GOE 276 (DAIMLER VII) AIRFOIL (goe276-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 276 (DAIMLER VII) AIRFOIL (goe276-il) Reynolds number: 1,000,000 Max Cl/Cd: 114.08 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe276-il-1000000.txt Download as CSV file: xf-goe276-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 276 (DAIMLER VII) 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3857 0.09186 0.09028 -0.0180 1.0000 0.0129
-7.750 -0.3932 0.08914 0.08760 -0.0179 1.0000 0.0136
-7.500 -0.3929 0.08560 0.08409 -0.0206 1.0000 0.0137
-7.250 -0.3894 0.08178 0.08029 -0.0238 1.0000 0.0137
-7.000 -0.3771 0.07731 0.07581 -0.0285 0.9994 0.0138
-6.750 -0.3482 0.07130 0.06976 -0.0374 0.9973 0.0138
-6.500 -0.3281 0.06623 0.06468 -0.0421 0.9954 0.0142
-6.250 -0.2999 0.06280 0.06121 -0.0474 0.9933 0.0146
-6.000 -0.2697 0.05913 0.05750 -0.0531 0.9895 0.0153
-5.750 -0.2183 0.05190 0.05009 -0.0654 0.9859 0.0175
-5.250 -0.1506 0.03863 0.03646 -0.0781 0.9737 0.0181
-5.000 -0.1218 0.03703 0.03483 -0.0804 0.9632 0.0185
-4.750 -0.0883 0.03514 0.03284 -0.0835 0.9509 0.0196
-4.500 -0.0440 0.03183 0.02921 -0.0865 0.9357 0.0223
-4.250 -0.0128 0.02845 0.02552 -0.0880 0.9183 0.0224
-4.000 0.0135 0.02297 0.01963 -0.0900 0.9042 0.0234
-3.750 0.0394 0.02212 0.01869 -0.0904 0.8924 0.0240
-3.500 0.0664 0.02104 0.01747 -0.0907 0.8817 0.0250
-3.250 0.0949 0.01992 0.01616 -0.0907 0.8712 0.0272
-3.000 0.1250 0.01993 0.01595 -0.0901 0.8611 0.0288
-2.750 0.1540 0.01851 0.01429 -0.0903 0.8519 0.0289
-2.500 0.1830 0.01455 0.00990 -0.0913 0.8421 0.0304
-2.250 0.2106 0.01385 0.00911 -0.0914 0.8286 0.0311
-2.000 0.2384 0.01324 0.00839 -0.0914 0.8120 0.0321
-1.750 0.2665 0.01268 0.00769 -0.0914 0.7929 0.0334
-1.500 0.2944 0.01238 0.00722 -0.0912 0.7693 0.0357
-0.750 0.3716 0.01249 0.00687 -0.0956 0.5789 0.0342
-0.500 0.4004 0.01072 0.00407 -0.0897 0.4801 0.0348
-0.250 0.4281 0.01033 0.00350 -0.0896 0.4530 0.0337
0.000 0.4562 0.01001 0.00310 -0.0896 0.4388 0.0332
0.250 0.4844 0.00983 0.00287 -0.0897 0.4287 0.0336
0.500 0.5128 0.00960 0.00259 -0.0897 0.4201 0.0336
0.750 0.5411 0.00946 0.00241 -0.0898 0.4129 0.0341
1.000 0.5696 0.00936 0.00229 -0.0899 0.4066 0.0353
1.250 0.5978 0.00933 0.00222 -0.0900 0.4002 0.0362
1.500 0.6262 0.00928 0.00216 -0.0901 0.3939 0.0370
1.750 0.6543 0.00929 0.00214 -0.0901 0.3874 0.0375
2.000 0.6827 0.00925 0.00209 -0.0902 0.3820 0.0386
2.250 0.7110 0.00923 0.00205 -0.0902 0.3759 0.0425
2.500 0.7391 0.00926 0.00207 -0.0903 0.3694 0.0485
2.750 0.7619 0.00738 0.00226 -0.0898 0.3630 1.0000
3.000 0.7898 0.00751 0.00232 -0.0898 0.3551 1.0000
3.250 0.8178 0.00760 0.00239 -0.0898 0.3474 1.0000
3.500 0.8457 0.00772 0.00247 -0.0898 0.3387 1.0000
3.750 0.8734 0.00785 0.00257 -0.0898 0.3288 1.0000
4.000 0.9012 0.00798 0.00267 -0.0898 0.3183 1.0000
4.250 0.9286 0.00814 0.00278 -0.0898 0.3052 1.0000
4.500 0.9554 0.00839 0.00291 -0.0897 0.2831 1.0000
4.750 0.9823 0.00863 0.00308 -0.0896 0.2641 1.0000
5.000 1.0087 0.00891 0.00327 -0.0894 0.2447 1.0000
5.250 1.0350 0.00922 0.00348 -0.0892 0.2244 1.0000
5.500 1.0612 0.00953 0.00370 -0.0891 0.2052 1.0000
5.750 1.0871 0.00988 0.00396 -0.0888 0.1830 1.0000
6.000 1.1120 0.01037 0.00427 -0.0885 0.1490 1.0000
6.250 1.1369 0.01086 0.00462 -0.0881 0.1256 1.0000
6.500 1.1624 0.01125 0.00496 -0.0878 0.1121 1.0000
6.750 1.1870 0.01176 0.00534 -0.0874 0.0901 1.0000
7.000 1.2051 0.01325 0.00638 -0.0861 0.0197 1.0000
7.250 1.2293 0.01381 0.00695 -0.0855 0.0153 1.0000
7.500 1.2543 0.01421 0.00739 -0.0851 0.0143 1.0000
7.750 1.2785 0.01469 0.00794 -0.0845 0.0132 1.0000
8.000 1.3018 0.01530 0.00860 -0.0838 0.0121 1.0000
8.250 1.3230 0.01618 0.00960 -0.0828 0.0111 1.0000
8.500 1.3465 0.01667 0.01014 -0.0822 0.0108 1.0000
8.750 1.3694 0.01722 0.01076 -0.0814 0.0102 1.0000
9.000 1.3917 0.01779 0.01139 -0.0807 0.0096 1.0000
9.250 1.4130 0.01845 0.01209 -0.0797 0.0091 1.0000
9.500 1.4327 0.01926 0.01297 -0.0786 0.0086 1.0000
9.750 1.4451 0.02079 0.01463 -0.0763 0.0081 1.0000
10.000 1.4593 0.02200 0.01597 -0.0743 0.0079 1.0000
10.250 1.4777 0.02274 0.01678 -0.0729 0.0077 1.0000
10.500 1.4937 0.02363 0.01776 -0.0712 0.0075 1.0000
10.750 1.5079 0.02462 0.01884 -0.0693 0.0072 1.0000
11.000 1.5196 0.02569 0.02001 -0.0670 0.0070 1.0000
11.250 1.5292 0.02680 0.02120 -0.0644 0.0069 1.0000
11.500 1.5356 0.02792 0.02242 -0.0613 0.0067 1.0000
11.750 1.5401 0.02906 0.02365 -0.0580 0.0065 1.0000
12.000 1.5454 0.03026 0.02493 -0.0552 0.0063 1.0000
12.250 1.5501 0.03160 0.02634 -0.0526 0.0061 1.0000
12.500 1.5521 0.03325 0.02807 -0.0501 0.0060 1.0000
12.750 1.5498 0.03543 0.03036 -0.0477 0.0058 1.0000
13.000 1.5404 0.03854 0.03360 -0.0453 0.0057 1.0000
13.250 1.5251 0.04258 0.03780 -0.0431 0.0056 1.0000
13.500 1.5100 0.04697 0.04235 -0.0415 0.0055 1.0000
13.750 1.5086 0.05000 0.04552 -0.0414 0.0054 1.0000
14.000 1.5058 0.05344 0.04909 -0.0416 0.0054 1.0000
14.250 1.5013 0.05732 0.05312 -0.0424 0.0053 1.0000
14.500 1.4937 0.06188 0.05782 -0.0435 0.0053 1.0000
14.750 1.4856 0.06679 0.06287 -0.0451 0.0053 1.0000
15.000 1.4759 0.07207 0.06829 -0.0471 0.0052 1.0000
15.250 1.4651 0.07773 0.07409 -0.0493 0.0052 1.0000
15.500 1.4533 0.08368 0.08018 -0.0518 0.0052 1.0000
15.750 1.4406 0.08982 0.08646 -0.0544 0.0051 1.0000
16.000 1.4270 0.09616 0.09293 -0.0571 0.0051 1.0000
16.250 1.4124 0.10278 0.09968 -0.0600 0.0051 1.0000
16.500 1.3979 0.10948 0.10650 -0.0630 0.0051 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 276 (DAIMLER VII) AIRFOIL (goe276-il)