Hypersonic Wind Tunnel Testing

AEDC TR-94-6

AD-A284 057

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Hypersonic Wind Tunnel Test Techniques

R. K. Matthews and R. W. Rhudy

Caispan Corporation/AEDC Operations

August 1994

Final Report for Period July 1992 - May 1993

DTIC

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ELECTE

SEP

₀₂

1994

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FApprvedfor

puW blimees; disriulon

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___

QUALITY

INSPECTED 5

0o)

ARNOLD ENGINEERING DEVELOPMENT CENTER

ARNOLD AIR FORCE BASE, TENNESSEE

0)

AIR FORCE MATERIEL COMMAND

NOInczs

When U. S. Government drawings, specifications, or other data are used for any purpose

other than a defite related Government procurement operation, the Govermneit thereby

incurs no responsibility nor any obligation whatsoever, and the fact that the Government

may have formulated, furnished, or in any way supplied the said drawings, specifications,

or other data, is not to be regarded by implication or otherwise, or in any manner licensing

the holder or any other person or corporation, or conveying any rights or permission to

manufacture, use, or sell any patented invention that may in any way be related thereto.

Qualified users may obtain copies of this report from the Defense Technical Information

Center.

References to named commercial products in this report are not to be considered in any

seine as an endorsement of the product by the United States Air Force or the Government.

This report has been reviewed by the Office of Public Affairs (PA) and is releasable to

the National Technical Information Service (NTIS). At NTIS, it will be available to the

general public, including foreign nations.

APPROVAL STATEMENT

This report has been reviewed and approved.

DENNIS N. HUPRICH, Major, USAF

Space and Missile Systems Test Division

Approved for publication:

FOR THE COMMANDER

CONRAD M. RITCHEY, Lt Col, USAF

Space and Missile Systems Test Division

Public reporting burden for this collection o1 information is es~timated tO average 1 hour per resoonse, including the time for reviewing instrfucions, fsearching eml$stilg data sources.

gathering and maint aining the data needed. and Ompl~eting and reviewing the collection of information Send comments regjarding this burden estimate Or any other aspe.ct: of this

collection of infor mation, including suggestions for reducing this burden. 10 Washington Headquarters Services. Directorate for Information Op~eratiOns and Reports. 121IS Jefferson Davis, Hhhwa. Suite 1204t Arlin ton. VA 22202-430•r and to the Office of Man Kement and Iud• et Papermorki Reduction Prolect 10704-0188r Washin tOn.OC 20503

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

Auc ust 1994

Final -- July 1992 -May 1993

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Hypersonic Wind Tunnel Test Techniques

JN

-

0979

6. AUTHOR(S)

Matthews, R. K. and Rhudy, R. W., Calspan Corporation/AEDC

Operations

____________

7 PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION

REPORT NUMBER

Arnold Engineering Development Center/DOF

A D T-4

Air Force Materiel Command

AD

-R9-Arnold Air Force Base, TN 37389-4000

9. SPONSORING/MONITORING AGENCY NAMES(S) AND ADDRESS(ES) 10. SPONSORING/'MONITORING AGENCY REPORT NUMBER

Arnold Engineering Development Center/DOF

Air Force Materiel Command

Arnold Air Force Base, TN 37389-4000

11. SUPPLEMENTARY NOTES

Available in Defense Technical Information Center (DTIC).

1 2a. DISTRIBUTION/AVAILABILITY STATEMENT 1 :•b. DISTRIBUTION CODE

Approved for public release; distribution is unlimited.

13. ABSTRACT (Maxim urn 00 words)

This report describes the procedures used in the continuous flow hypersonic tunnels of

the AEDC for static stability, pressure, heat transfer, materials/structures, boundary-layer

transition, and electromagnetic wave testing. Particular emphasis is placed on

heat-transfer techniques because of the importance of defining the thermal environment of

hypersonic vehicles. An overview of the materials/structures test methodology used in the

development of hypersonic vehicle components is presented. Unfortunately, the

methodology to predict transition has eluded the aerodynamicist for over three decades,

and there are still many unanswered questions. This report briefly touches on the many

parameters that affect transition and provides numerous references for those who are

interested in specializing in this topic. The methodology of using trip spheres is discussed,

and illustrative data are presented. Electromagnetic wave testing represents a relatively

new test technique that involves the union of several d isciplines: aerotherrnodynamics,

electromagnetics, materials/structures, and advanced diagnostics. The essence of this new

technique deals with the transmission and possible distortion of electromagnetic waves

(RF or ;R) as they pass through the bow shock, flow field, and electromagnetic (EM)

window of a missile flying at hypersonic speeds.

_________

14. SUBJECT TERMS 1 S, NUMBER OF PAGES

electromagnetic waves, missile seeker system, hypersonic vehicles,

59

boundary

layer,

boresight error, radomes

16. PRICE CODE

17. SECURITY CLASSIFICATION' 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT

OF REPORT OF THIS PAGE OF ABSTRACTSA

EASRP

T

AEDC-FR-94-6

FOREWORD

The hypersonic regime is the most severe of all flight regimes, and consequently demands smart

utili-zation of ground testing and evaluation, flight testing, ano computation/simulation methodologies. Because

of this challenge, von Karman Institute (VKI) asked the Arnold Engineering Development Center (AEDC)

to develop a comprehensive course to define the "Methodology of Hypersonic Testing." Seven American

scientists and engineers, representing AEDC and the University of Tennessee Space Institute (UTSI),

for-mulated this course from their background of over a century of combined experience in hypersonic testing.

The objective of the course v

i -nt

a comprehensive overview of the methods used in

hyper-sonic testing and evaluation, and

o,,!

Ah

he principles behind those test techniques. Topics covered

include an introduction to hypersonic aer-Aynamics with descriptions of chemical and gas-dynamic

phe-nomena associated with hypersonic fligN.,

-at,-ories and application of various hypersonic ground test

facilities; characterization of facility flow fields; measurement techniques (both intrusive and

non-intru-sive); hypersonic propulsion test principles and facilities; computational techniques and their integration

into test programs; ground-test-to-flight data correlation methods; and test program planning. The Lecture

Series begins at the introductory level and progressively increases in depth, culminating in a focus on

spe-cial test and evaluation issues in hypersonics such as boundary-layer transitie-i, shock interactions,

electro-magnetic wave testing, and propulsion integration test techniqaes.

To obtain a complete set of notes from this course write to:

Lecture Series Secretary

von Karman Institute

Charrissie de Waterloo, 72

B-16409 Rhode-Saint-Genese (Belgium)

The information contained in this report is a subset of the work described above.

Accesion For

NTIS

CRA&I

DTIC TAB

Unannounced

5

Justification ...

By ...

Distribution

I

Availability Codes

i

Avail and I or

Dist

I

Special

AEDC-TR-4-6

CONTENTS

bag

Test and Evaluation M ethods ...

5

M aterials/Structures Testing ...

29

Boundary-Layer Transition ...

41

AEDC-TR-94-6

TEST AND EVALUATION METHODS

by

R. K. MATTHEWS and R. W. RHUDY Senior Staff Engineers

Calspan Corporation/AEDC Operations Arnold Engineering Development Center

ABSTRACT E Heat gage output, mv Test facility selection is generally the first step FA Axial force

toward planning a wind tunnel test, either through FN Normal force availability, simulation, or test technique require- Fy Side Force ments. The operating characteristics of the chosen

facility will immediately restrict the choice of test h Heat-transfer coefficient techniques because of the tunnel operating mode. k Material conductivity Since the time response of the measurement technique

must be compatible with the run time of the facility. M Mach number some techniques are immediately ruled out for certain MI. 2. or x Balance moment

facilities. The aerodynamic and aerothermal test

tech-niques used in conventional hypersonic tunnels are

P

Pressure

generally similar throughout the world, varying only q Dynamics pressure

in the sophistication of the facilitys' instrumentation

and data acquisition systems. This section describes q Heat flux the procedures used in the continuous flow hyper- Re Reynolds number

sonic tunnels of the AEDC for static stability,

pressure, and heat transfer testing. Particular t Time emphasis will be on heat transfer techniques because T Temperature of the importance of defining the thermal

environ-ment of hypersonic vehicles and because the static S Reference area

stability and pressure techniques are very similar to S/R Surface distance-to-nose radius ratio those used in subsonic and supersonic facilities.

W Model weight

NOMENCLATURE x, y, z Coordinates of model CG in balance axis system

Ao Intercept of q versus Tw for heat gage

data a Angle of attack At Slope of h versus Tw for heat gage data / Angle of sideslip b Thin-skin wall thickness Q Density

c Specific heat AT Temperature difference

C, Static stability coefficients: e.g. lift CL, 0 Roll angle drag CD, pitching moment Cm Subscripts

Cp Pressure coefficient F Flight

CSF Heat gage calibration factor i Initial (time = 0)

C (tn) Coax gage calibration factor calculated at L Model length time tn

AEDC-T4"

T Tunnel dictate many of the pretest requirements/activities, r Recovery such as model scale, test techniques, etc. The

following sections cover the relatively standard static w Wall stability test model, as shown in Fig. 2.

0 Free stream

0 Stilling chamber conditions

INTRODUCTION .= :

One of the primary considerations in

k.

evaluating hypersonic facilities is the test CALCUTE: FORCE - • q,, , techniques (or test methods) available in a given NOTES: 10 (ITqS

facility. A thorough understanding of test N _{F - FU610}- FEHICL

techniques is very important in planning a test T - WIND TUNNEL _ . program to address a precisely defined test

objective. This section describes the aero- PAT ~8"1

dynamic and aerothermal test techniques that are in general use. The aerodynamic

metho-dology includes force balance measurements FCOTEE

and surface pressure measurements. Flow-field

CALCU(Ar:

T- q-_ and other aerodynamic techniques are discussed W N

in later sections. Aerothermal methodology WIND TUNNEL MOEL includes thermal mapping, discrete gage Figure 1. Wind tunnel/flight simulation. techniques and gage calibration. In general,

each specific test technique is described in terms NOE- i, ý, Ad i AE ALWAYS IN of (a) principle of operation, (b) apparatus, (c) TlHE IRIAKE AXIS SYSTEM,

data reduction and d) illustrative data. Lem FOIAID MOMENT NEVER IN THE MODEL AXIS.

REFEIENE( POINT

STATIC STABILITY TESTING

MODEL

EC-v

AXIS

I RALAE AXIS

Static stability tests in a hypersonic wind tunnel -are conducted in much the same manner and for the

same reasons as in lower-speedwind tunnels. Typical

tests are conducted to; verify that the performance :AXIS

of a particular design is accurate (parametric studies); JOO(L XIS 4W

verify theoretical codes (CFD); and/or prove that LEFT SIDE VIEW REAU VIEW

proposed modifications to existing flight hardware

will, in fact, improve the performance. In general, Figure 2. Model weight (tare) and center-of-gravity compared to low speed facilities the tunnels and the locations referenced to balance axis. test models are smaller and the test cxwironment

much more severe. These smaller models and the The procedures described are for tests in extreme environment (i.e., low static pressure and

extremely high total pressure and temperature lead conventional, relatively long duration wind tunnels. to special requirements not encountered in lower Impulse type, short-duration testing requires other speed wind tunnels. It is much cheaper and safer to special procedures to compensate for such things as build small-scale vehicles and test in the wind tunnel the inertia forces from model vibrations.

("Test before flight'7) than to build the real thing and

have it fail ("Build it and see if it works"). The Apparatus classical wind tunnel to flight correlation parameters

are shown in Fig. I. Once the need for a static in general, static stability data in a hypersonic stability wind tunnel test is determined and the test wind tunnel are obtained by use of a strain gage objectives clearly established, the choice of what balance, usually mounted internal to the test model. facility to use can be made. This choice naturally will Special circumstances may dictate deviations such as

AEDC-TR-944

extremely small models which require that the

OUTPUTS

OF

GAES

balance be mounted external to the model with the NO. I A use of a windshield, but these are so diverse that they A NJ -1

F

need to be addressed on a case-by-case basis and will - -Ih

I'

-not be covered here. Also, at times, requirements N - NJ OR

12

N " NO. I

arise for force/moment measurements to be made on control surfaces or other individua' parts of a vehicle;

however, the techniques used and

fabrication/cali-bration requirements are nearly identical to those

des-cribed below for the "main" balance.

OUTPUITS aGMES

Strain gage balances are constructed by machin- NO. I AND NO. 2

z,

--ing a thin section (called a flexure) in the balance and M F, + f2

bonding a strain gage to the surface. As loads are F - F, o F? NO.

U. I

applied, either by calibration or through the test it - F, (11 - 11) article, the elongation/compression of the flexure and

thereby the strain gage causes an electromotive force b. Force balance

(EMF) to be produced by the gage. This EMF can Figure 3. Force type and moment type balances. then be amplified and electronically processed in

conjunction with the balance SW M AN YAWNG WN

calibration to calculate the

applied load. Machining several flexures into a single balance at the proper locations and in the

proper plane with respect to the A balance centerline creates the TOP VXIEL WOtC

capability to resolve all six components, i.e., normal, side

and axial force and pitching, oO-MAL FOE AU PITMWNG MOMNIE N

yawing, and rolling moments. AttltttI The flexures and gages can be

arranged to measure two moments which are then resolved

into a force and a moment (e.g. / I--K PONIT FN and Mx) or two forces which . L 0R WATER JAE ATITACHIENT

can be summed to obtain the SIE VIEW

total force and, by using the Figure 4. Moment type internal balance. balance geometry, resolved to a

moment. These two types of balances and their

measurement resolutions are illustrated in Fig. 3. The Balance Calibration hypersonic wind tunnels at AEDC normally use

moment-type balances. A schematic showing the Once a balance has been designed and fabricated, normal arrangement of the flexures and gages for this it must be calibrated to determine the electrical output type balance is given in Fig. 4. As stated earlier, (EMF) from the strain gages for a given applied load. speciality balances measure from I to 6 components There are probably as many different techniques of either force-type or moment-type or a combination; calibrating a balance as there are test facilities. For however, the principle of operation, i.e., flexures and these notes, the procedures used to calibrate the strain gages, is the same. Balances are designed and balances used in the AEDC hypersonic tunnels will fabricated for specific maximum loads. Obviously, be covered. Other techniques are very similar, and the load range of a particular balance is determined the end result differs only by the desired precision by the size of the flexures and the overall strength of the test data.

of the balance.

AEDC-TR-9"

Balances are designed so that the flexures are the calibration loading is applied to the balance. The aligned to measure a single component, i.e., strictly measured balance outputs are reduced to forces and

FN, FA or Fy. However, it is impossible to machine moments using the previously calculated balance the balance and install the strain gages so that they constants in a "Balance Loading Program" (BLP) are perfectly aligned. Therefore, because of this and compared to the known applied loads. Since the misalignment a pure normal force will produce some maximum applied loads are chosen equal to the output from the side force and other gages; expected maximums during the test, this comparison conversely a pure side force will produce some output gives an insight into the test data precision. If the from the normal force and other gages; etc. These comparison 6f applied loads to calculated loads is secondary outputs are called interactions, and must outside established precision bounds for the

be accotnted for with the calibration, particular balance, the balance is recalibrated. The calibration of a balance consists of applying As shown in Fig. 6. a large number of balances loads and combinations of loads (i.e., pure FN, FN are available for use in the AEDC hypersonic wind

+ FA, FN + Fy, etc.) in increments up to the rated tunnels. These balances vary in design and cover a capacity of the balance and at several locations along wide range of load-carrying capabilities. For the high-the entire length of high-the balance. The electrical outputs est quality data, a balance should be selected that has from the strain gages are measured, amplified, and maximum rated load capability close to the maximum loaded into a very complex computer program. This expected test loads. Also shown at the top of Fig. program takes into account such things, in addition 6 is the water jacket which covers the balances when to the applied load and its location, as the weight of used in the hypersonic wind tunnels. Since the output the calibration equipment, balance deflections, bal- of the balance is highly temperature sensitive, it must ance roll angle, etc. After the entire set of calibration be kept cool (near room temperature) during data loads has been applied, the program calculates the taking. This can be accomplished by either getting balance constants to be used to resolve the forces and the data very quickly or by use of the water jacket moments sensed by the b-dance during a wind tunnel which shields the balance from both radiative and test. These constants, along with the other conductive heating.

information shown in Fig. 5 are loaded into the test

data reduction computer. A balance calibration is Model Fabrication

quite complicated and time consuming, typically

requiring 4 to 5 days; however, a single calibration The choice of the facility to be used to satisfy the will normally be used for several tests. As a check, test objective will dictate the size of the model that however, prior to each test a simplified version of can be tested. The model scale must be large enough to maintain the fidelity of the full-scale

configura-ATTITUDE

IAELA, AERO AND FLOW

CONSTANTS FLOW PARAMETERS COEF CONFIGURATION PARAMETERS AE RLO DYN A MI C (O ff qL a IA .

MDLCO _{ETC ET(} P._•

00 ONSTANTfS REDUCTIONET EC

+ COMPARISONS TO THEORY AND/OR PAEVIOUS DATA TUNNEL CONDITIONS

TUNNEL FLOW BIALANCE MEASUREMIENTS

AEDC-TR-94-6

_ __ disadvantages. In

the

point-pause method, data are

Swobtained

by positioning the model at a discrete angle of attack and angle of sideslip, usually waiting for the base pressure to stabilize, and recording the mea-surements electronically into the data reduction com-puter. This is repeated for each model attitude desired, usually in 2- or 3-deg increments over the

-

,•AERODYNAMIK .f

"

GENERAL PURPOSE -COMPONENT BALANKES (24 AVAILALE)

o NOMINAL BALANCE SIZES: LENGTH, IN. S.1 TO 13.8 (13-35 am) AANC REFERIENCE POINT

UDIAETER, IN. 0.6 ro 1.1 (QS-s3 cm) MONEL CENuTR OF GRAVITY ((I)6 N • FIAWE RANIGIES, Uh: 20 TO 1.500

SIM FOR CERANGES. LI: 2010 IO O STAINLESS STEEL MOOEL - AS LIGNT AS POSSI

.AXI L FORCE RANGES, LA: 4t0 300 * IDEALY M L €. 6. IN G M OF BALANC E lM GA O)

"

SPECIAL PURPOSE BALANCES (12 AVAJIAIL)*ELYML(..UCNE FRWIE(QIUSRS

* BALAICE AlSO LOCATED AS CLOSE AS POSSIILE TO THE

• TYS AVAILAILE: M MASS ADDTION (3 TO 6 COMPONENT) AEOYNMC CENTER OF PRSSURE

• MAGNUS FORCE AND MOME1NTS e DATA PMECISiON DEPENDENT ON ACCURATE MEASURENTS OF:

* REFERENCE

AREAS

Figure 6. AEDC tunnels A/B/C balances.

_*

( RKFERIENE _€.6.

_LOCATION

LENGTHR WlITl RESPECTO

TO

BLA

* MODEL COMPONENT WEIGHTS

tion; otherwise, the wind Figure 7. Model-balance arrangement. tunnel data will not

accurately predict the flight ADVANTAGES DISADVANTAGES

performance of the vehicle.

I. GREATER DATA PRECISION _I. 11UCN StOWER THAN CONTINUOUS SWEEP-MORE COSTLY

Model weight must be kept

low so that

it

is only a small

2. BAS PRESSURE MEASUREMENTS AT EVERY 21. ONE 1 DATA POINT PER RUN IN IMPULSE TUNNEL

percentage of the balance MODEL ATTITUDE

full-scale

capability.

3. LESS LIKELY TO HAUE ANOMAL NYSTMEIS 3. MM DSTORTION DUE TO NON-UNIFORUM HEATING

However, because of the EFFECTS

high-temperature environ-

4. USE PITCH AND RO MIECHIANISM TO SET 4. DATA NOT NECESSARILY AT SANME MOEL ALTITUDE

ment, models for hypersonic PREDETERMINED a AND 0 FOR EACH CONFIGURATION force testing are usually S. SIME DATA REDUCTION

fabricated from stainless

Figure 8. Point-pause data taking technique.

steel. After the model is

completed, the reference areas, lengths, weight, and

entire range of angle of attack and/or 5-deg

incre-c.g.

location must be accurately measured for input

ments over the sideslip range. The advantages and

to the data reduction program. These model dimens-

disadvantages of this technriue, are shown in Fig. 8.

ions and weights must be obtained for every

configuration to be tested. The test data pre-

In

the continuous sweep method, high-speed data

cision/imprecision is a direct function oi how

are taken as the model is pitched, rolled, or yawed

accurately these measurements are obtained and how

through an angle range. Angle change rates vary

they replicate the full-scale vehicle contours. The

from a few degrees a second to tens-of-degrees per

model is then assembled with the balance, water

second. The continuous data are "curve-fit" by a

jacket, and balance sting, and the relative location

computer routine, and finite data points are

tabu-of the model

c.g.

with respect to the balance center

lated at the desired angles. The data reduction

pro-is determined (see Fig. 7). It pro-is now ready to be

gram takes into account the same items as in the

installed in the wind tunnel.

point-pause method, i.e., balance/sting deflections,

model weight, etc. The advantages and disadvantages

Testing Methods of the continuous sweep data method are shown in

Fig. 9.

Wind tunnel force data are generally obtained in

one of two methods, either point-pause or continuous

During both the point-pause and the continuous

sweep. Each of these methods has it advantages and

sweep tests, data are usually repeated for a pitch

AEDC-TR-D-AIAN•WS DSADVANIA6$S quality final drag data

t. LiSS TUNNE riMIa. t1. wLEuSAA DATA require measurements

which will allow

calcula-20lBTINi ENTHIR10 RATNA 011211KI ON RUN IN 2. US PRESWKlt INt.AWKNENrIS At START ,/AO (NO ONLY

a2 " TINiI1i

tions

of the mass

flow

through the duct over the full range of test conditions and pressure measurements

4. MUCN LESS M00uL D6TOHTI0U Olt T0 t DATA SIUttON RIO~S M~lt (0iF1T1I STOW1 A[ TIME

from which velocity can be

DIFIEKHIIAL HATINGf

calculated. These measure-Figure 9. Continuous sweep data taking technique. ments can be obtained by separate test runs using mass meters and a set of series, but with the model rolled 180 deg. By com- well-designed pressure probes, or during the standard paring these data to the zero roll data, an evaluation_{of te efecs o nonnifrm M}

2,varitios, low point-pause force test runs using pressure probes and

of the effects of nonuniform (M a. variations, flow svrlsm lfi g as m to s angularity, etc.) tunnel flow can be made and

corrections programmed into the data reduction.

Great care must be taken to interpret these data * I MEASUM iFK A AND TM MORIENTS (UNW IUE UNW

correctly and to not make erroneous adjustments to MMEINT AN HDT W K NOET)

tWhW

AIN DATA R 0NM VERY SWiR TO SAMN SlUIE

the data. WHEN USE- WI MA IMAIK10 1MU B EXEtCIHOD TO NW YNT IT VIM 31 tLEAS 00 NOT NTEIKEM WITI WAE IMAINC

Additional Test Methods 511

RU-TMJNIIGJM SUDTIUSNINElUB

* TMTAL A11t-OH(E MiUST BE AAII IONU UCT NM

The preceding section addressed the standard six- 0 NET IHUM CALCUIATED 1 MASING MASS RFO AW MO V10-component force tests with base pressures. Very OTUM KE

often, additional measurements (Fig. 10) are required

_{* UE}

TM TJO SYSTEM (US)

IFO STORE

SEPAMi

in order to fulfill the test objective. Quite often, • 6g uNODUsJE1IOs

vehicles have control surfaces for which the loads at

various deflections are required in order to ensure Figure 10. Special test requirements. the structural integrity of the component, its

attach-ment, and the control system. These loads are mea- The AEDC Captive Trajectory System (CTS) is sured with a small balance ("fin balance"), usually shown in Fig. 1i. CTS tests usually have a parent three components, mounted internally to the model. vehicle mounted on a six-component balance sup-Great care must be taken during wind tunnel installa- ported on the tunnel standard pitch mechanism and tion of these models to ensure that the electrical leads an additional model of an "external store," (missile to the fin balances, which must "jumper" the main or bomb) mounted on a balance connected to the balance, do not restrict the main balance deflections CTS. The CTS is a mechanism which can produce and therefore cause erroneous readings. These six degrees of freedom independent of the main balances are fabricated and instrumented very similar model support. The attitude and position of the CTS to the main balance and require the same type cali- model relative to the parent vehicle can either be bration and data reduction. They normally measure preprogrammed points in the CTS computer (grid a normal force, hinge moment, and the root bending mode) or determined by the forces and moments moment. sensed by the CTS balance (trajectory mode). This is a highly complex type test and requires a large Another fairly common and yet non-standard amount of pretest set-up. However, once it is model for hypersonic wind tunnel test requires the operational, a large amount of data can be obtained simulation of engines which require ram air. To truly in a relatively short amount of wind tunnel time. The simulate the flight vehicle, the wind tunnel model AEDC/CTS is the only such system known in the must be built with a "flow through" duct. The world to operate at hypersonic speeds. The space internal part of these "engine simulators" cannot be shuttle solid rocket booster separation was tested at fabricated to truly reproduce the drag/thrust of the M = 4 in Tunnel A, and separation of the external full-scale vehicle; therefore, corrections to the mea- tank (E/T) from the orbiter was tested in Tunnel B

sured wind tunnel drag must be made. The data for at Mach 8.

these corrections is provided by one of several means, depending on the desired precision. The highest

AEDC-TR-94-6

temperature instrumentation, and other inputs are

amplified, converted to digital form, and fed to the

data reduction computer. These data are combined

(is SUPPORT

with the previously input balance calibration and

. j

o,

XYZ

VALAE

wind tunnel calibration data. The tunnel calibration

COMPUTER PROGRA MMED

AOM

CONTRIOLLED

_{data are used to calculate the test conditions. Model}

positioning readouts, are combined with the

S•

force/moment data from the balance to calculate the

aerodynamic coefficients and model attitude. These

data can thea be used to resolve the coefficients into

7any axis system (body, wind, pitch, etc.) desired.

MAIN

TUNNEL

MOIL

Total drag is made up of skin friction, base drag, and

SUPPO•T

VARIIi!

wave drag. Since the skir friction is generally a small

5011. DIAMETER

part of the total drag, no adjustments are made for

any difference between the wind tunnel value and

flight value. This is not, however, the case with base

a. Tunnel B installation sketch

axial force. Because of sting effects in the wind

tunnel, the base pressure may be very different from

the flight value. Therefore, an adjustment to the axial

force measured by the balance (CAT)T is made. As

noted in Fig. 12, the wind tunnel base axial force

(CAB)T, as calculated from the measured base

pressure is subtracted from the (CAT)-r to obtain the

forebody axial force (CAF) which is the same for

both wind tunnel and flight. The flight total axial

force (CAT)F is then derived by aLdng a predicted

flight base drag

(CAn)F

to the CAF. Other

adjustments may be applied to the data to

com-pensate for such things as wind tunnel flow

angularity, model/balance misalignments, etc. In the

special case of a model which has a simulated engine

b. Test of shuttle booster rocket separation in

duct, an adjustment to the drag data, similar to the

tunnel A.

base drag adjustment, must also be made. There are

Figure II. AEDC captive trajectory system.

several methods used to calculate the internal drag

of the wind tunnel duci with varying degrees of

ac-Data Reduction

curacy. The most precise method is to measure the

mass flow through the duct using a pre-calibrated

The data reduction program for point-pause type

mass flow-meter in addition to measurements (total

data is quite complicated, and yet much simpler than

and static pressures) from which velocity can be

that for the continuous sweep type. At each desired

calculated, and then to adjust the total axial force

model attitude (data point), the electrical outputs

by the momentum loss of the flow through the duct.

from the balance, model attitude sensors, base

pressure transducers, wind tunnel pressure and

The data reduction

program for continuous

sweep data is the same

as for point-pause,

NOTE:

ASSUMES

& dT -

(Cf),

except for the manner in

which the data points to

be tabulated are

gene--TOTAL AXIAL FORCE COEF rated. As explained

t oSE

AXIAL

FORCE

COEF

C)

= (ۥ7), -

(Ci0)•

(CA,), -

CaF

+

(C4),

earlier, instead of data

CA, - SKIN FRICTION C., (FROM MEASURED BASE PRESSURE) (PREDICED) at finite model pitch/ C4 -FORtEBODY AXIAL FOR(CE OEF WIND TUNNEL ( )I FLIGHT

( )'

yaw positions, a

conti-nuous stream of data is Figure 12. Axial force accounting. generated over a pitch,

AEDC-TR-94-6

yaw, or roll sweep. These data are then fit with a to the measured values provide a quick-look indica-computer-generated polynomial for each variable, tion of the data quality and, in the case of a and then the finite data points are generated at the parametric study helps determine the particular desired model attitudes. These finite points are then configuration that will provide the best flight results. operated on by the computer in the same way as for The comparisons can save a large amount of wind the point-pause data. If adjustments are to be made tunnel time and thereby sizably reduce the ovet all for base pressure, internal drag, or other items which cost of the program. The total data reduction flow require pressure measurement, separate test runs is illustrated in Fig. 13, and typical stability data are must be made in the point-pause mode to allow time shown in Fig. 14.

for the pressure instrumentation to respoud

accurately. These data can also be curve fit and fed SURFACE PRESSURE TESTING into the same adjustment routines as in the

point-pause mode. Surface pressure tests in hypersonic wind tunnels are primarily conducted in association with surface In addition to the calculations, adjustments, heat-transfer and/or flow-field probing measure-and/or corrections discussed above, the data ments to provide inputs to or to validate a CFD code. reduction computer is quite often preprogrammed Of current high interest are pressure tests defining with theoretical predictions (CFD) and/or previously the internal and external aerodynamics of scramjet measured values of the wind tunnel model aero- propulsion systems.

dynamic performance. Comparisons of these values

MEASURED TEST DATA DATA

DATA CONS .T ADJUSTMENTS OUTPUT

OUTPUT

UMODEL AMTTTU.

POSITION CONMD ON

'USEm IIALANE IN IoNTERNAL -NMIOF ATTITUDE

PUSS~~~DAT CNTIUOU W CDMPAIN

DATUTNNELEL

CONDITIONS

Figure 13. Data reduction.

. • 3.00 D.6 -0.01 2.00 e0.4 - 0 1-002 0.2 _{-4.03- 0} 0-0.04 -. 001 -S 0 5 10 1s 20 25 -S 0 S 10 Is 2 25 -5 1 5 10 is 20 2S

AlNGLE Of ATTACK A OF ATTiCK ANGLE OF ATTACK Figure 14. Example of on-line stability data.

AEDC-TR-94-6

Pressure Transducers Since it is important to obtain data as quickly as possible because oi the test environment and/or to Static pressure levels on test models are generally reduce costs, the transducer should be located close very low in hypersonic flow. However, in cases of to the point of measurement on the model to reduce shock wave interaction and/or impingement, they the pressure stabilization time.

can be orders of magnitude higher than the

free-stream static pressure. For this reason, great care Miniaturization has allowed large numbers of must be used in choosing the type of instrumentation individual transducers to be connected together into to be used to obtain the best possible precision, and what is called an ESP (Electronically Scanned also to protect the measurement device from Pressure) unit. These units can usually be housed overload. In general, model surface pressures in within the test model or mounted very close within hypersonic flow are measured with a pressure the mounting hardware. They must, however, be transducer similar to that shown schematicaly in Fig. cooled (usually wvith a water jacket) because the 15a. This differential transducer senses the difference transducers are highly temperature sensitive. The in pressure on the measuring side from that on the advantages of these modules, in addition to their reference side. As seen by the schematic, a small extremely small size and small volume of the overall difference in pressure will cause a deflection of the system, is that they require only one reference line, thin diaphragm, resulting in an electrical output from one calibration line, and one set of electrical leads the attached strain gage. The deflection and, for all of the transducers instead of individual lines. therefore, the output and maximum allowable The major disadvantage is that all surface pressures pressure differential, is a function of how rigid (thick) must be near the same value because all of the trans-the diaphragm is made. These transducers are ducers within a unit have the same maximum pressure manufactured in pressure ,atings from a few rating and, as stated, use a common reference hundredths of a psid to several thousand psid. pressure. They are, however, usually protected against a large overload. A typical unit capable of

su1muUM measuring up to 32 model pressures and manu-factured by Pressure Systems Incorporated is shown

MEASUREMENT in Fig. 15b. MODEL SURFACE SIDE REFERENCE

REFERENCE - _{Transducer Calibration}

PRIESSURE - SIDE

PRiESSURE-L _Pressure transducers must be calibrated, as in the

ELECTRICAL

LEADS

case of the force balances, to determine the electrical

FROM STRAIN GAGE output of the strain gage as a function of applied a. Pressure transducer schematic. pressure differential across the diaphragm. Unlike the balance calibrations, however, the pressure trans-ducers are usually calibrated at least once a day

during use. These calibrations take only a few

minutes because there is only one component and, therefore, no interactions. The calibration can be accomplished in one of several ways, depending on the type and magnitude of the rated pressure of the transducer. The most common method used for transducers rated up to atmospheric pressure is to reduce the pressure on the reference side of the diaphragm in increments by applying a known

(6.3 cm) .4.3cm) ( pressure less than atmosphere. The magnitude of the

2.5 3 [-0. 2 1.7 applied pressure is measured with accurate instru-__ • mentation (secondary standard) traceable to a

(2.S cm) 000000000primary meao standard. A secondary standard (c ry nr. ) a is l a field

1.0 R'REFERENCE measurement device traceable to a laboratory

pri-EASUEENT TU - -mary standard which is, in turn, traceable to the

STUBES 32 CALIBRATE TUBE National Institute of Standards and Technology. b. Electronically scanned pressure unit (ESP) Since most transducers are not linear through zero Figure 15. Pressure measurement device, pressure differential, if values of the test pressure to

AEDC-TR-94-6

be measured are expected to be both above and below with stainless steel tubing attached by one or more the value of the reference pressure to be used during of the methods shown in Fig. 16. When the confi-the test, confi-the applied calibration pressure should also guration is such that the backside or inside of the cover both cases. Quite often, several individual model is accessible, the type installation shown at the transducers being used for a test are connected to a left is the most desirable. In this type installation. common reference, as is the case for the ESP units. a small hole, usually less than 0.050 in. (1.3 mm) In these cases, all transducers on a common reference diam, is drilled through the model wall, a counter-can be calibrated at the same time. bore the size of the O.D. of the tube is drilled part way through from the backside, and the stainless steel If a transducer has a sealed vacuum on the tube is solderqd in place. Either of the two right-hand reference side of the diaphragm (an "absolute" installations can be used when there is not enough pressure transducer), or the maximum rated pressure room or access to solder the tubes on the backside. is greater than atmospheric pressure, or it is an ESP Care must be used in installation and/or handling unit of the type shown in Fig. 15, it must be calibrated this type model, or leaks can develop around the tube from the measurement side of the diaphragm. This or the tapered plug, resulting in measurement errors. is accomplished by sealing off the tube going to the Any of the three installations should be checked for model surface and applying the reference pressure leaks by applying a vacuum to the surface tap, sealing through a "T" as shown in Fig. 16. In the case of it off, and observing the test instrumentation over the ESP unit, this valve is internal to the unit; a few minutes to obtain a leak rate. The entire qystem however, for other types of systems, the valving and must be clean and free from foreign material (such calibration tubing must be added to the system and as oil), or the outgassing may appear to be a leak.

The number of pressure taps on a model can be limi-can become quite complicated to ensure there are no

leaks anywhere in the system. The calibration ted by available room for internal ESP units or for_{the tubes routed through the mounting hardware. As} pressure is applied and measured in the same manner the cas of the fre models, herscale ot

as te rferece idecaliraton.in the case of the force models, the scale of the

as the reference side calibration.

pressure models and the fabrication process must be MODEL SUUtKE such as to maintain the ,,_,:'",•/ f~•1Pl STANOAU0 fidelity of the full-scale "1 --- SEE SKETCHES uW T--RANSDUCER ROMA configuration. The "as-built"

REIFERENCI/AMIAATE

location of the surface

.. "

SSU SURrLY

_{pressure taps is critical to}

---- _•. i.-ALTERNATE -

M1ASUIRMENT

obtaining data which will

VA L'V EISE IAwIoATIIN SI.M accurately predict the flight MEASUREMENT SIO ST1-1ONO ANAO pressure loadings or which

CAUIIIATION SYSTEM IM can be used to validate

theoretical computer codes.

Testing Methods SWARED OR SWEAT SOKID t[ORED

AND MACHINED AFTERWARD SWAGA* _AND PLUG Testing of surface pressure

MACHINED AFTEIVWARO distribution models in

hyper-ACC

SUIE

INACESSILE IACiASSD

sonic wind tunnels can be verN

time consuming and therefore

Figure 16. Pressure measurement and calibration schematic. more expensive than static stability testing. In the intermittent tunnels, usually data for only one model Model Fabrication attitude per run can be obtained because of the time it takes for the pressure to stabilize. For the same As in the case of the static stability model, the reason, even in the continuous flow tunnels, it takes choice of test facility dictates the maximum size longer to get the pressure data over a pitch or yaw model. Weight is not a critical item in the pressure polar than even the "point-pause" type force data. models; therefore, reinforcing and other fabrication Even when the model and pressure systems are techniques can be used to reduce the thermal distor- optimized to reduce the required stabilization time, tion during testing. Pressure models for hypersonic it still requires up to minutes per model attitude. testing are usually fabricated from stainless steel, Because of the expense in time/money, pressure tests

AEDC-TR-946

are usually designed for data points in wider incre- AEROTHERMAL METHODOLOGY ments than the normal force test; i.e., where data for

a force test may obtained for - 15 < n < + 15 in Fundamentals and Simulation Parameters

2-deg increments, the pressure test would probably The requirements for developing hypersonic flight be 5-deg increments, vehicles place increasing demands on ground test Data Reduction capabilities. Of particular concern is the requirement to demonstrate that flight components such as leading edges, cowl lips, and structural panels will Test data are combined with previously input survive the aerothermal flight environment. Specific calibration data (transducers and wind tunnel), and components as shown in Fig. 18, can experience heat-the data output parameters i.e., P, P/PC., Cp etc. ing rates ranging from 200 to 2,000 Btu/ft2-sec and are calculated. A data reduction technique called the surface temperatures from 1100-19400C (2,000 to pressure prediction routine can be used to reduce the 3,500°F). Ground test of flight components have pre-amount of time required to obtain the pressure data viously been performed at test facilities like those at for a given model attitude. To apply this technique, NASA and AEDC. This section presents an overview the output of the pressure transducer is recorded in of the materials/structures test methodology and the uniform time increments (- I/sec) for a period of test techniques used in the development of hypersonic time (usually 30 sec) after the model has been set at vehicle components.

the desired attitude. The results of these samplings are

curve fit within the computer, and the FATNER WINGITAIL

results are extrapolated to the JOINTS AND SEALSN '(PROITUBERAN(ES) LEADING EDGES

asymptotic value of the actual pres- REWIEQOUIPMENT / -IOUNItY LAYER TRANSITION

sure. The technique, illustrated in Fig. WINDOWS / ___ __

17, takes into account such things as lox

pressure tube geometry, gas tempera- NOSE CONTROL

ture, viscosity, system geometry, etc. COOLING SURFACES

As shown by the illustration, the actual HOT STRUCTURE '/ U10LIM ENGINE

INUATO 111OIWING PANELS COOLING

equilibrium value of the pressure was AND FUEL

predicted very accurately by using the

input obtained during only 30 sec when Figure 18. Typical aerothermal structures/materials issues. it would have taken well over a minute

for the system to stabilize to the final value. The wind A review of some fundamental heat-transfer tunnel test time/costs can be greatly reduced with concepts is presented in Fig. 19. The typical textbooks very little sacrifice in accuracy by applying this discuss the convective heating to a wall and relates technique. the heat flux, q, to a temperature difference, Tg-Tw. For aerodynamic heating the heat flux is also

NOWUA ran- rMy proportional to the temperature gradient at the wall,

1.i1 = t.0.04+0W TaIM, - s.St m II and the heat-transfer coefficient, h, is used to relate 0.4482 0.06S - smot the heat flux, 4, and the temperature driving

09 0DImO FINA

1

msIU potential, Tr-Tw, where Tr is the recovery

.392 0.055 + 4s &1re i.t" 11.0as10 i nsa) temperature. The experimentalist often uses the

44 45 *.111IN gtS1.1n M

o (00 o .•8 t.e (01111 prA) facility total temperature, To, in place of the more

. . 0 . X. -0* .15 06 P.1167 rI _{elusive recovery temperature Tr.}

S0.2413

0.035 I%

Ir 'A common approach used to solve aerothermal 0.1724 0.02S •issues is based on combining analysis with

experi-0.11240.025mentation. It is imperative that analytical techniques 0.1034 0.015 , n i , r--- _{be used to plan the test and to analyze the final data.} 0 10 20 30 40 sO o 0 The two fundamental steps in the development

pro-TIME,

SEC

cess are:

Figure 17. Typical pressure stabilization curve. 1. defining the flight thermal environment 2. demonstrating hardware survivability

AEDC-TR-B44

is then used to extrapolate the results CONVECTION AEROHEATING to higher Mach numbers incorporating

TIS,

k is real-gas and viscous effects as

re-quired. This procedure is illustrated in

T

Fig. 21, which includes the primary test

facilities used to obtain heat-transfer

Yly- 0

data in the U.S. The AEDC Tunnels

WALL

(W)

_WAL

_4!1YT.

_WALL

_(W)

I

B and C are the national workhorse

_{facilities in this category, and it has}

q - 0(1 -

TQ"

q- 4 been estimated that 75 percent of

k'lp

I

- - existing hypersonic data defining

LET k- NTh',-W:] thermal environments have been

"WHERE:

.l - HEATING RATE, STU"FT•E 5C obtained in these tunnels.

h - HEAT TRWNSFER COEFFICIENT, The test techniques available

ITU/M SEC OR _{measure aero-heating are listed in I}

T, - RECOVERY TEMPERATURE, °1 22, along with a reference wh)

(V, T,) illustrates the use of the technique. T. - WAlL TEMPERATURE, -1 Thermal mapping techniques provide

a comprehensive look at the entire

Figure 19. Basic aeroheating concepts.

model and are often used to identify

PHASE

I - DEFINING

THEtML ENItOW TS

_PHASE

₁

(STEP

1)

the location of high heating rates (e.g.,

.

THERMALtiNn).NMHowev(STEthI

sEoNkG

* SCALE MODELS IN WINS TUNNELS * NAI RWSFEI TEST TECHNIQUE

shock interaction). However, the

uncertainty of thermal mapping data

is of the order of t 15 percent,

whereas the discrete measurement

Pt"tS.,

C OE %EMaIIo1 =: RAPW. TO FUMG techniques can produce

±

6-percent

SHEATING IPU ) data which are more reasonable for STHERML ENINMENT code validation tests. Additional details on heat-transfer measurement

PHASE 2 -DENONSTRATE HARDWARE SURWMABILITY (STEPS 2, 3, 4) techniques may be found in Refs. 4, 7,

"MATERIAL

TEST STRUCTURAL CO I FLIT HARDWARE and

10,

and a brief overview of each

TEST

DENO

TEST

technique is presented below.

(SAMLES)

(CMONENTS)

((COMPONENTS)

SDUPLICATE LOCAL T ARTICLES Phase Change Paint Technique ENVIPIUMENT tzfTT

(a.*. q LOCAL - q FUENT) The Phase Change Paint technique

of measuring the heat transfer to a

Figure 20. Methodology for aerothermal structures/materials

model surface was developed by Jones

development.

and Hunt.II This technique assumes

that the model wall temperature

re-These two fundamental phases are illustrated in

Fig. 20. In defining the thermal environment, the

STEP

(1)

versatility of analytical tools is combined with the

DEFINING

THERMAL

ENVIRONMENTS

experimentally measured heating distributions. These (I.E., WHAT HEATING RATES/ITEMPERATURES ARE ENCOUNTERED IN FLIGHT?)

data, obtained on scaled models in simulated flow APPROACH:

environments, are used to verify the accuracy of the ANALYSIS EXPERIMENTS

analytic tools. The important simulation parameters * ESTIMATE HEATING DISTRIBUTION * SCALE MODEL TESTS/SIMULATED o ENGINEERING CODES MACH AND REYNOLDS NUMBER

are Mach number and Reynolds number. However, 0 CGO e AEDC TUNNELS I & C

it may not be necessary to match the Mach number * EXTRAPOLATE DATA TO FLIGHT 9 NSW( TUNNEL 9 because of the "Mach number independence princi- * THERMAL RESPONSE CALC. * AMES * CALSPAN 3.S FT TUNNELSHOCK TUNNELS

pIe." A commonly used procedure to define the o Lot( TUNNELS

thermal environment is to use the data obtained at

Mach 8 or

10

to substantiate a code at precisely the

Figure 21. Development process, step I.

same conditions as the experimental data. The code

AEDC-TR-94-4

ANVAITAOEI INSIANTAUS TlF.

* PNASE-CmAMGE PAMN1 VIM ILLUSTRATION O NOT ISPOT MUST REAItY PANT. DATA PRESEIIATN 2

NI6N SATIAL RESOLUTION CAMN BE WNFUI

* I S0 liS CAMERA (DUPUIERIENERTme PKlTS Ai 0OlO SAITAL REISUTION 3

oMA, ANODNMUO US9W

* TiEID06APNK PNOSPINR 0UPFLEI M , GM SPATIAL MOEL PEPARATID1 AI NATA 4.11

1E1S1,1ON PIESENTATIN

"*

11N-Sm. NIGH N1ANlY MATA, MIE SPACING OXPHIIM _EFFT•MREL FAD, EU"I S

"*

WAX GAGE EASY TO INSTALL, ONTOUWU, KOURIM M LO OUTPUT, TST IESon 6

"*

SiOIMMT.4ELTER 6AE IG OUTPUT, SUMi i TOfuML, FAN AN WJLIATUN TIE REUiM 7

VERY WRINLE

"*

GWEAN WAS (III TEMP, LO TEMP) YEWAl OF EIPERIIEC. FAST RESPOINSE GAGE ATTIIOI RATI, NOT IUM I

"*

INI-FU DENSE SPACING, FAST RESS, CAN U RELATIVELY FFIWLT iNTALLATKII, 9

USED ON SMAIL RAN MATERIIAL ]M(36

Figure 22. Test techniques available for measurement of heat transfer.

PRINCIPLE OF OPERATION

SURFA(E TEMPERATURE MEASUREMENT OF SEMI-INFINITE SL

T,

~ith~~

GaTIME DATA REDUCTION EQUATIONS

TV - T,

I

T,-Ti (B)2ACKTE-

iUr#K

WHERE 0 - e (2)

Figure 23. Phase change paint technique. sponse is similar to that of a semi-infinite slab subjected to an instantaneous and constant

heat--transfer coefficient (see Fig. 23). The surface wall temperature rise for a semi-infinite slab is given by the equations shown in this figure.

A specific value of the wall temperature (Tpc) is indicated by a phase change paint (Tempilaq® ). These paints change from an opaque solid to a transparent liquid at a specified phase change temperature (Tp). For known values of Ti, Tpc, t,

and th hea-trnsfr coffiien (h)canI~e Figure 24. Typical examples of phase-change

and QV-eck, _{the heat-transfer coefficient (h) can be pitpoorps}

calculated as a function of the time required for the paint photographs.

phase change to occur by using mechanism at the desired test attitude, and the model

h

ck

(3)

initial temperature (Ti) is measured. The model is

A

then injected into the airstream for approximately 25

sec, and during this time the model surface

tempera-where P comes from the solution of Eq. (1) since the

ture rise produces isotherm melt lines. The

pro-left-hand side is known.

gression of the melt lines is photographed with

70-mm sequenced cameras operating at one or two

Prior to each run, the model is cleaned and cooled

frames per second. Typical examples of phase change

with alcohol and then spray-painted with Tempi-

paint photographs obtained during a run are

pre-laq® . The model is installed on the model injection

sented in Fig. 24, and an example of phase change

AEDC-TR4-96

SIM DATA Infrared

Sca"ning

OPE

_{a PAN, SIO}

111111-SO

MRl

_{MR FUSLG VENT,}

,O, FLAVAI PA1, PWUhAE 11101 111. PILOT WITA Thermal mapping

techni-Va - I NOiS WIL- q ques used in wind tunnel test

II W •• applications generally involve

5.6 _ the use of heat-sensitive model 0.1 surface coatings. The major

drawback to these methods has been the time required to

0

NkIGN.l:~

AIS~obtain

quantitative data from

"0.01

- %( SMED DOW TO SALE. t photographic test results.

00

0

oWith _{an infrared (IR)}

0 0 0 o o o scanning camera system,

heat-0.00 _____________________ transfer coefficient data in the

0 0.1 0.2 0.3 0.4 0.5 form of tabulations, plots,

IlL and surface maps are pro-Figure 25. Leeward centerline heat-transfer distributions at duced within minutes of test

ReL = 8.6 x 106, a = 30 deg. run completion.

A typical installation of the AGA Thermovision 680 paint data compared to thin-skin data is presented scanning camera for an aerodynamic heating test is in Fig. 25. This figure also illustrates a common sketched in Fig. 26. The camera is positioned outside technique used to extrapolate wind tunnel heating of the wind tunnel environment. The infrared distributions to flight. The wind tunnel data are radiation emitted by the test model within the field normalized by the Fay-Riddell stagnation point of view of the camera is collected by the system optics heating on a 1-ft-radius sphere scaled down to the and focused on the camera detector. The signal gen-model scale. To obtain heating rates in flight, the erated by the detector is proportional to the detected distribution is multiplied by the Fay-Riddell heating infrared radiation. Two rotating prisms form an on a I-ft nose radius sphere "flown along the flight optical scanner which controls the position of the trajectory." The basic assumption is that the distri- instantaneous field of view (IFOV) of the camera.

bution at M

=

8 is unchanged from tunnel to flight.

_

rd._-

L_.

A complete description of the phase change paint FLOW , =....

Oi1EB

technique as applied to a particular test situation is iE presented in Ref. 2.

OPIA

%.. TEST

MON a

R U M P

2

SC IU E M

TEST

DATA DATA

MODEL

-

SYM

DIGITIZER

11 CAMERA

WmaIND

_M

AWAM

I

~~~TUNNEL

ALTPOS

DATA

NOE Figure 27. Infrared system schematic. O.M (1.3.) The complete infrared system in use at AEDC is

schematically illustrated in Fig. 27. The system is

VIEW ILOIE NUIAUA1I K TMIL M composed of the test model, the AGA 680 camera, and a data system to collect, digitize, and convert the Figure 26. Sketch of typical IR camera installation, camera signal to the desired data output. For a typical

AEDC-TR-94-6

test, there are several modes of data output that can area in the object plane. Since the detector in the be selected to fulfill the test requirements. One type AGA 680 camera is circular in scope, a "spot of data output is a tabulated output of model surface diameter" is viewed in the object (test) plane (see Fig. temperature or heat-transfer coefficient for each 29a).

desired position within the total field of view. A

capability of presenting the temperature map of the The ability of the IR system to track a tempera-model surface in the form of a color plot is used, and ture profile across a "worst case" step-heating gradi-a sgradi-ample is presented in Fig. 28. Other forms of dgradi-atgradi-a ent will be discussed to aid in the understanding of presentation consist of both 70-mm photographs and the data. Assume that the IR camera is scanning 16-mm movies of the color video monitor. along the centerline of a target that is composed of a plate at a uni-form temperature, Tc, that has a circular protuberance at an elevated uniform temperature of

MO

Validat I ns• TH as shown in Fig. 29b. For a

Tis Esystem with perfect optics and electronics, the ability to track a step increase in temperature is a function of only the IFOV which determines the scanning spot diameter. If the camera had an infinitesimal IFOV, the tempera-ture profile would be tracked exactly as shown in Fig. 29b. This

an 3st 3g 41o 46s sis s64 1s Go,£ ,, rs, os ,9 ,s, ,,, less 112i, tii. case is of academic interest only, ' 273 2a 378 *419 467 S6, SI US 6 14 1'3 409 ese go? isz 95s i, toss but represents what would be re-quired to track a step change in Figure 28. Illustration of IR thermal mapping data. temperature exactly.

To aid in the understanding and interpretation of SPO _M

the data obtained with the IR scanning system - 01"

requires the following: (1) a definition of some basic

Dr-YN -

T

7

terminology associated with the IR system and the UNE .--- -data, and (2) a review of basic IR system operational O.4SCAN

characteristics and limitations. T_-. I The total field of view of the camera at the wind

tunnel centerline is a function of the lens selected for a. IR target the camera and the distance from the camera focal

point to the centerline. The desired total field of view

is determined based on the size of the test model to TN be viewed and the spatial resolution (to be discussed

later) that is desired. The total field of view is a rectangle as shown in Fig. 29a. One complete scan of this total field of view is defined as a frame of data. Each frame of data is composed of a matrix of 70 line scans, each containing 110 points, for the total of 7,700 discrete measurements. Each line/point

*combination identifies the location of the centerline b. IFOV = infinitesimal

of the IFOV of the camera as it scans the total field Figure 29. IR scanning of a step change in

of view. temperature.

The IFOV of the camera is specified as the angle Next, we will assume a practical spot diameter of in milliradians subtended by the projected detector 0.16 in. (4 mm) based on an 8-deg lens viewing the