HQ 1.0/9 AIRFOIL (hq109-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.0/9 AIRFOIL (hq109-il) Reynolds number: 500,000 Max Cl/Cd: 66.11 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq109-il-500000-n5.txt Download as CSV file: xf-hq109-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.0/9 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.5910 0.10026 0.09798 -0.0069 1.0000 0.0047
-10.500 -0.5974 0.09413 0.09190 -0.0099 1.0000 0.0046
-10.000 -0.7846 0.04367 0.04114 -0.0389 1.0000 0.0033
-9.750 -0.7892 0.04031 0.03758 -0.0386 1.0000 0.0034
-9.500 -0.7951 0.03587 0.03280 -0.0378 1.0000 0.0035
-9.250 -0.7921 0.03232 0.02892 -0.0368 1.0000 0.0035
-9.000 -0.7756 0.03117 0.02766 -0.0363 1.0000 0.0037
-8.750 -0.7596 0.02971 0.02603 -0.0357 1.0000 0.0039
-8.500 -0.7467 0.02722 0.02323 -0.0346 1.0000 0.0041
-8.250 -0.7326 0.02472 0.02038 -0.0335 1.0000 0.0043
-8.000 -0.7168 0.02243 0.01776 -0.0322 1.0000 0.0044
-7.750 -0.6995 0.02041 0.01543 -0.0309 1.0000 0.0045
-7.500 -0.6810 0.01878 0.01355 -0.0297 1.0000 0.0047
-7.250 -0.6617 0.01745 0.01201 -0.0284 1.0000 0.0049
-7.000 -0.6419 0.01632 0.01071 -0.0271 1.0000 0.0052
-6.750 -0.6127 0.01529 0.00949 -0.0278 0.9953 0.0055
-6.500 -0.5822 0.01421 0.00824 -0.0288 0.9892 0.0059
-6.250 -0.5508 0.01314 0.00705 -0.0300 0.9833 0.0068
-6.000 -0.5193 0.01253 0.00636 -0.0311 0.9760 0.0076
-5.750 -0.4871 0.01203 0.00580 -0.0323 0.9686 0.0089
-5.500 -0.4551 0.01158 0.00528 -0.0333 0.9599 0.0100
-5.250 -0.4250 0.01093 0.00453 -0.0340 0.9491 0.0124
-5.000 -0.3952 0.01054 0.00406 -0.0345 0.9374 0.0152
-4.750 -0.3668 0.01015 0.00368 -0.0346 0.9248 0.0228
-4.500 -0.3394 0.00986 0.00336 -0.0346 0.9126 0.0324
-4.250 -0.3125 0.00963 0.00309 -0.0344 0.9012 0.0403
-3.750 -0.2599 0.00914 0.00259 -0.0338 0.8789 0.0691
-3.500 -0.2339 0.00883 0.00236 -0.0335 0.8685 0.0998
-3.250 -0.2077 0.00854 0.00215 -0.0333 0.8589 0.1354
-3.000 -0.1815 0.00827 0.00197 -0.0331 0.8493 0.1768
-2.750 -0.1550 0.00803 0.00181 -0.0329 0.8395 0.2151
-2.500 -0.1290 0.00769 0.00161 -0.0327 0.8295 0.2734
-2.250 -0.1045 0.00710 0.00144 -0.0324 0.8192 0.4010
-2.000 -0.0790 0.00675 0.00134 -0.0320 0.8084 0.4876
-1.750 -0.0523 0.00659 0.00128 -0.0318 0.7983 0.5334
-1.500 -0.0260 0.00642 0.00124 -0.0314 0.7896 0.5900
-1.250 0.0008 0.00633 0.00123 -0.0311 0.7813 0.6285
-1.000 0.0280 0.00629 0.00121 -0.0309 0.7706 0.6527
-0.750 0.0550 0.00626 0.00118 -0.0306 0.7574 0.6723
-0.500 0.0821 0.00626 0.00116 -0.0303 0.7441 0.6887
-0.250 0.1094 0.00626 0.00115 -0.0301 0.7324 0.7029
0.000 0.1367 0.00626 0.00114 -0.0299 0.7187 0.7166
0.250 0.1638 0.00628 0.00114 -0.0297 0.7037 0.7307
0.500 0.1910 0.00629 0.00116 -0.0294 0.6909 0.7449
0.750 0.2182 0.00631 0.00118 -0.0292 0.6776 0.7572
1.000 0.2455 0.00635 0.00119 -0.0290 0.6617 0.7674
1.250 0.2726 0.00640 0.00122 -0.0288 0.6452 0.7776
1.750 0.3265 0.00651 0.00132 -0.0283 0.6076 0.7988
2.000 0.3529 0.00661 0.00137 -0.0280 0.5806 0.8104
2.250 0.3788 0.00675 0.00143 -0.0275 0.5455 0.8227
2.500 0.4043 0.00693 0.00154 -0.0270 0.5053 0.8359
2.750 0.4289 0.00720 0.00166 -0.0264 0.4523 0.8500
3.000 0.4526 0.00755 0.00182 -0.0257 0.3893 0.8661
3.250 0.4765 0.00787 0.00199 -0.0250 0.3315 0.8863
3.750 0.5326 0.00845 0.00242 -0.0253 0.2528 0.9482
4.000 0.5697 0.00876 0.00264 -0.0276 0.2202 0.9861
4.250 0.5983 0.00905 0.00286 -0.0280 0.1968 1.0000
4.500 0.6231 0.00944 0.00310 -0.0277 0.1606 1.0000
4.750 0.6475 0.00989 0.00340 -0.0273 0.1220 1.0000
5.000 0.6724 0.01029 0.00370 -0.0269 0.0968 1.0000
5.250 0.6970 0.01073 0.00404 -0.0265 0.0729 1.0000
5.500 0.7215 0.01120 0.00440 -0.0261 0.0525 1.0000
5.750 0.7463 0.01161 0.00476 -0.0257 0.0390 1.0000
6.000 0.7712 0.01201 0.00514 -0.0254 0.0297 1.0000
6.250 0.7958 0.01246 0.00559 -0.0249 0.0209 1.0000
6.500 0.8201 0.01294 0.00605 -0.0245 0.0144 1.0000
6.750 0.8441 0.01346 0.00657 -0.0240 0.0092 1.0000
7.000 0.8678 0.01403 0.00718 -0.0234 0.0067 1.0000
7.250 0.8915 0.01459 0.00781 -0.0228 0.0055 1.0000
7.500 0.9143 0.01530 0.00858 -0.0221 0.0047 1.0000
7.750 0.9363 0.01611 0.00951 -0.0213 0.0042 1.0000
8.000 0.9583 0.01687 0.01039 -0.0205 0.0040 1.0000
8.250 0.9795 0.01774 0.01141 -0.0196 0.0038 1.0000
8.500 0.9999 0.01869 0.01250 -0.0186 0.0037 1.0000
8.750 1.0193 0.01976 0.01370 -0.0175 0.0035 1.0000
9.000 1.0376 0.02095 0.01505 -0.0163 0.0034 1.0000
9.250 1.0551 0.02223 0.01650 -0.0151 0.0033 1.0000
9.500 1.0712 0.02367 0.01812 -0.0137 0.0033 1.0000
9.750 1.0866 0.02512 0.01976 -0.0123 0.0032 1.0000
10.000 1.1007 0.02663 0.02145 -0.0108 0.0031 1.0000
10.250 1.1142 0.02802 0.02301 -0.0093 0.0029 1.0000
10.500 1.1265 0.02936 0.02450 -0.0078 0.0028 1.0000
10.750 1.1376 0.03061 0.02587 -0.0063 0.0026 1.0000
11.000 1.1407 0.03232 0.02780 -0.0037 0.0026 1.0000
11.250 1.1380 0.03429 0.02996 -0.0008 0.0025 1.0000
11.500 1.1318 0.03670 0.03258 0.0017 0.0025 1.0000
11.750 1.1236 0.03944 0.03553 0.0035 0.0024 1.0000
12.000 1.1103 0.04303 0.03935 0.0045 0.0023 1.0000
12.250 1.0987 0.04685 0.04337 0.0044 0.0024 1.0000
12.500 1.0867 0.05117 0.04788 0.0033 0.0024 1.0000
12.750 1.0682 0.05704 0.05396 0.0006 0.0023 1.0000
13.000 1.0533 0.06324 0.06033 -0.0030 0.0024 1.0000
13.250 1.0396 0.07013 0.06737 -0.0076 0.0024 1.0000
13.500 1.0215 0.07894 0.07635 -0.0137 0.0024 1.0000
13.750 1.0041 0.08840 0.08595 -0.0198 0.0024 1.0000
14.000 0.9678 0.10302 0.10078 -0.0282 0.0026 1.0000
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Polar data table (+)
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