HQ 1.0/8 AIRFOIL (hq108-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 1.0/8 AIRFOIL (hq108-il) Reynolds number: 500,000 Max Cl/Cd: 78.58 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq108-il-500000.txt Download as CSV file: xf-hq108-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.0/8 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5722 0.08099 0.07886 -0.0167 1.0000 0.0138
-8.750 -0.5731 0.07740 0.07529 -0.0187 1.0000 0.0142
-8.500 -0.5862 0.06801 0.06593 -0.0284 1.0000 0.0137
-8.250 -0.5975 0.06270 0.06054 -0.0317 1.0000 0.0137
-8.000 -0.6007 0.05858 0.05635 -0.0333 1.0000 0.0141
-7.750 -0.5968 0.05599 0.05370 -0.0339 1.0000 0.0147
-7.500 -0.5936 0.05163 0.04920 -0.0349 1.0000 0.0150
-7.250 -0.5857 0.04827 0.04571 -0.0353 1.0000 0.0158
-7.000 -0.5756 0.04478 0.04201 -0.0354 1.0000 0.0173
-6.750 -0.5484 0.04439 0.04134 -0.0344 1.0000 0.0212
-6.500 -0.5366 0.04117 0.03784 -0.0336 1.0000 0.0213
-5.500 -0.4918 0.02110 0.01611 -0.0266 1.0000 0.0131
-5.250 -0.4763 0.01762 0.01230 -0.0244 1.0000 0.0118
-5.000 -0.4592 0.01555 0.00997 -0.0223 1.0000 0.0115
-4.750 -0.4289 0.01375 0.00795 -0.0228 0.9979 0.0117
-4.500 -0.3932 0.01262 0.00668 -0.0244 0.9946 0.0126
-4.250 -0.3607 0.01090 0.00477 -0.0256 0.9903 0.0147
-4.000 -0.3248 0.01037 0.00420 -0.0275 0.9861 0.0189
-3.750 -0.2878 0.00955 0.00330 -0.0294 0.9828 0.0287
-3.500 -0.2530 0.00894 0.00282 -0.0311 0.9779 0.0660
-3.250 -0.2184 0.00837 0.00250 -0.0329 0.9727 0.1274
-3.000 -0.1853 0.00737 0.00217 -0.0349 0.9676 0.2980
-2.750 -0.1573 0.00648 0.00194 -0.0354 0.9583 0.4881
-2.500 -0.1282 0.00612 0.00187 -0.0357 0.9491 0.5860
-2.250 -0.0993 0.00595 0.00180 -0.0358 0.9395 0.6343
-2.000 -0.0720 0.00586 0.00172 -0.0354 0.9282 0.6649
-1.750 -0.0459 0.00580 0.00166 -0.0347 0.9160 0.6957
-1.500 -0.0201 0.00575 0.00163 -0.0340 0.9041 0.7214
-1.250 0.0058 0.00571 0.00161 -0.0333 0.8929 0.7445
-1.000 0.0314 0.00569 0.00160 -0.0326 0.8815 0.7686
-0.750 0.0572 0.00568 0.00157 -0.0319 0.8688 0.7853
-0.500 0.0831 0.00566 0.00153 -0.0312 0.8557 0.8004
-0.250 0.1089 0.00565 0.00151 -0.0305 0.8432 0.8177
0.000 0.1345 0.00563 0.00149 -0.0298 0.8309 0.8355
0.250 0.1600 0.00559 0.00147 -0.0290 0.8180 0.8529
0.750 0.2117 0.00553 0.00145 -0.0278 0.7927 0.8862
1.000 0.2375 0.00549 0.00145 -0.0271 0.7805 0.9059
1.250 0.2650 0.00547 0.00146 -0.0267 0.7679 0.9302
1.500 0.2986 0.00547 0.00147 -0.0278 0.7541 0.9565
1.750 0.3373 0.00551 0.00149 -0.0301 0.7380 0.9779
2.000 0.3768 0.00555 0.00150 -0.0327 0.7169 0.9968
2.250 0.4039 0.00564 0.00152 -0.0326 0.6946 1.0000
2.500 0.4285 0.00577 0.00155 -0.0319 0.6649 1.0000
2.750 0.4535 0.00592 0.00163 -0.0313 0.6318 1.0000
3.000 0.4786 0.00611 0.00171 -0.0308 0.5942 1.0000
3.250 0.5029 0.00640 0.00181 -0.0301 0.5383 1.0000
3.500 0.5263 0.00685 0.00197 -0.0293 0.4631 1.0000
3.750 0.5496 0.00739 0.00220 -0.0287 0.3841 1.0000
4.000 0.5731 0.00798 0.00249 -0.0282 0.3054 1.0000
4.250 0.5969 0.00855 0.00280 -0.0277 0.2448 1.0000
4.500 0.6216 0.00902 0.00309 -0.0273 0.1960 1.0000
4.750 0.6456 0.00960 0.00342 -0.0269 0.1373 1.0000
5.000 0.6689 0.01031 0.00384 -0.0264 0.0809 1.0000
5.250 0.6930 0.01091 0.00429 -0.0259 0.0511 1.0000
5.500 0.7166 0.01166 0.00496 -0.0253 0.0250 1.0000
5.750 0.7400 0.01249 0.00581 -0.0245 0.0142 1.0000
6.000 0.7634 0.01329 0.00668 -0.0237 0.0111 1.0000
6.250 0.7821 0.01504 0.00863 -0.0221 0.0094 1.0000
6.500 0.8049 0.01605 0.00976 -0.0211 0.0088 1.0000
6.750 0.8264 0.01746 0.01133 -0.0200 0.0085 1.0000
7.000 0.8474 0.01928 0.01335 -0.0188 0.0084 1.0000
7.250 0.8679 0.02159 0.01591 -0.0175 0.0085 1.0000
7.500 0.8863 0.02457 0.01924 -0.0160 0.0088 1.0000
7.750 0.8986 0.02955 0.02468 -0.0141 0.0098 1.0000
12.500 0.6523 0.14015 0.13819 -0.0421 0.0166 1.0000
12.750 0.6490 0.14404 0.14209 -0.0443 0.0166 1.0000
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