NACA/LANGLEY SYMMETRICAL (n0011sc-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA/LANGLEY SYMMETRICAL (n0011sc-il) Reynolds number: 200,000 Max Cl/Cd: 41.53 at α=2° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n0011sc-il-200000.txt Download as CSV file: xf-n0011sc-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA/LANGLEY SYMMETRICAL
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
------ -------- --------- --------- -------- -------- --------
-10.500 -0.6096 0.09501 0.09099 -0.0304 1.0000 0.0837
-10.250 -0.6495 0.08491 0.08094 -0.0386 1.0000 0.0859
-10.000 -0.6991 0.07782 0.07377 -0.0409 1.0000 0.0864
-9.750 -0.7396 0.07415 0.07001 -0.0371 1.0000 0.0866
-9.500 -0.8962 0.05622 0.05111 -0.0194 1.0000 0.0570
-9.250 -0.9588 0.04751 0.04144 -0.0068 1.0000 0.0499
-9.000 -0.9600 0.04444 0.03819 -0.0028 1.0000 0.0495
-8.750 -0.9607 0.04164 0.03517 0.0015 1.0000 0.0492
-8.500 -0.9607 0.03891 0.03214 0.0059 1.0000 0.0489
-8.250 -0.9578 0.03646 0.02937 0.0100 1.0000 0.0488
-8.000 -0.9522 0.03419 0.02677 0.0139 1.0000 0.0489
-7.750 -0.9434 0.03224 0.02449 0.0173 1.0000 0.0492
-7.500 -0.9321 0.03057 0.02251 0.0204 1.0000 0.0496
-7.250 -0.9196 0.02905 0.02068 0.0233 1.0000 0.0504
-7.000 -0.9008 0.02748 0.01910 0.0247 1.0000 0.0517
-6.750 -0.8831 0.02659 0.01816 0.0265 1.0000 0.0531
-6.500 -0.8647 0.02562 0.01708 0.0283 1.0000 0.0545
-6.250 -0.8455 0.02460 0.01589 0.0300 1.0000 0.0559
-6.000 -0.8260 0.02373 0.01483 0.0318 1.0000 0.0573
-5.750 -0.8045 0.02251 0.01358 0.0330 1.0000 0.0592
-5.500 -0.7853 0.02183 0.01295 0.0345 1.0000 0.0616
-5.250 -0.7664 0.02129 0.01235 0.0363 1.0000 0.0648
-5.000 -0.7467 0.02052 0.01152 0.0379 1.0000 0.0677
-4.750 -0.7284 0.01981 0.01090 0.0397 1.0000 0.0710
-4.500 -0.7100 0.01936 0.01042 0.0415 1.0000 0.0753
-4.250 -0.6929 0.01867 0.00978 0.0436 1.0000 0.0801
-4.000 -0.6754 0.01827 0.00939 0.0455 1.0000 0.0860
-3.750 -0.6592 0.01769 0.00888 0.0478 1.0000 0.0931
-3.500 -0.6425 0.01726 0.00848 0.0499 1.0000 0.1027
-3.250 -0.6259 0.01687 0.00812 0.0521 1.0000 0.1146
-3.000 -0.6109 0.01635 0.00776 0.0545 1.0000 0.1316
-2.750 -0.5958 0.01588 0.00746 0.0568 1.0000 0.1596
-2.500 -0.5823 0.01531 0.00722 0.0594 1.0000 0.2098
-2.000 -0.5684 0.01367 0.00687 0.0671 1.0000 0.4462
-1.750 -0.4075 0.01441 0.00958 0.0421 1.0000 0.9115
-1.500 -0.3624 0.01548 0.01057 0.0395 1.0000 0.9353
-1.250 -0.3022 0.01653 0.01152 0.0335 1.0000 0.9508
-1.000 -0.2417 0.01739 0.01230 0.0273 1.0000 0.9637
-0.750 -0.1792 0.01808 0.01293 0.0204 1.0000 0.9750
-0.500 -0.1126 0.01859 0.01340 0.0125 1.0000 0.9855
-0.250 -0.0361 0.01893 0.01370 0.0024 1.0000 0.9962
0.000 0.0000 0.01904 0.01381 0.0000 1.0000 1.0000
0.250 0.0369 0.01892 0.01370 -0.0026 0.9961 1.0000
0.500 0.1143 0.01858 0.01339 -0.0128 0.9853 1.0000
0.750 0.1792 0.01807 0.01292 -0.0204 0.9750 1.0000
1.000 0.2408 0.01740 0.01231 -0.0271 0.9638 1.0000
1.250 0.3025 0.01652 0.01151 -0.0336 0.9508 1.0000
1.500 0.3621 0.01548 0.01057 -0.0394 0.9354 1.0000
1.750 0.4075 0.01441 0.00958 -0.0422 0.9117 1.0000
2.000 0.5681 0.01368 0.00687 -0.0670 0.4450 1.0000
2.250 0.5716 0.01461 0.00703 -0.0625 0.2992 1.0000
2.500 0.5822 0.01530 0.00722 -0.0594 0.2088 1.0000
2.750 0.5958 0.01587 0.00746 -0.0569 0.1599 1.0000
3.000 0.6110 0.01633 0.00775 -0.0545 0.1323 1.0000
3.250 0.6259 0.01686 0.00811 -0.0521 0.1142 1.0000
3.500 0.6424 0.01726 0.00847 -0.0499 0.1025 1.0000
3.750 0.6592 0.01769 0.00888 -0.0478 0.0932 1.0000
4.000 0.6753 0.01825 0.00937 -0.0456 0.0860 1.0000
4.250 0.6928 0.01868 0.00978 -0.0435 0.0800 1.0000
4.500 0.7099 0.01934 0.01040 -0.0415 0.0751 1.0000
4.750 0.7283 0.01980 0.01090 -0.0397 0.0709 1.0000
5.000 0.7466 0.02051 0.01151 -0.0379 0.0677 1.0000
5.250 0.7661 0.02125 0.01232 -0.0362 0.0644 1.0000
5.500 0.7852 0.02181 0.01294 -0.0345 0.0617 1.0000
5.750 0.8044 0.02250 0.01358 -0.0330 0.0592 1.0000
6.000 0.8260 0.02374 0.01484 -0.0318 0.0573 1.0000
6.250 0.8453 0.02460 0.01589 -0.0300 0.0558 1.0000
6.500 0.8644 0.02559 0.01704 -0.0282 0.0543 1.0000
6.750 0.8830 0.02655 0.01810 -0.0265 0.0529 1.0000
7.000 0.9007 0.02749 0.01909 -0.0247 0.0516 1.0000
7.250 0.9197 0.02941 0.02102 -0.0235 0.0502 1.0000
7.500 0.9321 0.03065 0.02258 -0.0204 0.0497 1.0000
7.750 0.9430 0.03225 0.02451 -0.0172 0.0492 1.0000
8.000 0.9515 0.03427 0.02687 -0.0137 0.0488 1.0000
8.250 0.9574 0.03645 0.02937 -0.0099 0.0488 1.0000
8.500 0.9605 0.03895 0.03219 -0.0058 0.0490 1.0000
8.750 0.9603 0.04170 0.03523 -0.0014 0.0492 1.0000
9.000 0.9600 0.04451 0.03826 0.0028 0.0496 1.0000
9.250 0.9591 0.04755 0.04147 0.0067 0.0499 1.0000
9.500 0.9560 0.05082 0.04492 0.0108 0.0501 1.0000
9.750 0.6049 0.06679 0.06303 0.0424 0.0911 1.0000
10.000 0.5710 0.07327 0.06954 0.0409 0.0907 1.0000
10.250 0.5534 0.07919 0.07546 0.0392 0.0900 1.0000
10.500 0.6326 0.07761 0.07368 0.0465 0.0880 1.0000
10.750 0.5859 0.08381 0.07997 0.0455 0.0880 1.0000
11.000 0.4795 0.10145 0.09762 0.0298 0.0858 1.0000
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Polar data table (+)
Polar graphs
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