USA 45 M AIRFOIL (usa45m-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 45 M AIRFOIL (usa45m-il) Reynolds number: 100,000 Max Cl/Cd: 47.96 at α=9° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa45m-il-100000-n5.txt Download as CSV file: xf-usa45m-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 45 M 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.3941 0.11197 0.10691 -0.0258 1.0000 0.0643
-10.000 -0.3995 0.10846 0.10347 -0.0291 1.0000 0.0646
-9.750 -0.4042 0.10466 0.09973 -0.0322 1.0000 0.0647
-9.500 -0.3937 0.09973 0.09482 -0.0307 1.0000 0.0655
-9.250 -0.3795 0.09651 0.09160 -0.0285 1.0000 0.0667
-9.000 -0.3716 0.09363 0.08874 -0.0280 1.0000 0.0682
-8.750 -0.3674 0.09079 0.08594 -0.0282 1.0000 0.0704
-8.500 -0.3686 0.08764 0.08285 -0.0294 1.0000 0.0728
-8.250 -0.3789 0.08412 0.07943 -0.0321 1.0000 0.0756
-8.000 -0.4014 0.08030 0.07570 -0.0357 1.0000 0.0769
-7.750 -0.4319 0.07665 0.07190 -0.0392 1.0000 0.0781
-7.500 -0.4446 0.07374 0.06883 -0.0387 1.0000 0.0783
-7.250 -0.4272 0.06886 0.06426 -0.0369 1.0000 0.0799
-7.000 -0.4245 0.06611 0.06154 -0.0352 1.0000 0.0807
-6.750 -0.4417 0.05695 0.05175 -0.0348 1.0000 0.0518
-6.500 -0.4309 0.05346 0.04816 -0.0349 0.9964 0.0514
-6.000 -0.3776 0.04498 0.03883 -0.0411 0.9736 0.0518
-5.750 -0.3501 0.04134 0.03495 -0.0430 0.9621 0.0513
-5.500 -0.3218 0.03815 0.03141 -0.0446 0.9499 0.0509
-5.250 -0.2926 0.03533 0.02821 -0.0459 0.9369 0.0508
-5.000 -0.2628 0.03318 0.02561 -0.0467 0.9226 0.0516
-4.750 -0.2326 0.03151 0.02347 -0.0473 0.9068 0.0525
-4.500 -0.2025 0.02941 0.02104 -0.0479 0.8902 0.0527
-4.250 -0.1717 0.02767 0.01899 -0.0484 0.8724 0.0527
-4.000 -0.1407 0.02598 0.01701 -0.0488 0.8529 0.0529
-3.750 -0.1116 0.02450 0.01529 -0.0488 0.8303 0.0532
-3.500 -0.0819 0.02323 0.01380 -0.0488 0.8065 0.0537
-3.250 -0.0538 0.02219 0.01258 -0.0486 0.7802 0.0546
-3.000 -0.0262 0.02139 0.01161 -0.0482 0.7525 0.0564
-2.750 0.0010 0.02071 0.01075 -0.0477 0.7244 0.0584
-2.500 0.0277 0.02006 0.00990 -0.0470 0.6963 0.0594
-2.250 0.0538 0.01946 0.00912 -0.0463 0.6688 0.0602
-2.000 0.0794 0.01894 0.00845 -0.0455 0.6431 0.0612
-1.750 0.1043 0.01851 0.00787 -0.0446 0.6192 0.0623
-1.500 0.1288 0.01816 0.00737 -0.0437 0.5975 0.0636
-1.250 0.1530 0.01787 0.00694 -0.0427 0.5778 0.0649
-1.000 0.1764 0.01748 0.00647 -0.0418 0.5604 0.0686
-0.750 0.2010 0.01729 0.00615 -0.0410 0.5448 0.0735
-0.500 0.2261 0.01716 0.00586 -0.0403 0.5309 0.0777
-0.250 0.2511 0.01697 0.00556 -0.0396 0.5184 0.0835
0.000 0.2764 0.01683 0.00533 -0.0389 0.5072 0.0953
0.250 0.2903 0.01460 0.00535 -0.0367 0.4979 0.6745
0.750 0.4429 0.01472 0.00545 -0.0545 0.4710 1.0000
1.000 0.4652 0.01489 0.00546 -0.0535 0.4631 1.0000
1.250 0.4877 0.01509 0.00551 -0.0525 0.4560 1.0000
1.500 0.5103 0.01528 0.00559 -0.0515 0.4488 1.0000
1.750 0.5330 0.01551 0.00567 -0.0505 0.4425 1.0000
2.000 0.5559 0.01572 0.00581 -0.0496 0.4356 1.0000
2.250 0.5789 0.01596 0.00593 -0.0487 0.4299 1.0000
2.500 0.6020 0.01621 0.00612 -0.0478 0.4238 1.0000
2.750 0.6250 0.01645 0.00631 -0.0470 0.4177 1.0000
3.000 0.6483 0.01673 0.00647 -0.0461 0.4127 1.0000
3.250 0.6714 0.01699 0.00675 -0.0453 0.4066 1.0000
3.500 0.6947 0.01727 0.00698 -0.0444 0.4011 1.0000
3.750 0.7182 0.01757 0.00719 -0.0436 0.3964 1.0000
4.000 0.7410 0.01786 0.00754 -0.0428 0.3903 1.0000
4.250 0.7643 0.01815 0.00781 -0.0419 0.3849 1.0000
4.500 0.7878 0.01847 0.00807 -0.0411 0.3803 1.0000
4.750 0.8104 0.01878 0.00847 -0.0402 0.3742 1.0000
5.000 0.8335 0.01909 0.00877 -0.0394 0.3689 1.0000
5.250 0.8568 0.01942 0.00908 -0.0386 0.3640 1.0000
5.500 0.8791 0.01975 0.00952 -0.0377 0.3577 1.0000
5.750 0.9022 0.02006 0.00983 -0.0369 0.3525 1.0000
6.000 0.9249 0.02041 0.01023 -0.0361 0.3470 1.0000
6.250 0.9470 0.02075 0.01068 -0.0352 0.3408 1.0000
6.500 0.9701 0.02106 0.01097 -0.0343 0.3356 1.0000
6.750 0.9917 0.02145 0.01150 -0.0334 0.3294 1.0000
7.000 1.0138 0.02179 0.01191 -0.0325 0.3234 1.0000
7.250 1.0362 0.02213 0.01229 -0.0316 0.3182 1.0000
7.500 1.0571 0.02257 0.01290 -0.0306 0.3117 1.0000
7.750 1.0791 0.02292 0.01330 -0.0297 0.3063 1.0000
8.000 1.0998 0.02336 0.01388 -0.0287 0.3002 1.0000
8.250 1.1199 0.02369 0.01433 -0.0275 0.2927 1.0000
8.500 1.1382 0.02397 0.01473 -0.0261 0.2829 1.0000
8.750 1.1555 0.02417 0.01497 -0.0246 0.2716 1.0000
9.000 1.1722 0.02444 0.01531 -0.0230 0.2603 1.0000
9.250 1.1885 0.02487 0.01590 -0.0214 0.2484 1.0000
9.500 1.2032 0.02531 0.01643 -0.0197 0.2348 1.0000
9.750 1.2169 0.02586 0.01704 -0.0178 0.2209 1.0000
10.000 1.2275 0.02653 0.01772 -0.0157 0.2041 1.0000
10.250 1.2362 0.02736 0.01857 -0.0134 0.1872 1.0000
10.500 1.2386 0.02844 0.01960 -0.0104 0.1661 1.0000
10.750 1.2342 0.02994 0.02102 -0.0069 0.1421 1.0000
11.250 1.2183 0.03413 0.02506 -0.0014 0.0939 1.0000
11.500 1.1938 0.03804 0.02877 0.0005 0.0561 1.0000
11.750 1.1712 0.04252 0.03321 0.0009 0.0459 1.0000
12.000 1.1533 0.04721 0.03799 0.0003 0.0421 1.0000
12.250 1.1386 0.05200 0.04291 -0.0010 0.0402 1.0000
12.500 1.1220 0.05736 0.04840 -0.0027 0.0385 1.0000
13.000 1.0929 0.06804 0.05936 -0.0062 0.0362 1.0000
13.250 1.0801 0.07325 0.06472 -0.0079 0.0353 1.0000
13.500 1.0671 0.07856 0.07017 -0.0097 0.0345 1.0000
13.750 1.0542 0.08401 0.07575 -0.0116 0.0337 1.0000
14.000 1.0426 0.08942 0.08130 -0.0136 0.0331 1.0000
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Polar data table (+)
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