NACA M14 AIRFOIL (m14-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M14 AIRFOIL (m14-il) Reynolds number: 200,000 Max Cl/Cd: 75.43 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m14-il-200000-n5.txt Download as CSV file: xf-m14-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M14 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4358 0.11723 0.11369 -0.0065 1.0000 0.0247
-9.500 -0.4326 0.11372 0.11021 -0.0089 1.0000 0.0248
-9.250 -0.4302 0.11024 0.10676 -0.0116 1.0000 0.0250
-9.000 -0.4267 0.10659 0.10314 -0.0143 1.0000 0.0251
-8.750 -0.4228 0.10281 0.09939 -0.0167 1.0000 0.0252
-8.500 -0.4190 0.09893 0.09554 -0.0188 1.0000 0.0252
-8.250 -0.4157 0.09504 0.09168 -0.0210 1.0000 0.0252
-8.000 -0.4125 0.09113 0.08780 -0.0230 1.0000 0.0251
-7.750 -0.4071 0.08727 0.08396 -0.0231 1.0000 0.0247
-7.250 -0.3936 0.07729 0.07399 -0.0339 1.0000 0.0257
-7.000 -0.3846 0.07697 0.07372 -0.0288 1.0000 0.0281
-6.750 -0.3743 0.07394 0.07070 -0.0312 1.0000 0.0306
-6.000 -0.2553 0.04123 0.03782 -0.0537 0.9125 0.0359
-5.500 -0.2530 0.04898 0.04501 -0.0574 0.9274 0.0330
-5.250 -0.2291 0.04468 0.04048 -0.0596 0.9103 0.0325
-5.000 -0.2047 0.04024 0.03572 -0.0616 0.8938 0.0361
-4.750 -0.1776 0.03532 0.03027 -0.0623 0.8784 0.0348
-4.500 -0.1538 0.03265 0.02726 -0.0621 0.8631 0.0346
-4.250 -0.1324 0.02903 0.02323 -0.0620 0.8488 0.0349
-4.000 -0.1103 0.02676 0.02071 -0.0619 0.8347 0.0362
-3.750 -0.0857 0.02443 0.01797 -0.0615 0.8214 0.0363
-3.500 -0.0602 0.02229 0.01539 -0.0610 0.8088 0.0361
-3.250 -0.0344 0.02068 0.01342 -0.0606 0.7966 0.0362
-3.000 -0.0082 0.01937 0.01178 -0.0602 0.7850 0.0365
-2.500 0.0450 0.01749 0.00943 -0.0596 0.7622 0.0378
-2.250 0.0720 0.01682 0.00858 -0.0594 0.7514 0.0390
-2.000 0.0991 0.01605 0.00758 -0.0591 0.7411 0.0391
-1.750 0.1263 0.01535 0.00671 -0.0588 0.7308 0.0391
-1.500 0.1536 0.01474 0.00597 -0.0585 0.7205 0.0391
-1.250 0.1806 0.01423 0.00534 -0.0582 0.7107 0.0394
-1.000 0.2076 0.01379 0.00481 -0.0579 0.7009 0.0398
-0.750 0.2347 0.01340 0.00438 -0.0577 0.6907 0.0403
-0.500 0.2617 0.01309 0.00400 -0.0575 0.6811 0.0411
-0.250 0.2887 0.01283 0.00368 -0.0573 0.6714 0.0420
0.000 0.3159 0.01262 0.00343 -0.0571 0.6614 0.0431
0.250 0.3432 0.01247 0.00321 -0.0569 0.6523 0.0442
0.500 0.3705 0.01235 0.00305 -0.0568 0.6428 0.0466
0.750 0.3978 0.01222 0.00291 -0.0567 0.6336 0.0509
1.000 0.4252 0.01216 0.00281 -0.0566 0.6250 0.0542
1.250 0.4527 0.01208 0.00275 -0.0565 0.6156 0.0625
1.500 0.4794 0.01183 0.00277 -0.0564 0.6070 0.1592
2.000 0.5487 0.01001 0.00287 -0.0594 0.5886 1.0000
2.500 0.6015 0.01025 0.00297 -0.0588 0.5710 1.0000
2.750 0.6279 0.01038 0.00304 -0.0585 0.5624 1.0000
3.000 0.6543 0.01052 0.00313 -0.0582 0.5535 1.0000
3.250 0.6809 0.01065 0.00326 -0.0580 0.5445 1.0000
3.500 0.7073 0.01081 0.00337 -0.0577 0.5362 1.0000
3.750 0.7339 0.01096 0.00353 -0.0575 0.5267 1.0000
4.000 0.7603 0.01112 0.00370 -0.0573 0.5178 1.0000
4.250 0.7867 0.01130 0.00387 -0.0571 0.5090 1.0000
4.500 0.8132 0.01147 0.00408 -0.0569 0.4995 1.0000
4.750 0.8395 0.01166 0.00428 -0.0566 0.4904 1.0000
5.000 0.8658 0.01185 0.00453 -0.0564 0.4808 1.0000
5.250 0.8920 0.01205 0.00479 -0.0562 0.4707 1.0000
5.500 0.9177 0.01226 0.00502 -0.0558 0.4557 1.0000
5.750 0.9423 0.01251 0.00524 -0.0553 0.4325 1.0000
6.000 0.9663 0.01281 0.00547 -0.0547 0.4010 1.0000
6.250 0.9890 0.01326 0.00576 -0.0540 0.3613 1.0000
6.500 1.0109 0.01383 0.00616 -0.0533 0.3149 1.0000
6.750 1.0313 0.01459 0.00670 -0.0524 0.2575 1.0000
7.000 1.0443 0.01625 0.00770 -0.0509 0.1455 1.0000
7.250 1.0548 0.01822 0.00906 -0.0490 0.0528 1.0000
7.500 1.0723 0.01933 0.01009 -0.0477 0.0326 1.0000
7.750 1.0909 0.02024 0.01105 -0.0466 0.0261 1.0000
8.000 1.1095 0.02110 0.01201 -0.0454 0.0226 1.0000
8.250 1.1272 0.02200 0.01305 -0.0441 0.0201 1.0000
8.500 1.1421 0.02311 0.01426 -0.0425 0.0185 1.0000
8.750 1.1515 0.02462 0.01589 -0.0403 0.0173 1.0000
9.000 1.1637 0.02578 0.01718 -0.0384 0.0166 1.0000
9.250 1.1752 0.02693 0.01845 -0.0364 0.0156 1.0000
9.500 1.1834 0.02813 0.01975 -0.0341 0.0147 1.0000
9.750 1.1896 0.02952 0.02127 -0.0316 0.0140 1.0000
10.000 1.1957 0.03107 0.02291 -0.0295 0.0135 1.0000
10.250 1.2018 0.03273 0.02467 -0.0276 0.0131 1.0000
10.500 1.2080 0.03453 0.02656 -0.0259 0.0127 1.0000
10.750 1.2140 0.03647 0.02859 -0.0243 0.0123 1.0000
11.000 1.2201 0.03864 0.03085 -0.0227 0.0119 1.0000
11.250 1.2271 0.04121 0.03353 -0.0209 0.0115 1.0000
11.500 1.2351 0.04303 0.03554 -0.0198 0.0111 1.0000
11.750 1.2420 0.04505 0.03776 -0.0188 0.0107 1.0000
12.000 1.2475 0.04732 0.04023 -0.0177 0.0103 1.0000
12.250 1.2512 0.04990 0.04305 -0.0168 0.0100 1.0000
12.500 1.2529 0.05276 0.04613 -0.0159 0.0098 1.0000
12.750 1.2522 0.05590 0.04949 -0.0152 0.0096 1.0000
13.000 1.2493 0.05933 0.05315 -0.0149 0.0095 1.0000
13.250 1.2440 0.06309 0.05714 -0.0148 0.0094 1.0000
13.500 1.2365 0.06720 0.06148 -0.0151 0.0093 1.0000
13.750 1.2268 0.07169 0.06620 -0.0158 0.0092 1.0000
14.000 1.2151 0.07661 0.07134 -0.0169 0.0092 1.0000
14.250 1.2015 0.08200 0.07695 -0.0186 0.0091 1.0000
14.500 1.1862 0.08786 0.08303 -0.0207 0.0091 1.0000
14.750 1.1695 0.09435 0.08974 -0.0235 0.0091 1.0000
15.000 1.1517 0.10148 0.09708 -0.0270 0.0092 1.0000
15.250 1.1325 0.10935 0.10515 -0.0312 0.0092 1.0000
15.500 1.1123 0.11809 0.11408 -0.0363 0.0093 1.0000
15.750 1.0907 0.12789 0.12406 -0.0422 0.0094 1.0000
16.000 1.0673 0.13908 0.13543 -0.0491 0.0096 1.0000
16.250 1.0407 0.15271 0.14919 -0.0574 0.0098 1.0000
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Polar data table (+)
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