BOEING VERTOL V43015-2.48 AIRFOIL (v43015-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: BOEING VERTOL V43015-2.48 AIRFOIL (v43015-il) Reynolds number: 100,000 Max Cl/Cd: 36.35 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-v43015-il-100000.txt Download as CSV file: xf-v43015-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING VERTOL V43015-2.48 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.3789 0.10386 0.09861 -0.0324 1.0000 0.1116
-10.500 -0.4313 0.09707 0.09194 -0.0421 1.0000 0.1142
-10.250 -0.4789 0.09261 0.08747 -0.0453 1.0000 0.1145
-10.000 -0.3907 0.09257 0.08752 -0.0348 1.0000 0.1203
-9.750 -0.4003 0.08875 0.08377 -0.0363 1.0000 0.1246
-9.500 -0.4380 0.08278 0.07786 -0.0410 1.0000 0.1265
-9.250 -0.4805 0.07946 0.07456 -0.0400 1.0000 0.1273
-9.000 -0.5168 0.07749 0.07259 -0.0355 1.0000 0.1276
-8.750 -0.5649 0.07632 0.07130 -0.0297 1.0000 0.1285
-8.500 -0.6097 0.07640 0.07110 -0.0228 1.0000 0.1292
-8.250 -0.5438 0.06928 0.06454 -0.0265 1.0000 0.1371
-8.000 -0.5674 0.06741 0.06260 -0.0220 1.0000 0.1408
-7.500 -0.6008 0.06264 0.05774 -0.0138 1.0000 0.1496
-7.250 -0.6018 0.06088 0.05600 -0.0106 1.0000 0.1558
-6.750 -0.6166 0.05678 0.05185 -0.0034 1.0000 0.1713
-6.500 -0.6264 0.05479 0.04980 0.0003 1.0000 0.1839
-5.750 -0.6062 0.04932 0.04443 0.0058 0.9940 0.2370
-5.500 -0.5711 0.04712 0.04231 0.0033 0.9836 0.2706
-5.250 -0.5382 0.04527 0.04031 0.0011 0.9715 0.2985
-5.000 -0.5083 0.04243 0.03740 -0.0004 0.9596 0.3197
-4.750 -0.4717 0.03908 0.03395 -0.0027 0.9488 0.3238
-4.500 -0.4070 0.03424 0.02669 -0.0071 0.9388 0.1820
-4.250 -0.3523 0.03326 0.02486 -0.0078 0.9267 0.1261
-4.000 -0.3015 0.03175 0.02291 -0.0102 0.9176 0.1126
-3.750 -0.2564 0.02978 0.02082 -0.0125 0.9066 0.1090
-3.500 -0.2115 0.02870 0.01935 -0.0140 0.8947 0.1031
-3.250 -0.1534 0.02733 0.01785 -0.0182 0.8867 0.1014
-3.000 -0.1046 0.02561 0.01616 -0.0210 0.8741 0.1025
-2.750 -0.0383 0.02334 0.01409 -0.0268 0.8660 0.1061
-2.500 0.0052 0.02196 0.01282 -0.0283 0.8489 0.1082
-2.250 0.0432 0.02080 0.01172 -0.0289 0.8304 0.1118
-2.000 0.0738 0.01971 0.01067 -0.0282 0.8104 0.1179
-1.750 0.0993 0.01894 0.00985 -0.0268 0.7872 0.1246
-1.500 0.1234 0.01831 0.00910 -0.0250 0.7617 0.1328
-1.250 0.1487 0.01758 0.00832 -0.0235 0.7363 0.1557
-1.000 0.1556 0.01517 0.00818 -0.0187 0.7142 0.6833
-0.750 0.2680 0.01735 0.01050 -0.0265 0.6679 0.9035
-0.500 0.3713 0.01881 0.01137 -0.0374 0.6183 0.9396
-0.250 0.4396 0.01911 0.01126 -0.0442 0.5773 0.9590
0.000 0.4999 0.01913 0.01089 -0.0502 0.5419 0.9772
0.250 0.5464 0.01898 0.01049 -0.0542 0.5116 0.9890
0.500 0.5894 0.01884 0.01006 -0.0576 0.4871 0.9983
0.750 0.6110 0.01895 0.00998 -0.0571 0.4690 1.0000
1.000 0.6271 0.01913 0.01002 -0.0555 0.4531 1.0000
1.250 0.6435 0.01934 0.01011 -0.0538 0.4386 1.0000
1.500 0.6608 0.01960 0.01017 -0.0523 0.4266 1.0000
1.750 0.6775 0.01983 0.01033 -0.0507 0.4142 1.0000
2.000 0.6948 0.02013 0.01055 -0.0492 0.4037 1.0000
2.250 0.7126 0.02041 0.01070 -0.0477 0.3938 1.0000
2.500 0.7297 0.02075 0.01100 -0.0460 0.3845 1.0000
2.750 0.7479 0.02105 0.01120 -0.0446 0.3760 1.0000
3.000 0.7659 0.02145 0.01156 -0.0430 0.3683 1.0000
3.250 0.7832 0.02178 0.01188 -0.0414 0.3604 1.0000
3.500 0.8039 0.02224 0.01213 -0.0403 0.3539 1.0000
3.750 0.8187 0.02263 0.01267 -0.0381 0.3466 1.0000
4.000 0.8371 0.02303 0.01301 -0.0366 0.3401 1.0000
4.250 0.8570 0.02358 0.01345 -0.0354 0.3345 1.0000
4.500 0.8716 0.02407 0.01408 -0.0332 0.3280 1.0000
4.750 0.8899 0.02454 0.01452 -0.0316 0.3223 1.0000
5.000 0.9114 0.02518 0.01499 -0.0307 0.3173 1.0000
5.250 0.9241 0.02579 0.01581 -0.0282 0.3117 1.0000
5.500 0.9407 0.02637 0.01644 -0.0263 0.3064 1.0000
5.750 0.9612 0.02696 0.01694 -0.0252 0.3020 1.0000
6.000 0.9774 0.02779 0.01783 -0.0234 0.2975 1.0000
6.250 0.9899 0.02857 0.01877 -0.0209 0.2924 1.0000
6.500 1.0066 0.02920 0.01943 -0.0192 0.2877 1.0000
6.750 1.0287 0.02986 0.01997 -0.0183 0.2836 1.0000
7.000 1.0391 0.03092 0.02121 -0.0157 0.2794 1.0000
7.250 1.0492 0.03194 0.02241 -0.0130 0.2750 1.0000
7.500 1.0642 0.03281 0.02335 -0.0111 0.2711 1.0000
7.750 1.0839 0.03358 0.02407 -0.0099 0.2677 1.0000
8.000 1.0997 0.03479 0.02531 -0.0083 0.2643 1.0000
8.250 1.0985 0.03627 0.02710 -0.0042 0.2605 1.0000
8.500 1.1050 0.03754 0.02853 -0.0013 0.2568 1.0000
8.750 1.1206 0.03846 0.02948 0.0003 0.2533 1.0000
9.000 1.1412 0.03938 0.03035 0.0012 0.2505 1.0000
9.250 1.1512 0.04108 0.03213 0.0032 0.2479 1.0000
9.500 1.1324 0.04358 0.03500 0.0089 0.2456 1.0000
9.750 1.1105 0.04633 0.03803 0.0144 0.2432 1.0000
10.000 1.0829 0.04934 0.04127 0.0202 0.2411 1.0000
10.250 1.0321 0.05317 0.04529 0.0277 0.2397 1.0000
10.500 0.6856 0.10076 0.09326 0.0162 0.2490 1.0000
10.750 1.1078 0.05351 0.04552 0.0259 0.2338 1.0000
11.000 1.1046 0.05617 0.04822 0.0283 0.2321 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING VERTOL V43015-2.48 AIRFOIL (v43015-il)