RAE 102 AIRFOIL (rae102-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAE 102 AIRFOIL (rae102-il) Reynolds number: 200,000 Max Cl/Cd: 41.45 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae102-il-200000-n5.txt Download as CSV file: xf-rae102-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.7606 0.08610 0.08241 -0.0120 1.0000 0.0130
-12.000 -0.7867 0.07523 0.07143 -0.0200 1.0000 0.0129
-11.750 -0.8049 0.06810 0.06420 -0.0252 1.0000 0.0127
-11.500 -0.8264 0.06155 0.05748 -0.0295 1.0000 0.0126
-11.250 -0.8469 0.05619 0.05193 -0.0320 1.0000 0.0126
-11.000 -0.8670 0.05170 0.04723 -0.0327 1.0000 0.0126
-10.750 -0.8845 0.04809 0.04338 -0.0315 1.0000 0.0126
-10.500 -0.9014 0.04483 0.03983 -0.0286 1.0000 0.0127
-10.250 -0.9112 0.04174 0.03641 -0.0257 1.0000 0.0128
-10.000 -0.9135 0.03873 0.03303 -0.0232 1.0000 0.0129
-9.750 -0.9119 0.03570 0.02966 -0.0210 1.0000 0.0132
-9.500 -0.9034 0.03353 0.02730 -0.0193 1.0000 0.0136
-9.250 -0.8914 0.03162 0.02518 -0.0178 1.0000 0.0139
-9.000 -0.8762 0.03038 0.02376 -0.0165 1.0000 0.0146
-8.750 -0.8601 0.02906 0.02225 -0.0152 1.0000 0.0154
-8.500 -0.8431 0.02758 0.02052 -0.0139 1.0000 0.0164
-8.250 -0.8252 0.02594 0.01862 -0.0125 1.0000 0.0172
-8.000 -0.8062 0.02445 0.01690 -0.0112 1.0000 0.0178
-7.750 -0.7880 0.02282 0.01512 -0.0099 1.0000 0.0185
-7.500 -0.7697 0.02159 0.01384 -0.0086 1.0000 0.0196
-7.250 -0.7501 0.02071 0.01287 -0.0074 1.0000 0.0208
-7.000 -0.7299 0.01994 0.01202 -0.0062 1.0000 0.0226
-6.750 -0.7097 0.01916 0.01113 -0.0050 1.0000 0.0244
-6.500 -0.6920 0.01810 0.01001 -0.0033 1.0000 0.0259
-6.250 -0.6736 0.01728 0.00916 -0.0018 1.0000 0.0279
-6.000 -0.6537 0.01667 0.00848 -0.0005 1.0000 0.0309
-5.750 -0.6335 0.01612 0.00785 0.0008 1.0000 0.0341
-5.500 -0.6150 0.01538 0.00712 0.0024 1.0000 0.0385
-5.250 -0.5947 0.01489 0.00656 0.0037 1.0000 0.0437
-5.000 -0.5753 0.01432 0.00602 0.0051 1.0000 0.0522
-4.750 -0.5556 0.01382 0.00552 0.0065 1.0000 0.0632
-4.500 -0.5361 0.01334 0.00510 0.0079 1.0000 0.0814
-4.250 -0.5138 0.01276 0.00470 0.0085 0.9984 0.1149
-4.000 -0.4814 0.01196 0.00424 0.0068 0.9903 0.1868
-3.750 -0.4507 0.01100 0.00379 0.0053 0.9818 0.3039
-3.500 -0.4219 0.01018 0.00358 0.0045 0.9720 0.4431
-3.250 -0.3903 0.00986 0.00346 0.0036 0.9620 0.5107
-3.000 -0.3570 0.00966 0.00335 0.0024 0.9521 0.5535
-2.750 -0.3232 0.00950 0.00323 0.0013 0.9418 0.5878
-2.500 -0.2909 0.00936 0.00314 0.0004 0.9295 0.6164
-2.250 -0.2586 0.00924 0.00306 -0.0003 0.9165 0.6420
-2.000 -0.2267 0.00914 0.00296 -0.0010 0.9026 0.6649
-1.750 -0.1956 0.00905 0.00288 -0.0014 0.8876 0.6839
-1.500 -0.1655 0.00897 0.00280 -0.0017 0.8721 0.7003
-1.250 -0.1365 0.00891 0.00272 -0.0017 0.8566 0.7149
-1.000 -0.1083 0.00887 0.00265 -0.0015 0.8407 0.7284
-0.750 -0.0808 0.00884 0.00259 -0.0012 0.8250 0.7417
-0.500 -0.0537 0.00882 0.00255 -0.0008 0.8099 0.7547
-0.250 -0.0268 0.00880 0.00253 -0.0004 0.7957 0.7679
0.000 0.0000 0.00880 0.00252 0.0000 0.7816 0.7816
0.250 0.0268 0.00880 0.00253 0.0004 0.7679 0.7957
0.500 0.0537 0.00882 0.00255 0.0008 0.7547 0.8099
0.750 0.0808 0.00884 0.00259 0.0012 0.7417 0.8250
1.000 0.1083 0.00887 0.00265 0.0015 0.7285 0.8407
1.250 0.1365 0.00891 0.00272 0.0017 0.7150 0.8566
1.500 0.1655 0.00897 0.00280 0.0017 0.7004 0.8722
1.750 0.1956 0.00905 0.00288 0.0014 0.6838 0.8877
2.000 0.2266 0.00914 0.00296 0.0010 0.6649 0.9027
2.250 0.2586 0.00924 0.00306 0.0003 0.6420 0.9166
2.500 0.2909 0.00936 0.00314 -0.0004 0.6164 0.9295
2.750 0.3232 0.00950 0.00323 -0.0013 0.5878 0.9419
3.000 0.3570 0.00966 0.00335 -0.0024 0.5537 0.9523
3.250 0.3902 0.00986 0.00346 -0.0036 0.5102 0.9622
3.500 0.4220 0.01018 0.00357 -0.0045 0.4421 0.9722
3.750 0.4508 0.01098 0.00379 -0.0053 0.3062 0.9821
4.000 0.4816 0.01196 0.00424 -0.0068 0.1877 0.9905
4.250 0.5138 0.01276 0.00469 -0.0085 0.1147 0.9986
4.500 0.5356 0.01333 0.00510 -0.0078 0.0817 1.0000
4.750 0.5551 0.01381 0.00552 -0.0064 0.0633 1.0000
5.000 0.5749 0.01431 0.00602 -0.0050 0.0523 1.0000
5.250 0.5944 0.01488 0.00654 -0.0036 0.0437 1.0000
5.500 0.6147 0.01537 0.00711 -0.0023 0.0386 1.0000
5.750 0.6331 0.01612 0.00785 -0.0008 0.0340 1.0000
6.000 0.6534 0.01667 0.00848 0.0005 0.0309 1.0000
6.250 0.6734 0.01727 0.00915 0.0019 0.0279 1.0000
6.500 0.6919 0.01810 0.01000 0.0033 0.0258 1.0000
6.750 0.7097 0.01915 0.01111 0.0050 0.0243 1.0000
7.000 0.7299 0.01995 0.01203 0.0062 0.0226 1.0000
7.250 0.7502 0.02069 0.01285 0.0074 0.0207 1.0000
7.500 0.7699 0.02157 0.01381 0.0085 0.0195 1.0000
7.750 0.7882 0.02283 0.01513 0.0098 0.0185 1.0000
8.000 0.8065 0.02444 0.01688 0.0112 0.0178 1.0000
8.250 0.8255 0.02595 0.01863 0.0125 0.0172 1.0000
8.500 0.8435 0.02760 0.02054 0.0138 0.0165 1.0000
8.750 0.8607 0.02903 0.02221 0.0151 0.0154 1.0000
9.000 0.8768 0.03042 0.02381 0.0164 0.0146 1.0000
9.250 0.8920 0.03170 0.02527 0.0176 0.0139 1.0000
9.500 0.9046 0.03336 0.02709 0.0191 0.0135 1.0000
9.750 0.9130 0.03572 0.02970 0.0208 0.0132 1.0000
10.000 0.9148 0.03875 0.03303 0.0230 0.0130 1.0000
10.250 0.9127 0.04171 0.03637 0.0254 0.0128 1.0000
10.500 0.9032 0.04483 0.03982 0.0283 0.0127 1.0000
10.750 0.8871 0.04800 0.04327 0.0312 0.0127 1.0000
11.000 0.8670 0.05189 0.04743 0.0324 0.0126 1.0000
11.250 0.8490 0.05613 0.05187 0.0318 0.0126 1.0000
11.500 0.8302 0.06122 0.05713 0.0295 0.0127 1.0000
11.750 0.8066 0.06814 0.06423 0.0249 0.0127 1.0000
12.000 0.7850 0.07608 0.07230 0.0190 0.0128 1.0000
12.250 0.7477 0.09153 0.08792 0.0081 0.0132 1.0000
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Polar data table (+)
Polar graphs
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