GOE 432 AIRFOIL (goe432-il) Xfoil prediction polar at RE=200,000 Ncrit=5
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Airfoil: GOE 432 AIRFOIL (goe432-il) Reynolds number: 200,000 Max Cl/Cd: 72 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe432-il-200000-n5.txt Download as CSV file: xf-goe432-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 432 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3480 0.10027 0.09683 -0.0273 1.0000 0.0221
-8.250 -0.3496 0.09761 0.09422 -0.0268 1.0000 0.0221
-8.000 -0.3532 0.09514 0.09181 -0.0259 1.0000 0.0221
-7.750 -0.2703 0.07827 0.07521 -0.0241 1.0000 0.0210
-7.500 -0.2735 0.07555 0.07253 -0.0226 1.0000 0.0197
-7.250 -0.2811 0.07289 0.06993 -0.0213 1.0000 0.0188
-6.750 -0.3701 0.07886 0.07575 -0.0228 0.9996 0.0157
-6.500 -0.3460 0.07398 0.07083 -0.0293 0.9953 0.0154
-6.250 -0.3197 0.06847 0.06529 -0.0371 0.9902 0.0152
-6.000 -0.2890 0.06257 0.05932 -0.0457 0.9855 0.0151
-5.750 -0.2552 0.05506 0.05168 -0.0558 0.9792 0.0156
-5.500 -0.2231 0.05031 0.04683 -0.0625 0.9745 0.0167
-5.250 -0.1876 0.04251 0.03876 -0.0706 0.9681 0.0166
-5.000 -0.1504 0.03493 0.03077 -0.0773 0.9635 0.0169
-4.750 -0.1206 0.02878 0.02411 -0.0804 0.9564 0.0174
-4.500 -0.0855 0.02407 0.01876 -0.0832 0.9525 0.0192
-4.250 -0.0585 0.01962 0.01339 -0.0833 0.9459 0.0218
-4.000 -0.0261 0.01807 0.01154 -0.0844 0.9406 0.0233
-3.750 0.0116 0.01735 0.01063 -0.0863 0.9368 0.0265
-3.500 0.0382 0.01617 0.00910 -0.0856 0.9271 0.0300
-3.250 0.0725 0.01557 0.00846 -0.0868 0.9208 0.0333
-3.000 0.1008 0.01511 0.00780 -0.0865 0.9101 0.0381
-2.750 0.1294 0.01432 0.00695 -0.0865 0.8996 0.0418
-2.500 0.1613 0.01392 0.00647 -0.0870 0.8899 0.0471
-2.250 0.1941 0.01342 0.00582 -0.0877 0.8789 0.0508
-2.000 0.2263 0.01267 0.00506 -0.0884 0.8648 0.0551
-1.750 0.2602 0.01216 0.00447 -0.0892 0.8449 0.0583
-1.500 0.2962 0.01176 0.00392 -0.0904 0.8202 0.0624
-1.250 0.3284 0.01142 0.00345 -0.0909 0.7909 0.0667
-1.000 0.3573 0.01123 0.00311 -0.0908 0.7603 0.0712
-0.750 0.3844 0.01115 0.00286 -0.0903 0.7292 0.0766
-0.500 0.4103 0.01109 0.00268 -0.0896 0.6963 0.0858
-0.250 0.4351 0.01102 0.00254 -0.0887 0.6619 0.1142
0.000 0.4579 0.01074 0.00254 -0.0877 0.6266 0.2527
0.250 0.4806 0.01066 0.00254 -0.0866 0.5928 0.3507
0.750 0.5509 0.00971 0.00259 -0.0895 0.5355 1.0000
1.000 0.5754 0.00992 0.00263 -0.0887 0.5190 1.0000
1.250 0.6002 0.01013 0.00268 -0.0879 0.5052 1.0000
1.500 0.6251 0.01033 0.00275 -0.0872 0.4925 1.0000
1.750 0.6503 0.01051 0.00284 -0.0866 0.4799 1.0000
2.000 0.6753 0.01070 0.00293 -0.0859 0.4670 1.0000
2.250 0.7003 0.01089 0.00304 -0.0852 0.4538 1.0000
2.500 0.7253 0.01109 0.00315 -0.0846 0.4423 1.0000
2.750 0.7503 0.01129 0.00329 -0.0839 0.4310 1.0000
3.000 0.7753 0.01148 0.00342 -0.0833 0.4185 1.0000
3.250 0.8001 0.01168 0.00357 -0.0826 0.4052 1.0000
3.500 0.8247 0.01189 0.00374 -0.0819 0.3917 1.0000
3.750 0.8494 0.01211 0.00391 -0.0812 0.3788 1.0000
4.000 0.8738 0.01234 0.00410 -0.0805 0.3654 1.0000
4.250 0.8980 0.01259 0.00432 -0.0798 0.3507 1.0000
4.500 0.9219 0.01286 0.00454 -0.0790 0.3344 1.0000
4.750 0.9456 0.01315 0.00478 -0.0782 0.3175 1.0000
5.000 0.9691 0.01346 0.00505 -0.0774 0.2996 1.0000
5.250 0.9925 0.01379 0.00534 -0.0766 0.2808 1.0000
5.500 1.0155 0.01414 0.00564 -0.0758 0.2580 1.0000
5.750 1.0376 0.01458 0.00598 -0.0748 0.2335 1.0000
6.000 1.0591 0.01508 0.00638 -0.0738 0.2124 1.0000
6.250 1.0809 0.01556 0.00681 -0.0728 0.1996 1.0000
6.500 1.1033 0.01599 0.00724 -0.0719 0.1924 1.0000
6.750 1.1251 0.01645 0.00772 -0.0709 0.1867 1.0000
7.000 1.1469 0.01691 0.00824 -0.0699 0.1820 1.0000
7.250 1.1690 0.01735 0.00875 -0.0690 0.1776 1.0000
7.500 1.1901 0.01786 0.00931 -0.0679 0.1733 1.0000
7.750 1.2102 0.01846 0.00993 -0.0667 0.1694 1.0000
8.000 1.2323 0.01889 0.01052 -0.0658 0.1658 1.0000
8.250 1.2532 0.01940 0.01114 -0.0647 0.1607 1.0000
8.500 1.2720 0.02005 0.01180 -0.0634 0.1532 1.0000
8.750 1.2929 0.02043 0.01230 -0.0624 0.1429 1.0000
9.000 1.3138 0.02082 0.01280 -0.0614 0.1325 1.0000
9.250 1.3344 0.02120 0.01325 -0.0603 0.1186 1.0000
9.500 1.3529 0.02174 0.01376 -0.0591 0.0976 1.0000
9.750 1.3665 0.02268 0.01457 -0.0572 0.0795 1.0000
10.000 1.3791 0.02371 0.01558 -0.0552 0.0578 1.0000
10.250 1.3807 0.02544 0.01707 -0.0517 0.0304 1.0000
10.500 1.3846 0.02693 0.01856 -0.0485 0.0215 1.0000
10.750 1.3896 0.02830 0.02003 -0.0455 0.0179 1.0000
11.000 1.3952 0.02963 0.02151 -0.0428 0.0160 1.0000
11.250 1.3989 0.03110 0.02312 -0.0401 0.0148 1.0000
11.500 1.3988 0.03290 0.02508 -0.0373 0.0137 1.0000
11.750 1.3954 0.03501 0.02739 -0.0346 0.0129 1.0000
12.000 1.3960 0.03692 0.02948 -0.0326 0.0123 1.0000
12.250 1.3948 0.03912 0.03186 -0.0309 0.0118 1.0000
12.500 1.3922 0.04161 0.03451 -0.0296 0.0111 1.0000
12.750 1.3869 0.04461 0.03769 -0.0288 0.0108 1.0000
13.000 1.3805 0.04799 0.04123 -0.0285 0.0104 1.0000
13.250 1.3725 0.05185 0.04526 -0.0288 0.0102 1.0000
13.500 1.3637 0.05615 0.04971 -0.0297 0.0100 1.0000
13.750 1.3536 0.06094 0.05467 -0.0312 0.0098 1.0000
14.000 1.3425 0.06610 0.05999 -0.0330 0.0097 1.0000
14.250 1.3312 0.07149 0.06552 -0.0350 0.0097 1.0000
14.500 1.3190 0.07713 0.07131 -0.0373 0.0096 1.0000
14.750 1.3056 0.08310 0.07741 -0.0397 0.0094 1.0000
15.000 1.2939 0.08893 0.08337 -0.0421 0.0093 1.0000
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Polar data table (+)
Polar graphs
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