BOEING-VERTOL VR-15 AIRFOIL (vr15-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: BOEING-VERTOL VR-15 AIRFOIL (vr15-il) Reynolds number: 500,000 Max Cl/Cd: 69.23 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-vr15-il-500000-n5.txt Download as CSV file: xf-vr15-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL VR-15 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.6487 0.09396 0.09180 0.0126 1.0000 0.0054
-9.000 -0.8635 0.02734 0.02318 -0.0096 1.0000 0.0049
-8.750 -0.8504 0.02385 0.01922 -0.0081 1.0000 0.0051
-8.500 -0.8307 0.02198 0.01709 -0.0072 1.0000 0.0054
-8.250 -0.8089 0.02058 0.01547 -0.0065 1.0000 0.0056
-8.000 -0.7861 0.01936 0.01406 -0.0059 1.0000 0.0060
-7.750 -0.7629 0.01821 0.01272 -0.0052 1.0000 0.0064
-7.500 -0.7393 0.01712 0.01143 -0.0046 1.0000 0.0070
-7.250 -0.7153 0.01619 0.01035 -0.0040 1.0000 0.0076
-7.000 -0.6902 0.01564 0.00975 -0.0037 1.0000 0.0083
-6.750 -0.6654 0.01506 0.00907 -0.0032 0.9997 0.0094
-6.500 -0.6357 0.01440 0.00828 -0.0038 0.9917 0.0104
-6.250 -0.6046 0.01417 0.00806 -0.0046 0.9852 0.0115
-6.000 -0.5746 0.01386 0.00769 -0.0052 0.9789 0.0127
-5.750 -0.5454 0.01347 0.00721 -0.0055 0.9728 0.0140
-5.500 -0.5162 0.01359 0.00736 -0.0059 0.9665 0.0154
-5.250 -0.4886 0.01359 0.00734 -0.0058 0.9596 0.0170
-5.000 -0.4623 0.01343 0.00709 -0.0054 0.9526 0.0184
-4.750 -0.4382 0.01284 0.00645 -0.0046 0.9454 0.0197
-4.500 -0.4131 0.01251 0.00609 -0.0040 0.9382 0.0207
-4.250 -0.3885 0.01217 0.00571 -0.0032 0.9308 0.0218
-4.000 -0.3633 0.01181 0.00530 -0.0025 0.9236 0.0229
-3.750 -0.3383 0.01153 0.00496 -0.0018 0.9162 0.0242
-3.500 -0.3126 0.01129 0.00467 -0.0012 0.9083 0.0251
-3.250 -0.2889 0.01075 0.00406 -0.0002 0.9005 0.0263
-3.000 -0.2634 0.01039 0.00368 0.0004 0.8925 0.0276
-2.750 -0.2379 0.01016 0.00342 0.0011 0.8854 0.0290
-2.500 -0.2118 0.00994 0.00317 0.0015 0.8753 0.0306
-2.250 -0.1864 0.00973 0.00288 0.0023 0.8615 0.0318
-2.000 -0.1608 0.00954 0.00262 0.0030 0.8455 0.0329
-1.750 -0.1348 0.00940 0.00241 0.0036 0.8293 0.0342
-1.500 -0.1089 0.00918 0.00211 0.0042 0.8118 0.0375
-1.250 -0.0824 0.00904 0.00192 0.0046 0.7940 0.0416
-1.000 -0.0555 0.00894 0.00176 0.0050 0.7770 0.0478
-0.750 -0.0291 0.00875 0.00162 0.0054 0.7609 0.0763
-0.500 -0.0037 0.00833 0.00149 0.0058 0.7464 0.1796
-0.250 0.0195 0.00764 0.00137 0.0064 0.7311 0.3656
0.000 0.0365 0.00651 0.00124 0.0084 0.7141 0.6617
0.250 0.0973 0.00606 0.00165 0.0021 0.6825 0.9640
0.500 0.1440 0.00646 0.00183 -0.0017 0.6330 0.9860
0.750 0.1840 0.00685 0.00185 -0.0045 0.5483 0.9917
1.000 0.2204 0.00733 0.00190 -0.0066 0.4520 0.9952
1.250 0.2556 0.00771 0.00195 -0.0086 0.3747 0.9971
1.500 0.2887 0.00799 0.00199 -0.0100 0.3254 0.9987
1.750 0.3211 0.00820 0.00204 -0.0112 0.2923 1.0000
2.000 0.3479 0.00835 0.00208 -0.0111 0.2704 1.0000
2.250 0.3745 0.00849 0.00213 -0.0110 0.2519 1.0000
2.500 0.4010 0.00864 0.00220 -0.0108 0.2351 1.0000
2.750 0.4271 0.00879 0.00227 -0.0106 0.2193 1.0000
3.000 0.4531 0.00894 0.00235 -0.0103 0.2049 1.0000
3.250 0.4790 0.00909 0.00244 -0.0100 0.1924 1.0000
3.500 0.5048 0.00923 0.00255 -0.0096 0.1830 1.0000
4.000 0.5563 0.00953 0.00280 -0.0089 0.1663 1.0000
4.250 0.5818 0.00971 0.00295 -0.0086 0.1584 1.0000
4.500 0.6073 0.00987 0.00310 -0.0082 0.1498 1.0000
4.750 0.6327 0.01006 0.00327 -0.0078 0.1428 1.0000
5.000 0.6581 0.01024 0.00345 -0.0074 0.1365 1.0000
5.250 0.6833 0.01045 0.00365 -0.0069 0.1289 1.0000
5.500 0.7082 0.01069 0.00385 -0.0065 0.1185 1.0000
5.750 0.7333 0.01090 0.00406 -0.0060 0.1111 1.0000
6.000 0.7580 0.01115 0.00430 -0.0056 0.1031 1.0000
6.250 0.7828 0.01140 0.00455 -0.0051 0.0943 1.0000
6.500 0.8073 0.01167 0.00482 -0.0046 0.0855 1.0000
6.750 0.8315 0.01201 0.00511 -0.0040 0.0727 1.0000
7.000 0.8552 0.01241 0.00545 -0.0035 0.0601 1.0000
7.250 0.8789 0.01281 0.00585 -0.0028 0.0523 1.0000
7.500 0.9022 0.01325 0.00627 -0.0022 0.0457 1.0000
7.750 0.9259 0.01362 0.00668 -0.0016 0.0405 1.0000
8.000 0.9487 0.01411 0.00716 -0.0009 0.0340 1.0000
8.250 0.9719 0.01453 0.00761 -0.0002 0.0303 1.0000
8.500 0.9944 0.01506 0.00816 0.0005 0.0267 1.0000
8.750 1.0170 0.01554 0.00870 0.0013 0.0244 1.0000
9.000 1.0393 0.01603 0.00925 0.0020 0.0224 1.0000
9.250 1.0610 0.01661 0.00987 0.0029 0.0205 1.0000
9.500 1.0815 0.01735 0.01067 0.0038 0.0188 1.0000
9.750 1.1030 0.01788 0.01131 0.0047 0.0181 1.0000
10.000 1.1239 0.01848 0.01200 0.0056 0.0172 1.0000
10.250 1.1443 0.01910 0.01272 0.0066 0.0163 1.0000
10.500 1.1641 0.01977 0.01346 0.0076 0.0153 1.0000
10.750 1.1825 0.02058 0.01434 0.0087 0.0144 1.0000
11.000 1.1982 0.02165 0.01551 0.0101 0.0136 1.0000
11.250 1.2163 0.02241 0.01640 0.0113 0.0132 1.0000
11.500 1.2333 0.02325 0.01736 0.0126 0.0128 1.0000
11.750 1.2492 0.02416 0.01839 0.0139 0.0124 1.0000
12.000 1.2639 0.02512 0.01948 0.0153 0.0119 1.0000
12.250 1.2774 0.02615 0.02063 0.0168 0.0115 1.0000
12.500 1.2896 0.02721 0.02181 0.0184 0.0112 1.0000
12.750 1.3000 0.02834 0.02304 0.0201 0.0108 1.0000
13.000 1.3052 0.02957 0.02437 0.0224 0.0105 1.0000
13.250 1.3066 0.03111 0.02603 0.0248 0.0102 1.0000
13.500 1.3034 0.03319 0.02826 0.0268 0.0099 1.0000
13.750 1.3012 0.03545 0.03067 0.0281 0.0096 1.0000
14.000 1.3026 0.03760 0.03298 0.0287 0.0095 1.0000
14.250 1.3023 0.04015 0.03568 0.0288 0.0093 1.0000
14.500 1.2982 0.04340 0.03910 0.0282 0.0092 1.0000
14.750 1.2911 0.04742 0.04328 0.0267 0.0090 1.0000
15.000 1.2807 0.05247 0.04851 0.0240 0.0089 1.0000
15.250 1.2651 0.05913 0.05536 0.0198 0.0089 1.0000
15.500 1.2437 0.06764 0.06406 0.0142 0.0089 1.0000
15.750 1.2164 0.07735 0.07395 0.0083 0.0090 1.0000
16.000 1.1859 0.08718 0.08391 0.0030 0.0091 1.0000
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Polar data table (+)
Polar graphs
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