HT22 (ht22-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: HT22 (ht22-il) Reynolds number: 50,000 Max Cl/Cd: 25.34 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ht22-il-50000.txt Download as CSV file: xf-ht22-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: HT22
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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.6710 0.11894 0.11231 0.0379 1.0001 0.1743
-8.250 -0.6515 0.11303 0.10638 0.0405 1.0001 0.1846
-8.000 -0.6685 0.11204 0.10554 0.0343 1.0001 0.1890
-7.750 -0.6485 0.10630 0.09978 0.0375 1.0001 0.2011
-7.500 -0.6407 0.10206 0.09559 0.0375 1.0001 0.2097
-7.250 -0.6472 0.09936 0.09300 0.0323 1.0001 0.2190
-7.000 -0.6445 0.09628 0.08995 0.0280 1.0001 0.2324
-6.750 -0.6358 0.09216 0.08588 0.0276 1.0001 0.2470
-6.500 -0.6227 0.08756 0.08131 0.0304 1.0001 0.2631
-6.250 -0.6101 0.08362 0.07738 0.0325 1.0001 0.2820
-5.500 -0.5843 0.07357 0.06749 0.0363 1.0001 0.3741
-5.000 -0.5640 0.06723 0.06128 0.0469 1.0001 0.4677
-4.250 -0.4260 0.04321 0.03538 -0.0121 1.0001 0.1933
-4.000 -0.3832 0.03738 0.02842 -0.0159 1.0001 0.1430
-3.750 -0.3530 0.03372 0.02426 -0.0166 1.0001 0.1366
-3.500 -0.3218 0.03068 0.02053 -0.0169 1.0001 0.1340
-3.250 -0.2901 0.02823 0.01737 -0.0166 1.0001 0.1316
-3.000 -0.2623 0.02569 0.01477 -0.0163 1.0001 0.1373
-2.750 -0.2324 0.02385 0.01250 -0.0157 1.0001 0.1466
-2.500 -0.2037 0.02195 0.01052 -0.0150 1.0001 0.1585
-2.250 -0.1759 0.02026 0.00888 -0.0141 1.0001 0.1817
-2.000 -0.1476 0.01859 0.00736 -0.0133 1.0001 0.2235
-1.750 -0.1117 0.01368 0.00536 -0.0116 1.0001 0.9999
-1.500 -0.0847 0.01362 0.00469 -0.0110 1.0001 0.9999
-1.250 -0.0583 0.01356 0.00426 -0.0105 1.0001 0.9999
-1.000 -0.0320 0.01353 0.00392 -0.0101 1.0001 0.9999
-0.750 -0.0057 0.01350 0.00368 -0.0098 1.0001 0.9999
-0.500 0.0205 0.01349 0.00351 -0.0094 1.0001 0.9999
-0.250 0.0467 0.01349 0.00337 -0.0091 1.0001 0.9999
0.000 0.0728 0.01351 0.00330 -0.0087 1.0001 0.9999
0.250 0.0988 0.01353 0.00328 -0.0084 1.0001 0.9999
0.500 0.1248 0.01357 0.00332 -0.0081 1.0001 0.9999
0.750 0.1505 0.01363 0.00340 -0.0078 1.0001 0.9999
1.000 0.1761 0.01371 0.00354 -0.0075 1.0001 0.9999
1.250 0.2015 0.01381 0.00374 -0.0073 1.0001 0.9999
1.500 0.2264 0.01394 0.00403 -0.0072 1.0001 0.9999
1.750 0.2505 0.01414 0.00445 -0.0072 1.0001 0.9999
2.000 0.2719 0.01455 0.00512 -0.0079 1.0001 0.9999
2.250 0.3706 0.01530 0.00631 -0.0232 0.8911 0.9999
2.500 0.4029 0.01590 0.00669 -0.0209 0.7807 0.9999
2.750 0.4206 0.01667 0.00712 -0.0168 0.7030 0.9999
3.000 0.4403 0.01749 0.00762 -0.0135 0.6373 0.9999
3.250 0.4620 0.01836 0.00825 -0.0112 0.5742 0.9999
3.500 0.4843 0.01926 0.00891 -0.0091 0.5139 0.9999
3.750 0.5068 0.02022 0.00968 -0.0072 0.4544 0.9999
4.000 0.5300 0.02126 0.01048 -0.0056 0.3952 0.9999
4.250 0.5538 0.02238 0.01137 -0.0042 0.3378 0.9999
4.500 0.5783 0.02366 0.01247 -0.0031 0.2883 0.9999
4.750 0.6030 0.02502 0.01365 -0.0021 0.2486 0.9999
5.000 0.6288 0.02672 0.01543 -0.0015 0.2186 0.9999
5.250 0.6546 0.02853 0.01736 -0.0010 0.1955 0.9999
5.500 0.6806 0.03080 0.01981 -0.0005 0.1805 0.9999
5.750 0.7054 0.03295 0.02211 -0.0001 0.1656 0.9999
6.000 0.7305 0.03664 0.02660 -0.0004 0.1589 0.9999
6.250 0.7540 0.04015 0.03054 -0.0005 0.1537 0.9999
6.500 0.7748 0.04425 0.03512 -0.0011 0.1480 0.9999
6.750 0.7924 0.04979 0.04135 -0.0028 0.1461 0.9999
7.000 0.8077 0.05592 0.04792 -0.0049 0.1486 0.9999
7.250 0.8123 0.06638 0.05905 -0.0123 0.1612 0.9999
7.500 0.8192 0.07797 0.07090 -0.0200 0.1925 0.9999
7.750 0.7648 0.09872 0.09164 -0.0589 0.3300 0.9999
8.000 0.7838 0.10366 0.09665 -0.0578 0.3209 0.9999
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Polar data table (+)
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