HQ 2.1/9.5 AIRFOIL (hq2195-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 2.1/9.5 AIRFOIL (hq2195-il) Reynolds number: 200,000 Max Cl/Cd: 74.51 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq2195-il-200000.txt Download as CSV file: xf-hq2195-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4696 0.08974 0.08616 -0.0391 1.0000 0.0537
-9.000 -0.4851 0.08392 0.08040 -0.0459 1.0000 0.0539
-8.750 -0.5015 0.07993 0.07639 -0.0491 1.0000 0.0540
-8.500 -0.5135 0.07637 0.07274 -0.0509 1.0000 0.0542
-8.250 -0.5080 0.06973 0.06628 -0.0496 1.0000 0.0556
-8.000 -0.4991 0.06765 0.06427 -0.0474 1.0000 0.0568
-7.750 -0.4997 0.06482 0.06144 -0.0470 1.0000 0.0578
-7.500 -0.5036 0.06186 0.05848 -0.0465 1.0000 0.0592
-7.250 -0.5107 0.05899 0.05557 -0.0454 1.0000 0.0606
-7.000 -0.5191 0.05628 0.05278 -0.0437 1.0000 0.0624
-6.750 -0.5328 0.05578 0.05161 -0.0421 1.0000 0.0668
-6.500 -0.5366 0.04949 0.04527 -0.0414 1.0000 0.0682
-6.250 -0.5170 0.04602 0.04192 -0.0423 0.9981 0.0701
-6.000 -0.4860 0.04299 0.03875 -0.0456 0.9944 0.0748
-5.750 -0.4566 0.03913 0.03451 -0.0492 0.9894 0.0845
-5.500 -0.4231 0.03642 0.03133 -0.0527 0.9851 0.0967
-5.250 -0.3840 0.02619 0.01985 -0.0533 0.9832 0.0488
-5.000 -0.3537 0.02319 0.01632 -0.0538 0.9788 0.0468
-4.750 -0.3202 0.02169 0.01457 -0.0552 0.9750 0.0488
-4.500 -0.2845 0.01992 0.01251 -0.0566 0.9723 0.0490
-4.250 -0.2472 0.01857 0.01098 -0.0584 0.9701 0.0502
-4.000 -0.2189 0.01764 0.00994 -0.0585 0.9647 0.0520
-3.750 -0.1862 0.01665 0.00887 -0.0596 0.9610 0.0557
-3.500 -0.1512 0.01571 0.00799 -0.0612 0.9581 0.0603
-3.250 -0.1132 0.01508 0.00731 -0.0633 0.9559 0.0670
-3.000 -0.0875 0.01440 0.00669 -0.0633 0.9498 0.0795
-2.750 -0.0548 0.01347 0.00599 -0.0647 0.9460 0.1256
-2.500 -0.0266 0.01145 0.00584 -0.0661 0.9432 0.5670
-2.250 0.0093 0.01139 0.00600 -0.0673 0.9410 0.6626
-2.000 0.0308 0.01150 0.00615 -0.0658 0.9340 0.7047
-1.750 0.0657 0.01145 0.00615 -0.0666 0.9297 0.7392
-1.500 0.1030 0.01129 0.00602 -0.0675 0.9234 0.7713
-1.250 0.1333 0.01113 0.00591 -0.0666 0.9142 0.8035
-1.000 0.1592 0.01106 0.00587 -0.0652 0.9064 0.8298
-0.750 0.1862 0.01098 0.00578 -0.0644 0.9002 0.8465
-0.500 0.2173 0.01084 0.00561 -0.0646 0.8959 0.8598
-0.250 0.2393 0.01080 0.00558 -0.0631 0.8874 0.8740
0.000 0.2681 0.01065 0.00541 -0.0628 0.8823 0.8879
0.250 0.2909 0.01057 0.00536 -0.0615 0.8741 0.9035
0.500 0.3191 0.01041 0.00519 -0.0610 0.8680 0.9193
0.750 0.3478 0.01030 0.00511 -0.0608 0.8601 0.9370
1.000 0.3851 0.01014 0.00495 -0.0624 0.8535 0.9537
1.250 0.4265 0.01003 0.00486 -0.0650 0.8452 0.9710
1.500 0.4709 0.00987 0.00468 -0.0682 0.8374 0.9891
1.750 0.4992 0.00984 0.00466 -0.0687 0.8264 1.0000
2.000 0.5238 0.00983 0.00462 -0.0681 0.8155 1.0000
2.250 0.5508 0.00980 0.00455 -0.0677 0.8043 1.0000
2.500 0.5779 0.00976 0.00449 -0.0673 0.7916 1.0000
2.750 0.6043 0.00973 0.00445 -0.0667 0.7766 1.0000
3.000 0.6296 0.00970 0.00441 -0.0659 0.7571 1.0000
3.250 0.6553 0.00962 0.00427 -0.0649 0.7327 1.0000
3.500 0.6798 0.00960 0.00420 -0.0637 0.6997 1.0000
3.750 0.7042 0.00964 0.00416 -0.0626 0.6602 1.0000
4.000 0.7280 0.00977 0.00415 -0.0615 0.6061 1.0000
4.250 0.7496 0.01017 0.00419 -0.0600 0.5384 1.0000
4.500 0.7709 0.01074 0.00446 -0.0588 0.4846 1.0000
4.750 0.7934 0.01127 0.00480 -0.0578 0.4446 1.0000
5.000 0.8162 0.01180 0.00517 -0.0570 0.4094 1.0000
5.250 0.8390 0.01230 0.00555 -0.0562 0.3745 1.0000
5.500 0.8611 0.01283 0.00595 -0.0553 0.3347 1.0000
5.750 0.8826 0.01340 0.00634 -0.0544 0.2858 1.0000
6.000 0.9028 0.01414 0.00680 -0.0534 0.2215 1.0000
6.250 0.9204 0.01524 0.00748 -0.0521 0.1548 1.0000
6.500 0.9388 0.01634 0.00832 -0.0509 0.1193 1.0000
6.750 0.9586 0.01729 0.00919 -0.0497 0.0975 1.0000
7.000 0.9770 0.01839 0.01019 -0.0484 0.0808 1.0000
7.250 0.9955 0.01950 0.01126 -0.0471 0.0681 1.0000
7.500 1.0145 0.02064 0.01245 -0.0458 0.0601 1.0000
7.750 1.0319 0.02215 0.01388 -0.0444 0.0547 1.0000
8.000 1.0532 0.02328 0.01513 -0.0433 0.0511 1.0000
8.250 1.0742 0.02427 0.01619 -0.0424 0.0474 1.0000
8.500 1.0946 0.02583 0.01772 -0.0416 0.0445 1.0000
8.750 1.1172 0.02790 0.01994 -0.0409 0.0426 1.0000
9.000 1.1387 0.02926 0.02152 -0.0400 0.0406 1.0000
9.250 1.1588 0.03060 0.02302 -0.0390 0.0384 1.0000
9.500 1.1781 0.03210 0.02463 -0.0381 0.0367 1.0000
9.750 1.1975 0.03428 0.02692 -0.0373 0.0354 1.0000
10.000 1.2124 0.03885 0.03181 -0.0364 0.0342 1.0000
10.250 1.2227 0.04045 0.03379 -0.0342 0.0334 1.0000
10.500 1.2298 0.04266 0.03637 -0.0319 0.0323 1.0000
10.750 1.2326 0.04556 0.03965 -0.0294 0.0316 1.0000
11.000 1.2293 0.04893 0.04339 -0.0265 0.0313 1.0000
11.250 1.2190 0.05226 0.04704 -0.0231 0.0312 1.0000
11.500 1.2030 0.05577 0.05084 -0.0197 0.0313 1.0000
11.750 1.1839 0.05958 0.05492 -0.0172 0.0314 1.0000
12.000 1.1622 0.06392 0.05951 -0.0158 0.0316 1.0000
12.250 1.1390 0.06883 0.06465 -0.0157 0.0318 1.0000
12.500 1.1145 0.07445 0.07048 -0.0169 0.0321 1.0000
12.750 1.0893 0.08089 0.07710 -0.0196 0.0324 1.0000
13.000 1.0626 0.08850 0.08487 -0.0240 0.0327 1.0000
13.250 1.0356 0.09748 0.09399 -0.0299 0.0332 1.0000
13.500 1.0084 0.10791 0.10450 -0.0369 0.0337 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 2.1/9.5 AIRFOIL (hq2195-il)