TSAGI 8% AIRFOIL (tsagi8-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: TSAGI 8% AIRFOIL (tsagi8-il) Reynolds number: 200,000 Max Cl/Cd: 57.66 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-tsagi8-il-200000.txt Download as CSV file: xf-tsagi8-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: TSAGI 8% 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.5926 0.07059 0.06686 -0.0227 1.0000 0.0409
-7.500 -0.5940 0.06388 0.06025 -0.0225 1.0000 0.0423
-7.250 -0.5853 0.06150 0.05792 -0.0212 1.0000 0.0442
-7.000 -0.5780 0.05887 0.05522 -0.0205 1.0000 0.0483
-6.750 -0.5741 0.05606 0.05186 -0.0197 1.0000 0.0529
-6.250 -0.5620 0.04699 0.04273 -0.0175 1.0000 0.0565
-6.000 -0.5504 0.04439 0.04004 -0.0159 1.0000 0.0593
-5.500 -0.5314 0.03831 0.03341 -0.0115 1.0000 0.0694
-5.250 -0.5173 0.03610 0.03111 -0.0096 1.0000 0.0733
-5.000 -0.5046 0.02770 0.02154 -0.0035 1.0000 0.0380
-4.750 -0.4899 0.02425 0.01763 -0.0005 1.0000 0.0353
-4.500 -0.4720 0.02204 0.01502 0.0019 1.0000 0.0357
-4.250 -0.4528 0.02056 0.01325 0.0040 1.0000 0.0381
-4.000 -0.4326 0.01880 0.01120 0.0060 1.0000 0.0385
-3.750 -0.4122 0.01748 0.00967 0.0079 1.0000 0.0399
-3.500 -0.3929 0.01585 0.00792 0.0098 1.0000 0.0425
-3.250 -0.3761 0.01512 0.00723 0.0116 1.0000 0.0482
-3.000 -0.3596 0.01449 0.00652 0.0138 1.0000 0.0527
-2.750 -0.3319 0.01350 0.00557 0.0135 0.9965 0.0626
-2.500 -0.2910 0.01264 0.00480 0.0106 0.9884 0.0881
-2.250 -0.2586 0.01080 0.00413 0.0087 0.9794 0.3198
-2.000 -0.1987 0.00920 0.00476 0.0041 0.9925 0.9322
-1.750 -0.0775 0.00973 0.00491 -0.0128 1.0000 0.9878
-1.500 -0.0015 0.00954 0.00450 -0.0227 1.0000 1.0000
-1.250 0.0056 0.00963 0.00458 -0.0196 1.0000 1.0000
-1.000 0.0471 0.00952 0.00440 -0.0231 0.9858 1.0000
-0.750 0.0986 0.00938 0.00417 -0.0284 0.9707 1.0000
-0.500 0.1492 0.00922 0.00395 -0.0334 0.9518 1.0000
-0.250 0.1981 0.00907 0.00373 -0.0379 0.9283 1.0000
0.000 0.2372 0.00897 0.00355 -0.0403 0.8996 1.0000
0.250 0.2643 0.00896 0.00343 -0.0400 0.8704 1.0000
0.500 0.2868 0.00900 0.00336 -0.0387 0.8441 1.0000
0.750 0.3083 0.00906 0.00333 -0.0373 0.8201 1.0000
1.000 0.3300 0.00914 0.00331 -0.0359 0.7991 1.0000
1.250 0.3517 0.00921 0.00334 -0.0346 0.7783 1.0000
1.500 0.3738 0.00930 0.00335 -0.0333 0.7596 1.0000
1.750 0.3961 0.00938 0.00340 -0.0321 0.7400 1.0000
2.000 0.4185 0.00948 0.00345 -0.0309 0.7213 1.0000
2.250 0.4411 0.00958 0.00354 -0.0297 0.7037 1.0000
2.500 0.4638 0.00968 0.00363 -0.0286 0.6844 1.0000
2.750 0.4864 0.00979 0.00372 -0.0275 0.6652 1.0000
3.000 0.5092 0.00991 0.00382 -0.0263 0.6466 1.0000
3.250 0.5318 0.01002 0.00395 -0.0252 0.6250 1.0000
3.500 0.5534 0.01011 0.00404 -0.0237 0.5965 1.0000
3.750 0.5737 0.01021 0.00405 -0.0220 0.5550 1.0000
4.000 0.5938 0.01038 0.00408 -0.0202 0.5053 1.0000
4.250 0.6141 0.01065 0.00422 -0.0187 0.4531 1.0000
4.500 0.6319 0.01121 0.00444 -0.0168 0.3566 1.0000
4.750 0.6348 0.01409 0.00566 -0.0134 0.0628 1.0000
5.000 0.6514 0.01531 0.00689 -0.0113 0.0467 1.0000
5.250 0.6703 0.01613 0.00782 -0.0096 0.0416 1.0000
5.500 0.6863 0.01733 0.00901 -0.0076 0.0360 1.0000
5.750 0.7027 0.01878 0.01051 -0.0054 0.0341 1.0000
6.000 0.7222 0.02005 0.01187 -0.0036 0.0326 1.0000
6.250 0.7426 0.02162 0.01355 -0.0019 0.0317 1.0000
6.500 0.7637 0.02327 0.01540 -0.0004 0.0307 1.0000
6.750 0.7834 0.02451 0.01675 0.0010 0.0277 1.0000
7.000 0.8037 0.02675 0.01924 0.0027 0.0277 1.0000
7.250 0.8218 0.02983 0.02280 0.0050 0.0292 1.0000
7.500 0.8358 0.03378 0.02719 0.0076 0.0321 1.0000
13.250 0.5812 0.13761 0.13411 0.0026 0.0466 1.0000
13.500 0.5756 0.14144 0.13794 0.0001 0.0464 1.0000
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