NACA M22 AIRFOIL (m22-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M22 AIRFOIL (m22-il) Reynolds number: 500,000 Max Cl/Cd: 100.27 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m22-il-500000.txt Download as CSV file: xf-m22-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.4463 0.10781 0.10503 0.0343 0.6819 0.0117
-7.500 -0.4400 0.10467 0.10188 0.0324 0.6764 0.0119
-7.250 -0.4346 0.10162 0.09879 0.0304 0.6715 0.0122
-7.000 -0.4253 0.09830 0.09548 0.0278 0.6658 0.0127
-6.750 -0.4135 0.09484 0.09197 0.0248 0.6606 0.0133
-6.500 -0.3968 0.09148 0.08856 0.0208 0.6558 0.0140
-6.250 -0.3757 0.08833 0.08537 0.0154 0.6504 0.0142
-6.000 -0.3533 0.08502 0.08197 0.0101 0.6456 0.0143
-5.750 -0.3305 0.08149 0.07835 0.0056 0.6408 0.0144
-5.500 -0.3194 0.07612 0.07298 0.0050 0.6357 0.0146
-5.250 -0.3085 0.07246 0.06928 0.0050 0.6307 0.0150
-5.000 -0.2895 0.06948 0.06625 0.0031 0.6255 0.0156
-4.750 -0.2671 0.06631 0.06301 0.0007 0.6202 0.0162
-4.500 -0.2429 0.06313 0.05972 -0.0018 0.6155 0.0171
-4.250 -0.2126 0.06019 0.05668 -0.0048 0.6103 0.0185
-4.000 -0.1724 0.05789 0.05419 -0.0089 0.6052 0.0191
-3.500 -0.1145 0.05157 0.04758 -0.0124 0.5960 0.0192
-3.250 -0.0998 0.04635 0.04231 -0.0131 0.5914 0.0199
-3.000 -0.0778 0.04395 0.03982 -0.0138 0.5869 0.0206
-2.750 -0.0515 0.04167 0.03745 -0.0147 0.5820 0.0217
-2.500 -0.0221 0.03949 0.03512 -0.0156 0.5770 0.0240
-2.250 0.0180 0.03854 0.03390 -0.0165 0.5726 0.0258
-2.000 0.0483 0.03641 0.03162 -0.0169 0.5680 0.0260
-1.750 0.0774 0.03397 0.02900 -0.0171 0.5634 0.0261
-1.500 0.0992 0.02945 0.02429 -0.0177 0.5595 0.0273
-1.250 0.1247 0.02794 0.02270 -0.0180 0.5550 0.0283
-1.000 0.1522 0.02650 0.02114 -0.0183 0.5503 0.0299
-0.750 0.1818 0.02519 0.01965 -0.0184 0.5460 0.0331
-0.500 0.2156 0.02487 0.01909 -0.0179 0.5417 0.0355
-0.250 0.2436 0.02129 0.01521 -0.0179 0.5375 0.0369
0.000 0.2700 0.02014 0.01400 -0.0182 0.5331 0.0383
0.250 0.2975 0.01929 0.01300 -0.0184 0.5291 0.0400
0.500 0.3262 0.01843 0.01205 -0.0185 0.5245 0.0435
0.750 0.3574 0.01868 0.01208 -0.0181 0.5199 0.0479
1.000 0.3840 0.01633 0.00956 -0.0184 0.5161 0.0511
1.250 0.4120 0.01576 0.00895 -0.0186 0.5119 0.0544
1.500 0.4412 0.01507 0.00806 -0.0185 0.5075 0.0636
1.750 0.4689 0.01454 0.00753 -0.0188 0.5032 0.0683
2.000 0.4972 0.01406 0.00694 -0.0189 0.4990 0.0794
2.250 0.5256 0.01364 0.00649 -0.0191 0.4946 0.0918
2.500 0.5540 0.01332 0.00613 -0.0193 0.4904 0.0989
2.750 0.5842 0.01172 0.00415 -0.0182 0.4869 0.0465
3.000 0.6118 0.01124 0.00369 -0.0181 0.4825 0.0445
3.250 0.6393 0.01095 0.00339 -0.0180 0.4780 0.0432
3.500 0.6668 0.01080 0.00321 -0.0180 0.4739 0.0427
3.750 0.6946 0.01069 0.00312 -0.0181 0.4698 0.0431
4.000 0.7226 0.01061 0.00307 -0.0183 0.4653 0.0449
4.250 0.7506 0.01060 0.00305 -0.0184 0.4610 0.0478
4.500 0.7787 0.01062 0.00309 -0.0186 0.4560 0.0559
4.750 0.8059 0.01032 0.00321 -0.0188 0.4457 0.2838
5.000 0.8770 0.00913 0.00343 -0.0288 0.4307 1.0000
5.250 0.9033 0.00919 0.00348 -0.0286 0.4188 1.0000
5.500 0.9295 0.00929 0.00355 -0.0286 0.4001 1.0000
5.750 0.9556 0.00953 0.00367 -0.0286 0.3664 1.0000
6.000 0.9803 0.01159 0.00473 -0.0303 0.1909 1.0000
6.250 1.0022 0.01424 0.00658 -0.0317 0.0223 1.0000
6.500 1.0268 0.01470 0.00714 -0.0315 0.0204 1.0000
6.750 1.0508 0.01529 0.00786 -0.0313 0.0188 1.0000
7.000 1.0738 0.01607 0.00874 -0.0311 0.0175 1.0000
7.250 1.0953 0.01710 0.00989 -0.0308 0.0163 1.0000
7.500 1.1110 0.01898 0.01195 -0.0305 0.0148 1.0000
7.750 1.1254 0.02056 0.01365 -0.0297 0.0144 1.0000
8.000 1.1413 0.02168 0.01485 -0.0289 0.0140 1.0000
8.250 1.1531 0.02311 0.01638 -0.0278 0.0137 1.0000
8.500 1.1610 0.02482 0.01818 -0.0267 0.0135 1.0000
8.750 1.1637 0.02684 0.02027 -0.0255 0.0135 1.0000
9.000 1.1644 0.02878 0.02226 -0.0233 0.0135 1.0000
9.250 1.1696 0.03060 0.02410 -0.0206 0.0139 1.0000
9.500 1.1862 0.03155 0.02504 -0.0188 0.0155 1.0000
9.750 1.2023 0.03269 0.02625 -0.0171 0.0160 1.0000
10.000 1.1179 0.02139 0.01554 -0.0086 0.0188 1.0000
10.250 1.1720 0.02400 0.01818 -0.0060 0.0204 1.0000
10.500 1.1844 0.02590 0.02020 -0.0047 0.0190 1.0000
10.750 1.1972 0.02819 0.02257 -0.0035 0.0177 1.0000
11.000 1.2208 0.03108 0.02547 -0.0028 0.0165 1.0000
11.250 1.3358 0.05296 0.04741 -0.0107 0.0144 1.0000
11.500 1.3347 0.05465 0.04933 -0.0087 0.0143 1.0000
11.750 1.3265 0.05495 0.04981 -0.0061 0.0140 1.0000
12.000 1.3180 0.05584 0.05084 -0.0042 0.0135 1.0000
12.250 1.3123 0.05793 0.05311 -0.0032 0.0131 1.0000
12.500 1.3058 0.06071 0.05608 -0.0026 0.0127 1.0000
12.750 1.2982 0.06391 0.05946 -0.0024 0.0124 1.0000
13.000 1.2890 0.06747 0.06319 -0.0024 0.0121 1.0000
13.250 1.2780 0.07137 0.06726 -0.0027 0.0119 1.0000
13.500 1.2654 0.07558 0.07162 -0.0034 0.0118 1.0000
13.750 1.2514 0.08012 0.07632 -0.0044 0.0117 1.0000
14.000 1.2363 0.08496 0.08133 -0.0057 0.0116 1.0000
14.250 1.2201 0.09015 0.08667 -0.0075 0.0115 1.0000
14.500 1.2031 0.09566 0.09233 -0.0096 0.0115 1.0000
14.750 1.1857 0.10161 0.09842 -0.0122 0.0115 1.0000
15.000 1.1682 0.10804 0.10500 -0.0153 0.0115 1.0000
15.250 1.1503 0.11500 0.11209 -0.0189 0.0116 1.0000
15.500 1.1326 0.12244 0.11967 -0.0230 0.0116 1.0000
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Polar data table (+)
Polar graphs
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