GOE 438 AIRFOIL (goe438-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 438 AIRFOIL (goe438-il) Reynolds number: 50,000 Max Cl/Cd: 24.65 at α=11.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe438-il-50000.txt Download as CSV file: xf-goe438-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 438 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3504 0.11003 0.10290 -0.0262 1.0000 0.2359
-8.750 -0.3259 0.10477 0.09760 -0.0244 1.0000 0.2479
-8.500 -0.3602 0.10588 0.09894 -0.0247 1.0000 0.2530
-8.250 -0.3352 0.10084 0.09386 -0.0229 1.0000 0.2678
-8.000 -0.3241 0.09715 0.09021 -0.0215 1.0000 0.2776
-7.750 -0.3557 0.09769 0.09096 -0.0200 1.0000 0.2860
-7.500 -0.3382 0.09370 0.08698 -0.0182 1.0000 0.3009
-7.250 -0.3327 0.09088 0.08423 -0.0160 1.0000 0.3154
-7.000 -0.3792 0.09201 0.08562 -0.0122 1.0000 0.3197
-6.750 -0.3425 0.08683 0.08039 -0.0107 1.0000 0.3421
-6.500 -0.3772 0.08713 0.08089 -0.0061 1.0000 0.3524
-6.250 -0.3792 0.08494 0.07879 -0.0028 1.0000 0.3688
-6.000 -0.3862 0.08311 0.07706 0.0007 1.0000 0.3856
-5.750 -0.3987 0.08167 0.07573 0.0044 1.0000 0.4030
-5.500 -0.3746 0.07822 0.07227 0.0080 1.0000 0.4302
-5.250 -0.3999 0.07769 0.07188 0.0129 1.0000 0.4528
-5.000 -0.3995 0.07591 0.07018 0.0180 1.0000 0.4865
-4.750 -0.3516 0.07195 0.06617 0.0209 1.0000 0.5450
-4.500 -0.3128 0.06894 0.06311 0.0236 1.0000 0.6172
-4.250 0.0956 0.04616 0.03921 -0.0330 1.0000 1.0000
-4.000 0.0986 0.04506 0.03820 -0.0321 1.0000 1.0000
-3.750 0.0340 0.04799 0.04143 -0.0158 1.0000 0.9713
-3.500 -0.0498 0.05127 0.04505 0.0026 1.0000 0.9244
-3.250 -0.1282 0.05352 0.04758 0.0173 1.0000 0.8753
-3.000 -0.1990 0.05475 0.04908 0.0287 1.0000 0.8293
-2.750 -0.3087 0.04475 0.03699 -0.0233 1.0000 0.2514
-2.500 -0.2845 0.04217 0.03402 -0.0247 1.0000 0.2303
-2.250 -0.2602 0.04025 0.03156 -0.0256 1.0000 0.2155
-2.000 -0.2386 0.03867 0.02957 -0.0259 1.0000 0.2073
-1.750 -0.2064 0.03789 0.02818 -0.0278 0.9961 0.2034
-1.500 -0.1525 0.03705 0.02690 -0.0334 0.9826 0.2028
-1.250 -0.1038 0.03618 0.02574 -0.0378 0.9683 0.2017
-1.000 -0.0560 0.03561 0.02486 -0.0419 0.9534 0.2020
-0.750 -0.0089 0.03504 0.02418 -0.0459 0.9383 0.2075
-0.500 0.0369 0.03470 0.02372 -0.0493 0.9228 0.2149
-0.250 0.0851 0.03441 0.02325 -0.0528 0.9073 0.2213
0.000 0.1332 0.03407 0.02289 -0.0562 0.8916 0.2326
0.250 0.1780 0.03371 0.02262 -0.0591 0.8758 0.2546
0.500 0.2216 0.03307 0.02234 -0.0617 0.8603 0.2967
0.750 0.2983 0.03062 0.02161 -0.0686 0.8466 1.0000
1.000 0.3416 0.03090 0.02153 -0.0708 0.8301 1.0000
1.250 0.3850 0.03105 0.02146 -0.0730 0.8135 1.0000
1.500 0.4266 0.03117 0.02141 -0.0748 0.7973 1.0000
1.750 0.4672 0.03120 0.02132 -0.0762 0.7810 1.0000
2.000 0.5035 0.03128 0.02129 -0.0769 0.7650 1.0000
2.250 0.5387 0.03132 0.02125 -0.0773 0.7490 1.0000
2.500 0.5713 0.03142 0.02128 -0.0773 0.7332 1.0000
2.750 0.6019 0.03159 0.02138 -0.0770 0.7177 1.0000
3.000 0.6312 0.03179 0.02154 -0.0765 0.7024 1.0000
3.250 0.6583 0.03212 0.02183 -0.0758 0.6874 1.0000
3.500 0.6833 0.03258 0.02227 -0.0749 0.6726 1.0000
3.750 0.7071 0.03315 0.02282 -0.0739 0.6583 1.0000
4.000 0.7302 0.03382 0.02348 -0.0729 0.6447 1.0000
4.250 0.7586 0.03423 0.02388 -0.0723 0.6324 1.0000
4.500 0.7971 0.03407 0.02371 -0.0727 0.6215 1.0000
4.750 0.8058 0.03569 0.02536 -0.0705 0.6077 1.0000
5.000 0.8160 0.03731 0.02702 -0.0686 0.5949 1.0000
5.250 0.8426 0.03797 0.02770 -0.0680 0.5844 1.0000
5.500 0.8653 0.03887 0.02866 -0.0671 0.5737 1.0000
5.750 0.8603 0.04164 0.03149 -0.0644 0.5615 1.0000
6.000 0.8801 0.04284 0.03274 -0.0633 0.5518 1.0000
6.250 0.8954 0.04434 0.03430 -0.0621 0.5418 1.0000
6.500 0.8642 0.04930 0.03929 -0.0585 0.5306 1.0000
6.750 0.9367 0.04668 0.03677 -0.0601 0.5230 1.0000
7.000 0.8495 0.05596 0.04600 -0.0545 0.5117 1.0000
7.250 0.8268 0.06116 0.05120 -0.0534 0.5029 1.0000
7.500 0.8475 0.06244 0.05257 -0.0525 0.4937 1.0000
7.750 0.8089 0.06925 0.05934 -0.0521 0.4863 1.0000
8.000 0.8699 0.06688 0.05713 -0.0508 0.4761 1.0000
8.250 0.8048 0.07631 0.06647 -0.0513 0.4704 1.0000
8.500 0.8598 0.07477 0.06510 -0.0500 0.4599 1.0000
8.750 0.8075 0.08317 0.07343 -0.0510 0.4562 1.0000
9.000 0.7911 0.08837 0.07866 -0.0516 0.4529 1.0000
9.250 0.8192 0.08965 0.08005 -0.0508 0.4426 1.0000
9.500 0.8001 0.09535 0.08577 -0.0519 0.4412 1.0000
9.750 0.7984 0.10022 0.09071 -0.0531 0.4427 1.0000
10.000 0.8045 0.10482 0.09539 -0.0542 0.4441 1.0000
11.500 1.3206 0.05357 0.04539 -0.0265 0.2559 1.0000
11.750 1.2728 0.05854 0.05062 -0.0209 0.2538 1.0000
12.000 1.2161 0.06578 0.05802 -0.0181 0.2539 1.0000
12.250 1.1452 0.07726 0.06953 -0.0196 0.2560 1.0000
12.500 1.1428 0.07967 0.07198 -0.0184 0.2410 1.0000
12.750 1.2253 0.07082 0.06311 -0.0125 0.2104 1.0000
13.000 0.8951 0.13318 0.12481 -0.0462 0.2535 1.0000
13.250 0.8746 0.14105 0.13267 -0.0497 0.2504 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 438 AIRFOIL (goe438-il)