HQ 2.5/10 AIRFOIL (hq2510-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 2.5/10 AIRFOIL (hq2510-il) Reynolds number: 1,000,000 Max Cl/Cd: 113.71 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq2510-il-1000000.txt Download as CSV file: xf-hq2510-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-14.000 -0.3571 0.14727 0.14568 -0.0167 1.0000 0.0069
-13.750 -0.3532 0.14425 0.14265 -0.0182 1.0000 0.0069
-7.500 -0.4397 0.02125 0.01792 -0.0819 0.9667 0.0064
-7.250 -0.4165 0.01904 0.01537 -0.0821 0.9575 0.0066
-7.000 -0.3938 0.01726 0.01331 -0.0817 0.9479 0.0067
-6.750 -0.3697 0.01604 0.01185 -0.0813 0.9392 0.0068
-6.500 -0.3458 0.01502 0.01064 -0.0808 0.9299 0.0069
-6.250 -0.3247 0.01288 0.00817 -0.0798 0.9205 0.0072
-6.000 -0.3016 0.01157 0.00665 -0.0791 0.9124 0.0075
-5.750 -0.2771 0.01071 0.00565 -0.0785 0.9041 0.0077
-5.500 -0.2519 0.01004 0.00486 -0.0781 0.8968 0.0083
-5.250 -0.2258 0.00959 0.00433 -0.0777 0.8892 0.0091
-5.000 -0.1993 0.00921 0.00386 -0.0774 0.8823 0.0101
-4.750 -0.1726 0.00891 0.00348 -0.0772 0.8751 0.0108
-4.500 -0.1461 0.00842 0.00287 -0.0769 0.8683 0.0131
-4.250 -0.1191 0.00815 0.00258 -0.0767 0.8613 0.0166
-4.000 -0.0921 0.00785 0.00235 -0.0765 0.8547 0.0327
-3.750 -0.0645 0.00774 0.00219 -0.0765 0.8476 0.0384
-3.500 -0.0370 0.00758 0.00201 -0.0764 0.8409 0.0460
-3.250 -0.0095 0.00743 0.00185 -0.0763 0.8333 0.0546
-3.000 0.0179 0.00725 0.00170 -0.0762 0.8257 0.0706
-2.750 0.0450 0.00701 0.00155 -0.0762 0.8178 0.1062
-2.500 0.0717 0.00658 0.00138 -0.0762 0.8101 0.1897
-2.250 0.0980 0.00613 0.00123 -0.0761 0.8029 0.2982
-2.000 0.1240 0.00558 0.00112 -0.0761 0.7952 0.4431
-1.750 0.1510 0.00537 0.00106 -0.0760 0.7874 0.5167
-1.500 0.1784 0.00526 0.00103 -0.0759 0.7779 0.5602
-1.250 0.2061 0.00521 0.00101 -0.0758 0.7687 0.5902
-1.000 0.2336 0.00518 0.00100 -0.0757 0.7592 0.6201
-0.750 0.2612 0.00516 0.00099 -0.0755 0.7491 0.6425
-0.500 0.2890 0.00516 0.00099 -0.0755 0.7400 0.6608
-0.250 0.3165 0.00516 0.00101 -0.0753 0.7315 0.6851
0.000 0.3442 0.00517 0.00102 -0.0752 0.7221 0.7021
0.250 0.3719 0.00518 0.00103 -0.0751 0.7120 0.7138
0.500 0.3995 0.00521 0.00105 -0.0750 0.7012 0.7258
0.750 0.4270 0.00525 0.00107 -0.0749 0.6905 0.7366
1.250 0.4823 0.00532 0.00112 -0.0747 0.6677 0.7543
1.500 0.5099 0.00537 0.00115 -0.0746 0.6554 0.7631
1.750 0.5371 0.00544 0.00119 -0.0745 0.6398 0.7726
2.000 0.5639 0.00551 0.00124 -0.0742 0.6221 0.7822
2.250 0.5909 0.00558 0.00130 -0.0740 0.6044 0.7925
2.500 0.6175 0.00568 0.00137 -0.0738 0.5835 0.8037
2.750 0.6436 0.00581 0.00144 -0.0734 0.5573 0.8162
3.000 0.6693 0.00595 0.00154 -0.0730 0.5300 0.8302
3.250 0.6941 0.00613 0.00166 -0.0724 0.4965 0.8470
3.500 0.7177 0.00635 0.00180 -0.0717 0.4569 0.8701
3.750 0.7391 0.00650 0.00196 -0.0703 0.4159 0.9258
4.000 0.7730 0.00684 0.00215 -0.0719 0.3683 1.0000
4.250 0.7979 0.00719 0.00235 -0.0715 0.3322 1.0000
4.500 0.8230 0.00752 0.00255 -0.0712 0.3021 1.0000
4.750 0.8476 0.00788 0.00277 -0.0708 0.2704 1.0000
5.000 0.8721 0.00826 0.00300 -0.0703 0.2394 1.0000
5.250 0.8959 0.00869 0.00325 -0.0698 0.2038 1.0000
5.500 0.9203 0.00906 0.00351 -0.0693 0.1776 1.0000
5.750 0.9445 0.00945 0.00378 -0.0689 0.1532 1.0000
6.000 0.9691 0.00979 0.00404 -0.0684 0.1343 1.0000
6.250 0.9919 0.01029 0.00439 -0.0677 0.1051 1.0000
6.500 1.0129 0.01098 0.00483 -0.0668 0.0677 1.0000
7.000 1.0507 0.01280 0.00626 -0.0641 0.0059 1.0000
7.250 1.0740 0.01324 0.00676 -0.0634 0.0051 1.0000
7.500 1.0971 0.01369 0.00728 -0.0626 0.0048 1.0000
7.750 1.1196 0.01418 0.00782 -0.0618 0.0046 1.0000
8.000 1.1415 0.01473 0.00843 -0.0609 0.0042 1.0000
8.250 1.1623 0.01536 0.00914 -0.0598 0.0039 1.0000
8.500 1.1820 0.01608 0.00994 -0.0586 0.0036 1.0000
8.750 1.2010 0.01682 0.01077 -0.0573 0.0035 1.0000
9.000 1.2168 0.01781 0.01187 -0.0555 0.0033 1.0000
9.250 1.2323 0.01876 0.01291 -0.0537 0.0033 1.0000
9.500 1.2464 0.01975 0.01399 -0.0517 0.0033 1.0000
9.750 1.2574 0.02089 0.01524 -0.0492 0.0032 1.0000
10.000 1.2654 0.02197 0.01642 -0.0462 0.0032 1.0000
10.250 1.2728 0.02305 0.01760 -0.0432 0.0033 1.0000
10.500 1.2780 0.02435 0.01902 -0.0401 0.0033 1.0000
10.750 1.2835 0.02574 0.02055 -0.0372 0.0034 1.0000
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Polar data table (+)
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