NACA 63012A AIRFOIL (n63012a-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 63012A AIRFOIL (n63012a-il) Reynolds number: 50,000 Max Cl/Cd: 28.13 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n63012a-il-50000-n5.txt Download as CSV file: xf-n63012a-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63012A AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.7313 0.09134 0.08404 -0.0199 1.0000 0.0498
-11.250 -0.7742 0.08112 0.07366 -0.0272 1.0000 0.0483
-11.000 -0.8097 0.07444 0.06673 -0.0303 1.0000 0.0475
-10.750 -0.8317 0.06993 0.06199 -0.0306 1.0000 0.0473
-10.500 -0.8475 0.06621 0.05804 -0.0294 1.0000 0.0473
-10.250 -0.8559 0.06275 0.05435 -0.0280 1.0000 0.0475
-10.000 -0.8574 0.05960 0.05100 -0.0268 1.0000 0.0483
-9.750 -0.8547 0.05675 0.04796 -0.0255 1.0000 0.0494
-9.500 -0.8512 0.05387 0.04482 -0.0241 1.0000 0.0508
-9.250 -0.8458 0.05092 0.04153 -0.0227 1.0000 0.0522
-9.000 -0.8373 0.04794 0.03813 -0.0212 1.0000 0.0536
-8.750 -0.8246 0.04500 0.03477 -0.0198 1.0000 0.0547
-8.500 -0.8078 0.04228 0.03161 -0.0186 1.0000 0.0560
-8.250 -0.7876 0.03984 0.02878 -0.0176 1.0000 0.0578
-8.000 -0.7664 0.03782 0.02679 -0.0171 1.0000 0.0611
-7.750 -0.7443 0.03610 0.02491 -0.0164 1.0000 0.0655
-7.500 -0.7183 0.03439 0.02284 -0.0156 1.0000 0.0697
-7.250 -0.6960 0.03277 0.02135 -0.0149 1.0000 0.0749
-7.000 -0.6742 0.03148 0.01987 -0.0138 1.0000 0.0829
-6.750 -0.6567 0.03005 0.01851 -0.0124 1.0000 0.0908
-6.500 -0.6409 0.02872 0.01716 -0.0108 1.0000 0.1024
-6.250 -0.6270 0.02734 0.01585 -0.0089 1.0000 0.1179
-6.000 -0.6153 0.02586 0.01455 -0.0068 1.0000 0.1425
-5.750 -0.6067 0.02415 0.01326 -0.0045 1.0000 0.1893
-5.500 -0.6029 0.02224 0.01222 -0.0015 1.0000 0.2936
-5.250 -0.5966 0.02121 0.01206 0.0024 1.0000 0.4374
-5.000 -0.5843 0.02099 0.01205 0.0059 1.0000 0.5237
-4.750 -0.5703 0.02103 0.01216 0.0093 1.0000 0.5809
-4.500 -0.5558 0.02121 0.01236 0.0128 1.0000 0.6267
-4.250 -0.5399 0.02153 0.01268 0.0163 1.0000 0.6640
-4.000 -0.5234 0.02178 0.01284 0.0196 1.0000 0.6943
-3.750 -0.5076 0.02186 0.01282 0.0225 1.0000 0.7198
-3.500 -0.4882 0.02196 0.01283 0.0248 1.0000 0.7394
-3.250 -0.4697 0.02192 0.01268 0.0268 1.0000 0.7572
-3.000 -0.4511 0.02183 0.01247 0.0285 1.0000 0.7733
-2.750 -0.4327 0.02171 0.01222 0.0300 1.0000 0.7882
-2.500 -0.4142 0.02159 0.01200 0.0314 1.0000 0.8025
-2.250 -0.3952 0.02149 0.01180 0.0326 1.0000 0.8167
-2.000 -0.3551 0.02152 0.01166 0.0297 0.9907 0.8299
-1.750 -0.3143 0.02156 0.01156 0.0266 0.9810 0.8429
-1.500 -0.2722 0.02161 0.01149 0.0233 0.9717 0.8553
-1.250 -0.2274 0.02169 0.01147 0.0197 0.9629 0.8659
-1.000 -0.1848 0.02173 0.01143 0.0163 0.9532 0.8770
-0.750 -0.1400 0.02177 0.01138 0.0125 0.9445 0.8884
-0.500 -0.0929 0.02182 0.01138 0.0082 0.9356 0.8981
-0.250 -0.0462 0.02184 0.01138 0.0041 0.9266 0.9076
0.000 0.0000 0.02184 0.01136 0.0000 0.9179 0.9179
0.250 0.0462 0.02184 0.01138 -0.0041 0.9076 0.9266
0.500 0.0929 0.02182 0.01138 -0.0083 0.8981 0.9356
0.750 0.1400 0.02177 0.01138 -0.0125 0.8884 0.9445
1.000 0.1848 0.02173 0.01142 -0.0163 0.8770 0.9532
1.250 0.2274 0.02169 0.01146 -0.0197 0.8659 0.9629
1.500 0.2722 0.02161 0.01149 -0.0234 0.8553 0.9717
1.750 0.3143 0.02156 0.01156 -0.0266 0.8429 0.9810
2.000 0.3552 0.02152 0.01166 -0.0297 0.8299 0.9908
2.250 0.3952 0.02149 0.01180 -0.0326 0.8167 1.0000
2.500 0.4142 0.02159 0.01200 -0.0314 0.8025 1.0000
2.750 0.4327 0.02171 0.01222 -0.0300 0.7882 1.0000
3.000 0.4512 0.02183 0.01247 -0.0285 0.7733 1.0000
3.250 0.4697 0.02192 0.01268 -0.0268 0.7572 1.0000
3.500 0.4882 0.02196 0.01283 -0.0248 0.7394 1.0000
3.750 0.5076 0.02186 0.01282 -0.0225 0.7198 1.0000
4.000 0.5235 0.02178 0.01284 -0.0196 0.6943 1.0000
4.250 0.5399 0.02153 0.01268 -0.0164 0.6640 1.0000
4.500 0.5559 0.02121 0.01236 -0.0128 0.6267 1.0000
4.750 0.5703 0.02103 0.01216 -0.0093 0.5808 1.0000
5.000 0.5844 0.02099 0.01205 -0.0059 0.5237 1.0000
5.250 0.5966 0.02121 0.01205 -0.0024 0.4373 1.0000
5.500 0.6029 0.02224 0.01222 0.0015 0.2935 1.0000
5.750 0.6068 0.02414 0.01326 0.0045 0.1892 1.0000
6.000 0.6153 0.02586 0.01455 0.0068 0.1425 1.0000
6.250 0.6270 0.02734 0.01585 0.0089 0.1178 1.0000
6.500 0.6409 0.02872 0.01716 0.0108 0.1024 1.0000
6.750 0.6567 0.03004 0.01851 0.0124 0.0908 1.0000
7.000 0.6742 0.03148 0.01986 0.0138 0.0830 1.0000
7.250 0.6960 0.03276 0.02134 0.0149 0.0750 1.0000
7.500 0.7183 0.03439 0.02284 0.0156 0.0697 1.0000
7.750 0.7443 0.03610 0.02490 0.0164 0.0655 1.0000
8.000 0.7663 0.03782 0.02679 0.0171 0.0611 1.0000
8.250 0.7876 0.03983 0.02877 0.0176 0.0577 1.0000
8.500 0.8078 0.04228 0.03161 0.0186 0.0560 1.0000
8.750 0.8246 0.04500 0.03477 0.0198 0.0547 1.0000
9.000 0.8373 0.04794 0.03812 0.0212 0.0536 1.0000
9.250 0.8458 0.05092 0.04153 0.0226 0.0522 1.0000
9.500 0.8512 0.05387 0.04482 0.0241 0.0508 1.0000
9.750 0.8548 0.05675 0.04796 0.0255 0.0494 1.0000
10.000 0.8576 0.05960 0.05100 0.0267 0.0483 1.0000
10.250 0.8561 0.06275 0.05435 0.0280 0.0475 1.0000
10.500 0.8478 0.06622 0.05805 0.0293 0.0472 1.0000
10.750 0.8320 0.06995 0.06201 0.0305 0.0473 1.0000
11.000 0.8101 0.07446 0.06674 0.0303 0.0475 1.0000
11.250 0.7748 0.08114 0.07368 0.0271 0.0483 1.0000
11.500 0.7319 0.09136 0.08407 0.0197 0.0498 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 63012A AIRFOIL (n63012a-il)