RAE 5214 AIRFOIL (rae5214-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAE 5214 AIRFOIL (rae5214-il) Reynolds number: 200,000 Max Cl/Cd: 49.13 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae5214-il-200000-n5.txt Download as CSV file: xf-rae5214-il-200000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5702 0.08385 0.07983 -0.0288 1.0000 0.0167
-10.000 -0.5852 0.07485 0.07088 -0.0348 1.0000 0.0165
-9.750 -0.6147 0.06338 0.05932 -0.0440 1.0000 0.0163
-9.500 -0.6415 0.05689 0.05271 -0.0482 1.0000 0.0161
-9.250 -0.6683 0.05161 0.04725 -0.0494 1.0000 0.0160
-9.000 -0.6859 0.04602 0.04133 -0.0495 1.0000 0.0161
-8.750 -0.6970 0.04048 0.03535 -0.0487 1.0000 0.0162
-8.500 -0.7016 0.03509 0.02932 -0.0473 1.0000 0.0166
-8.250 -0.6934 0.03202 0.02583 -0.0460 1.0000 0.0172
-8.000 -0.6765 0.03116 0.02495 -0.0450 1.0000 0.0177
-7.750 -0.6608 0.02977 0.02340 -0.0439 1.0000 0.0183
-7.500 -0.6461 0.02763 0.02090 -0.0425 1.0000 0.0191
-7.250 -0.6306 0.02516 0.01788 -0.0410 1.0000 0.0200
-7.000 -0.6126 0.02388 0.01651 -0.0397 1.0000 0.0204
-6.750 -0.5939 0.02287 0.01543 -0.0385 1.0000 0.0210
-6.500 -0.5748 0.02190 0.01435 -0.0372 1.0000 0.0218
-6.250 -0.5528 0.02085 0.01303 -0.0364 0.9994 0.0232
-6.000 -0.5208 0.01987 0.01203 -0.0377 0.9962 0.0244
-5.750 -0.4887 0.01892 0.01101 -0.0389 0.9930 0.0255
-5.500 -0.4572 0.01797 0.00993 -0.0398 0.9894 0.0268
-5.250 -0.4246 0.01721 0.00920 -0.0412 0.9862 0.0284
-5.000 -0.3925 0.01659 0.00854 -0.0423 0.9825 0.0304
-4.750 -0.3608 0.01587 0.00781 -0.0433 0.9782 0.0319
-4.500 -0.3272 0.01530 0.00726 -0.0448 0.9751 0.0338
-4.250 -0.2962 0.01478 0.00671 -0.0456 0.9703 0.0363
-4.000 -0.2638 0.01427 0.00622 -0.0468 0.9658 0.0387
-3.750 -0.2290 0.01381 0.00575 -0.0483 0.9628 0.0421
-3.500 -0.1984 0.01341 0.00535 -0.0490 0.9573 0.0462
-3.250 -0.1653 0.01299 0.00494 -0.0502 0.9529 0.0522
-3.000 -0.1307 0.01253 0.00455 -0.0516 0.9495 0.0662
-2.750 -0.1015 0.01199 0.00423 -0.0522 0.9432 0.1106
-2.500 -0.0711 0.01097 0.00385 -0.0535 0.9380 0.2636
-2.250 -0.0406 0.00999 0.00379 -0.0545 0.9340 0.4874
-2.000 -0.0139 0.00992 0.00394 -0.0540 0.9262 0.5670
-1.750 0.0170 0.01001 0.00414 -0.0541 0.9208 0.6213
-1.500 0.0463 0.01014 0.00434 -0.0538 0.9148 0.6541
-1.250 0.0753 0.01016 0.00438 -0.0536 0.9073 0.6682
-1.000 0.1074 0.01009 0.00423 -0.0540 0.8979 0.6753
-0.750 0.1382 0.00997 0.00408 -0.0540 0.8846 0.6798
-0.500 0.1666 0.00988 0.00395 -0.0535 0.8677 0.6851
-0.250 0.1950 0.00982 0.00380 -0.0531 0.8487 0.6907
0.000 0.2218 0.00975 0.00367 -0.0523 0.8250 0.6952
0.250 0.2486 0.00971 0.00357 -0.0515 0.7995 0.6999
0.500 0.2755 0.00972 0.00347 -0.0507 0.7715 0.7051
0.750 0.3024 0.00975 0.00340 -0.0501 0.7412 0.7104
1.000 0.3285 0.00980 0.00336 -0.0493 0.7074 0.7145
1.250 0.3541 0.00991 0.00332 -0.0484 0.6620 0.7194
1.500 0.3781 0.01018 0.00328 -0.0472 0.5879 0.7247
1.750 0.3985 0.01077 0.00333 -0.0456 0.4720 0.7293
2.000 0.4184 0.01153 0.00355 -0.0442 0.3497 0.7339
2.250 0.4402 0.01225 0.00381 -0.0433 0.2497 0.7394
2.750 0.4894 0.01309 0.00427 -0.0422 0.1759 0.7496
3.000 0.5148 0.01340 0.00451 -0.0417 0.1579 0.7555
3.250 0.5407 0.01369 0.00475 -0.0413 0.1439 0.7620
3.500 0.5662 0.01394 0.00501 -0.0408 0.1331 0.7676
3.750 0.5916 0.01423 0.00527 -0.0403 0.1244 0.7742
4.000 0.6171 0.01451 0.00556 -0.0398 0.1168 0.7814
4.250 0.6420 0.01480 0.00587 -0.0392 0.1108 0.7895
4.500 0.6671 0.01509 0.00619 -0.0387 0.1049 0.7983
4.750 0.6913 0.01545 0.00653 -0.0380 0.0996 0.8082
5.000 0.7164 0.01572 0.00689 -0.0374 0.0942 0.8194
5.250 0.7402 0.01609 0.00726 -0.0366 0.0889 0.8345
5.500 0.7645 0.01630 0.00760 -0.0358 0.0838 0.8560
5.750 0.7884 0.01655 0.00794 -0.0350 0.0792 0.8979
6.000 0.8185 0.01689 0.00836 -0.0356 0.0744 1.0000
6.250 0.8442 0.01731 0.00875 -0.0353 0.0701 1.0000
6.500 0.8695 0.01777 0.00925 -0.0350 0.0661 1.0000
6.750 0.8946 0.01821 0.00968 -0.0347 0.0622 1.0000
7.000 0.9193 0.01873 0.01022 -0.0342 0.0587 1.0000
7.250 0.9437 0.01923 0.01074 -0.0338 0.0555 1.0000
7.500 0.9675 0.01984 0.01136 -0.0333 0.0527 1.0000
7.750 0.9913 0.02042 0.01199 -0.0327 0.0498 1.0000
8.000 1.0143 0.02107 0.01263 -0.0321 0.0475 1.0000
8.250 1.0374 0.02177 0.01342 -0.0315 0.0451 1.0000
8.500 1.0599 0.02245 0.01414 -0.0308 0.0432 1.0000
8.750 1.0814 0.02327 0.01498 -0.0301 0.0418 1.0000
9.000 1.1030 0.02415 0.01599 -0.0292 0.0403 1.0000
9.250 1.1240 0.02503 0.01696 -0.0284 0.0390 1.0000
9.500 1.1444 0.02590 0.01787 -0.0275 0.0380 1.0000
9.750 1.1637 0.02686 0.01884 -0.0266 0.0371 1.0000
10.000 1.1831 0.02795 0.02013 -0.0256 0.0361 1.0000
10.250 1.2016 0.02906 0.02139 -0.0246 0.0351 1.0000
10.500 1.2192 0.03014 0.02259 -0.0235 0.0343 1.0000
10.750 1.2358 0.03124 0.02378 -0.0223 0.0336 1.0000
11.000 1.2513 0.03239 0.02501 -0.0210 0.0331 1.0000
11.250 1.2654 0.03365 0.02632 -0.0197 0.0326 1.0000
11.500 1.2769 0.03527 0.02817 -0.0181 0.0322 1.0000
11.750 1.2850 0.03698 0.03012 -0.0161 0.0317 1.0000
12.000 1.2896 0.03878 0.03215 -0.0138 0.0314 1.0000
12.250 1.2915 0.04073 0.03430 -0.0115 0.0310 1.0000
12.500 1.2912 0.04278 0.03657 -0.0094 0.0307 1.0000
12.750 1.2888 0.04506 0.03905 -0.0075 0.0304 1.0000
13.000 1.2846 0.04757 0.04175 -0.0061 0.0301 1.0000
13.250 1.2783 0.05043 0.04480 -0.0052 0.0299 1.0000
13.500 1.2700 0.05369 0.04825 -0.0048 0.0296 1.0000
13.750 1.2591 0.05754 0.05228 -0.0053 0.0295 1.0000
14.000 1.2450 0.06219 0.05712 -0.0067 0.0293 1.0000
14.250 1.2243 0.06836 0.06353 -0.0097 0.0293 1.0000
14.500 1.1787 0.08024 0.07578 -0.0177 0.0294 1.0000
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Polar data table (+)
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