GOE 415 AIRFOIL (goe415-il) Xfoil prediction polar at RE=200,000 Ncrit=5
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Airfoil: GOE 415 AIRFOIL (goe415-il) Reynolds number: 200,000 Max Cl/Cd: 69.55 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe415-il-200000-n5.txt Download as CSV file: xf-goe415-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 415 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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.2766 0.08471 0.08166 -0.0178 1.0000 0.0202
-8.000 -0.2757 0.08138 0.07837 -0.0195 1.0000 0.0213
-7.750 -0.2754 0.07808 0.07511 -0.0215 1.0000 0.0218
-7.500 -0.2752 0.07475 0.07182 -0.0232 1.0000 0.0220
-7.250 -0.2757 0.07141 0.06853 -0.0246 1.0000 0.0221
-7.000 -0.2785 0.06824 0.06543 -0.0255 1.0000 0.0222
-6.750 -0.2805 0.06496 0.06220 -0.0269 1.0000 0.0223
-6.500 -0.2821 0.06158 0.05887 -0.0285 1.0000 0.0223
-6.250 -0.2657 0.05612 0.05338 -0.0350 0.9928 0.0224
-6.000 -0.2444 0.05024 0.04743 -0.0419 0.9843 0.0224
-5.750 -0.2219 0.04432 0.04140 -0.0481 0.9751 0.0225
-5.500 -0.1981 0.03863 0.03555 -0.0533 0.9657 0.0225
-5.250 -0.1740 0.03326 0.02998 -0.0576 0.9554 0.0226
-5.000 -0.1494 0.02835 0.02486 -0.0610 0.9442 0.0226
-4.750 -0.1278 0.02298 0.01935 -0.0636 0.9330 0.0220
-4.500 -0.1360 0.03563 0.03161 -0.0655 0.9506 0.0159
-4.250 -0.0942 0.03064 0.02593 -0.0674 0.9402 0.0123
-3.750 -0.0318 0.02462 0.01905 -0.0697 0.9130 0.0122
-3.500 -0.0046 0.02131 0.01531 -0.0703 0.8974 0.0137
-3.250 0.0228 0.01962 0.01328 -0.0700 0.8788 0.0145
-3.000 0.0505 0.01789 0.01115 -0.0693 0.8583 0.0147
-2.750 0.0774 0.01649 0.00943 -0.0684 0.8328 0.0151
-2.500 0.1039 0.01535 0.00797 -0.0674 0.7995 0.0160
-2.250 0.1298 0.01463 0.00695 -0.0663 0.7598 0.0186
-2.000 0.1553 0.01390 0.00591 -0.0652 0.7223 0.0205
-1.750 0.1809 0.01334 0.00508 -0.0643 0.6889 0.0221
-1.500 0.2062 0.01271 0.00429 -0.0636 0.6582 0.0266
-1.250 0.2324 0.01238 0.00373 -0.0631 0.6295 0.0309
-1.000 0.2588 0.01211 0.00324 -0.0627 0.6039 0.0382
-0.750 0.2856 0.01199 0.00287 -0.0622 0.5820 0.0451
-0.500 0.3110 0.01118 0.00261 -0.0621 0.5638 0.2387
-0.250 0.3351 0.01054 0.00260 -0.0616 0.5477 0.4761
0.000 0.3704 0.00934 0.00252 -0.0626 0.5318 1.0000
0.250 0.3973 0.00950 0.00245 -0.0623 0.5178 1.0000
0.500 0.4242 0.00966 0.00242 -0.0621 0.5048 1.0000
0.750 0.4511 0.00982 0.00241 -0.0618 0.4928 1.0000
1.000 0.4781 0.00998 0.00242 -0.0616 0.4812 1.0000
1.250 0.5051 0.01013 0.00245 -0.0614 0.4691 1.0000
1.500 0.5321 0.01029 0.00249 -0.0612 0.4568 1.0000
1.750 0.5590 0.01045 0.00255 -0.0610 0.4443 1.0000
2.000 0.5859 0.01061 0.00262 -0.0608 0.4320 1.0000
2.250 0.6127 0.01079 0.00272 -0.0606 0.4203 1.0000
2.500 0.6395 0.01098 0.00282 -0.0604 0.4089 1.0000
2.750 0.6662 0.01116 0.00294 -0.0602 0.3969 1.0000
3.000 0.6928 0.01136 0.00309 -0.0600 0.3845 1.0000
3.250 0.7194 0.01157 0.00324 -0.0598 0.3720 1.0000
3.500 0.7457 0.01179 0.00341 -0.0595 0.3598 1.0000
3.750 0.7720 0.01203 0.00359 -0.0593 0.3481 1.0000
4.000 0.7982 0.01228 0.00382 -0.0591 0.3380 1.0000
4.250 0.8247 0.01251 0.00406 -0.0589 0.3287 1.0000
4.500 0.8509 0.01277 0.00431 -0.0586 0.3211 1.0000
4.750 0.8770 0.01303 0.00460 -0.0584 0.3121 1.0000
5.000 0.9031 0.01329 0.00489 -0.0582 0.3024 1.0000
5.250 0.9289 0.01359 0.00519 -0.0579 0.2939 1.0000
5.500 0.9549 0.01385 0.00554 -0.0577 0.2854 1.0000
5.750 0.9807 0.01414 0.00589 -0.0574 0.2767 1.0000
6.000 1.0057 0.01446 0.00619 -0.0571 0.2569 1.0000
6.250 1.0303 0.01484 0.00653 -0.0567 0.2353 1.0000
6.500 1.0534 0.01541 0.00692 -0.0563 0.1923 1.0000
6.750 1.0714 0.01679 0.00779 -0.0555 0.1132 1.0000
7.000 1.0838 0.01909 0.00949 -0.0539 0.0190 1.0000
7.250 1.1062 0.01985 0.01042 -0.0531 0.0152 1.0000
7.500 1.1278 0.02068 0.01146 -0.0523 0.0127 1.0000
7.750 1.1478 0.02169 0.01270 -0.0512 0.0110 1.0000
8.000 1.1655 0.02297 0.01424 -0.0500 0.0101 1.0000
8.250 1.1787 0.02470 0.01627 -0.0483 0.0094 1.0000
8.500 1.1938 0.02600 0.01779 -0.0469 0.0087 1.0000
8.750 1.2086 0.02723 0.01916 -0.0455 0.0079 1.0000
9.000 1.2187 0.02880 0.02087 -0.0437 0.0074 1.0000
9.250 1.2255 0.03054 0.02273 -0.0416 0.0072 1.0000
9.500 1.2288 0.03232 0.02461 -0.0391 0.0070 1.0000
9.750 1.2302 0.03423 0.02662 -0.0367 0.0068 1.0000
10.000 1.2311 0.03635 0.02884 -0.0347 0.0066 1.0000
10.250 1.2322 0.03865 0.03122 -0.0330 0.0065 1.0000
10.500 1.2341 0.04106 0.03371 -0.0317 0.0064 1.0000
10.750 1.2372 0.04355 0.03628 -0.0304 0.0063 1.0000
11.000 1.2421 0.04607 0.03885 -0.0290 0.0062 1.0000
11.250 1.2493 0.04879 0.04161 -0.0275 0.0060 1.0000
11.500 1.2518 0.05169 0.04468 -0.0266 0.0058 1.0000
11.750 1.2502 0.05462 0.04787 -0.0267 0.0056 1.0000
12.000 1.2485 0.05788 0.05137 -0.0271 0.0053 1.0000
12.250 1.2465 0.06141 0.05511 -0.0275 0.0051 1.0000
12.500 1.2430 0.06524 0.05915 -0.0282 0.0050 1.0000
12.750 1.2383 0.06938 0.06349 -0.0291 0.0049 1.0000
13.000 1.2320 0.07383 0.06814 -0.0303 0.0049 1.0000
13.250 1.2243 0.07861 0.07311 -0.0319 0.0049 1.0000
13.500 1.2152 0.08369 0.07839 -0.0338 0.0049 1.0000
13.750 1.2048 0.08905 0.08394 -0.0361 0.0049 1.0000
14.000 1.1936 0.09474 0.08980 -0.0387 0.0049 1.0000
14.250 1.1821 0.10070 0.09594 -0.0416 0.0049 1.0000
14.500 1.1701 0.10696 0.10237 -0.0449 0.0049 1.0000
14.750 1.1582 0.11349 0.10907 -0.0485 0.0050 1.0000
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Polar data table (+)
Polar graphs
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