HQ 1.5/9 B AIRFOIL (hq159b-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 1.5/9 B AIRFOIL (hq159b-il) Reynolds number: 100,000 Max Cl/Cd: 51.34 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq159b-il-100000.txt Download as CSV file: xf-hq159b-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/9 B AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4191 0.09993 0.09527 -0.0172 1.0000 0.1039
-9.500 -0.4370 0.09596 0.09138 -0.0216 1.0000 0.1061
-9.250 -0.5089 0.09884 0.09393 -0.0169 1.0000 0.0976
-9.000 -0.5096 0.09538 0.09052 -0.0188 1.0000 0.1038
-8.750 -0.5252 0.09122 0.08650 -0.0246 1.0000 0.1059
-8.500 -0.5091 0.08774 0.08299 -0.0214 1.0000 0.1129
-8.250 -0.5153 0.08398 0.07931 -0.0242 1.0000 0.1174
-8.000 -0.5347 0.07904 0.07447 -0.0310 1.0000 0.1192
-7.750 -0.5611 0.07440 0.06968 -0.0388 1.0000 0.1206
-7.500 -0.5355 0.07111 0.06657 -0.0339 1.0000 0.1324
-7.250 -0.5328 0.06716 0.06264 -0.0342 1.0000 0.1388
-7.000 -0.5427 0.06309 0.05840 -0.0378 1.0000 0.1496
-6.500 -0.5227 0.05694 0.05234 -0.0349 1.0000 0.1698
-6.250 -0.5178 0.05387 0.04923 -0.0343 1.0000 0.1839
-6.000 -0.5070 0.03807 0.03095 -0.0385 1.0000 0.0661
-5.750 -0.4916 0.03402 0.02675 -0.0372 1.0000 0.0635
-5.500 -0.4755 0.03079 0.02311 -0.0355 1.0000 0.0626
-5.250 -0.4572 0.02924 0.02091 -0.0335 1.0000 0.0656
-5.000 -0.4398 0.02594 0.01750 -0.0324 1.0000 0.0694
-4.750 -0.4196 0.02405 0.01537 -0.0309 1.0000 0.0735
-4.500 -0.3990 0.02242 0.01343 -0.0295 1.0000 0.0814
-4.250 -0.3781 0.02098 0.01192 -0.0283 1.0000 0.0888
-4.000 -0.3573 0.01964 0.01051 -0.0271 1.0000 0.0987
-3.750 -0.3360 0.01853 0.00932 -0.0257 1.0000 0.1079
-3.500 -0.3162 0.01747 0.00841 -0.0246 1.0000 0.1235
-3.250 -0.2959 0.01654 0.00764 -0.0234 1.0000 0.1435
-3.000 -0.2755 0.01546 0.00694 -0.0225 1.0000 0.2010
-2.750 -0.2669 0.01334 0.00706 -0.0188 1.0000 0.6483
-2.500 -0.2561 0.01346 0.00733 -0.0145 1.0000 0.7372
-2.250 -0.2447 0.01355 0.00745 -0.0107 1.0000 0.7893
-2.000 -0.2339 0.01359 0.00750 -0.0067 1.0000 0.8346
-1.750 -0.2229 0.01358 0.00751 -0.0027 1.0000 0.8822
-1.500 -0.1857 0.01365 0.00754 -0.0035 1.0000 0.9526
-1.250 -0.1249 0.01379 0.00745 -0.0109 1.0000 1.0000
-1.000 -0.1079 0.01376 0.00726 -0.0111 0.9979 1.0000
-0.750 -0.0630 0.01407 0.00736 -0.0159 0.9896 1.0000
-0.500 -0.0170 0.01446 0.00758 -0.0207 0.9824 1.0000
-0.250 0.0225 0.01472 0.00771 -0.0241 0.9735 1.0000
0.000 0.0646 0.01505 0.00792 -0.0278 0.9657 1.0000
0.250 0.1068 0.01536 0.00815 -0.0314 0.9574 1.0000
0.500 0.1445 0.01564 0.00836 -0.0341 0.9480 1.0000
0.750 0.1909 0.01592 0.00861 -0.0383 0.9404 1.0000
1.000 0.2368 0.01603 0.00872 -0.0420 0.9286 1.0000
1.250 0.2935 0.01586 0.00858 -0.0471 0.9145 1.0000
1.500 0.3471 0.01557 0.00836 -0.0513 0.9005 1.0000
1.750 0.3862 0.01548 0.00831 -0.0531 0.8877 1.0000
2.000 0.4215 0.01544 0.00832 -0.0541 0.8752 1.0000
2.250 0.4551 0.01538 0.00836 -0.0547 0.8624 1.0000
2.500 0.4868 0.01531 0.00836 -0.0548 0.8488 1.0000
2.750 0.5166 0.01522 0.00836 -0.0544 0.8341 1.0000
3.000 0.5451 0.01511 0.00832 -0.0536 0.8182 1.0000
3.250 0.5728 0.01497 0.00830 -0.0525 0.8016 1.0000
3.500 0.5977 0.01488 0.00830 -0.0510 0.7817 1.0000
3.750 0.6234 0.01462 0.00811 -0.0492 0.7591 1.0000
4.000 0.6475 0.01429 0.00780 -0.0469 0.7295 1.0000
4.250 0.6698 0.01409 0.00767 -0.0445 0.6921 1.0000
4.500 0.6927 0.01398 0.00756 -0.0424 0.6496 1.0000
4.750 0.7139 0.01396 0.00740 -0.0400 0.5884 1.0000
5.000 0.7326 0.01427 0.00736 -0.0375 0.5027 1.0000
5.250 0.7508 0.01502 0.00770 -0.0355 0.4226 1.0000
5.500 0.7700 0.01581 0.00823 -0.0340 0.3537 1.0000
5.750 0.7869 0.01688 0.00891 -0.0323 0.2779 1.0000
6.000 0.7975 0.01898 0.01008 -0.0300 0.1415 1.0000
6.250 0.8122 0.02102 0.01172 -0.0279 0.1040 1.0000
6.500 0.8315 0.02271 0.01331 -0.0265 0.0902 1.0000
6.750 0.8533 0.02434 0.01497 -0.0254 0.0796 1.0000
7.000 0.8779 0.02666 0.01716 -0.0248 0.0732 1.0000
7.250 0.9018 0.02814 0.01889 -0.0239 0.0660 1.0000
7.500 0.9278 0.03145 0.02212 -0.0237 0.0613 1.0000
7.750 0.9500 0.03351 0.02472 -0.0223 0.0583 1.0000
8.000 0.9701 0.03571 0.02726 -0.0211 0.0538 1.0000
8.250 0.9889 0.03889 0.03081 -0.0198 0.0526 1.0000
8.500 1.0023 0.04274 0.03526 -0.0179 0.0526 1.0000
8.750 1.0075 0.04749 0.04073 -0.0155 0.0541 1.0000
9.000 1.0088 0.05217 0.04594 -0.0133 0.0550 1.0000
9.250 1.0055 0.05683 0.05104 -0.0113 0.0557 1.0000
9.500 0.9972 0.06163 0.05620 -0.0094 0.0568 1.0000
9.750 0.9858 0.06639 0.06123 -0.0080 0.0582 1.0000
10.000 0.9727 0.07091 0.06590 -0.0066 0.0592 1.0000
10.250 0.8636 0.06865 0.06416 -0.0018 0.0641 1.0000
10.500 0.8070 0.07777 0.07349 -0.0067 0.0664 1.0000
10.750 0.7647 0.08878 0.08457 -0.0140 0.0679 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 1.5/9 B AIRFOIL (hq159b-il)