USA 25 AIRFOIL (usa25-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 25 AIRFOIL (usa25-il) Reynolds number: 200,000 Max Cl/Cd: 74.08 at α=8.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa25-il-200000.txt Download as CSV file: xf-usa25-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: USA 25 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
------ -------- --------- --------- -------- -------- --------
-7.250 -0.3614 0.09269 0.08959 -0.0164 1.0000 0.0466
-7.000 -0.3731 0.09115 0.08811 -0.0143 1.0000 0.0472
-6.750 -0.3804 0.08920 0.08620 -0.0131 1.0000 0.0484
-6.500 -0.3855 0.08706 0.08409 -0.0129 0.9998 0.0500
-6.250 -0.3430 0.08064 0.07750 -0.0318 0.9901 0.0531
-6.000 -0.3246 0.07659 0.07349 -0.0304 0.9868 0.0545
-5.750 -0.2959 0.07316 0.07003 -0.0342 0.9832 0.0573
-5.250 -0.2271 0.06409 0.06065 -0.0503 0.9704 0.0646
-5.000 -0.2097 0.06102 0.05760 -0.0502 0.9629 0.0660
-4.750 -0.1770 0.05778 0.05429 -0.0539 0.9584 0.0691
-4.500 -0.1333 0.05462 0.05071 -0.0614 0.9491 0.0763
-4.250 -0.1083 0.05039 0.04653 -0.0631 0.9446 0.0776
-4.000 -0.0826 0.04758 0.04370 -0.0643 0.9380 0.0794
-3.750 -0.0514 0.04486 0.04089 -0.0664 0.9311 0.0827
-3.500 -0.0069 0.04235 0.03784 -0.0704 0.9234 0.0899
-3.250 0.0177 0.03868 0.03425 -0.0715 0.9165 0.0916
-3.000 0.0460 0.03638 0.03193 -0.0723 0.9082 0.0941
-2.750 0.0813 0.03391 0.02930 -0.0741 0.9006 0.0976
-2.500 0.1111 0.02736 0.02201 -0.0738 0.8896 0.0776
-2.250 0.1421 0.01984 0.01349 -0.0735 0.8821 0.0606
-2.000 0.1808 0.01752 0.01071 -0.0750 0.8696 0.0615
-1.750 0.2184 0.01589 0.00874 -0.0763 0.8500 0.0621
-1.500 0.2685 0.01443 0.00696 -0.0801 0.8227 0.0630
-1.250 0.3124 0.01337 0.00565 -0.0828 0.7735 0.0654
-1.000 0.3422 0.01314 0.00508 -0.0826 0.6964 0.0688
-0.750 0.3633 0.01317 0.00467 -0.0807 0.6330 0.0735
-0.500 0.3842 0.01315 0.00443 -0.0790 0.5989 0.0812
-0.250 0.4072 0.01321 0.00435 -0.0778 0.5781 0.1018
0.000 0.4319 0.01344 0.00450 -0.0771 0.5635 0.1262
0.250 0.4572 0.01369 0.00461 -0.0765 0.5518 0.1406
0.500 0.4828 0.01387 0.00469 -0.0760 0.5419 0.1520
0.750 0.5086 0.01404 0.00474 -0.0755 0.5341 0.1603
1.000 0.5341 0.01408 0.00479 -0.0750 0.5269 0.1689
1.250 0.5604 0.01427 0.00491 -0.0748 0.5209 0.1823
1.500 0.5862 0.01436 0.00499 -0.0743 0.5151 0.1950
1.750 0.6116 0.01442 0.00508 -0.0739 0.5093 0.2099
2.000 0.6377 0.01450 0.00516 -0.0736 0.5047 0.2268
2.250 0.6623 0.01449 0.00523 -0.0729 0.5003 0.2443
2.500 0.6863 0.01449 0.00528 -0.0721 0.4953 0.2661
2.750 0.7105 0.01450 0.00534 -0.0714 0.4904 0.2967
3.000 0.8403 0.01358 0.00583 -0.0931 0.4787 1.0000
3.250 0.8639 0.01387 0.00598 -0.0922 0.4736 1.0000
3.500 0.8843 0.01404 0.00617 -0.0908 0.4678 1.0000
3.750 0.9059 0.01425 0.00633 -0.0895 0.4619 1.0000
4.000 0.9295 0.01457 0.00657 -0.0887 0.4571 1.0000
4.250 0.9501 0.01475 0.00683 -0.0873 0.4527 1.0000
4.500 0.9721 0.01499 0.00708 -0.0861 0.4482 1.0000
4.750 0.9955 0.01528 0.00733 -0.0853 0.4441 1.0000
5.000 1.0175 0.01558 0.00766 -0.0842 0.4397 1.0000
5.250 1.0379 0.01579 0.00795 -0.0827 0.4345 1.0000
5.500 1.0604 0.01605 0.00820 -0.0817 0.4294 1.0000
5.750 1.0829 0.01638 0.00854 -0.0807 0.4246 1.0000
6.000 1.1017 0.01655 0.00883 -0.0789 0.4181 1.0000
6.250 1.1223 0.01672 0.00894 -0.0774 0.4095 1.0000
6.500 1.1380 0.01670 0.00899 -0.0749 0.3992 1.0000
6.750 1.1556 0.01679 0.00914 -0.0729 0.3903 1.0000
7.000 1.1736 0.01684 0.00920 -0.0709 0.3814 1.0000
7.250 1.1887 0.01684 0.00932 -0.0683 0.3708 1.0000
7.500 1.2040 0.01682 0.00935 -0.0658 0.3596 1.0000
7.750 1.2196 0.01684 0.00940 -0.0633 0.3486 1.0000
8.000 1.2342 0.01686 0.00949 -0.0607 0.3350 1.0000
8.250 1.2486 0.01690 0.00963 -0.0581 0.3169 1.0000
8.500 1.2624 0.01704 0.00978 -0.0554 0.2946 1.0000
8.750 1.2745 0.01736 0.01003 -0.0525 0.2683 1.0000
9.000 1.2829 0.01793 0.01043 -0.0491 0.2375 1.0000
9.250 1.2892 0.01865 0.01100 -0.0454 0.2101 1.0000
9.500 1.2952 0.01947 0.01169 -0.0418 0.1858 1.0000
9.750 1.2997 0.02044 0.01254 -0.0381 0.1598 1.0000
10.000 1.3052 0.02143 0.01342 -0.0348 0.1300 1.0000
10.250 1.2962 0.02327 0.01482 -0.0297 0.0743 1.0000
10.500 1.2916 0.02495 0.01634 -0.0254 0.0490 1.0000
10.750 1.2914 0.02642 0.01777 -0.0218 0.0355 1.0000
11.000 1.2939 0.02779 0.01917 -0.0189 0.0313 1.0000
11.250 1.2962 0.02923 0.02070 -0.0161 0.0291 1.0000
11.500 1.2971 0.03083 0.02241 -0.0135 0.0278 1.0000
11.750 1.2948 0.03279 0.02451 -0.0110 0.0267 1.0000
12.000 1.2924 0.03488 0.02674 -0.0089 0.0262 1.0000
12.250 1.2892 0.03721 0.02923 -0.0072 0.0255 1.0000
12.500 1.2844 0.03989 0.03206 -0.0059 0.0248 1.0000
12.750 1.2787 0.04289 0.03521 -0.0050 0.0243 1.0000
13.000 1.2719 0.04626 0.03872 -0.0046 0.0241 1.0000
13.250 1.2642 0.04997 0.04256 -0.0047 0.0235 1.0000
13.500 1.2569 0.05384 0.04655 -0.0051 0.0232 1.0000
13.750 1.2500 0.05777 0.05059 -0.0056 0.0230 1.0000
14.000 1.2436 0.06167 0.05459 -0.0061 0.0228 1.0000
14.250 1.2394 0.06525 0.05827 -0.0065 0.0227 1.0000
14.500 1.2345 0.06890 0.06199 -0.0069 0.0224 1.0000
14.750 1.2333 0.07202 0.06520 -0.0070 0.0223 1.0000
15.000 1.2328 0.07499 0.06822 -0.0069 0.0220 1.0000
15.250 1.2352 0.07749 0.07078 -0.0065 0.0218 1.0000
15.500 1.2401 0.07963 0.07296 -0.0057 0.0216 1.0000
15.750 1.2475 0.08146 0.07486 -0.0048 0.0215 1.0000
16.000 1.2554 0.08338 0.07687 -0.0039 0.0215 1.0000
16.250 1.2625 0.08552 0.07912 -0.0031 0.0215 1.0000
16.500 1.2673 0.08812 0.08184 -0.0029 0.0214 1.0000
16.750 1.2703 0.09112 0.08500 -0.0028 0.0213 1.0000
17.000 1.2726 0.09439 0.08842 -0.0023 0.0209 1.0000
17.250 1.2688 0.09855 0.09276 -0.0036 0.0209 1.0000
17.500 1.2615 0.10338 0.09779 -0.0051 0.0208 1.0000
17.750 1.2543 0.10817 0.10276 -0.0072 0.0209 1.0000
18.000 1.2451 0.11339 0.10816 -0.0097 0.0210 1.0000
18.250 1.2340 0.11909 0.11407 -0.0130 0.0213 1.0000
18.500 1.2198 0.12557 0.12078 -0.0167 0.0215 1.0000
18.750 1.2011 0.13331 0.12877 -0.0213 0.0218 1.0000
19.000 1.1787 0.14230 0.13804 -0.0269 0.0223 1.0000
19.250 1.1504 0.15322 0.14921 -0.0341 0.0226 1.0000
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