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W IS020 CALSPAN F1ELD SERVICES INC ARNOLD AFS TN AgOC DIV F/6 22/2NASA/ROCKWELL INTERNATIONAL SPACE SHUITTLE ORBITER ABOPT t#ATINB--ETC CUl
r, UwW-7LASS-,,EDNOV &I L A TICATCH, K W UT I AJTR6-V6
O N IFE
ECS-B-3
Hill II*~L 111 1.Q_I..0
mill,
111111.25 1 .4
MICROCOPY RESOLUTION TEST CHARTNATIONAL BUREAU Of S ANDARDS 1963 A
AEDC-TSR-81-V36
NASA/ROCKWELL INTERNATIONAL SPACE SHUTTLE
ORBITER ABORT HEATING TEST (OH-ill)
~ L. A. Ticatch and K. W. Nutt
_____Caispan Field Services, Inc.
Novemiber 1981
Final Report for Period September.25-30, 1981
* Approved for public releese distribution unlimited.
DTrI CELECTE
JUNO 2198t
ARNOLD ENGINEERING DEVELOPMENT CENTERARNOLD AIR FORCE STATION, TENNESSEE
AIR FORCE SYSTEMS COMMANDUNITED STATES AIR FORCE
82 06 01 177
T"
NOTICES
When U. S. Government drawing, specifications, or other data are used for any purpose otherthan a definitely related Government procurement operation, the Government thereby incus noresponsibility not any obligation whatsoever, and the fact that the Government may haveformulated, furnished, or in any way supplied the said drawings, specifications, or other data, isnot to be regarded by implication or otherwise, or in any manner licensing the holder or anyother person or corporation, or conveying any rights or permission to manufacture, use, or sellmy patented invention that may in any way be related thereto.
References to named commercial products in this report are not to be considered in any senseas 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 NationalTechnical Information Service (NTIS). At NTIS, it will be available to the general public, includingforeign nations.
APPROVAL STATEMENT
This report has been reviewed and approved.
J. T. BESTAeronautical Syst -.as BranchDeputy for Operations
* Approved for publication:
FOR THE COMMANDER
r mo atp , :rrctorrospace Flight Dynamics Test
leputy for Operations
I I_________________________
UNCLASSIFIED ____
SECURITY CLASSIFICATION OF THIS PAGE (When Data FInt'red)
REPORT DOCUMENTATION PAGE READ INSTRUCTIONS1. REPORT NUMBER 2. GOVT ACCESSION NO. 3. RECIPIENT'S CATALOG NUMBER
AEDC-TSR-81-V36
4. TITLE (and Subtitle) S. TYPE OF REPORT & PERIOD COVERED
NASA/ROCKWELL INTERNATIONAL SPACE SHUTTLE Final Report
ORBITER ABORT HEATING TEST (OH-111) September 25-30, 19816. PERFORMING O1G. REPORT NUMBER
7. AUTHOR(&) S. CONTRACT OR GRANT NUMBER(s)
L. A. Ticatch and K. W. Nutt, Calspan FieldServices, Inc./AEDC Division
9. PERFORMING ORGANIZATION NAME AND ADDRESS 10. PROGRAM ELEMENT. PROJECT. TASKAREA & WORK UNIT NUMBERSArnold Engineering Development Center Program Element 921E01
Air Force Systems Command Control No. 9E01Arnold Air Force Station, TN 37389
!I. CONTROLLING OFFICE NAME AND ADDRESS 12. REPORT DATEJohnson Space Center (NASA-JSC(ES3)) November 1981Houston, Texas 77058 13. NUMBER OF PAGES
5314. MONITORING AGENCY NAME & ADDRESS(if different from Controlling Office) IS. SECURITY CLASS. (of this report)
UNCLASSIFIED
15a. DECLASSIFICATION/DOWNGRADINGSCHEDULE N/A
16. DISTRIBUTION STATEMENT (of this Report)
Approved for public release; distribution unlimited.
17. DISTRIBUTION STATEMENT (of the abstract entered in Block 20, if different from Report)
IS. SUPPLEMENTARY NOTES
Available in Defense Technical Information Center (DTIC).
19. KEY WORDS (Continue on reverse side if necessary and identify by block number)
heat transferthin skinspace shuttle orbiterhypersonic testing
20. ABSTRACT (Continue on reverse side If necessary end Identify by block number)
Thin-skin thermocouple heat transfer tests were conducted on two 0.0175 scaleand one 0.04 scale models of the Space Shuttle orbiter at attitudes that wouldbe encountered in a transatlantic abort maneuver. The model angles of attackranged from 40 to 55 degrees with yaw angle varying from -2 to 2 degrees. Datawere obtained at Mach 8 in the AEDC-VKF Hypersonic Wind Tunnel B at free-stream Reynolds numbers ranging from 0.5 x Tu/ to 1.5 x 106 per foot.
DD I FAN"73 1473 EDITION OF I NOV 65 IS OBSOLETE UNCLASSIFIEDSECURITY CLASSIFICATION OF THIS PAGE (When Date Entered)
: .4
CONTENTS
Page
NOMENCLATURE ......... ........................ 31.0 INTRODUCTION ........... ....................... 62.0 APPARATUS
2.1 Test Facility .......... .................... 62.2 Test Articles .......... .................... 72.3 Test Instrumentation
2.3.1 Test Conditions ........... ........ 72.3.2 Test Data .......................... 8
3.0 TEST DESCRIPTION3.1 Test Conditions ... ....... . ............ 83.2 Test Procedure
3.2.1 General ......... ................... 83.2.2 Thin-Skin Thermocouple ...... ............ 93.2.3 Oil-Flow ......... ... .............. 9
3.3 Data Reduction3.3.1 Thin-Skin Thermocouple Data ............ 9
3.4 Uncertainty of Measurements ... ............. ... 114.0 DATA PACKAGE PRESENTATION ..... ................ ... 12
REFERENCES .... .................. .......... . 12
APPENDIXES
I. ILLUSTRATIONS
Figure
1. Tunnel B ......... ......................... .... 142. Installation Photograph of 60-0 Model .. .......... ... 153. 60-0 Model Installation ................. 164. Basic Dimensions and Coordinate System for the 0.0175 *
Scale Orbiter Models ....... .................. .. 175. Installation Photograph of 56-0 Model ........... 186. 56-0 Model Installation .. 197. Installation Photograph of 83-0 Model .......... 208. 83-0 Model Installation ..... ................. ... 219. Basic Dimensions and Coordinate System for the 83-0 Model 22
10. Thermocouple Locations on 60-0 Model .. .......... . 2311. Thermocouple Locations on 56-0 Model .. .......... . 2912. Thermocouple Locations on 83-0 Model .. ............. 30
* 13. Thin-Skin Thermocouple Plotted Data .. ........... ... 3414. Oil-Flow Photographs on 60-0 Model (Run 262) ...... . 35
II. TABLES
Table
1. Data Transmittal Summary ................ 37
2. Estimated Uncertainties . . . . .......... .. 38
3. 60-0 Model Thermocouple Coordinates . . . . . . . . . .. 40
4. 56-0 Model Thermocouple Coordinates . . . . . . . . . .. 43
14I
Page
5. 83-0 Model Thermocouple Coordinates .. .. ........ 446. Test Data Summary. .. ...................7. Photographic Summary .. ................... 50
III REFERENCE HEAT-TRANSFER COEFFICIENTS. .. ........... 51
IV SAMPLE TABULATED DATA
1. Thin-Skin Thermocouple Tabulated Data. .. .......... 53
rAccession
For
DTIC TABiA AUnannouncedJustification E
Distribution/
AvailaLc41itY Codes
'Avail-adoDit Special 1
2
NOMENCLATURE
ALPI Indicated pitch angle, deg
ALPHA Angle of attack, deg
ALPPB Prebend angle, deg
b Model skin thickness, in.
B Wing span, in. (see Fig. 4)
BV Height of model vertical tail, in. (see Fig. 4)
c Model material specific heat, Btu/lbm-*R
C Local chord of wing or vertical tail, in. (see Fig. 4)
DELTAE Elevon deflection angle, deg
DELTASB Speed brake deflection angle, deg
DELTBF Body flap deflection angle, deg
DTW/DT Derivative of the model wall temperature
with respect to time, 0R/sec
H(REF) Reference heat transfer coefficient (seeAppendix III)
H(TR), H(TT) Heat-transfer coefficient based on recoveryH(O.9TT), H(O.85TT) temperature, TR (TR = TT, O.9TT, or 0.85TT assumed
for these data), QDOT/(TR-TW), Btu/ft2-sec-*R
L Reference length, in. (see Fig. 4).
rM, MACH NO. Free-stream Mach number
MU Dynamic viscosity based on free-stream
temperature, lbf-sec/ft2
I!3
MUTT Dynamic viscosity based on TT,lbf-sec/ft2
P Free-stream static pressure, psia
PT Tunnel stilling chamber pressure, psia
PT2 Stagnation pressure downstream of a normal shock,psia
PHI Radial angle location of thermocouple in modelcoordinates, deg (see Figs. 4 and 9)
PHI1 Indicated roll angle, deg
Q Free-stream dynamic pressure, psia
QDOT Heat-transfer rate, Btu/ft -sec
RE Fr-e-stream unit Reynolds number, ft-
-3RHO Free-stream density, lbm/ft
RN Reference nose radius, (0.0175 ft or 0.04 ft,determined by model scale)
RUN Data set identification number
STFR Stanton number based on reference conditions(see Appendix III)
T Free-stream static temperature, OR
TC NO Thermocouple identification number
TIME Elapsed time from lift-off, see
TR Assumed recovery temperature, OR
TT Tunnel stilling chamber temperature, OR
TW Model surface temperature, *R
V Free-stream velocity, ft/sec
X Model scale axial coordinate from model nose or
leading edge of wing or vertical tail (see Fig.
4 and 9) in.
XO Full scale axial coordinate from a point 235 in.ahead of the orbiter nose (see Fig. 9), in.
Y Model scale lateral coordinate (see Fig. 4), in.
YAW Yaw angle of model, deg
YO Full scale lateral coordinate, in.
Z Model scale vertical coordinate (see Fig. 4), in.
ZO Full scale vertical coordinate, in.
P Model material density, ibm/ft3
.1
-1
5
1.0 INTRODUCTION
The work reported herein was performed by the Arnold Engineering
Development Center (AEDC), Air Force Systems Command (AFSC), underProgram Element 921E01, Control Number 9E01, at the request of theJohnson Space Center (NASA-JSC(ES3)), Houston,Texas. The NASA-JSC (ES3)
program manager was Mrs. Dorothy B. Lee and the Rockwell International
project engineers were Mr. C. L. Berthold and Mr. J. Gee. The resultswere obtained by Calspan Field Services, Inc./AEDC Division, operatin,contractor for the Aerospace Flight Dynamics testing effort at the '.z'DC,AFSC, Arnold Air Force Station, Tennessee. The tests were perfor-.ed inthe von Karman Gas Dynamics Facility (VKF), under AEDC Project No. C628VB.
The test was performed in the 50-ii.-diam Hypersonic Wind Tunnel(B) at the von Karman Gas Dynamics Facility (VKF) during the periodSeptember 25, 1981 to September 30, 1981. Data were recorded at Mach
number 8 for nominal Reynolds numbers ranging from 0.5 x 106 to 1.5 x106 per foot. The nominal model angles of attack ranged from 40 to 55degrees with model yaw angles varying from -2 to 2 degrees. All thin-skin thermocouple data were obtained from three space shuttle orbitermodels designated 56-0, 60-0, and 83-0.
The test had a NASA/Rockwell designation of OH-111. The testobjective was to obtain thin-skin heat transfer data on the spaceshuttle orbiter model at attitudes that would be encountered in atransatlantic abort maneuver.
A summary of the test data transmitted is shown in Table 1.Inquiries to obtain copies of the test data should be directed toNASA-JSC (ES3), Houston, Texas 77058. A microfilm record has beenretained in the VKF at AEDC.
2.0 APPARATUS
2.1 TEST FACILITY
Tunnel B (Fig. 1), is a closed circuit hypersonic wind tunnelwith a 50-in. diam test section. Two axisymmetric contoured nozzlesare available to provide Mach numbers of 6 and 8 and the tunnel may
be operated continuously over a range of pressure levels from 20 to300 psia at Mach number 6, and 50 to 900 psia at Mach number 8, withair supplied by the VKF main compressor plant. Stagnation temperatures
sufficient to avoid air liquefaction in the test section (up to 1350*R)are obtained through the use of a natural gas fired combustion heater.The entire tunnel (throat, nozzle, test section, and diffuser) is cooledby integral, external water jackets. The tunnel is equipped with amodel injection system, which allows removal of the model from the testsection while the tunnel remains in operation. A description of thetunnel may be found in Ref. 1.
6b.
2.2 TEST ARTICLES
Three Space Shuttle orbiter models were used to obtain the thin-skin thermocouple data for this test. Two of the test articles were0.0175 scale models of the full orbiter and were designated as the60-0 and 56-0 models. The third model was a 0.04 scale of the fronthalf of the orbiter and was identified as the 83-0 model. All of themodels were supplied by Rockwell International.
The 60-0 model was a 0.0175 scale thin-skin thermocouple model ofthe Rockwell International Vehicle 5 configuration. The model was con-structed of 17-4 PH stainless steel with a nominal skin thickness of0.030 in. at the instrumented areas. All thermocouples were spot weldedto the thin-skin inner surface.
A photograph of the 60-0 model injected in the tunnel is shownin Fig. 2. A sketch of the 60-0 model installation in the tunnel isshown in Fig. 3. The basic dimensions and coordinate definitions forthe 0.0175 scale models are shown in the sketch presented in Fig. 4.The deflection angles of the speedbrake, elevons, and body flaps wereall set at zero throughout the test.
The 56-0 model used for this test was model number 2B of thematerial "LH" 56-0 phase change paint model series. This was a 0.0175scale model with the same external contour as the 60-0 model. Thepilot side of the fuselage consisted of a thin-skin thermocouple insertcontoured to the vehicle lines. This insert was constructed of 17-4stainless steel with a nominal skin thickness of 0.020 in. at thethermocouple locations. A photograph of the 56-0 model injected inthe tunnel is shown in Fig. 5. A sketch of the 56-0 model installationis shown in Fig. 6. The dimensions and coordinate system presented inFig. 4 also apply to the 0.0175 scale 56-0 model.
The 83-0 model was a 0.04 scale model of the forward half of theorbiter. This model was also constructed of 17-4 PH stainless steelwith a nominal skin thickness of 0.030 in. A photograph of the 83-0model in the installation tank beneath the test section is shown in Fig.
47. The installation sketch of the 83-0 model is shown in Fig. 8 andthe coordinate system and basic dimensions for the 83-0 model arepresented in Fig. 9.
2.3 TEST INSTRUMENTATION
2.3.1 Test Conditions
The instrumentation, recording devices, and calibration methodsused to measure the primary tunnel and test data parameters are listedin Table 2a along with the estimated measurement uncertainties. Therange and estimated uncertainties for primary parameters that were cal-culated from the measured parameters are listed in Table 2b.
7
2.3.2 Test Data
The 60-0 model was instrumented with 600 thirty-gauge iron-constantan
and Chromel®-constantan thermocouples. Only 250 of these thermocouples
were used on this test. Thermocouple locations for this model are
presented in Fig. 10; the dimensional locations and skin thickness for
the thermocouples connected on this test are listed in Table 3. Thethermocouples identified by a number only are iron-constantan. The
thermocouples identified by a number followed by the letter A or C are
Chromel-constantan that were added to the model. The letter D after a
thermocouple number designates an iron-constantan thermocouple in a new
location on the OMS pod.
The 56-0 model instrumentation consisted of 80 thirty-gauge Chromel-
constantan thermocouples located on the thin skin insert. All of these
thermocouples were connected on this test. The thermocouple locations
for this model are presented in Fig. 11. The dimensional locations
and skin thicknesses are listed in Table 4.
For this test only 250 of the 482 thirty-gauge Chromel-constantanthermocouples on the 83-0 model were connected. The thermocouple loca-
tions for this model are illustrated in Fig. 12. The dimensionallocations and skin thicknesses for the thermocouples used on this
test are listed in Table 5.
3.0 TEST DESCRIPTION
3.1 TEST CONDITIONS
A summary of the nominal test conditions at each Mach number is
given below:
M PT, psia TT, OR Q, psia P, psia RE x 10- 6 , ft- 1
8 100 1250 0.5 0.010 0.58 205 1250 1.0 0.02 1.08 .325 1300 1.5 0.035 1.5
A test summary showing the configurations tested and the variables
for each is presented in Table 6.
3.2 TEST PROCEDURE
3.2.1 General
In the VKF continuous flow wind tunnels (A, B, C), the model is
mounted on a sting support mechanism in an installation tank directly
underneath the tunnel test section. The tank is separated from the
tunnel by a pair of fairing doors and a safety door. When closed, the
fairing doors, except for a slot for the pitch sector, cover the openingto the Lank and the safety door seals the tunnel from the tank area.
After the model is prepared for a data run, the personnel access door to
8
the installation tank is closed, the tank is vented to the tunnel flow,
the safety and fairing doors are opened, and the model is injected into
the airstream. After the data are obtained, the model is retracted into
the tank and the sequence is reversed with the tank being vented toatmosphere to allow access to the model in preparation for the next run.
A given injection cycle is termed a run, and all the data obtained are
identified in the data tabulations by a run number.
3.2.2 Thin-Skin Thermocouple
Prior to each test run, the model temperatures were monitored toensure that the model was nominally isothermal. The model was then in-jected at the desired test attitude as the data acquisition sequencecommenced. The model remained on the tunnel centerline for about threeseconds and was then retracted into the installation tank. The modelwas then cooled while being repositioned for the next injection.
A 256 channel multiplexing analog-to-digital converter was used inconjunction with a Digital Equipment Corporation (DEC) PDP-11 computerand a DEC-10 computer to record the temperature data. The system sampledthe output of each thermocouple approximately .13 times per second.
3.2.3 Oil-Flow
Oil-flow testing was done on the 60-0 model and the 83-0 model.For oil-flow testing the models were painted black for contrast, andin general, a white oil with a viscosity of 25 centistokes was appliedto the surface with a sponge for each run. The oil was applied dif-ferently on the first two runs of the 83-0 model. On runs 77 and 78,coatings of 800-centistoke and 200-centistoke white oil were appliedover a coating of clear Dow Corning oil with a viscosity of 100
centistokes. The model was positioned to the test attitude and in-jected into the tunnel flow for about 20 sec. During this time, fourstill cameras photographed the model at 2-second intervals. Locotionsof the cameras and camera numbers are specified in Table 7. After themodel was retracted from the tunnel flow, it was cooled and cleanedbefore oil was reapplied for the next test run. Oil flow runs arespecified in the Test Data Summary, Table 6.
3.3 DATA REDUCTION
3.3.1 Thin-Skin Thermocouple Data
The reduction of thin skin temperature data to coefficient form
normally involves only the calorimeter heat balance for the thin skinas follows:
QDOT - pbc DTW/DT (1)
H(TR) = QDOT = pbc DTW/DT (2)TR-TW TR-TW
Thermal radiation and heat conduction effects on the thin-skinelement are neglected in the above relationship and the skin temperature
9
response is assumed to be due to convective heating only. It can beshown that for constant TR, the following relationship is true:
d n TR-TI _ DTW/DT (3)Tt_ R-TW TR-TW
Substituting Eq. (3) in Eq. (2) and rearranging terms yields:
H(TR) = d in [TR-TI (4)pbc dt TRTW
By assuming that the value of H(TR)/pbc is a constant, one can seethat the derivative (or slope) must also be constant. Hence, the term
in tTRTI]
is linear with time. This linearity assumes i:he validity of Eq. (2)which applies for convective heating only. The evaluation of conductioneffects will be discussed later.
The assumption that H(TR) ard c are constant are reasonable for thistest although small variations do occur in these parameters. The varia-tions of H(TR) caused by changing wall temperature and by transitionmovement with wall temperature are trivial for the small wall temperaturechanges that occur during data reduction. The value of the model materialspecific heat, c, was computed by the relation
c = 0.0797 + (5.556 x 10- 5)TW (17-4 PH stainless steel) (5)
The maximum variation of c over any curve fit was less than 1.5 percent.Thus, the assumption of constant c used to derive Equation 4 was reason-
able. The value of density used for the 17-4 PH stainless steel skinwas p = 490 lbm/ft3 , and the skin thickness, b, for each thermocoupleis listed in Tables 3, 4 or 5.
The right side of Equation 4 was evaluated using a linear least
L- squares curve fit of 15 consecutive data points to determine the slope.
The curve [it was started at approximately the time the model arrivedon the tunnel centerline. For each thermocouple the tabulated value ofH(TR) was calculated from the slope and the appropriate values of pbc;
' i.e.,
d [T-T lH(TR) pbc d [ in [TR-TW (6)
10
To investigate conduction effects a second value of H(TR) was calculated
at a time one second later. A comparison of these two values was usedto identify those thermocouples that were influenced by significant con-
duction (or system noise). The data for a given thermocouple were deleted*if these values of H(TR) differed by more than 35 percent. In general,
conduction and/or noise effects were found to be negligible.
Since the value of TR is not known at each thermocouple locationit has become standard procedure to use three assumed values of TR.The assumed values are 1.OTT, 0.9TT and 0.85TT. The use of theseassumed values of TR provides an indication of the sensitivity of theheat-transfer coefficients to the value of TR assumed. As can be
noted in the tabulated data, there are large percentage differencesin the values of the heat-transfer coefficients calculated from thethree assumed values. Therefore, if the data are to be used forflight predictions, the value selected for TR is obviously very
important and is a function of model location and boundary layerstate.
The heat-transfer coefficient calculated from Eq. 4 was normalized
using the Fay-Riddell stagnation point coefficient, H(REF), based on anose radius of 1.0 ft full scale (see Appendix III). The referencenose radius, RN, used to calculate H(REF) is either 0.0175 ft or 0.04ft as determined by the model scale.
3.4 UNCERTAINTY OF MEASUREMENTS
In general, instrumentation calibrations and data uncertaintyestimates were made using methods recognized by the National Bureau of
Standards (NBS). Measurement uncertainty is a combination of bias andprecision errors defined as:
U = ± (B = t 9 5S )
where B is the bias limit, S is the sample standard deviation and t?5 isthe 95th percentile point for the two-tailed Student's "t" distribu ion(95-percent confidence interval), which for sample sizes greater than 30
is taken equal to 2.
Estimates of the measured data uncertainties for this test aregiven in Table 2a. The data uncertainties for the measurements are
determined from in-place calibrations through the data recording systemand data reduction program.
* Propagation of the bias and precision errors of measured data
through the calculated data was made in accordance with Ref. 2 and
the results are given in Table 2b.
?*
The word DELETE is used on the tabulated data to identify these thermo-
couples.
4.0 DATA PACKAGE PRESENTATION
Heat-transfer coefficients were obtained at selected :ocations onthe 56-0, 60-0, and 83-0 models of the space shuttle orbiter. Sampletabulated data are presented in Appendix IV.
Representative data from the upper centerline (PHI = 180 deg) ofthe 83-0 model are presented in Fig. 13. Data from two runs are pre-sented as a sample of data repeatability.
Representative oil-flow data of the 60-0 model are shown in Fig.14.
REFERENCES
1. Test Facilities Handboc], (Eleventh Edition). "von Karman GasDynamics Facility, Vol. 3." Arnold Engineering Development Center,June 1979.
2. Thompson, J. W. and Abernethy, R. B. et al. "Handbook Uncertaintyin Gas Turbine Measurements." AEDC-TR-73-5 (AD755356), February1973.
1
12
APPENDIX I
ILLUSTRATIONS
1 13
Screen ~ ~ ~~ b Tunneln NozeScinTest Sect ionifuerScto
OeaigF ig. Tant Tunnel en
ILWI
I
L116L
0C)
'-1
0 c
-f-4
1-44
00
IDr
(116
CA
00-0 00
U) 4
'-44
4 I4
'00.A-4
gooC--4
P4 410 I03
0 0 0
0 44
r O
U):1n
000
17 C
H II
S1 cc~
C
IC
C
V
18
co
0 a
01-4
4J a
I00
00
I e.J
x 001
VI
I7
I2
a)o _ v j
-S-o
o) Co
00
-4-
J-o
0
U)
. ......
0)P4 4
04J
41
o 4)
00 0
0
41
cnn
C4J
220
0
rq 0)
*In 0
0) (A-40
0
23 4U
00
rI C-,
TO a0 0 = o
ate(Y - )IT) u**
g2.~~~1 12, ..- *
0
24
44
00
qN44
VIAI
\34 \bNo )'1
IB
25s.
t.1 W
ON c~IN < <
044
4,9
ONN
N 00<~ .
InI
4
N
00
-i - -. -C - - -- --
60,
ec, Q
-. "~Q r-~ - *et:
0 00~
a)
020
oInN 44
< 4-
- - - - - - - - -
I' r27
a OD 0 0O
0% 00
0~0
00 o(NJ
cu0 00
'a ~ca -
0 0 c 0
N o
-.000
-.0 r 0
'NJY m "
0 ~~J 90
28 NJ C
0 0
* 0 0
4
_.* . .
290
000
00.
-0 -0 j
0177 178 25
79
3 '
0 2Z. 2 I 2 3183 00
123+ _6 2 329'r- 6i e S62 7
22- 3230
0 97 20 !2100 2P7 0
002~ 20
0
--- --- -
00
4Cd
r 9 . 61,Z **no8
.9 z
0
4
0)
0 0
C4
C1 uI E 0 C,4
0
E-4~ -
0 0
-----------
31~
Forward Facin; Upward Facing
RCS Nozzles RCS Nozzle (Typ)
Center79 382 384 (cfii 38 5 386 387..... C- -- e ...... -e ... *(F +- e.......----
~161 162 16X.!/ 164 165 166
0 ® 0 0 0167 168 169 170 171 172 Pilots
_______-'-)- - - 0
- ...... ~r. '_
173 174 175 176
Enlargement
Nose View of-.--- ?A83-0 Model
'477 45051
Enlargement .29\3
474 449
303473 2
'7 * 281 (X/l,=0.025)
287(X/L=0.05)
290(X/L=0.2)
Centerline
c. Thermocouple Locations on Upper RCS Nozzles and
Lower Fuselage
Figure 12. Continued
32
*1!
CEl)
44
g. V
-HlIn
'-4
C4.
333
:j-4 0
V3 -r40 OD0o r-4-.H
Cl 0 0 -4r0 z w
0-I I~
En c
'~CC
IEM-ViI
oo ~~ ~ ~ ~ ~ ~ ~ ~ d I-DM-- n Cj1 0 o.: 3, c 10L n
34l
ConitRn:e = 8.x0 6 t
I Camera 3
Fitguro 1. Oil-Flow Photographs on n3O-O Model (Run 2632)
APPENDIX II
TABLES
36
to V
I I .0.
-0 0
-II
4 '
06
~0
1:1 1, 04 ; .4
a . 4.1) 0
'N 1( C) 0
I.. 0
0 Ccf O C'J 4 cl M
C. h
W %4
69 0
11) 4.
00 W
:a z .- o' A
0) 0
.4 Il 4
w 06N600
V 'AO CIO C
-C =l tav C -Cr4.. C- Q 94 a4 a
C, to "
A 0 400 '
4.,0
00 0 M L1 QA -4.4 'A.. C , " 2' C, , m -0
0 o9o 0 -
- 0 'n4
o0 ,,Cc 04-3 410. 0.;_ 2 4) -3
0~.u 0 L40 C, r2 S..-5'uo 0) 04) .. 0 0 0
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APPENDIX III
REFERENCE HEAT-TRANSFER COEFFICIENTS
In presenting heat-transfer coefficient results it is convenient
to use reference coefficients to normalize the data. Equilibrium stag-nation point values derived from the work of Fay and Riddell* were used
to normalize the data obtained in this test. These reference coefficients
are given by:
1/2 0.4 0.25
8.17173(PT2)1(MUTT)0[1 - p ] [0.2235 + (1.35 x 10 )(TT+560)]H(REF) =PT
(RN) /2(TT)0.5
and -
H (REF)STFR =(RHO)(V) [0.2235 + (1.35 x 10-5 )(TT + 560)]
where
PT2 Stagnation pressure downstream of anormal shock wave, psia
MUTT Air viscosity based on TT, lb -sec/ft 2
f
P Free-stream pressure, psia
TT Tunnel stilling chamber temperature, *R
RN Reference nose radius, (0.0175 ft or
0.04 ft determined by model scale)
3RHO Free-stream density, Ibm/.t
V Free-stream velocity, ft/sec
-
Fay, J. A. and Riddell, F. R. "Theory of Stagnation Point Heat Transfer
in Dissociated Air," Journal of the Aeronautical Sciences, Vol. 25, No.2, February 1958.
51
APPENDIX IV
SAM4PLE TABULATED DATA
ii 52
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A ~.43 353
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