F-15B/Flight Test Fixture 2: A Test Bed for Flight Research

APPENDIX A

STRUCTURAL LOAD CONSIDERATIONS

FOR ALTERNATE CONFIGURATIONS

Major configuration changes to the Flight Test Fixture II (FTF-II) could have an impact on FTF-II loads and the F-15B airplane, depending on a variety of factors. The four load cells located at the sway braces (fig. 4(b)) can be used to validate side loads for rapid envelope expansion of major configuration changes. When needed, in-flight loads can be monitored and compared to prior loads documented in this report for the baseline fin-like configuration.

Flat-plate approximations were used to estimate FTF-II side loads in terms of aircraft limits, where side force, Fy, is approximated by (sin _)*qo_*As.9 Ground tests of the FTF-II mounted to the F-15B pylon (fig. A-I) were then conducted to correlate known applied side loads on the fixture to the measured resultant loads at the sway braces.

By subtracting initial preloads (500-1000 lbf) at each sway pad and combining sway pad loads for each side, sway pad reaction loads, Fsw, were related directly to the applied moment, Mx, on the fixture. Figure A-2 shows this relationship was quite predictable and repeatable over several load cycles, where Fsw = 39.4 + 0.385"M x.

Initial load assumptions, estimates, and factors of safety for the FTF-II design were quite conservative. 9 As a result, a design limit equivalent to M x = 48,368 in-lbf (combined inertial and aerodynamic side loads 9) was used for analysis of the fixture. Initial estimates were that an M x equal to 48,368 in-lbf would result in an Fsw equal to 23,884 lbf. Determining actual loads using a linear extrapolation of ground test data indicates that the design limit Mx = 48,368 in-lbf would generate Fsw equal to 18,661 lbf. The difference between estimated and actual reaction loads is attributed to the indeterminate nature of the suspension lug/sway brace mounting structure. This result was also observed in the initial loads analysis and ground testing of a fixture test article. 9 The relevant aspect of this eval- uation with respect to a flight experiment is that in-flight side loads that exceed the design limit of the FTF-II will indicate Fsw > 18,661 lbf.

960782 Figure A-1. Load testing of the FrF-II mounted to the F-15B pylon.

15,000 ---4-- Baseline from ground calibration o Flat-plate approximation, flight o o Surface pressures, flight o o 0 • 10,000 D O0 M x , in-lbf 0 aid 0 FSW = " " * 5,000

t

I I 1 I I

0 1000 2000 3000 4000 5000 Fsw,lbf 960783 Figure A-2. The FTF-II reaction loads at the sway braces, flight- and ground-test data.

Further in-flight load evaluations were performed to gain additional confidence in estimated and actual loads.

Integrated surface pressures and the flat-plate approximation using z/b' = 0.50 were calculated for the matrix of flight conditions. These techniques (fig. A-2) show reasonable agreement with each other but tend to overestimate flight loads by more than 50 percent for a given reaction load at the load cells. At the same time, flight data (fig. A-2) also reveal that existing F-15B control system limits, as observed from flight data, limit aircraft aerodynamic loads to less than 30 percent of FTF-II design limit side loads. These data imply that aerodynamic loads for any reasonable FTF-II configuration are fairly insignificant with respect to exceeding the design load. The primary area of future concern would be to avoid very dynamic maneuvers where high inertial loads in the lateral axis can be coupled with aerodynamic loads. As a simple and conservative rule, inspection of the FTF-II should be conducted if the total Fsw > 10,000 lbf (greater than 53.6 percent of FrF-II design limit side load) during flight test.

Major configuration changes to the FTF-II could also have an impact on the FIT-II center of pressure and the resulting distribution of loads on the fixture. Initial side force estimations for the baseline fin-like configuration 9 used a center of pressure of x/c = 0.15 to x/c = 0.66 for the estimation of aerodynamic loading. Further in-flight load evaluations at the matrix of flight conditions were performed using integrated surface pressures to document the actual FFF-II center of pressure.

The flight data (fig. A-3) indicate that the center of pressure estimations were valid, over a range in Mach numbers of 0.50 to 1.28, for an expected range of x/c = 0.15 to 0.66. 9 At subsonic speeds, the center of pressure ranges between 0.15 < x/c < 0.40 with the farther aft locations at sideslip conditions where [3 > 6 °. Limited sideslip data at Mach 1.28 show the general trend for the center of pressure to move slightly aft at supersonic speeds as well.

Extrapolation of the flight results for the baseline fin-like configuration shows that development of a flight experi- ment requires only minimal consideration for the center of pressure location. In general, only major configuration changes or extensions of the nose section that would move the overall FTF-II center of pressure forward of the current x/c = 0.15 location would require further analysis.

1.4 m o z/b' = 0.16 [] z/b' = 0.48 1.3 -- 0 z/b' = 0.86 (1) Expected range for 1.2 ]dc (subsonic with 13= 0°) x/c increases as: 1.1 (1) Supersonlc _0 o (2) 13increases 1.0 Mach .9 number .8 • 7 B ,6 di>'_P _4° (2) .5

I I I I

.4 I 0 .1 .2 .3 .4 .5 xlc 960784 Figure A-3. Center of pressure locations determined from flight data as a function Mach number.

APPENDIX B

AIRCRAFT STABILITY CONSIDERATIONS

FOR ALTERNATE CONFIGURATIONS

Major configuration changes to the FTF-II could have an impact on aircraft stability, depending on the size and location of the configuration change. Initial envelope expansion of the F-15B/FTF-II configuration (vertical fin) consisted of flight test of the FTF-II uninstrumented on the F-15B aircraft to perform envelope expansion 9 to a max- imum of Mach 1.3 in the vertical fin configuration. Directional stability during phase one flight tests was evaluated by performing rudder doublets and making qualitative comparisons to the F-15B aircraft with and without a center- line tank. Comparisons at subsonic speeds indicated that the subsonic static directional stability, Cnf _, derivative lev- els for the F-15B/P'TF-II configuration were greater than 0.002/deg and had no adverse effects on aircraft stability. 9,10 Predictions also indicated that the Cnf _ would decrease with increasing supersonic Mach numbers and approach levels as low as approximately 0.0005/deg at Mach > 1.5. During phase two envelope expansion of the F-15B/FTF-U to Mach 1.8, several actions were taken by the project to understand and minimize the potential risk of reduced Cn_ caused by positioning the FIT-II at the centerline of the F-15B aircraft.

A study was performed to assess the risk of F-15 control augmentation system (CAS) disengagements and their impact on flight safety at reduced Cnfy The CAS disengagements were found to occur approximately once in every 200 flight hours and mostly during dynamic maneuvering. The CAS disengagements produced transients that were easily tolerated by the pilot and aircraft. In the case of a CAS failure, a 600-KCAS airspeed limit is imposed because of a mildly divergent longitudinal stability. In addition, a Mach 2.0 limit exists for the two-seat F- 15 aircraft because pilots comments were that uncomfortable amounts of sideslip are produced. Overall, CAS disengagements at high speeds were found to result in handling qualities undesirable for sustained operations but quite acceptable for safe aircraft maneuvering.

An analysis of F-15 lateral--directional stability without the CAS, using a NASA Dryden batch simulation, was also performed for airspeeds to a maximum of Mach 2.2 at an altitude of 45,000 ft with particular attention to the dutch roll mode. Three linear models of an F-15 aircraft at a gross weight of 36,200 Ibm were analyzed: one with predicted lateral-directional aerodynamics, one with aerodynamic increments caused by the lZTF-II added, and one with reduced increments to match Air Force flight test data. 10 From these models, the damping ratio, 4, and un- damped natural frequency, m n, were calculated for the dutch roll mode and compared to MIL-F-8785C for Class IV aircraft (fighter) category B (cruise) handling qualities requirements. As figure B-1 shows, the addition of the FI'F-II to the F-15B aircraft improved dutch roll damping at an altitude of 30,000 ft with sufficient stability at all flight conditions as predicted. Data at higher altitudes were also calculated and included with comparisons to flight test data.

F-15B/FTF-II envelope expansion was performed with rudder doublets to a maximum of Mach 1.8 with the CAS on and off for pilot evaluations and comparison to model predictions. Satisfactory-to-good damping was re- ported by the pilots with the CAS on or off at airspeeds to a maximum of Mach 1.8. Acceptable, but low-to- moderate, damping was noted at Mach> 1.5. With the CAS off at Mach > 1.5, pilot comments concurred with pre- dictions of handling qualities undesirable for sustained operations but quite acceptable for safe aircraft recovery.

Flight data from the rudder doublets were also analyzed to calculate _*c0 n for the dutch roll mode for comparison to predicted values (without the FI'F-II, CAS off) and MIL-F-8785C (fig. B-2). Figure B-2 shows the flight data with the FTF-II installed and the CAS off. The F-15 airplane with the FTF-II installed appears to be more stable than predictions of an F-15 aircraft without P-TF-II. As expected, _*e0 n decreases with increasing altitude. At Mach 1.8 and an altitude of 45,000 ft, _*m n is approximately 0.25 rad/sec, which is in the lower region of Level 3 handling qualities.

Thefinal phase of F-15B/FTF-IIenvelope expansion wasperformed with rudder doublets to a maximum of Mach2.0with theCASon.Maneuvers withtheCASoff werenotattempted at speeds greater thanMach1.8.No large reductions in stabilitywere observed, and pilotcomments indicated thathandling qualities wereacceptable for safe aircraftmaneuvering. Predicted trends indicate thattheF-15/F'IT-IIdutchroll modewiththeCASonor off is stable atMach_< 2.2.Additional evaluation of aircraftstabilityby performing rudder doublets with theCASon is recommended for significant changes to theFTF-II or if envelope expansion of thecurrentFTF-II configuration beyond Mach2.0atanaltitude of 45,000 ft isdesired.

/ ' ; : : : : : ' ' :/ ; : : , , __ , I, / ,FTF-Ilon--_ , , --H .... : .... _ - -,--,- - -"-T'-_-_-- - :-- - - : .... ;--A_ .... :---- -'- .... , ....

i ! /"_/MI'L-F-8785C i ! i//_ !

...: .... __..:. Level 2 --;---:---_'--,;,,_-TL---+ .... "--- .5 : : : handling : i /.,,_.' .,,"Z. ' i ! : : : : qualities : _...j1- ', : : : "i .... :: .... ....i....

•4 "'-i .... _iTL ! --I_;e; 3 i .... i---i .... i _* tOn .... i ---L- i i : /, handling : : : _.

" , --, .... , -"" T .... '- -1- '_ - - qualities ..... '_.... ?-; rad/iec ....... ..

' ' .- __ ' .)/ : (Category B) :. -" : .3 ........ ,--.,"f .... , ......... , .... r---7 .... ,----i .... ,----'T .... , ....

........ "_.." .... ;.. ........ : ............. :.._.,;K ...............

, , ', ; ,_,,: : ,,:1 , ', ,, : .... : .... _ .... : ......... : .... __.._,._,,_____._ .... L .............

.2 .... i i i i i..": i i ! : ---" .... ;-- -- ;-" -" _ :- :, -- --',_ - - ,_ - - - _,.... ,_--I .... _--; .... _--- 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 Mach number 960?85 Figure B-1. The F-15B dutch roll mode damping obtained from NASA Dryden batch simulation (GAS off) with and without FTF-II; Hp = 30,000 feet, aircraft weight of 36,200 Ibm.

i i ! , i/i i i/ i i i ---i ........ i:,-,.-.'------_--'-¥---_---_ .... i--/ .... "---i .... i....

' ' ' ', MIL-F-8785C ', / : ' ' ' ' /'_ ' . ' 30kft : ,/,,_',.,. ;._ :Level2handhng :/ : !j,_, , .5 ---_ .... ;--,_v.,_,[-./_---T-- qualities --J --- _. --- _ .... Y--- i i i / i i (Category B)/'.j_ Level 3 i ---! .... !-'-]_---! .... t .... ! .... !-7_'-- !" handling "-! ....

', ', /_ !/._ i : qualities ', t tOn .......... L__.: .... ;.._., .... L._[k _ • .... L---*' .... '_ - ---_- L0-k -/1--,, rad/sec ...... .' ...... - _ "f__. _: ....

: :"__--o_i ;/i : , _ .... _..__ .... .....,____ ........ :.-___,__..: , ---_ ...... • .... _- --_ .... , .... %;,.;.;.1, .... _ ---p .... _-- • ---_--- , , 4 ',J , , , _ q , Jq .2 .... _ .... L.__, .... ,.... _.... .... :___Jt : ; : : : .. _" - - : -.-0-- Flight, 30k/1 .... i .... ;---; .... _ - - - ; - - - -',_-_'- _- -- - -_ .... _. '--I_--" Flight, 45k ft ..... Dryden batch " : " : : _ ' , ', limulaUon 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 Mach number 960786 Figure B-2. Comparison of predicted and flight test values for F-15B dutch roll mode damping, CAS off, FI'F-II on for flight, FTF-II on for simulation.

APPENDIX C

CHORDWISE PRESSURE DISTRIBUTION DATA

Table C-1. Steady-state chotdwise pressure distribution data obtained in flight, Hp - 15,000 ft.

F.r_-$tream Mach number z/b" x/c 0.39 0.52 0.61 0.70 0.80 0.90 0.98 0.16 0.000 1.010 0.990 1.040 1.095 IAI4 1.204 0.980 0.025 -0.102 -0.112 -0.193 -0.248 -0.222 -0.108 0.000 0.050 -0183 -0.195 --0,285 --0.361 -0.370 -0.298 -0.169 0.075 -0,197 -0.210 -0.305 -0,388 -0,413 -0,365 -0,241 0.I00 -0.216 -0.227 -0.320 -0.404 -0.437 -0.416 -0.290 0.125 -0.202 --0.206 -0.299 -0.383 -0.413 -0,415 -0.297 0.150 -0.154 -0.165 -0.253 -0.333 -0.360 -0.483 --0.354 0.210 -0.086 --0.102 -0.184 -0.255 -0.274 -0,290 -0.203 0.260 -0.113 -0.131 -0,217 -0.302 -0.353 -0.303 -0.335 0.320 -0.088 --0.I01 -0.179 -0.257 -0.306 -0.389 -0.357 0.370 -0.072 -0.081 -0.155 -0.218 -0.24.4 -0.481 -0.410 0.420 -0.055 -0.069 -0.137 -0.197 -0,214 -0.410 -0,381 0.470 -0.065 -0.070 -0.139 -0.195 -0.212 -0.240 -0.422 0.540 -0.053 -0.061 -0.126 -0.181 -0.195 -0.182 -0.346 0.600 -0.036 -0.037 -0.102 -0.149 -0.151 -0.132 -0.220 0.700 -0.019 -0,018 -0.075 -0.116 -0.117 -0.093 -0.115 0.800 -0035 -0.029 -0.084 -0.125 -0.126 -0.100 -0.111 0.900 -0048 -0.040 -0.091 -0.131 -0.127 -0.091 -0.069 0960 -0.180 -0.178 -0.228 -0,318 -0.377 -0.282 -0.212 0.000 10l 9 1.007 1,040 1.096 I. 111 1.203 0.930 0.48 0.025 -0.119 -0,126 -0.209 -0.261 -0.225 -0.089 0.010 0.050 -0.225 -0.234 -0.330 -0,417 -0.427 -0.327 -0.205 0.075 -0242 -0.259 -0.366 -0.471 -0,524 -0.460 -0.329 0.I00 -0.248 -0.26,1 -0371 -0.477 -0.544 -0.576 -0.438 O. 125 -0.257 -0.265 -0.367 -0.464 -0.514 -0.656 -0.516 0.150 -0.196 -0.205 -0.293 --0.377 -0399 -0.654 -0.535 0.210 -0.088 -0.096 -0.168 --0.228 -0.243 -0.123 -0.341 0.260 -0.061 -0.069 -0.139 -0.202 -0230 -0.173 --0.173 0.320 -0.084 -0.079 -0.139 -0.198 -0.222 -0.277 -0.269 0.370 -0.043 -0.047 -0.114 -0.173 -0.192 -0.352 -0.270 0.420 -0.037 -0.041 -0.104 -0.156 -0.173 -0.361 -0.338 0,470 -0036 -0,037 -0.096 -0.150 -0.162 -0.144 -0.329 0540 -0,031 -0.033 -0.094 -0.141 -0.154 --0.143 --0.381 0.600 -0.030 -0.026 -0.081 -0.127 -0.131 -.0.115 -0.221 0.700 -0.030 -0.022 -0.077 -0.117 -0.117 -0.099 -0.133 0800 -0.032 -0.027 -0.077 -0.117 -0. t17 -0.094 -0.120 0.900 -0.063 -0.058 -0.112 -0.152 -0.149 -0.106 -0.079 0.960 -0106 -0.101 -0.167 -0.223 -0.239 -0.189 -0.185 0.000 0.980 0.970 1.010 1,062 1.074 1.157 0910 0.86 0,025 -0109 -0.131 -0.224 -0.300 -0.296 -0.182 -0.057 0.050 -0.195 -0.213 -0.312 -0.409 --0452 -0.425 -0.287 0.075 -0.212 -0.227 -0.327 -0.426 -0.485 -0.482 -0356 0.100 -0.210 -0.229 -0,325 -0.423 -0487 -0.560 -0433 0.125 -0.195 -0.197 -0.286 -0.375 -0.407 -0.611 -0.481 0.150 -0 153 -0.169 -0.254 -0.328 -0.351 -0.647 -0.523 0.210 4) 056 -0.065 -0.135 -0.190 -0.203 -0,084 -0.323 0.260 -0040 -0.043 -0.108 -0.160 -0.177 -0.143 -0117 0,320 -0.018 -0.028 -0.094 -0.147 -0.164 -0.206 -0.151 0.370 -0037 -0.037 -0.093 -0,145 -0.159 -0.242 -0203 0420 -0027 -.0.023 -0.084 -0.132 -0.144 -0.271 -0.208 0470 -0.023 -0.019 -0.076 -0.124 --0137 -0.146 -0.268 0.540 -0.020 -0.020 -0,078 -0.126 -0.133 -0.128 -0.276 0.600 -0.021 -0.016 -0.070 -0.113 -0120 -0.111 -0.306 0700 -0.019 -0.016 -0.075 -0117 -0.123 -0.110 -0.144 0.800 -0.023 -0.016 -0.073 -0.113 -0.114 -0.096 -0.108 0.900 -0.027 -0,027 -0.077 -0.107 -0,089 -0.010 0.020 0960 4).148 -0.125 -0.185 -0.231 -0.228 -0.177 -0.151 Table C-2. Steady-state chordwise pressure distribution data obtained ia flight, Hp _ 30,000 ft.

Free-stream Mach number z/b" x/c 0.51 0.62 0.70 0.79 0.89 1.00 1.05 1.15 1.28 1.39 1.48 0.16 0.000 1.005 1.045 1.092 1.132 1.207 1.202 1.246 1.330 1.307 1.149 !.045 0.025 --0.075 --0.162 -0.190 --0.186 -0.091 0.020 0.090 0.200 0.280 0.250 0.280 0.050 -0.172 -0.265 -0.315 -0.339 -0.277 -0.153 --0.098 -0.002 0.090 0.030 0.070 0.075 -0.183 -0.282 -0.342 -0.378 -0.335 -0.211 -0.156 -0.090 0.000 -0.072 -0.040 0.100 -0.196 -0.301 --0.357 -0.398 -0.394 -0.263 -0.206 -0.158 -0.071 -0.167 -0.123 0.125 -0.181 -0.279 -0.343 -0.379 -0.391 -0.269 -0.204 -0.147 -0.061 -0.218 -0.187 0.150 -0.136 -0.232 -0.289 -0.323 -0.448 -0.316 -0.243 -0.163 -0.065 -0.138 -0.229 0.210 -0.070 -0.161 -0.215 -0.243 -0.172 -0.168 -0.097 -0.010 0.090 0.090 0.050 0.260 -0. i 05 -0.200 -0.264 -0.313 -0.280 -0.311 -0.258 -0.204 -0. ! 00 -0.051 -0.042 0.320 -0.075 -0.166 -0.225 -0.262 -0.369 -0.360 -0.315 -0.309 -0.213 -0.153 -0.113 0.370 -0.057 -0.133 -0.183 -0.211 -0.289 -0.387 -0.312 -0.291 -0.199 -0.167 --0.127 0.420 -0.047 -0.123 -0.169 -0.186 -0.154 -0.350 -0.308 -0.265 -0.234 -0.227 -0.182 0.470 -0.050 -0.124 -0.169 -0.184 -0.156 --0.369 -0.313 -0.255 -0.213 -0.194 -0.200 0.540 -0.046 -0.116 -0.159 -0.172 -0.145 -0.313 -0.357 -0.354 -0.268 -0.237 -0.213 0.600 -0.020 -0.085 -0.122 -0.132 -0.093 -0.164 -0.215 -0.358 -0.287 -0.260 -0.229 0.700 -0.002 -0.067 -0.096 -0.098 -0.054 -0.070 -0.040 -0.127 -0.283 -0.324 -0.297 0.800 -0.024 -0.079 -0.108 -0.108 -0.063 -0.074 -0.034 -0.006 -0.078 -0.213 -0.204 0.900 -0.031 -0.083 -0.116 -0.109 -0.058 -0.035 0.010 0.030 0.010 -0.101 -0.243 0.960 -0.138 -0.221 -0.287 -0.339 -0.275 -0.171 -0.090 0.010 0.090 0.040 -0.006 0.48 0.000 1.011 1.048 1.094 1.131 1.203 1.198 1.251 1.327 1.252 1.102 1.004 0.025 -0.096 -0.178 -0.203 -0.189 -0.073 0.040 0.120 0.220 0.320 0.330 0.360 0.050 -0.206 -0.308 -0.367 -0.398 -0.315 -0.177 -0.104 0.000 0.100 0.110 0.150 0.075 -0.231 -0.339 -0.418 -0.479 -0.436 -0.295 -0.223 -0.120 -0.024 -0.025 0.010 0.100 -0.231 -0.347 -0.427 -0.501 -0.541 -0.398 -0.322 -0.229 -0.127 -0.132 -0.092 0.125 -0.237 -0.346 -0.420 -0.475 -0.614 -0.478 -0.405 -0.326 -0.215 -0.225 -0.184 0.150 -0.181 -0.274 -0.338 -0.374 -0.602 -0.496 -0.423 -0.374 -0.262 -0.274 -0.238 0.210 -0.079 -0.149 -0.199 -0.216 -0.122 -0.319 -0.270 -0.256 -0.185 -0.308 -0.283 0.260 -0.046 -0.126 -0.171 -0.199 -0.183 -0.132 -0.069 -0.105 -0.060 -0.077 -0.212 0.320 -0.070 -0.139 -0.179 -0.201 -0.252 -0.244 -0.154 -0.095 -0.056 -0.064 -0.079 0.370 -0.031 -0.103 -0.144 -0.163 -0.198 -0.232 -0.183 -0.155 -0.085 -0.056 -0.048 0.420 -0.021 -0.090 -0.129 -0.1 44 -0.122 -0.300 -0.225 -0.157 -0.074 -0.076 -0.043 0.470 -0.011 -0.086 -0.127 -0.139 -0.112 -0.273 -0.217 -0.202 -0.128 -0.102 -0.082 0.540 -0.017 -0.081 -0.118 -0.129 -0.101 -0.336 -0.280 -0.195 -0.158 -0.140 -0.108 0.600 -0.016 -0.073 -0.107 -0.112 -0.077 -0.141 -0.318 -0.246 -0.185 -0.163 -0.140 0.700 -0.005 -0.069 -0.097 -0.100 -0,058 -0.080 -0.054 -0.307 -0.218 -0.197 -0.205 0.800 -0.012 -0.069 -0.099 -0.098 -0.055 -0.074 -0.023 --0.013 -0.235 -0.235 -0.228 0.900 -0.050 -0.105 -0.130 -0. ! 27 -0.078 -0,036 0.030 0.070 0.000 -0.158 -0.214 0.960 -0.087 -0.160 -0.202 -0.216 -0.170 -0.134 -0.061 0.050 0.090 -0.085 -0.143 0.86 0.000 0.970 1.012 1.059 1.093 1.167 1.173 1.178 1.270 1.224 1.079 0.980 0.025 -0.090 -0.187 -0.239 -0.253 -0.159 -0.025 0.030 0.150 0.250 0.270 0.300 0.050 -0.186 -0.288 -0.359 -0.409 -0.388 -0.253 -0.175 -0.068 0.020 0.040 0.080 0.075 -0.207 -0.308 -0.384 -0.446 -0.451 -0.320 -0.242 -0.170 -0.084 -0.057 -0.017 0.100 -0.201 -0.308 -0.382 -0.449 -0.525 -0.393 -0.317 -0.246 -0.166 -0.168 -0.125 0.125 -0.186 -0.270 -0.339 -0.384 -0.558 -0.439 -0.365 -0.299 -0.218 -0.223 -0.182 0.150 -0.151 -0.242 -0.300 -0.331 -0.569 -0.482 -0.408 -0.350 -0.266 -0.271 -0.238 0.210 -0.043 -0.120 -0.163 -0.177 -0.119 -0.290 -0.249 -0.256 -0.180 -0.199 -0.208 0.260 -0.030 -0.102 -0.141 -0.156 -0.138 -0.082 -0.017 -0.165 -0,152 -0.191 --0.190 0.320 -0.007 -0.078 -0.122 -0.139 -0.149 -0.128 -0.072 -0.021 0.010 -0.126 -0.205 0.370 -0.026 -0.091 -0.125 -0.139 -0.138 -0.175 -0.113 -0.075 -0.008 0.010 -0.040 0.420 -0.013 -0,075 -0.110 -0.123 -0,106 -0.173 -0.117 -0.101 -0,034 -0,013 0.010 0.470 -0.005 -0.069 -0.104 -0.115 -0.091 -0.231 -0,162 -0.087 -0.067 -0.065 -0.030 0.540 -0.010 -0.071 -0.105 -0.113 -0.085 -0.235 -0.189 -0.152 -0.087 -0.077 -0.073 0.600 -0.006 -0.065 -0.098 -0.101 -0.067 -0.260 --0.202 --0.139 -0.104 -0.100 -0.078 0.700 0.000 -0.062 -0.094 -0.099 -0.063 -0.088 -0.088 -0.207 -0.143 -0.134 -0.112 0.800 -0.011 -0.063 -0.097 -0.094 -0.055 -0.066 -0.022 -0.171 -0.175 -0.156 -0.133 0.900 -0.016 -0.072 -0.094 -0.078 -0.001 0.060 0.120 0.160 0.040 -0.026 -0.023 0.960 -0.120 -0.182 -0.215 -0.212 -0.150 -0.112 -0.055 -0.012 -0.082 -0.241 -0.248 Table C-3. Steady-state chordwise pressure distribution data obtained in flight, lip - 45,000 ft.

Free-stream Mach number z, ao• x/c 0.70 0.79 0.90 0.98 1.06 1.16 1.28 1.39 1.49 1.60 1.69 1.80 1.90 2.00 0.16 0.000 1.110 1.147 1.220 1.220 1.232 1.339 1.447 1.450 1.431 1.370 1.461 1.499 1.424 1.317 0.025 -0.202 -0.190 -0.103 -0.006 0.060 0.190 0.260 0.260 0.270 0.290 0.360 0.390 0.420 0.420 0.050 -0.334 -0.351 -0.300 -0.202 -0.138 --0.017 0.060 0.040 0.070 0.110 0.170 0.220 0.250 0.270 0.075 -0.360 -0.397 -0.366 -0.261 -0.215 -0.107 -0.019 -0.058 -0.029 0.000 0.060 0.110 0.140 0.180 0.100 -0.368 -0.409 -0.414 -0.304 -0.268 -0.176 -0.080 -0.157 -0.119 -0.076 -0.023 0.020 0.060 0.090 0.125 -0.346 -0.387 -0.413 -0.309 -0.258 -0.165 -0.068 -0.215 -0.182 -0.141 -0.089 -0.038 -0.005 0.020 0.150 -0.294 -0.328 -0.465 -0.347 -0.282 --0.177 -0.070 -0.171 -0.224 --0.192 -0.143 --0.093 -0.057 -0.024 0.210 --0.206 -0.238 -0.201 --0.182 -0.121 -0.012 0.090 0.100 0.040 -0.003 0.010 0.030 0.030 0.020 0.260 -0.261 -0.322 --0.282 -0.349 -0.321 -0.223 -0.104 -0.045 -0.042 -0.043 0.000 0.040 0.070 0.080 0.320 -0.220 -0.263 -0.382 -0.375 -0.414 -0.340 --0.222 -0.157 -0.117 --0,113 -0.056 -0.004 0.030 0.050 0.370 -0.180 -0.213 -0.479 .-0.422 -0.381 -0.311 -0.206 -0.168 -0.133 -0.121 -0.074 -0.020 0.010 0.040 0.420 -0.164 -0.188 -0.181 -0.370 -0.361 -0.288 -0.240 -0.219 -0.174 --.0.143 -0.097 -0.053 -0.018 0.010 0.470 -0.162 -0.188 -0.166 -0.382 -0.353 -0.292 --0.206 --0.188 -0.193 -0.176 -0.127 -0.092 -0.058 -0.024 0.540 -0.161 --0.173 -0.165 -0.242 -0.438 -0.390 -0.286 -0.243 -0.215 -0.219 -0.186 -0.152 -0.117 -0.080 0.600 -0.125 -0.132 --0.110 -0.119 --0.290 -0.401 -0.315 -0.267 -0.234 --0.225 -0.194 --0.163 -0.141 -0.111 0.700 -0.100 -0.100 --0.076 -0.054 -0.110 -0.167 -0.311 -0.325 -0.290 --0.258 -0.224 -0.192 --0.171 -0.165 0.800 -0.111 -0.110 -0.086 -0.050 -0.059 -0.052 -0.088 -0.222 -0.205 -0.233 -0.195 -4).t61 -0.147 -0.146 0.900 -0.125 -0.122 -0.081 -0.025 -0.024 0.000 0.000 -0.087 -0.236 -0.244 -0.209 -0.166 -0.142 -0.123 0.960 -0.314 -0.363 --0.286 -0.193 -0.124 -0.001 0.090 0.050 --0.016 -0.144 -0.110 -0.075 --0.062 -0.053 0.48 0.000 1.120 1.153 1.225 1.226 1.251 1.335 1.426 1.431 1.424 1.505 1.503 1.443 1.372 1.270 0.025 -0.220 -0.189 -0.079 0,020 0.090 0.210 0.310 0.340 0.360 0.380 0.450 0.510 0.510 0.530 0.050 -0.387 -0.405 -0.326 -0.210 -0.134 ..-0.006 0.080 0.120 0.140 0.170 0.230 0.280 0.320 0.320 0.075 -0.428 -0.491 -0.452 -0.331 -0.260 -0.130 -0.034 -0.013 0.010 0.050 0.100 0.160 0.190 0.210 0.100 -0.431 -0.513 -0.561 -0.436 -0.368 -0.237 -0.135 -0.119 -0.085 --0.052 0,000 0.050 0.090 0.120 0.125 -0.436 -0.493 -0.643 -0.523 -0.462 -0.340 -0.231 -0.217 .-O. 180 -0.134 -0.082 -0.027 0.000 0.030 0.150 -0.345 -0.377 -0.636 --0.528 -0.493 -0.380 -0.273 -0.273 --0.234 -0.193 -0.143 --0.094 -0.055 -0.026 0.210 -0.195 -0.218 -0.100 -0.339 -0.340 -0.280 -0.196 -0.322 -0.277 -0.231 -0.180 -0.132 -0.101 -0.068 0.260 -0.174 -0.202 -0.164 -0.134 -0.141 -0,126 -0.065 -0.081 --.0.229 -0.256 -0.206 -0.148 -0.122 -0.089 0.320 -0.203 -0.219 -0.293 -0.260 -0.197 -0.135 -0.074 -0.063 -0.086 -0.148 -0.111 -0.074 -0.059 -0.079 0.370 -0.153 -0.170 -0.320 -0.240 -0.266 -0.183 -0.095 -0.052 -0.045 -0.132 -0.100 -0.071 -0.050 -0.029 0.420 -0.132 -0.149 -O. 143 -0.307 -0.267 --O. 180 -0.082 -0.053 -0.024 -0.103 -0.074 --0.056 -0.043 -0.027 0.470 .-O.130 -0.143 -0.121 --0.270 -0.278 -0.235 --0.136 -0.102 -0.079 -0.082 -0.064 --0.037 -0.022 -0.015 0.540 -0.122 -0.137 -0.119 --0.295 -0.303 -0.246 -0.172 --0.133 -0.101 -0.097 -0.063 -0.057 -0.036 -0.018 0.600 -0.116 -0.117 -0.093 --0.106 -0.364 -0.277 -0.195 -0.169 -0.143 -0.125 -0.090 -0.062 -0.056 -0.034 0.700 -0.111 -0.105 -0.080 -0.049 -0.114 -0.322 --0.242 -0.196 -0.196 -0.177 -0.134 -0.105 --0.087 -0.083 0.800 -0.109 -0.102 --0.073 --0.035 -0.046 -0.064 -0.260 -0.241 -0.230 -0.220 -0.176 --0.136 -0.110 -0.106 0.900 -0,152 -0.145 -0.096 -0.024 0.000 0.040 -0.008 -0.168 -0.199 -0.231 -0.196 --0.145 -0,127 -0.110 0.960 -0.225 -0.232 -0.171 -0.114 -0,066 0.030 0.080 -0.066 -0.144 --0.178 -0.157 -0.124 --0.104 -0.090 0.86 0.000 1.091 l.t21 1.196 1.198 1.187 1.313 1.433 1.435 1.398 1.400 1.472 1.414 1.347 1.247 0.025 -0.218 -0.229 -0.148 -0.044 0.030 0.150 0.260 0.290 0.300 0.300 0.370 0.370 0.400 0.420 0.050 -0.348 -0.399 --0.383 -0.278 -0.194 -0.074 0.020 0.060 0.090 0.120 0.180 0.230 0.260 0.290 0.075 -0.374 -0.442 -0.447 -0.340 -0.278 -0.174 -0.091 -0.052 -0.017 0.010 0.070 0.120 0.150 0.180 0.100 -0.374 -0,439 -0.526 -0.414 -0.355 -0.254 -0.173 -0.151 -0.115 -0.078 -0.023 0.020 0.060 0.090 0.125 -0.345 -0.403 -0,573 -0.459 -0.409 --0.305 -0.218 -0.210 -0.177 -0,146 -0.094 -0.043 -0.013 0.010 0.150 -0.289 -0.324 -0.597 --0.497 --0.452 -0.351 -0.266 -0.259 -0.230 -0.199 -0.149 -0.098 -0.067 -0.042 0.210 --0.153 -0.173 --0.063 -0.331 -0.336 -0.260 -0.187 -0,211 -0.205 -0.204 -0.161 -0.113 -0.087 -0.070 0.260 -0.139 -0.156 -0.132 -0.087 -0.148 -0.221 -0.157 -0.173 -0.174 -0.165 -0.131 -0.094 --0.072 -0.051 0.320 -0.t13 -0.136 --0.181 -0.137 -0.126 -0.031 0.000 -0.131 -0.197 -0.171 -0.130 -0.094 -0.072 -0.046 0.370 -0.129 -0.142 --0,203 -0A82 -0.151 -0.107 -0.014 0.010 -0.055 -0.184 -0.139 -0.098 -0.078 -0.060 0.420 -0.106 -0,125 -0.136 -0.170 -0.183 -0.120 -0.041 -0.010 0.010 -0.111 -0.119 -0.100 -0.080 -0.056 0.470 -0.102 -0.116 -0.109 -0.231 -0.181 -0.150 -0.072 -0.061 -0.023 --0.076 -0.048 -0.022 -0.019 -0.041 0.540 -0.102 -0.114 -0.103 -O.215 -0.219 -0,196 -0.089 -0.075 -0.076 -0.054 -0.038 -0.020 -0.001 0.010 0.600 -0.097 -0.105 -0.088 -0.150 -0.237 -0,166 -0.121 -0.091 --0.073 -0.08l -0.044 -0.035 -0.016 0.000 0.700 -0.083 -0.093 -0.075 -0.042 -0.322 -0.218 -0.157 -0.123 -0.106 -0.110 -0.077 -0.048 --0.042 -0.027 0.800 -0.095 -0.097 -0.072 --0.032 -0.067 -0.251 -0.173 -0.154 -0.135 -0.139 -0.100 -0.060 -0.050 -0.047 0.900 --0.085 -0.072 0.000 0.080 0.100 0.130 0.010 -0.041 --0.033 -0.149 -0.058 0.000 0.010 0.030 0.960 -0.228 -0.224 -0.165 -0.101 -0.070 -0.040 -0.105 -0.246 -0.244 -0.330 -0.282 -0.181 --O. 180 -0.183 Cpmax 1.12 1.17 1.22 1.26 1.31 1.37 IA3 1.48 1.52 1.56 1.59 1.62 1.64 1.66 Table C-4. Chordwise pressure distribution data, obtained in flight, as a function of angle of sideslip, Mach = 0.52 and 0.70.

M = 0.52, Hp = 15,000 ft M = 0.70, Hp = 45,000 ft 13 8.04 3.95 0.00 -3.98 -8.00 8.02 4.01 -0.06 -3.98 -7.99 z/b" x/c 0.16 0.000 0.800 0.950 i .007 0.940 0.790 0.920 1.046 1.089 1.058 0.910 -0.749 -0.487 -0.212 0.000 0.230 0.025 --0.564 --0.359 -0.136 0.060 0.260 0.050 -0.495 --0.363 -0.213 ---0.070 0.080 --0.667 -0.503 -0.323 -0.168 0.010 0.075 -0.446 --0.339 -0.226 -0.111 0.020 -0.598 -0.487 -0.342 -0.225 -0.058 0.100 -0.433 -0,341 -0.237 -0.138 -0.022 -0.591 -0.472 -0.350 -0.248 -0.092 0.125 --.0.403 -0.310 -0.218 -0.132 -0.022 -0.566 -0.446 -0.323 -0.227 -0.086 0.150 -0.341 -0.260 -0.173 -0.094 0.000 -0.496 -0.382 -0.276 -0.186 -0.053 0.210 -0.231 --0.173 -0.104 -0.043 0.040 -0.347 -0.259 -0.190 -0.124 0.000 -0.367 -0.304 -0.231 -0.179 -0.081 0.260 -0.233 -0.186 -0.133 -0.085 -0.010 0.320 -0.181 -0,143 -0.105 -0.065 0.000 -0.291 -0.237 -0.197 -0.154 -0.061 0.370 -0.137 -0. i 16 -0.083 -0.050 0.000 -0.230 -0.201 -0.166 -0.128 -0.053 -0.196 -0.170 -0.146 -0,119 -0.053 0.420 -0.108 -0,090 -0.067 -0.043 0.010 0.470 -0. ! 01 -0.089 -0.075 -0.045 0.000 -0.183 -0.180 -0.152 -0,125 -0.072 0.540 -0.089 -0.080 -0.059 -0.036 0.000 -0.175 -0.161 -0.155 -0.126 -0.064 0.600 -0.057 -0.056 -0.037 -0.016 0.020 -0.141 -0.121 -0.114 -0.088 -0.045 0.700 -0.032 -0.026 -0.018 -0.003 0.030 -0.106 -0.096 -0.096 -0.074 -0.034 0.800 -0.031 -0.030 -0.027 -0.019 0.000 -0.118 -0.112 -0.105 -0.099 -0.064 0.900 -0.042 -0.037 -0.038 -0.033 -0.015 -0.118 -0.121 -0.120 -0.115 -0.080 0.960 -0.193 -0,189 -0.174 -0.186 -0.176 -0.318 -0.304 -0.288 -0.286 -0.273 0.48 0.000 0.780 0.950 1.007 0.940 0.750 0.930 1.061 1.096 1.053 0.880 0.025 -0.628 -0.395 -0.152 0,060 0.270 -0.852 -0.529 -0.227 0.000 0.250 0,050 -0.542 -0.409 -0.250 -0.095 0.070 -0.775 -0.575 -0.374 -0.198 0.000 0.075 -0,490 -0.390 -0.272 -0.154 -0.019 -0.677 -0.550 -0.408 -0.277 -0.098 0.100 -0.434 -0.361 -0.276 --0.182 -0.061 -0.606 -0.509 -0.405 -0.306 -0.160 0.125 -0.395 -0.347 -0.273 -0,198 -0.095 -0564 -0.491 -0.412 -0.333 -0.189 0.150 -0.307 -0.262 -0.210 -0.148 -0,063 -0.442 -0.384 -0.325 -0.264 -0.155 0.210 -0.160 -0.137 -0.098 -0.057 0.000 -0.261 -0.235 -0.193 -0.137 -0.058 0.260 -0.117 -0.095 -0.066 -0.036 0.010 -0.218 -0.190 -0.162 -0.119 -0.042 0.320 -0.115 -0.100 -0.078 -0.046 -0.001 -0.235 -0.220 -0. ! 90 -0.163 -0.088 0.370 --0.076 -0.067 -0.047 -0.023 0.020 -0.169 -0.159 -0,134 -0.105 -0.053 0.420 -0.061 -0.054 -0.042 -0.015 0.010 -0.146 -0.140 -0.120 -0.097 -0.050 0.470 -0.055 -0.046 -0.037 -0.017 0.020 -0.138 -0.124 -0.118 -0.100 -0.034 0.540 -0.042 -0.039 -0.030 -0.017 0.010 -0.130 -0.129 -0,117 -0.090 -0.053 0.600 -0.028 -0.033 -0.021 -0.010 0.020 -0.123 -0.105 -0.110 -0.088 -0.045 0.700 -0.021 -0.027 -0.019 -0.010 0.010 -0,113 -0.118 -0.100 -0.083 -0.054 0.800 -0.023 -0.027 -0.021 -0.013 0.010 -0.100 -0.105 -0,093 -0.091 -0.055 0.900 -0.044 -0,052 -0.050 --0.045 --0.027 -0.135 -.-0.132 -0.140 -0.125 -0.107 0.960 -0,094 -0.101 -0.097 -0.098 -0.093 -0.198 -0.209 -0.207 -0.207 -0.200 0.86 0.000 0.830 0.950 0.970 0.900 0.710 0.930 1.063 1.074 1.009 0.8 I0 0.025 -0.502 -0,328 -0.154 0.010 0.190 -0.680 -0.442 -0.208 -0.023 0.200 0.050 -0.416 -0.337 -0.227 -0.117 0.000 -0.606 -0.474 -0.333 -0.208 -0.055 0.075 --0.362 -0,315 -0.232 -0,164 -0.062 -0.512 -0.446 -0.350 -0.261 -0.143 0.100 -0.332 -0,295 -0.234 -0.179 -0.093 -0.487 -0.425 -0,354 -0.292 -0.185 0.125 -0.273 -0.247 -0.199 -0.171 -0.104 -0.413 -0.364 -0,316 -0.273 -0.186 0.150 -0.224 -0.207 -0.171 -0,143 -0,084 -0.352 -0.319 -0.270 -0.244 -0.166 0.210 -0.116 -0.084 -0.066 -0.044 -0.008 -0.222 -0.168 -0.133 -0.117 -0.051 0.260 -0.080 -0.060 -0.043 -0,027 0.000 -0.188 -0.153 -0,127 -0.105 -0.057 0.320 -0.061 -0.045 -0.027 -0.013 0.010 -0.159 -0.127 -0.102 -0.085 -0.040 0.370 --0.067 -0.052 -0.033 -0.021 0.000 -0.183 -0.146 -0.126 -0.102 -0,070 0.420 -0.049 -0.032 -0.027 -0.012 0.010 -0.161 -0.124 -0.100 -0.088 -0.047 0.470 -0.043 -0.033 -0,014 -0.012 0.010 -0.158 -0.117 -0.096 -0.072 -0.056 0.540 -0.052 -0.032 -0.022 -0.011 0.000 -0.178 -0.114 -0.096 -0.084 -0.054 0.600 -0.051 -0.024 -0.013 -0.005 0,010 -0.183 -0.112 -0.089 -0.077 -0,052 0.700 -0.066 -0.029 -0.013 --0.011 0.000 -0.197 -0.107 -0.073 -0.067 -0.040 0.800 -0.073 -0.029 -0.012 -0.010 0.000 -0.201 -0.121 -0.091 -0.084 -0.062 0.900 -0.138 -0.039 -0.023 -0.053 -0.062 -0.286 -0.154 -0.083 -0.099 -0.119 0.960 -0,159 -0.142 -0,119 -0.124 -0.107 -0.284 --0.264 -0,224 -0.219 -0.t92 Table C-5. Chordwise pressure distribution data, obtained in flight, as a function of angle of sideslip, Mach = 0.89 and 1.28.

M = 1.28, lip = 45,000 ft M = 0.89, lip = 45,000 ft 3.57 1.93 -0.07 -2.02 -3.59 13 7.96 3,99 -0.03 -3.82 -7.97 z//r x/c 1.407 1.445 1.411 1.360 1.38 I 0.16 0.000 1.112 1.189 1.216 1.188 1.104 0.090 0.200 0.360 0.420 0.510 0.025 -0.537 -0358 -0.111 0.090 0.300 -0.104 0.000 O. 150 O. 190 0.280 0.050 -0.669 -0.524 -0.306 -0.140 0.050 -0.192 -0.077 0.040 0.070 O. 130 0.075 -0.641 -0.524 -0,375 -0.237 -0.061 -0.224 -0.1 ! 4 -0.056 -0.025 0.020 0.100 -0.641 -0.545 -0.415 -0.293 -0.124 -0.183 -0.098 -0.117 -0.027 0.040 0.125 -0.639 -0.552 -0.423 -0.281 -0.129 -0.149 -0.083 .-0. i 60 -0.002 0.050 0.150 -0.712 -0.596 -0.472 -0.320 -0.104 -0.054 0.060 O. 180 O. 120 O. 180 0.210 -0.606 -0.422 -0.186 -0.137 -0.046 -0.176 -0.105 0.020 -0.055 -0.001 0.260 -0.684 --0.393 -0.288 -0.260 --0.191 -0.294 -0.230 -0.063 -0.178 -0.127 0.320 -0.565 -0.324 -0.378 -0.323 -0.265 -0.277 -0.231 -0.083 -0.174 -0.122 0.370 -0.430 -0.503 -0.471 -0.286 -0.1 ! 8 -0.250 -0.241 -0.137 -0.218 ---0.175 0.420 -0.214 -0.195 -0.170 -0.138 -0.102 -0.242 -0.210 -0.139 -0.194 -0.139 0.470 -0.225 -0.197 -0.164 -0.147 -0.120 -0.324 -0.280 -0.170 -0.268 -0.224 0.540 -0.209 -0.180 -0.164 -0.14,4 -0.109 -0.353 -0.319 -0.194 -0.285 -0.239 0.600 -0.164 -0.130 -0.114 -0.094 -0.059 -0.322 -0.326 -0.253 -0.260 -0.2{}4 0.700 .-0.133 -0.091 -0.076 -0.057 -0.028 -0.122 -0.097 -0.160 -0.094 -0.062 0.800 -0.127 -0.098 -0.082 -0.075 -0.045 -0.031 -0.005 -0.177 0.000 0.010 0.900 -0.111 -0.100 -0.082 -0.064 -0.043 0.020 0.060 0.030 0.090 O. 110 0.960 -0.419 -0.351 -0.298 -0.280 -0.228 1.430 1.443 1.430 1.378 1.374 0.48 0.000 !.110 1.195 1.220 1.192 1,091 O. 170 0.270 0.430 0.450 0.530 0.025 -0.594 -0.322 -0.091 O. 110 0.330 -0.049 0.040 0.210 0.210 0.290 0.050 -0.690 -0.554 -0.341 -0.154 0.040 -0.172 -0.083 0.070 0.070 0.140 0.075 -0.808 -0.664 -0.468 -0.293 -0.110 -0.273 -0. ! 80 -0.027 -0.047 0.010 0.100 -0.892 -0.756 -0.570 -0.419 -0.227 -0.364 -0.265 -0.121 -0.152 -0.092 0.125 -0.950 -0.823 -0.654 -0.502 -.0.315 -0.413 -0.308 -0.176 -0.213 -0.160 0.150 -0.889 -0.780 -0.645 -0.497 -0.282 -0.256 -0.212 -0.216 -0.134 -0.073 0.210 -0.426 -0.206 -0.106 -0.129 -0.100 -0.127 -0.083 -0.028 -0.026 0.020 0.260 -0.198 -0.155 -0.179 -0.181 -0.133 -0.082 -0.063 -0.024 -0.064 -0.027 0.320 -0.311 -0.313 -0.296 -0.248 -0.186 -0.102 -0.090 -0.012 -0.094 -0.058 0.370 -0.241 -0.305 -0.304 -0.198 -0.106 -0.095 -0.076 -0.009 -0.088 -0.052 0.420 -0.145 -0.145 -0.133 -0.118 -0.087 -0.145 -0,140 -0.049 -0.150 -0.110 0.470 -0.153 -0.143 -0.130 -0.113 -0.091 -0.181 -0.168 -0.084 -0.168 -0.127 0.540 -0.138 -0.137 -0.119 -0.107 -0.085 -0.219 -0.192 -0.113 -0.199 -0.174 0.600 -0.107 -0.104 -0.099 -0.086 -0.056 -0.269 -0.242 -0.170 -0.226 -0 ! 97 0.700 -0.096 -0.089 -0.085 -0.070 -0.043 -0.315 -0.289 -0.193 -0.168 -0.115 0.800 -0.088 -0.087 -0.075 -0.061 -0.033 0.000 0.000 -0.124 -0.008 0.000 0.900 -0.104 -0.115 -0.102 -0.075 -0.033 0.070 0.070 -0.113 0.100 0.140 0.960 -0.163 -0.160 -0.184 -0.185 -0.189 1.429 1.450 1.418 1.332 1.326 0.000 1.104 1.183 1.188 1.138 1.025 0.86 O. 130 0.230 0.370 0.370 0.460 0.025 -0.517 -0.357 -0.159 0.020 0.240 -0.077 -0.002 O. 150 O. 130 0.200 0.050 -0.768 -0.593 -0.398 -0.232 -0.054 -0.184 -0.120 0.030 0.000 0.060 0.075 -0.714 -0.603 -0.456 -0.343 -0.206 -0258 -0.205 -0.063 -0.104 -0.051 O. 1O0 -0.696 -0.633 -0.535 -0.450 -0.322 -0298 -0.244 -0.129 -0.177 -0.131 O, 125 -0.717 -0.654 -0.579 -0.498 -0.378 -0.329 -0.280 -0.189 -0.238 -0.194 0.150 -0.740 -0.687 -0.585 -0.459 -0.329 -0.270 -0.192 -0.167 -0.209 -0.173 0.210 -0.321 -0.097 -0.085 -0.126 -0.140 -0.258 -0.179 -0.111 -0.146 -0.072 0.260 -0.125 -0.124 -0.145 -0.146 -0.143 0.000 0.000 -0.119 0.000 0.030 0.320 -0.196 -0.181 -0.176 -0.153 -0.117 -0.008 -0.001 0.040 -0.042 -0.018 0.370 -0.222 -0.205 -0.184 -0.147 -0.115 -0.038 -0.030 0.030 -0.082 -0.068 0.420 -0.160 -0.152 -0.125 -0.108 -0.094 0.470 -0.147 -0.131 -0.106 -0.100 -0.090 -0.061 -0.064 -0.018 -0.105 -0.084 -0.106 -0.088 -0.056 -0.117 -0.114 0.540 -0.156 -0.129 -0.098 -0.091 -0.089 -0.126 -0.119 -0.077 -0.128 -0.099 0.600 -0.153 -0.112 -0.089 -0.071 -0.071 -0.170 -0.149 -0.091 -0.149 -0.136 0.700 -0,169 -0.107 -0.076 -0.065 -0.061 -0.219 -0.191 -0.116 -0.153 -0.136 0.800 -0,177 -0.102 -0.068 -0.063 -0.059 0.900 -0.122 -0.027 -0.004 -0.018 -0.027 -0.054 0.010 -0.068 0.060 0.080 -0.109 -0.100 -0.206 -0.100 -0.051 0.960 -0.274 -0.208 -0.172 -0.166 -0.177 Table C-6. Chordwise pressure distribution data, obtained in flight, as a function of angle of attack.

M = 0.54, lip = 30,000 ft -1.85 0.30 2.19 4.05 13 -0.21 -0.42 -0.27 -0.26 z//r x/c 0.000 1.029 1,034 1.023 1.000 0.16 0,025 -0,127 -0.104 -0,104 -0.103 0.050 -0.210 -0.197 -0.189 -0.182 0.075 -0.221 --0.196 -0.197 -0.192 0.100 -0.238 -0.223 -0.214 -0.202 0,125 -0.215 -0.203 -0.189 --0,186 0,150 -0.167 -0.159 -0.145 -0.148 0.210 -0.103 -0.087 -0.081 -0.069 0.260 -0.134 -0.119 -0.112 -0.099 0.320 -0,102 -0.086 -0.084 -0.080 0.370 ---0,075 --0.067 --0,058 -0.054 0.420 --0,059 --0,053 -0,048 -0.050 0.470 -0.068 -0.054 -..0.050 -0.055 0,540 -0,062 --0,055 -0.047 -0.051 0.600 -0.033 -0,025 -0,027 -0,028 0,700 -0,011 -0.011 -0.011 -0.020 0.800 -0.024 -0.020 -0.030 -0,047 0.900 -0,028 -0,030 -0.043 -0,063 0.960 -0,136 -0,148 -0.166 -0.189 0.000 1,033 1.039 1,029 1.007 0.48 0,025 -0,141 -0.124 -0.120 -0.124 0.050 -0.252 -0.231 -0.232 -0,223 0.075 -0.271 -0.248 -0,248 --0.247 0.100 -0.276 -0.256 --0.250 -0,241 0.125 -0.276 -0.258 -0.257 -0.252 0.150 -0,208 -0.201 -0,189 -0,189 0.210 -0,095 -0,087 -0,074 -0.082 0.260 -0,071 --0,060 --0,054 -0,053 0,320 -0,095 -0,080 -0,079 -0.083 0.370 -0.047 --0.039 -0,037 -0.042 0,420 -0,034 -0,023 -0,025 -0,026 0,470 -0.034 -0.027 -0.027 -0,033 0,540 -0.026 -0.022 -0.026 -0.022 0.600 -0,016 --0.014 -0.018 -0,031 0.700 -0.015 -0.013 -0.015 -0.026 0,800 -0.017 -0.019 --0.024 --0.036 0,900 -0,046 -0,047 -0,059 -0,073 0,960 --0,093 -0,090 -0,110 -0,135 0.000 0,990 0.990 0.980 0.920 0.86 0.025 -0.151 --0.125 -0.106 -0,098 0,050 -0,240 --0.213 --0.207 -0.197 0.075 -0,255 -0,229 -0,217 -0.197 0.100 -0.250 -0,230 --0,211 -0,213 0,125 -0,222 -0.208 -0.197 -0.187 0.150 -0.182 -0.169 -0.166 -0,163 0,210 -0,071 --0.057 -0,061 --0.060 0,260 -0,053 -0,046 -0.041 -0,044 0.320 -0.025 --0,020 -0,023 -0.026 0.370 -0,047 -0.041 -0.042 -0,051 0.420 -0.026 -0,015 -0.026 -0.028 0,470 -0.025 -0.019 --0.022 -0,034 0.540 -0.020 -0.013 -0.026 -0.045 0.600 -0,016 -0.011 -0.023 -0.034 0,700 -0,010 --0.005 -0,014 -0,037 0.800 --0.017 -0.010 -0.030 -0.049 0.900 -0.022 -0.022 -0.033 --0.046 0.960 -0.120 -0.125 -0.142 -0.176

APPENDIX D

BOUNDARY-LAYER PROFILE DATA

Table D-I. Steady-state boundary-layer profile data obtained in flight, Hp - 15,000 ft.

Free-stream Math number y (m.) 0.39 0.52 0.61 0.70 0.80 0.90 0.98 0.025 0.583 0,592 0.599 0.613 0.596 0.606 0.590 0.040 0.633 0.652 0,655 0.653 0.658 0,638 0,635 0.095 0.691 0.694 0.705 0.714 0.709 0.701 0.689 O. 150 0.709 0.714 0.728 0.740 0.743 0.738 0.713 O. 190 0.744 0.747 0.756 0.764 0.773 0,762 0.746 0.250 0.768 0,762 0.780 0.787 0.790 0.793 0.768 0.290 0.794 0.788 0.794 0.797 0.809 0.806 0.775 0.350 0.800 0.791 0.807 0.817 0.823 0.823 0.804 0.400 0.826 0.823 0.821 0.829 O.839 0.833 0.813 0.500 0.847 0.848 0.851 0.855 0.860 0,862 0.844 0.650 0.909 0.919 0.909 0.907 0.911 0.916 0.900 0.850 0.945 0.954 0.956 0.954 0.955 0.956 0.941 1.000 0.968 0.979 0.981 0.981 0,981 0.985 0.981 1.250 0.984 0,988 0.992 0.995 0,996 0.996 0,996 1.450 0.994 0,998 0.996 0,998 0.999 0,997 1.000 1.650 0.991 0.994 0.995 0.998 0.999 0,997 0,999 1.900 0,992 0.996 0.996 0.998 0,999 0.998 0.998 2.100 0.994 0.997 0.997 0.999 1.000 0.998 0.998 2.300 0.989 0.996 0.997 0,998 0,998 0,999 0.998 2.500 0.983 0.994 0.995 0.998 0.999 1.000 0.997 M r 0.42 0.53 0.64 0.76 0.85 0.95 0.99 1"/ 497.7 506.6 517.1 530.6 543.5 559.7 569.5 T,, 527.1 535.0 543.6 548.8 552.6 555.2 556.3 P,, 1194.3 1180.7 1168.2 1157.5 1153.5 1166.7 1205.5 Table D-2. Steady-state boundary-layer profile data obtained in flight. Hp = 30,000 fl.

Free-stream Mach number y (in.) 0.51 0.62 0.70 0.79 0.89 1.00 1.05 1.15 1.28 1.39 1.48 0.025 0.566 0.582 0.580 0.583 0.576 0.562 0.535 0.551 O. 156 0.070 O. 188 0,040 0.625 0.618 0.645 0,635 0.635 0.617 0,593 0.588 0.285 0,276 0.306 0.095 0.682 0,677 0,696 0.691 0,687 0,664 0,649 0,631 0,375 0.409 0.450 0.150 0.711 0.706 0,726 0,730 0.712 0.694 0,676 0,654 0,441 0.500 0.504 0.190 0,742 0.739 0.751 0,754 0.746 0.722 0,705 0.684 0.523 0.599 0,599 0.250 0.758 0.762 0.772 0,773 0.771 0,750 0.726 0,706 0.568 0.656 0.663 0,290 0,767 0.774 0,790 0.789 0.792 0,757 0.738 0.718 0.603 0.695 0.703 0.350 0.787 0.796 0.811 0.807 0,810 0.778 0.766 0.745 0.659 0.738 0,757 0,400 0,808 0,813 0.823 0,816 0.820 0.787 0,777 0,759 0,683 0.764 0,780 0.500 0.835 0,841 0.852 0,848 0.849 0.818 0.822 0.796 0.761 0.813 0.841 0,650 0.893 0,900 0,908 0.901 0.902 0.885 0,876 0,859 0.851 0,886 0,901 0.850 0,942 0.948 0,950 0.949 0,950 0.929 0,940 0.915 0.914 0,926 0,934 1.000 0.961 0.977 0.980 0.977 0.971 0.970 0,973 0.957 0,955 0.955 0,963 1.250 0.979 0,990 0.993 0,994 0.991 0,993 0.995 0.988 0.989 0,981 0.987 1.450 0.993 0,994 0,999 0.999 0.997 0,999 0.999 0,999 1.000 1.000 0.992 1.650 0.989 0,993 0.997 0.997 0.997 0,999 1,000 1.000 0,998 0.989 0.994 1.900 0,994 0.996 0.998 0,998 0.998 0.999 0.999 0.999 0,995 0.986 0,997 2.100 1.000 1.000 1.000 1.000 0.998 1.000 0.999 0,999 0.991 0,982 1,000 2.300 0.989 0,994 0.997 0.997 1.000 0.998 0,998 0,997 0.979 0.981 0,999 2,500 0.979 0.994 0,997 0,999 0.999 0,998 0.997 0,997 0,977 0.982 0,998 M r 0.53 0.65 0.75 0.84 0.93 1.00 1.01 1.09 1.27 1.55 1.71 T/ 435.2 445.1 455.4 465.6 477.5 490.6 503.8 526.6 548.7 558.9 593.1 T K 486.2 494.7 500.3 504.6 508.9 515.1 521.5 526.7 528.4 526.5 523.1 Pw 647.6 638.8 634.4 629.2 635.0 663.7 711.3 743.1 645.0 606, t 603.4 Table D-3. Steady-state boundary-layer profile data obtained in flight, Hp _ 45,000 ft.

Free-stream Mach number y (in.) 0.70 0.79 0.90 0.98 1.06 1.16 1.28 1.39 1.49 1.60 1.69 1.80 1.90 2.00 0,025 0.568 0.573 0.584 0.551 0.519 0.512 0.180 0,177 0.207 0.444 0.507 0.533 0.514 0.491 0,040 0.621 0.621 0.630 0.601 0,583 0.559 0.146 0,317 0.335 0.506 0,564 0.585 0.581 0.568 0,095 0.669 0,677 0.681 0.652 0.643 0.575 0,312 0.443 0.476 0,570 0.628 0.653 0.667 0.660 0.150 0.705 0.702 0.708 0.673 0,676 0.602 0.387 0.516 0.539 0.613 0,669 0.706 0.727 0.724 0.190 0.738 0.739 0.745 0.714 0.717 0.627 0.477 0,596 0.621 0.669 0.728 0,762 0.801 0.793 0.250 0.758 0.764 0.765 0.736 0,742 0,653 0,539 0,653 0,686 0.712 0.776 0.809 0,857 0,847 0,290 0,771 0.778 0,782 0.751 0,765 0.669 0,599 0,698 0.728 0.735 0.804 0.825 0.882 0.875 0,350 0.793 0.800 0,800 0.780 0.784 0,690 0.646 0.731 0.774 0,766 0,832 0.854 0.914 0.907 0,400 0.807 0.816 0.815 0,794 0,804 0.707 0.679 0.760 0,804 0,785 0,848 0,867 0.924 0,926 0.500 0.836 0,849 0.840 0.823 0.842 0,746 0.756 0,808 0,856 0.823 0,886 0.895 0.957 0.961 0.650 0.898 0.906 0.890 0.879 0,905 0.832 0.854 0.877 0.913 0.870 0.935 0.931 0,984 0,995 0,850 0,941 0,948 0,929 0.930 0,955 0.905 0.920 0,919 0.945 0.904 0.952 0,955 0.985 0.998 1,000 0.966 0,973 0,957 0,967 0.984 0.955 0.963 0.948 0.972 0.917 0,952 0,962 0,978 0,990 1.250 0.982 0,989 0,981 0,988 0,995 0.986 0,992 0.979 0.996 0.917 0,970 0,974 0.972 0,984 1.450 0.992 0,996 0.997 0.995 0.999 1.000 1,000 1.000 1.000 0.919 0.987 0.987 0.977 0.985 1.650 0.990 0.997 0.997 0.995 0.998 0.998 0.999 0.993 0.994 0.927 1.000 0.990 0.986 0_994 1.900 0.991 0.998 0,998 0,997 0.998 0.998 0.989 0.992 0.996 0.951 0.997 0.995 0.993 0,996 2.100 0.992 0,998 0.999 0.998 0.998 0.998 0.979 0.990 0.998 0.975 0.995 0.999 1,000 0.999 2.300 0.990 0.999 0.999 0.997 0.997 0,996 0,975 0.990 0.998 0.993 0.993 1.000 0.998 0.997 2.500 0,989 0.997 0.999 0,995 0.997 0.996 0.975 0.989 0,997 1.000 0.999 0,996 0.995 0,996 Mr 0.78 0.87 0.96 1.00 1.04 1.12 1.29 1.56 1.70 -- 1.91 2.00 2.09 2.17 T/ 427.2 438.2 451.2 461.7 475.4 496.4 510.5 530.9 553.3 602.7 620.5 635.3 652.9 679.9 T_, 446.7 445.5 445.8 446.6 449.3 454.6 459.7 467.6 476.1 586.7 587.5 587.3 585.8 582.5 P_,. 317.1 314.7 319.9 339.3 363.5 365.4 314.4 290.9 293.2 238.7 282.1 307.4 312.5 330.9 Table D-4. Boundary-layer profile data, obtained in flight, as a function of sideslip, Mach = 0.52 and 0.70.

M = 0.52, lip _ 15,000 ft M = 0.70, tip = 45,000 ft 13 8.04 3.95 0.00 -3.98 -8.00 8.02 4.01 -0.06 -3.98 -7.99 y (in.)

0.025 0.686 0.626 0.577 0.629 0,688 0.712 0.631 0.555 0.629 0.654 0.040 0.729 0,681 0,628 0.675 0.707 0.761 0.683 0.614 0.650 0.688 0.095 0.775 0,775 0,684 0.705 0.739 0.817 0.751 0.670 0.704 0,737 0.150 0,834 0.804 0.709 0.731 0,783 0,839 0.765 0.693 0,730 0,752 0,190 0,863 0,840 0.732 0,778 0.795 0.884 0.810 0.746 0,759 0.781 0.250 0.893 0.860 0.772 0.796 0.822 0.908 0.821 0.754 0.785 0.812 0.290 0.910 0.884 0,796 0.810 0.852 0.935 0844 0.773 0,796 0,832 0.350 0.924 0,893 0.819 0.832 0.872 0.955 0,875 0.795 0.819 0.839 0.400 0,947 0,897 0.831 0.842 0,880 0,973 0.877 0.814 0.837 0.864 0.500 0,975 0.928 0.863 0.863 0,905 0.985 0,907 0,843 0,875 0,899 0.650 0,998 0.981 0,902 0.921 0.957 0.999 0.952 0.909 0,921 0.950 0.850 0,989 0.985 0.951 0.964 0.987 1.000 0.955 0.947 0.963 0.973 1.000 0.996 0.965 0.974 0.986 0,997 0.993 0.944 0.972 0.982 0.991 1.250 0,989 0.971 0.988 0.995 0.996 0.996 0.945 0.992 0.988 0.987 1.450 1.000 0.963 0.996 0,995 0,997 0,999 0.970 0.993 0,989 0.999 1.650 0,996 0.969 0.997 0.995 0,997 0.985 0.981 0.991 0,988 0.990 1,900 0,980 0.976 0,997 0,996 0.997 0.966 0.986 0,989 0,989 0.991 2.100 0.965 0.983 0,998 0.997 0,996 0.947 0,990 0,986 0.991 0,991 2.300 0.952 0.996 0,998 0,998 0.997 0,935 0,960 0.973 0,990 0.997 2,500 0.928 1.000 1.000 0,996 0,994 0,922 0.936 0,960 0.991 0.994 Table D-5. Boundary-layer profile data, obtained in flight, as a function of angle of sideslip, Mach= 0.89 and 1.28.

M = 0.89, lip _ 45,000 fi M = 1.28, Hp _ 45,000 fl 13 7.96 3.99 -0.03 -3.82 -7.97 3.57 !.93 -0.07 -2.02 -3.59 y On.)

-- 0.331 0.276 0.514 0.025 0.693 0.637 0.561 0.620 0.64,4 -- 0.414 0.339 0.553 0.040 0.740 0.704 0.617 0.650 0.669 0.157 0.159 0.497 0.413 0.598 0.095 0.798 0.767 0.659 0.689 0.715 0.270 O. 147 0.552 0.456 0.639 O. 150 0.834 0.799 0.685 0.7 i 8 0.739 0.316 0.631 0.526 0.681 0.411 0.190 0.880 0.848 0.734 0.750 0.773 0.417 0.679 0.581 0.720 0.504 0.250 0.916 0.878 0.763 0.779 0.800 0.481 0.706 0.627 0.742 0.582 0.290 0.938 0.899 0.779 0.792 0.819 0.550 0.739 0.675 0.775 0.653 0.350 0.961 0.913 0.795 0.809 0.838 0.604 0.763 0.720 0.799 0.692 0.400 0.973 0.928 0.811 0.826 0.853 0.703 0.803 0.793 0.835 0.788 0.500 0.986 0.957 0.834 0.858 0.889 0.825 0.865 0.877 0.911 0.909 0.650 0.999 0.988 0.882 0.905 0.935 0.901 0.907 0.935 0.962 0.976 0.850 1.000 0.989 0.925 0.947 0.977 0.954 0.937 0.972 0.964 0.993 1.000 0.998 0.983 0.952 0.975 0.990 0.989 0.965 0.987 0.962 1.000 1.250 0.998 0.981 0.978 0.990 0.995 1.000 0.977 0.997 0.987 0.995 1.450 0.993 0.986 0.993 1.000 0.997 0.999 0.987 0.997 0.996 0.976 1.650 0.983 0.994 0.993 0.996 0.997 0.989 0.993 0.996 0.995 0.956 1.900 0.959 0.996 0.996 0.997 0.997 0.978 0.999 0.994 0.995 0.936 2.100 0.935 0.999 0.998 0.998 0.998 0.976 0.998 0.984 0.994 0.929 2.300 0.905 0.997 1.000 0.997 0.996 0.976 1.000 0.980 0.997 0.924 2.500 0.858 0.992 0.998 0.999 0.996 Table D-6. Boundary-layer profile data, obtained in flight, as a function of angle of attack, Mach _ 0.54, Hp = 30,000 ft.

M = 0.54, lip _ 30,000 ft ot -1.85 0.30 2.19 4.05 l_ -0.21 -0.42 --0.27 -0.26 y (in.)

0.025 0.588 0.583 0.572 0.599 0.040 0.636 0.631 0.631 0.68 I 0.095 0.687 0.682 0.708 0.736 0.150 0.722 0.702 0.746 0.778 0.190 0.746 0.752 0.773 0.805 0.250 0.773 0.764 0.793 0.833 0.290 0.780 0.793 0.833 0.837 0.350 0.803 0.809 0.846 0.865 0.400 0.818 0.817 0.855 0.882 0.500 0.849 0.836 0.881 0.912 0.650 0.904 0.903 0.954 0.952 0.850 0.951 0.958 0.989 0.953 1.000 0.980 0.974 0.995 0.944 1.250 0.994 0.991 1.000 0.940 1.450 1.000 0.995 0.999 0.940 1.650 0.994 1.000 0.984 0.946 1.900 0.997 1.000 0.982 0.964 2.100 0.999 1.000 0.980 0.982 2.300 0.992 0.996 0.962 0.994 2.500 0.999 0.991 0.960 1.000

REPORT DOCUMENTATION PAGE Form Approved

OMB NO. 0704-0188

Publicreporting burden k)rthiscollectionofinformatk_ is estimatedtoaverage 1 hour per response,including thetime for rmcewing instructions, searching exist=ng data sources,gathering and maintaining the data needed, end completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection o_ information, including suggestK)ns for reducing thisburden,to Washington HeadquartersServices,Directorate for In|orrnation Operations and Reports, 1215 JeffersonDavis Highway,Suite1204, Arhngton, VA 222024302, endto the Office ofManagementand Budget, Paperwork ReductionProject(0704-0188), Washington,DC 20503.

I. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT'rYPE AND DATES COVERED

December 1996 Technical Memorandum

4.TITLE AND SUBTITLE : S. FUNDING NUMBERS

F-15B/Flight Test Fixture II: A Test Bed for Flight Research

WU 505-68-52

6. AUTHOR(S)

David. M Richwine

8. PERFORMING ORGANIZATION

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

REPORT NUMBER

NASA Dryden Flight Research Center

P.O. Box 273

H-2113

Edwards, California 93523-0273

10. SPONSORING/MONrrORING

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AGENCY REPORT NUMBER

National Aeronautics and Space Administration

NASA TM-4782

Washington, DC 20546-0001

11. SUPPLEMENTARY NOTES 12b. DISTRIBUTION CODE 1211.DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified--Unlimited

Subject Category 02

13. ABSTRACT (Maximum 200 words) NASA Dryden Flight Research Center has developed a second-generation flight test fixture for use as a generic test bed for aerodynamic and fluid mechanics research. The Flight Test Fixture II (FTF-II) is a low- aspect-ratio vertical fin-like shape that is mounted on the centerline of the F-15B lower fuselage. The fixture is designed for flight research at Mach numbers to a maximum of 2.0. The FFF-II is a composite structure with a modular configuration and removable components for functional flexibility. This report documents the flow environment of the fixture, such as surface pressure distributions and boundary-layer profiles, throughout a matrix of conditions within the F-15B/FTF-II flight envelope. Environmental conditions within the fixture are presented to assist in the design and testing of future avionics and instrumentation. The intent of this document is to serve as a user's guide and assist in the development of future flight experiments that use the FTF-II as a test bed. Additional information enclosed in the appendices has been included to assist with more detailed analyses, if required.

14. SUBJECTTERMS 15. NUMBER OF PAGES Aerodynamic characterization, Boundary-layer pressures, F-I5B aerodynamic test bed, F-15 directional stability, Flight test facility, Flight test fixture, Flow visualization, 16. PRICE CODE

A04

Surface pressure distributions

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Unclassified Unclassified Unclassified Unlimited

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