USA 5 AIRFOIL (usa5-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: USA 5 AIRFOIL (usa5-il) Reynolds number: 200,000 Max Cl/Cd: 72.37 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa5-il-200000.txt Download as CSV file: xf-usa5-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: USA 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
------ -------- --------- --------- -------- -------- --------
-6.250 -0.3575 0.08076 0.07779 -0.0179 1.0000 0.0327
-6.000 -0.3582 0.07847 0.07553 -0.0173 1.0000 0.0338
-5.750 -0.3576 0.07612 0.07320 -0.0171 1.0000 0.0347
-5.500 -0.3553 0.07367 0.07077 -0.0172 1.0000 0.0357
-5.250 -0.3507 0.07104 0.06813 -0.0177 1.0000 0.0371
-5.000 -0.3414 0.06840 0.06547 -0.0194 1.0000 0.0389
-4.750 -0.2736 0.06393 0.06067 -0.0347 0.9940 0.0408
-4.500 -0.2501 0.05686 0.05358 -0.0390 0.9879 0.0420
-4.250 -0.2172 0.05294 0.04964 -0.0427 0.9832 0.0440
-4.000 -0.1823 0.04932 0.04590 -0.0470 0.9761 0.0476
-3.750 -0.1248 0.04419 0.04026 -0.0556 0.9705 0.0538
-3.500 -0.1002 0.04067 0.03680 -0.0572 0.9635 0.0564
-2.500 0.0493 0.02103 0.01498 -0.0644 0.9417 0.0403
-2.250 0.0764 0.01811 0.01153 -0.0639 0.9335 0.0431
-2.000 0.1171 0.01678 0.00994 -0.0660 0.9296 0.0497
-1.750 0.1547 0.01515 0.00799 -0.0673 0.9243 0.0566
-1.500 0.1908 0.01469 0.00748 -0.0685 0.9168 0.0703
-1.250 0.2255 0.01433 0.00705 -0.0695 0.9086 0.0865
-1.000 0.2618 0.01371 0.00629 -0.0706 0.9010 0.1052
-0.750 0.2912 0.01307 0.00570 -0.0706 0.8899 0.1257
-0.500 0.3217 0.01255 0.00516 -0.0708 0.8784 0.1411
-0.250 0.3519 0.01212 0.00472 -0.0708 0.8648 0.1552
0.000 0.3812 0.01175 0.00435 -0.0706 0.8485 0.1698
0.250 0.4105 0.01135 0.00390 -0.0704 0.8307 0.1800
0.500 0.4363 0.01108 0.00361 -0.0695 0.8082 0.1912
0.750 0.4626 0.01084 0.00333 -0.0688 0.7859 0.2034
1.000 0.4875 0.01062 0.00311 -0.0678 0.7622 0.2233
1.250 0.5988 0.00885 0.00285 -0.0864 0.7244 1.0000
1.500 0.6221 0.00903 0.00279 -0.0850 0.6900 1.0000
1.750 0.6445 0.00927 0.00278 -0.0836 0.6530 1.0000
2.000 0.6665 0.00956 0.00284 -0.0822 0.6155 1.0000
2.250 0.6884 0.00988 0.00292 -0.0808 0.5797 1.0000
2.500 0.7106 0.01018 0.00303 -0.0795 0.5460 1.0000
2.750 0.7332 0.01049 0.00316 -0.0784 0.5195 1.0000
3.000 0.7561 0.01080 0.00333 -0.0774 0.4985 1.0000
3.250 0.7794 0.01112 0.00355 -0.0765 0.4814 1.0000
3.500 0.8030 0.01143 0.00376 -0.0756 0.4673 1.0000
3.750 0.8268 0.01174 0.00400 -0.0748 0.4552 1.0000
4.000 0.8510 0.01202 0.00427 -0.0741 0.4441 1.0000
4.250 0.8751 0.01232 0.00456 -0.0734 0.4338 1.0000
4.500 0.8986 0.01262 0.00487 -0.0725 0.4212 1.0000
4.750 0.9192 0.01289 0.00505 -0.0711 0.3953 1.0000
5.000 0.9416 0.01314 0.00530 -0.0701 0.3775 1.0000
5.250 0.9640 0.01339 0.00555 -0.0691 0.3609 1.0000
5.500 0.9856 0.01364 0.00579 -0.0679 0.3377 1.0000
5.750 1.0067 0.01391 0.00606 -0.0666 0.3065 1.0000
6.000 1.0270 0.01430 0.00632 -0.0652 0.2570 1.0000
6.250 1.0414 0.01536 0.00685 -0.0631 0.1738 1.0000
6.500 1.0531 0.01687 0.00778 -0.0607 0.0796 1.0000
6.750 1.0642 0.01843 0.00901 -0.0579 0.0295 1.0000
7.000 1.0827 0.01922 0.00996 -0.0561 0.0267 1.0000
7.250 1.1006 0.02006 0.01102 -0.0542 0.0253 1.0000
7.500 1.1166 0.02103 0.01218 -0.0521 0.0244 1.0000
7.750 1.1300 0.02214 0.01348 -0.0496 0.0239 1.0000
8.000 1.1361 0.02377 0.01525 -0.0461 0.0225 1.0000
8.250 1.1418 0.02543 0.01699 -0.0426 0.0215 1.0000
8.500 1.1508 0.02694 0.01859 -0.0397 0.0214 1.0000
8.750 1.1612 0.02867 0.02038 -0.0370 0.0214 1.0000
9.000 1.1763 0.03082 0.02258 -0.0350 0.0215 1.0000
9.250 1.1987 0.03304 0.02486 -0.0341 0.0221 1.0000
9.500 1.2181 0.03418 0.02618 -0.0325 0.0229 1.0000
9.750 1.2423 0.03597 0.02835 -0.0309 0.0258 1.0000
10.000 1.2828 0.04198 0.03457 -0.0324 0.0315 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 5 AIRFOIL (usa5-il)