.900 in

(Cylinder Diameter)

Moment Reference Point (MRP)

(0.60 1B Aft of Nose)

N1 Cl 1262 in 2.8650 in

N2 C1 1189 in 2.6994 in

N3 C2 1146 in 2.6022 in

N3 C2 1384 in 3.1416in

NI1 C3 1739 in 3.9468 in

N2 C3 1665 in 3.7806 in

N3 C3 1622 in 3.6828 in


Table 11

0.00227-SCALE PARAMETRIC PRESSURE-FED BOOSTER BASE AREAS

WITH CAVITY PRESSURE

Frustum Full Scale Model Scale

Base Area Cavity Area Base Area Cavity Area2 .2(ft ) (in)

F0 334 523 0.248 0.388

F1 1080 523 0.801 0.388

F2 2121 523 1.574 0.388

F3 3023 523 2.243 0.388

WITHOUT CAVITY PRESSURE

72

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LMSC-HREC TR D30646

2.6 REENTRY STUDY OF A 0.00513-SCALE SOLID ROCKET BOOSTER(TWT 541)

2.6.1 Test Purpose

The current space shuttle concept calls for the solid rocket boosters to

be recovered for reuse. The purpose of this test was to give preliminary in-

formation on the booster aerodynamic characteristics so that recovery tech-

niques might be devised.

Upon separation the booster follows a ballistic trajectory. Since separa-

tion will be made with no particular regard to booster attitude, the booster may

well be tumbling upon reentry. To measure the effect of all possible reentry

attitudes, the model was tested at angles of attack from -10 to 190 degrees. A

range of Mach numbers from .6 to 4.96 was covered for the test. The nozzle

was tested in three different roll positions. A pretest report for this study was

published in April 1972 (Ref. 7).

2.6.2 Test Facility


For a description of this tunnel see Section 2.1.2.


The model is a 0.00513-scale version of a 156-inch solid rocket booster.

Figure 18 shows general arrangements of the model. To test the model for the

wide range of angles of attack, special mounting arrangements had to be pro-

vided. Figure 19 shows the mounting systems used to provide the full angle

of attack range. Figure 20 shows the model installed in the tunnel.

The model is constructed entirely of stainless steel. Model design was

done by MSFC. Model nomenclature for the test was:

84


LMSC-.HREC TR D306346

Symbol Definition

B 1 body

N1 full nose (a = -10 to 130 deg)

N2 cut nose (a = 130 to 190 deg)

E1 cut rocket nozzle (a = -10 to 50 deg

S1 one caliber strake with 0-degdihedralangle, located at 0-deg with respect tothe side of the body

S2A two caliber strake with 15-deg dihedralangle, located at 0-deg with respect tothe side of the body

S2B identical to SZA except has 30-degdihedral

S3B same as S2A except has 0-deg dihedral


Model forces and moments were measured using MSFC balance 201.

All forces and moments were reduced to coefficient form by use of the refer-

ence dimensions shown in Table 13. Data were corrected for sting deflections

and weight tares.



2.6.5 Discussion

Configurations tested are shown in Table 14. The test was straight-

forward with few problem areas. Data discrepancies existed in the form of

shifts when the straight sting mounting system (a = -10 to 50 ° and 130 to 190 ° )

was changed to the dogleg mounting system (a = 50 to 90 ° and 90 to 130 ° ) and

when the nose and nozzle were interchanged on the dogleg mount. The dis-

crepancies are believed to be caused by the cutouts in the model to accomodate

85



the mounting systems in the case of the pitch plane data and due to model-

balance misalignment or a possible tunnel crossflow angularity problem in the

case of yaw plane data. The test required 140 hours of wind tunnel occupancy

time to complete.

86



Table 13

0.00513-SCALE 156-INCH SRM REFERENCE DIMENSIONS

156 in.

156 in.

645 in.

645 in.

869 in.

869 in.

0.8 in.

0.8 in.

3.213 ir.

3.213 in.

4.213 in.

3.213 in.

Parameter Full Scale Model ScaleII

0.503 in 219104 in 2Reference Area (S re f )

Reference Length (fref)

Reference Span (bref)

Balance Location(from body base)

a = - 1 0 ° to 50°

a= 50 ° to 90 0

a= 90 ° to 130 °

a = 1300 to 1900

Moment Reference Center(from body base)

a = - 10° to 50 °

0 0a = 50 ° to 90

a= 90 ° to 130 °

a= 1300 to 1900

Base Area (Ab)

Nozzle Area (A 1 )

(A 2 )

3.310 in.

3.310 in.

4.456 in.

4.456 in.

20.503 in

.052 in 2

.173 in 2

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2.7 STATIC STABILITY AND CONTROL STUDY OF A 0.004-SCALEPARAMETRIC LAUNCH VEHICLE (TWT 544 AND TWT 544X)

2.7.1 Test Purpose

The current configuration of the space shuttle is a parallel burn vehicle

with large SRMs arranged around a large liquid hydrogen-oxygen tank. The

purpose of these tests was to gain data to optimize the configuration. Param-

eters varied during the test included HO tank diameter, HO nose shape, SRM

radial and longitudinal location, SRM diameter, SRM length, SRM nose shape,

orbiter-to-HO tank incidence, and HO tank fins. Orbiter control surface

effectiveness while in the launch configuration was also investigated. Tunnel

parameters included a Mach number range of 0.6 to 4.96, angles of attack

from -10 to 10 degrees, and angles of sideslip from -6 to 6 degrees while at

0 degrees angle of attack. A pretest report was published in May 1972 (Ref. 11).

2.7.2 Test Facility


For a description of this facility see Section 2.1.2.


The orbiter is the same configuration as described in 2.2.3. Figures 21, 22

and 23 show the HO tanks and fins and SRMs used to make up the launch configuration.

Figure 24 shows the launch configuration. Figure 25 shows installation photo-

graphs of the launch configuration. The model nomenclature for this test is:

Symbol Definition

F HO tank vertical fin (S = 1.21 in )

O1 baseline orbiter (including the abortsolid rocket motors)

02 baseline orbiter less abort solid rocketmotors

97



03 baseline orbiter without wings

T 1 346-in. diam. HO tank with 22-degnosecone (baseline)

T2 346-in. diam. HO tank with 22-degnosecone-and retro rocket

T3 346-in. diam. HO tank with 17-degnosecone

T4 346-in. diam. HO tank with 22-degnosecone and one body diam. lengthextension


T6 400-in. diam. HO tank with 22-degnosecone and retro rocket

T7 same as T1 without structural rings

T8 346-in. diam. HO tank with generalizednosecone

T9 312-in. diam. HO tank with 17-deg nosecone

T10 346-in. diam. HO tank with modifiedApollo nosecone

S1 156-in. diam. solid rocket motor with17-deg nosecone (baseline)

S2 156-in. diam. solid rocket motor with17-deg nosecone and one body diameterlength extension

S3 156-in. diam. solid rocket motor withskewed nose tangent to HO tank

S4 156-in. diam. solid rocket motor withskewed nose turned 180 deg relativeto S3 position

S5 178-in. diam. solid rocket motor with17-deg nosecone

S6 156-in. diam. solid rocket motor with17-deg nosecone less rocket nozzle

All model parts are made of aluminum.

98




Model forces and moments were measured by MSFC balance 232. After

corrections were made for sting deflections and weight tares, the forces and

moments were reduced to coefficient form by the reference dimensions shown

in Table 15.

The data were entered in the Chrysler Corporation System for Analysis

and Development of Static Aerothermodynamic Criteria (SADSAC) program,

and published as two data reports (Ref. 28 for TWT 544 and Ref. 29 for TWT

544X).

2.7.5 Discussion

An extremely large number of configurations were investigated during

these tests. Tables 16 and 17 show the variety of these configurations. The

first test (TWT 544) suggested configurations that were later investigated

during the second test (TWT 544X).

One hundred thirty-one hours of wind tunnel occupancy time were re-

quired for TWT 544 and one hundred sixteen hours were required for TWT 544X.

99



Table 15

0.004-SCALE SPACE SHUTTLE LAUNCH CONFIGURATIONREFERENCE DIMENSIONS


Reference Area (Sref)

(Wing Theoretical Area)

Reference Length (ref)(M.A.C.)

Reference Span (bref)(Wing Span)

Balance Location (BMC)(Balance Moment Center)

(from base of HO tank)

Moment Reference Point(MRP)

Base Area (Ab)

Q1

S1, S2, S3, S4, S6

S5

T1, T2, T3, T4, T

T5, T6

T9

'7, T8, TlO

3420.0 ft2

507.0 in.

1115.0 in.

840.0 in.

318

133

173

653

872

530

ft 2

ft 2

ft 2

ft

ft 2

ft 2

100


7.880 in 2

2.028 in.

4.460 in.

3.293 in.

3.360 in.

0.732

0.306

0.398

1.504

2.010

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2.8 ORBITER PRESSURE DISTRIBUTION OF A 0.004-SCALE ORBITERWHILE MOUNTED IN THE LAUNCH CONFIGURATION (TWT 550)

2.8.1 Test Purpose

Mounting of the orbiter in proximity to the external HO tank and the solid

rocket motors imposes interference loads upon the orbiter. To determine the

localized loads on the panels of the orbiter wing, static pressure taps were

located on one wing panel on both upper and lower surfaces and down the lower

surface centerline of the body. The orbiter was then tested in the launch con-

figuration to determine local panel loads.

The test Mach number range was from 0.6 to 4.96, the angle of attack

range was -8 to 8 degrees and the angle of sideslip was -6 to 6 degrees. A

pretest report for this test was published in July 1972 (Ref. 30).

2.8.2 Test Facility

The test was conducted in the MSFC 14 x 14-Inch Trisonic Wind Tunnel

This facility is described in Section 2.1.2.


The HO tank and the solid rocket motors for the launch configuration

are the same as those described in Section 2.7.3. The orbiter is the same

configuration as the one described in Section 2.2.3 except that it lacks the

wingtip pods of the attitude control propulsion system. The orbiter is physic-

ally different in that the wing is made of aluminum and has static pressure

taps on the upper and lower surface of the left-hand panel. A row of taps is

also located on the lower surface longitudinal centerline. Table 18 and Fig. 26

show the location of these static pressure taps. The body of the orbiter is made

of Stycast and the vertical fins are made of aluminum.

121



Figure 27 shows the model installed in the tunnel. The model nomen-

clature for the test is:

Symbol Definition

01 baseline orbiter (including the abortsolid rocket motors)

02 baseline orbiter less abort solidrocket motors





Static pressure data were recorded by the use of Scanivalves containing

5 psid transducers for Mach numbers of 0.6 to 1.96 and 50 psid transducers

for Mach numbers of 2.74 and 4.96. The data were referenced to tunnel static

pressure and reduced as nondimensional pressure coefficients.

The data were entered into the Chrysler SADSAC program and will soon

be published as a data report (Ref. 31).

2.8.5 Discussion

This test was a follow-on of an earlier orbiter pressure test of the

same configuration. Orbiter-to-tank incidence angle, SRM radial location,

and a new HO tank nose were evaluated during this test. Table 19 shows

a listing of all configurations tested.

Sixty-nine hours of wind tunnel occupancy time were required to complete

the study.

122



Table 18

ORBITER SURFACE PRESSURE TAP LOCATIONS

Tap Number X/Llocal 1 Y/0.5 brefLlocal 0.5 bref(inref .local (inref.____________j (in.) { in.)

1 .230 .844 .944 2.23

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3 .243 .564 1.623

4 .497 - -

5 .750

6 .134 .294 3.225

7 .292 -

8 .466

9 .647

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11 .044 0 5.275

12 .091 - - -

13 .148 - - -

14 .225 - - -

15 .304 - - -

16 .455 -

17 .944 - -

18 .135 .286 3.225

19 .293 - -

20 .469 -

21 .646 -

22 .823 -

23 .223 .555 1.623

24 .498 - -

25 .749 - -

26 .246 .846 .944

27 .768 - -

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2.9 PRESSURE DISTRIBUTION ON 0.004-SCALE HO TANK AND SRMSWHILE MOUNTED IN THE LAUNCH CONFIGURATION (TWT 543)

2.9.1 Test Purpose

Just as local loaddistributions were needed for the orbiter, local load

distributions were also needed for the HO tank and the SRMs. To accom-

plish this, static pressure taps were located in the HO tank and the SRMs. Test

parameters for this study included Mach numbers from 0.6 to 1.96, angles of

attack from -10 to 10 degrees and angles of sideslip from -10 to 10 degrees.

A pretest report was published in June 1972 for this study (Ref. 12).

2.9.2 Test Facility

The test took place in the MSFC 14 x 14-Inch Trisonic Wind Tunnel.



The orbiter used for this test is the same as described in Section 2.2.3.

The HO tank and SRMs have the same general dimensions as those described

in Section 2.7.3. However, to provide for surface static pressure locations,

new HO tank and SRMs were constructed of Stycast in which the static pres-

sure orifices were molded. Figures 28 through 30 show the general arrange-

ment of the HO tank and the SRMs and the location of the pressure taps. The

sting was permanently molded into the tank, as shown in Fig. 31. The HO tank

had 152 pressure orifices and each SRM had 38 orifices for a total of 230 orifices.

Figure 32 shows the model as installed in the wind tunnel

129



The model nomenclature for this test was:

Symbol Definition

01 baseline orbiter (including the abortsolid rocket motors)

T1 346-in. diam. HO tank with 22-degnosecone (baseline)



The local surface static pressures were measured by 50 psid transducers

contained in Scanivalves. The static pressures were referenced to tunnel free-

stream static pressure and presented in the form of nondimensional pressure

coefficients.

The data are currently being incorporated in the Chrysler SADSAC pro-

gram and will be published as a data report (Ref. 32).

2.9.5 Discussion

The test was run to give a preliminary look at local loads on the HO tank

and SRMs while in the launch configuration. The small model scale (0.004)

prevented locating as many pressure locations as is usually felt to be desirable

on the bodies. This problem was alleviated somewhat by providing three different

mounting positions, each spaced 15 degrees radially apart for each SRM. By

rotating the SRMs through four positions and then combining the data from

both SRMs, pressure distributions at 15-degree intervals could be determined

for the entire diameter of an SRM.

Pressure distributions obtained were sufficient for determining load

distributions, but more pressure taps would have been necessary to deter-

mine local flow singularities such as shock impingement.

The test program is outlined in Table 20. The test program required

132 hours to complete.

130


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2.10 MUTUAL INTERFERENCE STUDY OF THE ORBITER, HO TANK ANDSOLID ROCKET BOOSTER OF A 0.004-SCALE SPACE SHUTTLELAUNCH CONFIGURATION

2.10.1 Test Purpose

To optimize the space shuttle launch configuration and to size the attach-

ment points of the various components it was necessary to know the individual

loads of each component. To determine these loads, the North American Space

Shuttle launch configuration was tested on the MSFC Dual-Sting Mounting System.

Model parameters included orbiter-to-HO tank incidence angle, orbiter-to-HO

tank separation distance, SRB longitudinal position, SRBs attached to the HO

tank (metric) and SRBs attached to the lower sting (nonmetric). Test param-

eters included a Mach number range from 0.6 to 4.96, an angle of attack range

from -5 to 10 degrees, and an angle of sideslip range from -6 to 6 degrees

at zero degrees angle of attack. A pretest report for this test was published in

September 1972 (Ref. 13) and an addendum was published in October 1972 (Ref.14).

2.10.2 Test Facility




The orbiter was a 0.004-scale model of the North American Space Shuttle

Orbiter. Figure 33 shows a general arrangement of the orbiter. All model

parts are made of stainless steel except for the wing and orbital maneuvering

system pods which are made of aluminum. The model was designed and con-

structed at Lockheed-Huntsville.

The 0.004-scale HO tank and booster are shown in Figs. 34 and 35.

The tank body, SRB aft body and SRB fins are made of stainless steel. The

SRB nozzles are made of brass. All other parts are made of aluminum.

The tank and SRBs were designed by MSFC. The model nomenclature for

this test was:

140



Symbol Definition

O1 baseline orbiter less abort solidrocket booster

T3 318-in. diam. tank with ogivenosecone

S1 156-in. diam. solid rocket motorwith 17-deg nosecone and stabilizingfins.

For testing, the orbiter was mounted on the top sting of the MSFC Dual

Sting Mounting System. The tank was mounted on the lower sting. Figure 36

shows the mounting system. The SRBs were attached to the tank (metric) or

supported by separate stings attached to the lower sting (nonmetric). The

SRB nonmetric sting mount is shown in Fig. 37.

Figure 38 shows all components in proximity as for the test. Figure 39a

is an installation photograph showing the left SRB mounted in the metric posi-

tion. Figure 39b is an installation photograph showing the right SRB in the

nonmetric position.


Orbiter forces and moments were measured by MSFC balance 231 and

tank forces and moments were measured by MSFC balance 232. The data

were corrected for weight tares and reduced to nondimensional coefficients

by the use of the reference dimensions shown in Table 21. Initially all data

were reduced about the moment reference points of the respective bodies as

shown in Table 21.

Since the deflection constants and loads are different for the two bodies,

angular and linear displacements are different for the two bodies when pitched

through an angle-of-attack sweep. To make consistent comparisons, the bodies

must remain in the same relation to each other regardless of configuration.

141



To nominalize the relative positions of the orbiter and tank, a grid of three

incidence angles and two separation distances was tested. After initial data

reduction, a Northrop interpolation program (Ref. 33) was used to produce

nominalized data at the proper incidence angles and separation distances.

The program also transferred orbiter moments to the moment reference

point of the tank.

These data are currently being entered into the Chrysler SADSAC pro-

gram and will be published as a data report (Ref. 34).

2.10.5 Discussion

Data from this test can be used to determine the individual loads of the

launch vehicle components. Although the SRBs were not fitted with balances,

comparison of the metric and nonmetric SRB data for the tank balance allows

an SRB load increment to be computed. Comparison of the load data of a body

while in proximity with data obtained for the body while being tested alone will

allow computation of the interference loads.

Table 22 is a listing of the configurations tested. The test required 223

hours of wind tunnel occupancy time to complete.

142



Table 21

/NORTH AMERICAN LAUNCH CONFIGURATION REFERENCE DIMENSIONS

143



Reference Area (Sref) 3220 ft 2 7.419 in 2

(Orbiter Wing Area)

Reference Length (2refrReference Length (If 1328.0 in. 5.312 in.

(Orbiter Body Length)

Reference Span (bref)1328.0 in. 5.312 in.

(Orbiter Body Length)

Balance Location

Orbiter (aft of nose) 3.719 in.

HO-Tank (forward of base) 3.113 in.

Moment Reference Point

Orbiter

XMRP (aft of nose) 863.2 in. 3.453 in.

HO Tank

XMRP (forward of base) 1205 in. 4.820 in.

Base Area (Ab)

Orbiter 382 ft 2 0.878 in 2

SRB (one) 132.8 ft 2 .306 in 2

HO Tank 553 ft 2 1.271 in

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Section 3

REFERENCES

1. Bradley, Dale, "Interim Results of Parametric Studies of Space ShuttleBooster and Orbiter Concepts," LMSC-HREC D225587, Lockheed Missiles& Space Company, Huntsville, Ala., January 1972.

2. Bradley, Dale, "Pretest Report for a 0.015-Scale Parametric SpaceShuttle Booster Model in the NSRDC 7x10-Foot Transonic Wind Tunnel,"LMSC-HREC D162856, Lockheed Missiles & Space Company, Huntsville,Ala., February 1971.

3. Bradley, Dale, "Pretest Report for Cruise Engine Interference Effects ona 0.015-Scale Space Shuttle Booster Model in the NSRDC 7x10-Foot Tran-sonic Wind Tunnel," LMSC-HREC D162968, Lockheed Missiles & SpaceCompany, Huntsville, Ala., April 1971.

4. Buchholz, Robert E., "Pretest Report for Directional Stability, LateralStability, and Cruise Engine Placement Studies of a 0.015-Scale SpaceShuttle Booster Model in the NSRDC 7x10-Foot Transonic Wind Tunnel,"LMSC-HREC D225297, Lockheed Missiles & Space Company, Huntsville,Ala., August 1971.

5. Bradley, Dale, "Pretest Report for Minimization of Base Drag on a 0.015-Scale Space Shuttle Booster Model in the CAL 8x8-Foot Transonic WindTunnel," LMSC-HREC D225393, Lockheed Missiles & Space Company,Huntsville, Ala., September 1971.

6. Bradley, Dale, "Pretest Report for Tests of 0.00227-Scale ParametricSpace Shuttle Pressure-Fed Booster Models in the MSFC 14x14-InchSupersonic Wind Tunnel," LMSC-HREC D225623, Lockheed Missiles &Space Company, Huntsville, Ala., January 1972.

7. Buchholz, Robert E., "Pretest Report for a Force Test of a 0.00513-Scale 156-Inch Solid Rocket Motor at Angles of Attack from -10 to 190Degrees in the NASA-MSFC 14x14-Inch Trisonic Wind Tunnel," LMSC-HREC D225833, Lockheed Missiles & Space Company, Huntsville, Ala.,April 1972.

8. Ellis, Roger R., "Pretest Report for a 0.00325-Scale Parametric SpaceShuttle Orbiter Model in the NASA-MSFC 14x14-Inch Trisonic WindTunnel, LMSC-HREC D225042, Lockheed Missiles & Space Company,Huntsville, Ala., April 1971.

155



9. Ellis, Roger R., "Pretest Report for a Force Test of a 0.004-Scale SpaceShuttle Orbiter Configuration in the MSFC 14xl4-Inch Trisonic WindTunnel," LMSC-HREC D225857, Lockheed Missiles & Space Company,Huntsville, Ala., May 1972.

10. Gamble, M. D., "Pretest Report for a Wing Panel Loads Study of a 0.004-Scale Space Shuttle Orbiter When Mounted as Part of the Launch Vehicle,"LMSC-HREC D206148, Lockheed Missiles & Space Company, Inc., Huntsville,Ala., August 1972.

11. Ellis, Roger R., "Pretest Report for a Force Test of a Parametric SpaceShuttle Launch Configuration in the MSFC 14x14-Inch Trisonic Wind Tunnel,"LMSC-HREC D225929, Lockheed Missiles & Space Company, Huntsville,Ala., May 1972.

12. Lott, Robert A., "Pretest Report: Investigation of the Pressure Distri-bution on the HO Tank and SRMS of a 0.004-Scale Model of the Space ShuttleLaunch Configuration," LMSC-HRECD306027, Lockheed Missiles & SpaceCompany, Huntsville, Ala., June 1972.

13. Buchholz, Robert E., "Pretest Report for a Dual Balance Force Test toDetermine the Interference Effects Due to Proximity of the Space ShuttleOrbiter and HO Tank-Solid Rocket Motor Combination (Booster)," LMSC-HREC D306153, Lockheed Missiles & Space Company, Huntsville, Ala.,September 1972.

14. Buchholz, Robert E., "Addendum to Pretest Report for a Dual BalanceForce Test to Determine the Interference Effects Due to Proximity ofthe Space Shuttle Orbiter and HO Tank-Solid Rocket Motor Combination(Booster)," LMSC-HREC D306153-A, Lockheed Missiles & Space Company,Huntsville, Ala., October 1972.

15. Ellis, Roger R.,Robert E. Buchholz, and John W. Warr, III, "Results ofan Analytical Wing-Canard Study for the LOX-RP Booster," LMSC-HRECD225585, Lockheed Missiles & Space Company, Huntsville, Ala., January1972.

16. Ellis, Roger R., "Results of the MSFC 14x14-Inch Trisonic Wind TunnelChoked Flow Investigation During MSFC TWT 498 at Mach 1.1 and 1.3,"LMSC-HREC D225520, Lockheed Missiles & Space Company, Huntsville,Ala., December 1971.

17. Ellis, Roger R., "Results of the Sting Interference Study During MSFCWind Tunnel Test TWT 498/499," LMSC-HREC D306055, LockheedMissiles & Space Company, Huntsville, Ala., July 1972.

18. Simon, Erwin, "The George C. Marshall Space Flight Center's 14x14-InchTrisonic Wind Tunnel Technical Handbook," NASA TMX-64624, George C.Marshall Space Flight Center, Huntsville, Ala., November 1971.

19. "Static Aerodynamic Characteristics and Control Effectiveness of Two DeltaWing Orbiter Configurations at Mach' Numbers from 0.6 to 4.96," CR-120,020,MSFC 14x14-Inch Trisonic Wind Tunnel Test 498, George C. Marshall SpaceFlight Center, Huntsville, Ala., March 1972.

156



20. "Aerodynamic Characteristics of a Double Delta Wing Space ShuttleOrbiter at Mach Numbers from 0.6 to 4.96," CR-120,057, MSFC TrisonicWind Tunnel, Test 542, George C. Marshall Space Flight Center, Huntsville,Ala., August 1972.

21. Thomas, Walter S., "The David Taylor Model Basin 7x10-Foot TransonicWind Tunnel Facility," Aero Report 985, Naval Ship Research and Develop-ment Center, Carderock, Md., July 1970.

22. "Directional and Lateral Stability and Interference Effects of Cruise EngineLocation on a 0.015-Scale Space Shuttle-Booster," CR120,019, NSRDC 7x10-Foor Transonic Wind Tunnel Test 3310, Naval Ship Research and Develop-ment Center, Carderock, Md., May 1972.

23. "8-Foot Transonic Wind Tunnel," WTO-300, Cornell Aeronautical Labora-tory, Inc., Buffalo, N. Y., March 1968.

24. "Experimental Investigations for Base Drag Reduction on a 0.015-ScaleModel MSFC Proposed Space Shuttle Booster at Mach Numbers from 0.4to 1.10," CR120,030, Cornell 8x8-Foot Transonic Wind Tunnel Test CAL-T-18-063, Cornell Aeronautical Laboratory, Inc., Buffalo, N.Y., February1972.

25. "Transonic Wind Tunnel Tests of a 0.015-Scale Space Shuttle BoosterModel," CAL No. AA-4018-W-7, Aerosciences Division 8-Foot TransonicWind Tunnel, Cornell Aeronautical Laboratory, Inc., Buffalo, N. Y.,January 1972.

26. "Aerodynamic Characteristics of Cone-Cylinder-Flare-Fin Configurationat Mach Numbers of 1.96, 2.74 and 4.96 and Angles of Attack from 50 to90 Degrees," CR120,042, MSFC 14x14-Inch Trisonic Wind Tunnel Test526, Marshall Space Flight Center, Huntsville, Ala., June 1972.

27. "Aerodynamic Characteristics of a 156-Inch Solid Rocket Motor at Anglesof Attack from -10 to 190 Degress," CR120,056, MSFC 14x14-Inch TrisonicWind Tunnel Test 541, Marshall Space Flight Center, Huntsville, Ala.,August 1972.

28. "Static Stability and Control Effectiveness of a Parametric Launch Vehicle,"CR120,059, MSFC 14x14-Inch Trisonic Wind Tunnel Test 544, MarshallSpace Flight Center, Huntsville, Ala., July 1972.

29. "Performance, Static Stability and Control Effectiveness of a ParametricSpace Shuttle Launch Vehicle," CR120,074, MSFC 14x14-Inch TrisonicWind Tunnel Test 544X, Marshall Space Flight Center, Huntsville, Ala.,October 1972.

30. Buchholz, Robert E., "Pretest Report for a Pressure Test of a 0.004-ScaleSpace Shuttle Orbiter in the Launch Configuration to be Conducted in theNASA-MSFC 14x14-Inch Trisonic Wind Tunnel," LMSC-HREC D306060,Lockheed Missiles & Space Company, Huntsville, Ala., July 1972.

15'7



31. "Static Surface Pressures of the 0.004-Scale 049 Orbiter in the LaunchConfiguration," CR120,075, MSFC 14x14-Inch Trisonic Wind Tunnel TestTWT-550, George C. Marshall Space Flight Center, Huntsville, Ala.,(to be published).

32. "An Investigation of the Load Distribution over the SRM and HO Tank of a0.004-Scale Space Shuttle 049 Launch Configuration," CR120,058, MSFC14x14-Inch Trisonic Wind Tunnel Test TWT-543, George C. MarshallSpace Flight Center, Huntsville, Ala. (to be published).

33. Cole, Paul, "Double Interpolation Program for Use with MSFC StagingMechanism," M-9241-72-67, Northrop Services, Inc., Huntsville, Ala.,May 1972.

34. "Mutual Aerodynamic Interference Effects Due to Proximity of the ModifiedATP Space-Shuttle Orbiter, External Tank and Solid Rocket Boosters,"CR120,060, MSFC 14x14-Inch Trisonic Wind Tunnel Test TWT-545, GeorgeC. Marshall Space Flight Center, Huntsville, Ala., (to be published).

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