GOE 775 AIRFOIL (goe775-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 775 AIRFOIL (goe775-il) Reynolds number: 50,000 Max Cl/Cd: 18.85 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe775-il-50000.txt Download as CSV file: xf-goe775-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 775 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.2941 0.11861 0.10934 -0.0089 1.0000 0.4776
-9.500 -0.2700 0.11551 0.10623 -0.0093 1.0000 0.4889
-9.250 -0.3055 0.11536 0.10611 -0.0069 1.0000 0.4958
-9.000 -0.2573 0.11013 0.10087 -0.0089 1.0000 0.5048
-8.750 -0.2718 0.10903 0.09980 -0.0072 1.0000 0.5156
-8.500 -0.2788 0.10260 0.09335 -0.0096 1.0000 0.4824
-8.250 -0.3976 0.08671 0.07741 -0.0149 1.0000 0.4026
-8.000 -0.4322 0.08147 0.07222 -0.0136 1.0000 0.4080
-7.750 -0.6257 0.06372 0.05465 -0.0082 1.0000 0.4114
-7.500 -0.8111 0.05270 0.04363 0.0100 1.0000 0.4114
-7.250 -0.5891 0.06355 0.05453 -0.0036 1.0000 0.4300
-7.000 -0.5999 0.06164 0.05268 0.0004 1.0000 0.4380
-6.750 -0.7164 0.05460 0.04573 0.0131 1.0000 0.4424
-6.500 -0.7982 0.04916 0.04019 0.0237 1.0000 0.4493
-6.250 -0.8066 0.04763 0.03865 0.0281 1.0000 0.4586
-6.000 -0.8015 0.04704 0.03807 0.0313 1.0000 0.4693
-5.750 -0.8102 0.04548 0.03648 0.0356 1.0000 0.4794
-5.500 -0.8019 0.04510 0.03612 0.0384 1.0000 0.4908
-5.250 -0.7984 0.04430 0.03532 0.0415 1.0000 0.5017
-5.000 -0.7991 0.04329 0.03426 0.0449 1.0000 0.5142
-4.750 -0.7825 0.04340 0.03445 0.0469 1.0000 0.5251
-4.500 -0.7927 0.04157 0.03247 0.0506 1.0000 0.5393
-4.250 -0.7660 0.04251 0.03358 0.0518 1.0000 0.5497
-4.000 -0.7665 0.04130 0.03230 0.0547 1.0000 0.5635
-3.750 -0.7508 0.04146 0.03252 0.0565 1.0000 0.5759
-3.500 -0.7417 0.04100 0.03207 0.0587 1.0000 0.5888
-3.250 -0.7393 0.03998 0.03095 0.0610 1.0000 0.6047
-3.000 -0.7184 0.04066 0.03176 0.0625 1.0000 0.6158
-2.750 -0.6825 0.04092 0.03201 0.0599 0.9922 0.6331
-2.500 -0.6320 0.04154 0.03262 0.0551 0.9781 0.6526
-2.250 -0.5891 0.04185 0.03288 0.0514 0.9635 0.6720
-2.000 -0.5325 0.04336 0.03447 0.0474 0.9486 0.6866
-1.750 -0.4872 0.04407 0.03520 0.0443 0.9334 0.7029
-1.500 -0.4485 0.04449 0.03560 0.0420 0.9188 0.7205
-1.250 -0.4045 0.04497 0.03605 0.0389 0.9046 0.7392
-1.000 -0.3166 0.04710 0.03822 0.0305 0.8895 0.7521
-0.750 -0.2767 0.04770 0.03882 0.0285 0.8735 0.7673
-0.500 -0.2367 0.04806 0.03916 0.0261 0.8586 0.7845
-0.250 -0.1100 0.04948 0.04056 0.0112 0.8453 0.7987
0.000 -0.0265 0.05005 0.04115 0.0021 0.8285 0.8118
0.250 0.0137 0.05006 0.04116 -0.0005 0.8124 0.8280
0.500 0.1178 0.04942 0.04050 -0.0126 0.7998 0.8445
0.750 0.2298 0.04822 0.03932 -0.0255 0.7849 0.8580
1.000 0.2692 0.04787 0.03899 -0.0278 0.7679 0.8730
1.250 0.3131 0.04717 0.03828 -0.0304 0.7532 0.8893
1.500 0.4006 0.04513 0.03621 -0.0388 0.7398 0.9041
1.750 0.4445 0.04465 0.03576 -0.0419 0.7213 0.9185
2.000 0.4841 0.04420 0.03532 -0.0443 0.7039 0.9332
2.250 0.5321 0.04333 0.03443 -0.0476 0.6882 0.9484
2.500 0.5859 0.04201 0.03306 -0.0515 0.6728 0.9634
2.750 0.6293 0.04169 0.03278 -0.0552 0.6536 0.9780
3.000 0.6806 0.04103 0.03213 -0.0602 0.6343 0.9921
3.250 0.7176 0.04068 0.03179 -0.0628 0.6173 1.0000
3.500 0.7368 0.04013 0.03113 -0.0612 0.6057 1.0000
3.750 0.7397 0.04112 0.03219 -0.0589 0.5901 1.0000
4.000 0.7518 0.04132 0.03235 -0.0570 0.5778 1.0000
4.250 0.7633 0.04149 0.03251 -0.0549 0.5647 1.0000
4.500 0.7649 0.04254 0.03361 -0.0521 0.5513 1.0000
4.750 0.7884 0.04182 0.03275 -0.0507 0.5404 1.0000
5.000 0.7792 0.04357 0.03463 -0.0470 0.5266 1.0000
5.250 0.8050 0.04281 0.03372 -0.0457 0.5161 1.0000
5.500 0.7924 0.04463 0.03568 -0.0414 0.5032 1.0000
5.750 0.8067 0.04472 0.03570 -0.0392 0.4925 1.0000
6.000 0.8002 0.04604 0.03707 -0.0351 0.4810 1.0000
6.250 0.8074 0.04661 0.03761 -0.0322 0.4708 1.0000
6.500 0.7902 0.04851 0.03957 -0.0271 0.4604 1.0000
6.750 0.8073 0.04857 0.03956 -0.0249 0.4505 1.0000
7.000 0.7074 0.05503 0.04614 -0.0127 0.4442 1.0000
7.250 0.6283 0.05997 0.05103 -0.0027 0.4383 1.0000
7.500 0.6314 0.06091 0.05192 0.0007 0.4298 1.0000
7.750 0.8257 0.05187 0.04275 -0.0117 0.4120 1.0000
8.000 0.6591 0.06164 0.05258 0.0058 0.4114 1.0000
8.250 0.4385 0.08085 0.07158 0.0134 0.4087 1.0000
8.500 0.3993 0.08647 0.07715 0.0146 0.4035 1.0000
8.750 0.2858 0.10634 0.09713 0.0061 0.5168 1.0000
9.000 0.2752 0.10897 0.09973 0.0069 0.5160 1.0000
9.250 0.2575 0.10999 0.10072 0.0087 0.5058 1.0000
9.500 0.3088 0.11527 0.10602 0.0066 0.4960 1.0000
9.750 0.2727 0.11551 0.10622 0.0089 0.4902 1.0000
10.000 0.2916 0.11825 0.10896 0.0088 0.4784 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 775 AIRFOIL (goe775-il)