GOE 114 (MVA MK.1) AIRFOIL (goe114-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 114 (MVA MK.1) AIRFOIL (goe114-il) Reynolds number: 500,000 Max Cl/Cd: 93.42 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe114-il-500000.txt Download as CSV file: xf-goe114-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 114 (MVA MK.1) 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.3647 0.09703 0.09480 -0.0426 1.0000 0.0090
-9.250 -0.3636 0.09340 0.09121 -0.0436 1.0000 0.0090
-7.750 -0.3767 0.06681 0.06479 -0.0569 0.9937 0.0090
-7.500 -0.3527 0.05621 0.05399 -0.0780 0.9838 0.0090
-7.250 -0.3337 0.04964 0.04711 -0.0886 0.9757 0.0090
-7.000 -0.3177 0.04045 0.03745 -0.0981 0.9699 0.0094
-6.750 -0.3030 0.03329 0.02989 -0.1022 0.9615 0.0101
-6.500 -0.2786 0.02970 0.02619 -0.1051 0.9565 0.0116
-6.250 -0.2559 0.02704 0.02325 -0.1061 0.9478 0.0126
-6.000 -0.2300 0.02460 0.02048 -0.1070 0.9411 0.0136
-5.750 -0.2074 0.02270 0.01828 -0.1066 0.9314 0.0148
-5.500 -0.1832 0.02121 0.01650 -0.1060 0.9223 0.0164
-5.250 -0.1540 0.02183 0.01689 -0.1057 0.9147 0.0188
-5.000 -0.1305 0.02079 0.01553 -0.1048 0.9052 0.0190
-4.750 -0.1121 0.01604 0.01045 -0.1040 0.8967 0.0227
-4.500 -0.0870 0.01476 0.00907 -0.1036 0.8895 0.0262
-4.250 -0.0615 0.01388 0.00805 -0.1029 0.8813 0.0294
-4.000 -0.0340 0.01250 0.00653 -0.1014 0.8747 0.0153
-3.750 -0.0099 0.01144 0.00537 -0.1003 0.8667 0.0139
-3.500 0.0152 0.01067 0.00445 -0.0996 0.8600 0.0152
-3.250 0.0410 0.01023 0.00392 -0.0992 0.8525 0.0178
-3.000 0.0673 0.00975 0.00326 -0.0988 0.8459 0.0185
-2.750 0.0938 0.00938 0.00273 -0.0984 0.8389 0.0189
-2.500 0.1209 0.00909 0.00221 -0.0981 0.8325 0.0232
-2.250 0.1452 0.00823 0.00175 -0.0977 0.8256 0.1706
-2.000 0.1708 0.00793 0.00189 -0.0976 0.8192 0.3128
-1.750 0.1984 0.00805 0.00195 -0.0975 0.8124 0.3330
-1.500 0.2259 0.00815 0.00198 -0.0975 0.8061 0.3470
-1.250 0.2531 0.00822 0.00204 -0.0974 0.7993 0.3585
-1.000 0.2803 0.00818 0.00197 -0.0973 0.7930 0.3622
-0.750 0.3074 0.00815 0.00193 -0.0972 0.7860 0.3658
-0.500 0.3347 0.00815 0.00190 -0.0971 0.7796 0.3698
-0.250 0.3618 0.00814 0.00186 -0.0970 0.7724 0.3734
0.000 0.3888 0.00809 0.00184 -0.0969 0.7657 0.3775
0.250 0.4157 0.00808 0.00183 -0.0968 0.7578 0.3821
0.500 0.4424 0.00809 0.00182 -0.0966 0.7479 0.3866
0.750 0.4660 0.00801 0.00169 -0.0955 0.7194 0.3911
1.000 0.4909 0.00802 0.00168 -0.0949 0.6965 0.3961
1.250 0.5153 0.00808 0.00164 -0.0941 0.6669 0.4008
1.500 0.5380 0.00820 0.00165 -0.0931 0.6236 0.4057
1.750 0.5516 0.00897 0.00179 -0.0903 0.4926 0.4103
2.000 0.5557 0.01089 0.00245 -0.0865 0.2192 0.4140
2.250 0.5700 0.01209 0.00305 -0.0843 0.0211 0.4189
2.500 0.5952 0.01227 0.00336 -0.0838 0.0198 0.4256
2.750 0.6192 0.01253 0.00382 -0.0830 0.0165 0.4331
3.000 0.6421 0.01291 0.00436 -0.0820 0.0148 0.4424
3.250 0.6629 0.01347 0.00509 -0.0805 0.0141 0.4548
3.500 0.6824 0.01407 0.00591 -0.0788 0.0152 0.4762
3.750 0.6985 0.01469 0.00693 -0.0764 0.0179 0.5633
5.000 0.7978 0.00854 0.00216 -0.0688 0.0158 1.0000
5.250 0.8234 0.00995 0.00360 -0.0685 0.0134 1.0000
5.500 0.8527 0.01296 0.00660 -0.0690 0.0112 1.0000
5.750 0.9030 0.02984 0.02343 -0.0720 0.0103 1.0000
6.000 0.9239 0.03211 0.02594 -0.0703 0.0103 1.0000
6.250 0.9430 0.03425 0.02832 -0.0685 0.0102 1.0000
6.500 0.9610 0.03609 0.03040 -0.0664 0.0101 1.0000
6.750 0.9821 0.03596 0.03051 -0.0645 0.0093 1.0000
7.000 1.0025 0.03714 0.03190 -0.0626 0.0083 1.0000
7.250 1.0187 0.03928 0.03428 -0.0603 0.0075 1.0000
7.500 1.0324 0.04174 0.03697 -0.0580 0.0071 1.0000
7.750 1.0433 0.04442 0.03988 -0.0554 0.0067 1.0000
8.000 1.0518 0.04720 0.04288 -0.0528 0.0064 1.0000
8.250 1.0577 0.05006 0.04596 -0.0501 0.0062 1.0000
8.500 1.0613 0.05297 0.04908 -0.0474 0.0060 1.0000
8.750 1.0613 0.05602 0.05232 -0.0444 0.0058 1.0000
9.000 1.0587 0.05902 0.05549 -0.0414 0.0057 1.0000
9.250 1.0532 0.06197 0.05861 -0.0383 0.0056 1.0000
9.500 1.0422 0.06464 0.06142 -0.0346 0.0056 1.0000
9.750 1.0275 0.06728 0.06419 -0.0309 0.0055 1.0000
10.000 1.0122 0.07015 0.06719 -0.0281 0.0055 1.0000
10.250 0.9957 0.07346 0.07063 -0.0261 0.0055 1.0000
10.500 0.9787 0.07710 0.07439 -0.0251 0.0055 1.0000
10.750 0.9604 0.08133 0.07874 -0.0250 0.0055 1.0000
11.000 0.9423 0.08606 0.08360 -0.0260 0.0055 1.0000
11.250 0.9236 0.09153 0.08918 -0.0280 0.0056 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 114 (MVA MK.1) AIRFOIL (goe114-il)