PT40 (pt40-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: PT40 (pt40-il) Reynolds number: 200,000 Max Cl/Cd: 64.3 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-pt40-il-200000.txt Download as CSV file: xf-pt40-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: PT40
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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3274 0.10088 0.09670 -0.0275 1.0000 0.0724
-8.250 -0.3320 0.09784 0.09370 -0.0294 1.0000 0.0744
-8.000 -0.3677 0.09357 0.08952 -0.0346 1.0000 0.0755
-7.750 -0.3505 0.09118 0.08716 -0.0320 1.0000 0.0759
-7.500 -0.3422 0.08902 0.08504 -0.0302 1.0000 0.0764
-7.250 -0.3408 0.08696 0.08304 -0.0287 1.0000 0.0771
-7.000 -0.3466 0.08505 0.08120 -0.0270 1.0000 0.0777
-6.750 -0.3613 0.08345 0.07969 -0.0243 1.0000 0.0783
-6.500 -0.3772 0.08166 0.07799 -0.0223 1.0000 0.0790
-6.250 -0.4232 0.07352 0.06959 -0.0366 0.9971 0.0828
-6.000 -0.3960 0.07071 0.06689 -0.0367 0.9944 0.0832
-5.750 -0.3689 0.06820 0.06444 -0.0378 0.9908 0.0839
-5.500 -0.3400 0.06547 0.06171 -0.0406 0.9860 0.0851
-5.250 -0.3068 0.05749 0.05327 -0.0544 0.9769 0.0916
-5.000 -0.2763 0.05497 0.05084 -0.0562 0.9729 0.0924
-4.750 -0.2442 0.05249 0.04837 -0.0590 0.9686 0.0940
-4.500 -0.2108 0.04740 0.04276 -0.0661 0.9600 0.1011
-4.250 -0.1756 0.04484 0.04031 -0.0689 0.9573 0.1023
-4.000 -0.1479 0.04293 0.03841 -0.0701 0.9489 0.1046
-3.750 -0.1093 0.03951 0.03463 -0.0748 0.9445 0.1126
-3.500 -0.0723 0.03747 0.03264 -0.0776 0.9410 0.1153
-3.250 -0.0415 0.03524 0.03005 -0.0795 0.9321 0.1244
-3.000 -0.0051 0.03317 0.02806 -0.0818 0.9283 0.1273
-2.750 0.0249 0.03166 0.02625 -0.0828 0.9193 0.1382
-2.500 0.0580 0.02983 0.02452 -0.0842 0.9132 0.1420
-2.250 0.0888 0.02843 0.02297 -0.0852 0.9053 0.1550
-2.000 0.1189 0.02734 0.02189 -0.0855 0.8963 0.1628
-1.750 0.1605 0.02210 0.01511 -0.0851 0.8891 0.0997
-1.500 0.1889 0.02077 0.01368 -0.0848 0.8784 0.0991
-1.250 0.2188 0.01960 0.01238 -0.0846 0.8681 0.0985
-1.000 0.2484 0.01854 0.01117 -0.0840 0.8551 0.0979
-0.750 0.2752 0.01772 0.01020 -0.0829 0.8383 0.0979
-0.500 0.3022 0.01710 0.00943 -0.0818 0.8210 0.0989
-0.250 0.3298 0.01675 0.00890 -0.0808 0.8052 0.1003
0.000 0.3577 0.01588 0.00800 -0.0802 0.7920 0.1017
0.250 0.3837 0.01540 0.00754 -0.0793 0.7767 0.1031
0.500 0.4106 0.01502 0.00716 -0.0785 0.7623 0.1050
0.750 0.4379 0.01473 0.00681 -0.0778 0.7482 0.1079
1.000 0.4646 0.01452 0.00655 -0.0769 0.7327 0.1104
1.250 0.4903 0.01405 0.00615 -0.0760 0.7168 0.1130
1.500 0.5166 0.01378 0.00592 -0.0752 0.7010 0.1164
1.750 0.5430 0.01362 0.00573 -0.0743 0.6847 0.1211
2.000 0.5688 0.01335 0.00548 -0.0734 0.6673 0.1261
2.250 0.5945 0.01320 0.00534 -0.0725 0.6483 0.1323
2.500 0.6201 0.01305 0.00520 -0.0716 0.6280 0.1419
2.750 0.6456 0.01293 0.00508 -0.0707 0.6063 0.1603
3.000 0.6799 0.01085 0.00517 -0.0713 0.5802 1.0000
3.250 0.7036 0.01106 0.00515 -0.0699 0.5519 1.0000
3.500 0.7265 0.01131 0.00519 -0.0685 0.5193 1.0000
3.750 0.7485 0.01164 0.00529 -0.0669 0.4797 1.0000
4.000 0.7688 0.01209 0.00544 -0.0652 0.4293 1.0000
4.250 0.7878 0.01268 0.00568 -0.0633 0.3728 1.0000
4.500 0.8070 0.01335 0.00600 -0.0616 0.3297 1.0000
4.750 0.8279 0.01397 0.00637 -0.0602 0.3004 1.0000
5.000 0.8494 0.01455 0.00676 -0.0589 0.2794 1.0000
5.250 0.8719 0.01507 0.00713 -0.0578 0.2626 1.0000
5.500 0.8949 0.01555 0.00752 -0.0568 0.2488 1.0000
5.750 0.9173 0.01609 0.00792 -0.0557 0.2372 1.0000
6.000 0.9407 0.01651 0.00828 -0.0548 0.2262 1.0000
6.250 0.9639 0.01696 0.00868 -0.0539 0.2164 1.0000
6.500 0.9867 0.01743 0.00907 -0.0529 0.2078 1.0000
6.750 1.0100 0.01786 0.00949 -0.0519 0.2000 1.0000
7.000 1.0327 0.01830 0.00989 -0.0510 0.1923 1.0000
7.250 1.0555 0.01874 0.01032 -0.0500 0.1847 1.0000
7.500 1.0777 0.01915 0.01070 -0.0489 0.1769 1.0000
7.750 1.0998 0.01958 0.01115 -0.0479 0.1690 1.0000
8.000 1.1207 0.02003 0.01155 -0.0467 0.1608 1.0000
8.250 1.1424 0.02038 0.01196 -0.0456 0.1511 1.0000
8.500 1.1626 0.02081 0.01239 -0.0443 0.1396 1.0000
8.750 1.1811 0.02134 0.01286 -0.0428 0.1250 1.0000
9.000 1.1978 0.02205 0.01349 -0.0411 0.1083 1.0000
9.250 1.2124 0.02292 0.01424 -0.0391 0.0961 1.0000
9.500 1.2253 0.02380 0.01508 -0.0368 0.0885 1.0000
9.750 1.2374 0.02475 0.01599 -0.0345 0.0835 1.0000
10.000 1.2496 0.02571 0.01691 -0.0323 0.0799 1.0000
10.250 1.2615 0.02675 0.01796 -0.0301 0.0770 1.0000
10.500 1.2741 0.02775 0.01899 -0.0281 0.0745 1.0000
10.750 1.2851 0.02889 0.02008 -0.0261 0.0726 1.0000
11.000 1.2971 0.03006 0.02128 -0.0243 0.0708 1.0000
11.250 1.3097 0.03118 0.02245 -0.0226 0.0690 1.0000
11.500 1.3218 0.03235 0.02364 -0.0210 0.0675 1.0000
11.750 1.3337 0.03362 0.02486 -0.0194 0.0663 1.0000
12.000 1.3471 0.03497 0.02620 -0.0180 0.0652 1.0000
12.250 1.3595 0.03627 0.02761 -0.0165 0.0643 1.0000
12.500 1.3720 0.03762 0.02904 -0.0152 0.0633 1.0000
12.750 1.3845 0.03899 0.03048 -0.0139 0.0624 1.0000
13.000 1.3975 0.04038 0.03190 -0.0127 0.0616 1.0000
13.250 1.4113 0.04177 0.03330 -0.0115 0.0609 1.0000
13.500 1.4280 0.04314 0.03464 -0.0106 0.0602 1.0000
13.750 1.4463 0.04470 0.03618 -0.0098 0.0595 1.0000
14.000 1.4510 0.04659 0.03827 -0.0083 0.0591 1.0000
14.250 1.4549 0.04863 0.04049 -0.0070 0.0585 1.0000
14.500 1.4583 0.05079 0.04283 -0.0059 0.0580 1.0000
14.750 1.4609 0.05308 0.04528 -0.0048 0.0575 1.0000
15.000 1.4627 0.05550 0.04786 -0.0040 0.0571 1.0000
15.250 1.4635 0.05809 0.05060 -0.0032 0.0567 1.0000
15.500 1.4629 0.06088 0.05354 -0.0027 0.0564 1.0000
15.750 1.4605 0.06389 0.05672 -0.0023 0.0562 1.0000
16.000 1.4555 0.06724 0.06024 -0.0022 0.0559 1.0000
16.250 1.4476 0.07099 0.06417 -0.0024 0.0558 1.0000
16.500 1.4362 0.07524 0.06862 -0.0030 0.0556 1.0000
16.750 1.4201 0.08024 0.07384 -0.0042 0.0556 1.0000
17.000 1.3951 0.08667 0.08054 -0.0067 0.0556 1.0000
17.250 1.3451 0.09753 0.09184 -0.0126 0.0560 1.0000
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Polar data table (+)
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