HQ 1.5/8 AIRFOIL (hq158-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 1.5/8 AIRFOIL (hq158-il) Reynolds number: 1,000,000 Max Cl/Cd: 91.97 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq158-il-1000000.txt Download as CSV file: xf-hq158-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/8 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4277 0.08266 0.08117 -0.0178 1.0000 0.0080
-9.000 -0.4295 0.07786 0.07638 -0.0195 1.0000 0.0081
-8.750 -0.4326 0.07293 0.07146 -0.0214 1.0000 0.0082
-8.500 -0.4369 0.06793 0.06648 -0.0234 1.0000 0.0083
-8.250 -0.4437 0.06270 0.06127 -0.0258 1.0000 0.0083
-8.000 -0.4532 0.05734 0.05593 -0.0287 1.0000 0.0083
-7.750 -0.4755 0.05125 0.04987 -0.0347 1.0000 0.0079
-7.500 -0.4863 0.04488 0.04341 -0.0391 1.0000 0.0078
-7.250 -0.4904 0.03918 0.03761 -0.0412 1.0000 0.0079
-7.000 -0.4905 0.03414 0.03245 -0.0418 1.0000 0.0081
-6.750 -0.4883 0.02940 0.02755 -0.0416 0.9996 0.0082
-6.500 -0.4671 0.02335 0.02122 -0.0450 0.9951 0.0087
-6.250 -0.4431 0.01871 0.01631 -0.0468 0.9906 0.0096
-6.000 -0.4110 0.01744 0.01488 -0.0478 0.9863 0.0106
-5.750 -0.4235 0.01790 0.01391 -0.0468 0.9852 0.0063
-5.500 -0.3918 0.01473 0.01037 -0.0480 0.9822 0.0057
-5.250 -0.3628 0.01296 0.00834 -0.0483 0.9755 0.0058
-5.000 -0.3330 0.01079 0.00588 -0.0488 0.9696 0.0065
-4.750 -0.3051 0.00989 0.00489 -0.0490 0.9606 0.0076
-4.500 -0.2770 0.00936 0.00427 -0.0490 0.9511 0.0083
-4.250 -0.2504 0.00887 0.00368 -0.0487 0.9400 0.0089
-4.000 -0.2243 0.00857 0.00331 -0.0482 0.9284 0.0099
-3.750 -0.1983 0.00822 0.00286 -0.0477 0.9168 0.0106
-3.500 -0.1727 0.00763 0.00211 -0.0470 0.9053 0.0147
-3.250 -0.1466 0.00729 0.00180 -0.0466 0.8943 0.0352
-3.000 -0.1202 0.00710 0.00161 -0.0462 0.8828 0.0544
-2.750 -0.0941 0.00684 0.00144 -0.0459 0.8704 0.0900
-2.500 -0.0684 0.00649 0.00127 -0.0456 0.8581 0.1581
-2.250 -0.0439 0.00579 0.00108 -0.0453 0.8471 0.3178
-2.000 -0.0192 0.00516 0.00096 -0.0450 0.8375 0.4844
-1.750 0.0072 0.00495 0.00089 -0.0448 0.8278 0.5526
-1.500 0.0339 0.00485 0.00084 -0.0445 0.8162 0.5963
-1.250 0.0609 0.00480 0.00081 -0.0442 0.8042 0.6267
-1.000 0.0880 0.00474 0.00079 -0.0440 0.7933 0.6575
-0.750 0.1152 0.00469 0.00077 -0.0438 0.7833 0.6856
-0.500 0.1423 0.00466 0.00077 -0.0435 0.7724 0.7110
-0.250 0.1692 0.00465 0.00077 -0.0432 0.7603 0.7364
0.000 0.1963 0.00464 0.00077 -0.0430 0.7481 0.7540
0.500 0.2506 0.00465 0.00079 -0.0425 0.7244 0.7882
1.000 0.3046 0.00468 0.00083 -0.0420 0.6970 0.8185
1.250 0.3315 0.00471 0.00085 -0.0417 0.6809 0.8321
1.500 0.3582 0.00473 0.00088 -0.0414 0.6626 0.8469
2.000 0.4096 0.00480 0.00094 -0.0404 0.6125 0.8890
2.250 0.4384 0.00485 0.00100 -0.0405 0.5691 0.9443
2.500 0.4801 0.00522 0.00112 -0.0439 0.4908 1.0000
2.750 0.5043 0.00564 0.00126 -0.0433 0.4226 1.0000
3.000 0.5287 0.00608 0.00142 -0.0428 0.3551 1.0000
3.250 0.5540 0.00641 0.00157 -0.0425 0.3133 1.0000
3.500 0.5802 0.00664 0.00174 -0.0422 0.2981 1.0000
3.750 0.6064 0.00686 0.00189 -0.0420 0.2611 1.0000
4.000 0.6318 0.00719 0.00206 -0.0417 0.2262 1.0000
4.250 0.6575 0.00750 0.00226 -0.0414 0.1970 1.0000
4.500 0.6821 0.00793 0.00249 -0.0410 0.1489 1.0000
4.750 0.7057 0.00852 0.00281 -0.0405 0.0942 1.0000
5.000 0.7301 0.00902 0.00315 -0.0400 0.0598 1.0000
5.250 0.7553 0.00939 0.00345 -0.0397 0.0432 1.0000
5.500 0.7807 0.00973 0.00374 -0.0393 0.0322 1.0000
5.750 0.8058 0.01013 0.00407 -0.0389 0.0187 1.0000
6.000 0.8295 0.01078 0.00466 -0.0382 0.0069 1.0000
6.250 0.8541 0.01130 0.00527 -0.0376 0.0059 1.0000
6.500 0.8790 0.01173 0.00575 -0.0371 0.0053 1.0000
6.750 0.9033 0.01224 0.00636 -0.0366 0.0049 1.0000
7.000 0.9270 0.01285 0.00706 -0.0359 0.0047 1.0000
7.250 0.9505 0.01347 0.00775 -0.0352 0.0044 1.0000
7.500 0.9731 0.01421 0.00859 -0.0344 0.0041 1.0000
7.750 0.9932 0.01533 0.00984 -0.0332 0.0036 1.0000
8.000 1.0067 0.01772 0.01253 -0.0309 0.0032 1.0000
8.250 1.0280 0.01870 0.01364 -0.0300 0.0031 1.0000
8.500 1.0462 0.02030 0.01544 -0.0285 0.0030 1.0000
8.750 1.0624 0.02232 0.01769 -0.0269 0.0029 1.0000
9.000 1.0748 0.02502 0.02072 -0.0248 0.0029 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 1.5/8 AIRFOIL (hq158-il)