RAE 104 AIRFOIL (rae104-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAE 104 AIRFOIL (rae104-il) Reynolds number: 50,000 Max Cl/Cd: 26.21 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae104-il-50000-n5.txt Download as CSV file: xf-rae104-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAE 104 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
------ -------- --------- --------- -------- -------- --------
-11.250 -0.6355 0.10303 0.09577 -0.0242 1.0000 0.0432
-11.000 -0.6393 0.09748 0.09025 -0.0269 1.0000 0.0429
-10.750 -0.6478 0.09165 0.08443 -0.0301 1.0000 0.0426
-10.500 -0.6590 0.08618 0.07897 -0.0330 1.0000 0.0423
-10.250 -0.6732 0.08110 0.07386 -0.0353 1.0000 0.0420
-10.000 -0.6889 0.07662 0.06932 -0.0364 1.0000 0.0417
-9.750 -0.7055 0.07271 0.06534 -0.0361 1.0000 0.0415
-9.500 -0.7218 0.06921 0.06174 -0.0345 1.0000 0.0413
-9.250 -0.7346 0.06565 0.05802 -0.0326 1.0000 0.0412
-9.000 -0.7443 0.06202 0.05416 -0.0306 1.0000 0.0412
-8.750 -0.7507 0.05849 0.05034 -0.0283 1.0000 0.0414
-8.500 -0.7544 0.05504 0.04652 -0.0257 1.0000 0.0419
-8.250 -0.7559 0.05185 0.04281 -0.0229 1.0000 0.0430
-8.000 -0.7514 0.04855 0.03928 -0.0207 1.0000 0.0444
-7.750 -0.7420 0.04593 0.03648 -0.0189 1.0000 0.0460
-7.500 -0.7312 0.04320 0.03343 -0.0168 1.0000 0.0473
-7.250 -0.7179 0.04043 0.03028 -0.0148 1.0000 0.0486
-7.000 -0.7015 0.03780 0.02721 -0.0129 1.0000 0.0503
-6.750 -0.6829 0.03560 0.02456 -0.0112 1.0000 0.0537
-6.500 -0.6629 0.03352 0.02229 -0.0100 1.0000 0.0581
-6.250 -0.6396 0.03166 0.02027 -0.0090 1.0000 0.0625
-6.000 -0.6128 0.02997 0.01826 -0.0082 1.0000 0.0680
-5.750 -0.5916 0.02857 0.01691 -0.0071 1.0000 0.0772
-5.500 -0.5702 0.02720 0.01542 -0.0057 1.0000 0.0862
-5.250 -0.5536 0.02601 0.01421 -0.0037 1.0000 0.0995
-5.000 -0.5401 0.02485 0.01310 -0.0014 1.0000 0.1196
-4.750 -0.5295 0.02356 0.01202 0.0013 1.0000 0.1506
-4.500 -0.5237 0.02197 0.01103 0.0045 1.0000 0.2218
-4.250 -0.5253 0.02008 0.01027 0.0090 1.0000 0.3867
-4.000 -0.5148 0.01965 0.01119 0.0150 1.0000 0.6543
-3.750 -0.5042 0.02001 0.01146 0.0199 1.0000 0.7389
-3.500 -0.4859 0.02066 0.01198 0.0241 1.0000 0.7993
-3.250 -0.4378 0.02196 0.01296 0.0240 1.0000 0.8530
-3.000 -0.3814 0.02273 0.01329 0.0206 1.0000 0.8855
-2.750 -0.3226 0.02296 0.01314 0.0153 1.0000 0.9032
-2.500 -0.2815 0.02289 0.01282 0.0125 1.0000 0.9170
-2.250 -0.2472 0.02275 0.01246 0.0106 1.0000 0.9293
-2.000 -0.2148 0.02260 0.01214 0.0089 1.0000 0.9406
-1.750 -0.1785 0.02240 0.01180 0.0064 1.0000 0.9495
-1.500 -0.1481 0.02224 0.01153 0.0048 1.0000 0.9592
-1.250 -0.1186 0.02211 0.01128 0.0034 1.0000 0.9688
-1.000 -0.0855 0.02193 0.01104 0.0011 1.0000 0.9769
-0.750 -0.0551 0.02182 0.01087 -0.0007 1.0000 0.9856
-0.500 -0.0236 0.02172 0.01074 -0.0029 1.0000 0.9941
-0.250 0.0001 0.02167 0.01066 -0.0036 1.0000 1.0000
0.000 0.0000 0.02167 0.01067 0.0000 1.0000 1.0000
0.250 -0.0001 0.02167 0.01066 0.0036 1.0000 1.0000
0.500 0.0235 0.02172 0.01074 0.0029 0.9941 1.0000
0.750 0.0550 0.02182 0.01087 0.0007 0.9856 1.0000
1.000 0.0854 0.02193 0.01103 -0.0011 0.9769 1.0000
1.250 0.1185 0.02211 0.01128 -0.0033 0.9689 1.0000
1.500 0.1479 0.02224 0.01153 -0.0048 0.9592 1.0000
1.750 0.1784 0.02240 0.01180 -0.0064 0.9495 1.0000
2.000 0.2148 0.02260 0.01214 -0.0089 0.9406 1.0000
2.250 0.2471 0.02275 0.01245 -0.0106 0.9293 1.0000
2.500 0.2814 0.02289 0.01282 -0.0124 0.9170 1.0000
2.750 0.3224 0.02295 0.01314 -0.0153 0.9032 1.0000
3.000 0.3808 0.02273 0.01329 -0.0205 0.8856 1.0000
3.250 0.4377 0.02195 0.01296 -0.0240 0.8529 1.0000
3.500 0.4846 0.02071 0.01203 -0.0240 0.8004 1.0000
3.750 0.5042 0.02001 0.01146 -0.0199 0.7389 1.0000
4.000 0.5147 0.01964 0.01118 -0.0149 0.6537 1.0000
4.250 0.5251 0.02009 0.01027 -0.0090 0.3854 1.0000
4.500 0.5236 0.02197 0.01103 -0.0045 0.2216 1.0000
4.750 0.5295 0.02356 0.01202 -0.0013 0.1507 1.0000
5.000 0.5401 0.02485 0.01309 0.0014 0.1198 1.0000
5.250 0.5537 0.02601 0.01421 0.0037 0.1000 1.0000
5.500 0.5702 0.02719 0.01541 0.0057 0.0863 1.0000
5.750 0.5916 0.02856 0.01690 0.0071 0.0774 1.0000
6.000 0.6128 0.02997 0.01825 0.0082 0.0681 1.0000
6.250 0.6396 0.03166 0.02028 0.0090 0.0624 1.0000
6.500 0.6629 0.03352 0.02228 0.0100 0.0581 1.0000
6.750 0.6829 0.03559 0.02455 0.0112 0.0537 1.0000
7.000 0.7015 0.03782 0.02722 0.0129 0.0502 1.0000
7.250 0.7180 0.04043 0.03028 0.0148 0.0486 1.0000
7.500 0.7313 0.04318 0.03341 0.0168 0.0474 1.0000
7.750 0.7420 0.04594 0.03650 0.0189 0.0461 1.0000
8.000 0.7514 0.04855 0.03928 0.0207 0.0444 1.0000
8.250 0.7554 0.05183 0.04281 0.0229 0.0428 1.0000
8.500 0.7546 0.05502 0.04649 0.0257 0.0420 1.0000
8.750 0.7508 0.05849 0.05033 0.0283 0.0414 1.0000
9.000 0.7442 0.06205 0.05419 0.0306 0.0412 1.0000
9.250 0.7348 0.06566 0.05802 0.0326 0.0412 1.0000
9.500 0.7226 0.06917 0.06169 0.0344 0.0414 1.0000
9.750 0.7057 0.07270 0.06533 0.0361 0.0415 1.0000
10.000 0.6900 0.07656 0.06926 0.0364 0.0417 1.0000
10.250 0.6739 0.08107 0.07383 0.0352 0.0420 1.0000
10.500 0.6601 0.08607 0.07886 0.0331 0.0423 1.0000
10.750 0.6481 0.09169 0.08447 0.0300 0.0426 1.0000
11.000 0.6406 0.09733 0.09009 0.0270 0.0429 1.0000
11.250 0.6368 0.10285 0.09559 0.0244 0.0432 1.0000
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Polar data table (+)
Polar graphs
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