RAE 100 AIRFOIL (rae100-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAE 100 AIRFOIL (rae100-il) Reynolds number: 100,000 Max Cl/Cd: 35.24 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae100-il-100000-n5.txt Download as CSV file: xf-rae100-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAE 100 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.750 -0.7911 0.09882 0.09340 -0.0011 1.0000 0.0333
-12.500 -0.8322 0.08482 0.07932 -0.0107 1.0000 0.0328
-12.250 -0.8675 0.07390 0.06826 -0.0191 1.0000 0.0323
-12.000 -0.8957 0.06595 0.06013 -0.0251 1.0000 0.0321
-11.750 -0.9205 0.05982 0.05378 -0.0288 1.0000 0.0320
-11.500 -0.9406 0.05520 0.04892 -0.0301 1.0000 0.0320
-11.250 -0.9584 0.05143 0.04488 -0.0294 1.0000 0.0322
-11.000 -0.9716 0.04826 0.04142 -0.0274 1.0000 0.0326
-10.750 -0.9770 0.04512 0.03790 -0.0256 1.0000 0.0330
-10.500 -0.9771 0.04222 0.03459 -0.0239 1.0000 0.0339
-10.250 -0.9661 0.04044 0.03278 -0.0228 1.0000 0.0351
-10.000 -0.9545 0.03878 0.03100 -0.0217 1.0000 0.0364
-9.750 -0.9424 0.03686 0.02885 -0.0204 1.0000 0.0379
-9.500 -0.9287 0.03478 0.02646 -0.0191 1.0000 0.0395
-9.250 -0.9129 0.03280 0.02413 -0.0178 1.0000 0.0411
-9.000 -0.8962 0.03120 0.02247 -0.0168 1.0000 0.0431
-8.750 -0.8781 0.02998 0.02119 -0.0159 1.0000 0.0455
-8.500 -0.8591 0.02862 0.01964 -0.0148 1.0000 0.0484
-8.250 -0.8395 0.02722 0.01804 -0.0137 1.0000 0.0511
-8.000 -0.8207 0.02606 0.01690 -0.0127 1.0000 0.0544
-7.750 -0.8005 0.02502 0.01574 -0.0117 1.0000 0.0587
-7.500 -0.7807 0.02388 0.01452 -0.0105 1.0000 0.0628
-7.250 -0.7605 0.02296 0.01357 -0.0095 1.0000 0.0682
-7.000 -0.7400 0.02201 0.01256 -0.0084 1.0000 0.0745
-6.750 -0.7193 0.02116 0.01167 -0.0074 1.0000 0.0821
-6.500 -0.6987 0.02031 0.01083 -0.0063 1.0000 0.0911
-6.250 -0.6777 0.01953 0.01005 -0.0052 1.0000 0.1023
-6.000 -0.6565 0.01879 0.00934 -0.0042 1.0000 0.1163
-5.750 -0.6352 0.01809 0.00868 -0.0032 1.0000 0.1335
-5.500 -0.6135 0.01745 0.00809 -0.0022 1.0000 0.1550
-5.250 -0.5922 0.01681 0.00757 -0.0012 1.0000 0.1803
-5.000 -0.5707 0.01623 0.00711 -0.0002 1.0000 0.2089
-4.750 -0.5491 0.01571 0.00669 0.0009 1.0000 0.2412
-4.500 -0.5274 0.01523 0.00633 0.0019 1.0000 0.2749
-4.250 -0.5059 0.01478 0.00602 0.0031 1.0000 0.3099
-4.000 -0.4842 0.01438 0.00574 0.0042 1.0000 0.3455
-3.750 -0.4625 0.01403 0.00550 0.0054 1.0000 0.3811
-3.500 -0.4409 0.01370 0.00530 0.0066 1.0000 0.4160
-3.250 -0.4193 0.01341 0.00514 0.0078 1.0000 0.4503
-3.000 -0.3978 0.01315 0.00499 0.0091 1.0000 0.4843
-2.750 -0.3764 0.01292 0.00487 0.0104 1.0000 0.5175
-2.500 -0.3552 0.01272 0.00479 0.0118 1.0000 0.5500
-2.250 -0.3343 0.01255 0.00474 0.0132 1.0000 0.5820
-2.000 -0.3137 0.01241 0.00470 0.0147 1.0000 0.6136
-1.750 -0.2795 0.01229 0.00472 0.0134 0.9908 0.6489
-1.500 -0.2399 0.01219 0.00473 0.0112 0.9774 0.6850
-1.250 -0.2002 0.01210 0.00474 0.0091 0.9634 0.7199
-0.750 -0.1217 0.01193 0.00476 0.0055 0.9321 0.7859
-0.500 -0.0820 0.01188 0.00478 0.0038 0.9139 0.8168
-0.250 -0.0408 0.01184 0.00479 0.0018 0.8944 0.8453
0.000 0.0000 0.01183 0.00481 0.0000 0.8709 0.8709
0.250 0.0408 0.01184 0.00480 -0.0018 0.8453 0.8944
0.500 0.0820 0.01188 0.00478 -0.0038 0.8168 0.9139
0.750 0.1217 0.01193 0.00476 -0.0055 0.7860 0.9320
1.250 0.2002 0.01210 0.00474 -0.0091 0.7200 0.9634
1.500 0.2399 0.01219 0.00473 -0.0112 0.6849 0.9774
1.750 0.2795 0.01229 0.00472 -0.0134 0.6490 0.9908
2.000 0.3138 0.01241 0.00470 -0.0147 0.6137 1.0000
2.250 0.3344 0.01255 0.00472 -0.0132 0.5820 1.0000
2.500 0.3554 0.01272 0.00479 -0.0118 0.5500 1.0000
2.750 0.3765 0.01292 0.00487 -0.0105 0.5174 1.0000
3.000 0.3979 0.01315 0.00499 -0.0092 0.4843 1.0000
3.250 0.4194 0.01341 0.00514 -0.0079 0.4503 1.0000
3.500 0.4410 0.01370 0.00530 -0.0066 0.4158 1.0000
3.750 0.4626 0.01403 0.00550 -0.0054 0.3811 1.0000
4.000 0.4844 0.01438 0.00574 -0.0042 0.3455 1.0000
4.250 0.5060 0.01478 0.00602 -0.0031 0.3098 1.0000
4.500 0.5275 0.01523 0.00633 -0.0020 0.2750 1.0000
4.750 0.5492 0.01571 0.00669 -0.0009 0.2412 1.0000
5.000 0.5708 0.01623 0.00711 0.0001 0.2090 1.0000
5.250 0.5923 0.01681 0.00757 0.0012 0.1802 1.0000
5.500 0.6136 0.01745 0.00809 0.0022 0.1549 1.0000
5.750 0.6353 0.01809 0.00868 0.0032 0.1336 1.0000
6.000 0.6565 0.01879 0.00934 0.0042 0.1162 1.0000
6.250 0.6777 0.01953 0.01005 0.0052 0.1023 1.0000
6.500 0.6987 0.02031 0.01083 0.0063 0.0911 1.0000
6.750 0.7193 0.02116 0.01167 0.0074 0.0820 1.0000
7.000 0.7399 0.02201 0.01256 0.0084 0.0745 1.0000
7.250 0.7604 0.02296 0.01357 0.0095 0.0682 1.0000
7.500 0.7806 0.02388 0.01451 0.0105 0.0628 1.0000
7.750 0.8004 0.02502 0.01573 0.0117 0.0587 1.0000
8.000 0.8206 0.02605 0.01690 0.0127 0.0543 1.0000
8.250 0.8393 0.02721 0.01803 0.0137 0.0510 1.0000
8.500 0.8589 0.02862 0.01964 0.0148 0.0484 1.0000
8.750 0.8779 0.02999 0.02120 0.0159 0.0457 1.0000
9.000 0.8959 0.03120 0.02248 0.0169 0.0431 1.0000
9.250 0.9127 0.03280 0.02412 0.0179 0.0412 1.0000
9.500 0.9285 0.03477 0.02644 0.0192 0.0396 1.0000
9.750 0.9421 0.03687 0.02885 0.0205 0.0380 1.0000
10.000 0.9542 0.03880 0.03102 0.0217 0.0365 1.0000
10.250 0.9660 0.04039 0.03273 0.0229 0.0350 1.0000
10.500 0.9766 0.04223 0.03462 0.0240 0.0340 1.0000
10.750 0.9771 0.04508 0.03785 0.0256 0.0331 1.0000
11.000 0.9713 0.04826 0.04141 0.0274 0.0325 1.0000
11.250 0.9587 0.05139 0.04484 0.0295 0.0323 1.0000
11.500 0.9416 0.05509 0.04880 0.0301 0.0321 1.0000
11.750 0.9205 0.05980 0.05376 0.0288 0.0320 1.0000
12.000 0.8961 0.06589 0.06008 0.0251 0.0321 1.0000
12.250 0.8690 0.07366 0.06801 0.0193 0.0324 1.0000
12.500 0.8351 0.08426 0.07876 0.0112 0.0328 1.0000
12.750 0.7937 0.09825 0.09283 0.0015 0.0333 1.0000
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Polar data table (+)
Polar graphs
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