AIRFOIL PROFILE12A 9.00% (rg12a-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: AIRFOIL PROFILE12A 9.00% (rg12a-il) Reynolds number: 50,000 Max Cl/Cd: 33.99 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rg12a-il-50000.txt Download as CSV file: xf-rg12a-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4618 0.10713 0.10013 -0.0110 1.0000 0.2412
-8.750 -0.4612 0.10397 0.09705 -0.0110 1.0000 0.2516
-8.500 -0.4838 0.10315 0.09639 -0.0124 1.0000 0.2611
-8.250 -0.4567 0.09777 0.09097 -0.0098 1.0000 0.2800
-8.000 -0.4518 0.09466 0.08793 -0.0088 1.0000 0.2966
-7.750 -0.4500 0.09183 0.08517 -0.0077 1.0000 0.3132
-7.500 -0.4635 0.09008 0.08356 -0.0065 1.0000 0.3317
-7.250 -0.4519 0.08624 0.07976 -0.0048 1.0000 0.3494
-7.000 -0.4423 0.08290 0.07643 -0.0030 1.0000 0.3678
-6.750 -0.4665 0.08210 0.07582 0.0000 1.0000 0.3896
-5.750 -0.4915 0.05073 0.04331 -0.0352 1.0000 0.1312
-5.500 -0.4766 0.04575 0.03765 -0.0355 1.0000 0.1187
-5.250 -0.4594 0.04189 0.03355 -0.0347 1.0000 0.1150
-5.000 -0.4397 0.03787 0.02888 -0.0341 1.0000 0.1094
-4.750 -0.4169 0.03464 0.02480 -0.0331 1.0000 0.1067
-4.500 -0.3949 0.03184 0.02186 -0.0321 1.0000 0.1099
-4.250 -0.3714 0.02973 0.01927 -0.0310 1.0000 0.1183
-4.000 -0.3474 0.02742 0.01682 -0.0298 1.0000 0.1272
-3.750 -0.3231 0.02551 0.01474 -0.0285 1.0000 0.1445
-3.500 -0.2989 0.02361 0.01287 -0.0270 1.0000 0.1769
-3.250 -0.2769 0.02158 0.01135 -0.0254 1.0000 0.2506
-3.000 -0.1105 0.01900 0.01072 -0.0364 1.0000 1.0000
-2.750 -0.1061 0.01861 0.01013 -0.0337 1.0000 1.0000
-2.500 -0.1020 0.01831 0.00965 -0.0308 1.0000 1.0000
-2.250 -0.0985 0.01806 0.00925 -0.0276 1.0000 1.0000
-2.000 -0.0957 0.01786 0.00887 -0.0243 1.0000 1.0000
-1.750 -0.0932 0.01770 0.00858 -0.0208 1.0000 1.0000
-1.500 -0.0899 0.01759 0.00833 -0.0176 1.0000 1.0000
-1.250 -0.0836 0.01754 0.00813 -0.0148 1.0000 1.0000
-1.000 -0.0729 0.01757 0.00799 -0.0129 1.0000 1.0000
-0.750 -0.0591 0.01769 0.00793 -0.0115 1.0000 1.0000
-0.500 -0.0433 0.01787 0.00792 -0.0105 1.0000 1.0000
-0.250 -0.0265 0.01811 0.00802 -0.0097 1.0000 1.0000
0.000 -0.0092 0.01841 0.00819 -0.0091 1.0000 1.0000
0.250 0.0084 0.01878 0.00844 -0.0087 1.0000 1.0000
0.500 0.0259 0.01921 0.00877 -0.0083 1.0000 1.0000
0.750 0.0433 0.01970 0.00916 -0.0080 1.0000 1.0000
1.000 0.0606 0.02026 0.00965 -0.0078 1.0000 1.0000
1.250 0.0776 0.02088 0.01022 -0.0078 1.0000 1.0000
1.500 0.0942 0.02157 0.01087 -0.0078 1.0000 1.0000
1.750 0.1318 0.02267 0.01196 -0.0118 0.9899 1.0000
2.000 0.1922 0.02402 0.01333 -0.0196 0.9687 1.0000
2.250 0.2468 0.02501 0.01438 -0.0260 0.9462 1.0000
2.500 0.2986 0.02579 0.01526 -0.0315 0.9231 1.0000
2.750 0.3565 0.02640 0.01601 -0.0374 0.9009 1.0000
3.000 0.4045 0.02678 0.01654 -0.0413 0.8764 1.0000
3.250 0.4531 0.02700 0.01697 -0.0448 0.8519 1.0000
3.500 0.5117 0.02680 0.01703 -0.0492 0.8282 1.0000
3.750 0.5699 0.02628 0.01683 -0.0528 0.8039 1.0000
4.000 0.6154 0.02586 0.01664 -0.0540 0.7769 1.0000
4.250 0.6594 0.02531 0.01631 -0.0544 0.7485 1.0000
4.500 0.7014 0.02469 0.01593 -0.0541 0.7182 1.0000
4.750 0.7298 0.02462 0.01600 -0.0523 0.6835 1.0000
5.000 0.7630 0.02429 0.01575 -0.0506 0.6477 1.0000
5.250 0.7876 0.02439 0.01591 -0.0482 0.6081 1.0000
5.500 0.8118 0.02453 0.01609 -0.0456 0.5662 1.0000
5.750 0.8349 0.02475 0.01625 -0.0430 0.5215 1.0000
6.000 0.8550 0.02516 0.01657 -0.0402 0.4734 1.0000
6.250 0.8740 0.02571 0.01693 -0.0374 0.4216 1.0000
6.500 0.8899 0.02654 0.01755 -0.0344 0.3665 1.0000
6.750 0.9046 0.02769 0.01846 -0.0314 0.3101 1.0000
7.000 0.9177 0.02918 0.01964 -0.0285 0.2546 1.0000
7.250 0.9298 0.03095 0.02113 -0.0258 0.2044 1.0000
7.500 0.9449 0.03319 0.02313 -0.0237 0.1645 1.0000
7.750 0.9633 0.03571 0.02558 -0.0222 0.1378 1.0000
8.000 0.9851 0.03888 0.02897 -0.0209 0.1232 1.0000
8.250 1.0071 0.04224 0.03241 -0.0201 0.1135 1.0000
8.500 1.0190 0.04535 0.03613 -0.0179 0.1065 1.0000
8.750 1.0373 0.04906 0.03993 -0.0171 0.1008 1.0000
9.000 1.0423 0.05320 0.04472 -0.0148 0.0997 1.0000
9.250 1.0434 0.05766 0.04971 -0.0126 0.0993 1.0000
9.500 1.0415 0.06230 0.05478 -0.0107 0.0996 1.0000
9.750 1.0365 0.06711 0.05992 -0.0091 0.1002 1.0000
10.000 1.0300 0.07209 0.06515 -0.0078 0.1008 1.0000
10.250 0.9840 0.07641 0.06994 -0.0055 0.1038 1.0000
10.500 0.9356 0.08268 0.07638 -0.0069 0.1071 1.0000
10.750 0.9028 0.09032 0.08401 -0.0113 0.1101 1.0000
11.000 0.8861 0.09783 0.09150 -0.0153 0.1125 1.0000
11.250 0.8877 0.10381 0.09748 -0.0166 0.1142 1.0000
11.500 0.8100 0.12623 0.11959 -0.0376 0.1489 1.0000
11.750 0.8245 0.13201 0.12545 -0.0376 0.1552 1.0000
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Polar data table (+)
Polar graphs
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