HQ 1.0/8 AIRFOIL (hq108-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 1.0/8 AIRFOIL (hq108-il) Reynolds number: 1,000,000 Max Cl/Cd: 85.04 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq108-il-1000000.txt Download as CSV file: xf-hq108-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.250 -0.6035 0.15564 0.15391 0.0192 1.0000 0.0048
-13.000 -0.5982 0.15178 0.15005 0.0175 1.0000 0.0048
-5.750 -0.5139 0.01651 0.01214 -0.0272 0.9984 0.0077
-5.500 -0.4805 0.01493 0.01040 -0.0285 0.9956 0.0069
-5.250 -0.4493 0.01251 0.00769 -0.0293 0.9921 0.0067
-5.000 -0.4175 0.01105 0.00606 -0.0302 0.9878 0.0067
-4.750 -0.3837 0.01011 0.00500 -0.0316 0.9843 0.0070
-4.500 -0.3517 0.00884 0.00352 -0.0326 0.9788 0.0087
-4.250 -0.3192 0.00839 0.00302 -0.0337 0.9716 0.0104
-4.000 -0.2878 0.00806 0.00261 -0.0344 0.9626 0.0122
-3.750 -0.2589 0.00758 0.00215 -0.0346 0.9513 0.0272
-3.500 -0.2318 0.00730 0.00193 -0.0344 0.9379 0.0495
-3.250 -0.2056 0.00709 0.00174 -0.0340 0.9240 0.0701
-3.000 -0.1800 0.00677 0.00155 -0.0336 0.9105 0.1145
-2.750 -0.1552 0.00627 0.00134 -0.0332 0.8970 0.2116
-2.500 -0.1308 0.00568 0.00115 -0.0328 0.8839 0.3414
-2.250 -0.1057 0.00522 0.00102 -0.0325 0.8719 0.4562
-2.000 -0.0800 0.00492 0.00095 -0.0321 0.8610 0.5432
-1.750 -0.0532 0.00479 0.00089 -0.0319 0.8502 0.5902
-1.500 -0.0263 0.00472 0.00085 -0.0316 0.8379 0.6213
-1.250 0.0006 0.00467 0.00082 -0.0313 0.8251 0.6543
-1.000 0.0277 0.00463 0.00080 -0.0311 0.8130 0.6788
-0.750 0.0548 0.00459 0.00079 -0.0308 0.8010 0.7031
-0.500 0.0818 0.00457 0.00078 -0.0305 0.7888 0.7272
-0.250 0.1092 0.00457 0.00077 -0.0303 0.7760 0.7426
0.000 0.1366 0.00457 0.00076 -0.0301 0.7630 0.7559
0.250 0.1639 0.00458 0.00077 -0.0300 0.7512 0.7713
0.500 0.1912 0.00457 0.00078 -0.0298 0.7391 0.7877
0.750 0.2185 0.00458 0.00080 -0.0295 0.7261 0.8041
1.500 0.2995 0.00463 0.00086 -0.0288 0.6773 0.8462
1.750 0.3259 0.00467 0.00090 -0.0284 0.6557 0.8622
2.000 0.3517 0.00471 0.00094 -0.0279 0.6287 0.8824
2.250 0.3763 0.00479 0.00099 -0.0270 0.5886 0.9098
2.500 0.4040 0.00489 0.00106 -0.0269 0.5495 0.9527
2.750 0.4422 0.00520 0.00117 -0.0294 0.4829 0.9854
3.000 0.4732 0.00562 0.00132 -0.0304 0.4086 1.0000
3.250 0.4977 0.00605 0.00148 -0.0300 0.3395 1.0000
3.500 0.5226 0.00645 0.00166 -0.0296 0.2829 1.0000
3.750 0.5476 0.00686 0.00186 -0.0292 0.2333 1.0000
4.000 0.5732 0.00718 0.00207 -0.0290 0.1961 1.0000
4.250 0.5984 0.00758 0.00228 -0.0287 0.1502 1.0000
4.500 0.6228 0.00812 0.00257 -0.0283 0.0972 1.0000
4.750 0.6480 0.00855 0.00285 -0.0280 0.0663 1.0000
5.000 0.6734 0.00894 0.00315 -0.0276 0.0439 1.0000
5.250 0.6987 0.00936 0.00346 -0.0273 0.0257 1.0000
5.500 0.7230 0.01001 0.00407 -0.0267 0.0094 1.0000
5.750 0.7486 0.01039 0.00449 -0.0263 0.0075 1.0000
6.000 0.7711 0.01144 0.00570 -0.0253 0.0053 1.0000
6.250 0.7955 0.01208 0.00642 -0.0247 0.0050 1.0000
6.500 0.8187 0.01292 0.00738 -0.0238 0.0048 1.0000
6.750 0.8409 0.01396 0.00855 -0.0229 0.0048 1.0000
7.000 0.8620 0.01529 0.01004 -0.0217 0.0048 1.0000
7.250 0.8822 0.01698 0.01192 -0.0205 0.0048 1.0000
7.500 0.9010 0.01930 0.01451 -0.0190 0.0049 1.0000
7.750 0.9197 0.02158 0.01705 -0.0177 0.0050 1.0000
8.000 0.9407 0.02288 0.01852 -0.0168 0.0048 1.0000
8.250 0.9592 0.02472 0.02059 -0.0156 0.0047 1.0000
8.500 0.9712 0.02792 0.02414 -0.0138 0.0047 1.0000
8.750 0.9761 0.03209 0.02873 -0.0114 0.0049 1.0000
9.000 0.9587 0.04223 0.03965 -0.0064 0.0073 1.0000
9.250 0.9605 0.04512 0.04275 -0.0047 0.0069 1.0000
9.500 0.9484 0.04978 0.04766 -0.0023 0.0069 1.0000
9.750 0.9381 0.05220 0.05020 0.0003 0.0067 1.0000
10.000 0.9242 0.05463 0.05273 0.0021 0.0065 1.0000
10.250 0.8936 0.06059 0.05887 0.0014 0.0067 1.0000
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