USA 41 AIRFOIL (usa41-il) Xfoil prediction polar at RE=200,000 Ncrit=5
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Airfoil: USA 41 AIRFOIL (usa41-il) Reynolds number: 200,000 Max Cl/Cd: 79.89 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa41-il-200000-n5.txt Download as CSV file: xf-usa41-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 41 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3388 0.10212 0.09857 -0.0275 1.0000 0.0198
-8.500 -0.3311 0.09836 0.09482 -0.0261 1.0000 0.0188
-8.250 -0.3357 0.09620 0.09274 -0.0287 1.0000 0.0199
-8.000 -0.3277 0.09280 0.08935 -0.0268 1.0000 0.0188
-7.750 -0.3358 0.09065 0.08730 -0.0279 1.0000 0.0197
-7.500 -0.3334 0.08791 0.08460 -0.0263 1.0000 0.0192
-7.250 -0.3380 0.08569 0.08242 -0.0248 1.0000 0.0188
-7.000 -0.3417 0.08337 0.08017 -0.0239 1.0000 0.0185
-6.750 -0.3461 0.08112 0.07798 -0.0231 1.0000 0.0181
-6.500 -0.3277 0.07657 0.07343 -0.0290 0.9956 0.0178
-6.250 -0.2976 0.07076 0.06759 -0.0384 0.9882 0.0177
-6.000 -0.2663 0.06485 0.06163 -0.0480 0.9803 0.0178
-5.750 -0.2279 0.05716 0.05383 -0.0607 0.9723 0.0193
-5.500 -0.1810 0.04731 0.04374 -0.0751 0.9661 0.0201
-5.250 -0.1502 0.04143 0.03769 -0.0818 0.9587 0.0208
-5.000 -0.1169 0.03895 0.03509 -0.0858 0.9541 0.0220
-4.750 -0.0821 0.03480 0.03069 -0.0905 0.9477 0.0249
-4.500 -0.0425 0.02544 0.02044 -0.0970 0.9416 0.0268
-4.250 -0.0087 0.02165 0.01587 -0.0995 0.9360 0.0296
-4.000 0.0212 0.01977 0.01368 -0.1006 0.9288 0.0313
-3.750 0.0556 0.01821 0.01179 -0.1022 0.9238 0.0330
-3.500 0.0849 0.01718 0.01049 -0.1026 0.9151 0.0355
-3.250 0.1184 0.01600 0.00897 -0.1036 0.9087 0.0367
-3.000 0.1479 0.01511 0.00783 -0.1037 0.8992 0.0374
-2.750 0.1780 0.01436 0.00690 -0.1040 0.8899 0.0380
-2.500 0.2085 0.01385 0.00623 -0.1043 0.8804 0.0393
-2.250 0.2358 0.01307 0.00534 -0.1040 0.8686 0.0401
-2.000 0.2631 0.01257 0.00477 -0.1037 0.8561 0.0403
-1.750 0.2903 0.01214 0.00428 -0.1033 0.8428 0.0406
-1.500 0.3172 0.01180 0.00384 -0.1028 0.8283 0.0412
-1.250 0.3442 0.01152 0.00348 -0.1024 0.8128 0.0420
-1.000 0.3712 0.01130 0.00316 -0.1019 0.7961 0.0433
-0.750 0.3978 0.01113 0.00289 -0.1014 0.7779 0.0451
-0.500 0.4244 0.01100 0.00265 -0.1008 0.7587 0.0474
-0.250 0.4509 0.01092 0.00245 -0.1002 0.7387 0.0507
0.000 0.4769 0.01080 0.00232 -0.0996 0.7182 0.0633
0.250 0.5023 0.01063 0.00240 -0.0991 0.6973 0.1635
0.500 0.5282 0.01072 0.00240 -0.0984 0.6764 0.1892
0.750 0.5539 0.01079 0.00239 -0.0978 0.6554 0.2015
1.000 0.5794 0.01087 0.00238 -0.0972 0.6350 0.2158
1.500 0.6302 0.01099 0.00242 -0.0960 0.5966 0.2478
1.750 0.6554 0.01099 0.00249 -0.0955 0.5793 0.2875
2.000 0.6947 0.00979 0.00264 -0.0982 0.5620 1.0000
2.250 0.7202 0.01000 0.00274 -0.0976 0.5470 1.0000
2.500 0.7457 0.01021 0.00287 -0.0970 0.5328 1.0000
2.750 0.7712 0.01042 0.00300 -0.0965 0.5193 1.0000
3.000 0.7967 0.01064 0.00316 -0.0959 0.5063 1.0000
3.250 0.8221 0.01085 0.00335 -0.0954 0.4937 1.0000
3.500 0.8475 0.01108 0.00353 -0.0948 0.4812 1.0000
3.750 0.8727 0.01131 0.00374 -0.0943 0.4686 1.0000
4.000 0.8978 0.01155 0.00397 -0.0937 0.4561 1.0000
4.500 0.9479 0.01204 0.00445 -0.0925 0.4308 1.0000
4.750 0.9727 0.01229 0.00472 -0.0919 0.4173 1.0000
5.000 0.9974 0.01255 0.00502 -0.0913 0.4038 1.0000
5.250 1.0221 0.01282 0.00533 -0.0907 0.3921 1.0000
5.500 1.0465 0.01310 0.00566 -0.0900 0.3790 1.0000
5.750 1.0689 0.01347 0.00596 -0.0890 0.3508 1.0000
6.000 1.0894 0.01398 0.00632 -0.0878 0.3086 1.0000
6.250 1.1101 0.01454 0.00674 -0.0866 0.2679 1.0000
6.500 1.1274 0.01548 0.00732 -0.0851 0.1982 1.0000
6.750 1.1372 0.01738 0.00849 -0.0829 0.0859 1.0000
7.000 1.1501 0.01895 0.00972 -0.0808 0.0336 1.0000
7.250 1.1680 0.01993 0.01073 -0.0792 0.0241 1.0000
7.500 1.1851 0.02095 0.01191 -0.0775 0.0205 1.0000
7.750 1.2028 0.02185 0.01297 -0.0759 0.0185 1.0000
8.000 1.2192 0.02279 0.01405 -0.0742 0.0165 1.0000
8.250 1.2335 0.02388 0.01524 -0.0723 0.0148 1.0000
8.500 1.2432 0.02531 0.01680 -0.0697 0.0138 1.0000
8.750 1.2478 0.02713 0.01875 -0.0665 0.0130 1.0000
9.000 1.2532 0.02879 0.02052 -0.0634 0.0125 1.0000
9.250 1.2624 0.03004 0.02189 -0.0608 0.0120 1.0000
9.500 1.2711 0.03141 0.02343 -0.0583 0.0114 1.0000
9.750 1.2798 0.03276 0.02490 -0.0560 0.0107 1.0000
10.000 1.2877 0.03429 0.02655 -0.0538 0.0101 1.0000
10.250 1.2948 0.03605 0.02843 -0.0517 0.0097 1.0000
10.500 1.3017 0.03795 0.03047 -0.0497 0.0094 1.0000
10.750 1.3079 0.03993 0.03258 -0.0478 0.0091 1.0000
11.000 1.3135 0.04213 0.03493 -0.0460 0.0089 1.0000
11.250 1.3174 0.04465 0.03759 -0.0444 0.0087 1.0000
11.500 1.3201 0.04750 0.04062 -0.0428 0.0085 1.0000
11.750 1.3200 0.05095 0.04429 -0.0413 0.0084 1.0000
12.000 1.3169 0.05454 0.04812 -0.0400 0.0083 1.0000
12.250 1.3112 0.05788 0.05172 -0.0390 0.0082 1.0000
12.500 1.3035 0.06146 0.05557 -0.0386 0.0081 1.0000
12.750 1.2941 0.06556 0.05996 -0.0387 0.0080 1.0000
13.000 1.2838 0.06988 0.06453 -0.0394 0.0079 1.0000
13.250 1.2714 0.07486 0.06976 -0.0407 0.0078 1.0000
13.500 1.2576 0.08037 0.07550 -0.0426 0.0077 1.0000
13.750 1.2432 0.08628 0.08163 -0.0451 0.0077 1.0000
14.000 1.2277 0.09275 0.08831 -0.0483 0.0076 1.0000
14.250 1.2121 0.09964 0.09538 -0.0519 0.0077 1.0000
14.500 1.1955 0.10716 0.10309 -0.0561 0.0077 1.0000
14.750 1.1795 0.11498 0.11108 -0.0608 0.0077 1.0000
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Polar data table (+)
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