RAE6-9CK AIRFOIL (rae69ck-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: RAE6-9CK AIRFOIL (rae69ck-il) Reynolds number: 100,000 Max Cl/Cd: 37.2 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae69ck-il-100000.txt Download as CSV file: xf-rae69ck-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: RAE6-9CK AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.4490 0.09853 0.09362 -0.0362 1.0000 0.1700
-10.500 -0.5973 0.06491 0.05992 -0.0585 1.0000 0.0811
-10.250 -0.6171 0.05952 0.05451 -0.0589 1.0000 0.0789
-10.000 -0.6511 0.05456 0.04947 -0.0585 1.0000 0.0760
-9.750 -0.7465 0.06409 0.05868 -0.0537 1.0000 0.0796
-9.500 -0.7942 0.05950 0.05341 -0.0493 1.0000 0.0694
-9.250 -0.7919 0.05471 0.04858 -0.0478 1.0000 0.0672
-9.000 -0.7962 0.05092 0.04456 -0.0454 1.0000 0.0653
-8.750 -0.7987 0.04754 0.04087 -0.0428 1.0000 0.0646
-8.500 -0.7972 0.04452 0.03751 -0.0402 1.0000 0.0652
-8.250 -0.7920 0.04155 0.03417 -0.0378 1.0000 0.0655
-8.000 -0.7833 0.03856 0.03078 -0.0356 1.0000 0.0654
-7.750 -0.7709 0.03578 0.02759 -0.0336 1.0000 0.0655
-7.500 -0.7560 0.03336 0.02480 -0.0317 1.0000 0.0667
-7.250 -0.7399 0.03175 0.02277 -0.0298 1.0000 0.0694
-7.000 -0.7218 0.02932 0.02028 -0.0285 1.0000 0.0730
-6.750 -0.7027 0.02771 0.01860 -0.0271 1.0000 0.0768
-6.500 -0.6831 0.02639 0.01710 -0.0255 1.0000 0.0818
-6.250 -0.6649 0.02487 0.01565 -0.0241 1.0000 0.0888
-6.000 -0.6459 0.02380 0.01448 -0.0225 1.0000 0.0964
-5.750 -0.6290 0.02252 0.01338 -0.0209 1.0000 0.1079
-5.500 -0.6123 0.02138 0.01235 -0.0193 1.0000 0.1239
-5.250 -0.5962 0.02002 0.01125 -0.0180 1.0000 0.1521
-5.000 -0.5825 0.01658 0.00940 -0.0185 1.0000 0.3411
-4.750 -0.5689 0.01736 0.01086 -0.0145 1.0000 0.5572
-4.500 -0.5527 0.01833 0.01178 -0.0111 1.0000 0.5933
-4.250 -0.5365 0.01914 0.01253 -0.0079 1.0000 0.6184
-4.000 -0.5211 0.01987 0.01322 -0.0045 1.0000 0.6374
-3.750 -0.5061 0.02051 0.01381 -0.0011 1.0000 0.6539
-3.500 -0.4910 0.02102 0.01427 0.0020 1.0000 0.6692
-3.250 -0.4763 0.02146 0.01467 0.0052 1.0000 0.6841
-3.000 -0.4648 0.02205 0.01526 0.0094 1.0000 0.7008
-2.750 -0.4550 0.02262 0.01583 0.0140 1.0000 0.7191
-2.500 -0.4433 0.02294 0.01611 0.0178 1.0000 0.7356
-2.250 -0.4263 0.02285 0.01595 0.0195 1.0000 0.7448
-2.000 -0.4035 0.02256 0.01553 0.0192 1.0000 0.7506
-1.750 -0.3794 0.02224 0.01510 0.0184 1.0000 0.7559
-1.500 -0.3596 0.02208 0.01486 0.0191 1.0000 0.7604
-1.250 -0.3359 0.02188 0.01457 0.0185 1.0000 0.7658
-1.000 -0.3116 0.02171 0.01431 0.0178 0.9999 0.7708
-0.750 -0.2756 0.02196 0.01450 0.0153 0.9944 0.7753
-0.500 -0.2394 0.02207 0.01454 0.0123 0.9885 0.7805
0.000 -0.1703 0.02241 0.01481 0.0076 0.9765 0.7906
0.250 -0.1306 0.02270 0.01506 0.0041 0.9708 0.7961
0.500 -0.1009 0.02275 0.01511 0.0027 0.9637 0.8012
0.750 -0.0645 0.02303 0.01540 0.0003 0.9575 0.8069
1.000 -0.0328 0.02314 0.01551 -0.0016 0.9499 0.8135
1.250 0.0028 0.02339 0.01582 -0.0037 0.9431 0.8186
1.500 0.0320 0.02351 0.01598 -0.0049 0.9347 0.8254
1.750 0.0712 0.02376 0.01631 -0.0076 0.9279 0.8318
2.000 0.0978 0.02385 0.01647 -0.0082 0.9183 0.8390
2.250 0.1367 0.02406 0.01677 -0.0108 0.9108 0.8457
2.500 0.1672 0.02414 0.01695 -0.0118 0.9007 0.8537
2.750 0.1973 0.02421 0.01714 -0.0126 0.8901 0.8618
3.000 0.2380 0.02424 0.01732 -0.0151 0.8804 0.8708
3.250 0.2779 0.02408 0.01733 -0.0170 0.8694 0.8790
3.500 0.3144 0.02383 0.01724 -0.0182 0.8559 0.8887
3.750 0.3656 0.02298 0.01663 -0.0208 0.8404 0.8978
4.000 0.4396 0.02052 0.01446 -0.0247 0.8142 0.9037
4.250 0.4992 0.01829 0.01249 -0.0263 0.7878 0.9110
4.500 0.5435 0.01662 0.01102 -0.0260 0.7531 0.9201
4.750 0.5739 0.01548 0.00997 -0.0238 0.6926 0.9323
5.000 0.5862 0.01576 0.00889 -0.0182 0.3942 0.9466
5.250 0.5894 0.01884 0.01019 -0.0153 0.1553 0.9654
5.500 0.6225 0.02062 0.01164 -0.0173 0.1192 0.9834
5.750 0.6527 0.02211 0.01297 -0.0187 0.1027 1.0000
6.000 0.6786 0.02340 0.01420 -0.0193 0.0924 1.0000
6.250 0.7116 0.02507 0.01592 -0.0207 0.0838 1.0000
6.500 0.7477 0.02735 0.01809 -0.0228 0.0778 1.0000
6.750 0.7791 0.02927 0.02031 -0.0236 0.0725 1.0000
7.000 0.8099 0.03161 0.02294 -0.0242 0.0697 1.0000
7.250 0.8377 0.03432 0.02599 -0.0243 0.0684 1.0000
7.500 0.8632 0.03698 0.02876 -0.0246 0.0657 1.0000
7.750 0.8827 0.04184 0.03387 -0.0244 0.0638 1.0000
8.000 0.9017 0.04416 0.03662 -0.0229 0.0646 1.0000
8.250 0.9024 0.04913 0.04270 -0.0185 0.0721 1.0000
10.000 0.7625 0.09613 0.09200 -0.0160 0.1594 1.0000
10.250 0.7248 0.10380 0.09955 -0.0244 0.1528 1.0000
10.500 0.6114 0.10749 0.10359 -0.0212 0.1593 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAE6-9CK AIRFOIL (rae69ck-il)