GOE 442 AIRFOIL (goe442-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 442 AIRFOIL (goe442-il) Reynolds number: 500,000 Max Cl/Cd: 97.61 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe442-il-500000.txt Download as CSV file: xf-goe442-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 442 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.500 -0.3731 0.08544 0.08333 -0.0186 1.0000 0.0200
-7.250 -0.3858 0.08380 0.08175 -0.0162 1.0000 0.0203
-7.000 -0.3922 0.08109 0.07908 -0.0165 1.0000 0.0211
-6.750 -0.3949 0.07732 0.07533 -0.0205 0.9999 0.0216
-6.500 -0.3578 0.07031 0.06826 -0.0333 0.9963 0.0219
-6.250 -0.3388 0.06465 0.06259 -0.0382 0.9936 0.0226
-6.000 -0.3126 0.06135 0.05925 -0.0423 0.9901 0.0231
-5.750 -0.2798 0.05705 0.05489 -0.0488 0.9868 0.0241
-5.500 -0.2400 0.05145 0.04918 -0.0577 0.9842 0.0256
-5.250 -0.1915 0.04233 0.03966 -0.0685 0.9783 0.0278
-5.000 -0.1656 0.03003 0.02685 -0.0752 0.9737 0.0229
-4.750 -0.1292 0.02903 0.02579 -0.0781 0.9718 0.0244
-4.500 -0.1108 0.01558 0.01055 -0.0778 0.9651 0.0226
-4.250 -0.0758 0.01472 0.00959 -0.0793 0.9618 0.0238
-4.000 -0.0384 0.01384 0.00855 -0.0811 0.9595 0.0256
-3.750 0.0012 0.01335 0.00788 -0.0832 0.9578 0.0279
-3.500 0.0315 0.01226 0.00673 -0.0837 0.9529 0.0308
-3.250 0.0642 0.01179 0.00618 -0.0844 0.9478 0.0336
-3.000 0.0994 0.01095 0.00524 -0.0856 0.9439 0.0373
-2.750 0.1286 0.01074 0.00504 -0.0856 0.9364 0.0407
-2.500 0.1612 0.01043 0.00467 -0.0862 0.9295 0.0442
-2.250 0.1886 0.00976 0.00396 -0.0859 0.9192 0.0481
-2.000 0.2185 0.00945 0.00364 -0.0860 0.9082 0.0518
-1.750 0.2479 0.00923 0.00337 -0.0859 0.8949 0.0556
-1.500 0.2759 0.00882 0.00290 -0.0856 0.8778 0.0586
-1.250 0.3039 0.00847 0.00250 -0.0853 0.8564 0.0635
-1.000 0.3310 0.00830 0.00224 -0.0847 0.8293 0.0677
-0.750 0.3572 0.00824 0.00203 -0.0840 0.7976 0.0710
-0.500 0.3815 0.00811 0.00175 -0.0829 0.7593 0.0789
-0.250 0.4047 0.00821 0.00161 -0.0816 0.7114 0.0851
0.000 0.4269 0.00827 0.00150 -0.0802 0.6592 0.1027
0.250 0.4474 0.00795 0.00156 -0.0788 0.6103 0.3042
0.500 0.5049 0.00655 0.00170 -0.0857 0.5689 1.0000
1.000 0.5522 0.00698 0.00178 -0.0837 0.5248 1.0000
1.250 0.5767 0.00715 0.00183 -0.0828 0.5098 1.0000
1.500 0.6013 0.00732 0.00189 -0.0820 0.4947 1.0000
1.750 0.6260 0.00748 0.00196 -0.0812 0.4799 1.0000
2.000 0.6508 0.00764 0.00203 -0.0805 0.4658 1.0000
2.250 0.6757 0.00779 0.00210 -0.0797 0.4517 1.0000
2.750 0.7259 0.00808 0.00226 -0.0784 0.4251 1.0000
3.000 0.7511 0.00822 0.00236 -0.0777 0.4133 1.0000
3.250 0.7763 0.00837 0.00246 -0.0771 0.4016 1.0000
3.500 0.8013 0.00854 0.00257 -0.0764 0.3897 1.0000
3.750 0.8261 0.00872 0.00271 -0.0757 0.3776 1.0000
4.000 0.8512 0.00889 0.00284 -0.0751 0.3651 1.0000
4.250 0.8762 0.00906 0.00298 -0.0745 0.3519 1.0000
4.500 0.9009 0.00926 0.00313 -0.0738 0.3371 1.0000
4.750 0.9254 0.00949 0.00331 -0.0731 0.3211 1.0000
5.000 0.9497 0.00973 0.00349 -0.0724 0.3034 1.0000
5.250 0.9737 0.01000 0.00369 -0.0716 0.2796 1.0000
5.500 0.9964 0.01042 0.00395 -0.0707 0.2425 1.0000
5.750 1.0169 0.01107 0.00433 -0.0695 0.1926 1.0000
6.000 1.0385 0.01163 0.00473 -0.0684 0.1685 1.0000
6.250 1.0613 0.01205 0.00511 -0.0675 0.1578 1.0000
6.500 1.0837 0.01250 0.00553 -0.0665 0.1505 1.0000
6.750 1.1071 0.01284 0.00590 -0.0657 0.1458 1.0000
7.000 1.1303 0.01319 0.00628 -0.0649 0.1414 1.0000
7.250 1.1517 0.01370 0.00676 -0.0638 0.1357 1.0000
7.500 1.1757 0.01395 0.00708 -0.0632 0.1308 1.0000
7.750 1.1984 0.01431 0.00744 -0.0623 0.1248 1.0000
8.000 1.2202 0.01475 0.00789 -0.0614 0.1199 1.0000
8.250 1.2441 0.01497 0.00820 -0.0607 0.1158 1.0000
8.500 1.2662 0.01535 0.00860 -0.0598 0.1110 1.0000
8.750 1.2880 0.01575 0.00904 -0.0589 0.1061 1.0000
9.000 1.3120 0.01594 0.00931 -0.0583 0.1008 1.0000
9.250 1.3335 0.01633 0.00970 -0.0574 0.0934 1.0000
9.500 1.3569 0.01658 0.00995 -0.0567 0.0786 1.0000
9.750 1.3657 0.01810 0.01106 -0.0541 0.0334 1.0000
10.000 1.3775 0.01931 0.01225 -0.0516 0.0238 1.0000
10.250 1.3921 0.02019 0.01321 -0.0496 0.0207 1.0000
10.500 1.4021 0.02138 0.01448 -0.0470 0.0184 1.0000
10.750 1.4152 0.02215 0.01536 -0.0447 0.0173 1.0000
11.000 1.4247 0.02303 0.01633 -0.0420 0.0161 1.0000
11.250 1.4310 0.02408 0.01746 -0.0389 0.0151 1.0000
11.500 1.4303 0.02556 0.01905 -0.0351 0.0143 1.0000
11.750 1.4299 0.02709 0.02071 -0.0316 0.0138 1.0000
12.000 1.4347 0.02835 0.02208 -0.0290 0.0134 1.0000
12.250 1.4373 0.02982 0.02368 -0.0265 0.0130 1.0000
12.500 1.4383 0.03150 0.02548 -0.0242 0.0126 1.0000
12.750 1.4371 0.03350 0.02759 -0.0221 0.0122 1.0000
13.000 1.4357 0.03567 0.02988 -0.0205 0.0119 1.0000
13.250 1.4325 0.03821 0.03254 -0.0192 0.0117 1.0000
13.500 1.4286 0.04105 0.03551 -0.0185 0.0114 1.0000
13.750 1.4234 0.04429 0.03886 -0.0183 0.0112 1.0000
14.000 1.4135 0.04837 0.04305 -0.0187 0.0109 1.0000
14.250 1.4061 0.05240 0.04721 -0.0194 0.0108 1.0000
14.500 1.3957 0.05703 0.05197 -0.0205 0.0107 1.0000
14.750 1.3775 0.06280 0.05788 -0.0218 0.0103 1.0000
15.000 1.3722 0.06707 0.06228 -0.0233 0.0102 1.0000
15.250 1.3674 0.07148 0.06683 -0.0251 0.0101 1.0000
15.500 1.3605 0.07620 0.07168 -0.0270 0.0100 1.0000
15.750 1.3518 0.08126 0.07688 -0.0290 0.0099 1.0000
16.000 1.3435 0.08648 0.08223 -0.0312 0.0098 1.0000
16.250 1.3349 0.09187 0.08775 -0.0336 0.0096 1.0000
16.500 1.3253 0.09746 0.09347 -0.0361 0.0096 1.0000
16.750 1.3159 0.10326 0.09940 -0.0389 0.0095 1.0000
17.000 1.3064 0.10918 0.10545 -0.0417 0.0094 1.0000
17.250 1.2968 0.11527 0.11167 -0.0449 0.0093 1.0000
17.500 1.2875 0.12149 0.11800 -0.0482 0.0092 1.0000
17.750 1.2771 0.12799 0.12464 -0.0516 0.0093 1.0000
18.000 1.2673 0.13462 0.13140 -0.0554 0.0092 1.0000
18.250 1.2558 0.14171 0.13862 -0.0595 0.0092 1.0000
18.500 1.2456 0.14875 0.14578 -0.0637 0.0091 1.0000
18.750 1.2337 0.15639 0.15356 -0.0684 0.0091 1.0000
19.000 1.2209 0.16454 0.16184 -0.0735 0.0091 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 442 AIRFOIL (goe442-il)