HQ 1.0/10 AIRFOIL (hq1010-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.0/10 AIRFOIL (hq1010-il) Reynolds number: 50,000 Max Cl/Cd: 32.91 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq1010-il-50000-n5.txt Download as CSV file: xf-hq1010-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.0/10 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5697 0.09381 0.08691 -0.0209 1.0000 0.0431
-9.750 -0.5764 0.08739 0.08055 -0.0250 1.0000 0.0424
-9.500 -0.5884 0.08006 0.07327 -0.0304 1.0000 0.0417
-9.250 -0.6055 0.07371 0.06694 -0.0350 1.0000 0.0410
-9.000 -0.6258 0.06856 0.06174 -0.0374 1.0000 0.0405
-8.750 -0.6408 0.06382 0.05686 -0.0385 1.0000 0.0400
-8.500 -0.6500 0.05933 0.05215 -0.0388 1.0000 0.0400
-8.250 -0.6537 0.05512 0.04765 -0.0386 1.0000 0.0402
-8.000 -0.6522 0.05129 0.04350 -0.0380 1.0000 0.0411
-7.750 -0.6468 0.04782 0.03967 -0.0371 1.0000 0.0429
-7.500 -0.6389 0.04435 0.03570 -0.0361 1.0000 0.0454
-7.250 -0.6280 0.04094 0.03160 -0.0347 1.0000 0.0480
-7.000 -0.6130 0.03785 0.02828 -0.0336 1.0000 0.0502
-6.750 -0.5956 0.03545 0.02565 -0.0324 1.0000 0.0532
-6.500 -0.5771 0.03340 0.02319 -0.0311 1.0000 0.0596
-6.250 -0.5581 0.03138 0.02106 -0.0299 1.0000 0.0657
-6.000 -0.5370 0.02953 0.01891 -0.0284 1.0000 0.0731
-5.750 -0.5182 0.02810 0.01741 -0.0271 1.0000 0.0840
-5.500 -0.4987 0.02666 0.01592 -0.0255 1.0000 0.0943
-5.250 -0.4805 0.02548 0.01469 -0.0240 1.0000 0.1095
-5.000 -0.4627 0.02426 0.01345 -0.0222 1.0000 0.1230
-4.750 -0.4459 0.02317 0.01240 -0.0206 1.0000 0.1444
-4.500 -0.4299 0.02204 0.01145 -0.0191 1.0000 0.1763
-4.250 -0.4138 0.02078 0.01050 -0.0178 1.0000 0.2325
-4.000 -0.4016 0.01942 0.00978 -0.0159 1.0000 0.3365
-3.750 -0.3904 0.01862 0.00972 -0.0128 1.0000 0.4807
-3.500 -0.3778 0.01847 0.00985 -0.0092 1.0000 0.5832
-3.250 -0.3647 0.01851 0.00993 -0.0057 1.0000 0.6519
-3.000 -0.3521 0.01860 0.01001 -0.0020 1.0000 0.7058
-2.750 -0.3399 0.01870 0.01008 0.0018 1.0000 0.7505
-2.500 -0.3268 0.01874 0.01005 0.0052 1.0000 0.7891
-2.250 -0.3113 0.01873 0.00993 0.0081 1.0000 0.8210
-2.000 -0.2834 0.01878 0.00983 0.0083 0.9952 0.8510
-1.750 -0.2448 0.01890 0.00980 0.0068 0.9880 0.8842
-1.500 -0.1955 0.01910 0.00978 0.0029 0.9830 0.9160
-1.250 -0.1420 0.01921 0.00969 -0.0024 0.9774 0.9366
-1.000 -0.0875 0.01931 0.00960 -0.0084 0.9719 0.9527
-0.750 -0.0350 0.01935 0.00945 -0.0141 0.9650 0.9676
-0.500 0.0204 0.01940 0.00938 -0.0204 0.9589 0.9817
-0.250 0.0717 0.01941 0.00929 -0.0261 0.9512 0.9982
0.000 0.1079 0.01946 0.00927 -0.0290 0.9407 1.0000
0.250 0.1437 0.01951 0.00928 -0.0317 0.9300 1.0000
0.500 0.1752 0.01957 0.00932 -0.0333 0.9181 1.0000
0.750 0.2064 0.01965 0.00938 -0.0348 0.9063 1.0000
1.000 0.2367 0.01976 0.00947 -0.0359 0.8945 1.0000
1.250 0.2672 0.01989 0.00960 -0.0369 0.8830 1.0000
1.500 0.2988 0.02001 0.00976 -0.0380 0.8715 1.0000
1.750 0.3292 0.02009 0.00987 -0.0385 0.8574 1.0000
2.000 0.3579 0.02010 0.00991 -0.0383 0.8395 1.0000
2.250 0.3880 0.02002 0.00991 -0.0381 0.8210 1.0000
2.500 0.4178 0.01995 0.00989 -0.0377 0.8035 1.0000
2.750 0.4409 0.02005 0.01006 -0.0365 0.7855 1.0000
3.000 0.4657 0.02012 0.01021 -0.0356 0.7679 1.0000
3.250 0.4916 0.02015 0.01037 -0.0346 0.7501 1.0000
3.500 0.5165 0.02017 0.01050 -0.0334 0.7308 1.0000
3.750 0.5403 0.02021 0.01065 -0.0320 0.7089 1.0000
4.000 0.5641 0.02022 0.01077 -0.0306 0.6848 1.0000
4.250 0.5877 0.02020 0.01086 -0.0290 0.6574 1.0000
4.500 0.6108 0.02019 0.01098 -0.0272 0.6247 1.0000
4.750 0.6331 0.02022 0.01107 -0.0253 0.5851 1.0000
5.000 0.6550 0.02031 0.01113 -0.0233 0.5383 1.0000
5.250 0.6757 0.02059 0.01128 -0.0213 0.4835 1.0000
5.500 0.6950 0.02112 0.01157 -0.0193 0.4259 1.0000
5.750 0.7129 0.02190 0.01216 -0.0175 0.3699 1.0000
6.000 0.7304 0.02284 0.01289 -0.0159 0.3208 1.0000
6.250 0.7476 0.02392 0.01378 -0.0145 0.2781 1.0000
6.500 0.7647 0.02508 0.01482 -0.0132 0.2397 1.0000
6.750 0.7819 0.02628 0.01596 -0.0119 0.2035 1.0000
7.000 0.7989 0.02753 0.01712 -0.0107 0.1715 1.0000
7.250 0.8165 0.02890 0.01850 -0.0095 0.1467 1.0000
7.500 0.8344 0.03032 0.01993 -0.0083 0.1237 1.0000
7.750 0.8524 0.03192 0.02151 -0.0072 0.1072 1.0000
8.000 0.8698 0.03358 0.02314 -0.0062 0.0922 1.0000
8.250 0.8891 0.03545 0.02516 -0.0051 0.0804 1.0000
8.500 0.9070 0.03739 0.02733 -0.0040 0.0695 1.0000
8.750 0.9255 0.03982 0.03000 -0.0030 0.0618 1.0000
9.000 0.9419 0.04226 0.03259 -0.0019 0.0559 1.0000
9.250 0.9539 0.04509 0.03579 -0.0006 0.0503 1.0000
9.500 0.9634 0.04792 0.03897 0.0008 0.0459 1.0000
9.750 0.9717 0.05056 0.04178 0.0019 0.0430 1.0000
10.000 0.9715 0.05412 0.04575 0.0036 0.0409 1.0000
10.250 0.9642 0.05775 0.04983 0.0055 0.0393 1.0000
10.500 0.9519 0.06134 0.05374 0.0072 0.0382 1.0000
10.750 0.9370 0.06534 0.05801 0.0079 0.0376 1.0000
11.000 0.9192 0.06997 0.06287 0.0074 0.0375 1.0000
11.250 0.8989 0.07542 0.06852 0.0054 0.0378 1.0000
11.500 0.8767 0.08193 0.07519 0.0020 0.0384 1.0000
11.750 0.8547 0.08954 0.08290 -0.0028 0.0394 1.0000
12.000 0.8346 0.09792 0.09127 -0.0082 0.0403 1.0000
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Polar data table (+)
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