AIRFOIL PROFILE12A 9.00% (rg12a-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: AIRFOIL PROFILE12A 9.00% (rg12a-il) Reynolds number: 1,000,000 Max Cl/Cd: 96.02 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rg12a-il-1000000.txt Download as CSV file: xf-rg12a-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: AIRFOIL PROFILE12A 9.00%
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4930 0.08801 0.08638 -0.0253 1.0000 0.0096
-9.250 -0.4982 0.08248 0.08088 -0.0277 1.0000 0.0096
-9.000 -0.5022 0.07814 0.07656 -0.0296 1.0000 0.0096
-8.750 -0.5091 0.07333 0.07179 -0.0320 1.0000 0.0096
-8.500 -0.5248 0.06691 0.06539 -0.0371 1.0000 0.0096
-8.250 -0.5448 0.06175 0.06020 -0.0385 1.0000 0.0096
-8.000 -0.5567 0.05600 0.05436 -0.0391 1.0000 0.0096
-7.750 -0.5673 0.05057 0.04880 -0.0382 1.0000 0.0096
-7.500 -0.5774 0.04555 0.04358 -0.0363 1.0000 0.0096
-6.250 -0.4791 0.01919 0.01502 -0.0455 0.9881 0.0101
-6.000 -0.4501 0.01586 0.01137 -0.0462 0.9865 0.0083
-5.750 -0.4189 0.01337 0.00852 -0.0469 0.9851 0.0077
-5.500 -0.3855 0.01205 0.00704 -0.0482 0.9840 0.0076
-5.250 -0.3543 0.01118 0.00607 -0.0490 0.9818 0.0079
-5.000 -0.3252 0.01052 0.00534 -0.0493 0.9780 0.0085
-4.750 -0.2922 0.00996 0.00472 -0.0505 0.9755 0.0092
-4.500 -0.2579 0.00940 0.00408 -0.0519 0.9735 0.0100
-4.250 -0.2228 0.00875 0.00336 -0.0536 0.9718 0.0114
-4.000 -0.1864 0.00839 0.00298 -0.0555 0.9706 0.0133
-3.750 -0.1598 0.00799 0.00255 -0.0551 0.9644 0.0175
-3.500 -0.1269 0.00747 0.00218 -0.0562 0.9608 0.0483
-3.250 -0.0927 0.00711 0.00194 -0.0577 0.9576 0.0809
-3.000 -0.0649 0.00675 0.00174 -0.0578 0.9494 0.1247
-2.750 -0.0312 0.00630 0.00154 -0.0592 0.9436 0.1934
-2.500 -0.0020 0.00582 0.00135 -0.0598 0.9330 0.2893
-2.250 0.0274 0.00523 0.00118 -0.0604 0.9208 0.4226
-2.000 0.0555 0.00474 0.00104 -0.0607 0.9057 0.5474
-1.750 0.0819 0.00436 0.00099 -0.0604 0.8884 0.6717
-1.500 0.1090 0.00433 0.00097 -0.0601 0.8687 0.7130
-1.250 0.1359 0.00435 0.00095 -0.0596 0.8475 0.7380
-1.000 0.1623 0.00441 0.00095 -0.0591 0.8251 0.7607
-0.750 0.1887 0.00449 0.00094 -0.0587 0.8020 0.7731
-0.500 0.2149 0.00457 0.00094 -0.0582 0.7776 0.7837
-0.250 0.2410 0.00466 0.00095 -0.0577 0.7530 0.7935
0.000 0.2673 0.00477 0.00096 -0.0572 0.7290 0.8028
0.250 0.2934 0.00486 0.00099 -0.0567 0.7057 0.8111
0.500 0.3198 0.00497 0.00102 -0.0563 0.6828 0.8188
0.750 0.3460 0.00508 0.00106 -0.0559 0.6604 0.8265
1.000 0.3724 0.00518 0.00109 -0.0555 0.6371 0.8341
1.250 0.3987 0.00529 0.00115 -0.0551 0.6147 0.8418
1.500 0.4247 0.00540 0.00120 -0.0547 0.5919 0.8499
1.750 0.4511 0.00553 0.00126 -0.0543 0.5685 0.8590
2.000 0.4767 0.00565 0.00133 -0.0538 0.5448 0.8667
2.250 0.5030 0.00580 0.00141 -0.0535 0.5197 0.8733
2.500 0.5288 0.00593 0.00148 -0.0530 0.4934 0.8787
2.750 0.5551 0.00609 0.00156 -0.0527 0.4675 0.8843
3.000 0.5810 0.00625 0.00165 -0.0524 0.4413 0.8899
3.250 0.6065 0.00643 0.00177 -0.0519 0.4137 0.8963
3.500 0.6318 0.00665 0.00189 -0.0515 0.3808 0.9031
3.750 0.6568 0.00684 0.00202 -0.0509 0.3542 0.9107
4.000 0.6807 0.00711 0.00217 -0.0502 0.3169 0.9196
4.250 0.7037 0.00741 0.00235 -0.0493 0.2772 0.9309
4.500 0.7261 0.00766 0.00251 -0.0482 0.2446 0.9454
4.750 0.7499 0.00793 0.00269 -0.0474 0.2143 0.9653
5.000 0.7813 0.00833 0.00294 -0.0485 0.1744 0.9880
5.250 0.8087 0.00877 0.00321 -0.0488 0.1406 1.0000
5.500 0.8338 0.00917 0.00349 -0.0485 0.1131 1.0000
5.750 0.8583 0.00962 0.00380 -0.0481 0.0861 1.0000
6.000 0.8824 0.01011 0.00415 -0.0477 0.0600 1.0000
6.250 0.9050 0.01076 0.00463 -0.0470 0.0306 1.0000
6.500 0.9288 0.01128 0.00510 -0.0463 0.0192 1.0000
6.750 0.9530 0.01174 0.00556 -0.0458 0.0164 1.0000
7.000 0.9762 0.01235 0.00626 -0.0449 0.0144 1.0000
7.250 1.0006 0.01275 0.00671 -0.0444 0.0139 1.0000
7.500 1.0247 0.01318 0.00719 -0.0439 0.0132 1.0000
7.750 1.0480 0.01368 0.00776 -0.0432 0.0124 1.0000
8.000 1.0708 0.01422 0.00837 -0.0425 0.0117 1.0000
8.250 1.0927 0.01484 0.00903 -0.0416 0.0108 1.0000
8.500 1.1103 0.01593 0.01024 -0.0400 0.0099 1.0000
8.750 1.1264 0.01715 0.01159 -0.0383 0.0094 1.0000
9.000 1.1492 0.01759 0.01209 -0.0376 0.0091 1.0000
9.250 1.1703 0.01820 0.01278 -0.0367 0.0087 1.0000
9.500 1.1907 0.01885 0.01351 -0.0357 0.0082 1.0000
9.750 1.2108 0.01949 0.01420 -0.0347 0.0077 1.0000
10.000 1.2301 0.02015 0.01492 -0.0336 0.0072 1.0000
10.250 1.2481 0.02090 0.01571 -0.0324 0.0068 1.0000
10.500 1.2475 0.02358 0.01862 -0.0286 0.0061 1.0000
10.750 1.2667 0.02404 0.01914 -0.0275 0.0059 1.0000
11.000 1.2809 0.02487 0.02008 -0.0257 0.0058 1.0000
11.250 1.2921 0.02570 0.02101 -0.0235 0.0055 1.0000
11.500 1.3018 0.02658 0.02200 -0.0211 0.0052 1.0000
11.750 1.3086 0.02771 0.02323 -0.0184 0.0050 1.0000
12.000 1.3164 0.02872 0.02434 -0.0162 0.0048 1.0000
12.250 1.3195 0.03014 0.02588 -0.0136 0.0047 1.0000
12.500 1.3223 0.03163 0.02749 -0.0114 0.0046 1.0000
12.750 1.3277 0.03292 0.02885 -0.0097 0.0044 1.0000
13.000 1.3304 0.03455 0.03057 -0.0082 0.0043 1.0000
13.250 1.3336 0.03621 0.03230 -0.0070 0.0041 1.0000
13.500 1.3233 0.03949 0.03575 -0.0057 0.0041 1.0000
13.750 1.3124 0.04313 0.03956 -0.0054 0.0040 1.0000
14.000 1.2956 0.04793 0.04454 -0.0060 0.0039 1.0000
14.250 1.2801 0.05313 0.04993 -0.0075 0.0039 1.0000
14.500 1.2606 0.05941 0.05640 -0.0101 0.0039 1.0000
14.750 1.2380 0.06682 0.06401 -0.0138 0.0038 1.0000
15.000 1.2313 0.07199 0.06929 -0.0167 0.0039 1.0000
15.250 1.2069 0.08074 0.07822 -0.0218 0.0038 1.0000
15.500 1.1855 0.08949 0.08712 -0.0270 0.0038 1.0000
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Polar data table (+)
Polar graphs
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