RAE 5215 AIRFOIL (rae5215-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAE 5215 AIRFOIL (rae5215-il) Reynolds number: 200,000 Max Cl/Cd: 49.81 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae5215-il-200000-n5.txt Download as CSV file: xf-rae5215-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAE 5215 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.6050 0.08988 0.08597 -0.0171 1.0000 0.0172
-9.750 -0.6142 0.08229 0.07841 -0.0222 1.0000 0.0170
-9.500 -0.6372 0.06946 0.06555 -0.0332 1.0000 0.0166
-9.250 -0.6666 0.06017 0.05612 -0.0413 1.0000 0.0163
-9.000 -0.6868 0.05268 0.04833 -0.0460 1.0000 0.0162
-8.750 -0.6969 0.04581 0.04104 -0.0480 1.0000 0.0162
-8.500 -0.6980 0.03975 0.03446 -0.0488 1.0000 0.0163
-8.250 -0.6910 0.03476 0.02890 -0.0488 1.0000 0.0167
-8.000 -0.6789 0.03034 0.02380 -0.0484 1.0000 0.0173
-7.750 -0.6613 0.02768 0.02072 -0.0479 1.0000 0.0178
-7.500 -0.6406 0.02670 0.01968 -0.0473 1.0000 0.0183
-7.250 -0.6198 0.02563 0.01850 -0.0467 1.0000 0.0192
-7.000 -0.5990 0.02399 0.01648 -0.0459 1.0000 0.0204
-6.750 -0.5780 0.02277 0.01513 -0.0451 1.0000 0.0212
-6.500 -0.5569 0.02187 0.01417 -0.0442 1.0000 0.0220
-6.250 -0.5358 0.02093 0.01311 -0.0433 1.0000 0.0230
-6.000 -0.5145 0.01996 0.01196 -0.0423 1.0000 0.0244
-5.750 -0.4918 0.01923 0.01124 -0.0417 0.9996 0.0254
-5.500 -0.4591 0.01828 0.01023 -0.0430 0.9967 0.0266
-5.250 -0.4261 0.01739 0.00922 -0.0442 0.9941 0.0280
-5.000 -0.3942 0.01664 0.00853 -0.0455 0.9906 0.0294
-4.750 -0.3608 0.01606 0.00789 -0.0469 0.9874 0.0317
-4.500 -0.3265 0.01543 0.00730 -0.0486 0.9848 0.0338
-4.250 -0.2938 0.01493 0.00677 -0.0498 0.9808 0.0364
-4.000 -0.2603 0.01438 0.00625 -0.0513 0.9773 0.0388
-3.750 -0.2257 0.01392 0.00576 -0.0529 0.9746 0.0418
-3.500 -0.1919 0.01349 0.00535 -0.0544 0.9713 0.0464
-3.250 -0.1596 0.01309 0.00496 -0.0554 0.9666 0.0527
-3.000 -0.1256 0.01265 0.00458 -0.0569 0.9630 0.0660
-2.750 -0.0910 0.01205 0.00423 -0.0587 0.9602 0.1182
-2.500 -0.0601 0.01092 0.00387 -0.0603 0.9554 0.2986
-2.250 -0.0292 0.01014 0.00389 -0.0614 0.9508 0.4971
-2.000 0.0029 0.01009 0.00407 -0.0620 0.9468 0.5777
-1.750 0.0330 0.01021 0.00431 -0.0621 0.9416 0.6272
-1.500 0.0629 0.01034 0.00446 -0.0620 0.9339 0.6576
-1.250 0.0935 0.01031 0.00444 -0.0620 0.9247 0.6695
-1.000 0.1239 0.01024 0.00430 -0.0621 0.9140 0.6759
-0.750 0.1522 0.01018 0.00423 -0.0617 0.9028 0.6805
-0.500 0.1806 0.01011 0.00412 -0.0613 0.8898 0.6858
-0.250 0.2075 0.01004 0.00397 -0.0604 0.8691 0.6914
0.000 0.2336 0.00997 0.00384 -0.0593 0.8458 0.6958
0.250 0.2595 0.00993 0.00373 -0.0583 0.8192 0.7006
0.500 0.2858 0.00993 0.00362 -0.0574 0.7879 0.7058
0.750 0.3129 0.00995 0.00356 -0.0568 0.7567 0.7110
1.000 0.3392 0.01000 0.00351 -0.0560 0.7169 0.7152
1.250 0.3635 0.01024 0.00340 -0.0547 0.6254 0.7200
1.500 0.3829 0.01125 0.00349 -0.0529 0.4340 0.7252
1.750 0.4051 0.01218 0.00376 -0.0522 0.2901 0.7298
2.000 0.4299 0.01276 0.00400 -0.0518 0.2165 0.7344
2.250 0.4561 0.01316 0.00422 -0.0516 0.1800 0.7398
2.500 0.4830 0.01348 0.00443 -0.0514 0.1604 0.7454
2.750 0.5093 0.01375 0.00466 -0.0511 0.1480 0.7501
3.000 0.5358 0.01405 0.00491 -0.0508 0.1383 0.7559
3.250 0.5628 0.01431 0.00517 -0.0507 0.1309 0.7624
3.500 0.5889 0.01459 0.00545 -0.0503 0.1244 0.7680
3.750 0.6152 0.01490 0.00576 -0.0500 0.1190 0.7745
4.000 0.6417 0.01516 0.00606 -0.0497 0.1137 0.7816
4.250 0.6672 0.01551 0.00641 -0.0493 0.1087 0.7896
4.500 0.6933 0.01581 0.00676 -0.0489 0.1042 0.7984
4.750 0.7189 0.01606 0.00708 -0.0484 0.0992 0.8079
5.000 0.7440 0.01643 0.00744 -0.0479 0.0943 0.8190
5.250 0.7696 0.01663 0.00775 -0.0474 0.0890 0.8334
5.500 0.7935 0.01689 0.00803 -0.0466 0.0842 0.8537
5.750 0.8163 0.01703 0.00835 -0.0453 0.0794 0.8917
6.000 0.8424 0.01725 0.00861 -0.0450 0.0750 1.0000
6.250 0.8692 0.01766 0.00905 -0.0449 0.0703 1.0000
6.500 0.8955 0.01808 0.00944 -0.0448 0.0662 1.0000
6.750 0.9217 0.01854 0.00995 -0.0446 0.0619 1.0000
7.000 0.9473 0.01902 0.01038 -0.0444 0.0584 1.0000
7.250 0.9727 0.01956 0.01099 -0.0441 0.0546 1.0000
7.500 0.9974 0.02015 0.01154 -0.0438 0.0520 1.0000
7.750 1.0220 0.02081 0.01229 -0.0433 0.0491 1.0000
8.000 1.0461 0.02145 0.01296 -0.0429 0.0468 1.0000
8.250 1.0696 0.02221 0.01377 -0.0423 0.0447 1.0000
8.500 1.0930 0.02295 0.01458 -0.0418 0.0426 1.0000
8.750 1.1157 0.02371 0.01536 -0.0412 0.0411 1.0000
9.000 1.1378 0.02463 0.01637 -0.0405 0.0397 1.0000
9.250 1.1596 0.02558 0.01745 -0.0398 0.0385 1.0000
9.500 1.1809 0.02652 0.01849 -0.0390 0.0374 1.0000
9.750 1.2015 0.02747 0.01950 -0.0382 0.0365 1.0000
10.000 1.2210 0.02853 0.02058 -0.0373 0.0358 1.0000
10.250 1.2402 0.02977 0.02203 -0.0363 0.0350 1.0000
10.500 1.2584 0.03102 0.02348 -0.0352 0.0341 1.0000
10.750 1.2757 0.03226 0.02487 -0.0340 0.0333 1.0000
11.000 1.2920 0.03347 0.02620 -0.0328 0.0327 1.0000
11.250 1.3073 0.03471 0.02754 -0.0316 0.0321 1.0000
11.500 1.3210 0.03604 0.02895 -0.0302 0.0317 1.0000
11.750 1.3327 0.03759 0.03061 -0.0286 0.0314 1.0000
12.000 1.3398 0.03959 0.03290 -0.0266 0.0310 1.0000
12.250 1.3418 0.04165 0.03523 -0.0240 0.0307 1.0000
12.500 1.3402 0.04392 0.03775 -0.0214 0.0304 1.0000
12.750 1.3357 0.04648 0.04057 -0.0190 0.0301 1.0000
13.000 1.3282 0.04942 0.04376 -0.0171 0.0299 1.0000
13.250 1.3174 0.05287 0.04747 -0.0159 0.0297 1.0000
13.500 1.3027 0.05703 0.05189 -0.0155 0.0296 1.0000
13.750 1.2832 0.06225 0.05737 -0.0164 0.0295 1.0000
14.000 1.2561 0.06935 0.06477 -0.0195 0.0294 1.0000
14.250 1.2085 0.08209 0.07788 -0.0282 0.0296 1.0000
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Polar data table (+)
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