HQ 2.5/8 AIRFOIL (hq258-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 2.5/8 AIRFOIL (hq258-il) Reynolds number: 1,000,000 Max Cl/Cd: 124.09 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq258-il-1000000.txt Download as CSV file: xf-hq258-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.5/8 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3970 0.08225 0.08070 -0.0307 1.0000 0.0073
-7.750 -0.3986 0.07897 0.07746 -0.0315 1.0000 0.0073
-7.500 -0.4057 0.07624 0.07477 -0.0311 1.0000 0.0073
-7.250 -0.4182 0.07395 0.07252 -0.0298 1.0000 0.0073
-7.000 -0.3926 0.06659 0.06515 -0.0428 0.9965 0.0073
-6.750 -0.3658 0.05929 0.05777 -0.0541 0.9917 0.0073
-6.500 -0.3392 0.05283 0.05119 -0.0623 0.9868 0.0073
-6.250 -0.3101 0.04651 0.04471 -0.0695 0.9835 0.0073
-6.000 -0.2804 0.04081 0.03881 -0.0749 0.9799 0.0073
-5.750 -0.2539 0.03590 0.03368 -0.0782 0.9730 0.0073
-5.500 -0.2344 0.02593 0.02312 -0.0818 0.9629 0.0079
-5.250 -0.2109 0.02242 0.01927 -0.0826 0.9550 0.0084
-5.000 -0.1882 0.02031 0.01692 -0.0823 0.9454 0.0089
-4.750 -0.1640 0.01840 0.01477 -0.0820 0.9372 0.0094
-4.250 -0.1135 0.01213 0.00765 -0.0797 0.9209 0.0062
-4.000 -0.0879 0.01038 0.00563 -0.0789 0.9138 0.0061
-3.750 -0.0618 0.00942 0.00452 -0.0784 0.9061 0.0067
-3.500 -0.0354 0.00880 0.00377 -0.0780 0.8991 0.0078
-3.250 -0.0092 0.00801 0.00281 -0.0776 0.8915 0.0085
-3.000 0.0178 0.00759 0.00229 -0.0774 0.8843 0.0098
-2.750 0.0450 0.00732 0.00194 -0.0771 0.8761 0.0113
-2.500 0.0725 0.00706 0.00157 -0.0769 0.8674 0.0135
-2.250 0.0997 0.00674 0.00133 -0.0767 0.8592 0.0513
-2.000 0.1271 0.00655 0.00119 -0.0766 0.8511 0.0814
-1.750 0.1537 0.00601 0.00105 -0.0767 0.8436 0.2183
-1.500 0.1786 0.00504 0.00095 -0.0768 0.8354 0.5115
-1.250 0.2056 0.00487 0.00092 -0.0766 0.8254 0.5804
-1.000 0.2328 0.00480 0.00088 -0.0764 0.8141 0.6201
-0.750 0.2601 0.00476 0.00085 -0.0762 0.8031 0.6498
-0.500 0.2873 0.00472 0.00084 -0.0760 0.7925 0.6801
-0.250 0.3143 0.00468 0.00085 -0.0757 0.7819 0.7166
0.000 0.3411 0.00465 0.00087 -0.0754 0.7705 0.7523
0.250 0.3681 0.00464 0.00088 -0.0751 0.7581 0.7723
0.500 0.3953 0.00466 0.00089 -0.0749 0.7456 0.7872
0.750 0.4226 0.00468 0.00089 -0.0748 0.7329 0.8000
1.000 0.4496 0.00470 0.00091 -0.0746 0.7192 0.8139
1.250 0.4765 0.00472 0.00093 -0.0743 0.7043 0.8297
1.500 0.5027 0.00475 0.00096 -0.0739 0.6867 0.8487
1.750 0.5281 0.00475 0.00100 -0.0733 0.6668 0.8765
2.000 0.5561 0.00466 0.00102 -0.0732 0.6465 0.9616
2.250 0.5874 0.00477 0.00106 -0.0740 0.6226 1.0000
2.500 0.6139 0.00495 0.00113 -0.0738 0.5933 1.0000
2.750 0.6403 0.00516 0.00124 -0.0736 0.5607 1.0000
3.000 0.6654 0.00548 0.00136 -0.0732 0.5113 1.0000
3.250 0.6903 0.00586 0.00152 -0.0727 0.4576 1.0000
3.500 0.7152 0.00625 0.00170 -0.0723 0.4055 1.0000
4.000 0.7625 0.00735 0.00219 -0.0713 0.2689 1.0000
4.250 0.7878 0.00771 0.00243 -0.0710 0.2331 1.0000
4.500 0.8121 0.00819 0.00268 -0.0706 0.1862 1.0000
4.750 0.8348 0.00887 0.00303 -0.0699 0.1226 1.0000
5.000 0.8557 0.00983 0.00355 -0.0690 0.0450 1.0000
5.250 0.8778 0.01066 0.00416 -0.0682 0.0064 1.0000
5.500 0.9034 0.01101 0.00457 -0.0677 0.0050 1.0000
5.750 0.9280 0.01151 0.00522 -0.0671 0.0040 1.0000
6.000 0.9523 0.01204 0.00583 -0.0664 0.0036 1.0000
6.250 0.9757 0.01272 0.00662 -0.0656 0.0034 1.0000
6.500 0.9978 0.01358 0.00762 -0.0645 0.0034 1.0000
6.750 1.0185 0.01462 0.00879 -0.0632 0.0034 1.0000
7.000 1.0372 0.01598 0.01031 -0.0615 0.0035 1.0000
7.250 1.0543 0.01776 0.01225 -0.0596 0.0038 1.0000
7.500 1.0741 0.01930 0.01388 -0.0582 0.0046 1.0000
15.000 0.9395 0.18834 0.18685 -0.0948 0.0057 1.0000
15.250 0.9441 0.19318 0.19169 -0.0974 0.0055 1.0000
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Polar data table (+)
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