GOE 177 AIRFOIL (goe177-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 177 AIRFOIL (goe177-il) Reynolds number: 500,000 Max Cl/Cd: 79.78 at α=2.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe177-il-500000-n5.txt Download as CSV file: xf-goe177-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 177 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3227 0.10203 0.09990 -0.0184 1.0000 0.0066
-8.250 -0.3167 0.09873 0.09663 -0.0198 1.0000 0.0066
-8.000 -0.3159 0.09447 0.09241 -0.0209 1.0000 0.0068
-7.750 -0.3082 0.09235 0.09031 -0.0210 1.0000 0.0073
-7.500 -0.3024 0.09007 0.08805 -0.0216 1.0000 0.0078
-7.250 -0.2985 0.08766 0.08569 -0.0222 0.9990 0.0085
-7.000 -0.2778 0.08379 0.08181 -0.0284 0.9875 0.0099
-6.750 -0.2514 0.08015 0.07817 -0.0355 0.9755 0.0106
-6.500 -0.2292 0.07584 0.07384 -0.0416 0.9600 0.0107
-6.000 -0.1965 0.06336 0.06126 -0.0538 0.9153 0.0062
-5.750 -0.1759 0.05902 0.05677 -0.0587 0.8722 0.0059
-5.500 -0.1529 0.05448 0.05201 -0.0638 0.8270 0.0057
-5.250 -0.1252 0.04932 0.04662 -0.0699 0.7933 0.0055
-4.750 -0.0392 0.01953 0.01532 -0.0937 0.7607 0.0049
-4.500 -0.0109 0.01633 0.01142 -0.0946 0.7394 0.0052
-4.250 0.0171 0.01457 0.00916 -0.0948 0.7188 0.0056
-4.000 0.0451 0.01343 0.00761 -0.0948 0.7006 0.0060
-3.750 0.0729 0.01274 0.00663 -0.0948 0.6831 0.0067
-3.500 0.1008 0.01176 0.00538 -0.0949 0.6661 0.0075
-3.250 0.1288 0.01121 0.00462 -0.0948 0.6490 0.0081
-3.000 0.1569 0.01076 0.00398 -0.0947 0.6337 0.0091
-2.750 0.1851 0.01041 0.00344 -0.0946 0.6180 0.0102
-2.500 0.2132 0.01004 0.00287 -0.0946 0.6011 0.0134
-2.250 0.2411 0.00981 0.00249 -0.0944 0.5818 0.0182
-2.000 0.2690 0.00961 0.00220 -0.0944 0.5635 0.0325
-1.750 0.2967 0.00953 0.00216 -0.0944 0.5458 0.0590
-1.500 0.3244 0.00961 0.00212 -0.0943 0.5293 0.0696
-1.250 0.3522 0.00968 0.00210 -0.0942 0.5148 0.0787
-1.000 0.3801 0.00972 0.00202 -0.0942 0.5016 0.0810
-0.750 0.4080 0.00965 0.00187 -0.0942 0.4889 0.0841
-0.500 0.4356 0.00967 0.00178 -0.0941 0.4697 0.0872
-0.250 0.4631 0.00973 0.00172 -0.0941 0.4492 0.0908
0.000 0.4906 0.00980 0.00167 -0.0940 0.4277 0.0930
0.250 0.5180 0.00989 0.00164 -0.0939 0.4086 0.0949
0.500 0.5453 0.00998 0.00163 -0.0939 0.3851 0.0996
0.750 0.5724 0.01011 0.00165 -0.0938 0.3617 0.1070
1.000 0.5998 0.01020 0.00169 -0.0938 0.3447 0.1227
1.250 0.6272 0.01019 0.00176 -0.0939 0.3312 0.1783
1.500 0.6546 0.01019 0.00187 -0.0939 0.3178 0.2528
1.750 0.6818 0.01024 0.00199 -0.0940 0.3029 0.3194
2.000 0.7089 0.01029 0.00211 -0.0940 0.2859 0.3874
2.500 0.7555 0.00947 0.00237 -0.0923 0.2434 1.0000
2.750 0.7794 0.01021 0.00270 -0.0920 0.1684 1.0000
3.000 0.8035 0.01093 0.00311 -0.0916 0.1100 1.0000
3.250 0.8302 0.01117 0.00331 -0.0915 0.1011 1.0000
3.500 0.8570 0.01139 0.00349 -0.0914 0.0935 1.0000
3.750 0.8821 0.01190 0.00383 -0.0911 0.0577 1.0000
4.000 0.9086 0.01215 0.00407 -0.0909 0.0525 1.0000
4.250 0.9352 0.01237 0.00431 -0.0908 0.0496 1.0000
4.500 0.9617 0.01260 0.00456 -0.0907 0.0461 1.0000
4.750 0.9876 0.01291 0.00485 -0.0905 0.0398 1.0000
5.000 1.0124 0.01339 0.00515 -0.0901 0.0176 1.0000
5.250 1.0387 0.01362 0.00548 -0.0899 0.0124 1.0000
5.500 1.0642 0.01398 0.00595 -0.0896 0.0101 1.0000
5.750 1.0889 0.01447 0.00657 -0.0891 0.0085 1.0000
6.000 1.1133 0.01496 0.00714 -0.0886 0.0072 1.0000
6.250 1.1376 0.01544 0.00772 -0.0881 0.0062 1.0000
6.500 1.1610 0.01603 0.00839 -0.0875 0.0056 1.0000
6.750 1.1818 0.01695 0.00943 -0.0865 0.0051 1.0000
7.000 1.2034 0.01771 0.01028 -0.0856 0.0049 1.0000
7.250 1.2251 0.01841 0.01106 -0.0848 0.0043 1.0000
7.500 1.2460 0.01913 0.01184 -0.0840 0.0038 1.0000
7.750 1.2644 0.02010 0.01292 -0.0828 0.0035 1.0000
8.000 1.2779 0.02151 0.01443 -0.0809 0.0033 1.0000
8.250 1.2917 0.02278 0.01582 -0.0790 0.0032 1.0000
8.500 1.3037 0.02415 0.01730 -0.0768 0.0030 1.0000
8.750 1.3133 0.02566 0.01891 -0.0744 0.0029 1.0000
9.000 1.3205 0.02724 0.02060 -0.0716 0.0028 1.0000
9.250 1.3251 0.02885 0.02231 -0.0685 0.0028 1.0000
9.500 1.3300 0.03060 0.02416 -0.0656 0.0028 1.0000
9.750 1.3359 0.03246 0.02613 -0.0631 0.0027 1.0000
10.000 1.3427 0.03443 0.02822 -0.0608 0.0027 1.0000
10.250 1.3499 0.03645 0.03038 -0.0587 0.0027 1.0000
10.500 1.3564 0.03844 0.03252 -0.0568 0.0025 1.0000
10.750 1.3616 0.04047 0.03468 -0.0552 0.0024 1.0000
11.000 1.3656 0.04259 0.03692 -0.0538 0.0022 1.0000
11.250 1.3686 0.04480 0.03924 -0.0527 0.0021 1.0000
11.500 1.3700 0.04729 0.04185 -0.0519 0.0021 1.0000
11.750 1.3693 0.05021 0.04490 -0.0511 0.0020 1.0000
12.000 1.3658 0.05383 0.04868 -0.0501 0.0019 1.0000
12.250 1.3611 0.05756 0.05267 -0.0496 0.0019 1.0000
12.500 1.3549 0.06149 0.05683 -0.0496 0.0019 1.0000
12.750 1.3470 0.06581 0.06136 -0.0500 0.0019 1.0000
13.000 1.3373 0.07048 0.06623 -0.0507 0.0018 1.0000
13.250 1.3263 0.07549 0.07144 -0.0519 0.0018 1.0000
13.500 1.3145 0.08072 0.07685 -0.0534 0.0018 1.0000
13.750 1.3010 0.08644 0.08275 -0.0554 0.0018 1.0000
14.000 1.2873 0.09246 0.08895 -0.0578 0.0018 1.0000
14.250 1.2732 0.09878 0.09544 -0.0606 0.0018 1.0000
14.500 1.2595 0.10530 0.10210 -0.0637 0.0018 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 177 AIRFOIL (goe177-il)