GOE 255 (MVA CA.6) AIRFOIL (goe255-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 255 (MVA CA.6) AIRFOIL (goe255-il) Reynolds number: 500,000 Max Cl/Cd: 90.51 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe255-il-500000-n5.txt Download as CSV file: xf-goe255-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 255 (MVA CA.6) 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
------ -------- --------- --------- -------- -------- --------
-16.500 -0.6573 0.08941 0.08572 -0.0753 1.0000 0.0087
-16.250 -0.6657 0.08453 0.08077 -0.0777 1.0000 0.0088
-16.000 -0.6740 0.07979 0.07597 -0.0800 1.0000 0.0089
-15.750 -0.6816 0.07509 0.07119 -0.0826 0.9984 0.0089
-15.500 -0.6849 0.07012 0.06613 -0.0865 0.9937 0.0090
-15.250 -0.6885 0.06533 0.06125 -0.0901 0.9878 0.0091
-15.000 -0.6928 0.06071 0.05653 -0.0937 0.9823 0.0092
-14.750 -0.6978 0.05636 0.05208 -0.0968 0.9755 0.0092
-14.500 -0.7030 0.05225 0.04787 -0.0995 0.9672 0.0093
-14.250 -0.7080 0.04843 0.04395 -0.1016 0.9588 0.0094
-14.000 -0.7138 0.04477 0.04018 -0.1034 0.9484 0.0095
-13.750 -0.7184 0.04120 0.03651 -0.1051 0.9363 0.0096
-13.500 -0.7208 0.03779 0.03298 -0.1069 0.9203 0.0098
-13.250 -0.7220 0.03452 0.02954 -0.1085 0.8881 0.0099
-13.000 -0.7221 0.03151 0.02623 -0.1099 0.8431 0.0100
-12.750 -0.7170 0.02874 0.02319 -0.1124 0.8178 0.0102
-12.500 -0.7101 0.02660 0.02084 -0.1140 0.8027 0.0103
-12.250 -0.7052 0.02519 0.01924 -0.1133 0.7917 0.0105
-12.000 -0.6972 0.02377 0.01770 -0.1127 0.7822 0.0108
-11.750 -0.6832 0.02271 0.01652 -0.1121 0.7741 0.0110
-11.500 -0.6660 0.02182 0.01553 -0.1117 0.7671 0.0113
-11.250 -0.6470 0.02104 0.01464 -0.1113 0.7613 0.0117
-11.000 -0.6268 0.02031 0.01382 -0.1110 0.7564 0.0120
-10.750 -0.6054 0.01964 0.01306 -0.1107 0.7512 0.0124
-10.500 -0.5833 0.01902 0.01234 -0.1104 0.7461 0.0128
-10.250 -0.5603 0.01847 0.01168 -0.1101 0.7415 0.0132
-10.000 -0.5373 0.01783 0.01097 -0.1099 0.7367 0.0138
-9.750 -0.5140 0.01723 0.01030 -0.1097 0.7317 0.0145
-9.500 -0.4897 0.01675 0.00974 -0.1096 0.7271 0.0153
-9.250 -0.4644 0.01630 0.00921 -0.1095 0.7228 0.0162
-9.000 -0.4388 0.01587 0.00871 -0.1094 0.7184 0.0172
-8.750 -0.4136 0.01539 0.00818 -0.1093 0.7139 0.0185
-8.500 -0.3878 0.01501 0.00772 -0.1093 0.7097 0.0202
-8.250 -0.3616 0.01459 0.00727 -0.1093 0.7055 0.0228
-8.000 -0.3354 0.01415 0.00683 -0.1093 0.7008 0.0280
-7.750 -0.3097 0.01368 0.00638 -0.1093 0.6962 0.0389
-7.500 -0.2835 0.01327 0.00600 -0.1093 0.6919 0.0537
-7.250 -0.2564 0.01291 0.00568 -0.1095 0.6872 0.0658
-7.000 -0.2289 0.01265 0.00540 -0.1096 0.6821 0.0754
-6.750 -0.2016 0.01237 0.00511 -0.1097 0.6772 0.0857
-6.500 -0.1740 0.01208 0.00484 -0.1099 0.6722 0.0978
-6.250 -0.1466 0.01175 0.00459 -0.1100 0.6665 0.1179
-6.000 -0.1192 0.01149 0.00437 -0.1102 0.6609 0.1423
-5.750 -0.0911 0.01126 0.00419 -0.1104 0.6550 0.1605
-5.500 -0.0630 0.01109 0.00402 -0.1106 0.6481 0.1759
-5.250 -0.0348 0.01094 0.00386 -0.1108 0.6416 0.1900
-5.000 -0.0066 0.01078 0.00369 -0.1110 0.6336 0.2026
-4.750 0.0214 0.01066 0.00354 -0.1111 0.6248 0.2128
-4.500 0.0494 0.01055 0.00339 -0.1112 0.6146 0.2247
-4.250 0.0774 0.01044 0.00326 -0.1114 0.6047 0.2383
-4.000 0.1051 0.01032 0.00314 -0.1115 0.5961 0.2562
-3.750 0.1331 0.01019 0.00307 -0.1117 0.5880 0.2842
-3.500 0.1609 0.01014 0.00305 -0.1118 0.5801 0.3143
-3.250 0.1893 0.01016 0.00305 -0.1120 0.5718 0.3326
-3.000 0.2174 0.01021 0.00304 -0.1121 0.5645 0.3449
-2.750 0.2461 0.01023 0.00304 -0.1123 0.5587 0.3540
-2.500 0.2747 0.01029 0.00303 -0.1125 0.5538 0.3628
-2.000 0.3318 0.01036 0.00304 -0.1129 0.5458 0.3750
-1.750 0.3606 0.01041 0.00303 -0.1131 0.5418 0.3792
-1.500 0.3889 0.01044 0.00304 -0.1133 0.5377 0.3829
-1.000 0.4456 0.01054 0.00308 -0.1137 0.5309 0.3912
-0.750 0.4744 0.01059 0.00309 -0.1139 0.5279 0.3950
-0.500 0.5028 0.01062 0.00311 -0.1141 0.5244 0.3982
-0.250 0.5308 0.01067 0.00315 -0.1143 0.5210 0.4021
0.000 0.5589 0.01074 0.00320 -0.1144 0.5181 0.4072
0.250 0.5871 0.01081 0.00324 -0.1146 0.5154 0.4113
0.500 0.6156 0.01083 0.00328 -0.1148 0.5130 0.4153
0.750 0.6439 0.01085 0.00334 -0.1151 0.5105 0.4200
1.000 0.6721 0.01091 0.00340 -0.1152 0.5078 0.4249
1.250 0.7000 0.01098 0.00346 -0.1154 0.5049 0.4292
1.500 0.7274 0.01105 0.00353 -0.1155 0.5018 0.4329
1.750 0.7550 0.01111 0.00361 -0.1156 0.4985 0.4362
2.000 0.7831 0.01115 0.00368 -0.1158 0.4955 0.4400
2.250 0.8109 0.01121 0.00376 -0.1159 0.4926 0.4441
2.500 0.8384 0.01128 0.00385 -0.1160 0.4895 0.4482
2.750 0.8654 0.01135 0.00394 -0.1160 0.4864 0.4529
3.000 0.8921 0.01145 0.00405 -0.1160 0.4833 0.4583
3.500 0.9463 0.01155 0.00426 -0.1161 0.4737 0.4718
3.750 0.9718 0.01166 0.00439 -0.1158 0.4680 0.4810
4.000 0.9984 0.01171 0.00451 -0.1158 0.4626 0.4934
4.250 1.0242 0.01178 0.00466 -0.1157 0.4561 0.5105
4.500 1.0489 0.01186 0.00481 -0.1153 0.4480 0.5362
4.750 1.0723 0.01194 0.00499 -0.1147 0.4356 0.5741
5.000 1.0934 0.01208 0.00521 -0.1137 0.4118 0.6222
5.500 1.1091 0.01314 0.00620 -0.1070 0.3390 0.7400
5.750 1.1177 0.01324 0.00669 -0.1032 0.3284 0.9029
6.000 1.1386 0.01357 0.00704 -0.1023 0.3212 1.0000
6.250 1.1562 0.01402 0.00746 -0.1009 0.3154 1.0000
6.500 1.1756 0.01441 0.00785 -0.0997 0.3112 1.0000
6.750 1.1953 0.01479 0.00825 -0.0987 0.3075 1.0000
7.000 1.2135 0.01525 0.00870 -0.0975 0.3030 1.0000
7.500 1.2481 0.01629 0.00973 -0.0949 0.2917 1.0000
7.750 1.2652 0.01684 0.01029 -0.0937 0.2861 1.0000
8.000 1.2792 0.01757 0.01098 -0.0921 0.2801 1.0000
8.250 1.2985 0.01804 0.01149 -0.0912 0.2737 1.0000
8.500 1.3121 0.01883 0.01224 -0.0897 0.2654 1.0000
8.750 1.3294 0.01945 0.01289 -0.0887 0.2585 1.0000
9.000 1.3432 0.02029 0.01371 -0.0874 0.2501 1.0000
9.250 1.3582 0.02109 0.01452 -0.0863 0.2408 1.0000
9.500 1.3702 0.02213 0.01552 -0.0849 0.2292 1.0000
9.750 1.3781 0.02349 0.01680 -0.0833 0.2098 1.0000
10.000 1.3755 0.02566 0.01881 -0.0809 0.1804 1.0000
10.250 1.3726 0.02796 0.02100 -0.0786 0.1600 1.0000
10.500 1.3740 0.03000 0.02299 -0.0769 0.1469 1.0000
10.750 1.3746 0.03215 0.02508 -0.0751 0.1331 1.0000
11.000 1.3744 0.03442 0.02729 -0.0735 0.1165 1.0000
11.250 1.3604 0.03790 0.03057 -0.0711 0.0854 1.0000
11.500 1.3523 0.04101 0.03359 -0.0692 0.0669 1.0000
11.750 1.3326 0.04532 0.03773 -0.0669 0.0322 1.0000
12.000 1.3243 0.04872 0.04112 -0.0655 0.0163 1.0000
12.250 1.3289 0.05097 0.04343 -0.0648 0.0144 1.0000
12.500 1.3335 0.05325 0.04577 -0.0641 0.0132 1.0000
12.750 1.3393 0.05544 0.04803 -0.0636 0.0124 1.0000
13.000 1.3449 0.05767 0.05034 -0.0632 0.0118 1.0000
13.250 1.3498 0.06000 0.05274 -0.0628 0.0113 1.0000
13.500 1.3535 0.06252 0.05533 -0.0624 0.0109 1.0000
13.750 1.3562 0.06518 0.05806 -0.0620 0.0104 1.0000
14.000 1.3580 0.06796 0.06094 -0.0617 0.0100 1.0000
14.250 1.3618 0.07054 0.06359 -0.0615 0.0097 1.0000
14.500 1.3650 0.07320 0.06632 -0.0613 0.0094 1.0000
14.750 1.3674 0.07599 0.06920 -0.0612 0.0090 1.0000
15.000 1.3689 0.07892 0.07220 -0.0611 0.0087 1.0000
15.250 1.3698 0.08195 0.07531 -0.0611 0.0084 1.0000
15.500 1.3692 0.08519 0.07863 -0.0611 0.0082 1.0000
15.750 1.3679 0.08853 0.08205 -0.0612 0.0080 1.0000
16.000 1.3643 0.09223 0.08585 -0.0614 0.0079 1.0000
16.250 1.3586 0.09627 0.08999 -0.0617 0.0077 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 255 (MVA CA.6) AIRFOIL (goe255-il)