USA 51 AIRFOIL (usa51-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 51 AIRFOIL (usa51-il) Reynolds number: 100,000 Max Cl/Cd: 48.85 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa51-il-100000-n5.txt Download as CSV file: xf-usa51-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 51 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4255 0.09901 0.09406 -0.0247 1.0000 0.0344
-9.000 -0.4240 0.09530 0.09039 -0.0252 1.0000 0.0320
-8.750 -0.4263 0.09119 0.08635 -0.0266 1.0000 0.0302
-8.250 -0.4456 0.08018 0.07551 -0.0339 1.0000 0.0252
-8.000 -0.4542 0.07648 0.07187 -0.0347 1.0000 0.0249
-7.750 -0.4595 0.07240 0.06782 -0.0358 1.0000 0.0245
-7.500 -0.4640 0.06822 0.06364 -0.0366 1.0000 0.0243
-7.250 -0.4678 0.06390 0.05928 -0.0371 1.0000 0.0241
-7.000 -0.4699 0.05962 0.05493 -0.0371 1.0000 0.0238
-6.750 -0.4705 0.05522 0.05039 -0.0367 1.0000 0.0236
-6.500 -0.4688 0.05099 0.04598 -0.0359 1.0000 0.0234
-6.250 -0.4656 0.04660 0.04133 -0.0347 1.0000 0.0234
-6.000 -0.4598 0.04243 0.03684 -0.0332 1.0000 0.0234
-5.750 -0.4513 0.03855 0.03253 -0.0315 1.0000 0.0240
-5.500 -0.4396 0.03529 0.02872 -0.0296 1.0000 0.0250
-5.250 -0.4264 0.03235 0.02527 -0.0277 1.0000 0.0259
-5.000 -0.4121 0.02941 0.02196 -0.0261 1.0000 0.0267
-4.750 -0.3953 0.02722 0.01947 -0.0246 1.0000 0.0278
-4.500 -0.3766 0.02546 0.01741 -0.0231 1.0000 0.0290
-4.250 -0.3417 0.02389 0.01548 -0.0247 0.9950 0.0330
-4.000 -0.3019 0.02213 0.01324 -0.0268 0.9882 0.0361
-3.750 -0.2643 0.02042 0.01145 -0.0289 0.9806 0.0409
-3.500 -0.2255 0.01929 0.01015 -0.0310 0.9719 0.0472
-3.250 -0.1879 0.01826 0.00910 -0.0331 0.9636 0.0552
-3.000 -0.1535 0.01748 0.00818 -0.0343 0.9543 0.0629
-2.750 -0.1211 0.01684 0.00753 -0.0352 0.9447 0.0764
-2.500 -0.0855 0.01607 0.00684 -0.0367 0.9367 0.1022
-2.250 -0.0547 0.01530 0.00635 -0.0374 0.9257 0.1735
-2.000 -0.0250 0.01456 0.00598 -0.0379 0.9138 0.2769
-1.750 0.0001 0.01318 0.00580 -0.0375 0.9018 0.5672
-1.500 0.0754 0.01248 0.00601 -0.0447 0.9010 0.9414
-1.250 0.1426 0.01241 0.00570 -0.0524 0.8926 0.9895
-1.000 0.1857 0.01229 0.00536 -0.0555 0.8757 1.0000
-0.750 0.2215 0.01213 0.00501 -0.0569 0.8554 1.0000
-0.500 0.2564 0.01200 0.00469 -0.0581 0.8326 1.0000
-0.250 0.2908 0.01191 0.00440 -0.0591 0.8081 1.0000
0.000 0.3221 0.01189 0.00418 -0.0595 0.7821 1.0000
0.250 0.3499 0.01193 0.00404 -0.0592 0.7549 1.0000
0.500 0.3759 0.01202 0.00395 -0.0586 0.7272 1.0000
0.750 0.4007 0.01214 0.00391 -0.0578 0.6992 1.0000
1.000 0.4246 0.01228 0.00389 -0.0568 0.6698 1.0000
1.250 0.4477 0.01246 0.00389 -0.0557 0.6388 1.0000
1.750 0.4928 0.01288 0.00395 -0.0533 0.5769 1.0000
2.000 0.5150 0.01313 0.00401 -0.0521 0.5484 1.0000
2.250 0.5372 0.01340 0.00410 -0.0509 0.5236 1.0000
2.500 0.5597 0.01367 0.00426 -0.0498 0.5029 1.0000
2.750 0.5824 0.01395 0.00442 -0.0488 0.4850 1.0000
3.000 0.6054 0.01423 0.00462 -0.0478 0.4696 1.0000
3.250 0.6286 0.01450 0.00487 -0.0469 0.4559 1.0000
3.500 0.6517 0.01479 0.00511 -0.0460 0.4426 1.0000
3.750 0.6747 0.01508 0.00537 -0.0451 0.4282 1.0000
4.000 0.6974 0.01537 0.00566 -0.0441 0.4123 1.0000
4.250 0.7199 0.01565 0.00594 -0.0431 0.3957 1.0000
4.500 0.7428 0.01593 0.00625 -0.0422 0.3807 1.0000
4.750 0.7656 0.01620 0.00657 -0.0412 0.3651 1.0000
5.000 0.7882 0.01648 0.00692 -0.0403 0.3482 1.0000
5.250 0.8107 0.01677 0.00727 -0.0393 0.3292 1.0000
5.500 0.8328 0.01709 0.00764 -0.0382 0.3051 1.0000
5.750 0.8539 0.01748 0.00803 -0.0371 0.2749 1.0000
6.000 0.8740 0.01800 0.00851 -0.0358 0.2392 1.0000
6.500 0.9093 0.01965 0.00973 -0.0329 0.1713 1.0000
6.750 0.9274 0.02047 0.01045 -0.0316 0.1513 1.0000
7.000 0.9451 0.02132 0.01122 -0.0303 0.1335 1.0000
7.250 0.9643 0.02202 0.01200 -0.0291 0.1174 1.0000
7.500 0.9842 0.02268 0.01268 -0.0280 0.0932 1.0000
7.750 0.9964 0.02405 0.01366 -0.0260 0.0368 1.0000
8.000 1.0075 0.02551 0.01521 -0.0235 0.0312 1.0000
8.250 1.0201 0.02680 0.01668 -0.0213 0.0276 1.0000
8.500 1.0331 0.02800 0.01810 -0.0193 0.0256 1.0000
8.750 1.0438 0.02932 0.01964 -0.0170 0.0241 1.0000
9.000 1.0514 0.03082 0.02133 -0.0144 0.0227 1.0000
9.250 1.0539 0.03250 0.02318 -0.0114 0.0214 1.0000
9.500 1.0498 0.03447 0.02531 -0.0078 0.0203 1.0000
9.750 1.0478 0.03641 0.02735 -0.0047 0.0196 1.0000
10.000 1.0515 0.03805 0.02921 -0.0025 0.0191 1.0000
10.250 1.0541 0.03991 0.03121 -0.0005 0.0188 1.0000
10.500 1.0572 0.04192 0.03335 0.0013 0.0184 1.0000
10.750 1.0615 0.04400 0.03555 0.0027 0.0181 1.0000
11.000 1.0665 0.04623 0.03791 0.0040 0.0177 1.0000
11.250 1.0713 0.04850 0.04033 0.0050 0.0172 1.0000
11.500 1.0754 0.05098 0.04298 0.0059 0.0166 1.0000
11.750 1.0772 0.05367 0.04584 0.0065 0.0159 1.0000
12.000 1.0787 0.05654 0.04889 0.0068 0.0155 1.0000
12.250 1.0779 0.05968 0.05220 0.0068 0.0150 1.0000
12.500 1.0762 0.06313 0.05585 0.0067 0.0147 1.0000
12.750 1.0728 0.06685 0.05977 0.0061 0.0147 1.0000
13.000 1.0670 0.07097 0.06409 0.0050 0.0146 1.0000
13.250 1.0591 0.07552 0.06886 0.0035 0.0145 1.0000
13.500 1.0491 0.08058 0.07413 0.0013 0.0144 1.0000
13.750 1.0373 0.08613 0.07988 -0.0013 0.0144 1.0000
14.000 1.0246 0.09205 0.08601 -0.0046 0.0145 1.0000
14.250 1.0105 0.09853 0.09267 -0.0083 0.0146 1.0000
14.500 0.9953 0.10559 0.09991 -0.0125 0.0147 1.0000
14.750 0.9787 0.11336 0.10786 -0.0173 0.0148 1.0000
15.000 0.9625 0.12155 0.11620 -0.0225 0.0151 1.0000
15.250 0.9452 0.13066 0.12545 -0.0281 0.0153 1.0000
15.500 0.9261 0.14104 0.13594 -0.0344 0.0155 1.0000
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Polar data table (+)
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