MH 22 7.2% (mh22-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: MH 22 7.2% (mh22-il) Reynolds number: 500,000 Max Cl/Cd: 82.76 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-mh22-il-500000.txt Download as CSV file: xf-mh22-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: MH 22 7.2%
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 -0.5151 0.08344 0.08138 -0.0010 1.0000 0.0106
-7.500 -0.5171 0.07975 0.07772 -0.0031 1.0000 0.0106
-7.250 -0.5190 0.07576 0.07377 -0.0061 1.0000 0.0108
-7.000 -0.5150 0.07105 0.06905 -0.0106 1.0000 0.0112
-6.750 -0.5084 0.06593 0.06391 -0.0150 1.0000 0.0113
-6.500 -0.4989 0.06071 0.05863 -0.0191 1.0000 0.0118
-6.250 -0.4871 0.05529 0.05313 -0.0225 1.0000 0.0123
-6.000 -0.4728 0.05004 0.04774 -0.0250 1.0000 0.0129
-5.750 -0.4564 0.04477 0.04229 -0.0267 1.0000 0.0138
-5.500 -0.4351 0.04026 0.03754 -0.0272 1.0000 0.0154
-5.250 -0.4106 0.03769 0.03470 -0.0268 1.0000 0.0164
-3.500 -0.2350 0.01140 0.00549 -0.0249 0.9031 0.0073
-3.250 -0.2129 0.01032 0.00415 -0.0232 0.8696 0.0088
-3.000 -0.1885 0.00999 0.00369 -0.0223 0.8434 0.0123
-2.750 -0.1633 0.00949 0.00296 -0.0213 0.8226 0.0134
-2.500 -0.1371 0.00906 0.00232 -0.0204 0.8048 0.0200
-2.250 -0.1132 0.00821 0.00199 -0.0199 0.7896 0.1548
-2.000 -0.0884 0.00772 0.00184 -0.0195 0.7762 0.2648
-1.750 -0.0639 0.00720 0.00167 -0.0190 0.7642 0.3923
-1.500 -0.0414 0.00653 0.00156 -0.0181 0.7535 0.5746
-1.250 -0.0228 0.00578 0.00151 -0.0159 0.7437 0.7831
-1.000 0.0278 0.00548 0.00154 -0.0201 0.7339 0.9515
-0.750 0.0787 0.00557 0.00150 -0.0249 0.7242 0.9871
-0.500 0.1284 0.00562 0.00137 -0.0297 0.7149 1.0000
-0.250 0.1536 0.00565 0.00131 -0.0291 0.7054 1.0000
0.000 0.1791 0.00568 0.00126 -0.0286 0.6961 1.0000
0.250 0.2045 0.00573 0.00123 -0.0281 0.6873 1.0000
0.500 0.2301 0.00578 0.00121 -0.0275 0.6781 1.0000
0.750 0.2559 0.00582 0.00121 -0.0271 0.6686 1.0000
1.000 0.2816 0.00587 0.00122 -0.0266 0.6592 1.0000
1.250 0.3074 0.00594 0.00122 -0.0261 0.6497 1.0000
1.500 0.3333 0.00599 0.00124 -0.0256 0.6392 1.0000
1.750 0.3593 0.00604 0.00127 -0.0252 0.6280 1.0000
2.000 0.3853 0.00610 0.00131 -0.0248 0.6162 1.0000
2.250 0.4113 0.00616 0.00137 -0.0243 0.6033 1.0000
2.500 0.4373 0.00623 0.00141 -0.0239 0.5889 1.0000
2.750 0.4632 0.00631 0.00146 -0.0234 0.5726 1.0000
3.000 0.4892 0.00639 0.00152 -0.0230 0.5532 1.0000
3.250 0.5150 0.00650 0.00159 -0.0226 0.5308 1.0000
3.500 0.5406 0.00665 0.00174 -0.0221 0.5035 1.0000
3.750 0.5661 0.00684 0.00186 -0.0216 0.4743 1.0000
4.000 0.5898 0.00728 0.00201 -0.0210 0.4082 1.0000
4.250 0.6126 0.00793 0.00224 -0.0203 0.3186 1.0000
4.500 0.6362 0.00849 0.00252 -0.0198 0.2493 1.0000
4.750 0.6552 0.00991 0.00313 -0.0190 0.0934 1.0000
5.000 0.6757 0.01124 0.00402 -0.0180 0.0055 1.0000
5.250 0.7005 0.01175 0.00470 -0.0172 0.0046 1.0000
5.500 0.7244 0.01244 0.00556 -0.0164 0.0044 1.0000
5.750 0.7469 0.01341 0.00678 -0.0153 0.0043 1.0000
6.000 0.7678 0.01466 0.00820 -0.0140 0.0044 1.0000
6.250 0.7875 0.01618 0.00989 -0.0125 0.0046 1.0000
6.500 0.8066 0.01813 0.01201 -0.0108 0.0049 1.0000
6.750 0.8257 0.02064 0.01472 -0.0092 0.0053 1.0000
7.000 0.8466 0.02311 0.01734 -0.0078 0.0065 1.0000
8.000 0.8927 0.04393 0.03982 -0.0004 0.0139 1.0000
8.250 0.8984 0.04742 0.04360 0.0010 0.0139 1.0000
8.500 0.9005 0.05088 0.04732 0.0025 0.0139 1.0000
8.750 0.8987 0.05425 0.05094 0.0038 0.0138 1.0000
9.000 0.8928 0.05752 0.05443 0.0051 0.0138 1.0000
9.250 0.8814 0.06039 0.05747 0.0068 0.0138 1.0000
9.500 0.8640 0.06354 0.06075 0.0077 0.0137 1.0000
9.750 0.8474 0.06752 0.06485 0.0063 0.0138 1.0000
10.000 0.8319 0.07215 0.06958 0.0034 0.0138 1.0000
10.250 0.8166 0.07785 0.07537 -0.0007 0.0138 1.0000
10.500 0.8015 0.08464 0.08224 -0.0058 0.0138 1.0000
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Polar data table (+)
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