GOE 625 AIRFOIL (goe625-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 625 AIRFOIL (goe625-il) Reynolds number: 200,000 Max Cl/Cd: 58.08 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe625-il-200000.txt Download as CSV file: xf-goe625-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 625 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.0201 0.09365 0.08974 -0.1119 0.9203 0.1145
-10.000 -0.1787 0.05859 0.05443 -0.1375 0.9058 0.0947
-9.750 -0.1943 0.05442 0.05015 -0.1378 0.8981 0.0951
-9.500 -0.0980 0.06631 0.06226 -0.1297 0.8952 0.1040
-9.250 -0.0917 0.06376 0.05970 -0.1295 0.8865 0.1031
-9.000 -0.1047 0.05859 0.05443 -0.1317 0.8791 0.1021
-8.750 -0.2133 0.04207 0.03710 -0.1349 0.8651 0.0996
-8.500 -0.2276 0.03616 0.03046 -0.1335 0.8594 0.1019
-8.250 -0.2327 0.03284 0.02648 -0.1306 0.8497 0.1038
-8.000 -0.2069 0.03187 0.02558 -0.1305 0.8442 0.1055
-7.750 -0.1762 0.03167 0.02544 -0.1307 0.8400 0.1072
-7.500 -0.1587 0.03088 0.02457 -0.1294 0.8313 0.1091
-7.250 -0.1389 0.02930 0.02267 -0.1285 0.8252 0.1114
-7.000 -0.1167 0.02763 0.02052 -0.1280 0.8205 0.1135
-6.750 -0.0979 0.02653 0.01936 -0.1268 0.8114 0.1152
-6.500 -0.0704 0.02587 0.01869 -0.1266 0.8050 0.1169
-6.250 -0.0438 0.02523 0.01797 -0.1263 0.7984 0.1187
-6.000 -0.0203 0.02448 0.01710 -0.1255 0.7894 0.1206
-5.750 0.0076 0.02355 0.01592 -0.1254 0.7830 0.1229
-5.500 0.0307 0.02284 0.01499 -0.1244 0.7734 0.1245
-5.250 0.0580 0.02184 0.01386 -0.1242 0.7656 0.1262
-5.000 0.0840 0.02130 0.01333 -0.1238 0.7562 0.1282
-4.750 0.1114 0.02081 0.01277 -0.1235 0.7468 0.1304
-4.500 0.1377 0.02030 0.01217 -0.1230 0.7362 0.1326
-4.250 0.1657 0.01974 0.01143 -0.1228 0.7259 0.1349
-4.000 0.1911 0.01935 0.01089 -0.1220 0.7127 0.1372
-3.750 0.2178 0.01872 0.01024 -0.1217 0.7001 0.1400
-3.250 0.2703 0.01807 0.00943 -0.1207 0.6698 0.1464
-3.000 0.2958 0.01783 0.00902 -0.1200 0.6529 0.1502
-2.750 0.3209 0.01745 0.00859 -0.1194 0.6363 0.1543
-2.500 0.3462 0.01732 0.00840 -0.1188 0.6206 0.1597
-2.250 0.3721 0.01730 0.00817 -0.1182 0.6066 0.1665
-2.000 0.3971 0.01703 0.00791 -0.1177 0.5939 0.1745
-1.750 0.4224 0.01693 0.00774 -0.1172 0.5820 0.1850
-1.500 0.4487 0.01687 0.00760 -0.1169 0.5728 0.1969
-1.250 0.4736 0.01675 0.00753 -0.1165 0.5634 0.2089
-1.000 0.5007 0.01681 0.00745 -0.1163 0.5558 0.2211
-0.750 0.5260 0.01673 0.00744 -0.1160 0.5483 0.2330
-0.500 0.5521 0.01671 0.00740 -0.1157 0.5416 0.2446
-0.250 0.5803 0.01678 0.00739 -0.1159 0.5362 0.2575
0.000 0.6059 0.01674 0.00746 -0.1156 0.5309 0.2720
0.250 0.6319 0.01673 0.00753 -0.1154 0.5258 0.2918
0.500 0.6584 0.01666 0.00761 -0.1153 0.5214 0.3300
1.000 0.6972 0.01559 0.00816 -0.1117 0.5138 0.7866
1.250 0.7480 0.01576 0.00855 -0.1152 0.5087 0.9644
1.500 0.8212 0.01613 0.00876 -0.1246 0.5037 1.0000
1.750 0.8442 0.01637 0.00886 -0.1239 0.5005 1.0000
2.000 0.8697 0.01671 0.00905 -0.1236 0.4974 1.0000
2.250 0.8885 0.01693 0.00927 -0.1221 0.4942 1.0000
2.500 0.9090 0.01717 0.00948 -0.1209 0.4907 1.0000
2.750 0.9312 0.01742 0.00967 -0.1200 0.4872 1.0000
3.000 0.9551 0.01767 0.00981 -0.1194 0.4834 1.0000
3.250 0.9833 0.01805 0.01001 -0.1197 0.4791 1.0000
3.500 1.0013 0.01828 0.01028 -0.1180 0.4750 1.0000
3.750 1.0221 0.01852 0.01052 -0.1168 0.4707 1.0000
4.000 1.0461 0.01879 0.01074 -0.1163 0.4671 1.0000
4.250 1.0723 0.01908 0.01095 -0.1162 0.4639 1.0000
4.500 1.1015 0.01949 0.01123 -0.1167 0.4607 1.0000
4.750 1.1239 0.01987 0.01163 -0.1160 0.4577 1.0000
5.000 1.1443 0.02020 0.01201 -0.1149 0.4546 1.0000
5.250 1.1667 0.02053 0.01236 -0.1142 0.4513 1.0000
5.500 1.1910 0.02085 0.01265 -0.1138 0.4481 1.0000
5.750 1.2179 0.02116 0.01289 -0.1139 0.4448 1.0000
6.000 1.2489 0.02162 0.01323 -0.1148 0.4415 1.0000
6.250 1.2672 0.02202 0.01372 -0.1134 0.4383 1.0000
6.500 1.2849 0.02239 0.01416 -0.1119 0.4348 1.0000
6.750 1.3059 0.02274 0.01454 -0.1110 0.4312 1.0000
7.000 1.3302 0.02306 0.01484 -0.1107 0.4279 1.0000
7.250 1.3590 0.02340 0.01509 -0.1112 0.4247 1.0000
7.500 1.3869 0.02394 0.01559 -0.1117 0.4213 1.0000
7.750 1.3971 0.02436 0.01616 -0.1089 0.4177 1.0000
8.000 1.4123 0.02476 0.01662 -0.1071 0.4136 1.0000
8.250 1.4331 0.02505 0.01691 -0.1062 0.4097 1.0000
8.500 1.4614 0.02531 0.01709 -0.1066 0.4061 1.0000
8.750 1.4859 0.02582 0.01758 -0.1065 0.4023 1.0000
9.000 1.4883 0.02634 0.01826 -0.1026 0.3983 1.0000
9.250 1.4972 0.02676 0.01875 -0.0997 0.3942 1.0000
9.500 1.5161 0.02706 0.01903 -0.0986 0.3903 1.0000
9.750 1.5478 0.02729 0.01916 -0.0997 0.3865 1.0000
10.000 1.5500 0.02798 0.01997 -0.0960 0.3824 1.0000
10.250 1.5487 0.02870 0.02083 -0.0919 0.3777 1.0000
10.500 1.5613 0.02912 0.02126 -0.0900 0.3732 1.0000
10.750 1.5868 0.02925 0.02130 -0.0900 0.3691 1.0000
11.000 1.5876 0.03016 0.02232 -0.0867 0.3644 1.0000
11.250 1.5846 0.03117 0.02345 -0.0831 0.3591 1.0000
11.500 1.5972 0.03166 0.02393 -0.0815 0.3544 1.0000
11.750 1.6258 0.03174 0.02388 -0.0820 0.3503 1.0000
12.000 1.6094 0.03357 0.02596 -0.0773 0.3451 1.0000
12.250 1.6131 0.03469 0.02715 -0.0752 0.3401 1.0000
12.500 1.6284 0.03527 0.02770 -0.0742 0.3361 1.0000
12.750 1.6406 0.03615 0.02859 -0.0731 0.3319 1.0000
13.000 1.6315 0.03822 0.03083 -0.0701 0.3267 1.0000
13.250 1.6383 0.03941 0.03206 -0.0687 0.3223 1.0000
13.500 1.6577 0.03986 0.03245 -0.0683 0.3186 1.0000
13.750 1.6566 0.04175 0.03445 -0.0664 0.3143 1.0000
14.000 1.6508 0.04404 0.03687 -0.0644 0.3097 1.0000
14.250 1.6584 0.04537 0.03823 -0.0634 0.3057 1.0000
14.500 1.6806 0.04564 0.03842 -0.0632 0.3023 1.0000
14.750 1.6726 0.04836 0.04128 -0.0615 0.2982 1.0000
15.000 1.6651 0.05118 0.04424 -0.0600 0.2940 1.0000
15.250 1.6711 0.05287 0.04596 -0.0592 0.2903 1.0000
15.500 1.6925 0.05315 0.04617 -0.0590 0.2870 1.0000
15.750 1.6901 0.05570 0.04883 -0.0580 0.2835 1.0000
16.000 1.6730 0.05979 0.05310 -0.0568 0.2796 1.0000
16.250 1.6733 0.06221 0.05559 -0.0562 0.2759 1.0000
16.500 1.6911 0.06279 0.05611 -0.0560 0.2724 1.0000
16.750 1.7025 0.06409 0.05742 -0.0556 0.2690 1.0000
17.000 1.6705 0.07019 0.06377 -0.0547 0.2654 1.0000
17.250 1.6590 0.07418 0.06788 -0.0544 0.2616 1.0000
17.500 1.6717 0.07534 0.06903 -0.0543 0.2584 1.0000
17.750 1.7135 0.07306 0.06658 -0.0542 0.2548 1.0000
18.000 1.6578 0.08245 0.07632 -0.0541 0.2515 1.0000
18.250 1.6106 0.09134 0.08546 -0.0546 0.2476 1.0000
18.500 1.6248 0.09232 0.08644 -0.0547 0.2442 1.0000
18.750 1.6768 0.08829 0.08224 -0.0542 0.2409 1.0000
19.000 1.6188 0.09878 0.09303 -0.0554 0.2371 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 625 AIRFOIL (goe625-il)