GOE 522 AIRFOIL (goe522-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 522 AIRFOIL (goe522-il) Reynolds number: 100,000 Max Cl/Cd: 37.59 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe522-il-100000-n5.txt Download as CSV file: xf-goe522-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 522 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-14.000 -0.2315 0.12613 0.12028 -0.0776 0.9745 0.0543
-13.750 -0.4084 0.08833 0.08184 -0.0944 0.9715 0.0546
-13.500 -0.4171 0.08236 0.07572 -0.0988 0.9675 0.0549
-13.250 -0.4185 0.07740 0.07063 -0.1031 0.9644 0.0553
-13.000 -0.4205 0.07331 0.06643 -0.1060 0.9600 0.0557
-12.750 -0.4234 0.06954 0.06255 -0.1084 0.9539 0.0562
-12.500 -0.4188 0.06580 0.05874 -0.1119 0.9496 0.0566
-12.250 -0.4102 0.06207 0.05494 -0.1162 0.9463 0.0572
-12.000 -0.4180 0.05888 0.05171 -0.1171 0.9372 0.0577
-11.750 -0.4134 0.05528 0.04803 -0.1206 0.9319 0.0584
-11.250 -0.4134 0.04870 0.04126 -0.1252 0.9166 0.0599
-11.000 -0.4039 0.04530 0.03774 -0.1291 0.9114 0.0612
-10.750 -0.4113 0.04259 0.03495 -0.1295 0.9009 0.0619
-10.500 -0.4074 0.03965 0.03191 -0.1320 0.8932 0.0634
-10.250 -0.4117 0.03743 0.02958 -0.1320 0.8839 0.0649
-10.000 -0.4191 0.03580 0.02784 -0.1302 0.8740 0.0663
-9.750 -0.4060 0.03416 0.02612 -0.1307 0.8696 0.0698
-9.500 -0.4267 0.03358 0.02546 -0.1242 0.8560 0.0711
-9.250 -0.4074 0.03230 0.02411 -0.1246 0.8514 0.0769
-9.000 -0.4085 0.03162 0.02337 -0.1209 0.8420 0.0810
-8.750 -0.3971 0.03079 0.02248 -0.1191 0.8347 0.0871
-8.500 -0.3700 0.02978 0.02142 -0.1200 0.8307 0.0952
-8.250 -0.3654 0.02931 0.02090 -0.1166 0.8215 0.1011
-8.000 -0.3478 0.02868 0.02022 -0.1154 0.8144 0.1081
-7.750 -0.3157 0.02794 0.01940 -0.1166 0.8105 0.1171
-7.500 -0.3026 0.02753 0.01895 -0.1144 0.8025 0.1233
-7.250 -0.2850 0.02712 0.01842 -0.1127 0.7945 0.1293
-7.000 -0.2515 0.02644 0.01767 -0.1139 0.7901 0.1359
-6.750 -0.2325 0.02605 0.01719 -0.1125 0.7824 0.1414
-6.500 -0.2156 0.02563 0.01674 -0.1107 0.7737 0.1459
-6.250 -0.1787 0.02503 0.01599 -0.1123 0.7691 0.1529
-6.000 -0.1657 0.02469 0.01566 -0.1098 0.7598 0.1578
-5.750 -0.1409 0.02425 0.01514 -0.1092 0.7520 0.1647
-5.500 -0.1030 0.02357 0.01444 -0.1110 0.7470 0.1723
-5.250 -0.0942 0.02339 0.01422 -0.1077 0.7355 0.1779
-5.000 -0.0619 0.02284 0.01367 -0.1084 0.7284 0.1865
-4.750 -0.0375 0.02249 0.01327 -0.1078 0.7198 0.1952
-4.500 -0.0167 0.02218 0.01301 -0.1065 0.7101 0.2044
-4.250 0.0233 0.02169 0.01250 -0.1086 0.7039 0.2199
-4.000 0.0329 0.02163 0.01248 -0.1053 0.6914 0.2320
-3.750 0.0697 0.02129 0.01210 -0.1067 0.6839 0.2521
-3.500 0.0844 0.02126 0.01206 -0.1043 0.6723 0.2674
-3.250 0.1170 0.02106 0.01179 -0.1050 0.6640 0.2868
-3.000 0.1370 0.02104 0.01170 -0.1035 0.6539 0.3027
-2.750 0.1643 0.02094 0.01154 -0.1032 0.6451 0.3185
-2.500 0.1895 0.02087 0.01144 -0.1026 0.6362 0.3318
-2.250 0.2121 0.02086 0.01130 -0.1015 0.6268 0.3455
-2.000 0.2441 0.02074 0.01117 -0.1021 0.6196 0.3582
-1.750 0.2613 0.02082 0.01116 -0.1001 0.6105 0.3694
-1.500 0.2900 0.02076 0.01110 -0.1001 0.6037 0.3801
-1.250 0.3165 0.02078 0.01099 -0.0998 0.5968 0.3910
-1.000 0.3368 0.02082 0.01105 -0.0984 0.5892 0.3993
-0.750 0.3664 0.02081 0.01094 -0.0985 0.5830 0.4098
-0.500 0.3926 0.02085 0.01093 -0.0982 0.5769 0.4196
-0.250 0.4124 0.02096 0.01106 -0.0967 0.5704 0.4301
0.000 0.4391 0.02103 0.01107 -0.0964 0.5650 0.4421
0.250 0.4731 0.02106 0.01103 -0.0974 0.5603 0.4546
0.500 0.4909 0.02122 0.01120 -0.0957 0.5545 0.4659
0.750 0.5114 0.02135 0.01137 -0.0943 0.5488 0.4779
1.000 0.5381 0.02145 0.01142 -0.0941 0.5439 0.4908
1.250 0.5726 0.02153 0.01144 -0.0952 0.5399 0.5039
1.500 0.5924 0.02172 0.01165 -0.0939 0.5351 0.5155
1.750 0.6102 0.02191 0.01190 -0.0921 0.5300 0.5277
2.000 0.6336 0.02207 0.01206 -0.0914 0.5251 0.5422
2.250 0.6631 0.02218 0.01212 -0.0916 0.5205 0.5597
2.500 0.6873 0.02233 0.01229 -0.0909 0.5154 0.5762
2.750 0.6975 0.02259 0.01263 -0.0879 0.5093 0.5921
3.000 0.7177 0.02276 0.01279 -0.0865 0.5028 0.6112
3.250 0.7511 0.02277 0.01272 -0.0874 0.4968 0.6337
3.500 0.7555 0.02312 0.01321 -0.0834 0.4901 0.6542
3.750 0.7698 0.02336 0.01354 -0.0811 0.4839 0.6797
4.000 0.7929 0.02346 0.01371 -0.0802 0.4788 0.7115
4.500 0.8328 0.02387 0.01442 -0.0774 0.4699 0.8075
4.750 0.8700 0.02418 0.01495 -0.0798 0.4641 0.9068
5.000 0.9004 0.02446 0.01516 -0.0808 0.4591 1.0000
5.250 0.9278 0.02468 0.01522 -0.0809 0.4549 1.0000
5.500 0.9362 0.02524 0.01579 -0.0782 0.4499 1.0000
5.750 0.9440 0.02584 0.01641 -0.0755 0.4444 1.0000
6.000 0.9599 0.02627 0.01679 -0.0739 0.4394 1.0000
6.250 0.9833 0.02653 0.01694 -0.0734 0.4351 1.0000
6.500 0.9942 0.02714 0.01755 -0.0713 0.4301 1.0000
6.750 0.9981 0.02796 0.01842 -0.0683 0.4244 1.0000
7.000 1.0117 0.02852 0.01895 -0.0666 0.4193 1.0000
7.250 1.0341 0.02879 0.01912 -0.0660 0.4149 1.0000
7.500 1.0389 0.02971 0.02009 -0.0634 0.4095 1.0000
7.750 1.0419 0.03072 0.02116 -0.0607 0.4033 1.0000
8.000 1.0564 0.03128 0.02168 -0.0593 0.3981 1.0000
8.250 1.0738 0.03176 0.02209 -0.0582 0.3932 1.0000
8.500 1.0672 0.03336 0.02382 -0.0549 0.3860 1.0000
8.750 1.0777 0.03418 0.02461 -0.0533 0.3802 1.0000
9.000 1.0957 0.03465 0.02500 -0.0524 0.3752 1.0000
9.250 1.0861 0.03670 0.02719 -0.0494 0.3674 1.0000
9.500 1.0970 0.03759 0.02805 -0.0480 0.3615 1.0000
9.750 1.1063 0.03866 0.02910 -0.0467 0.3557 1.0000
10.000 1.1012 0.04079 0.03131 -0.0445 0.3484 1.0000
10.250 1.1163 0.04153 0.03200 -0.0437 0.3434 1.0000
10.500 1.1182 0.04332 0.03383 -0.0422 0.3373 1.0000
10.750 1.1164 0.04549 0.03605 -0.0407 0.3306 1.0000
11.000 1.1340 0.04611 0.03661 -0.0401 0.3262 1.0000
11.250 1.1338 0.04831 0.03885 -0.0389 0.3205 1.0000
11.500 1.1314 0.05079 0.04140 -0.0377 0.3147 1.0000
11.750 1.1468 0.05169 0.04225 -0.0372 0.3107 1.0000
12.000 1.1668 0.05220 0.04269 -0.0368 0.3074 1.0000
12.250 1.1481 0.05645 0.04711 -0.0356 0.3012 1.0000
12.500 1.1534 0.05843 0.04910 -0.0350 0.2969 1.0000
12.750 1.1727 0.05900 0.04962 -0.0347 0.2940 1.0000
13.000 1.1945 0.05936 0.04992 -0.0344 0.2915 1.0000
13.250 1.1575 0.06602 0.05683 -0.0336 0.2854 1.0000
13.500 1.1560 0.06901 0.05988 -0.0334 0.2818 1.0000
13.750 1.1702 0.07020 0.06107 -0.0331 0.2794 1.0000
14.000 1.1920 0.07054 0.06138 -0.0329 0.2777 1.0000
14.250 1.2181 0.07041 0.06121 -0.0326 0.2763 1.0000
14.500 1.0587 0.09415 0.08550 -0.0352 0.2632 1.0000
14.750 1.0782 0.09465 0.08600 -0.0351 0.2620 1.0000
15.000 1.1026 0.09447 0.08582 -0.0347 0.2612 1.0000
15.250 1.1289 0.09404 0.08537 -0.0344 0.2604 1.0000
16.750 0.8743 0.15843 0.15053 -0.0544 0.2205 1.0000
17.000 0.8847 0.16028 0.15240 -0.0551 0.2177 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 522 AIRFOIL (goe522-il)