HQ 2.1/9.5 AIRFOIL (hq2195-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 2.1/9.5 AIRFOIL (hq2195-il) Reynolds number: 100,000 Max Cl/Cd: 55.4 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq2195-il-100000.txt Download as CSV file: xf-hq2195-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.4596 0.11642 0.11113 -0.0214 1.0000 0.0919
-10.250 -0.4755 0.11405 0.10886 -0.0270 1.0000 0.0942
-10.000 -0.4936 0.11126 0.10619 -0.0328 1.0000 0.0947
-9.750 -0.4525 0.10537 0.10020 -0.0251 1.0000 0.0991
-9.500 -0.4480 0.10227 0.09711 -0.0258 1.0000 0.1041
-9.250 -0.4596 0.09915 0.09410 -0.0300 1.0000 0.1079
-9.000 -0.4861 0.09587 0.09098 -0.0368 1.0000 0.1089
-8.750 -0.4460 0.09195 0.08695 -0.0291 1.0000 0.1164
-8.500 -0.4551 0.08874 0.08383 -0.0319 1.0000 0.1214
-8.250 -0.4847 0.08515 0.08042 -0.0376 1.0000 0.1230
-8.000 -0.4635 0.08160 0.07686 -0.0332 1.0000 0.1278
-7.750 -0.4640 0.07887 0.07420 -0.0327 1.0000 0.1323
-7.500 -0.4990 0.07415 0.06952 -0.0417 1.0000 0.1377
-7.250 -0.4938 0.07022 0.06567 -0.0398 1.0000 0.1412
-7.000 -0.4859 0.06799 0.06351 -0.0371 1.0000 0.1469
-6.750 -0.5098 0.06404 0.05943 -0.0403 1.0000 0.1549
-6.500 -0.4981 0.06190 0.05742 -0.0362 1.0000 0.1600
-6.250 -0.5082 0.05869 0.05410 -0.0367 1.0000 0.1708
-6.000 -0.5028 0.05659 0.05203 -0.0343 1.0000 0.1792
-5.750 -0.5004 0.05391 0.04934 -0.0328 1.0000 0.1900
-5.500 -0.4961 0.05150 0.04690 -0.0315 1.0000 0.2051
-5.000 -0.4450 0.03626 0.02959 -0.0396 1.0000 0.1048
-4.750 -0.4157 0.03219 0.02419 -0.0388 1.0000 0.0808
-4.500 -0.3939 0.02904 0.02084 -0.0384 1.0000 0.0789
-4.250 -0.3706 0.02689 0.01834 -0.0378 1.0000 0.0780
-4.000 -0.3467 0.02549 0.01657 -0.0371 1.0000 0.0795
-3.750 -0.3225 0.02390 0.01471 -0.0366 1.0000 0.0818
-3.500 -0.2989 0.02243 0.01319 -0.0359 1.0000 0.0839
-3.250 -0.2757 0.02138 0.01206 -0.0352 1.0000 0.0873
-3.000 -0.2527 0.02059 0.01118 -0.0343 1.0000 0.0927
-2.750 -0.2302 0.01965 0.01036 -0.0337 1.0000 0.1017
-2.500 -0.2076 0.01884 0.00960 -0.0329 1.0000 0.1128
-2.250 -0.1845 0.01811 0.00899 -0.0324 1.0000 0.1386
-2.000 -0.1648 0.01532 0.00872 -0.0313 1.0000 0.6187
-1.750 -0.1556 0.01545 0.00915 -0.0264 1.0000 0.7479
-1.500 -0.1458 0.01559 0.00938 -0.0220 1.0000 0.8090
-1.250 -0.1354 0.01565 0.00950 -0.0178 0.9988 0.8650
-1.000 -0.0913 0.01565 0.00956 -0.0198 0.9924 0.9681
-0.750 -0.0368 0.01601 0.00967 -0.0264 0.9839 1.0000
-0.500 0.0103 0.01633 0.00977 -0.0313 0.9739 1.0000
-0.250 0.0571 0.01666 0.00992 -0.0359 0.9640 1.0000
0.000 0.1091 0.01705 0.01016 -0.0413 0.9557 1.0000
0.250 0.1457 0.01724 0.01023 -0.0437 0.9453 1.0000
0.500 0.1878 0.01750 0.01040 -0.0470 0.9363 1.0000
0.750 0.2328 0.01772 0.01056 -0.0507 0.9277 1.0000
1.000 0.2692 0.01790 0.01069 -0.0527 0.9170 1.0000
1.250 0.3138 0.01803 0.01080 -0.0559 0.9077 1.0000
1.500 0.3590 0.01805 0.01083 -0.0592 0.8980 1.0000
1.750 0.3973 0.01807 0.01087 -0.0610 0.8862 1.0000
2.000 0.4423 0.01796 0.01081 -0.0639 0.8755 1.0000
2.250 0.4937 0.01761 0.01054 -0.0674 0.8666 1.0000
2.500 0.5297 0.01740 0.01040 -0.0682 0.8530 1.0000
2.750 0.5658 0.01710 0.01019 -0.0687 0.8388 1.0000
3.000 0.6010 0.01673 0.00990 -0.0689 0.8236 1.0000
3.250 0.6353 0.01628 0.00953 -0.0686 0.8072 1.0000
3.500 0.6627 0.01596 0.00930 -0.0673 0.7853 1.0000
3.750 0.6940 0.01540 0.00877 -0.0661 0.7626 1.0000
4.000 0.7207 0.01493 0.00831 -0.0642 0.7325 1.0000
4.250 0.7442 0.01462 0.00800 -0.0621 0.6934 1.0000
4.500 0.7680 0.01442 0.00777 -0.0602 0.6474 1.0000
4.750 0.7915 0.01442 0.00763 -0.0584 0.5951 1.0000
5.000 0.8144 0.01470 0.00763 -0.0568 0.5427 1.0000
5.250 0.8364 0.01526 0.00792 -0.0553 0.4947 1.0000
5.500 0.8576 0.01594 0.00841 -0.0539 0.4501 1.0000
5.750 0.8779 0.01665 0.00894 -0.0525 0.4061 1.0000
6.000 0.8964 0.01739 0.00953 -0.0509 0.3539 1.0000
6.250 0.9117 0.01834 0.01018 -0.0488 0.2850 1.0000
6.500 0.9250 0.01969 0.01104 -0.0467 0.2090 1.0000
6.750 0.9402 0.02119 0.01214 -0.0450 0.1636 1.0000
7.000 0.9572 0.02274 0.01345 -0.0435 0.1356 1.0000
7.250 0.9756 0.02443 0.01499 -0.0422 0.1152 1.0000
7.500 0.9965 0.02615 0.01665 -0.0412 0.1000 1.0000
7.750 1.0209 0.02828 0.01866 -0.0407 0.0902 1.0000
8.000 1.0442 0.02985 0.02039 -0.0399 0.0822 1.0000
8.250 1.0696 0.03228 0.02286 -0.0396 0.0766 1.0000
8.500 1.0932 0.03453 0.02544 -0.0387 0.0729 1.0000
8.750 1.1149 0.03672 0.02782 -0.0380 0.0693 1.0000
9.000 1.1352 0.04028 0.03147 -0.0376 0.0656 1.0000
9.250 1.1484 0.04286 0.03462 -0.0356 0.0643 1.0000
9.500 1.1587 0.04625 0.03853 -0.0335 0.0635 1.0000
9.750 1.1647 0.04996 0.04274 -0.0313 0.0630 1.0000
10.000 1.1651 0.05372 0.04697 -0.0288 0.0622 1.0000
10.250 1.1605 0.05756 0.05124 -0.0262 0.0616 1.0000
10.500 1.1508 0.06156 0.05561 -0.0237 0.0614 1.0000
10.750 1.1350 0.06563 0.05998 -0.0210 0.0616 1.0000
11.000 1.1136 0.06963 0.06422 -0.0183 0.0621 1.0000
11.250 1.0913 0.07409 0.06886 -0.0171 0.0628 1.0000
11.500 1.0680 0.07915 0.07409 -0.0172 0.0635 1.0000
11.750 1.0459 0.08489 0.07996 -0.0186 0.0643 1.0000
12.000 1.0275 0.09134 0.08649 -0.0209 0.0651 1.0000
12.250 0.9441 0.11314 0.10854 -0.0369 0.0797 1.0000
12.500 0.9561 0.11646 0.11188 -0.0354 0.0788 1.0000
12.750 0.8578 0.15295 0.14810 -0.0647 0.1352 1.0000
13.000 0.8948 0.15748 0.15276 -0.0596 0.1321 1.0000
13.250 0.8559 0.16298 0.15809 -0.0698 0.1287 1.0000
13.500 0.7275 0.16067 0.15639 -0.0588 0.1319 1.0000
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Polar data table (+)
Polar graphs
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