BOEING-VERTOL VR-15 AIRFOIL (vr15-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL VR-15 AIRFOIL (vr15-il) Reynolds number: 200,000 Max Cl/Cd: 52 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-vr15-il-200000-n5.txt Download as CSV file: xf-vr15-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.6131 0.08712 0.08375 0.0030 1.0000 0.0117
-8.500 -0.6194 0.08100 0.07767 -0.0028 1.0000 0.0115
-8.250 -0.6263 0.07484 0.07148 -0.0076 1.0000 0.0114
-8.000 -0.6290 0.06843 0.06498 -0.0115 1.0000 0.0114
-7.750 -0.6290 0.06214 0.05855 -0.0140 1.0000 0.0114
-7.500 -0.6262 0.05614 0.05235 -0.0154 1.0000 0.0115
-7.250 -0.6208 0.05030 0.04626 -0.0157 1.0000 0.0118
-7.000 -0.6139 0.04421 0.03980 -0.0152 1.0000 0.0126
-6.750 -0.6137 0.03453 0.02929 -0.0126 1.0000 0.0141
-6.500 -0.5930 0.03401 0.02871 -0.0123 1.0000 0.0148
-6.250 -0.5760 0.03143 0.02584 -0.0110 1.0000 0.0156
-6.000 -0.5609 0.02756 0.02140 -0.0089 1.0000 0.0171
-5.750 -0.5437 0.02475 0.01801 -0.0069 1.0000 0.0193
-5.500 -0.5236 0.02422 0.01745 -0.0059 1.0000 0.0206
-5.250 -0.5054 0.02284 0.01584 -0.0041 1.0000 0.0221
-5.000 -0.4868 0.02161 0.01428 -0.0022 1.0000 0.0242
-4.750 -0.4620 0.02004 0.01241 -0.0016 0.9977 0.0259
-4.500 -0.4292 0.01882 0.01107 -0.0029 0.9940 0.0275
-4.250 -0.3967 0.01802 0.01016 -0.0040 0.9891 0.0297
-4.000 -0.3632 0.01719 0.00917 -0.0052 0.9848 0.0321
-3.750 -0.3308 0.01636 0.00821 -0.0061 0.9799 0.0335
-3.500 -0.2997 0.01528 0.00709 -0.0069 0.9746 0.0355
-3.250 -0.2669 0.01463 0.00645 -0.0081 0.9704 0.0383
-3.000 -0.2375 0.01407 0.00587 -0.0084 0.9633 0.0403
-2.750 -0.2062 0.01357 0.00533 -0.0092 0.9577 0.0424
-2.500 -0.1779 0.01324 0.00493 -0.0092 0.9498 0.0450
-2.250 -0.1493 0.01274 0.00442 -0.0093 0.9432 0.0480
-2.000 -0.1232 0.01241 0.00408 -0.0088 0.9343 0.0518
-1.750 -0.0959 0.01216 0.00378 -0.0085 0.9270 0.0573
-1.500 -0.0709 0.01185 0.00351 -0.0077 0.9160 0.0715
-1.250 -0.0499 0.01116 0.00326 -0.0062 0.9015 0.1812
-1.000 -0.0371 0.00972 0.00300 -0.0032 0.8857 0.4981
-0.750 0.0406 0.00888 0.00362 -0.0115 0.8731 0.9697
-0.500 0.1043 0.00901 0.00360 -0.0186 0.8565 1.0000
-0.250 0.1278 0.00896 0.00342 -0.0175 0.8379 1.0000
0.000 0.1519 0.00891 0.00327 -0.0166 0.8202 1.0000
0.250 0.1763 0.00889 0.00314 -0.0156 0.8019 1.0000
0.500 0.2008 0.00887 0.00301 -0.0147 0.7824 1.0000
0.750 0.2258 0.00888 0.00291 -0.0140 0.7598 1.0000
1.000 0.2509 0.00891 0.00282 -0.0132 0.7340 1.0000
1.250 0.2761 0.00896 0.00274 -0.0125 0.7005 1.0000
1.500 0.3010 0.00907 0.00266 -0.0117 0.6546 1.0000
1.750 0.3256 0.00930 0.00258 -0.0109 0.5824 1.0000
2.000 0.3502 0.00975 0.00256 -0.0103 0.4861 1.0000
2.250 0.3751 0.01026 0.00266 -0.0100 0.4036 1.0000
2.500 0.4001 0.01069 0.00281 -0.0097 0.3445 1.0000
2.750 0.4253 0.01103 0.00296 -0.0094 0.3087 1.0000
3.000 0.4505 0.01132 0.00312 -0.0090 0.2839 1.0000
3.250 0.4756 0.01159 0.00330 -0.0086 0.2640 1.0000
3.500 0.5006 0.01187 0.00352 -0.0082 0.2470 1.0000
3.750 0.5255 0.01214 0.00373 -0.0078 0.2320 1.0000
4.000 0.5504 0.01240 0.00395 -0.0073 0.2192 1.0000
4.250 0.5752 0.01267 0.00419 -0.0068 0.2071 1.0000
4.500 0.5999 0.01295 0.00446 -0.0063 0.1957 1.0000
4.750 0.6244 0.01324 0.00472 -0.0058 0.1849 1.0000
5.000 0.6488 0.01354 0.00500 -0.0053 0.1747 1.0000
5.250 0.6734 0.01381 0.00532 -0.0047 0.1657 1.0000
5.500 0.6975 0.01415 0.00565 -0.0042 0.1584 1.0000
5.750 0.7219 0.01444 0.00600 -0.0036 0.1508 1.0000
6.000 0.7459 0.01479 0.00635 -0.0030 0.1428 1.0000
6.250 0.7699 0.01511 0.00671 -0.0024 0.1328 1.0000
6.500 0.7939 0.01543 0.00707 -0.0018 0.1222 1.0000
6.750 0.8176 0.01579 0.00746 -0.0012 0.1123 1.0000
7.000 0.8411 0.01619 0.00787 -0.0006 0.1021 1.0000
7.250 0.8642 0.01662 0.00831 0.0001 0.0905 1.0000
7.500 0.8872 0.01709 0.00882 0.0007 0.0783 1.0000
8.000 0.9313 0.01829 0.01003 0.0023 0.0571 1.0000
8.250 0.9526 0.01900 0.01079 0.0032 0.0499 1.0000
8.500 0.9723 0.01992 0.01175 0.0042 0.0441 1.0000
8.750 0.9933 0.02062 0.01260 0.0052 0.0400 1.0000
9.000 1.0132 0.02145 0.01350 0.0061 0.0360 1.0000
9.250 1.0316 0.02243 0.01457 0.0073 0.0329 1.0000
9.500 1.0506 0.02332 0.01561 0.0084 0.0307 1.0000
9.750 1.0687 0.02429 0.01671 0.0096 0.0288 1.0000
10.000 1.0853 0.02536 0.01790 0.0109 0.0271 1.0000
10.250 1.0984 0.02683 0.01944 0.0125 0.0255 1.0000
10.500 1.1146 0.02791 0.02072 0.0138 0.0243 1.0000
10.750 1.1290 0.02914 0.02213 0.0152 0.0230 1.0000
11.000 1.1415 0.03051 0.02367 0.0168 0.0221 1.0000
11.250 1.1524 0.03194 0.02526 0.0184 0.0212 1.0000
11.500 1.1612 0.03347 0.02693 0.0202 0.0205 1.0000
11.750 1.1667 0.03513 0.02874 0.0221 0.0199 1.0000
12.000 1.1664 0.03703 0.03078 0.0246 0.0194 1.0000
12.250 1.1616 0.03955 0.03345 0.0269 0.0189 1.0000
12.500 1.1603 0.04181 0.03592 0.0283 0.0185 1.0000
12.750 1.1571 0.04442 0.03877 0.0291 0.0182 1.0000
13.000 1.1514 0.04755 0.04213 0.0293 0.0178 1.0000
13.250 1.1432 0.05130 0.04611 0.0286 0.0175 1.0000
13.500 1.1324 0.05581 0.05084 0.0268 0.0173 1.0000
13.750 1.1189 0.06130 0.05655 0.0239 0.0171 1.0000
14.000 1.1024 0.06801 0.06348 0.0195 0.0171 1.0000
14.250 1.0827 0.07609 0.07175 0.0140 0.0171 1.0000
14.500 1.0590 0.08538 0.08123 0.0079 0.0172 1.0000
14.750 1.0312 0.09567 0.09167 0.0016 0.0175 1.0000
15.000 0.9986 0.10705 0.10317 -0.0049 0.0178 1.0000
15.250 0.9615 0.11978 0.11598 -0.0119 0.0183 1.0000
15.500 0.9154 0.13576 0.13202 -0.0202 0.0187 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL VR-15 AIRFOIL (vr15-il)