RAE 5214 AIRFOIL (rae5214-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RAE 5214 AIRFOIL (rae5214-il) Reynolds number: 500,000 Max Cl/Cd: 68.4 at α=8° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae5214-il-500000-n5.txt Download as CSV file: xf-rae5214-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.8651 0.04842 0.04518 -0.0534 1.0000 0.0089
-12.000 -0.9047 0.04072 0.03712 -0.0573 1.0000 0.0089
-11.750 -0.9292 0.03632 0.03245 -0.0576 1.0000 0.0089
-11.500 -0.9398 0.03324 0.02909 -0.0562 1.0000 0.0090
-11.250 -0.9419 0.03015 0.02567 -0.0550 1.0000 0.0091
-11.000 -0.9355 0.02783 0.02305 -0.0538 1.0000 0.0092
-10.750 -0.9245 0.02602 0.02103 -0.0527 1.0000 0.0094
-10.500 -0.9091 0.02484 0.01974 -0.0517 1.0000 0.0095
-10.250 -0.8918 0.02391 0.01872 -0.0507 1.0000 0.0097
-10.000 -0.8736 0.02306 0.01780 -0.0497 1.0000 0.0099
-9.750 -0.8549 0.02225 0.01690 -0.0487 1.0000 0.0100
-9.500 -0.8363 0.02139 0.01593 -0.0476 1.0000 0.0102
-9.250 -0.8174 0.02055 0.01499 -0.0465 1.0000 0.0105
-9.000 -0.7987 0.01966 0.01398 -0.0454 1.0000 0.0108
-8.750 -0.7800 0.01876 0.01295 -0.0441 1.0000 0.0112
-8.500 -0.7611 0.01792 0.01198 -0.0428 1.0000 0.0115
-8.250 -0.7313 0.01723 0.01123 -0.0438 0.9981 0.0119
-8.000 -0.6996 0.01672 0.01068 -0.0450 0.9957 0.0122
-7.750 -0.6682 0.01619 0.01010 -0.0460 0.9933 0.0127
-7.500 -0.6376 0.01557 0.00941 -0.0469 0.9904 0.0133
-7.250 -0.6062 0.01496 0.00869 -0.0480 0.9878 0.0140
-7.000 -0.5739 0.01441 0.00811 -0.0492 0.9858 0.0145
-6.750 -0.5424 0.01403 0.00772 -0.0502 0.9830 0.0152
-6.500 -0.5114 0.01360 0.00725 -0.0510 0.9796 0.0161
-6.250 -0.4791 0.01320 0.00679 -0.0521 0.9769 0.0171
-6.000 -0.4463 0.01280 0.00641 -0.0533 0.9746 0.0181
-5.750 -0.4171 0.01246 0.00604 -0.0537 0.9693 0.0193
-5.500 -0.3862 0.01205 0.00559 -0.0544 0.9648 0.0203
-5.250 -0.3555 0.01161 0.00516 -0.0552 0.9604 0.0213
-5.000 -0.3260 0.01125 0.00479 -0.0555 0.9547 0.0223
-4.750 -0.2944 0.01095 0.00445 -0.0563 0.9502 0.0234
-4.500 -0.2653 0.01056 0.00407 -0.0566 0.9441 0.0247
-4.250 -0.2356 0.01027 0.00377 -0.0570 0.9381 0.0260
-4.000 -0.2058 0.01000 0.00348 -0.0573 0.9323 0.0273
-3.750 -0.1778 0.00973 0.00322 -0.0572 0.9243 0.0291
-3.500 -0.1487 0.00951 0.00298 -0.0574 0.9177 0.0311
-3.250 -0.1211 0.00929 0.00278 -0.0572 0.9093 0.0341
-3.000 -0.0927 0.00909 0.00257 -0.0571 0.9021 0.0392
-2.750 -0.0652 0.00888 0.00241 -0.0570 0.8942 0.0504
-2.500 -0.0374 0.00865 0.00224 -0.0568 0.8867 0.0712
-2.250 -0.0102 0.00831 0.00207 -0.0567 0.8781 0.1192
-2.000 0.0161 0.00781 0.00186 -0.0565 0.8657 0.2136
-1.750 0.0412 0.00709 0.00162 -0.0563 0.8492 0.3680
-1.500 0.0665 0.00664 0.00153 -0.0558 0.8306 0.4897
-1.250 0.0927 0.00653 0.00156 -0.0551 0.8104 0.5580
-1.000 0.1193 0.00655 0.00161 -0.0545 0.7895 0.5949
-0.750 0.1463 0.00663 0.00164 -0.0540 0.7677 0.6180
-0.500 0.1729 0.00672 0.00165 -0.0534 0.7410 0.6297
-0.250 0.1993 0.00685 0.00161 -0.0528 0.7036 0.6352
0.000 0.2256 0.00701 0.00159 -0.0522 0.6616 0.6394
0.250 0.2515 0.00719 0.00160 -0.0516 0.6143 0.6433
0.500 0.2760 0.00754 0.00165 -0.0508 0.5391 0.6475
0.750 0.2995 0.00807 0.00176 -0.0499 0.4409 0.6519
1.000 0.3240 0.00858 0.00190 -0.0493 0.3511 0.6561
1.250 0.3483 0.00910 0.00206 -0.0488 0.2604 0.6598
1.500 0.3739 0.00948 0.00222 -0.0484 0.2046 0.6638
1.750 0.4004 0.00975 0.00235 -0.0480 0.1750 0.6680
2.000 0.4274 0.00997 0.00246 -0.0478 0.1553 0.6721
2.250 0.4544 0.01016 0.00259 -0.0476 0.1399 0.6758
2.500 0.4813 0.01034 0.00272 -0.0473 0.1266 0.6794
2.750 0.5082 0.01052 0.00286 -0.0470 0.1161 0.6834
3.000 0.5353 0.01070 0.00301 -0.0468 0.1088 0.6877
3.250 0.5623 0.01090 0.00317 -0.0465 0.1023 0.6917
3.500 0.5892 0.01107 0.00334 -0.0462 0.0974 0.6953
3.750 0.6159 0.01125 0.00352 -0.0459 0.0923 0.6995
4.000 0.6425 0.01146 0.00373 -0.0456 0.0874 0.7042
4.250 0.6693 0.01166 0.00391 -0.0453 0.0821 0.7088
4.500 0.6957 0.01187 0.00411 -0.0450 0.0774 0.7130
4.750 0.7223 0.01205 0.00432 -0.0447 0.0732 0.7177
5.000 0.7485 0.01229 0.00454 -0.0443 0.0690 0.7232
5.250 0.7750 0.01248 0.00477 -0.0440 0.0655 0.7287
5.500 0.8009 0.01272 0.00501 -0.0436 0.0617 0.7346
5.750 0.8271 0.01294 0.00527 -0.0433 0.0582 0.7409
6.000 0.8528 0.01320 0.00553 -0.0429 0.0544 0.7473
6.250 0.8785 0.01345 0.00582 -0.0424 0.0511 0.7549
6.500 0.9038 0.01375 0.00612 -0.0420 0.0477 0.7636
6.750 0.9290 0.01400 0.00645 -0.0415 0.0449 0.7740
7.000 0.9537 0.01433 0.00680 -0.0409 0.0421 0.7859
7.250 0.9784 0.01460 0.00716 -0.0403 0.0396 0.8005
7.500 1.0021 0.01492 0.00755 -0.0395 0.0374 0.8205
7.750 1.0245 0.01515 0.00794 -0.0384 0.0355 0.8570
8.000 1.0513 0.01537 0.00833 -0.0383 0.0335 1.0000
8.250 1.0756 0.01582 0.00879 -0.0378 0.0320 1.0000
8.500 1.1000 0.01625 0.00925 -0.0372 0.0305 1.0000
8.750 1.1239 0.01673 0.00974 -0.0367 0.0294 1.0000
9.000 1.1471 0.01727 0.01030 -0.0360 0.0285 1.0000
9.250 1.1704 0.01777 0.01085 -0.0354 0.0277 1.0000
9.500 1.1934 0.01830 0.01143 -0.0347 0.0270 1.0000
9.750 1.2159 0.01886 0.01204 -0.0340 0.0264 1.0000
10.000 1.2379 0.01945 0.01267 -0.0332 0.0258 1.0000
10.250 1.2590 0.02011 0.01337 -0.0323 0.0253 1.0000
10.500 1.2790 0.02087 0.01416 -0.0313 0.0248 1.0000
10.750 1.2998 0.02151 0.01489 -0.0304 0.0244 1.0000
11.000 1.3199 0.02219 0.01565 -0.0294 0.0240 1.0000
11.250 1.3392 0.02291 0.01645 -0.0283 0.0236 1.0000
11.500 1.3577 0.02366 0.01728 -0.0271 0.0232 1.0000
11.750 1.3752 0.02445 0.01815 -0.0259 0.0229 1.0000
12.000 1.3917 0.02529 0.01906 -0.0245 0.0226 1.0000
12.250 1.4067 0.02617 0.02003 -0.0230 0.0223 1.0000
12.500 1.4189 0.02709 0.02102 -0.0211 0.0221 1.0000
12.750 1.4281 0.02810 0.02210 -0.0188 0.0219 1.0000
13.000 1.4347 0.02925 0.02333 -0.0163 0.0217 1.0000
13.250 1.4394 0.03058 0.02474 -0.0139 0.0215 1.0000
13.500 1.4463 0.03181 0.02609 -0.0119 0.0214 1.0000
13.750 1.4516 0.03321 0.02762 -0.0100 0.0212 1.0000
14.000 1.4556 0.03478 0.02932 -0.0082 0.0211 1.0000
14.250 1.4585 0.03652 0.03119 -0.0067 0.0210 1.0000
14.500 1.4594 0.03852 0.03333 -0.0054 0.0208 1.0000
14.750 1.4590 0.04077 0.03573 -0.0044 0.0207 1.0000
15.000 1.4568 0.04336 0.03846 -0.0038 0.0206 1.0000
15.250 1.4529 0.04630 0.04154 -0.0036 0.0205 1.0000
15.500 1.4469 0.04972 0.04512 -0.0040 0.0204 1.0000
15.750 1.4388 0.05368 0.04923 -0.0050 0.0203 1.0000
16.000 1.4281 0.05837 0.05408 -0.0068 0.0202 1.0000
16.250 1.4141 0.06400 0.05989 -0.0096 0.0202 1.0000
16.500 1.3960 0.07092 0.06700 -0.0136 0.0201 1.0000
16.750 1.3727 0.07950 0.07578 -0.0191 0.0201 1.0000
17.000 1.3431 0.09003 0.08652 -0.0260 0.0201 1.0000
17.250 1.3029 0.10337 0.10008 -0.0347 0.0202 1.0000
17.500 1.2541 0.11897 0.11592 -0.0446 0.0202 1.0000
17.750 1.1770 0.14205 0.13928 -0.0596 0.0203 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAE 5214 AIRFOIL (rae5214-il)