GOE 624 AIRFOIL (goe624-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 624 AIRFOIL (goe624-il) Reynolds number: 500,000 Max Cl/Cd: 97.55 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe624-il-500000-n5.txt Download as CSV file: xf-goe624-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 624 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-14.000 -0.7846 0.04559 0.04198 -0.0967 0.9709 0.0252
-13.750 -0.7977 0.03868 0.03478 -0.1051 0.9437 0.0252
-13.500 -0.8044 0.03555 0.03144 -0.1065 0.9229 0.0253
-13.000 -0.8036 0.03195 0.02748 -0.1043 0.8923 0.0257
-12.750 -0.7944 0.03048 0.02584 -0.1034 0.8818 0.0259
-12.500 -0.7826 0.02910 0.02429 -0.1026 0.8721 0.0260
-12.250 -0.7682 0.02783 0.02285 -0.1018 0.8638 0.0262
-12.000 -0.7521 0.02667 0.02151 -0.1011 0.8554 0.0265
-11.750 -0.7343 0.02559 0.02027 -0.1004 0.8479 0.0267
-11.500 -0.7150 0.02459 0.01910 -0.0998 0.8400 0.0269
-11.000 -0.6733 0.02282 0.01703 -0.0987 0.8265 0.0273
-10.750 -0.6510 0.02205 0.01610 -0.0982 0.8199 0.0275
-10.500 -0.6279 0.02138 0.01528 -0.0978 0.8139 0.0277
-10.250 -0.6052 0.02052 0.01434 -0.0974 0.8075 0.0281
-10.000 -0.5820 0.01977 0.01351 -0.0970 0.8007 0.0284
-9.750 -0.5579 0.01916 0.01281 -0.0966 0.7945 0.0286
-9.500 -0.5329 0.01856 0.01215 -0.0963 0.7880 0.0289
-9.250 -0.5078 0.01802 0.01153 -0.0961 0.7817 0.0292
-9.000 -0.4823 0.01751 0.01094 -0.0958 0.7753 0.0295
-8.750 -0.4564 0.01702 0.01037 -0.0955 0.7669 0.0298
-8.500 -0.4306 0.01657 0.00981 -0.0952 0.7585 0.0301
-8.250 -0.4042 0.01612 0.00929 -0.0950 0.7490 0.0304
-8.000 -0.3779 0.01572 0.00878 -0.0948 0.7396 0.0308
-7.750 -0.3512 0.01532 0.00830 -0.0946 0.7294 0.0312
-7.500 -0.3243 0.01497 0.00785 -0.0944 0.7210 0.0315
-7.250 -0.2970 0.01464 0.00744 -0.0943 0.7130 0.0318
-7.000 -0.2704 0.01422 0.00697 -0.0942 0.7053 0.0324
-6.750 -0.2431 0.01386 0.00658 -0.0941 0.6971 0.0330
-6.500 -0.2156 0.01358 0.00623 -0.0940 0.6894 0.0336
-6.250 -0.1878 0.01330 0.00591 -0.0940 0.6826 0.0342
-6.000 -0.1599 0.01305 0.00559 -0.0939 0.6752 0.0348
-5.750 -0.1320 0.01283 0.00530 -0.0939 0.6683 0.0355
-5.500 -0.1038 0.01262 0.00503 -0.0939 0.6607 0.0361
-5.250 -0.0760 0.01238 0.00475 -0.0938 0.6537 0.0371
-5.000 -0.0477 0.01217 0.00451 -0.0939 0.6475 0.0382
-4.750 -0.0194 0.01198 0.00429 -0.0939 0.6411 0.0396
-4.250 0.0373 0.01164 0.00389 -0.0940 0.6290 0.0442
-4.000 0.0656 0.01147 0.00372 -0.0940 0.6222 0.0491
-3.750 0.0939 0.01134 0.00359 -0.0940 0.6159 0.0572
-3.500 0.1226 0.01122 0.00349 -0.0942 0.6096 0.0659
-3.250 0.1512 0.01114 0.00340 -0.0942 0.6033 0.0729
-3.000 0.1797 0.01107 0.00332 -0.0943 0.5975 0.0785
-2.750 0.2087 0.01101 0.00324 -0.0945 0.5913 0.0831
-2.500 0.2372 0.01094 0.00316 -0.0945 0.5844 0.0881
-2.250 0.2658 0.01090 0.00309 -0.0946 0.5778 0.0929
-2.000 0.2944 0.01082 0.00302 -0.0947 0.5700 0.0983
-1.750 0.3227 0.01080 0.00296 -0.0948 0.5627 0.1044
-1.500 0.3514 0.01073 0.00291 -0.0949 0.5552 0.1110
-1.250 0.3796 0.01070 0.00286 -0.0950 0.5474 0.1188
-1.000 0.4081 0.01064 0.00283 -0.0951 0.5396 0.1307
-0.750 0.4360 0.01059 0.00280 -0.0951 0.5305 0.1491
-0.500 0.4640 0.01051 0.00280 -0.0952 0.5214 0.1787
-0.250 0.4916 0.01047 0.00280 -0.0952 0.5117 0.2101
0.000 0.5196 0.01041 0.00281 -0.0953 0.5027 0.2421
0.250 0.5470 0.01039 0.00284 -0.0953 0.4941 0.2786
0.500 0.5747 0.01032 0.00288 -0.0954 0.4855 0.3238
0.750 0.6016 0.01025 0.00294 -0.0953 0.4771 0.3841
1.000 0.6285 0.01009 0.00300 -0.0953 0.4693 0.4717
1.250 0.6544 0.00990 0.00310 -0.0951 0.4623 0.5812
1.500 0.6797 0.00968 0.00321 -0.0947 0.4566 0.6943
1.750 0.7030 0.00950 0.00334 -0.0936 0.4505 0.8057
2.000 0.7245 0.00943 0.00350 -0.0918 0.4447 0.9181
2.250 0.7647 0.00954 0.00361 -0.0943 0.4393 0.9774
2.500 0.8064 0.00968 0.00371 -0.0974 0.4338 1.0000
2.750 0.8312 0.00985 0.00381 -0.0969 0.4289 1.0000
3.000 0.8566 0.01000 0.00391 -0.0964 0.4247 1.0000
3.250 0.8827 0.01012 0.00401 -0.0962 0.4206 1.0000
3.500 0.9083 0.01028 0.00413 -0.0958 0.4155 1.0000
3.750 0.9332 0.01049 0.00427 -0.0954 0.4098 1.0000
4.000 0.9593 0.01063 0.00440 -0.0951 0.4045 1.0000
4.250 0.9850 0.01080 0.00453 -0.0948 0.3984 1.0000
4.500 1.0098 0.01102 0.00469 -0.0944 0.3930 1.0000
4.750 1.0358 0.01118 0.00484 -0.0942 0.3891 1.0000
5.000 1.0619 0.01133 0.00500 -0.0940 0.3852 1.0000
5.250 1.0874 0.01151 0.00516 -0.0937 0.3809 1.0000
5.500 1.1121 0.01173 0.00535 -0.0933 0.3766 1.0000
5.750 1.1372 0.01192 0.00554 -0.0929 0.3730 1.0000
6.000 1.1628 0.01209 0.00572 -0.0927 0.3694 1.0000
6.250 1.1878 0.01228 0.00592 -0.0923 0.3654 1.0000
6.500 1.2119 0.01251 0.00613 -0.0919 0.3613 1.0000
6.750 1.2352 0.01276 0.00637 -0.0913 0.3574 1.0000
7.000 1.2600 0.01294 0.00659 -0.0909 0.3536 1.0000
7.250 1.2838 0.01316 0.00682 -0.0904 0.3490 1.0000
7.500 1.3057 0.01344 0.00709 -0.0896 0.3436 1.0000
7.750 1.3276 0.01371 0.00736 -0.0888 0.3379 1.0000
8.000 1.3480 0.01398 0.00764 -0.0877 0.3310 1.0000
8.250 1.3655 0.01434 0.00798 -0.0861 0.3249 1.0000
8.500 1.3863 0.01463 0.00830 -0.0852 0.3190 1.0000
8.750 1.4038 0.01504 0.00870 -0.0838 0.3113 1.0000
9.000 1.4221 0.01546 0.00912 -0.0826 0.3031 1.0000
9.250 1.4379 0.01601 0.00964 -0.0811 0.2942 1.0000
9.500 1.4560 0.01648 0.01013 -0.0799 0.2866 1.0000
10.000 1.4877 0.01772 0.01136 -0.0773 0.2703 1.0000
10.250 1.5014 0.01849 0.01211 -0.0758 0.2621 1.0000
10.500 1.5156 0.01926 0.01288 -0.0745 0.2527 1.0000
10.750 1.5276 0.02020 0.01380 -0.0730 0.2433 1.0000
11.000 1.5374 0.02131 0.01488 -0.0714 0.2334 1.0000
11.250 1.5497 0.02230 0.01588 -0.0701 0.2246 1.0000
11.500 1.5597 0.02348 0.01705 -0.0687 0.2177 1.0000
11.750 1.5720 0.02453 0.01813 -0.0676 0.2115 1.0000
12.000 1.5805 0.02587 0.01948 -0.0663 0.2051 1.0000
12.250 1.5902 0.02716 0.02079 -0.0651 0.1984 1.0000
12.500 1.5965 0.02874 0.02237 -0.0637 0.1914 1.0000
12.750 1.6043 0.03025 0.02390 -0.0626 0.1852 1.0000
13.000 1.6102 0.03194 0.02561 -0.0614 0.1788 1.0000
13.250 1.6152 0.03373 0.02743 -0.0603 0.1738 1.0000
13.500 1.6227 0.03536 0.02910 -0.0593 0.1691 1.0000
13.750 1.6270 0.03732 0.03110 -0.0584 0.1643 1.0000
14.000 1.6304 0.03943 0.03324 -0.0575 0.1599 1.0000
14.250 1.6366 0.04131 0.03518 -0.0567 0.1550 1.0000
14.500 1.6380 0.04370 0.03760 -0.0559 0.1498 1.0000
14.750 1.6403 0.04607 0.04002 -0.0553 0.1452 1.0000
15.000 1.6432 0.04843 0.04242 -0.0547 0.1397 1.0000
15.250 1.6405 0.05139 0.04541 -0.0541 0.1342 1.0000
15.500 1.6416 0.05400 0.04808 -0.0537 0.1278 1.0000
15.750 1.6371 0.05729 0.05138 -0.0532 0.1208 1.0000
16.000 1.6312 0.06079 0.05490 -0.0529 0.1125 1.0000
16.250 1.6229 0.06461 0.05873 -0.0526 0.1044 1.0000
16.500 1.6129 0.06873 0.06287 -0.0525 0.0977 1.0000
16.750 1.6042 0.07278 0.06695 -0.0525 0.0925 1.0000
17.000 1.5952 0.07691 0.07112 -0.0526 0.0883 1.0000
17.250 1.5877 0.08091 0.07518 -0.0528 0.0846 1.0000
17.500 1.5774 0.08534 0.07966 -0.0531 0.0812 1.0000
17.750 1.5721 0.08917 0.08356 -0.0535 0.0784 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 624 AIRFOIL (goe624-il)