GOE 550 AIRFOIL (goe550-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 550 AIRFOIL (goe550-il) Reynolds number: 1,000,000 Max Cl/Cd: 114.84 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe550-il-1000000-n5.txt Download as CSV file: xf-goe550-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 550 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-16.250 -0.8487 0.08435 0.08149 -0.0638 1.0000 0.0175
-16.000 -0.8667 0.07858 0.07566 -0.0668 1.0000 0.0176
-15.750 -0.8891 0.07242 0.06942 -0.0701 1.0000 0.0177
-15.500 -0.9192 0.06566 0.06258 -0.0738 1.0000 0.0178
-15.250 -0.9707 0.05695 0.05376 -0.0786 1.0000 0.0174
-15.000 -1.1435 0.02997 0.02625 -0.1004 0.9976 0.0148
-14.750 -1.1264 0.02632 0.02242 -0.1054 0.9950 0.0155
-14.500 -1.1046 0.02433 0.02032 -0.1079 0.9931 0.0165
-14.250 -1.0835 0.02293 0.01882 -0.1090 0.9906 0.0174
-14.000 -1.0595 0.02170 0.01750 -0.1102 0.9883 0.0183
-13.750 -1.0334 0.02059 0.01634 -0.1115 0.9863 0.0196
-13.500 -1.0052 0.01968 0.01537 -0.1129 0.9848 0.0208
-13.250 -0.9759 0.01887 0.01448 -0.1143 0.9836 0.0219
-13.000 -0.9534 0.01810 0.01367 -0.1141 0.9801 0.0231
-12.750 -0.9270 0.01744 0.01296 -0.1146 0.9772 0.0243
-12.500 -0.8981 0.01685 0.01231 -0.1154 0.9750 0.0253
-12.250 -0.8682 0.01624 0.01164 -0.1165 0.9733 0.0262
-12.000 -0.8375 0.01561 0.01098 -0.1177 0.9718 0.0274
-11.750 -0.8154 0.01513 0.01046 -0.1169 0.9660 0.0284
-11.500 -0.7863 0.01466 0.00993 -0.1175 0.9626 0.0295
-11.250 -0.7534 0.01424 0.00946 -0.1188 0.9602 0.0304
-11.000 -0.7241 0.01372 0.00890 -0.1195 0.9557 0.0316
-10.750 -0.6937 0.01326 0.00840 -0.1204 0.9500 0.0329
-10.500 -0.6588 0.01283 0.00794 -0.1221 0.9456 0.0342
-10.250 -0.6266 0.01247 0.00752 -0.1231 0.9371 0.0353
-10.000 -0.5910 0.01211 0.00708 -0.1249 0.9262 0.0364
-9.750 -0.5585 0.01170 0.00660 -0.1260 0.9099 0.0381
-9.500 -0.5294 0.01141 0.00623 -0.1264 0.8912 0.0398
-9.250 -0.5023 0.01119 0.00592 -0.1262 0.8744 0.0413
-9.000 -0.4760 0.01102 0.00563 -0.1259 0.8589 0.0422
-8.750 -0.4506 0.01077 0.00529 -0.1254 0.8454 0.0437
-8.250 -0.3993 0.01032 0.00472 -0.1245 0.8218 0.0475
-8.000 -0.3732 0.01015 0.00447 -0.1241 0.8121 0.0490
-7.750 -0.3468 0.00998 0.00424 -0.1237 0.8025 0.0504
-7.500 -0.3207 0.00976 0.00397 -0.1233 0.7939 0.0532
-7.250 -0.2943 0.00957 0.00375 -0.1230 0.7854 0.0559
-7.000 -0.2675 0.00942 0.00354 -0.1227 0.7779 0.0584
-6.750 -0.2411 0.00921 0.00332 -0.1223 0.7697 0.0636
-6.500 -0.2144 0.00904 0.00313 -0.1220 0.7622 0.0691
-6.250 -0.1876 0.00885 0.00294 -0.1217 0.7548 0.0766
-6.000 -0.1609 0.00870 0.00278 -0.1214 0.7479 0.0834
-5.750 -0.1335 0.00858 0.00265 -0.1212 0.7408 0.0889
-5.500 -0.1064 0.00848 0.00251 -0.1209 0.7331 0.0933
-5.250 -0.0791 0.00835 0.00238 -0.1207 0.7262 0.0984
-5.000 -0.0517 0.00828 0.00226 -0.1204 0.7189 0.1022
-4.750 -0.0244 0.00820 0.00215 -0.1202 0.7121 0.1060
-4.500 0.0030 0.00809 0.00203 -0.1199 0.7044 0.1106
-4.250 0.0301 0.00802 0.00193 -0.1196 0.6965 0.1153
-4.000 0.0573 0.00795 0.00183 -0.1193 0.6867 0.1204
-3.750 0.0842 0.00787 0.00174 -0.1190 0.6756 0.1273
-3.500 0.1106 0.00781 0.00165 -0.1186 0.6622 0.1365
-3.000 0.1639 0.00770 0.00154 -0.1178 0.6356 0.1624
-2.750 0.1912 0.00768 0.00151 -0.1175 0.6258 0.1733
-2.500 0.2181 0.00767 0.00149 -0.1172 0.6153 0.1815
-2.250 0.2455 0.00768 0.00146 -0.1169 0.6042 0.1874
-2.000 0.2727 0.00769 0.00144 -0.1166 0.5943 0.1912
-1.750 0.2998 0.00770 0.00143 -0.1163 0.5843 0.1964
-1.500 0.3272 0.00772 0.00142 -0.1160 0.5749 0.2006
-1.250 0.3542 0.00776 0.00141 -0.1157 0.5648 0.2034
-1.000 0.3814 0.00780 0.00141 -0.1154 0.5542 0.2058
-0.750 0.4086 0.00782 0.00142 -0.1151 0.5458 0.2096
-0.500 0.4358 0.00785 0.00143 -0.1148 0.5373 0.2129
-0.250 0.4629 0.00789 0.00144 -0.1145 0.5294 0.2161
0.000 0.4900 0.00793 0.00146 -0.1141 0.5204 0.2190
0.250 0.5171 0.00799 0.00149 -0.1138 0.5116 0.2213
0.500 0.5440 0.00803 0.00152 -0.1135 0.5032 0.2259
0.750 0.5711 0.00806 0.00156 -0.1132 0.4957 0.2313
1.000 0.5977 0.00814 0.00160 -0.1128 0.4856 0.2351
1.250 0.6243 0.00821 0.00165 -0.1124 0.4737 0.2377
1.500 0.6507 0.00829 0.00170 -0.1120 0.4622 0.2417
1.750 0.6768 0.00838 0.00177 -0.1115 0.4508 0.2471
2.000 0.7034 0.00845 0.00183 -0.1111 0.4399 0.2516
2.500 0.7555 0.00865 0.00198 -0.1102 0.4181 0.2600
2.750 0.7811 0.00877 0.00207 -0.1097 0.4042 0.2656
3.000 0.8061 0.00893 0.00218 -0.1090 0.3880 0.2702
3.500 0.8560 0.00924 0.00242 -0.1077 0.3586 0.2814
3.750 0.8814 0.00937 0.00254 -0.1071 0.3489 0.2869
4.000 0.9065 0.00950 0.00266 -0.1065 0.3402 0.2942
4.250 0.9320 0.00961 0.00278 -0.1060 0.3333 0.3020
4.500 0.9571 0.00973 0.00292 -0.1054 0.3259 0.3115
4.750 0.9822 0.00984 0.00306 -0.1048 0.3197 0.3239
5.000 1.0071 0.00992 0.00320 -0.1041 0.3131 0.3473
5.250 1.0307 0.00997 0.00341 -0.1034 0.3057 0.4222
5.500 1.0552 0.01002 0.00358 -0.1027 0.2993 0.4751
5.750 1.0789 0.01012 0.00376 -0.1019 0.2926 0.5156
6.000 1.0989 0.00993 0.00402 -0.1004 0.2820 0.7170
6.500 1.1553 0.01006 0.00454 -0.1008 0.2500 1.0000
6.750 1.1769 0.01036 0.00478 -0.0996 0.2353 1.0000
7.000 1.1958 0.01078 0.00507 -0.0980 0.2131 1.0000
7.250 1.2114 0.01136 0.00547 -0.0958 0.1803 1.0000
7.500 1.2200 0.01222 0.00607 -0.0924 0.1358 1.0000
7.750 1.2343 0.01271 0.00647 -0.0899 0.1202 1.0000
8.000 1.2504 0.01311 0.00683 -0.0877 0.1102 1.0000
8.250 1.2668 0.01350 0.00719 -0.0857 0.1028 1.0000
8.500 1.2848 0.01384 0.00751 -0.0839 0.0979 1.0000
8.750 1.3022 0.01419 0.00786 -0.0821 0.0914 1.0000
9.000 1.3181 0.01463 0.00826 -0.0800 0.0835 1.0000
9.250 1.3319 0.01516 0.00873 -0.0777 0.0718 1.0000
9.500 1.3445 0.01576 0.00926 -0.0753 0.0604 1.0000
9.750 1.3577 0.01634 0.00980 -0.0731 0.0531 1.0000
10.250 1.3845 0.01752 0.01096 -0.0688 0.0427 1.0000
10.500 1.3966 0.01820 0.01163 -0.0666 0.0382 1.0000
10.750 1.4098 0.01884 0.01227 -0.0646 0.0350 1.0000
11.000 1.4218 0.01956 0.01300 -0.0626 0.0320 1.0000
11.250 1.4334 0.02033 0.01378 -0.0606 0.0289 1.0000
11.500 1.4447 0.02116 0.01462 -0.0587 0.0266 1.0000
11.750 1.4560 0.02200 0.01549 -0.0569 0.0243 1.0000
12.000 1.4659 0.02298 0.01648 -0.0550 0.0223 1.0000
12.250 1.4764 0.02396 0.01749 -0.0533 0.0203 1.0000
12.500 1.4851 0.02511 0.01865 -0.0516 0.0186 1.0000
12.750 1.4947 0.02623 0.01982 -0.0501 0.0174 1.0000
13.000 1.5033 0.02748 0.02110 -0.0486 0.0162 1.0000
13.250 1.5104 0.02889 0.02254 -0.0471 0.0150 1.0000
13.500 1.5186 0.03025 0.02395 -0.0458 0.0144 1.0000
13.750 1.5261 0.03173 0.02548 -0.0446 0.0138 1.0000
14.000 1.5318 0.03340 0.02719 -0.0434 0.0130 1.0000
14.250 1.5370 0.03516 0.02900 -0.0422 0.0125 1.0000
14.500 1.5415 0.03705 0.03094 -0.0412 0.0119 1.0000
14.750 1.5470 0.03891 0.03287 -0.0404 0.0117 1.0000
15.000 1.5516 0.04090 0.03493 -0.0396 0.0113 1.0000
15.250 1.5552 0.04305 0.03714 -0.0389 0.0109 1.0000
15.500 1.5577 0.04537 0.03953 -0.0384 0.0105 1.0000
15.750 1.5596 0.04784 0.04206 -0.0379 0.0103 1.0000
16.000 1.5605 0.05048 0.04476 -0.0376 0.0100 1.0000
16.250 1.5597 0.05337 0.04773 -0.0373 0.0097 1.0000
16.500 1.5584 0.05642 0.05085 -0.0373 0.0095 1.0000
16.750 1.5586 0.05936 0.05387 -0.0373 0.0093 1.0000
17.000 1.5581 0.06244 0.05703 -0.0375 0.0091 1.0000
17.250 1.5563 0.06576 0.06043 -0.0378 0.0090 1.0000
17.500 1.5548 0.06909 0.06384 -0.0382 0.0087 1.0000
17.750 1.5516 0.07274 0.06757 -0.0388 0.0085 1.0000
18.000 1.5477 0.07653 0.07145 -0.0395 0.0085 1.0000
18.250 1.5439 0.08036 0.07535 -0.0403 0.0082 1.0000
18.500 1.5388 0.08447 0.07955 -0.0413 0.0080 1.0000
18.750 1.5337 0.08867 0.08384 -0.0424 0.0079 1.0000
19.000 1.5278 0.09304 0.08830 -0.0437 0.0078 1.0000
19.250 1.5208 0.09766 0.09300 -0.0452 0.0077 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 550 AIRFOIL (goe550-il)