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TM 54/20-320LMSC-HREC D225144

LOCKHEED MISSILES & SPACE COMPANY

HUNTSVILLE RESEARCH & ENGINEERING CENTER

HUNTSVILLE RESEARCH PARK

4800 BRADFORD DRIVE, HUNTSVILLE, ALABAMA

CALIBRATION OF PROPULSIONSIMULATION NOZZLES FOR

SPACE SHUTTLE BOOSTER ANDORBITER MODELS FOR THE

ABORT/SEPARATION STAGINGEXPERIMENTAL PROGRAM

July 1971

Contract NAS8-20082

by

L. Ray Baker, Jr.

APPROVED:R V: _ _John W. Benefield, OSupervisor

Fluid Mechanics Section

George D. Reny, Ma agerAeromechanics Dept.

LMSC-HREC D225144

FOREWORD

This document presents the results of work performed by

Lockheed's Huntsville Research & Engineering Center while

under subcontract to Northrop Nortronics (NSL PO 5-09287) for

the Aero-Astrodynamics Laboratory of Marshall Space Flight

Center, Contract NAS8-20082. This task was performed in re-

sponse to the requirement of Appendix B, Schedule Order 178, at

the request of Mr. V. Keith Henson, S&E-AERO-AAV.

ii

LOCKHEED- HUNTSVILLE RESEARCH & ENGINEERING CENTER

LMSC-HREC D225144

CONTENTS

Section Page

FOREWORD ii

1 INTRODUCTION AND SUMMARY 1

2 CALIBRATION TEST DESCRIPTIONS 4

2.1 Model Nozzles 4

2.2 Test Program 5

2.3 Analytical Study 6

3 EXPERIMENTAL AND ANALYTICAL RESULTS 8

3.1 Booster Nozzle 8

3.2 Orbiter Nozzle 9

3.3 Nozzle Setting Curve 9

4 C ONC LUSIONS 12

REFERENCES 13

iii


LMSC-HREC D225144

Section 1

INTRODUCTION AND SUMMARY

One of the primary areas of concern in current studies of the Space

Shuttle configurations is the phenomenon surrounding the abort/separation

maneuver. During this maneuver, the booster and orbiter vehicles are sub-

jected to a combination of aerodynamic and plume gas dynamic effects, the

magnitude of which are a function of vehicle orientation, trajectory conditions

and vehicle main propulsion system power levels. The complexity of such a

problem makes it virtually impossible to assess, analytically, the combined

effects on the individual vehicles. Therefore, an experimental program was

conceived to determine the power-on aerodynamic flight characteristics

of the booster and orbiter vehicles during the abort separation maneuver.

An essential contribution to such a test program is the gasdynamic sim-

ulation of the size and shape of the plumes emitting from the full-scale orbiter

and booster main propulsion systems. The two major effects of the plume

which must be simulated are: the interaction of the plume with the external

flow field, and the direct impingement of the plumes on surfaces which are

enveloped by the plume. The model rocket exhaust plume will produce the

correct effects on the external flow if the model plume is the correct size and

shape and is in the correct location relative to the model configuration. Correct

scaling of the direct impingement of the plumes on surfaces will be obtained,

if in addition to the above, the model exhaust plume matches the momentum

flux per unit area of the full-scale exhaust plume.

To accomplish the objective of correct plume simulation, the similarity

parameters developed by Herron (Ref. 1) and the requirement to match momen-

tum flux per unit area between the model and the full-scale systems, were

applied to the orbiter and booster propulsion systems by Sims (Ref. 2). Wind

tunnel facility requirements dictated the use of air or dry nitrogen ( y = 1.4)


LMSC-HREC D225144

as the working gas (for the model nozzles) and the use of a model sting support

passing through the model base. The necessity of using a gas with a Y different

from the full-scale system resulted in a mismatch in the momentum flux when

the plume size and shape were correctly simulated. Supporting the model with

a base sting system necessitated a large departure from a true scale model

base for both orbiter and booster. A single annular (plug) variable area ratio

conical nozzle placed concentrically around the support sting was selected as

a practical means to achieve plume simulation within the contraints of the test.

The similarity parameters used in the analysis were evaluated at the plume

boundary. Therefore, the use of conical nozzles instead of scaled space

shuttle contoured nozzles should not affect the validity of similarity parameters.

In addition, it was assumed that the similarity parameters apply to plumes

with boundary conditions other than quiescent air (a Newtonian pressure dis-

tribution on the plume boundary,for example. The detailed results of the simu-

lation analysis are presented in Ref. 2.

Two plug nozzle systems have been designed and fabricated to simulate

the multiple nozzle rocket exhaust plume emitting from the booster and orbiter

main propulsion systems during the abort/separation maneuver. The variable

area ratio capability was incorporated into both nozzle systems to permit the

gasdynamic simulation of the full-scale rocket exhaust plume at the various

trajectory conditions of interest.

Calibration testing of the booster and orbiter plug nozzles was accom-

plished in Tunnel C of the Arnold Engineering Development Center's Von

Karman Gas Dynamics Facility (Ref. 3). The nozzles were tested individually

at a series of area ratio settings. Nozzle operating conditions (chamber pres-

sure, Po, and chamber temperature, To ) were maintained in ranges com-

patible with the abort staging test conditions. A quiescent low back pressure,

Pb' condition was maintained in the test cell. Data recorded at each area

ratio setting included: optical data to determine plume shapes; static pressure

measurements on the sting surface at the nozzle exit; nozzle mass flow mea-

surements; and pitot pressure surveys in the plume at several axial locations

downstream of the nozzle exit plane.

2


LMSC-HREC D225144

In a parallel effort, analytical solutions of the nozzle flow field and

associated plume were generated for various area ratio settings of the booster

and orbiter models. A method-of-characteristics solution employing real gas

thermodynamic data for air was utilized in the calculations. Analytical re-

sults for each area ratio setting included: plume shape; static pressure distri-

bution along sting surface; and plots of constant Mach number and constant

pitot pressure contours in the plume flow field. These results formed a base-

line for evaluating the experimentally measured performance of the plug nozzles.

This document presents the results of a detailed evaluation of the nozzle

calibration study. A description of the test hardware is presented, and the

calibration procedures are discussed. Analytical and experimental results

are presented and compared.

3


LMSC-HREC D225144

Section 2

CALIBRATION TEST DESCRIPTIONS

The technical aspects of the calibration tests of the nozzles to be used

in the Space Shuttle abort/separation test are discussed in this section. The

model nozzles are described and the calibration test program outlined. A

discussion of the analytical study conducted in support of the test program

concludes this section.

2.1 MODEL NOZZLES

Two "plug" nozzle systems were designed and fabricated to simulate

the multiple nozzle main propulsion systems of the booster and orbiter vehicles.

Each nozzle is designed so that the model vehicle support sting forms the

"plug" portion of the nozzle (Fig. 1). A contoured region on each sting forms

the inner nozzle wall as shown in Figs. 2 and 3. The outer wall of the nozzles

is formed by a sleeve contoured internally. The contoured regions on the

sting and sleeve form a contoured converging/diverging axisymmetric "plug"

nozzle. The sleeve is moved fore and aft axially to change the effective throat

area and thereby change the nozzle area ratio.

A desired nozzle area ratio is obtained by adjusting the outer sleeve

to give the correct axial dimension, Ep, between the nozzle lip and the sur-

face of a reference block. The reference block location is fixed on the sting

surface by a reference pin (Figs. 4 and 5).

The working gas for the. nozzles is room temperature air. The air is

introduced into the nozzle chambers through supply lines that are integral

to the model support sting. Total pressure of the air, Poj, was varied over

a range of 400 to 1300 psia during the test.

4


LMSC-HREC D225144

Actual nozzle dimensions were established by a post-test inspection

conducted by personnel of the MSFC Manufacturing Engineering Labora-

tory. These dimensions are presented in Fig. 4 for the booster nozzle and

Fig. 5 for the orbiter nozzle. In addition, nominal dimensions for each of the

nozzles are included on each of the figures for comparison purposes. The

location of the static pressure port and the reference pin for setting nozzle

area ratio is also specified on each figure.

2.2 TEST PROGRAM

Calibration testing of the booster and orbiter "plug" nozzles was ac-

complished in Tunnel C of the Arnold Engineering Development Center's

Von Karman Gas Dynamic Facility. The objectives of the nozzle calibration

test were:

* Establish, experimentally, nozzle performance characteristics -for the range of area ratio settings to be used with the boosterand orbiter nozzles, respectively.

* Establish a curve (based on experimental results) of nozzle exitconditions as a function of nozzle area ratio setting.

Each nozzle was tested at a series of area ratio settings. The test

procedure consisted of flowing high pressure ambient temperature air

through the nozzle. Data recording was then initiated after continuous,

steady flow was established in the nozzle. Nozzle operating conditions (chamber

pressure, PO., and chamber temperature, To.) were maintained in ranges

compatible with the abort staging test conditions. Tunnel C was operated as

an evacuated test cell. A quiescent, low pressure, Pb, condition was main-

tained in the tunnel during the test program. A summary of test conditions

for the booster and orbiter nozzles is presented in Table 1.

The data recorded at each area ratio setting included:

* Holographic interferograms for determining initial plume boundaryangle and plume shape

5


LMSC-HREC D225144

* Static pressure measurements on the sting surface at thenozzle exit

* Mass flow rate in nozzle supply line using venturi measuringsystem

* Pitot pressure surveys in the plume at specified axial loca-tions downstream of the nozzle exit plane

* Nozzle total conditions, Poj and To

* Test cell ambient pressure, Pb.

2.3 ANALYTICAL STUDY

An analytical study of the booster and orbiter "plug" nozzle configurations

was conducted parallel with preparations for calibration test program. The

objectives of this study were to: (1) establish a baseline for evaluating the

experimental results; (2) provide preliminary information on nozzle performance

characteristics such as thrust and mass flow rate that might affect hardware

design; and (3) provide a rapid means for determining nozzle performance

characteristics (if reasonable agreement is obtained between analytical and

experimental results).

The initial analytical study was conducted using nominal dimensions

(see Figs. 4 and 5) for the booster and orbiter nozzles. Prior to calculating

nozzle performance, it was necessary to determine the variations of nozzle

area ratio with nozzle setting. This was accomplished by calculating the

minimum geometric throat area for each nozzle setting using an automated

technique. The variation of area ratio with nozzle setting are presented in

Fig. 6 for the booster nozzle and Fig. 7 for the orbiter nozzle. This informa-

tion was used to determine nozzle settings for both the calibration tests and

the analytical study.

Nozzle performance characteristics and internal and plume flow field

properties were calculated using Lockheed's variable mixture ratio method-

of-characteristics (VOFMOC) program, Ref. 4. The program was modified

6


LMSC-HREC D225144

to internally shift the nozzle walls to the correct relative position for a specified

nozzle setting (i.e., area ratio). Real gas thermodynamic properties of air

were utilized in the flowfield calculations. The gas properties were calculated

from equations derived from the Beattie-Bridgeman equation of state assuming

variable specific heats, Ref. 5. Flowfield calculations were initiated from

the previously calculated location of the geometric throat using a constant Mach

number start line. Results of these calculations are presented in Figs. 8

through 11.

Following the calibration test program, the analytical calculations were

repeated using actual measured dimensions of the booster and orbiter nozzles.

This information is also presented in Figs. 8 through 11.

A detailed discussion of the results of the analytical and experimental

studies is presented in the following section.

7


LMSC-HREC D225144

Section 3

EXPERIMENTAL AND ANALYTICAL RESULTS

The nozzle performance characteristics of the booster and orbiter nozzles

were evaluated on the basis of comparison of analytical and measured values

of mass flow rate, nozzle pressure ratio, nozzle chamber pressure, and plume

shape information. The experimental and analytical results were correlated

on the basis of nozzle setting, E . Nozzle setting is defined as the measuredp

gap between the nozzle exit lip and the surface of a reference block that is

mounted on the sting during the setting operation. Actual setting of a specific

nozzle gap dimension was accomplished with the use of standard gage blocks

("Jo" blocks) to within +0.0005 in. Figures 6 and 7 show that nozzle area

ratio is extremely sensitive to the setting gap, especially for the orbiter.

Also it is evident from these figures that nozzle dimensions significantly af-

fect the distribution of area ratio with nozzle setting. The degree of success

in obtaining agreement between experimental and analytical results was hinged

directly on actual knowledge of the nozzle dimensions.

3.1 BOOSTER NOZZLE

Performance characteristics of the booster nozzle are presented in

Figs. 8 and 10. The experimental and analytical mass flow rate values shown

in Fig. 8 have been adjusted to a common chamber pressure, Poj, base of

1000 psia for the purpose of comparison. Nozzle pressure ratio data were

determined from the ratio of the static pressure, pj, on the sting surface at

the nozzle exit plane to the nozzle chamber pressure.

In general, both the mass flow and pressure ratio data suggest that the

booster nozzle operates at a lower area ratio experimentally for a specific

nozzle setting than it does analytically. The analytical results obtained with

the measured nozzle dimensions did, however, shift closer to the experimental

8


LMSC-HREC D225144

results. These calculations were run with the same gasdynamic chamber

conditions (Po0 . To ) that were measured experimentally. A comprehensive

review of both sets of data revealed no gasdynamic explanation for the differences

in the test and analytical results.

Because of the difficulty in measuring the nozzle contours, it is suspected

that the actual nozzle dimensions are still not known. In addition, the inspec-

tion results showed that the reference block dimensions and the distance

between reference pin location and nozzle lip at the "reference" condition

vary from the normn enough to contribute to a lack in certainty of the knowledge

of the variation of nozzle area ratio with actual gap setting.

3.2 ORBITER NOZZLE

Performance characteristics of the orbiter nozzle are presented in Figs. 9

and 11. These data were prepared in the same manner as they were for the

booster nozzle data.

Comparison of the analytical and experimental results for the orbiter

nozzle revealed a different trend than was encountered with the booster nozzle.

The initial analytical results (based on nominal nozzle dimensions) fell well

below the experimental results for both mass flow rate and nozzle pressure

ratio. Nozzle performance calculations made using measured nozzle dimen-

sions and calibration test gasdynamic operating conditions were above the

experimental data curves but followed, in general, the contour of the experi-

mental data curve. These results indicate again the dependency of the analyti-

cal calculation of nozzle performance on quality of the knowledge of the nozzle

dimensions. As with the booster, no gasdynamic explanation of the differences

in the experimental analytical results could be determined.

3.3 NOZZLE SETTING CURVE

The disagreement between the experimental and analytical results negated

use of the analytical results to obtain nozzle setting for the abort staging test.

9


LMSC-HREC D225144

Therefore, experimental data were used exclusively to determine the required

setting curves of nozzle lip Mach number as a function of the nozzle setting.

The basic equations used to determine the desired data can be found in most

gasdynamic texts.

Initially, the Mach number of the limiting Prandtl-Meyer fan is calcu-

lated from the following relation:

72-1

M 2 P

1/2- 1.0] 2

Y2-l(1)

P to.

where: -' is ratio of the ambient back pressure, Pb, to the chamber pres-Pb

sure, P . Using this Mach number and the nozzle half angle, 8 Ns the re-J

sulting turning angle of the flow, V2

is calculated from the following equation:

V2 = ( 2 -l )arctan [1/2

arctan f72+1 (M2 - ) - arctan((M2 - 1)1/2)((

To complete the calculations, the initial plume expansion angle, 6.j, was

determined from the experimental optical data. Examples of these data,

holographic interferograms, are shown in Figs. 12 and 13 for the booster and

orbiter, respectively. Values of 6. were obtained by measuring the initialJ

plume angle directly from the photographs. This information was then used

to obtain one expression for the flow turning angle at the nozzle lip

v1 = Vz -6jV1 2 j

A second expression for V1 was then written as,

10


(3)

(2)

LMSC-HREC D225144

V ( I + 1/21 Vrl

arctan 1 1Y (M2 - ][1 '1 +---

- arctan [(M _ - 1)1 (4)

Combining Eqs. (3) and (4), a function was formed from which the Mach number

at the nozzle lip could be calculated using iterative techniques.

F(M1 ) = (Y 1 -l1 arctan{| [Y (M1 ) -······ Y 1+1 IM

- arctan _(M -]1( -(V 2 - 6)

The results of these calculations are presented in Figs. 14 and 15 for the booster

and orbiter, respectively. The curve for the booster was faired through data

that were generally representative of the norm of the nozzle operating condi-

tions. The orbiter data were faired through the points associated with high cham-

ber pressures. This was done to account for the flexing of the nozzle struc-

ture which was noted at these pressures and because the high pressure was

representative of most of the abort staging test conditions. These curves

were utilized with plume scaling information of Ref. 2 to obtain the nozzle

settings shown in Table 2.

11

LOCKHEED - HUNTSVILLE RESEARCH & ENGINEERING CENTER

I

(5)

LMSC-HREC D225144

Section 4

CONCLUSIONS

The following conclusions were reached during the course of the nozzle

calibration study:

* The use of the concept of variable area ratio nozzles for thisapplication is sound and provides a means to satisfy plumegasdynamic simulations over a wide range of conditions.

* The ability to predict analytically the performance of nozzlesof this design and size is associated directly with accurateknowledge of the nozzle dimensions.

* Although, neither the booster nor the orbiter nozzles operatedprecisely as anticipated, the variable area ratio capability per-mitted the simulation criterion to be satisfied.

12


LMSC-HREC D225144

REFERENCES

1. Herron, R. D., "Investigation of Jet Boundary Simulation Parametersfor Underexpanded Jets in a Quiescent Atmosphere, " AEDC TR-68-108,Arnold Air Force Station, Tenn., September 1968.

2. Sims, Joseph L., "Plume Simulation for Space Shuttle Abort StagingAerodynamic Testing, " NASA-MSFC S&E-AERO-AF-70-6, 16 December1970.

3. Test Facilities Handbook, 8th ed., Arnold Engineering DevelopmentCenter, Arnold Air Force Station, Tenn., December 1969.

4. Smith, S.D. and A.W. Ratliff, "User's Manual -Description of theVariable O/F Ratio Method-of-Characteristics Program for Nozzleand Plume Analysis," LMSC-HREC D162220-I, Lockheed Missiles &Space Company, Huntsville, Ala., July 1971.

5. Randall, R. E., "Thermodynamic Properties of Gases: Equations Derivedfrom the Beattie-Bridgeman Equation of State Assuming Variable SpecificHeats," AEDC TR-57-10, Arnold Air Force Station, Tenn., August 1957.

13


LMSC-HREC D225144

Table 1

TEST CONDITIONS FOR SPACE SHUTTLE NOZZLEPLUME CALIBRATION TEST

Inter-Run/Group Configuration Nom. P T E PO fero- Sta.

b o p gram Sta.M (psia) (oF) (in. ) (psia) (in. )

___ __ __ __ __ __ _ __ __ __ __ __ __ __ __ --

- _ __ __ __ _

123456789

1011121313131415161718192021222324252627

2829292930

31

i

Booster

Booster

Orbiter

OrbiterOrbiter

0

0

00

0.36

0.400.400.33

0.340.310.33

0.26

0.260.33

0.330.28

0.28

0.26

0.260.26

100

100

100100100

0.389

0.411

0.426

0.542

0.5420.592

0.452

0.504

0.478

0.512

0.5120.545

0.495

I

505640

500460550550430572

660250450

405330430

1200390485485485485430415

8001200415

I800

12401200800415415800

1200

XX

XXX

X

XXXXX

XXXXXX

XX

XXX

XXXXXXXXX

1.531.533.062.24

1.532.28

1.522.283.03

1.502.280.71

0.71

1.520.710.711.52

0.470.94

0.470.94

0.47

14


L I I I

LMSC-HREC D225144

Table 2

NOZZLE SETTINGS FOR SPACE SHUTTLEABORT SEPARATION TEST PROGRAM

Thrust Level Booster Orbiter

E EBooster Orbiter Mach Altitude p P /Pp(%) (%) No. (ft) (in.) (in.)

0 0 5 153,000 .0 0 2 60,500 - .0 0 2 60,500 . .

50 100 2 60,500 .560 .0115 .502 .005480 0 3 93,700 - - - -

50 100 3 93,700 .519 .0172 .485 .007540 0 5 153,000 - - -

50 100 5 .491 .0218 .471 .010240 100 5 - - .471 .01024

100 100 5 .473 .0251 .471 .0102450 50 5 .491 .0218 .473 .00974

0 0 5 .50 100 5 .491 .0218 .471 .01024

0 0 5 _ . .50 100 5 .491 .0218 .471 .010240 0 2 60,500 - - - -

0 20 20 3 93,700 -

0 5 153,000 - - -

100 2 60,500 .522 .0167 - -

100 3 93,700 .495 .0211 - -

100 5 153,000 .473 .0251 - -

0 2 60,500 - - -

0 3 93,700 .0 5 153,000 - - - -

100 2 60,500 .522 .0167 - -100 l 3 93,700 .495 .0211 - -100 0 5 153,000 .473 .0251 - -

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