NASA SC(2)-0012 AIRFOIL (sc20012-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA SC(2)-0012 AIRFOIL (sc20012-il) Reynolds number: 100,000 Max Cl/Cd: 29.67 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-sc20012-il-100000-n5.txt Download as CSV file: xf-sc20012-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0012 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.8509 0.07995 0.07399 -0.0256 1.0000 0.0396
-12.250 -0.8938 0.07036 0.06414 -0.0311 1.0000 0.0392
-12.000 -0.9261 0.06384 0.05734 -0.0330 1.0000 0.0392
-11.750 -0.9515 0.05890 0.05211 -0.0328 1.0000 0.0394
-11.500 -0.9727 0.05487 0.04776 -0.0310 1.0000 0.0398
-11.250 -0.9908 0.05139 0.04392 -0.0279 1.0000 0.0403
-11.000 -1.0027 0.04774 0.03978 -0.0250 1.0000 0.0410
-10.750 -0.9993 0.04523 0.03702 -0.0231 1.0000 0.0418
-10.500 -0.9879 0.04353 0.03522 -0.0218 1.0000 0.0425
-10.250 -0.9765 0.04168 0.03321 -0.0205 1.0000 0.0435
-10.000 -0.9644 0.03969 0.03097 -0.0191 1.0000 0.0446
-9.750 -0.9514 0.03761 0.02856 -0.0176 1.0000 0.0463
-9.500 -0.9366 0.03568 0.02630 -0.0163 1.0000 0.0482
-9.250 -0.9178 0.03458 0.02521 -0.0156 1.0000 0.0499
-9.000 -0.8990 0.03329 0.02376 -0.0147 1.0000 0.0520
-8.750 -0.8795 0.03176 0.02192 -0.0137 1.0000 0.0546
-8.500 -0.8590 0.03055 0.02070 -0.0130 1.0000 0.0568
-8.250 -0.8381 0.02958 0.01969 -0.0122 1.0000 0.0594
-8.000 -0.8165 0.02847 0.01836 -0.0114 1.0000 0.0627
-7.750 -0.7950 0.02741 0.01730 -0.0107 1.0000 0.0655
-7.500 -0.7734 0.02654 0.01643 -0.0100 1.0000 0.0687
-7.250 -0.7512 0.02568 0.01541 -0.0092 1.0000 0.0725
-7.000 -0.7300 0.02476 0.01454 -0.0083 1.0000 0.0758
-6.750 -0.7089 0.02401 0.01380 -0.0074 1.0000 0.0795
-6.500 -0.6875 0.02331 0.01301 -0.0064 1.0000 0.0838
-6.250 -0.6683 0.02253 0.01230 -0.0051 1.0000 0.0876
-6.000 -0.6494 0.02191 0.01170 -0.0038 1.0000 0.0921
-5.750 -0.6314 0.02134 0.01111 -0.0022 1.0000 0.0971
-5.500 -0.6156 0.02075 0.01059 -0.0003 1.0000 0.1021
-5.250 -0.6003 0.02031 0.01013 0.0017 1.0000 0.1082
-5.000 -0.5814 0.01976 0.00967 0.0028 0.9980 0.1156
-4.750 -0.5485 0.01912 0.00908 0.0012 0.9912 0.1280
-4.500 -0.5157 0.01848 0.00853 -0.0004 0.9843 0.1459
-4.250 -0.4838 0.01782 0.00804 -0.0018 0.9773 0.1740
-4.000 -0.4520 0.01708 0.00759 -0.0033 0.9711 0.2240
-3.750 -0.4240 0.01619 0.00715 -0.0041 0.9641 0.3091
-3.500 -0.3994 0.01510 0.00683 -0.0042 0.9572 0.4490
-3.250 -0.3739 0.01470 0.00723 -0.0031 0.9516 0.6030
-3.000 -0.3460 0.01476 0.00733 -0.0028 0.9443 0.6699
-2.750 -0.3154 0.01485 0.00741 -0.0029 0.9388 0.7098
-2.500 -0.2870 0.01498 0.00756 -0.0025 0.9325 0.7388
-2.250 -0.2589 0.01511 0.00767 -0.0020 0.9259 0.7617
-2.000 -0.2287 0.01522 0.00777 -0.0019 0.9209 0.7812
-1.750 -0.2017 0.01540 0.00794 -0.0010 0.9151 0.8014
-1.500 -0.1754 0.01559 0.00815 0.0001 0.9092 0.8200
-1.250 -0.1459 0.01575 0.00830 0.0005 0.9043 0.8353
-1.000 -0.1161 0.01583 0.00834 0.0006 0.8995 0.8479
-0.750 -0.0890 0.01593 0.00844 0.0010 0.8932 0.8579
-0.500 -0.0602 0.01594 0.00843 0.0008 0.8882 0.8657
-0.250 -0.0277 0.01594 0.00842 0.0001 0.8837 0.8709
0.000 0.0000 0.01595 0.00842 0.0000 0.8774 0.8774
0.250 0.0277 0.01594 0.00842 -0.0001 0.8709 0.8838
0.500 0.0602 0.01594 0.00843 -0.0008 0.8657 0.8882
0.750 0.0890 0.01593 0.00844 -0.0010 0.8579 0.8932
1.000 0.1161 0.01582 0.00834 -0.0006 0.8479 0.8995
1.250 0.1459 0.01575 0.00830 -0.0005 0.8353 0.9043
1.500 0.1754 0.01559 0.00815 -0.0001 0.8200 0.9092
1.750 0.2017 0.01540 0.00794 0.0010 0.8013 0.9151
2.000 0.2287 0.01522 0.00777 0.0019 0.7812 0.9210
2.250 0.2589 0.01511 0.00767 0.0020 0.7617 0.9259
2.500 0.2870 0.01498 0.00756 0.0025 0.7388 0.9325
2.750 0.3154 0.01485 0.00740 0.0029 0.7099 0.9388
3.000 0.3460 0.01476 0.00733 0.0028 0.6700 0.9443
3.250 0.3739 0.01469 0.00723 0.0031 0.6030 0.9517
3.500 0.3994 0.01510 0.00682 0.0042 0.4488 0.9572
3.750 0.4240 0.01619 0.00715 0.0041 0.3090 0.9641
4.000 0.4521 0.01708 0.00759 0.0033 0.2239 0.9711
4.250 0.4839 0.01782 0.00804 0.0018 0.1740 0.9774
4.500 0.5158 0.01848 0.00853 0.0004 0.1459 0.9843
4.750 0.5486 0.01912 0.00907 -0.0012 0.1280 0.9912
5.000 0.5815 0.01976 0.00966 -0.0028 0.1156 0.9980
5.250 0.6002 0.02030 0.01013 -0.0016 0.1082 1.0000
5.500 0.6156 0.02075 0.01059 0.0004 0.1021 1.0000
5.750 0.6314 0.02134 0.01110 0.0022 0.0971 1.0000
6.000 0.6494 0.02191 0.01170 0.0038 0.0921 1.0000
6.250 0.6683 0.02253 0.01230 0.0052 0.0876 1.0000
6.500 0.6875 0.02331 0.01300 0.0064 0.0838 1.0000
6.750 0.7089 0.02401 0.01380 0.0074 0.0795 1.0000
7.000 0.7300 0.02476 0.01454 0.0083 0.0758 1.0000
7.250 0.7513 0.02568 0.01541 0.0092 0.0725 1.0000
7.500 0.7735 0.02654 0.01643 0.0100 0.0687 1.0000
7.750 0.7951 0.02741 0.01730 0.0107 0.0656 1.0000
8.000 0.8166 0.02847 0.01836 0.0114 0.0627 1.0000
8.250 0.8382 0.02958 0.01968 0.0122 0.0594 1.0000
8.500 0.8591 0.03055 0.02069 0.0129 0.0568 1.0000
8.750 0.8796 0.03177 0.02193 0.0136 0.0545 1.0000
9.000 0.8991 0.03330 0.02377 0.0146 0.0520 1.0000
9.250 0.9179 0.03458 0.02521 0.0155 0.0499 1.0000
9.500 0.9367 0.03568 0.02629 0.0163 0.0482 1.0000
9.750 0.9516 0.03761 0.02856 0.0176 0.0463 1.0000
10.000 0.9646 0.03970 0.03097 0.0190 0.0446 1.0000
10.250 0.9767 0.04169 0.03321 0.0204 0.0435 1.0000
10.500 0.9881 0.04353 0.03523 0.0218 0.0425 1.0000
10.750 0.9997 0.04523 0.03701 0.0230 0.0417 1.0000
11.000 1.0030 0.04775 0.03979 0.0249 0.0410 1.0000
11.250 0.9910 0.05142 0.04395 0.0278 0.0403 1.0000
11.500 0.9729 0.05491 0.04780 0.0309 0.0398 1.0000
11.750 0.9520 0.05894 0.05215 0.0327 0.0394 1.0000
12.000 0.9265 0.06389 0.05739 0.0329 0.0392 1.0000
12.250 0.8942 0.07045 0.06423 0.0309 0.0392 1.0000
12.500 0.8513 0.08007 0.07412 0.0254 0.0397 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA SC(2)-0012 AIRFOIL (sc20012-il)