HQ 1.0/8 AIRFOIL (hq108-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 1.0/8 AIRFOIL (hq108-il) Reynolds number: 200,000 Max Cl/Cd: 62.38 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq108-il-200000.txt Download as CSV file: xf-hq108-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5547 0.08797 0.08458 -0.0149 1.0000 0.0430
-8.500 -0.5610 0.08251 0.07917 -0.0212 1.0000 0.0436
-8.250 -0.5723 0.07707 0.07374 -0.0272 1.0000 0.0438
-7.250 -0.5802 0.05697 0.05323 -0.0350 1.0000 0.0464
-7.000 -0.5698 0.05368 0.04993 -0.0348 1.0000 0.0478
-6.750 -0.5595 0.05031 0.04644 -0.0349 1.0000 0.0497
-6.500 -0.5481 0.04676 0.04268 -0.0350 1.0000 0.0526
-6.250 -0.5397 0.04272 0.03798 -0.0349 1.0000 0.0587
-6.000 -0.5242 0.03916 0.03445 -0.0344 1.0000 0.0605
-5.750 -0.5033 0.03130 0.02566 -0.0313 1.0000 0.0310
-5.500 -0.4879 0.02684 0.02078 -0.0298 1.0000 0.0277
-5.250 -0.4685 0.02306 0.01642 -0.0278 1.0000 0.0254
-5.000 -0.4474 0.02050 0.01347 -0.0261 1.0000 0.0250
-4.750 -0.4265 0.01896 0.01172 -0.0247 1.0000 0.0267
-4.500 -0.4054 0.01812 0.01071 -0.0233 1.0000 0.0307
-4.250 -0.3846 0.01678 0.00920 -0.0216 1.0000 0.0319
-4.000 -0.3660 0.01488 0.00722 -0.0199 1.0000 0.0348
-3.750 -0.3470 0.01394 0.00627 -0.0183 1.0000 0.0408
-3.500 -0.3277 0.01298 0.00533 -0.0169 1.0000 0.0597
-3.250 -0.3080 0.01202 0.00454 -0.0158 1.0000 0.1010
-3.000 -0.2896 0.01047 0.00401 -0.0152 1.0000 0.2953
-2.750 -0.2749 0.00937 0.00406 -0.0133 1.0000 0.5741
-2.500 -0.2524 0.00924 0.00420 -0.0122 0.9981 0.6741
-2.250 -0.2156 0.00927 0.00423 -0.0139 0.9915 0.7297
-2.000 -0.1793 0.00930 0.00429 -0.0154 0.9846 0.7745
-1.750 -0.1443 0.00930 0.00431 -0.0165 0.9777 0.8106
-1.500 -0.1093 0.00927 0.00431 -0.0176 0.9708 0.8467
-1.250 -0.0769 0.00922 0.00428 -0.0179 0.9637 0.8796
-1.000 -0.0386 0.00918 0.00423 -0.0196 0.9577 0.9044
-0.750 0.0059 0.00916 0.00419 -0.0226 0.9538 0.9297
-0.500 0.0522 0.00915 0.00414 -0.0261 0.9481 0.9507
-0.250 0.1034 0.00909 0.00403 -0.0308 0.9440 0.9622
0.000 0.1530 0.00900 0.00391 -0.0353 0.9382 0.9720
0.250 0.2018 0.00883 0.00373 -0.0394 0.9292 0.9806
0.500 0.2489 0.00864 0.00353 -0.0432 0.9171 0.9889
0.750 0.2933 0.00847 0.00334 -0.0465 0.9035 0.9998
1.000 0.3121 0.00839 0.00322 -0.0448 0.8868 1.0000
1.250 0.3293 0.00840 0.00320 -0.0428 0.8702 1.0000
1.500 0.3492 0.00844 0.00323 -0.0412 0.8550 1.0000
1.750 0.3708 0.00849 0.00325 -0.0398 0.8398 1.0000
2.000 0.3935 0.00855 0.00329 -0.0385 0.8242 1.0000
2.250 0.4169 0.00860 0.00333 -0.0373 0.8077 1.0000
2.500 0.4405 0.00865 0.00336 -0.0361 0.7891 1.0000
2.750 0.4643 0.00869 0.00343 -0.0349 0.7671 1.0000
3.000 0.4881 0.00873 0.00346 -0.0337 0.7414 1.0000
3.250 0.5120 0.00880 0.00348 -0.0324 0.7108 1.0000
3.500 0.5360 0.00890 0.00355 -0.0313 0.6750 1.0000
3.750 0.5598 0.00907 0.00364 -0.0301 0.6314 1.0000
4.000 0.5826 0.00934 0.00378 -0.0288 0.5683 1.0000
4.250 0.6036 0.00987 0.00396 -0.0273 0.4770 1.0000
4.500 0.6235 0.01066 0.00430 -0.0260 0.3750 1.0000
4.750 0.6439 0.01156 0.00477 -0.0250 0.2841 1.0000
5.000 0.6650 0.01247 0.00534 -0.0242 0.2046 1.0000
5.250 0.6848 0.01369 0.00604 -0.0233 0.1147 1.0000
5.500 0.7049 0.01504 0.00717 -0.0222 0.0665 1.0000
5.750 0.7255 0.01634 0.00837 -0.0211 0.0402 1.0000
6.000 0.7458 0.01788 0.00998 -0.0197 0.0308 1.0000
6.250 0.7640 0.02055 0.01266 -0.0181 0.0264 1.0000
6.500 0.7886 0.02145 0.01374 -0.0172 0.0230 1.0000
6.750 0.8117 0.02297 0.01542 -0.0164 0.0204 1.0000
7.000 0.8341 0.02511 0.01778 -0.0153 0.0192 1.0000
7.250 0.8555 0.02779 0.02079 -0.0141 0.0189 1.0000
7.500 0.8746 0.03131 0.02481 -0.0125 0.0195 1.0000
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Polar data table (+)
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