HQ 2.5/10 AIRFOIL (hq2510-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 2.5/10 AIRFOIL (hq2510-il) Reynolds number: 500,000 Max Cl/Cd: 105.63 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq2510-il-500000.txt Download as CSV file: xf-hq2510-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.5/10 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3324 0.08432 0.08225 -0.0397 1.0000 0.0174
-9.500 -0.3334 0.08056 0.07851 -0.0400 1.0000 0.0179
-9.250 -0.3338 0.07698 0.07494 -0.0405 1.0000 0.0182
-9.000 -0.3349 0.07342 0.07141 -0.0410 1.0000 0.0185
-8.750 -0.3386 0.06956 0.06759 -0.0416 1.0000 0.0188
-8.500 -0.3434 0.06591 0.06397 -0.0420 1.0000 0.0192
-8.250 -0.3545 0.06195 0.06006 -0.0423 1.0000 0.0193
-8.000 -0.3746 0.05855 0.05673 -0.0412 0.9995 0.0191
-7.750 -0.4576 0.05609 0.05396 -0.0585 0.9943 0.0174
-7.500 -0.4430 0.04976 0.04744 -0.0655 0.9879 0.0179
-7.250 -0.4187 0.04680 0.04438 -0.0698 0.9843 0.0188
-7.000 -0.3934 0.04278 0.04019 -0.0743 0.9804 0.0200
-6.750 -0.3678 0.03854 0.03568 -0.0778 0.9749 0.0222
-6.500 -0.3451 0.02662 0.02269 -0.0820 0.9699 0.0183
-6.250 -0.3232 0.01974 0.01504 -0.0819 0.9626 0.0133
-6.000 -0.2948 0.01607 0.01081 -0.0822 0.9577 0.0125
-5.750 -0.2663 0.01451 0.00903 -0.0824 0.9524 0.0127
-5.500 -0.2395 0.01349 0.00788 -0.0822 0.9455 0.0136
-5.250 -0.2116 0.01262 0.00690 -0.0822 0.9398 0.0148
-5.000 -0.1863 0.01194 0.00609 -0.0816 0.9319 0.0161
-4.750 -0.1606 0.01109 0.00511 -0.0811 0.9254 0.0178
-4.500 -0.1348 0.01066 0.00467 -0.0807 0.9177 0.0225
-4.250 -0.1088 0.01005 0.00401 -0.0803 0.9111 0.0312
-4.000 -0.0824 0.00974 0.00368 -0.0799 0.9036 0.0445
-3.750 -0.0556 0.00950 0.00339 -0.0797 0.8969 0.0540
-3.500 -0.0289 0.00931 0.00315 -0.0794 0.8895 0.0630
-3.250 -0.0022 0.00901 0.00289 -0.0792 0.8827 0.0803
-3.000 0.0240 0.00860 0.00261 -0.0789 0.8753 0.1203
-2.750 0.0488 0.00784 0.00235 -0.0788 0.8681 0.2648
-2.500 0.0727 0.00703 0.00216 -0.0785 0.8605 0.4634
-2.250 0.0986 0.00678 0.00212 -0.0781 0.8527 0.5501
-2.000 0.1254 0.00669 0.00208 -0.0777 0.8453 0.5954
-1.750 0.1522 0.00664 0.00206 -0.0774 0.8373 0.6300
-1.500 0.1795 0.00662 0.00202 -0.0770 0.8306 0.6555
-1.250 0.2067 0.00660 0.00202 -0.0768 0.8226 0.6770
-1.000 0.2339 0.00660 0.00199 -0.0764 0.8153 0.6955
-0.750 0.2604 0.00659 0.00200 -0.0759 0.8058 0.7200
-0.500 0.2868 0.00658 0.00201 -0.0754 0.7961 0.7402
-0.250 0.3138 0.00658 0.00197 -0.0750 0.7867 0.7537
0.000 0.3407 0.00658 0.00196 -0.0747 0.7769 0.7666
0.250 0.3680 0.00657 0.00195 -0.0744 0.7680 0.7774
0.500 0.3954 0.00659 0.00193 -0.0742 0.7595 0.7875
0.750 0.4224 0.00658 0.00192 -0.0740 0.7496 0.7968
1.000 0.4495 0.00657 0.00192 -0.0737 0.7392 0.8069
1.250 0.4764 0.00658 0.00193 -0.0734 0.7285 0.8178
1.500 0.5032 0.00659 0.00193 -0.0731 0.7178 0.8299
1.750 0.5296 0.00659 0.00194 -0.0727 0.7064 0.8428
2.000 0.5556 0.00657 0.00196 -0.0722 0.6941 0.8577
2.250 0.5810 0.00655 0.00200 -0.0715 0.6809 0.8763
2.500 0.6055 0.00651 0.00202 -0.0706 0.6657 0.9046
2.750 0.6404 0.00648 0.00204 -0.0720 0.6481 0.9625
3.000 0.6731 0.00657 0.00209 -0.0732 0.6292 1.0000
3.250 0.6995 0.00673 0.00217 -0.0730 0.6081 1.0000
3.500 0.7256 0.00689 0.00226 -0.0727 0.5829 1.0000
3.750 0.7510 0.00711 0.00238 -0.0722 0.5538 1.0000
4.000 0.7754 0.00740 0.00252 -0.0716 0.5157 1.0000
4.250 0.7988 0.00778 0.00273 -0.0709 0.4713 1.0000
4.500 0.8218 0.00822 0.00296 -0.0702 0.4252 1.0000
4.750 0.8449 0.00867 0.00323 -0.0695 0.3819 1.0000
5.000 0.8683 0.00911 0.00351 -0.0688 0.3439 1.0000
5.250 0.8918 0.00953 0.00380 -0.0682 0.3103 1.0000
5.500 0.9148 0.01002 0.00415 -0.0675 0.2775 1.0000
5.750 0.9365 0.01061 0.00450 -0.0667 0.2345 1.0000
6.000 0.9597 0.01106 0.00483 -0.0661 0.2042 1.0000
6.250 0.9830 0.01150 0.00517 -0.0655 0.1799 1.0000
6.500 1.0061 0.01195 0.00553 -0.0649 0.1566 1.0000
6.750 1.0284 0.01247 0.00595 -0.0642 0.1309 1.0000
7.000 1.0490 0.01315 0.00644 -0.0632 0.0982 1.0000
7.250 1.0662 0.01417 0.00712 -0.0618 0.0511 1.0000
7.500 1.0802 0.01554 0.00820 -0.0597 0.0116 1.0000
7.750 1.1004 0.01626 0.00897 -0.0585 0.0090 1.0000
8.000 1.1200 0.01703 0.00987 -0.0572 0.0080 1.0000
8.250 1.1398 0.01773 0.01067 -0.0560 0.0076 1.0000
8.500 1.1582 0.01852 0.01157 -0.0546 0.0071 1.0000
8.750 1.1751 0.01941 0.01259 -0.0529 0.0068 1.0000
9.000 1.1897 0.02043 0.01371 -0.0510 0.0063 1.0000
9.250 1.2028 0.02150 0.01489 -0.0489 0.0061 1.0000
9.500 1.2125 0.02261 0.01611 -0.0462 0.0060 1.0000
9.750 1.2195 0.02384 0.01744 -0.0432 0.0059 1.0000
10.000 1.2254 0.02520 0.01893 -0.0402 0.0058 1.0000
10.250 1.2314 0.02664 0.02049 -0.0375 0.0058 1.0000
10.500 1.2364 0.02829 0.02227 -0.0348 0.0058 1.0000
10.750 1.2416 0.03009 0.02423 -0.0324 0.0059 1.0000
11.000 1.2462 0.03214 0.02645 -0.0300 0.0060 1.0000
11.250 1.2497 0.03450 0.02901 -0.0278 0.0061 1.0000
11.500 1.2517 0.03718 0.03194 -0.0256 0.0064 1.0000
11.750 1.2494 0.04058 0.03561 -0.0234 0.0067 1.0000
12.000 1.2441 0.04407 0.03940 -0.0215 0.0070 1.0000
12.250 1.2351 0.04792 0.04350 -0.0200 0.0072 1.0000
12.500 1.2234 0.05209 0.04790 -0.0190 0.0074 1.0000
12.750 1.2073 0.05699 0.05304 -0.0186 0.0077 1.0000
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Polar data table (+)
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