RAE 103 AIRFOIL (rae103-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAE 103 AIRFOIL (rae103-il) Reynolds number: 100,000 Max Cl/Cd: 36.48 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae103-il-100000-n5.txt Download as CSV file: xf-rae103-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAE 103 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.6653 0.09185 0.08676 -0.0171 1.0000 0.0254
-10.750 -0.6826 0.08198 0.07691 -0.0239 1.0000 0.0244
-10.500 -0.7092 0.07326 0.06814 -0.0299 1.0000 0.0236
-10.250 -0.7376 0.06654 0.06127 -0.0330 1.0000 0.0229
-10.000 -0.7603 0.06194 0.05651 -0.0332 1.0000 0.0227
-9.750 -0.7825 0.05795 0.05231 -0.0312 1.0000 0.0225
-9.500 -0.7932 0.05487 0.04904 -0.0290 1.0000 0.0228
-9.250 -0.7988 0.05176 0.04569 -0.0270 1.0000 0.0232
-9.000 -0.8005 0.04878 0.04244 -0.0249 1.0000 0.0240
-8.750 -0.8003 0.04551 0.03881 -0.0226 1.0000 0.0246
-8.500 -0.7970 0.04215 0.03504 -0.0203 1.0000 0.0249
-8.250 -0.7899 0.03891 0.03133 -0.0180 1.0000 0.0253
-8.000 -0.7789 0.03602 0.02799 -0.0159 1.0000 0.0259
-7.750 -0.7645 0.03336 0.02491 -0.0140 1.0000 0.0267
-7.500 -0.7472 0.03115 0.02228 -0.0124 1.0000 0.0279
-7.250 -0.7292 0.02908 0.01999 -0.0111 1.0000 0.0297
-7.000 -0.7107 0.02769 0.01853 -0.0099 1.0000 0.0321
-6.750 -0.6903 0.02612 0.01680 -0.0087 1.0000 0.0339
-6.500 -0.6698 0.02464 0.01512 -0.0073 1.0000 0.0361
-6.250 -0.6497 0.02350 0.01382 -0.0059 1.0000 0.0390
-6.000 -0.6331 0.02225 0.01263 -0.0043 1.0000 0.0432
-5.750 -0.6156 0.02125 0.01156 -0.0025 1.0000 0.0473
-5.500 -0.5992 0.02027 0.01050 -0.0004 1.0000 0.0517
-5.250 -0.5828 0.01946 0.00970 0.0014 1.0000 0.0599
-5.000 -0.5672 0.01861 0.00883 0.0035 1.0000 0.0689
-4.750 -0.5516 0.01784 0.00810 0.0055 1.0000 0.0835
-4.500 -0.5367 0.01704 0.00746 0.0076 1.0000 0.1097
-4.250 -0.5235 0.01616 0.00688 0.0098 1.0000 0.1625
-4.000 -0.5133 0.01506 0.00636 0.0123 1.0000 0.2691
-3.750 -0.5064 0.01388 0.00604 0.0156 1.0000 0.4281
-3.500 -0.4790 0.01338 0.00617 0.0157 0.9912 0.5952
-3.250 -0.4461 0.01335 0.00623 0.0150 0.9830 0.6605
-3.000 -0.4146 0.01338 0.00625 0.0146 0.9738 0.7035
-2.750 -0.3820 0.01343 0.00625 0.0141 0.9656 0.7380
-2.500 -0.3485 0.01350 0.00629 0.0134 0.9580 0.7663
-2.250 -0.3155 0.01357 0.00631 0.0128 0.9502 0.7884
-2.000 -0.2797 0.01360 0.00628 0.0115 0.9436 0.8056
-1.750 -0.2458 0.01360 0.00618 0.0104 0.9354 0.8187
-1.500 -0.2094 0.01358 0.00609 0.0087 0.9285 0.8299
-1.250 -0.1771 0.01354 0.00599 0.0078 0.9197 0.8408
-1.000 -0.1388 0.01354 0.00592 0.0057 0.9136 0.8485
-0.750 -0.1062 0.01351 0.00585 0.0047 0.9046 0.8577
-0.500 -0.0684 0.01349 0.00580 0.0027 0.8984 0.8657
-0.250 -0.0354 0.01349 0.00579 0.0016 0.8894 0.8738
0.000 0.0000 0.01347 0.00576 0.0000 0.8824 0.8824
0.250 0.0354 0.01349 0.00579 -0.0016 0.8738 0.8894
0.500 0.0684 0.01349 0.00580 -0.0027 0.8656 0.8984
0.750 0.1062 0.01351 0.00586 -0.0047 0.8577 0.9046
1.000 0.1388 0.01354 0.00592 -0.0057 0.8485 0.9137
1.250 0.1772 0.01354 0.00599 -0.0078 0.8408 0.9198
1.500 0.2095 0.01358 0.00609 -0.0087 0.8299 0.9285
1.750 0.2458 0.01360 0.00618 -0.0104 0.8187 0.9354
2.000 0.2797 0.01360 0.00628 -0.0115 0.8055 0.9436
2.250 0.3155 0.01357 0.00631 -0.0128 0.7883 0.9502
2.500 0.3485 0.01350 0.00629 -0.0134 0.7663 0.9580
2.750 0.3820 0.01343 0.00625 -0.0141 0.7380 0.9657
3.000 0.4146 0.01337 0.00625 -0.0146 0.7030 0.9739
3.250 0.4460 0.01335 0.00622 -0.0150 0.6603 0.9830
3.500 0.4789 0.01338 0.00617 -0.0157 0.5941 0.9912
3.750 0.5063 0.01388 0.00604 -0.0155 0.4271 1.0000
4.000 0.5131 0.01507 0.00637 -0.0123 0.2676 1.0000
4.250 0.5235 0.01616 0.00688 -0.0098 0.1625 1.0000
4.500 0.5367 0.01705 0.00746 -0.0076 0.1097 1.0000
4.750 0.5515 0.01784 0.00811 -0.0055 0.0836 1.0000
5.000 0.5672 0.01861 0.00883 -0.0034 0.0688 1.0000
5.250 0.5828 0.01947 0.00971 -0.0014 0.0599 1.0000
5.500 0.5992 0.02027 0.01050 0.0005 0.0518 1.0000
5.750 0.6155 0.02126 0.01157 0.0025 0.0473 1.0000
6.000 0.6331 0.02225 0.01263 0.0043 0.0431 1.0000
6.250 0.6497 0.02350 0.01381 0.0059 0.0389 1.0000
6.500 0.6698 0.02465 0.01513 0.0073 0.0360 1.0000
6.750 0.6903 0.02611 0.01679 0.0087 0.0340 1.0000
7.000 0.7107 0.02768 0.01853 0.0099 0.0320 1.0000
7.250 0.7296 0.02912 0.02005 0.0111 0.0300 1.0000
7.500 0.7472 0.03117 0.02229 0.0124 0.0279 1.0000
7.750 0.7645 0.03339 0.02493 0.0141 0.0267 1.0000
8.000 0.7790 0.03600 0.02797 0.0159 0.0259 1.0000
8.250 0.7899 0.03894 0.03136 0.0180 0.0253 1.0000
8.500 0.7972 0.04212 0.03501 0.0202 0.0250 1.0000
8.750 0.8003 0.04553 0.03883 0.0226 0.0246 1.0000
9.000 0.8002 0.04887 0.04253 0.0249 0.0241 1.0000
9.250 0.7982 0.05191 0.04585 0.0270 0.0233 1.0000
9.500 0.7936 0.05483 0.04900 0.0290 0.0227 1.0000
9.750 0.7946 0.05653 0.05075 0.0304 0.0217 1.0000
10.000 0.7635 0.06155 0.05610 0.0332 0.0225 1.0000
10.250 0.7339 0.06722 0.06198 0.0327 0.0233 1.0000
10.500 0.7105 0.07310 0.06797 0.0299 0.0234 1.0000
10.750 0.6832 0.08205 0.07698 0.0238 0.0248 1.0000
11.000 0.6656 0.09200 0.08692 0.0169 0.0255 1.0000
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Polar data table (+)
Polar graphs
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