NACA M22 AIRFOIL (m22-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M22 AIRFOIL (m22-il) Reynolds number: 100,000 Max Cl/Cd: 55.14 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m22-il-100000-n5.txt Download as CSV file: xf-m22-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M22 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
------ -------- --------- --------- -------- -------- --------
-7.000 -0.3648 0.10126 0.09683 0.0094 0.7303 0.0235
-6.750 -0.3568 0.09693 0.09246 0.0097 0.7245 0.0244
-6.500 -0.3455 0.09350 0.08900 0.0080 0.7182 0.0254
-6.250 -0.3323 0.09023 0.08568 0.0056 0.7120 0.0265
-6.000 -0.3176 0.08699 0.08237 0.0030 0.7067 0.0276
-5.750 -0.3001 0.08369 0.07902 -0.0001 0.7002 0.0288
-5.500 -0.2805 0.08063 0.07583 -0.0033 0.6950 0.0304
-5.250 -0.2509 0.07848 0.07355 -0.0088 0.6888 0.0319
-5.000 -0.2212 0.07628 0.07116 -0.0135 0.6830 0.0324
-4.750 -0.1568 0.05478 0.04988 -0.0200 0.6647 0.0331
-4.250 -0.1654 0.06393 0.05857 -0.0179 0.6673 0.0345
-4.000 -0.1452 0.06063 0.05517 -0.0190 0.6620 0.0362
-3.750 -0.1202 0.05770 0.05210 -0.0209 0.6562 0.0386
-3.500 -0.0724 0.05759 0.05150 -0.0252 0.6513 0.0443
-3.250 -0.0575 0.05249 0.04642 -0.0257 0.6459 0.0460
-3.000 -0.0377 0.04942 0.04327 -0.0260 0.6405 0.0494
-2.750 0.0025 0.04888 0.04227 -0.0281 0.6362 0.0570
-2.500 0.0265 0.04517 0.03847 -0.0291 0.6303 0.0590
-2.250 0.0464 0.04252 0.03575 -0.0292 0.6252 0.0657
-2.000 0.0841 0.04150 0.03423 -0.0304 0.6209 0.0719
-1.750 0.1054 0.03843 0.03115 -0.0308 0.6150 0.0746
-1.500 0.1330 0.03671 0.02920 -0.0312 0.6102 0.0806
-1.250 0.1608 0.03496 0.02722 -0.0316 0.6057 0.0895
-1.000 0.1955 0.03462 0.02646 -0.0321 0.5999 0.0994
-0.750 0.2194 0.03208 0.02384 -0.0323 0.5955 0.1017
-0.500 0.2468 0.03061 0.02217 -0.0324 0.5909 0.1052
0.000 0.3145 0.02680 0.01743 -0.0318 0.5813 0.0561
0.250 0.3432 0.02564 0.01595 -0.0317 0.5771 0.0589
0.500 0.3706 0.02487 0.01517 -0.0322 0.5712 0.0626
0.750 0.4016 0.02399 0.01378 -0.0317 0.5669 0.0577
1.000 0.4295 0.02296 0.01258 -0.0317 0.5628 0.0566
1.250 0.4583 0.02221 0.01172 -0.0320 0.5571 0.0556
1.500 0.4866 0.02152 0.01088 -0.0319 0.5526 0.0547
1.750 0.5149 0.02093 0.01014 -0.0318 0.5487 0.0540
2.000 0.5438 0.02049 0.00965 -0.0322 0.5430 0.0534
2.250 0.5732 0.02007 0.00913 -0.0324 0.5384 0.0531
2.500 0.6026 0.01970 0.00863 -0.0325 0.5346 0.0532
2.750 0.6302 0.01954 0.00846 -0.0327 0.5288 0.0536
3.000 0.6564 0.01939 0.00824 -0.0324 0.5244 0.0545
3.250 0.6823 0.01927 0.00799 -0.0319 0.5208 0.0560
3.500 0.7093 0.01936 0.00811 -0.0321 0.5150 0.0587
3.750 0.7359 0.01941 0.00818 -0.0320 0.5104 0.0652
4.000 0.7623 0.01940 0.00819 -0.0317 0.5069 0.0894
4.500 0.8507 0.01855 0.00901 -0.0396 0.4955 1.0000
4.750 0.8759 0.01874 0.00914 -0.0391 0.4921 1.0000
5.000 0.9018 0.01916 0.00968 -0.0392 0.4865 1.0000
5.250 0.9272 0.01945 0.01004 -0.0390 0.4818 1.0000
5.500 0.9524 0.01966 0.01030 -0.0386 0.4781 1.0000
5.750 0.9776 0.01983 0.01062 -0.0384 0.4680 1.0000
6.000 1.0021 0.01934 0.01008 -0.0375 0.4493 1.0000
6.250 1.0272 0.01951 0.01042 -0.0373 0.4374 1.0000
6.500 1.0513 0.01927 0.01020 -0.0367 0.4026 1.0000
6.750 1.0752 0.01950 0.01047 -0.0364 0.3670 1.0000
7.000 1.0881 0.02154 0.01149 -0.0364 0.2080 1.0000
7.250 1.0825 0.02632 0.01511 -0.0361 0.0342 1.0000
7.500 1.0963 0.02781 0.01672 -0.0352 0.0290 1.0000
7.750 1.1077 0.02940 0.01852 -0.0344 0.0256 1.0000
8.000 1.1131 0.03147 0.02079 -0.0334 0.0230 1.0000
8.500 1.1100 0.03653 0.02628 -0.0316 0.0215 1.0000
8.750 1.1078 0.03927 0.02921 -0.0309 0.0212 1.0000
9.000 1.1050 0.04223 0.03234 -0.0305 0.0209 1.0000
9.250 1.1017 0.04532 0.03558 -0.0301 0.0206 1.0000
9.500 1.0986 0.04840 0.03882 -0.0297 0.0202 1.0000
9.750 1.0967 0.05134 0.04188 -0.0292 0.0197 1.0000
10.000 1.0965 0.05399 0.04464 -0.0286 0.0189 1.0000
10.250 1.0987 0.05623 0.04696 -0.0275 0.0179 1.0000
10.500 1.1049 0.05783 0.04861 -0.0259 0.0170 1.0000
10.750 1.1182 0.05848 0.04925 -0.0232 0.0166 1.0000
11.000 1.1395 0.05856 0.04939 -0.0199 0.0162 1.0000
11.250 1.1667 0.05875 0.04963 -0.0164 0.0159 1.0000
11.500 1.1933 0.05976 0.05077 -0.0136 0.0157 1.0000
11.750 1.2121 0.06171 0.05287 -0.0118 0.0149 1.0000
12.000 1.2239 0.06440 0.05572 -0.0106 0.0141 1.0000
12.250 1.2344 0.06772 0.05923 -0.0094 0.0135 1.0000
12.500 1.2391 0.07125 0.06300 -0.0086 0.0133 1.0000
12.750 1.2386 0.07472 0.06673 -0.0083 0.0134 1.0000
13.000 1.2351 0.07836 0.07062 -0.0081 0.0135 1.0000
13.250 1.2293 0.08220 0.07472 -0.0083 0.0136 1.0000
13.500 1.2213 0.08631 0.07908 -0.0088 0.0137 1.0000
13.750 1.2114 0.09070 0.08372 -0.0096 0.0139 1.0000
14.000 1.1998 0.09539 0.08866 -0.0108 0.0141 1.0000
14.250 1.1869 0.10042 0.09392 -0.0124 0.0143 1.0000
14.500 1.1709 0.10595 0.09970 -0.0146 0.0145 1.0000
14.750 1.1546 0.11198 0.10597 -0.0174 0.0147 1.0000
15.000 1.1384 0.11848 0.11269 -0.0207 0.0149 1.0000
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Polar data table (+)
Polar graphs
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