GOE 574 AIRFOIL (goe574-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 574 AIRFOIL (goe574-il) Reynolds number: 1,000,000 Max Cl/Cd: 121.85 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe574-il-1000000.txt Download as CSV file: xf-goe574-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 574 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.2097 0.10103 0.09955 -0.0431 0.9935 0.0105
-9.750 -0.2038 0.09634 0.09486 -0.0465 0.9920 0.0114
-9.500 -0.1977 0.09185 0.09037 -0.0495 0.9905 0.0114
-9.250 -0.1888 0.08734 0.08586 -0.0528 0.9890 0.0115
-9.000 -0.1797 0.08243 0.08095 -0.0563 0.9877 0.0115
-8.750 -0.1691 0.07759 0.07611 -0.0601 0.9865 0.0115
-8.500 -0.1597 0.07252 0.07104 -0.0639 0.9856 0.0116
-8.250 -0.1567 0.06841 0.06694 -0.0656 0.9819 0.0116
-8.000 -0.1621 0.06199 0.06054 -0.0685 0.9782 0.0119
-7.750 -0.1489 0.05824 0.05679 -0.0718 0.9755 0.0121
-7.500 -0.1476 0.05589 0.05445 -0.0717 0.9676 0.0125
-6.750 -0.2065 0.05877 0.05719 -0.0792 0.9514 0.0122
-6.500 -0.2004 0.05611 0.05449 -0.0788 0.9214 0.0124
-6.250 -0.1536 0.05117 0.04942 -0.0886 0.9064 0.0137
-5.750 -0.0767 0.01993 0.01645 -0.1103 0.8541 0.0127
-5.500 -0.0426 0.01839 0.01451 -0.1125 0.8219 0.0137
-5.250 -0.0322 0.01528 0.01077 -0.1090 0.8000 0.0149
-5.000 -0.0103 0.01488 0.01011 -0.1076 0.7816 0.0158
-4.750 0.0042 0.01324 0.00816 -0.1050 0.7654 0.0169
-4.500 0.0267 0.01337 0.00823 -0.1040 0.7484 0.0176
-4.250 0.0486 0.01337 0.00816 -0.1028 0.7309 0.0185
-4.000 0.0693 0.01297 0.00759 -0.1013 0.7143 0.0197
-3.750 0.0910 0.01273 0.00718 -0.0998 0.6978 0.0209
-3.250 0.1319 0.01126 0.00535 -0.0967 0.6657 0.0234
-3.000 0.1546 0.01116 0.00518 -0.0957 0.6490 0.0248
-2.750 0.1775 0.01083 0.00473 -0.0946 0.6325 0.0259
-2.500 0.2006 0.01053 0.00432 -0.0936 0.6160 0.0271
-2.250 0.2240 0.01032 0.00399 -0.0926 0.5998 0.0282
-2.000 0.2481 0.01028 0.00386 -0.0918 0.5834 0.0293
-1.750 0.2718 0.01012 0.00358 -0.0909 0.5673 0.0297
-1.500 0.2955 0.00996 0.00332 -0.0900 0.5513 0.0301
-1.250 0.3174 0.00955 0.00282 -0.0887 0.5361 0.0304
-1.000 0.3385 0.00920 0.00237 -0.0873 0.5220 0.0311
-0.500 0.3834 0.00885 0.00185 -0.0849 0.4984 0.0335
-0.250 0.4070 0.00878 0.00172 -0.0840 0.4890 0.0349
0.000 0.4306 0.00876 0.00164 -0.0832 0.4806 0.0358
0.250 0.4549 0.00872 0.00156 -0.0824 0.4739 0.0366
0.500 0.4788 0.00872 0.00151 -0.0816 0.4676 0.0372
0.750 0.5032 0.00871 0.00148 -0.0809 0.4626 0.0381
1.000 0.5275 0.00871 0.00145 -0.0801 0.4578 0.0401
1.250 0.5512 0.00873 0.00146 -0.0793 0.4531 0.0506
1.500 0.5700 0.00835 0.00155 -0.0775 0.4497 0.2436
1.750 0.5602 0.00703 0.00169 -0.0695 0.4471 0.7853
2.000 0.8285 0.00780 0.00265 -0.1250 0.4358 1.0000
2.250 0.8516 0.00785 0.00271 -0.1241 0.4319 1.0000
2.500 0.8742 0.00793 0.00277 -0.1230 0.4283 1.0000
2.750 0.8961 0.00804 0.00285 -0.1219 0.4245 1.0000
3.000 0.9190 0.00809 0.00289 -0.1209 0.4180 1.0000
3.250 0.9412 0.00817 0.00297 -0.1198 0.4131 1.0000
3.500 0.9626 0.00829 0.00307 -0.1185 0.4083 1.0000
3.750 0.9856 0.00833 0.00312 -0.1176 0.4023 1.0000
4.000 1.0066 0.00846 0.00321 -0.1163 0.3949 1.0000
4.250 1.0291 0.00851 0.00329 -0.1152 0.3868 1.0000
4.500 1.0503 0.00862 0.00338 -0.1139 0.3777 1.0000
4.750 1.0701 0.00880 0.00346 -0.1124 0.3532 1.0000
5.000 1.0763 0.00977 0.00387 -0.1083 0.2495 1.0000
5.250 1.0657 0.01160 0.00490 -0.1009 0.1005 1.0000
5.500 1.0667 0.01266 0.00564 -0.0955 0.0192 1.0000
5.750 1.0824 0.01296 0.00597 -0.0931 0.0168 1.0000
6.000 1.0966 0.01331 0.00637 -0.0903 0.0145 1.0000
6.250 1.1088 0.01367 0.00680 -0.0871 0.0132 1.0000
6.500 1.1207 0.01393 0.00709 -0.0839 0.0125 1.0000
6.750 1.1310 0.01426 0.00747 -0.0803 0.0119 1.0000
7.000 1.1415 0.01461 0.00788 -0.0769 0.0113 1.0000
7.250 1.1507 0.01503 0.00835 -0.0732 0.0108 1.0000
7.500 1.1605 0.01547 0.00884 -0.0698 0.0103 1.0000
7.750 1.1682 0.01603 0.00946 -0.0660 0.0099 1.0000
8.000 1.1699 0.01684 0.01037 -0.0611 0.0094 1.0000
8.250 1.1623 0.01808 0.01174 -0.0547 0.0090 1.0000
8.500 1.1698 0.01876 0.01249 -0.0513 0.0087 1.0000
8.750 1.1797 0.01936 0.01314 -0.0484 0.0085 1.0000
9.000 1.1901 0.01998 0.01382 -0.0457 0.0081 1.0000
9.250 1.1897 0.02116 0.01508 -0.0414 0.0079 1.0000
9.500 1.1981 0.02197 0.01594 -0.0388 0.0077 1.0000
9.750 1.1983 0.02330 0.01736 -0.0352 0.0076 1.0000
10.000 1.2044 0.02440 0.01851 -0.0327 0.0072 1.0000
10.250 1.2088 0.02569 0.01986 -0.0303 0.0069 1.0000
10.500 1.2033 0.02787 0.02215 -0.0269 0.0071 1.0000
10.750 1.2112 0.02913 0.02345 -0.0255 0.0068 1.0000
11.000 1.2176 0.03059 0.02497 -0.0241 0.0065 1.0000
11.250 1.2184 0.03261 0.02704 -0.0222 0.0063 1.0000
11.500 1.2202 0.03464 0.02917 -0.0201 0.0065 1.0000
11.750 1.2232 0.03658 0.03112 -0.0185 0.0061 1.0000
12.000 1.2295 0.03868 0.03333 -0.0157 0.0057 1.0000
12.250 1.2363 0.04029 0.03504 -0.0147 0.0055 1.0000
12.500 1.2471 0.04200 0.03685 -0.0131 0.0055 1.0000
12.750 1.2554 0.04374 0.03869 -0.0119 0.0054 1.0000
13.000 1.2641 0.04573 0.04083 -0.0104 0.0051 1.0000
13.250 1.2724 0.04808 0.04333 -0.0088 0.0052 1.0000
13.500 1.2759 0.05058 0.04599 -0.0076 0.0052 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 574 AIRFOIL (goe574-il)