RAE 102 AIRFOIL (rae102-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RAE 102 AIRFOIL (rae102-il) Reynolds number: 50,000 Max Cl/Cd: 27.39 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae102-il-50000-n5.txt Download as CSV file: xf-rae102-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAE 102 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.7472 0.07692 0.06954 -0.0272 1.0000 0.0445
-10.000 -0.7632 0.07278 0.06530 -0.0276 1.0000 0.0443
-9.750 -0.7788 0.06905 0.06142 -0.0267 1.0000 0.0443
-9.500 -0.7904 0.06511 0.05725 -0.0258 1.0000 0.0443
-9.250 -0.7972 0.06125 0.05311 -0.0246 1.0000 0.0443
-9.000 -0.8007 0.05742 0.04893 -0.0232 1.0000 0.0445
-8.750 -0.7993 0.05360 0.04480 -0.0217 1.0000 0.0448
-8.500 -0.7928 0.05000 0.04093 -0.0205 1.0000 0.0455
-8.250 -0.7829 0.04693 0.03761 -0.0192 1.0000 0.0467
-8.000 -0.7702 0.04447 0.03488 -0.0181 1.0000 0.0491
-7.750 -0.7565 0.04188 0.03192 -0.0167 1.0000 0.0520
-7.500 -0.7400 0.03908 0.02861 -0.0152 1.0000 0.0546
-7.250 -0.7194 0.03648 0.02547 -0.0140 1.0000 0.0565
-7.000 -0.6980 0.03422 0.02326 -0.0133 1.0000 0.0602
-6.750 -0.6760 0.03251 0.02132 -0.0123 1.0000 0.0665
-6.500 -0.6517 0.03063 0.01930 -0.0114 1.0000 0.0717
-6.250 -0.6296 0.02914 0.01768 -0.0103 1.0000 0.0792
-6.000 -0.6096 0.02773 0.01626 -0.0090 1.0000 0.0895
-5.750 -0.5908 0.02634 0.01482 -0.0074 1.0000 0.1006
-5.500 -0.5737 0.02502 0.01354 -0.0058 1.0000 0.1196
-5.250 -0.5583 0.02355 0.01223 -0.0041 1.0000 0.1467
-5.000 -0.5453 0.02184 0.01095 -0.0023 1.0000 0.2057
-4.750 -0.5369 0.01995 0.01007 0.0003 1.0000 0.3414
-4.500 -0.5264 0.01921 0.01005 0.0041 1.0000 0.4962
-4.250 -0.5121 0.01902 0.01000 0.0075 1.0000 0.5760
-4.000 -0.4963 0.01898 0.01001 0.0109 1.0000 0.6330
-3.750 -0.4800 0.01899 0.00998 0.0143 1.0000 0.6789
-3.500 -0.4616 0.01906 0.01003 0.0175 1.0000 0.7165
-3.250 -0.4428 0.01907 0.00999 0.0203 1.0000 0.7489
-3.000 -0.4217 0.01907 0.00990 0.0225 1.0000 0.7759
-2.750 -0.3998 0.01900 0.00973 0.0243 1.0000 0.7993
-2.500 -0.3756 0.01891 0.00951 0.0254 1.0000 0.8197
-2.250 -0.3526 0.01877 0.00927 0.0263 1.0000 0.8396
-2.000 -0.3237 0.01870 0.00909 0.0261 1.0000 0.8567
-1.750 -0.2933 0.01863 0.00890 0.0254 1.0000 0.8735
-1.500 -0.2610 0.01859 0.00876 0.0242 1.0000 0.8902
-1.250 -0.2267 0.01856 0.00865 0.0223 1.0000 0.9068
-1.000 -0.1904 0.01855 0.00857 0.0199 1.0000 0.9233
-0.750 -0.1518 0.01856 0.00850 0.0168 1.0000 0.9400
-0.500 -0.1104 0.01859 0.00847 0.0130 1.0000 0.9560
-0.250 -0.0564 0.01861 0.00845 0.0067 0.9905 0.9675
0.000 0.0000 0.01863 0.00845 0.0000 0.9788 0.9789
0.250 0.0565 0.01861 0.00845 -0.0067 0.9675 0.9905
0.500 0.1103 0.01858 0.00847 -0.0130 0.9560 1.0000
0.750 0.1518 0.01856 0.00849 -0.0168 0.9400 1.0000
1.000 0.1904 0.01855 0.00856 -0.0199 0.9234 1.0000
1.250 0.2266 0.01856 0.00865 -0.0223 0.9068 1.0000
1.500 0.2609 0.01858 0.00876 -0.0242 0.8902 1.0000
1.750 0.2932 0.01863 0.00890 -0.0254 0.8735 1.0000
2.000 0.3236 0.01869 0.00909 -0.0261 0.8567 1.0000
2.250 0.3525 0.01877 0.00926 -0.0263 0.8396 1.0000
2.500 0.3754 0.01891 0.00950 -0.0253 0.8197 1.0000
2.750 0.3996 0.01899 0.00973 -0.0242 0.7993 1.0000
3.000 0.4215 0.01906 0.00990 -0.0225 0.7759 1.0000
3.250 0.4426 0.01907 0.00998 -0.0203 0.7490 1.0000
3.500 0.4614 0.01906 0.01003 -0.0174 0.7167 1.0000
3.750 0.4798 0.01899 0.00997 -0.0143 0.6791 1.0000
4.000 0.4961 0.01898 0.01000 -0.0109 0.6332 1.0000
4.250 0.5119 0.01902 0.01000 -0.0075 0.5763 1.0000
4.500 0.5262 0.01921 0.01005 -0.0041 0.4967 1.0000
4.750 0.5367 0.01995 0.01007 -0.0003 0.3407 1.0000
5.000 0.5452 0.02184 0.01095 0.0023 0.2054 1.0000
5.250 0.5583 0.02355 0.01223 0.0041 0.1469 1.0000
5.500 0.5738 0.02502 0.01353 0.0058 0.1197 1.0000
5.750 0.5908 0.02634 0.01482 0.0074 0.1004 1.0000
6.000 0.6097 0.02774 0.01626 0.0090 0.0894 1.0000
6.250 0.6297 0.02914 0.01768 0.0102 0.0791 1.0000
6.500 0.6519 0.03063 0.01930 0.0113 0.0717 1.0000
6.750 0.6761 0.03252 0.02132 0.0123 0.0663 1.0000
7.000 0.6981 0.03423 0.02326 0.0133 0.0602 1.0000
7.250 0.7196 0.03649 0.02548 0.0139 0.0565 1.0000
7.500 0.7402 0.03910 0.02863 0.0152 0.0546 1.0000
7.750 0.7568 0.04188 0.03192 0.0166 0.0520 1.0000
8.000 0.7706 0.04450 0.03492 0.0180 0.0492 1.0000
8.250 0.7833 0.04695 0.03762 0.0192 0.0467 1.0000
8.500 0.7933 0.05002 0.04095 0.0204 0.0455 1.0000
8.750 0.7997 0.05361 0.04482 0.0217 0.0448 1.0000
9.000 0.8013 0.05745 0.04896 0.0231 0.0445 1.0000
9.250 0.7981 0.06125 0.05311 0.0245 0.0444 1.0000
9.500 0.7907 0.06518 0.05733 0.0257 0.0443 1.0000
9.750 0.7798 0.06909 0.06147 0.0266 0.0443 1.0000
10.000 0.7650 0.07284 0.06536 0.0274 0.0444 1.0000
10.250 0.7500 0.07692 0.06953 0.0271 0.0445 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAE 102 AIRFOIL (rae102-il)