GOE 257 AIRFOIL (goe257-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 257 AIRFOIL (goe257-il) Reynolds number: 100,000 Max Cl/Cd: 56.76 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe257-il-100000.txt Download as CSV file: xf-goe257-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 257 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4215 0.11825 0.11338 -0.0138 1.0000 0.0545
-9.000 -0.4255 0.11662 0.11186 -0.0183 1.0000 0.0549
-8.750 -0.4290 0.11450 0.10981 -0.0228 1.0000 0.0551
-8.500 -0.4022 0.10553 0.10077 -0.0155 1.0000 0.0578
-8.250 -0.3951 0.10211 0.09737 -0.0159 1.0000 0.0598
-8.000 -0.3910 0.09897 0.09427 -0.0171 1.0000 0.0620
-7.750 -0.3895 0.09601 0.09137 -0.0188 1.0000 0.0643
-7.500 -0.3011 0.08120 0.07694 -0.0272 1.0000 0.0720
-7.250 -0.3001 0.07809 0.07389 -0.0268 1.0000 0.0752
-7.000 -0.3049 0.07521 0.07107 -0.0274 1.0000 0.0779
-6.750 -0.3108 0.07237 0.06831 -0.0305 1.0000 0.0805
-6.500 -0.3167 0.07008 0.06598 -0.0375 1.0000 0.0821
-6.250 -0.3176 0.06542 0.06141 -0.0363 1.0000 0.0834
-6.000 -0.3170 0.06238 0.05848 -0.0309 1.0000 0.0859
-5.750 -0.3229 0.06051 0.05670 -0.0285 1.0000 0.0879
-5.500 -0.3362 0.05970 0.05596 -0.0260 1.0000 0.0894
-5.250 -0.2902 0.06092 0.05624 -0.0455 1.0000 0.0989
-5.000 -0.2851 0.05847 0.05389 -0.0431 1.0000 0.1026
-4.750 -0.2686 0.05642 0.05153 -0.0466 1.0000 0.1123
-4.500 -0.2752 0.05446 0.04977 -0.0428 1.0000 0.1141
-4.250 -0.2437 0.05114 0.04630 -0.0474 0.9953 0.1287
-4.000 -0.2010 0.04725 0.04228 -0.0530 0.9876 0.1465
-3.750 -0.1547 0.04368 0.03847 -0.0596 0.9805 0.1733
-3.500 -0.1140 0.04050 0.03519 -0.0642 0.9726 0.2038
-3.250 -0.0769 0.03770 0.03236 -0.0677 0.9649 0.2493
-2.250 0.1361 0.02615 0.01850 -0.0875 0.9338 0.1508
-2.000 0.1837 0.02401 0.01573 -0.0896 0.9261 0.1194
-1.750 0.2208 0.02267 0.01397 -0.0903 0.9138 0.1063
-1.500 0.2556 0.02163 0.01254 -0.0907 0.9015 0.0987
-1.250 0.2874 0.02092 0.01164 -0.0907 0.8889 0.0961
-1.000 0.3171 0.02000 0.01065 -0.0904 0.8764 0.0952
-0.750 0.3455 0.01933 0.00991 -0.0899 0.8640 0.0956
-0.500 0.3726 0.01862 0.00923 -0.0892 0.8521 0.0992
-0.250 0.3986 0.01811 0.00875 -0.0884 0.8391 0.1023
0.000 0.4249 0.01778 0.00838 -0.0876 0.8260 0.1050
0.250 0.4514 0.01756 0.00806 -0.0869 0.8131 0.1095
0.500 0.4784 0.01735 0.00778 -0.0862 0.8003 0.1190
0.750 0.5042 0.01590 0.00773 -0.0861 0.7884 0.5785
1.000 0.5309 0.01512 0.00744 -0.0844 0.7766 1.0000
1.250 0.5575 0.01535 0.00740 -0.0837 0.7647 1.0000
1.500 0.5839 0.01561 0.00747 -0.0831 0.7513 1.0000
1.750 0.6101 0.01590 0.00759 -0.0824 0.7371 1.0000
2.000 0.6364 0.01620 0.00776 -0.0819 0.7231 1.0000
2.250 0.6626 0.01651 0.00795 -0.0814 0.7090 1.0000
2.500 0.6890 0.01683 0.00820 -0.0809 0.6953 1.0000
2.750 0.7154 0.01715 0.00845 -0.0805 0.6819 1.0000
3.000 0.7417 0.01744 0.00868 -0.0801 0.6685 1.0000
3.250 0.7682 0.01772 0.00892 -0.0796 0.6558 1.0000
3.500 0.7950 0.01801 0.00919 -0.0793 0.6448 1.0000
3.750 0.8219 0.01828 0.00943 -0.0790 0.6339 1.0000
4.000 0.8485 0.01859 0.00980 -0.0788 0.6219 1.0000
4.250 0.8752 0.01888 0.01016 -0.0786 0.6105 1.0000
4.500 0.9019 0.01911 0.01042 -0.0783 0.5994 1.0000
4.750 0.9288 0.01928 0.01057 -0.0778 0.5885 1.0000
5.000 0.9551 0.01947 0.01084 -0.0774 0.5753 1.0000
5.250 0.9813 0.01959 0.01105 -0.0769 0.5611 1.0000
5.500 1.0074 0.01966 0.01116 -0.0763 0.5461 1.0000
5.750 1.0332 0.01973 0.01128 -0.0757 0.5295 1.0000
6.000 1.0585 0.01983 0.01149 -0.0751 0.5105 1.0000
6.250 1.0838 0.01993 0.01165 -0.0743 0.4910 1.0000
6.500 1.1081 0.02003 0.01176 -0.0734 0.4669 1.0000
6.750 1.1314 0.02018 0.01188 -0.0724 0.4395 1.0000
7.000 1.1537 0.02043 0.01213 -0.0713 0.4099 1.0000
7.250 1.1761 0.02078 0.01256 -0.0704 0.3819 1.0000
7.500 1.1977 0.02110 0.01299 -0.0694 0.3492 1.0000
7.750 1.2171 0.02154 0.01339 -0.0681 0.3016 1.0000
8.000 1.2245 0.02343 0.01459 -0.0655 0.1963 1.0000
8.250 1.2192 0.02729 0.01728 -0.0617 0.0749 1.0000
8.500 1.2255 0.02952 0.01961 -0.0588 0.0646 1.0000
8.750 1.2259 0.03198 0.02211 -0.0553 0.0595 1.0000
9.000 1.2345 0.03361 0.02395 -0.0527 0.0553 1.0000
9.250 1.2400 0.03547 0.02586 -0.0499 0.0515 1.0000
9.500 1.2454 0.03760 0.02796 -0.0471 0.0490 1.0000
9.750 1.2700 0.04045 0.03066 -0.0465 0.0468 1.0000
10.000 1.2958 0.04253 0.03289 -0.0458 0.0457 1.0000
10.250 1.3219 0.04499 0.03555 -0.0454 0.0446 1.0000
10.500 1.3408 0.04741 0.03826 -0.0442 0.0429 1.0000
10.750 1.3570 0.04999 0.04108 -0.0429 0.0413 1.0000
11.000 1.3709 0.05310 0.04450 -0.0415 0.0409 1.0000
11.250 1.3777 0.05647 0.04826 -0.0393 0.0412 1.0000
11.500 1.3773 0.05993 0.05211 -0.0365 0.0417 1.0000
11.750 1.3698 0.06333 0.05587 -0.0333 0.0423 1.0000
12.000 1.3584 0.06690 0.05977 -0.0302 0.0429 1.0000
12.250 1.3451 0.07081 0.06397 -0.0278 0.0435 1.0000
12.500 1.3298 0.07506 0.06850 -0.0262 0.0441 1.0000
12.750 1.3129 0.07971 0.07340 -0.0254 0.0447 1.0000
13.000 1.2949 0.08480 0.07873 -0.0254 0.0452 1.0000
13.250 1.2763 0.09040 0.08453 -0.0261 0.0458 1.0000
13.500 1.1086 0.08807 0.08251 -0.0137 0.0446 1.0000
13.750 1.0793 0.09345 0.08808 -0.0148 0.0450 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 257 AIRFOIL (goe257-il)