NASA SC(2)-0612 AIRFOIL (sc20612-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA SC(2)-0612 AIRFOIL (sc20612-il) Reynolds number: 50,000 Max Cl/Cd: 23.25 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-sc20612-il-50000-n5.txt Download as CSV file: xf-sc20612-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0612 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
------ -------- --------- --------- -------- -------- --------
-11.750 -0.6129 0.09741 0.08891 -0.0415 1.0000 0.0755
-11.500 -0.6443 0.08701 0.07850 -0.0480 1.0000 0.0748
-11.250 -0.6821 0.07821 0.06965 -0.0530 1.0000 0.0740
-11.000 -0.7174 0.07190 0.06324 -0.0550 1.0000 0.0735
-10.750 -0.7489 0.06741 0.05867 -0.0545 1.0000 0.0732
-10.500 -0.7702 0.06323 0.05431 -0.0545 1.0000 0.0734
-10.250 -0.7828 0.05940 0.05024 -0.0541 1.0000 0.0740
-10.000 -0.7887 0.05573 0.04625 -0.0535 1.0000 0.0749
-9.750 -0.7889 0.05218 0.04233 -0.0530 1.0000 0.0760
-9.500 -0.7850 0.04877 0.03843 -0.0526 1.0000 0.0779
-9.250 -0.7731 0.04631 0.03578 -0.0520 1.0000 0.0802
-9.000 -0.7577 0.04451 0.03391 -0.0513 1.0000 0.0829
-8.750 -0.7418 0.04241 0.03158 -0.0507 1.0000 0.0858
-8.500 -0.7245 0.04019 0.02896 -0.0502 1.0000 0.0890
-8.250 -0.7060 0.03839 0.02700 -0.0495 1.0000 0.0927
-8.000 -0.6867 0.03702 0.02560 -0.0486 1.0000 0.0968
-7.750 -0.6664 0.03556 0.02389 -0.0478 1.0000 0.1017
-7.500 -0.6463 0.03426 0.02253 -0.0467 1.0000 0.1066
-7.250 -0.6261 0.03320 0.02145 -0.0455 1.0000 0.1121
-7.000 -0.6051 0.03218 0.02025 -0.0443 1.0000 0.1187
-6.750 -0.5856 0.03129 0.01948 -0.0427 1.0000 0.1249
-6.500 -0.5649 0.03051 0.01853 -0.0412 1.0000 0.1329
-6.250 -0.5456 0.02968 0.01785 -0.0398 1.0000 0.1407
-6.000 -0.5255 0.02892 0.01706 -0.0385 1.0000 0.1503
-5.750 -0.5051 0.02811 0.01631 -0.0375 1.0000 0.1625
-5.500 -0.4841 0.02722 0.01553 -0.0369 1.0000 0.1771
-5.250 -0.4616 0.02622 0.01473 -0.0369 1.0000 0.1990
-5.000 -0.4365 0.02497 0.01386 -0.0379 1.0000 0.2387
-4.750 -0.4097 0.02341 0.01308 -0.0397 1.0000 0.3357
-4.500 -0.4020 0.02361 0.01437 -0.0338 1.0000 0.4607
-4.250 -0.3899 0.02456 0.01554 -0.0287 1.0000 0.5487
-3.750 -0.3431 0.02577 0.01640 -0.0264 1.0000 0.6398
-3.500 -0.3275 0.02650 0.01706 -0.0228 1.0000 0.6632
-3.250 -0.3083 0.02711 0.01756 -0.0205 1.0000 0.6892
-3.000 -0.2971 0.02779 0.01822 -0.0158 1.0000 0.7099
-2.750 -0.2865 0.02834 0.01874 -0.0110 1.0000 0.7306
-2.500 -0.2743 0.02869 0.01906 -0.0069 1.0000 0.7502
-2.250 -0.2586 0.02882 0.01913 -0.0041 1.0000 0.7670
-2.000 -0.2370 0.02880 0.01903 -0.0033 1.0000 0.7805
-1.750 -0.2203 0.02865 0.01883 -0.0011 1.0000 0.7891
-1.500 -0.1971 0.02855 0.01866 -0.0010 1.0000 0.7989
-1.250 -0.1781 0.02841 0.01848 0.0004 1.0000 0.8075
-1.000 -0.1534 0.02833 0.01836 0.0000 1.0000 0.8152
-0.750 -0.1320 0.02819 0.01820 0.0005 1.0000 0.8208
-0.500 -0.1061 0.02817 0.01815 -0.0002 1.0000 0.8265
-0.250 -0.0805 0.02817 0.01814 -0.0009 1.0000 0.8317
0.000 -0.0515 0.02819 0.01817 -0.0021 0.9968 0.8363
0.250 -0.0126 0.02836 0.01834 -0.0053 0.9902 0.8402
0.500 0.0294 0.02860 0.01860 -0.0092 0.9832 0.8438
0.750 0.0678 0.02875 0.01880 -0.0122 0.9745 0.8472
1.000 0.1062 0.02885 0.01897 -0.0149 0.9648 0.8508
1.250 0.1490 0.02900 0.01920 -0.0184 0.9548 0.8545
1.500 0.1911 0.02908 0.01939 -0.0218 0.9424 0.8579
1.750 0.2338 0.02912 0.01956 -0.0252 0.9285 0.8608
2.000 0.2807 0.02883 0.01943 -0.0285 0.9108 0.8631
2.250 0.3218 0.02820 0.01896 -0.0301 0.8844 0.8657
2.500 0.3688 0.02718 0.01813 -0.0318 0.8531 0.8685
2.750 0.4129 0.02602 0.01714 -0.0327 0.8180 0.8714
3.000 0.4455 0.02518 0.01647 -0.0320 0.7753 0.8748
3.250 0.4765 0.02436 0.01578 -0.0308 0.7144 0.8780
3.500 0.5307 0.02333 0.01404 -0.0315 0.5239 0.8801
3.750 0.5480 0.02457 0.01413 -0.0292 0.3543 0.8833
4.000 0.5648 0.02589 0.01480 -0.0280 0.2624 0.8868
4.250 0.5869 0.02705 0.01559 -0.0277 0.2141 0.8902
4.500 0.6096 0.02798 0.01633 -0.0272 0.1876 0.8936
4.750 0.6347 0.02881 0.01710 -0.0270 0.1681 0.8973
5.000 0.6626 0.02972 0.01795 -0.0274 0.1538 0.9009
5.250 0.6935 0.03070 0.01889 -0.0282 0.1416 0.9043
5.500 0.7243 0.03159 0.01990 -0.0288 0.1312 0.9076
5.750 0.7542 0.03265 0.02092 -0.0292 0.1230 0.9114
6.000 0.7841 0.03373 0.02216 -0.0298 0.1148 0.9156
6.250 0.8142 0.03508 0.02352 -0.0305 0.1083 0.9196
6.500 0.8410 0.03632 0.02500 -0.0304 0.1018 0.9239
6.750 0.8678 0.03771 0.02632 -0.0308 0.0969 0.9284
7.250 0.9171 0.04116 0.03045 -0.0305 0.0879 0.9383
7.500 0.9409 0.04276 0.03207 -0.0306 0.0844 0.9440
7.750 0.9596 0.04511 0.03492 -0.0297 0.0810 0.9499
8.000 0.9768 0.04775 0.03803 -0.0289 0.0785 0.9564
8.250 0.9933 0.05024 0.04086 -0.0283 0.0761 0.9639
8.500 1.0116 0.05237 0.04314 -0.0281 0.0738 0.9755
8.750 1.0250 0.05527 0.04632 -0.0277 0.0717 1.0000
9.000 1.0252 0.05942 0.05111 -0.0262 0.0702 1.0000
9.250 1.0220 0.06374 0.05593 -0.0249 0.0694 1.0000
9.500 1.0131 0.06826 0.06086 -0.0235 0.0689 1.0000
9.750 0.9965 0.07282 0.06578 -0.0220 0.0686 1.0000
10.000 0.9741 0.07796 0.07122 -0.0215 0.0686 1.0000
10.250 0.9461 0.08424 0.07775 -0.0228 0.0688 1.0000
10.500 0.9132 0.09236 0.08609 -0.0269 0.0693 1.0000
10.750 0.8806 0.10276 0.09663 -0.0343 0.0700 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA SC(2)-0612 AIRFOIL (sc20612-il)