GOE 490 AIRFOIL (goe490-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 490 AIRFOIL (goe490-il) Reynolds number: 50,000 Max Cl/Cd: 38.25 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe490-il-50000-n5.txt Download as CSV file: xf-goe490-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 490 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3644 0.09746 0.09068 -0.0310 1.0000 0.0722
-8.000 -0.3686 0.09457 0.08788 -0.0309 1.0000 0.0718
-7.750 -0.3750 0.09182 0.08524 -0.0306 1.0000 0.0716
-7.500 -0.3806 0.08890 0.08242 -0.0303 1.0000 0.0704
-7.250 -0.3850 0.08567 0.07928 -0.0307 1.0000 0.0699
-7.000 -0.3892 0.08229 0.07599 -0.0313 1.0000 0.0685
-6.750 -0.3934 0.07858 0.07233 -0.0322 1.0000 0.0671
-6.500 -0.3967 0.07453 0.06830 -0.0335 1.0000 0.0657
-6.250 -0.3983 0.06998 0.06372 -0.0351 1.0000 0.0646
-6.000 -0.3950 0.06704 0.06077 -0.0346 1.0000 0.0658
-5.750 -0.3902 0.06425 0.05794 -0.0342 1.0000 0.0681
-5.500 -0.3843 0.06069 0.05426 -0.0346 1.0000 0.0698
-5.250 -0.3765 0.05664 0.05005 -0.0353 1.0000 0.0706
-5.000 -0.3664 0.05242 0.04559 -0.0359 1.0000 0.0708
-4.500 -0.3390 0.04355 0.03579 -0.0372 1.0000 0.0745
-4.250 -0.3236 0.04070 0.03261 -0.0368 1.0000 0.0772
-4.000 -0.3071 0.03856 0.03021 -0.0360 1.0000 0.0789
-3.750 -0.2850 0.03608 0.02732 -0.0362 0.9986 0.0805
-3.500 -0.2498 0.03344 0.02410 -0.0385 0.9924 0.0833
-3.250 -0.2129 0.03107 0.02092 -0.0407 0.9863 0.0893
-3.000 -0.1788 0.03024 0.02003 -0.0426 0.9784 0.0956
-2.750 -0.1421 0.02868 0.01798 -0.0444 0.9708 0.1007
-2.500 -0.1023 0.02748 0.01650 -0.0468 0.9620 0.1071
-2.250 -0.0598 0.02641 0.01509 -0.0494 0.9526 0.1182
-2.000 -0.0210 0.02572 0.01433 -0.0515 0.9413 0.1372
-1.750 0.0182 0.02512 0.01366 -0.0537 0.9305 0.1657
-1.500 0.0578 0.02480 0.01335 -0.0561 0.9208 0.2035
-1.250 0.0936 0.02462 0.01320 -0.0577 0.9102 0.2406
-1.000 0.1285 0.02426 0.01277 -0.0590 0.8993 0.2702
-0.750 0.1670 0.02375 0.01222 -0.0609 0.8899 0.3012
-0.500 0.2047 0.02311 0.01168 -0.0627 0.8803 0.3394
-0.250 0.2768 0.02060 0.01121 -0.0711 0.8763 0.9993
0.000 0.3166 0.02059 0.01082 -0.0730 0.8651 1.0000
0.250 0.3499 0.02062 0.01058 -0.0738 0.8519 1.0000
0.500 0.3827 0.02063 0.01039 -0.0744 0.8384 1.0000
0.750 0.4150 0.02064 0.01022 -0.0748 0.8246 1.0000
1.000 0.4467 0.02065 0.01009 -0.0751 0.8105 1.0000
1.250 0.4783 0.02064 0.00997 -0.0754 0.7958 1.0000
1.500 0.5094 0.02064 0.00987 -0.0755 0.7806 1.0000
1.750 0.5402 0.02065 0.00979 -0.0755 0.7649 1.0000
2.000 0.5707 0.02065 0.00971 -0.0754 0.7486 1.0000
2.250 0.6013 0.02065 0.00964 -0.0753 0.7319 1.0000
2.500 0.6269 0.02076 0.00969 -0.0743 0.7124 1.0000
2.750 0.6539 0.02085 0.00973 -0.0736 0.6926 1.0000
3.000 0.6819 0.02091 0.00971 -0.0729 0.6733 1.0000
3.250 0.7061 0.02110 0.00985 -0.0718 0.6525 1.0000
3.500 0.7309 0.02130 0.01002 -0.0707 0.6324 1.0000
3.750 0.7566 0.02150 0.01015 -0.0698 0.6134 1.0000
4.000 0.7794 0.02180 0.01044 -0.0686 0.5936 1.0000
4.250 0.8025 0.02211 0.01074 -0.0674 0.5744 1.0000
4.500 0.8262 0.02242 0.01105 -0.0663 0.5557 1.0000
4.750 0.8490 0.02276 0.01137 -0.0651 0.5370 1.0000
5.000 0.8704 0.02316 0.01179 -0.0637 0.5173 1.0000
5.250 0.8923 0.02355 0.01219 -0.0623 0.4981 1.0000
5.500 0.9143 0.02396 0.01256 -0.0610 0.4790 1.0000
5.750 0.9344 0.02446 0.01309 -0.0595 0.4589 1.0000
6.000 0.9552 0.02497 0.01362 -0.0580 0.4400 1.0000
6.250 0.9761 0.02553 0.01422 -0.0567 0.4218 1.0000
6.500 0.9967 0.02613 0.01485 -0.0553 0.4046 1.0000
6.750 1.0159 0.02680 0.01562 -0.0539 0.3871 1.0000
7.000 1.0353 0.02749 0.01642 -0.0524 0.3706 1.0000
7.250 1.0545 0.02820 0.01725 -0.0510 0.3552 1.0000
7.500 1.0734 0.02894 0.01816 -0.0495 0.3405 1.0000
7.750 1.0918 0.02969 0.01904 -0.0480 0.3263 1.0000
8.000 1.1058 0.03035 0.01980 -0.0458 0.3091 1.0000
8.250 1.1167 0.03095 0.02043 -0.0432 0.2906 1.0000
8.500 1.1276 0.03162 0.02113 -0.0406 0.2736 1.0000
8.750 1.1402 0.03244 0.02208 -0.0385 0.2597 1.0000
9.000 1.1527 0.03331 0.02313 -0.0363 0.2469 1.0000
9.250 1.1577 0.03418 0.02409 -0.0333 0.2297 1.0000
9.500 1.1631 0.03516 0.02520 -0.0304 0.2152 1.0000
9.750 1.1676 0.03624 0.02645 -0.0274 0.2010 1.0000
10.000 1.1640 0.03761 0.02787 -0.0241 0.1765 1.0000
10.250 1.1596 0.03936 0.02958 -0.0213 0.1471 1.0000
10.500 1.1571 0.04142 0.03164 -0.0191 0.1185 1.0000
10.750 1.1503 0.04414 0.03421 -0.0172 0.0736 1.0000
11.000 1.1370 0.04775 0.03752 -0.0158 0.0596 1.0000
11.250 1.1253 0.05158 0.04125 -0.0150 0.0531 1.0000
11.500 1.1179 0.05521 0.04504 -0.0146 0.0489 1.0000
11.750 1.1087 0.05927 0.04925 -0.0146 0.0460 1.0000
12.000 1.0985 0.06369 0.05381 -0.0152 0.0441 1.0000
12.250 1.0881 0.06833 0.05859 -0.0161 0.0428 1.0000
12.500 1.0778 0.07318 0.06354 -0.0173 0.0415 1.0000
12.750 1.0710 0.07766 0.06821 -0.0184 0.0405 1.0000
13.000 1.0644 0.08220 0.07293 -0.0195 0.0395 1.0000
13.250 1.0581 0.08677 0.07766 -0.0208 0.0384 1.0000
13.500 1.0523 0.09134 0.08237 -0.0221 0.0373 1.0000
13.750 1.0467 0.09593 0.08708 -0.0234 0.0361 1.0000
14.000 1.0423 0.10037 0.09162 -0.0248 0.0352 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 490 AIRFOIL (goe490-il)