GOE 275 (DAIMLER VI) AIRFOIL (goe275-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 275 (DAIMLER VI) AIRFOIL (goe275-il) Reynolds number: 500,000 Max Cl/Cd: 99.72 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe275-il-500000.txt Download as CSV file: xf-goe275-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 275 (DAIMLER VI) 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3323 0.09046 0.08831 -0.0197 1.0000 0.0178
-7.500 -0.3425 0.08898 0.08690 -0.0172 1.0000 0.0183
-7.250 -0.3496 0.08725 0.08520 -0.0161 0.9991 0.0185
-7.000 -0.3145 0.08133 0.07926 -0.0292 0.9942 0.0195
-6.750 -0.2754 0.07439 0.07224 -0.0462 0.9879 0.0197
-6.500 -0.2519 0.06923 0.06706 -0.0511 0.9853 0.0198
-6.250 -0.2329 0.06594 0.06378 -0.0525 0.9822 0.0201
-6.000 -0.2094 0.06244 0.06026 -0.0564 0.9774 0.0205
-5.750 -0.1773 0.05831 0.05608 -0.0629 0.9741 0.0211
-5.500 -0.1456 0.05414 0.05183 -0.0691 0.9696 0.0220
-5.250 -0.0884 0.04718 0.04446 -0.0815 0.9632 0.0236
-5.000 -0.0624 0.04261 0.03988 -0.0850 0.9604 0.0239
-4.750 -0.0383 0.03998 0.03722 -0.0866 0.9538 0.0242
-4.500 -0.0078 0.03725 0.03441 -0.0895 0.9492 0.0249
-4.250 0.0294 0.03416 0.03117 -0.0933 0.9460 0.0260
-4.000 0.0719 0.03130 0.02767 -0.0952 0.9386 0.0283
-3.750 0.1004 0.02711 0.02347 -0.0978 0.9335 0.0289
-3.500 0.1300 0.02518 0.02149 -0.0995 0.9265 0.0294
-3.250 0.1626 0.02337 0.01954 -0.1012 0.9190 0.0304
-3.000 0.1939 0.02170 0.01765 -0.1019 0.9098 0.0323
-2.750 0.2281 0.01981 0.01524 -0.1024 0.9005 0.0347
-2.500 0.2540 0.01828 0.01366 -0.1025 0.8874 0.0355
-2.250 0.2803 0.01721 0.01246 -0.1023 0.8717 0.0367
-2.000 0.3070 0.01644 0.01147 -0.1017 0.8547 0.0392
-1.750 0.3331 0.01557 0.01022 -0.1007 0.8369 0.0422
-1.500 0.3574 0.01447 0.00908 -0.1001 0.8170 0.0436
-1.250 0.3824 0.01390 0.00837 -0.0993 0.7929 0.0464
-1.000 0.4071 0.01340 0.00755 -0.0982 0.7621 0.0520
-0.750 0.4301 0.01277 0.00678 -0.0972 0.7224 0.0547
-0.500 0.4536 0.01373 0.00741 -0.0956 0.6766 0.0610
-0.250 0.4748 0.01216 0.00568 -0.0946 0.6391 0.0683
0.000 0.4974 0.01183 0.00517 -0.0936 0.6039 0.0809
1.250 0.6192 0.01138 0.00364 -0.0877 0.4372 0.0503
1.500 0.6427 0.01093 0.00312 -0.0867 0.4146 0.0467
1.750 0.6672 0.01086 0.00296 -0.0859 0.3961 0.0453
2.000 0.6919 0.01085 0.00286 -0.0852 0.3808 0.0450
2.250 0.7170 0.01087 0.00281 -0.0845 0.3687 0.0453
2.500 0.7423 0.01091 0.00281 -0.0839 0.3600 0.0469
2.750 0.7681 0.01093 0.00282 -0.0834 0.3527 0.0480
3.000 0.7937 0.01100 0.00285 -0.0829 0.3460 0.0485
3.250 0.8193 0.01109 0.00290 -0.0824 0.3398 0.0498
3.500 0.8454 0.01116 0.00295 -0.0819 0.3341 0.0533
3.750 0.8781 0.00962 0.00330 -0.0835 0.3276 1.0000
4.000 0.9038 0.00975 0.00340 -0.0830 0.3216 1.0000
4.250 0.9291 0.00992 0.00352 -0.0824 0.3151 1.0000
4.500 0.9542 0.01012 0.00367 -0.0819 0.3092 1.0000
4.750 0.9799 0.01025 0.00381 -0.0814 0.3032 1.0000
5.000 1.0047 0.01046 0.00398 -0.0808 0.2966 1.0000
5.250 1.0303 0.01059 0.00413 -0.0803 0.2901 1.0000
5.500 1.0551 0.01079 0.00429 -0.0797 0.2812 1.0000
5.750 1.0805 0.01093 0.00446 -0.0792 0.2717 1.0000
6.000 1.1053 0.01112 0.00463 -0.0786 0.2608 1.0000
6.250 1.1298 0.01133 0.00481 -0.0780 0.2473 1.0000
6.500 1.1539 0.01158 0.00502 -0.0773 0.2300 1.0000
7.000 1.1990 0.01235 0.00559 -0.0756 0.1904 1.0000
7.250 1.2210 0.01279 0.00596 -0.0746 0.1775 1.0000
7.500 1.2427 0.01324 0.00638 -0.0736 0.1666 1.0000
7.750 1.2647 0.01366 0.00678 -0.0727 0.1545 1.0000
8.000 1.2862 0.01412 0.00720 -0.0717 0.1387 1.0000
8.250 1.3075 0.01461 0.00760 -0.0707 0.1172 1.0000
8.750 1.3465 0.01583 0.00866 -0.0682 0.0921 1.0000
9.000 1.3649 0.01648 0.00929 -0.0667 0.0864 1.0000
9.250 1.3853 0.01693 0.00981 -0.0656 0.0838 1.0000
9.500 1.4044 0.01745 0.01041 -0.0642 0.0810 1.0000
9.750 1.4217 0.01808 0.01108 -0.0627 0.0781 1.0000
10.000 1.4361 0.01888 0.01190 -0.0606 0.0745 1.0000
10.250 1.4558 0.01926 0.01237 -0.0594 0.0721 1.0000
10.500 1.4749 0.01963 0.01284 -0.0582 0.0692 1.0000
10.750 1.4909 0.02013 0.01338 -0.0564 0.0660 1.0000
11.000 1.5014 0.02092 0.01420 -0.0538 0.0622 1.0000
11.250 1.5237 0.02103 0.01441 -0.0530 0.0580 1.0000
11.500 1.5399 0.02151 0.01487 -0.0514 0.0502 1.0000
11.750 1.5488 0.02245 0.01564 -0.0489 0.0321 1.0000
12.000 1.5493 0.02390 0.01701 -0.0452 0.0227 1.0000
12.250 1.5541 0.02508 0.01826 -0.0424 0.0203 1.0000
12.500 1.5556 0.02653 0.01978 -0.0393 0.0185 1.0000
12.750 1.5580 0.02798 0.02135 -0.0366 0.0175 1.0000
13.000 1.5607 0.02947 0.02296 -0.0342 0.0169 1.0000
13.250 1.5610 0.03122 0.02484 -0.0320 0.0162 1.0000
13.500 1.5585 0.03332 0.02705 -0.0299 0.0156 1.0000
13.750 1.5533 0.03582 0.02968 -0.0281 0.0151 1.0000
14.000 1.5445 0.03889 0.03287 -0.0267 0.0147 1.0000
14.250 1.5332 0.04254 0.03667 -0.0260 0.0144 1.0000
14.500 1.5219 0.04660 0.04087 -0.0260 0.0141 1.0000
14.750 1.5160 0.05036 0.04478 -0.0265 0.0139 1.0000
15.000 1.5078 0.05467 0.04923 -0.0275 0.0137 1.0000
15.250 1.4986 0.05933 0.05404 -0.0289 0.0134 1.0000
15.500 1.4867 0.06462 0.05947 -0.0307 0.0133 1.0000
15.750 1.4734 0.07028 0.06528 -0.0329 0.0132 1.0000
16.000 1.4613 0.07594 0.07107 -0.0351 0.0130 1.0000
16.250 1.4476 0.08196 0.07721 -0.0376 0.0129 1.0000
16.500 1.4342 0.08806 0.08344 -0.0402 0.0127 1.0000
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Polar data table (+)
Polar graphs
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