HQ 2.1/9.5 AIRFOIL (hq2195-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 2.1/9.5 AIRFOIL (hq2195-il) Reynolds number: 500,000 Max Cl/Cd: 96.22 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq2195-il-500000.txt Download as CSV file: xf-hq2195-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.1/9.5 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.3814 0.08977 0.08754 -0.0329 1.0000 0.0255
-10.000 -0.3842 0.08468 0.08247 -0.0352 1.0000 0.0257
-9.750 -0.3889 0.07916 0.07697 -0.0376 1.0000 0.0258
-9.500 -0.3949 0.07337 0.07119 -0.0402 1.0000 0.0258
-9.250 -0.4036 0.06680 0.06464 -0.0435 1.0000 0.0259
-9.000 -0.4230 0.05794 0.05576 -0.0500 1.0000 0.0259
-8.750 -0.4454 0.05262 0.05040 -0.0536 1.0000 0.0259
-8.500 -0.4697 0.04955 0.04730 -0.0534 1.0000 0.0258
-8.250 -0.4886 0.04629 0.04397 -0.0519 1.0000 0.0259
-8.000 -0.5154 0.03696 0.03449 -0.0542 0.9967 0.0265
-7.750 -0.4950 0.03324 0.03072 -0.0573 0.9931 0.0271
-7.500 -0.4734 0.03004 0.02743 -0.0602 0.9884 0.0278
-7.250 -0.4493 0.02641 0.02363 -0.0635 0.9848 0.0292
-7.000 -0.4180 0.02443 0.02109 -0.0660 0.9797 0.0333
-6.750 -0.4491 0.02535 0.02096 -0.0662 0.9792 0.0242
-6.500 -0.4194 0.02098 0.01604 -0.0679 0.9751 0.0234
-6.250 -0.3856 0.01863 0.01335 -0.0696 0.9724 0.0239
-6.000 -0.3503 0.01736 0.01183 -0.0713 0.9700 0.0250
-5.750 -0.3209 0.01608 0.01037 -0.0718 0.9632 0.0254
-5.500 -0.2880 0.01521 0.00934 -0.0728 0.9585 0.0258
-5.250 -0.2591 0.01327 0.00728 -0.0734 0.9527 0.0267
-5.000 -0.2313 0.01241 0.00639 -0.0735 0.9452 0.0280
-4.750 -0.2022 0.01192 0.00587 -0.0738 0.9389 0.0296
-4.500 -0.1758 0.01139 0.00529 -0.0735 0.9305 0.0309
-4.250 -0.1486 0.01093 0.00475 -0.0733 0.9235 0.0323
-4.000 -0.1221 0.01060 0.00436 -0.0729 0.9154 0.0334
-3.750 -0.0963 0.00990 0.00361 -0.0726 0.9085 0.0368
-3.500 -0.0695 0.00965 0.00334 -0.0724 0.9012 0.0410
-3.250 -0.0424 0.00927 0.00293 -0.0722 0.8954 0.0485
-3.000 -0.0157 0.00894 0.00262 -0.0719 0.8881 0.0668
-2.750 0.0105 0.00845 0.00236 -0.0717 0.8808 0.1295
-2.500 0.0341 0.00744 0.00207 -0.0715 0.8710 0.3342
-2.250 0.0587 0.00683 0.00197 -0.0711 0.8620 0.5040
-2.000 0.0848 0.00662 0.00193 -0.0706 0.8535 0.5719
-1.750 0.1112 0.00652 0.00194 -0.0701 0.8454 0.6226
-1.500 0.1382 0.00649 0.00193 -0.0697 0.8379 0.6573
-1.250 0.1655 0.00648 0.00193 -0.0694 0.8308 0.6792
-1.000 0.1928 0.00647 0.00193 -0.0691 0.8246 0.7013
-0.750 0.2198 0.00646 0.00196 -0.0687 0.8187 0.7250
-0.500 0.2466 0.00646 0.00199 -0.0682 0.8120 0.7490
-0.250 0.2739 0.00646 0.00198 -0.0679 0.8063 0.7633
0.000 0.3012 0.00643 0.00199 -0.0676 0.7992 0.7737
0.250 0.3287 0.00643 0.00198 -0.0674 0.7928 0.7841
0.750 0.3835 0.00642 0.00200 -0.0668 0.7782 0.8053
1.000 0.4106 0.00641 0.00202 -0.0665 0.7701 0.8160
1.250 0.4377 0.00643 0.00202 -0.0662 0.7620 0.8276
1.500 0.4646 0.00641 0.00205 -0.0658 0.7522 0.8399
1.750 0.4911 0.00641 0.00208 -0.0653 0.7421 0.8525
2.000 0.5172 0.00640 0.00210 -0.0648 0.7304 0.8664
2.250 0.5427 0.00639 0.00211 -0.0640 0.7154 0.8820
2.500 0.5669 0.00636 0.00209 -0.0630 0.6915 0.9002
2.750 0.5901 0.00635 0.00206 -0.0617 0.6580 0.9249
3.000 0.6206 0.00645 0.00203 -0.0621 0.6011 0.9633
3.250 0.6485 0.00693 0.00213 -0.0624 0.5187 1.0000
3.500 0.6724 0.00740 0.00235 -0.0618 0.4664 1.0000
3.750 0.6975 0.00778 0.00256 -0.0615 0.4282 1.0000
4.000 0.7229 0.00812 0.00276 -0.0611 0.3974 1.0000
4.250 0.7482 0.00847 0.00298 -0.0608 0.3664 1.0000
4.500 0.7733 0.00882 0.00321 -0.0604 0.3351 1.0000
4.750 0.7982 0.00919 0.00344 -0.0600 0.3027 1.0000
5.000 0.8229 0.00958 0.00368 -0.0596 0.2649 1.0000
5.250 0.8471 0.01003 0.00397 -0.0591 0.2246 1.0000
5.500 0.8697 0.01066 0.00433 -0.0585 0.1735 1.0000
5.750 0.8920 0.01133 0.00476 -0.0578 0.1283 1.0000
6.000 0.9151 0.01192 0.00520 -0.0572 0.0998 1.0000
6.250 0.9387 0.01245 0.00563 -0.0566 0.0816 1.0000
6.500 0.9623 0.01296 0.00608 -0.0560 0.0678 1.0000
6.750 0.9852 0.01356 0.00662 -0.0553 0.0542 1.0000
7.000 1.0078 0.01419 0.00721 -0.0545 0.0431 1.0000
7.250 1.0313 0.01468 0.00771 -0.0539 0.0373 1.0000
7.500 1.0527 0.01542 0.00848 -0.0529 0.0333 1.0000
7.750 1.0755 0.01596 0.00907 -0.0522 0.0311 1.0000
8.000 1.0974 0.01658 0.00971 -0.0513 0.0290 1.0000
8.250 1.1140 0.01776 0.01095 -0.0497 0.0267 1.0000
8.500 1.1357 0.01834 0.01161 -0.0488 0.0257 1.0000
8.750 1.1561 0.01905 0.01241 -0.0478 0.0246 1.0000
9.000 1.1759 0.01979 0.01320 -0.0467 0.0234 1.0000
9.250 1.1951 0.02054 0.01399 -0.0455 0.0222 1.0000
9.500 1.2093 0.02186 0.01536 -0.0438 0.0210 1.0000
9.750 1.2232 0.02337 0.01701 -0.0420 0.0202 1.0000
10.000 1.2412 0.02420 0.01795 -0.0407 0.0196 1.0000
10.250 1.2574 0.02523 0.01909 -0.0392 0.0189 1.0000
10.500 1.2724 0.02624 0.02020 -0.0376 0.0181 1.0000
10.750 1.2851 0.02720 0.02124 -0.0356 0.0174 1.0000
11.000 1.2959 0.02815 0.02224 -0.0335 0.0167 1.0000
11.250 1.3034 0.02981 0.02398 -0.0313 0.0161 1.0000
11.500 1.3067 0.03272 0.02712 -0.0288 0.0155 1.0000
11.750 1.3139 0.03378 0.02833 -0.0266 0.0151 1.0000
12.000 1.3191 0.03534 0.03006 -0.0245 0.0148 1.0000
12.250 1.3224 0.03715 0.03205 -0.0224 0.0144 1.0000
12.500 1.3237 0.03919 0.03427 -0.0205 0.0141 1.0000
12.750 1.3237 0.04138 0.03663 -0.0189 0.0138 1.0000
13.000 1.3213 0.04393 0.03935 -0.0175 0.0136 1.0000
13.250 1.3187 0.04652 0.04208 -0.0164 0.0133 1.0000
13.500 1.3140 0.04942 0.04514 -0.0157 0.0131 1.0000
13.750 1.3106 0.05223 0.04807 -0.0155 0.0128 1.0000
14.000 1.3018 0.05605 0.05206 -0.0157 0.0127 1.0000
14.250 1.2953 0.05969 0.05583 -0.0163 0.0125 1.0000
14.500 1.2862 0.06399 0.06025 -0.0175 0.0123 1.0000
14.750 1.2726 0.06932 0.06575 -0.0195 0.0122 1.0000
15.000 1.2581 0.07518 0.07176 -0.0223 0.0121 1.0000
15.250 1.2411 0.08207 0.07883 -0.0261 0.0120 1.0000
15.500 1.2200 0.09047 0.08744 -0.0312 0.0121 1.0000
15.750 1.1899 0.10165 0.09889 -0.0385 0.0123 1.0000
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