GOE 547 AIRFOIL (goe547-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 547 AIRFOIL (goe547-il) Reynolds number: 500,000 Max Cl/Cd: 105.79 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe547-il-500000-n5.txt Download as CSV file: xf-goe547-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 547 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
------ -------- --------- --------- -------- -------- --------
-12.250 -0.8004 0.03682 0.03411 -0.0988 0.9895 0.0139
-12.000 -0.7954 0.03107 0.02788 -0.1046 0.9833 0.0143
-11.750 -0.7668 0.03110 0.02796 -0.1053 0.9804 0.0149
-11.500 -0.7367 0.03101 0.02787 -0.1065 0.9781 0.0153
-11.250 -0.7076 0.03041 0.02721 -0.1081 0.9761 0.0158
-11.000 -0.6868 0.02930 0.02598 -0.1087 0.9711 0.0165
-10.750 -0.6651 0.02758 0.02402 -0.1099 0.9669 0.0174
-10.500 -0.6422 0.02531 0.02140 -0.1117 0.9641 0.0184
-10.250 -0.6238 0.02403 0.01988 -0.1113 0.9573 0.0190
-10.000 -0.5971 0.02320 0.01899 -0.1122 0.9539 0.0195
-9.750 -0.5655 0.02274 0.01847 -0.1137 0.9514 0.0200
-9.500 -0.5402 0.02236 0.01805 -0.1138 0.9442 0.0206
-9.250 -0.5078 0.02185 0.01744 -0.1154 0.9404 0.0214
-9.000 -0.4747 0.02096 0.01638 -0.1175 0.9376 0.0224
-8.750 -0.4476 0.02021 0.01546 -0.1180 0.9320 0.0233
-8.500 -0.4173 0.01942 0.01450 -0.1192 0.9275 0.0241
-8.250 -0.3846 0.01883 0.01375 -0.1208 0.9240 0.0247
-8.000 -0.3555 0.01816 0.01291 -0.1216 0.9188 0.0251
-7.750 -0.3307 0.01643 0.01095 -0.1221 0.9127 0.0262
-7.500 -0.3002 0.01571 0.01012 -0.1231 0.9080 0.0269
-7.250 -0.2741 0.01512 0.00943 -0.1231 0.9015 0.0275
-7.000 -0.2459 0.01464 0.00885 -0.1235 0.8960 0.0282
-6.750 -0.2183 0.01412 0.00822 -0.1237 0.8907 0.0288
-6.500 -0.1929 0.01358 0.00758 -0.1234 0.8843 0.0292
-6.250 -0.1657 0.01308 0.00696 -0.1235 0.8787 0.0297
-6.000 -0.1407 0.01262 0.00641 -0.1230 0.8727 0.0301
-5.750 -0.1148 0.01220 0.00590 -0.1227 0.8669 0.0306
-5.500 -0.0881 0.01184 0.00545 -0.1226 0.8618 0.0312
-5.250 -0.0631 0.01153 0.00507 -0.1221 0.8557 0.0318
-5.000 -0.0368 0.01121 0.00467 -0.1218 0.8499 0.0322
-4.750 -0.0111 0.01092 0.00431 -0.1214 0.8443 0.0326
-4.500 0.0146 0.01066 0.00399 -0.1210 0.8385 0.0328
-4.250 0.0412 0.01042 0.00368 -0.1208 0.8333 0.0331
-4.000 0.0662 0.01010 0.00331 -0.1202 0.8273 0.0339
-3.750 0.0922 0.00986 0.00300 -0.1198 0.8211 0.0348
-3.500 0.1185 0.00967 0.00277 -0.1195 0.8155 0.0356
-3.250 0.1445 0.00950 0.00256 -0.1191 0.8092 0.0367
-3.000 0.1713 0.00937 0.00238 -0.1188 0.8035 0.0378
-2.750 0.1973 0.00923 0.00223 -0.1184 0.7971 0.0392
-2.500 0.2237 0.00912 0.00209 -0.1181 0.7904 0.0415
-2.250 0.2500 0.00900 0.00198 -0.1177 0.7842 0.0478
-2.000 0.2761 0.00887 0.00190 -0.1174 0.7773 0.0653
-1.750 0.3025 0.00880 0.00181 -0.1170 0.7707 0.0761
-1.500 0.3286 0.00871 0.00173 -0.1166 0.7634 0.0853
-1.250 0.3547 0.00861 0.00165 -0.1163 0.7564 0.0995
-1.000 0.3793 0.00831 0.00160 -0.1157 0.7487 0.1822
-0.750 0.4053 0.00824 0.00157 -0.1154 0.7415 0.2099
-0.500 0.4311 0.00816 0.00155 -0.1149 0.7335 0.2355
-0.250 0.4563 0.00805 0.00154 -0.1144 0.7245 0.2738
0.000 0.4806 0.00798 0.00153 -0.1137 0.7112 0.3154
0.250 0.5046 0.00787 0.00154 -0.1129 0.6990 0.3763
0.750 0.5502 0.00769 0.00164 -0.1108 0.6682 0.5166
1.000 0.5735 0.00769 0.00169 -0.1098 0.6533 0.5527
1.250 0.5972 0.00770 0.00174 -0.1088 0.6402 0.5882
1.500 0.6205 0.00769 0.00181 -0.1079 0.6277 0.6303
1.750 0.6431 0.00761 0.00187 -0.1067 0.6172 0.6890
2.250 0.7439 0.00731 0.00218 -0.1167 0.5979 1.0000
2.500 0.7671 0.00742 0.00226 -0.1157 0.5892 1.0000
3.000 0.8113 0.00770 0.00244 -0.1132 0.5632 1.0000
3.250 0.8326 0.00787 0.00255 -0.1118 0.5442 1.0000
3.500 0.8523 0.00808 0.00267 -0.1100 0.5197 1.0000
3.750 0.8723 0.00830 0.00280 -0.1084 0.4950 1.0000
4.000 0.8907 0.00859 0.00298 -0.1065 0.4649 1.0000
4.250 0.9075 0.00897 0.00319 -0.1042 0.4286 1.0000
4.500 0.9217 0.00949 0.00348 -0.1016 0.3807 1.0000
4.750 0.9353 0.01008 0.00382 -0.0989 0.3319 1.0000
5.000 0.9508 0.01060 0.00415 -0.0966 0.2950 1.0000
5.250 0.9670 0.01108 0.00448 -0.0944 0.2621 1.0000
5.500 0.9832 0.01156 0.00481 -0.0923 0.2319 1.0000
5.750 0.9991 0.01201 0.00512 -0.0901 0.2052 1.0000
6.000 1.0144 0.01245 0.00544 -0.0878 0.1768 1.0000
6.250 1.0239 0.01317 0.00591 -0.0845 0.1294 1.0000
6.500 1.0385 0.01367 0.00632 -0.0821 0.1124 1.0000
6.750 1.0556 0.01406 0.00669 -0.0802 0.1029 1.0000
7.500 1.1011 0.01559 0.00801 -0.0737 0.0547 1.0000
7.750 1.1163 0.01612 0.00852 -0.0717 0.0455 1.0000
8.000 1.1300 0.01675 0.00909 -0.0694 0.0301 1.0000
8.250 1.1414 0.01753 0.00976 -0.0669 0.0186 1.0000
8.500 1.1557 0.01815 0.01040 -0.0648 0.0155 1.0000
8.750 1.1714 0.01870 0.01100 -0.0630 0.0144 1.0000
9.000 1.1857 0.01934 0.01170 -0.0610 0.0131 1.0000
9.250 1.1993 0.02004 0.01245 -0.0590 0.0123 1.0000
9.500 1.2109 0.02089 0.01336 -0.0568 0.0113 1.0000
9.750 1.2242 0.02163 0.01418 -0.0548 0.0109 1.0000
10.000 1.2374 0.02240 0.01501 -0.0530 0.0105 1.0000
10.250 1.2503 0.02320 0.01587 -0.0512 0.0099 1.0000
10.500 1.2620 0.02410 0.01685 -0.0493 0.0096 1.0000
10.750 1.2731 0.02506 0.01787 -0.0474 0.0091 1.0000
11.000 1.2836 0.02608 0.01895 -0.0456 0.0086 1.0000
11.250 1.2917 0.02731 0.02024 -0.0436 0.0083 1.0000
11.500 1.2963 0.02883 0.02185 -0.0414 0.0080 1.0000
11.750 1.3046 0.03011 0.02321 -0.0396 0.0078 1.0000
12.000 1.3129 0.03142 0.02461 -0.0379 0.0076 1.0000
12.250 1.3197 0.03288 0.02616 -0.0362 0.0075 1.0000
12.500 1.3267 0.03437 0.02776 -0.0347 0.0072 1.0000
12.750 1.3324 0.03601 0.02949 -0.0332 0.0071 1.0000
13.000 1.3372 0.03777 0.03134 -0.0317 0.0069 1.0000
13.250 1.3417 0.03961 0.03327 -0.0304 0.0067 1.0000
13.500 1.3448 0.04164 0.03541 -0.0291 0.0066 1.0000
13.750 1.3497 0.04354 0.03739 -0.0281 0.0064 1.0000
14.000 1.3515 0.04582 0.03977 -0.0271 0.0063 1.0000
14.250 1.3550 0.04801 0.04204 -0.0264 0.0061 1.0000
14.500 1.3574 0.05037 0.04449 -0.0257 0.0060 1.0000
14.750 1.3575 0.05305 0.04727 -0.0252 0.0059 1.0000
15.000 1.3545 0.05619 0.05050 -0.0248 0.0057 1.0000
15.250 1.3512 0.05949 0.05394 -0.0245 0.0057 1.0000
15.500 1.3463 0.06310 0.05766 -0.0245 0.0056 1.0000
15.750 1.3453 0.06633 0.06102 -0.0247 0.0055 1.0000
16.000 1.3424 0.06991 0.06473 -0.0250 0.0054 1.0000
16.250 1.3388 0.07370 0.06866 -0.0256 0.0054 1.0000
16.500 1.3340 0.07779 0.07290 -0.0264 0.0053 1.0000
16.750 1.3269 0.08230 0.07754 -0.0275 0.0053 1.0000
17.000 1.3209 0.08686 0.08224 -0.0289 0.0052 1.0000
17.250 1.3127 0.09188 0.08741 -0.0305 0.0051 1.0000
17.500 1.3053 0.09696 0.09263 -0.0325 0.0051 1.0000
17.750 1.2946 0.10273 0.09855 -0.0348 0.0051 1.0000
18.000 1.2860 0.10835 0.10432 -0.0374 0.0050 1.0000
18.250 1.2753 0.11446 0.11057 -0.0402 0.0050 1.0000
18.500 1.2636 0.12094 0.11720 -0.0435 0.0049 1.0000
18.750 1.2525 0.12745 0.12385 -0.0469 0.0049 1.0000
19.000 1.2411 0.13414 0.13068 -0.0506 0.0049 1.0000
19.250 1.2292 0.14111 0.13779 -0.0546 0.0049 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 547 AIRFOIL (goe547-il)