NPL 9510 AIRFOIL (npl9510-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NPL 9510 AIRFOIL (npl9510-il) Reynolds number: 1,000,000 Max Cl/Cd: 57.16 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-npl9510-il-1000000.txt Download as CSV file: xf-npl9510-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NPL 9510 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-19.750 -1.1438 0.12886 0.12607 -0.0083 1.0000 0.0159
-19.500 -1.1563 0.12248 0.11959 -0.0118 1.0000 0.0160
-19.250 -1.1681 0.11632 0.11333 -0.0150 1.0000 0.0161
-19.000 -1.1795 0.11037 0.10729 -0.0182 1.0000 0.0162
-18.750 -1.1903 0.10464 0.10146 -0.0211 1.0000 0.0163
-18.500 -1.2001 0.09911 0.09584 -0.0239 1.0000 0.0164
-18.250 -1.2093 0.09381 0.09045 -0.0265 1.0000 0.0164
-18.000 -1.2176 0.08879 0.08533 -0.0290 1.0000 0.0165
-17.750 -1.2248 0.08396 0.08041 -0.0313 1.0000 0.0166
-17.500 -1.2310 0.07940 0.07576 -0.0334 1.0000 0.0168
-17.250 -1.2364 0.07509 0.07135 -0.0354 1.0000 0.0169
-17.000 -1.2405 0.07099 0.06716 -0.0373 1.0000 0.0170
-16.750 -1.2437 0.06711 0.06319 -0.0389 1.0000 0.0171
-16.500 -1.2461 0.06345 0.05944 -0.0405 1.0000 0.0173
-16.250 -1.2473 0.05998 0.05587 -0.0419 1.0000 0.0174
-16.000 -1.2477 0.05672 0.05253 -0.0432 1.0000 0.0175
-15.750 -1.2467 0.05369 0.04941 -0.0444 1.0000 0.0177
-15.500 -1.2445 0.05088 0.04652 -0.0454 1.0000 0.0178
-15.250 -1.2413 0.04831 0.04386 -0.0463 1.0000 0.0179
-15.000 -1.2372 0.04590 0.04137 -0.0471 1.0000 0.0180
-14.750 -1.2406 0.04285 0.03824 -0.0476 1.0000 0.0183
-14.500 -1.2438 0.03984 0.03516 -0.0482 1.0000 0.0186
-14.250 -1.2420 0.03734 0.03260 -0.0488 1.0000 0.0188
-14.000 -1.2373 0.03516 0.03037 -0.0493 1.0000 0.0191
-13.750 -1.2304 0.03321 0.02836 -0.0497 1.0000 0.0193
-13.500 -1.2226 0.03139 0.02650 -0.0500 1.0000 0.0196
-13.250 -1.2147 0.02965 0.02471 -0.0502 1.0000 0.0198
-13.000 -1.2077 0.02799 0.02299 -0.0501 1.0000 0.0200
-12.750 -1.2048 0.02644 0.02139 -0.0490 1.0000 0.0203
-12.500 -1.2005 0.02498 0.01988 -0.0481 1.0000 0.0205
-12.250 -1.1911 0.02363 0.01847 -0.0479 1.0000 0.0208
-12.000 -1.1781 0.02243 0.01721 -0.0480 1.0000 0.0211
-11.750 -1.1617 0.02139 0.01611 -0.0483 1.0000 0.0214
-11.500 -1.1437 0.02053 0.01518 -0.0484 1.0000 0.0216
-11.250 -1.1284 0.01908 0.01366 -0.0486 1.0000 0.0222
-11.000 -1.1089 0.01816 0.01269 -0.0487 1.0000 0.0227
-10.750 -1.0877 0.01747 0.01198 -0.0486 1.0000 0.0232
-10.500 -1.0655 0.01689 0.01136 -0.0485 1.0000 0.0238
-10.250 -1.0428 0.01634 0.01078 -0.0485 1.0000 0.0244
-10.000 -1.0120 0.01581 0.01020 -0.0499 0.9994 0.0251
-9.750 -0.9803 0.01527 0.00962 -0.0516 0.9987 0.0258
-9.500 -0.9486 0.01449 0.00881 -0.0535 0.9980 0.0273
-9.250 -0.9165 0.01401 0.00832 -0.0551 0.9972 0.0286
-9.000 -0.8843 0.01363 0.00790 -0.0565 0.9965 0.0299
-8.750 -0.8518 0.01304 0.00732 -0.0583 0.9958 0.0324
-8.500 -0.8190 0.01266 0.00692 -0.0599 0.9952 0.0348
-8.250 -0.7860 0.01218 0.00645 -0.0616 0.9947 0.0385
-8.000 -0.7527 0.01178 0.00606 -0.0633 0.9942 0.0430
-7.750 -0.7207 0.01135 0.00566 -0.0647 0.9933 0.0498
-7.500 -0.6889 0.01087 0.00527 -0.0661 0.9921 0.0618
-7.250 -0.6565 0.01041 0.00491 -0.0676 0.9911 0.0793
-7.000 -0.6241 0.01004 0.00463 -0.0690 0.9901 0.0963
-6.750 -0.5914 0.00973 0.00437 -0.0704 0.9892 0.1099
-6.500 -0.5585 0.00943 0.00414 -0.0719 0.9883 0.1240
-6.250 -0.5252 0.00915 0.00392 -0.0734 0.9875 0.1393
-6.000 -0.4916 0.00884 0.00369 -0.0749 0.9869 0.1583
-5.750 -0.4576 0.00850 0.00346 -0.0766 0.9863 0.1820
-5.500 -0.4234 0.00807 0.00319 -0.0784 0.9858 0.2192
-5.250 -0.3890 0.00748 0.00289 -0.0805 0.9854 0.2840
-5.000 -0.3546 0.00703 0.00267 -0.0824 0.9850 0.3430
-4.750 -0.3203 0.00678 0.00255 -0.0840 0.9846 0.3790
-4.500 -0.2859 0.00661 0.00247 -0.0855 0.9842 0.4094
-4.250 -0.2514 0.00648 0.00242 -0.0871 0.9838 0.4303
-4.000 -0.2223 0.00643 0.00238 -0.0873 0.9814 0.4426
-3.750 -0.1906 0.00637 0.00235 -0.0881 0.9798 0.4540
-3.500 -0.1579 0.00632 0.00231 -0.0891 0.9787 0.4619
-3.250 -0.1251 0.00627 0.00228 -0.0901 0.9775 0.4692
-3.000 -0.0923 0.00623 0.00225 -0.0911 0.9764 0.4753
-2.750 -0.0594 0.00618 0.00224 -0.0921 0.9754 0.4823
-2.500 -0.0265 0.00616 0.00222 -0.0931 0.9743 0.4880
-2.250 0.0080 0.00607 0.00216 -0.0944 0.9728 0.4936
-2.000 0.0374 0.00598 0.00210 -0.0944 0.9673 0.5013
-1.750 0.0649 0.00581 0.00192 -0.0936 0.9562 0.5085
-1.500 0.0896 0.00568 0.00181 -0.0923 0.9432 0.5146
-1.250 0.1163 0.00564 0.00177 -0.0917 0.9320 0.5201
-1.000 0.1438 0.00561 0.00172 -0.0912 0.9195 0.5245
-0.750 0.1708 0.00556 0.00166 -0.0907 0.9003 0.5287
-0.500 0.1976 0.00559 0.00161 -0.0900 0.8684 0.5323
-0.250 0.2222 0.00585 0.00157 -0.0888 0.7862 0.5359
0.000 0.2331 0.00827 0.00200 -0.0860 0.2712 0.5387
0.250 0.2583 0.00898 0.00216 -0.0856 0.1185 0.5421
0.500 0.2855 0.00936 0.00231 -0.0855 0.0607 0.5459
0.750 0.3138 0.00954 0.00246 -0.0854 0.0506 0.5494
1.000 0.3421 0.00974 0.00262 -0.0854 0.0454 0.5529
1.250 0.3705 0.00992 0.00278 -0.0853 0.0428 0.5561
1.500 0.3989 0.01006 0.00292 -0.0853 0.0405 0.5596
1.750 0.4266 0.01033 0.00321 -0.0851 0.0380 0.5633
2.000 0.4550 0.01049 0.00338 -0.0850 0.0369 0.5668
2.250 0.4831 0.01067 0.00357 -0.0849 0.0356 0.5702
2.500 0.5111 0.01087 0.00375 -0.0848 0.0342 0.5732
2.750 0.5380 0.01127 0.00415 -0.0845 0.0326 0.5768
3.000 0.5659 0.01147 0.00438 -0.0844 0.0319 0.5805
3.250 0.5937 0.01169 0.00463 -0.0842 0.0310 0.5840
3.500 0.6213 0.01194 0.00489 -0.0840 0.0301 0.5873
3.750 0.6489 0.01218 0.00512 -0.0838 0.0291 0.5905
4.000 0.6750 0.01270 0.00563 -0.0834 0.0279 0.5941
4.250 0.7021 0.01301 0.00600 -0.0831 0.0273 0.5980
4.500 0.7295 0.01328 0.00632 -0.0828 0.0266 0.6017
4.750 0.7566 0.01360 0.00666 -0.0825 0.0258 0.6053
5.000 0.7836 0.01392 0.00698 -0.0823 0.0252 0.6085
5.250 0.8105 0.01423 0.00731 -0.0820 0.0245 0.6125
5.500 0.8358 0.01494 0.00805 -0.0814 0.0237 0.6165
5.750 0.8618 0.01553 0.00871 -0.0809 0.0232 0.6206
6.000 0.8887 0.01584 0.00907 -0.0806 0.0227 0.6244
6.250 0.9153 0.01620 0.00949 -0.0802 0.0221 0.6281
6.500 0.9417 0.01659 0.00994 -0.0798 0.0216 0.6325
6.750 0.9679 0.01698 0.01038 -0.0794 0.0211 0.6371
7.000 0.9940 0.01739 0.01081 -0.0790 0.0207 0.6416
7.250 1.0193 0.01794 0.01139 -0.0785 0.0203 0.6459
7.500 1.0420 0.01936 0.01295 -0.0776 0.0198 0.6504
7.750 1.0674 0.01985 0.01356 -0.0770 0.0195 0.6554
8.000 1.0921 0.02052 0.01434 -0.0764 0.0192 0.6603
8.250 1.1162 0.02131 0.01527 -0.0756 0.0189 0.6656
8.500 1.1396 0.02216 0.01626 -0.0748 0.0185 0.6711
8.750 1.1627 0.02301 0.01723 -0.0740 0.0182 0.6766
9.000 1.1854 0.02387 0.01821 -0.0731 0.0180 0.6821
9.250 1.2078 0.02466 0.01913 -0.0722 0.0177 0.6884
9.500 1.2297 0.02547 0.02004 -0.0712 0.0175 0.6948
9.750 1.2516 0.02617 0.02083 -0.0702 0.0173 0.7011
10.000 1.2729 0.02688 0.02161 -0.0692 0.0171 0.7079
10.250 1.2912 0.02812 0.02296 -0.0679 0.0168 0.7143
10.500 1.2972 0.03143 0.02669 -0.0650 0.0165 0.7205
10.750 1.3115 0.03302 0.02850 -0.0631 0.0164 0.7279
11.000 1.3199 0.03537 0.03117 -0.0606 0.0162 0.7346
11.250 1.3151 0.03933 0.03557 -0.0568 0.0160 0.7413
11.500 0.6645 0.13141 0.12928 -0.0649 0.0204 0.6747
|
Polar data table (+)
Polar graphs
<< Back to NPL 9510 AIRFOIL (npl9510-il)