GOE 591 AIRFOIL (goe591-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 591 AIRFOIL (goe591-il) Reynolds number: 100,000 Max Cl/Cd: 52.75 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe591-il-100000.txt Download as CSV file: xf-goe591-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 591 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
------ -------- --------- --------- -------- -------- --------
-7.500 -0.3588 0.09947 0.09541 -0.0218 1.0000 0.0945
-7.250 -0.3740 0.09845 0.09447 -0.0190 1.0000 0.0957
-7.000 -0.3922 0.09754 0.09365 -0.0161 1.0000 0.0968
-6.750 -0.3833 0.09408 0.09019 -0.0229 0.9939 0.1013
-6.500 -0.3626 0.08771 0.08373 -0.0408 0.9815 0.1052
-6.250 -0.3416 0.08415 0.08020 -0.0375 0.9767 0.1081
-6.000 -0.3134 0.08006 0.07603 -0.0448 0.9688 0.1165
-5.750 -0.2930 0.07495 0.07085 -0.0522 0.9590 0.1220
-5.500 -0.2699 0.07179 0.06766 -0.0539 0.9514 0.1282
-5.250 -0.2401 0.06666 0.06235 -0.0630 0.9428 0.1373
-5.000 -0.2199 0.06392 0.05958 -0.0637 0.9336 0.1441
-4.750 -0.1864 0.05987 0.05540 -0.0693 0.9274 0.1566
-4.500 -0.1657 0.05675 0.05214 -0.0720 0.9168 0.1694
-4.250 -0.1275 0.05314 0.04835 -0.0770 0.9117 0.1851
-4.000 -0.1075 0.05064 0.04572 -0.0782 0.9008 0.1996
-3.750 -0.0688 0.04760 0.04260 -0.0816 0.8966 0.2188
-3.500 -0.0494 0.04577 0.04059 -0.0821 0.8853 0.2454
-3.250 0.0191 0.03301 0.02551 -0.0909 0.8825 0.1015
-3.000 0.0450 0.03116 0.02329 -0.0905 0.8727 0.1012
-2.750 0.0904 0.02914 0.02071 -0.0932 0.8682 0.1033
-2.500 0.1204 0.02786 0.01906 -0.0931 0.8591 0.1042
-2.250 0.1645 0.02593 0.01703 -0.0958 0.8543 0.1082
-2.000 0.2145 0.02461 0.01550 -0.0990 0.8511 0.1185
-1.750 0.2385 0.02382 0.01479 -0.0981 0.8400 0.1267
-1.500 0.2844 0.02246 0.01342 -0.1005 0.8361 0.1502
-1.250 0.3073 0.02189 0.01308 -0.0993 0.8248 0.1875
-1.000 0.3492 0.02056 0.01208 -0.1012 0.8205 0.2663
-0.750 0.4343 0.01784 0.01138 -0.1114 0.8152 1.0000
-0.500 0.4688 0.01766 0.01089 -0.1120 0.8068 1.0000
-0.250 0.4920 0.01777 0.01080 -0.1108 0.7951 1.0000
0.000 0.5247 0.01765 0.01047 -0.1111 0.7866 1.0000
0.250 0.5515 0.01766 0.01031 -0.1104 0.7760 1.0000
0.500 0.5741 0.01783 0.01034 -0.1091 0.7642 1.0000
0.750 0.6033 0.01781 0.01017 -0.1089 0.7550 1.0000
1.000 0.6306 0.01783 0.01006 -0.1084 0.7447 1.0000
1.250 0.6525 0.01805 0.01018 -0.1070 0.7329 1.0000
1.500 0.6792 0.01813 0.01014 -0.1064 0.7230 1.0000
1.750 0.7080 0.01812 0.01001 -0.1061 0.7135 1.0000
2.000 0.7293 0.01837 0.01021 -0.1047 0.7017 1.0000
2.250 0.7547 0.01850 0.01026 -0.1040 0.6915 1.0000
2.500 0.7846 0.01848 0.01012 -0.1038 0.6825 1.0000
2.750 0.8057 0.01876 0.01036 -0.1025 0.6707 1.0000
3.000 0.8300 0.01893 0.01050 -0.1016 0.6601 1.0000
3.250 0.8609 0.01890 0.01034 -0.1016 0.6513 1.0000
3.500 0.8814 0.01922 0.01066 -0.1002 0.6396 1.0000
3.750 0.9048 0.01948 0.01091 -0.0993 0.6292 1.0000
4.000 0.9353 0.01953 0.01085 -0.0993 0.6208 1.0000
4.250 0.9551 0.01993 0.01129 -0.0979 0.6095 1.0000
4.500 0.9788 0.02024 0.01161 -0.0970 0.5996 1.0000
4.750 1.0077 0.02037 0.01166 -0.0969 0.5907 1.0000
5.000 1.0274 0.02081 0.01218 -0.0955 0.5798 1.0000
5.250 1.0519 0.02114 0.01250 -0.0948 0.5703 1.0000
5.500 1.0789 0.02136 0.01271 -0.0944 0.5608 1.0000
5.750 1.0987 0.02183 0.01325 -0.0930 0.5498 1.0000
6.000 1.1228 0.02216 0.01359 -0.0923 0.5396 1.0000
6.250 1.1512 0.02231 0.01369 -0.0921 0.5293 1.0000
6.500 1.1704 0.02266 0.01412 -0.0904 0.5165 1.0000
6.750 1.1911 0.02296 0.01446 -0.0891 0.5037 1.0000
7.000 1.2130 0.02326 0.01478 -0.0879 0.4914 1.0000
7.250 1.2374 0.02357 0.01510 -0.0871 0.4802 1.0000
7.500 1.2612 0.02391 0.01546 -0.0863 0.4690 1.0000
7.750 1.2787 0.02442 0.01609 -0.0846 0.4568 1.0000
8.000 1.2986 0.02490 0.01664 -0.0832 0.4449 1.0000
8.250 1.3207 0.02532 0.01711 -0.0822 0.4334 1.0000
8.500 1.3462 0.02567 0.01743 -0.0817 0.4220 1.0000
8.750 1.3612 0.02616 0.01804 -0.0796 0.4090 1.0000
9.000 1.3767 0.02652 0.01848 -0.0774 0.3950 1.0000
9.250 1.3910 0.02678 0.01881 -0.0751 0.3803 1.0000
9.500 1.4039 0.02710 0.01919 -0.0726 0.3659 1.0000
9.750 1.4155 0.02750 0.01967 -0.0700 0.3521 1.0000
10.000 1.4249 0.02790 0.02013 -0.0670 0.3381 1.0000
10.250 1.4312 0.02832 0.02061 -0.0636 0.3239 1.0000
10.500 1.4349 0.02882 0.02118 -0.0599 0.3106 1.0000
10.750 1.4365 0.02942 0.02182 -0.0560 0.2976 1.0000
11.000 1.4364 0.03015 0.02259 -0.0521 0.2843 1.0000
11.250 1.4361 0.03104 0.02352 -0.0486 0.2711 1.0000
11.500 1.4337 0.03214 0.02467 -0.0451 0.2576 1.0000
11.750 1.4285 0.03351 0.02610 -0.0418 0.2430 1.0000
12.000 1.4240 0.03510 0.02776 -0.0389 0.2290 1.0000
12.250 1.4159 0.03711 0.02988 -0.0362 0.2123 1.0000
12.500 1.4064 0.03954 0.03241 -0.0340 0.1937 1.0000
12.750 1.3926 0.04261 0.03548 -0.0321 0.1719 1.0000
13.000 1.3750 0.04645 0.03925 -0.0307 0.1461 1.0000
13.250 1.3538 0.05108 0.04375 -0.0298 0.1215 1.0000
13.500 1.3317 0.05626 0.04882 -0.0295 0.0983 1.0000
13.750 1.3117 0.06158 0.05405 -0.0296 0.0815 1.0000
14.000 1.2947 0.06683 0.05925 -0.0301 0.0728 1.0000
14.250 1.2804 0.07195 0.06439 -0.0308 0.0675 1.0000
14.500 1.2658 0.07727 0.06972 -0.0318 0.0635 1.0000
14.750 1.2538 0.08231 0.07476 -0.0327 0.0605 1.0000
15.000 1.2475 0.08658 0.07914 -0.0334 0.0570 1.0000
15.250 1.2427 0.09059 0.08320 -0.0340 0.0543 1.0000
15.500 1.2426 0.09343 0.08592 -0.0336 0.0510 1.0000
15.750 1.2470 0.09613 0.08878 -0.0335 0.0487 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 591 AIRFOIL (goe591-il)