RAE 5214 AIRFOIL (rae5214-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAE 5214 AIRFOIL (rae5214-il) Reynolds number: 200,000 Max Cl/Cd: 46.43 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae5214-il-200000.txt Download as CSV file: xf-rae5214-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: RAE 5214 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5073 0.10529 0.10134 -0.0183 1.0000 0.0513
-10.000 -0.5223 0.09858 0.09469 -0.0252 1.0000 0.0536
-9.750 -0.5579 0.08780 0.08395 -0.0385 1.0000 0.0540
-9.500 -0.5273 0.08878 0.08497 -0.0297 1.0000 0.0550
-9.250 -0.5175 0.08676 0.08298 -0.0283 1.0000 0.0561
-9.000 -0.5165 0.08312 0.07936 -0.0297 1.0000 0.0573
-8.750 -0.5226 0.07805 0.07434 -0.0332 1.0000 0.0584
-8.500 -0.5430 0.07116 0.06748 -0.0395 1.0000 0.0593
-8.250 -0.5656 0.06600 0.06227 -0.0432 1.0000 0.0598
-7.750 -0.5994 0.05615 0.05192 -0.0460 1.0000 0.0632
-7.500 -0.5883 0.05355 0.04939 -0.0450 1.0000 0.0642
-7.250 -0.5804 0.05111 0.04693 -0.0440 1.0000 0.0662
-7.000 -0.5875 0.04790 0.04306 -0.0431 1.0000 0.0721
-6.750 -0.5742 0.04473 0.04006 -0.0420 1.0000 0.0735
-6.500 -0.5625 0.04266 0.03798 -0.0406 1.0000 0.0759
-6.250 -0.5554 0.04019 0.03516 -0.0394 1.0000 0.0840
-6.000 -0.5391 0.02941 0.02296 -0.0367 1.0000 0.0457
-5.750 -0.5202 0.02552 0.01814 -0.0348 1.0000 0.0412
-5.500 -0.5005 0.02346 0.01605 -0.0338 1.0000 0.0421
-5.250 -0.4805 0.02247 0.01507 -0.0328 1.0000 0.0439
-5.000 -0.4592 0.02121 0.01359 -0.0316 1.0000 0.0454
-4.750 -0.4371 0.01995 0.01212 -0.0304 1.0000 0.0462
-4.500 -0.4147 0.01894 0.01092 -0.0293 1.0000 0.0473
-4.250 -0.3932 0.01797 0.01007 -0.0285 1.0000 0.0500
-4.000 -0.3708 0.01732 0.00937 -0.0275 1.0000 0.0525
-3.750 -0.3482 0.01666 0.00864 -0.0266 1.0000 0.0543
-3.500 -0.3265 0.01588 0.00800 -0.0258 1.0000 0.0575
-3.250 -0.2890 0.01540 0.00749 -0.0279 0.9963 0.0619
-3.000 -0.2527 0.01466 0.00687 -0.0300 0.9926 0.0679
-2.750 -0.2159 0.01407 0.00634 -0.0321 0.9882 0.0766
-2.500 -0.1771 0.01331 0.00578 -0.0347 0.9849 0.1120
-2.250 -0.1502 0.01137 0.00585 -0.0358 0.9803 0.5845
-2.000 -0.1135 0.01166 0.00620 -0.0372 0.9755 0.6322
-1.750 -0.0803 0.01190 0.00643 -0.0379 0.9694 0.6607
-1.500 -0.0453 0.01222 0.00683 -0.0387 0.9641 0.6943
-1.250 -0.0150 0.01251 0.00723 -0.0383 0.9570 0.7261
-1.000 0.0214 0.01261 0.00740 -0.0392 0.9512 0.7473
-0.750 0.0592 0.01251 0.00731 -0.0407 0.9433 0.7564
-0.500 0.1092 0.01227 0.00701 -0.0447 0.9364 0.7630
-0.250 0.1516 0.01195 0.00672 -0.0469 0.9267 0.7683
0.000 0.1998 0.01149 0.00624 -0.0499 0.9164 0.7743
0.250 0.2367 0.01111 0.00584 -0.0507 0.9024 0.7803
0.500 0.2661 0.01081 0.00556 -0.0500 0.8867 0.7856
0.750 0.2944 0.01061 0.00536 -0.0494 0.8710 0.7919
1.000 0.3224 0.01044 0.00518 -0.0488 0.8548 0.7980
1.250 0.3489 0.01026 0.00503 -0.0477 0.8366 0.8038
1.500 0.3760 0.01012 0.00487 -0.0469 0.8155 0.8104
1.750 0.4019 0.00999 0.00471 -0.0458 0.7890 0.8165
2.000 0.4269 0.00989 0.00456 -0.0444 0.7546 0.8232
2.250 0.4518 0.00990 0.00442 -0.0431 0.7015 0.8306
2.500 0.4713 0.01015 0.00424 -0.0406 0.5854 0.8374
2.750 0.4812 0.01155 0.00442 -0.0373 0.3424 0.8462
3.250 0.5203 0.01320 0.00521 -0.0347 0.1953 0.8648
3.500 0.5422 0.01364 0.00556 -0.0335 0.1771 0.8761
3.750 0.5654 0.01397 0.00589 -0.0325 0.1642 0.8899
4.000 0.5890 0.01438 0.00628 -0.0316 0.1538 0.9083
4.250 0.6170 0.01478 0.00668 -0.0316 0.1443 0.9386
4.500 0.6528 0.01528 0.00719 -0.0336 0.1349 1.0000
4.750 0.6809 0.01607 0.00780 -0.0341 0.1270 1.0000
5.000 0.7098 0.01643 0.00821 -0.0346 0.1194 1.0000
5.250 0.7366 0.01719 0.00883 -0.0347 0.1125 1.0000
5.500 0.7638 0.01760 0.00928 -0.0347 0.1057 1.0000
5.750 0.7898 0.01848 0.01004 -0.0347 0.0998 1.0000
6.000 0.8161 0.01893 0.01057 -0.0345 0.0939 1.0000
6.250 0.8416 0.01993 0.01143 -0.0343 0.0884 1.0000
6.500 0.8672 0.02039 0.01203 -0.0339 0.0830 1.0000
6.750 0.8923 0.02135 0.01285 -0.0337 0.0785 1.0000
7.000 0.9173 0.02210 0.01378 -0.0331 0.0740 1.0000
7.250 0.9421 0.02279 0.01442 -0.0328 0.0704 1.0000
7.500 0.9664 0.02393 0.01568 -0.0323 0.0669 1.0000
7.750 0.9904 0.02473 0.01656 -0.0317 0.0637 1.0000
8.000 1.0148 0.02581 0.01756 -0.0314 0.0616 1.0000
8.250 1.0373 0.02740 0.01943 -0.0306 0.0595 1.0000
8.500 1.0593 0.02891 0.02117 -0.0298 0.0577 1.0000
8.750 1.0814 0.03012 0.02249 -0.0292 0.0561 1.0000
9.000 1.1037 0.03132 0.02367 -0.0287 0.0547 1.0000
9.250 1.1212 0.03354 0.02618 -0.0276 0.0534 1.0000
9.500 1.1344 0.03618 0.02929 -0.0259 0.0524 1.0000
9.750 1.1443 0.03933 0.03289 -0.0241 0.0517 1.0000
10.000 1.1489 0.04300 0.03700 -0.0221 0.0513 1.0000
10.250 1.1454 0.04752 0.04199 -0.0197 0.0512 1.0000
10.500 1.1313 0.05284 0.04778 -0.0169 0.0514 1.0000
10.750 1.1084 0.05824 0.05356 -0.0141 0.0518 1.0000
11.000 1.0773 0.06310 0.05868 -0.0110 0.0523 1.0000
11.250 1.0466 0.06831 0.06409 -0.0102 0.0527 1.0000
11.500 1.0175 0.07429 0.07022 -0.0117 0.0532 1.0000
11.750 0.9957 0.08053 0.07655 -0.0146 0.0536 1.0000
12.000 0.9791 0.08711 0.08319 -0.0182 0.0539 1.0000
12.250 0.8477 0.13609 0.13235 -0.0465 0.0784 1.0000
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Polar data table (+)
Polar graphs
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