NACA M2 AIRFOIL (m2-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M2 AIRFOIL (m2-il) Reynolds number: 500,000 Max Cl/Cd: 45.94 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m2-il-500000.txt Download as CSV file: xf-m2-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M2 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.8926 0.05734 0.05486 -0.0160 1.0000 0.0152
-10.250 -0.9149 0.05221 0.04948 -0.0155 1.0000 0.0151
-10.000 -0.9346 0.04600 0.04289 -0.0144 1.0000 0.0151
-9.750 -0.9532 0.03873 0.03505 -0.0125 1.0000 0.0152
-9.500 -0.9549 0.03361 0.02948 -0.0108 1.0000 0.0155
-9.250 -0.9381 0.03209 0.02783 -0.0101 1.0000 0.0160
-9.000 -0.9190 0.03088 0.02651 -0.0096 1.0000 0.0166
-8.750 -0.9001 0.02933 0.02479 -0.0089 1.0000 0.0172
-8.500 -0.8807 0.02767 0.02288 -0.0081 1.0000 0.0182
-8.250 -0.8601 0.02606 0.02095 -0.0073 1.0000 0.0193
-8.000 -0.8363 0.02550 0.02015 -0.0067 1.0000 0.0200
-7.750 -0.8209 0.02137 0.01565 -0.0056 1.0000 0.0212
-7.500 -0.7971 0.02049 0.01471 -0.0053 1.0000 0.0222
-7.250 -0.7726 0.01965 0.01378 -0.0050 1.0000 0.0235
-7.000 -0.7475 0.01897 0.01297 -0.0046 1.0000 0.0250
-6.750 -0.7216 0.01861 0.01248 -0.0043 1.0000 0.0263
-6.500 -0.6991 0.01644 0.01010 -0.0036 1.0000 0.0280
-6.250 -0.6750 0.01537 0.00900 -0.0032 1.0000 0.0296
-6.000 -0.6497 0.01477 0.00836 -0.0028 1.0000 0.0315
-5.750 -0.6244 0.01420 0.00772 -0.0024 1.0000 0.0334
-5.500 -0.5991 0.01364 0.00710 -0.0020 1.0000 0.0347
-5.250 -0.5738 0.01321 0.00660 -0.0015 1.0000 0.0355
-5.000 -0.5518 0.01201 0.00534 -0.0006 1.0000 0.0382
-4.750 -0.5279 0.01149 0.00480 0.0001 1.0000 0.0402
-4.500 -0.5040 0.01107 0.00435 0.0008 1.0000 0.0420
-4.250 -0.4800 0.01072 0.00397 0.0016 1.0000 0.0440
-4.000 -0.4560 0.01043 0.00364 0.0023 1.0000 0.0456
-3.750 -0.4321 0.01003 0.00322 0.0031 1.0000 0.0491
-3.500 -0.4080 0.00973 0.00296 0.0038 1.0000 0.0548
-3.250 -0.3838 0.00943 0.00280 0.0044 1.0000 0.0773
-3.000 -0.3593 0.00925 0.00272 0.0050 1.0000 0.1014
-2.750 -0.3346 0.00914 0.00261 0.0056 1.0000 0.1127
-2.500 -0.3100 0.00900 0.00251 0.0061 1.0000 0.1223
-2.250 -0.2852 0.00886 0.00241 0.0066 1.0000 0.1322
-2.000 -0.2513 0.00865 0.00228 0.0051 0.9979 0.1479
-1.750 -0.2117 0.00811 0.00207 0.0021 0.9947 0.2257
-1.500 -0.1773 0.00670 0.00188 -0.0003 0.9910 0.5375
-1.250 -0.1390 0.00611 0.00181 -0.0027 0.9866 0.6675
-1.000 -0.1012 0.00577 0.00179 -0.0048 0.9813 0.7492
-0.750 -0.0689 0.00552 0.00180 -0.0054 0.9730 0.8132
-0.500 -0.0436 0.00542 0.00193 -0.0041 0.9627 0.8733
-0.250 -0.0212 0.00545 0.00202 -0.0021 0.9492 0.9108
0.000 0.0000 0.00546 0.00205 0.0000 0.9326 0.9326
0.250 0.0212 0.00545 0.00203 0.0021 0.9109 0.9493
0.500 0.0437 0.00542 0.00193 0.0041 0.8737 0.9627
0.750 0.0690 0.00552 0.00180 0.0053 0.8140 0.9730
1.000 0.1012 0.00577 0.00179 0.0048 0.7492 0.9813
1.250 0.1390 0.00611 0.00181 0.0027 0.6676 0.9866
1.500 0.1773 0.00670 0.00188 0.0003 0.5375 0.9910
1.750 0.2118 0.00811 0.00207 -0.0021 0.2252 0.9947
2.000 0.2513 0.00865 0.00228 -0.0051 0.1479 0.9979
2.250 0.2851 0.00886 0.00241 -0.0066 0.1321 1.0000
2.500 0.3099 0.00900 0.00251 -0.0061 0.1224 1.0000
2.750 0.3346 0.00914 0.00261 -0.0056 0.1127 1.0000
3.000 0.3592 0.00925 0.00272 -0.0050 0.1014 1.0000
3.250 0.3837 0.00942 0.00280 -0.0044 0.0775 1.0000
3.500 0.4079 0.00973 0.00296 -0.0038 0.0549 1.0000
3.750 0.4321 0.01003 0.00322 -0.0031 0.0491 1.0000
4.000 0.4559 0.01043 0.00364 -0.0023 0.0456 1.0000
4.250 0.4800 0.01072 0.00397 -0.0016 0.0440 1.0000
4.500 0.5039 0.01107 0.00435 -0.0008 0.0420 1.0000
4.750 0.5278 0.01149 0.00480 -0.0001 0.0402 1.0000
5.000 0.5517 0.01201 0.00534 0.0006 0.0382 1.0000
5.250 0.5738 0.01319 0.00658 0.0015 0.0355 1.0000
5.500 0.5991 0.01365 0.00710 0.0020 0.0347 1.0000
5.750 0.6244 0.01420 0.00772 0.0024 0.0334 1.0000
6.000 0.6497 0.01477 0.00835 0.0028 0.0315 1.0000
6.250 0.6750 0.01537 0.00899 0.0031 0.0296 1.0000
6.500 0.6991 0.01644 0.01009 0.0036 0.0280 1.0000
6.750 0.7216 0.01860 0.01248 0.0043 0.0263 1.0000
7.000 0.7475 0.01899 0.01299 0.0046 0.0251 1.0000
7.250 0.7726 0.01968 0.01381 0.0050 0.0235 1.0000
7.500 0.7971 0.02050 0.01472 0.0053 0.0223 1.0000
7.750 0.8210 0.02138 0.01566 0.0056 0.0212 1.0000
8.000 0.8364 0.02549 0.02013 0.0067 0.0200 1.0000
8.250 0.8601 0.02606 0.02096 0.0073 0.0193 1.0000
8.500 0.8808 0.02763 0.02284 0.0081 0.0182 1.0000
8.750 0.9003 0.02931 0.02476 0.0089 0.0172 1.0000
9.000 0.9192 0.03086 0.02648 0.0095 0.0166 1.0000
9.250 0.9383 0.03208 0.02782 0.0101 0.0160 1.0000
9.500 0.9551 0.03359 0.02946 0.0108 0.0155 1.0000
9.750 0.9533 0.03876 0.03507 0.0125 0.0152 1.0000
10.000 0.9343 0.04609 0.04298 0.0144 0.0151 1.0000
10.250 0.9151 0.05224 0.04951 0.0154 0.0151 1.0000
10.500 0.8925 0.05744 0.05496 0.0159 0.0152 1.0000
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Polar data table (+)
Polar graphs
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