BOEING-VERTOL VR-8 AIRFOIL (vr8-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL VR-8 AIRFOIL (vr8-il) Reynolds number: 500,000 Max Cl/Cd: 56.23 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-vr8-il-500000.txt Download as CSV file: xf-vr8-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL VR-8 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5542 0.10968 0.10734 0.0029 1.0000 0.0129
-10.000 -0.5528 0.10494 0.10262 0.0000 1.0000 0.0138
-7.500 -0.5654 0.05617 0.05357 -0.0284 1.0000 0.0143
-7.250 -0.5617 0.05188 0.04907 -0.0283 1.0000 0.0144
-7.000 -0.5561 0.04783 0.04477 -0.0277 1.0000 0.0144
-6.750 -0.5593 0.04080 0.03759 -0.0273 1.0000 0.0149
-6.500 -0.5476 0.03811 0.03480 -0.0265 1.0000 0.0153
-6.250 -0.5346 0.03537 0.03189 -0.0254 1.0000 0.0156
-6.000 -0.5207 0.03268 0.02901 -0.0241 1.0000 0.0160
-5.750 -0.5056 0.03007 0.02617 -0.0226 1.0000 0.0167
-5.500 -0.4890 0.02753 0.02337 -0.0209 1.0000 0.0177
-5.250 -0.4660 0.02698 0.02248 -0.0189 1.0000 0.0200
-5.000 -0.4554 0.02194 0.01699 -0.0169 1.0000 0.0215
-4.750 -0.4362 0.02072 0.01572 -0.0158 1.0000 0.0235
-4.250 -0.3256 0.00841 0.00342 -0.0255 0.9726 0.0464
-4.000 -0.3197 0.01488 0.00890 -0.0217 0.9806 0.0273
-3.750 -0.2849 0.01278 0.00670 -0.0228 0.9626 0.0263
-3.500 -0.2465 0.01171 0.00553 -0.0248 0.9400 0.0271
-3.250 -0.2148 0.01046 0.00416 -0.0255 0.9016 0.0283
-3.000 -0.1899 0.01000 0.00351 -0.0246 0.8535 0.0286
-2.750 -0.1662 0.00973 0.00303 -0.0236 0.8101 0.0297
-2.500 -0.1424 0.00956 0.00260 -0.0226 0.7632 0.0300
-2.250 -0.1200 0.00961 0.00222 -0.0212 0.6755 0.0301
-2.000 -0.0958 0.00967 0.00191 -0.0204 0.6010 0.0306
-1.750 -0.0705 0.00973 0.00165 -0.0200 0.5369 0.0318
-1.500 -0.0447 0.00982 0.00144 -0.0196 0.4708 0.0341
-1.250 -0.0245 0.01106 0.00151 -0.0189 0.1200 0.0419
-1.000 0.0009 0.01082 0.00139 -0.0187 0.0455 0.1437
-0.750 0.0231 0.00949 0.00122 -0.0185 0.0448 0.4969
-0.500 0.0489 0.00918 0.00120 -0.0182 0.0440 0.5982
-0.250 0.0741 0.00889 0.00121 -0.0177 0.0435 0.6888
0.000 0.1002 0.00877 0.00122 -0.0173 0.0432 0.7370
0.250 0.1264 0.00872 0.00124 -0.0169 0.0406 0.7740
0.500 0.1514 0.00863 0.00128 -0.0162 0.0394 0.8197
0.750 0.1743 0.00857 0.00135 -0.0148 0.0386 0.8693
1.000 0.1987 0.00860 0.00147 -0.0136 0.0382 0.9225
1.250 0.2297 0.00875 0.00162 -0.0139 0.0382 0.9657
1.500 0.2828 0.00885 0.00168 -0.0194 0.0380 0.9892
1.750 0.3394 0.00891 0.00168 -0.0258 0.0365 0.9982
2.000 0.3701 0.00899 0.00174 -0.0265 0.0363 1.0000
2.250 0.3941 0.00909 0.00183 -0.0258 0.0362 1.0000
2.500 0.4181 0.00920 0.00196 -0.0251 0.0357 1.0000
2.750 0.4423 0.00932 0.00209 -0.0244 0.0350 1.0000
3.000 0.4665 0.00945 0.00226 -0.0237 0.0342 1.0000
3.250 0.4907 0.00961 0.00246 -0.0229 0.0331 1.0000
3.500 0.5148 0.00979 0.00270 -0.0222 0.0313 1.0000
3.750 0.5388 0.01000 0.00295 -0.0214 0.0299 1.0000
4.000 0.5627 0.01024 0.00322 -0.0206 0.0293 1.0000
4.250 0.5863 0.01052 0.00354 -0.0198 0.0291 1.0000
4.500 0.6096 0.01086 0.00392 -0.0189 0.0291 1.0000
4.750 0.6326 0.01125 0.00439 -0.0180 0.0293 1.0000
5.000 0.6551 0.01169 0.00489 -0.0170 0.0295 1.0000
5.250 0.6769 0.01223 0.00548 -0.0159 0.0295 1.0000
5.500 0.6976 0.01292 0.00621 -0.0146 0.0287 1.0000
5.750 0.7192 0.01352 0.00689 -0.0136 0.0281 1.0000
6.000 0.7403 0.01428 0.00771 -0.0124 0.0278 1.0000
6.250 0.7611 0.01521 0.00871 -0.0111 0.0274 1.0000
6.500 0.7824 0.01625 0.00981 -0.0100 0.0263 1.0000
6.750 0.8039 0.01752 0.01112 -0.0089 0.0248 1.0000
7.000 0.8270 0.02008 0.01392 -0.0074 0.0261 1.0000
7.250 0.8498 0.02180 0.01575 -0.0067 0.0246 1.0000
7.500 0.8681 0.02407 0.01837 -0.0054 0.0210 1.0000
7.750 0.8832 0.03005 0.02438 -0.0051 0.0184 1.0000
8.000 0.9029 0.02817 0.02295 -0.0034 0.0162 1.0000
8.250 0.9205 0.02995 0.02488 -0.0025 0.0149 1.0000
8.500 0.9340 0.03316 0.02820 -0.0018 0.0131 1.0000
8.750 0.9425 0.03579 0.03138 0.0005 0.0127 1.0000
9.000 0.9470 0.04020 0.03611 0.0021 0.0131 1.0000
9.250 0.9525 0.04778 0.04373 0.0026 0.0136 1.0000
9.500 0.9525 0.05165 0.04789 0.0041 0.0136 1.0000
9.750 0.9466 0.05548 0.05200 0.0058 0.0136 1.0000
10.000 0.9384 0.05910 0.05585 0.0072 0.0136 1.0000
10.250 0.9232 0.06226 0.05920 0.0092 0.0136 1.0000
10.500 0.9060 0.06540 0.06248 0.0102 0.0136 1.0000
10.750 0.8880 0.06923 0.06645 0.0093 0.0136 1.0000
11.000 0.8718 0.07369 0.07102 0.0069 0.0136 1.0000
11.250 0.8563 0.07890 0.07634 0.0034 0.0136 1.0000
11.500 0.8382 0.08521 0.08276 -0.0015 0.0136 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL VR-8 AIRFOIL (vr8-il)