WHITCOMB INTEGRAL SUPERCRITICAL AIRFOIL (whitcomb-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: WHITCOMB INTEGRAL SUPERCRITICAL AIRFOIL (whitcomb-il) Reynolds number: 100,000 Max Cl/Cd: 22.87 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-whitcomb-il-100000.txt Download as CSV file: xf-whitcomb-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: WHITCOMB INTEGRAL SUPERCRITICAL AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.6043 0.10549 0.09987 -0.0247 1.0000 0.1743
-9.500 -0.7838 0.07041 0.06444 -0.0433 1.0000 0.0955
-9.250 -0.7902 0.06598 0.05996 -0.0432 1.0000 0.0944
-9.000 -0.8014 0.06111 0.05493 -0.0436 1.0000 0.0929
-8.750 -0.8116 0.05560 0.04911 -0.0443 1.0000 0.0910
-8.500 -0.8142 0.04975 0.04278 -0.0455 1.0000 0.0892
-8.250 -0.8066 0.04420 0.03660 -0.0471 1.0000 0.0878
-8.000 -0.7895 0.04049 0.03246 -0.0479 1.0000 0.0881
-7.750 -0.7688 0.03778 0.02946 -0.0483 1.0000 0.0896
-7.500 -0.7458 0.03549 0.02685 -0.0489 1.0000 0.0925
-7.250 -0.7200 0.03320 0.02407 -0.0498 1.0000 0.0955
-7.000 -0.6949 0.03081 0.02140 -0.0501 1.0000 0.0980
-6.750 -0.6710 0.02937 0.02000 -0.0499 1.0000 0.1021
-6.500 -0.6449 0.02811 0.01857 -0.0499 1.0000 0.1075
-6.250 -0.6194 0.02656 0.01688 -0.0496 1.0000 0.1123
-6.000 -0.5944 0.02554 0.01591 -0.0493 1.0000 0.1189
-5.750 -0.5680 0.02444 0.01469 -0.0490 1.0000 0.1263
-5.500 -0.5430 0.02349 0.01389 -0.0486 1.0000 0.1354
-5.250 -0.5176 0.02252 0.01299 -0.0482 1.0000 0.1463
-5.000 -0.4915 0.02164 0.01221 -0.0479 1.0000 0.1610
-4.750 -0.4646 0.02080 0.01158 -0.0479 1.0000 0.1813
-4.500 -0.4369 0.01993 0.01096 -0.0481 1.0000 0.2123
-4.250 -0.4078 0.01898 0.01051 -0.0487 1.0000 0.2657
-4.000 -0.3740 0.01759 0.01065 -0.0505 1.0000 0.4498
-3.750 -0.3918 0.01950 0.01325 -0.0366 1.0000 0.5318
-3.500 -0.3698 0.02120 0.01487 -0.0337 1.0000 0.6259
-3.250 -0.3518 0.02263 0.01618 -0.0301 1.0000 0.6591
-3.000 -0.3408 0.02375 0.01726 -0.0247 1.0000 0.6791
-2.750 -0.3295 0.02458 0.01805 -0.0196 1.0000 0.6968
-2.500 -0.3186 0.02518 0.01860 -0.0146 1.0000 0.7133
-2.250 -0.3074 0.02557 0.01897 -0.0098 1.0000 0.7292
-2.000 -0.2959 0.02580 0.01916 -0.0053 1.0000 0.7450
-1.750 -0.2842 0.02589 0.01923 -0.0009 1.0000 0.7612
-1.500 -0.2729 0.02588 0.01920 0.0034 1.0000 0.7780
-1.250 -0.2614 0.02577 0.01907 0.0075 1.0000 0.7954
-1.000 -0.2490 0.02557 0.01885 0.0112 1.0000 0.8123
-0.750 -0.2361 0.02527 0.01854 0.0146 1.0000 0.8284
-0.500 -0.2230 0.02486 0.01813 0.0178 1.0000 0.8425
-0.250 -0.2095 0.02443 0.01770 0.0208 1.0000 0.8568
0.000 -0.1966 0.02401 0.01728 0.0240 1.0000 0.8737
0.250 -0.1850 0.02350 0.01678 0.0273 1.0000 0.8905
0.500 -0.1700 0.02302 0.01631 0.0296 1.0000 0.9037
0.750 -0.1517 0.02253 0.01583 0.0310 1.0000 0.9109
1.000 -0.1247 0.02244 0.01576 0.0300 1.0000 0.9150
1.250 -0.0986 0.02231 0.01566 0.0292 1.0000 0.9181
1.500 -0.0736 0.02218 0.01557 0.0287 1.0000 0.9210
1.750 -0.0466 0.02219 0.01564 0.0277 1.0000 0.9233
2.000 -0.0185 0.02230 0.01582 0.0263 1.0000 0.9252
2.250 0.0247 0.02274 0.01635 0.0221 0.9940 0.9265
2.500 0.0998 0.02314 0.01688 0.0130 0.9658 0.9277
2.750 0.1676 0.02282 0.01670 0.0061 0.9383 0.9286
3.000 0.2313 0.02178 0.01584 0.0009 0.9066 0.9291
3.250 0.2792 0.02021 0.01446 -0.0006 0.8646 0.9301
3.500 0.3113 0.01877 0.01315 0.0008 0.7956 0.9310
3.750 0.4191 0.02000 0.01147 -0.0096 0.2661 0.9325
4.000 0.4408 0.02104 0.01202 -0.0090 0.2040 0.9327
4.250 0.4663 0.02199 0.01271 -0.0090 0.1723 0.9326
4.500 0.4941 0.02285 0.01343 -0.0094 0.1521 0.9331
4.750 0.5239 0.02378 0.01426 -0.0100 0.1379 0.9339
5.000 0.5560 0.02495 0.01527 -0.0112 0.1273 0.9343
5.250 0.5868 0.02581 0.01623 -0.0120 0.1184 0.9342
5.500 0.6201 0.02729 0.01767 -0.0135 0.1117 0.9342
5.750 0.6503 0.02844 0.01901 -0.0140 0.1062 0.9345
6.000 0.6797 0.02974 0.02028 -0.0149 0.1009 0.9351
6.250 0.7081 0.03164 0.02245 -0.0152 0.0978 0.9356
6.500 0.7343 0.03344 0.02466 -0.0151 0.0954 0.9355
6.750 0.7594 0.03527 0.02680 -0.0150 0.0925 0.9354
7.000 0.7854 0.03703 0.02871 -0.0153 0.0897 0.9354
7.250 0.8092 0.03948 0.03141 -0.0153 0.0887 0.9355
7.500 0.8237 0.04289 0.03547 -0.0138 0.0895 0.9359
7.750 0.8288 0.04796 0.04133 -0.0114 0.0928 0.9363
8.000 0.8354 0.05264 0.04650 -0.0100 0.0953 0.9365
8.250 0.8442 0.05715 0.05130 -0.0095 0.0973 0.9363
8.500 0.8708 0.06184 0.05590 -0.0109 0.0999 0.9364
8.750 0.7693 0.08708 0.08285 -0.0136 0.1779 0.9355
9.000 0.6968 0.09559 0.09153 -0.0261 0.1703 0.9346
9.250 0.7196 0.09777 0.09373 -0.0243 0.1646 0.9347
9.500 0.7945 0.10043 0.09625 -0.0163 0.1614 0.9352
9.750 0.6935 0.11082 0.10674 -0.0370 0.1543 0.9343
10.000 0.7009 0.11474 0.11066 -0.0387 0.1497 0.9345
|
Polar data table (+)
Polar graphs
<< Back to WHITCOMB INTEGRAL SUPERCRITICAL AIRFOIL (whitcomb-il)