USA 27 AIRFOIL (usa27-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 27 AIRFOIL (usa27-il) Reynolds number: 500,000 Max Cl/Cd: 92.03 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa27-il-500000-n5.txt Download as CSV file: xf-usa27-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 27 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.2608 0.10114 0.09881 -0.0513 0.9910 0.0142
-9.750 -0.2526 0.09688 0.09454 -0.0546 0.9888 0.0146
-9.500 -0.2529 0.08946 0.08713 -0.0610 0.9861 0.0157
-9.250 -0.2362 0.08721 0.08488 -0.0634 0.9844 0.0160
-9.000 -0.2241 0.08468 0.08235 -0.0656 0.9802 0.0163
-8.750 -0.2111 0.08124 0.07891 -0.0693 0.9770 0.0168
-8.500 -0.1984 0.07692 0.07459 -0.0744 0.9742 0.0171
-8.250 -0.1933 0.07274 0.07042 -0.0784 0.9654 0.0175
-8.000 -0.1870 0.06788 0.06554 -0.0842 0.9561 0.0178
-7.750 -0.1876 0.05608 0.05361 -0.0968 0.9436 0.0193
-7.500 -0.1693 0.05389 0.05137 -0.0988 0.9370 0.0196
-7.250 -0.1519 0.05110 0.04852 -0.1010 0.9282 0.0198
-7.000 -0.1326 0.04742 0.04473 -0.1040 0.9189 0.0201
-6.750 -0.1170 0.04313 0.04028 -0.1061 0.9064 0.0205
-6.500 -0.1021 0.03715 0.03405 -0.1083 0.8941 0.0210
-6.250 -0.0928 0.02968 0.02603 -0.1088 0.8823 0.0221
-6.000 -0.0801 0.02532 0.02110 -0.1076 0.8699 0.0225
-5.750 -0.0624 0.02276 0.01807 -0.1063 0.8569 0.0229
-5.500 -0.0425 0.02122 0.01605 -0.1049 0.8442 0.0235
-5.250 -0.0223 0.01956 0.01421 -0.1040 0.8329 0.0240
-5.000 0.0005 0.01854 0.01301 -0.1033 0.8217 0.0242
-4.750 0.0239 0.01765 0.01195 -0.1026 0.8100 0.0244
-4.500 0.0474 0.01684 0.01094 -0.1017 0.7965 0.0247
-4.250 0.0709 0.01610 0.01003 -0.1009 0.7806 0.0249
-4.000 0.0945 0.01544 0.00917 -0.1000 0.7618 0.0251
-3.750 0.1178 0.01484 0.00837 -0.0990 0.7386 0.0253
-3.500 0.1404 0.01435 0.00765 -0.0978 0.7100 0.0255
-3.250 0.1624 0.01397 0.00704 -0.0964 0.6804 0.0258
-2.500 0.2327 0.01300 0.00557 -0.0935 0.6270 0.0267
-2.250 0.2575 0.01266 0.00513 -0.0928 0.6162 0.0269
-2.000 0.2822 0.01237 0.00475 -0.0921 0.6063 0.0271
-1.750 0.3072 0.01210 0.00442 -0.0914 0.5964 0.0274
-1.500 0.3322 0.01188 0.00413 -0.0908 0.5865 0.0277
-1.250 0.3568 0.01171 0.00389 -0.0900 0.5759 0.0280
-1.000 0.3819 0.01154 0.00368 -0.0894 0.5644 0.0282
-0.750 0.4067 0.01141 0.00350 -0.0888 0.5529 0.0284
-0.500 0.4313 0.01130 0.00334 -0.0881 0.5407 0.0286
-0.250 0.4550 0.01110 0.00310 -0.0872 0.5266 0.0290
0.000 0.4784 0.01096 0.00290 -0.0863 0.5082 0.0298
0.250 0.5016 0.01094 0.00279 -0.0854 0.4870 0.0306
0.750 0.5480 0.01102 0.00267 -0.0835 0.4484 0.0316
1.000 0.5718 0.01108 0.00264 -0.0827 0.4330 0.0321
1.500 0.6207 0.01119 0.00263 -0.0813 0.4104 0.0336
1.750 0.6453 0.01127 0.00265 -0.0807 0.4021 0.0345
2.000 0.6704 0.01132 0.00267 -0.0802 0.3942 0.0354
2.250 0.6950 0.01141 0.00272 -0.0795 0.3864 0.0376
2.750 0.7422 0.01122 0.00294 -0.0781 0.3725 0.2275
3.000 0.7667 0.01120 0.00307 -0.0776 0.3667 0.2932
3.250 0.7908 0.01120 0.00321 -0.0770 0.3599 0.3568
3.750 0.8842 0.01041 0.00384 -0.0857 0.3456 0.9902
4.000 0.9427 0.01069 0.00405 -0.0928 0.3387 1.0000
4.250 0.9641 0.01082 0.00416 -0.0916 0.3337 1.0000
4.500 0.9850 0.01096 0.00430 -0.0902 0.3292 1.0000
4.750 1.0051 0.01113 0.00444 -0.0887 0.3248 1.0000
5.000 1.0260 0.01127 0.00458 -0.0874 0.3204 1.0000
5.250 1.0466 0.01142 0.00473 -0.0860 0.3152 1.0000
5.500 1.0657 0.01163 0.00490 -0.0844 0.3085 1.0000
5.750 1.0859 0.01180 0.00505 -0.0829 0.2997 1.0000
6.000 1.1047 0.01202 0.00524 -0.0812 0.2898 1.0000
6.250 1.1235 0.01224 0.00543 -0.0796 0.2793 1.0000
6.500 1.1421 0.01247 0.00563 -0.0779 0.2673 1.0000
6.750 1.1597 0.01276 0.00585 -0.0760 0.2526 1.0000
7.000 1.1740 0.01319 0.00615 -0.0736 0.2258 1.0000
7.250 1.1820 0.01392 0.00662 -0.0701 0.1868 1.0000
7.500 1.1938 0.01451 0.00709 -0.0674 0.1673 1.0000
7.750 1.2087 0.01499 0.00751 -0.0652 0.1552 1.0000
8.000 1.2238 0.01549 0.00796 -0.0632 0.1421 1.0000
8.250 1.2369 0.01613 0.00849 -0.0609 0.1221 1.0000
8.500 1.2328 0.01779 0.00972 -0.0560 0.0597 1.0000
8.750 1.2335 0.01921 0.01097 -0.0520 0.0190 1.0000
9.000 1.2488 0.01978 0.01159 -0.0503 0.0168 1.0000
9.250 1.2630 0.02043 0.01229 -0.0485 0.0153 1.0000
9.500 1.2764 0.02115 0.01308 -0.0466 0.0139 1.0000
9.750 1.2910 0.02180 0.01380 -0.0450 0.0133 1.0000
10.000 1.3045 0.02253 0.01460 -0.0433 0.0125 1.0000
10.250 1.3170 0.02335 0.01549 -0.0416 0.0118 1.0000
10.500 1.3282 0.02428 0.01649 -0.0398 0.0113 1.0000
10.750 1.3373 0.02538 0.01767 -0.0379 0.0108 1.0000
11.000 1.3430 0.02676 0.01916 -0.0358 0.0103 1.0000
11.250 1.3528 0.02788 0.02036 -0.0342 0.0101 1.0000
11.500 1.3614 0.02914 0.02169 -0.0326 0.0097 1.0000
11.750 1.3689 0.03053 0.02317 -0.0311 0.0094 1.0000
12.000 1.3752 0.03206 0.02479 -0.0297 0.0090 1.0000
12.250 1.3804 0.03376 0.02657 -0.0283 0.0087 1.0000
12.500 1.3842 0.03566 0.02855 -0.0271 0.0084 1.0000
12.750 1.3864 0.03779 0.03078 -0.0260 0.0082 1.0000
13.000 1.3869 0.04020 0.03328 -0.0251 0.0080 1.0000
13.250 1.3852 0.04300 0.03619 -0.0244 0.0078 1.0000
13.500 1.3801 0.04635 0.03965 -0.0239 0.0077 1.0000
13.750 1.3713 0.05031 0.04374 -0.0238 0.0075 1.0000
14.000 1.3683 0.05371 0.04726 -0.0239 0.0075 1.0000
14.250 1.3637 0.05739 0.05106 -0.0241 0.0074 1.0000
14.500 1.3574 0.06135 0.05513 -0.0245 0.0073 1.0000
14.750 1.3501 0.06555 0.05945 -0.0251 0.0072 1.0000
15.000 1.3419 0.06999 0.06400 -0.0258 0.0071 1.0000
15.250 1.3329 0.07460 0.06873 -0.0267 0.0070 1.0000
15.500 1.3236 0.07934 0.07359 -0.0277 0.0069 1.0000
15.750 1.3146 0.08412 0.07848 -0.0287 0.0069 1.0000
16.000 1.3059 0.08892 0.08338 -0.0299 0.0068 1.0000
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Polar data table (+)
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