RAF 26 AIRFOIL (raf26-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 26 AIRFOIL (raf26-il) Reynolds number: 200,000 Max Cl/Cd: 71.28 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf26-il-200000.txt Download as CSV file: xf-raf26-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: RAF 26 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4960 0.10001 0.09635 -0.0167 1.0000 0.0384
-8.750 -0.4979 0.09650 0.09289 -0.0197 1.0000 0.0401
-8.500 -0.5009 0.09313 0.08959 -0.0230 1.0000 0.0408
-8.250 -0.4222 0.07576 0.07247 -0.0275 1.0000 0.0433
-8.000 -0.4192 0.07249 0.06922 -0.0267 1.0000 0.0444
-7.750 -0.4210 0.06903 0.06580 -0.0266 1.0000 0.0454
-7.500 -0.4265 0.06560 0.06241 -0.0268 1.0000 0.0463
-7.250 -0.5103 0.07042 0.06704 -0.0363 1.0000 0.0429
-7.000 -0.5046 0.06792 0.06457 -0.0344 1.0000 0.0441
-6.750 -0.4981 0.06463 0.06126 -0.0348 1.0000 0.0456
-6.500 -0.4905 0.06063 0.05721 -0.0365 1.0000 0.0473
-6.250 -0.4813 0.05605 0.05252 -0.0388 1.0000 0.0499
-6.000 -0.4659 0.05252 0.04824 -0.0426 1.0000 0.0542
-5.750 -0.4614 0.04565 0.04156 -0.0425 1.0000 0.0562
-5.500 -0.4485 0.04306 0.03897 -0.0415 1.0000 0.0582
-5.250 -0.4342 0.04024 0.03598 -0.0408 1.0000 0.0621
-5.000 -0.4206 0.03645 0.03174 -0.0404 1.0000 0.0696
-4.750 -0.3988 0.02906 0.02349 -0.0381 1.0000 0.0400
-4.500 -0.3819 0.02465 0.01860 -0.0365 1.0000 0.0354
-4.250 -0.3621 0.02266 0.01623 -0.0350 1.0000 0.0376
-4.000 -0.3408 0.02020 0.01329 -0.0334 1.0000 0.0367
-3.750 -0.3182 0.01818 0.01088 -0.0319 1.0000 0.0366
-3.500 -0.2953 0.01677 0.00921 -0.0306 1.0000 0.0379
-3.250 -0.2728 0.01578 0.00804 -0.0293 1.0000 0.0404
-3.000 -0.2515 0.01463 0.00686 -0.0281 1.0000 0.0465
-2.750 -0.2303 0.01395 0.00612 -0.0267 1.0000 0.0522
-2.500 -0.2099 0.01324 0.00543 -0.0254 1.0000 0.0640
-2.250 -0.1896 0.01265 0.00488 -0.0241 1.0000 0.0878
-2.000 -0.1705 0.01160 0.00449 -0.0230 1.0000 0.2181
-1.750 -0.1383 0.01072 0.00466 -0.0246 0.9955 0.4871
-1.500 -0.1072 0.01008 0.00478 -0.0252 0.9895 0.6879
-1.250 -0.0396 0.00944 0.00462 -0.0332 0.9935 1.0000
-1.000 0.0026 0.00956 0.00456 -0.0365 0.9872 1.0000
-0.750 0.0451 0.00970 0.00455 -0.0399 0.9818 1.0000
-0.500 0.0843 0.00975 0.00448 -0.0425 0.9748 1.0000
-0.250 0.1256 0.00981 0.00445 -0.0455 0.9691 1.0000
0.000 0.1653 0.00977 0.00436 -0.0481 0.9612 1.0000
0.250 0.2085 0.00963 0.00417 -0.0512 0.9517 1.0000
0.500 0.2563 0.00944 0.00397 -0.0552 0.9454 1.0000
0.750 0.2944 0.00928 0.00380 -0.0572 0.9358 1.0000
1.000 0.3392 0.00901 0.00355 -0.0604 0.9270 1.0000
1.250 0.3895 0.00869 0.00328 -0.0647 0.9172 1.0000
1.500 0.4353 0.00846 0.00308 -0.0680 0.9023 1.0000
1.750 0.4758 0.00834 0.00299 -0.0703 0.8832 1.0000
2.000 0.5122 0.00832 0.00301 -0.0717 0.8642 1.0000
2.250 0.5399 0.00838 0.00310 -0.0713 0.8415 1.0000
2.500 0.5665 0.00844 0.00315 -0.0705 0.8135 1.0000
2.750 0.5908 0.00852 0.00319 -0.0691 0.7774 1.0000
3.000 0.6138 0.00866 0.00328 -0.0675 0.7333 1.0000
3.250 0.6351 0.00891 0.00336 -0.0655 0.6761 1.0000
3.500 0.6527 0.00936 0.00350 -0.0628 0.5901 1.0000
3.750 0.6600 0.01045 0.00373 -0.0583 0.4177 1.0000
4.000 0.6647 0.01233 0.00438 -0.0544 0.1786 1.0000
4.250 0.6825 0.01329 0.00496 -0.0527 0.1130 1.0000
4.500 0.7014 0.01416 0.00576 -0.0510 0.0885 1.0000
4.750 0.7174 0.01550 0.00704 -0.0487 0.0582 1.0000
5.000 0.7376 0.01640 0.00788 -0.0471 0.0436 1.0000
5.250 0.7573 0.01763 0.00918 -0.0453 0.0381 1.0000
5.500 0.7787 0.01876 0.01035 -0.0438 0.0343 1.0000
5.750 0.8000 0.02063 0.01229 -0.0425 0.0305 1.0000
6.000 0.8244 0.02240 0.01427 -0.0414 0.0288 1.0000
6.250 0.8487 0.02479 0.01698 -0.0402 0.0281 1.0000
6.500 0.8711 0.02784 0.02043 -0.0386 0.0282 1.0000
6.750 0.8896 0.03160 0.02464 -0.0365 0.0293 1.0000
7.000 0.9042 0.03546 0.02895 -0.0340 0.0298 1.0000
7.250 0.9170 0.03884 0.03276 -0.0314 0.0295 1.0000
15.750 0.6963 0.18805 0.18490 -0.0628 0.0310 1.0000
16.000 0.6940 0.19040 0.18726 -0.0657 0.0308 1.0000
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