USA 40 AIRFOIL (usa40-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 40 AIRFOIL (usa40-il) Reynolds number: 200,000 Max Cl/Cd: 71.09 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa40-il-200000.txt Download as CSV file: xf-usa40-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: USA 40 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3548 0.09394 0.09048 -0.0432 1.0000 0.0789
-8.500 -0.3934 0.09252 0.08919 -0.0402 1.0000 0.0790
-8.250 -0.3553 0.08888 0.08553 -0.0370 1.0000 0.0800
-8.000 -0.3583 0.08778 0.08449 -0.0333 1.0000 0.0806
-7.750 -0.3698 0.08700 0.08378 -0.0294 1.0000 0.0811
-7.500 -0.3466 0.08432 0.08110 -0.0318 0.9966 0.0830
-7.250 -0.3315 0.06780 0.06444 -0.0671 0.9805 0.0913
-7.000 -0.3019 0.06679 0.06350 -0.0650 0.9758 0.0923
-6.750 -0.2675 0.06444 0.06114 -0.0674 0.9715 0.0941
-6.500 -0.2515 0.03543 0.03043 -0.1017 0.9567 0.0673
-6.250 -0.2188 0.03101 0.02557 -0.1054 0.9480 0.0653
-6.000 -0.1766 0.02727 0.02131 -0.1100 0.9436 0.0646
-5.750 -0.1327 0.02435 0.01790 -0.1139 0.9407 0.0641
-5.500 -0.1014 0.02264 0.01592 -0.1148 0.9309 0.0644
-5.250 -0.0613 0.02105 0.01409 -0.1171 0.9259 0.0654
-5.000 -0.0296 0.01993 0.01277 -0.1176 0.9160 0.0667
-4.750 0.0075 0.01904 0.01164 -0.1189 0.9090 0.0688
-4.500 0.0360 0.01779 0.01037 -0.1189 0.8975 0.0708
-4.250 0.0684 0.01691 0.00949 -0.1194 0.8884 0.0731
-4.000 0.0970 0.01623 0.00876 -0.1191 0.8758 0.0759
-3.750 0.1255 0.01570 0.00812 -0.1187 0.8628 0.0789
-3.500 0.1536 0.01484 0.00735 -0.1185 0.8505 0.0842
-3.250 0.1823 0.01434 0.00677 -0.1182 0.8377 0.0910
-3.000 0.2091 0.01373 0.00621 -0.1178 0.8229 0.1010
-2.750 0.2368 0.01318 0.00569 -0.1175 0.8085 0.1175
-2.500 0.2648 0.01267 0.00534 -0.1174 0.7945 0.1493
-2.250 0.2928 0.01227 0.00511 -0.1172 0.7803 0.2146
-2.000 0.3204 0.01192 0.00491 -0.1171 0.7656 0.2732
-1.750 0.3475 0.01164 0.00482 -0.1168 0.7499 0.3350
-1.500 0.3745 0.01139 0.00476 -0.1165 0.7342 0.4100
-1.250 0.4008 0.01113 0.00475 -0.1159 0.7183 0.4958
-1.000 0.4263 0.01100 0.00483 -0.1150 0.7025 0.5829
-0.750 0.4517 0.01103 0.00493 -0.1139 0.6866 0.6469
-0.500 0.4769 0.01114 0.00502 -0.1127 0.6704 0.6959
-0.250 0.5018 0.01127 0.00510 -0.1115 0.6539 0.7332
0.000 0.5260 0.01140 0.00516 -0.1102 0.6374 0.7633
0.250 0.5499 0.01152 0.00520 -0.1088 0.6210 0.7913
0.500 0.5734 0.01160 0.00522 -0.1074 0.6044 0.8161
0.750 0.5969 0.01165 0.00521 -0.1060 0.5873 0.8390
1.000 0.6195 0.01166 0.00518 -0.1045 0.5706 0.8602
1.250 0.6421 0.01165 0.00514 -0.1031 0.5545 0.8840
1.500 0.6655 0.01163 0.00508 -0.1017 0.5389 0.9130
1.750 0.6987 0.01164 0.00503 -0.1026 0.5227 1.0000
2.000 0.7284 0.01189 0.00513 -0.1032 0.5081 1.0000
2.250 0.7575 0.01216 0.00523 -0.1036 0.4943 1.0000
2.500 0.7860 0.01244 0.00534 -0.1038 0.4808 1.0000
2.750 0.8142 0.01269 0.00550 -0.1040 0.4672 1.0000
3.000 0.8421 0.01297 0.00569 -0.1041 0.4553 1.0000
3.250 0.8698 0.01331 0.00586 -0.1041 0.4448 1.0000
3.500 0.8971 0.01356 0.00607 -0.1040 0.4332 1.0000
3.750 0.9240 0.01386 0.00628 -0.1039 0.4219 1.0000
4.000 0.9505 0.01420 0.00647 -0.1037 0.4114 1.0000
4.250 0.9772 0.01443 0.00671 -0.1035 0.4011 1.0000
4.500 1.0037 0.01480 0.00696 -0.1033 0.3927 1.0000
4.750 1.0298 0.01501 0.00719 -0.1030 0.3823 1.0000
5.000 1.0554 0.01532 0.00742 -0.1027 0.3726 1.0000
5.250 1.0811 0.01558 0.00767 -0.1023 0.3634 1.0000
5.500 1.1069 0.01592 0.00796 -0.1020 0.3562 1.0000
5.750 1.1325 0.01617 0.00825 -0.1016 0.3482 1.0000
6.000 1.1575 0.01654 0.00853 -0.1012 0.3409 1.0000
6.250 1.1826 0.01677 0.00884 -0.1008 0.3323 1.0000
6.500 1.2066 0.01712 0.00911 -0.1002 0.3241 1.0000
6.750 1.2310 0.01735 0.00942 -0.0996 0.3148 1.0000
7.000 1.2545 0.01773 0.00973 -0.0990 0.3076 1.0000
7.250 1.2789 0.01799 0.01009 -0.0985 0.2995 1.0000
7.500 1.3014 0.01838 0.01042 -0.0977 0.2918 1.0000
7.750 1.3250 0.01865 0.01078 -0.0970 0.2826 1.0000
8.000 1.3466 0.01905 0.01115 -0.0961 0.2743 1.0000
8.250 1.3690 0.01938 0.01153 -0.0953 0.2643 1.0000
8.500 1.3902 0.01978 0.01195 -0.0944 0.2550 1.0000
8.750 1.4104 0.02022 0.01238 -0.0933 0.2455 1.0000
9.000 1.4311 0.02065 0.01284 -0.0923 0.2354 1.0000
9.250 1.4499 0.02116 0.01336 -0.0910 0.2257 1.0000
9.500 1.4679 0.02169 0.01390 -0.0896 0.2155 1.0000
9.750 1.4847 0.02223 0.01446 -0.0881 0.2054 1.0000
10.000 1.4988 0.02291 0.01512 -0.0862 0.1925 1.0000
10.250 1.5119 0.02369 0.01588 -0.0843 0.1795 1.0000
10.500 1.5236 0.02457 0.01673 -0.0823 0.1656 1.0000
10.750 1.5333 0.02561 0.01773 -0.0802 0.1502 1.0000
11.000 1.5399 0.02690 0.01896 -0.0780 0.1326 1.0000
11.250 1.5428 0.02851 0.02048 -0.0756 0.1132 1.0000
11.500 1.5441 0.03034 0.02227 -0.0733 0.0995 1.0000
11.750 1.5448 0.03233 0.02423 -0.0713 0.0902 1.0000
12.000 1.5472 0.03428 0.02623 -0.0696 0.0835 1.0000
12.250 1.5485 0.03644 0.02844 -0.0682 0.0776 1.0000
12.500 1.5527 0.03842 0.03051 -0.0670 0.0707 1.0000
12.750 1.5545 0.04073 0.03291 -0.0660 0.0597 1.0000
13.000 1.5470 0.04410 0.03625 -0.0651 0.0517 1.0000
13.250 1.5413 0.04746 0.03966 -0.0646 0.0493 1.0000
13.500 1.5357 0.05098 0.04325 -0.0644 0.0477 1.0000
13.750 1.5287 0.05479 0.04718 -0.0644 0.0468 1.0000
14.000 1.5202 0.05894 0.05146 -0.0646 0.0461 1.0000
14.250 1.5104 0.06341 0.05606 -0.0652 0.0456 1.0000
14.500 1.4988 0.06826 0.06106 -0.0660 0.0451 1.0000
14.750 1.4860 0.07346 0.06640 -0.0670 0.0447 1.0000
15.000 1.4721 0.07894 0.07202 -0.0683 0.0443 1.0000
15.250 1.4579 0.08459 0.07780 -0.0697 0.0440 1.0000
15.500 1.4435 0.09038 0.08373 -0.0713 0.0437 1.0000
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