EPPLER 637 AIRFOIL (e637-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 637 AIRFOIL (e637-il) Reynolds number: 500,000 Max Cl/Cd: 98.37 at α=8.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e637-il-500000-n5.txt Download as CSV file: xf-e637-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 637 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.3480 0.09538 0.09227 -0.0228 0.7081 0.0101
-9.250 -0.3430 0.09278 0.08960 -0.0233 0.6877 0.0103
-9.000 -0.3450 0.08849 0.08527 -0.0255 0.6733 0.0102
-8.750 -0.3379 0.08662 0.08333 -0.0256 0.6558 0.0105
-8.500 -0.3379 0.08312 0.07982 -0.0272 0.6434 0.0106
-8.250 -0.3391 0.07962 0.07629 -0.0291 0.6320 0.0108
-8.000 -0.3476 0.07524 0.07191 -0.0323 0.6223 0.0106
-7.750 -0.3584 0.07137 0.06801 -0.0339 0.6131 0.0106
-7.500 -0.3688 0.06725 0.06383 -0.0344 0.6057 0.0105
-7.250 -0.3679 0.06413 0.06062 -0.0348 0.5957 0.0107
-7.000 -0.3645 0.06103 0.05743 -0.0349 0.5860 0.0112
-6.750 -0.3686 0.05444 0.05064 -0.0348 0.5804 0.0096
-6.500 -0.3616 0.05128 0.04734 -0.0341 0.5718 0.0094
-6.250 -0.3536 0.04771 0.04359 -0.0331 0.5638 0.0093
-6.000 -0.3438 0.04418 0.03986 -0.0319 0.5559 0.0092
-5.750 -0.3330 0.04006 0.03549 -0.0301 0.5487 0.0094
-5.500 -0.3210 0.03606 0.03119 -0.0280 0.5420 0.0095
-5.250 -0.3068 0.03283 0.02770 -0.0261 0.5344 0.0094
-5.000 -0.2917 0.02975 0.02432 -0.0240 0.5271 0.0092
-4.750 -0.2766 0.02599 0.02016 -0.0214 0.5198 0.0089
-4.250 -0.2398 0.01914 0.01230 -0.0166 0.5059 0.0085
-4.000 -0.2162 0.01736 0.01015 -0.0154 0.4979 0.0086
-3.750 -0.1911 0.01615 0.00868 -0.0146 0.4908 0.0089
-3.500 -0.1655 0.01520 0.00751 -0.0138 0.4834 0.0092
-3.250 -0.1397 0.01448 0.00661 -0.0132 0.4765 0.0096
-3.000 -0.1140 0.01394 0.00593 -0.0126 0.4688 0.0098
-2.750 -0.0887 0.01342 0.00531 -0.0119 0.4620 0.0100
-2.500 -0.0639 0.01292 0.00477 -0.0112 0.4553 0.0105
-2.000 -0.0128 0.01235 0.00409 -0.0101 0.4435 0.0119
-1.750 0.0125 0.01205 0.00373 -0.0094 0.4373 0.0124
-1.500 0.0378 0.01179 0.00340 -0.0088 0.4314 0.0128
-1.250 0.0631 0.01151 0.00309 -0.0081 0.4259 0.0134
-1.000 0.0888 0.01132 0.00286 -0.0076 0.4206 0.0143
-0.750 0.1148 0.01119 0.00267 -0.0071 0.4159 0.0155
-0.500 0.1410 0.01104 0.00250 -0.0066 0.4108 0.0175
-0.250 0.1670 0.01093 0.00237 -0.0062 0.4056 0.0227
0.000 0.1924 0.01076 0.00226 -0.0056 0.4011 0.0468
0.250 0.2175 0.01053 0.00220 -0.0050 0.3969 0.1049
0.500 0.2272 0.00907 0.00207 -0.0020 0.3934 0.5298
0.750 0.2837 0.00829 0.00258 -0.0072 0.3877 0.9401
1.000 0.3272 0.00862 0.00283 -0.0100 0.3831 0.9547
1.250 0.3689 0.00887 0.00299 -0.0127 0.3783 0.9636
1.500 0.4031 0.00898 0.00301 -0.0140 0.3742 0.9660
1.750 0.4331 0.00907 0.00302 -0.0145 0.3704 0.9683
2.000 0.4609 0.00913 0.00304 -0.0145 0.3668 0.9712
2.250 0.4880 0.00921 0.00308 -0.0143 0.3629 0.9739
2.500 0.5202 0.00928 0.00309 -0.0153 0.3592 0.9749
2.750 0.5518 0.00937 0.00313 -0.0162 0.3558 0.9762
3.000 0.5832 0.00943 0.00317 -0.0170 0.3524 0.9777
3.250 0.6135 0.00950 0.00321 -0.0177 0.3487 0.9794
3.500 0.6428 0.00959 0.00328 -0.0181 0.3451 0.9815
3.750 0.6703 0.00971 0.00336 -0.0181 0.3420 0.9838
4.000 0.6997 0.00980 0.00344 -0.0185 0.3389 0.9853
4.250 0.7316 0.00986 0.00351 -0.0195 0.3354 0.9863
4.500 0.7629 0.00994 0.00358 -0.0204 0.3318 0.9875
4.750 0.7934 0.01005 0.00367 -0.0212 0.3286 0.9888
5.000 0.8230 0.01018 0.00378 -0.0217 0.3255 0.9903
5.250 0.8530 0.01025 0.00389 -0.0224 0.3223 0.9918
5.500 0.8825 0.01036 0.00401 -0.0229 0.3188 0.9933
5.750 0.9116 0.01048 0.00415 -0.0233 0.3153 0.9946
6.000 0.9417 0.01061 0.00427 -0.0241 0.3118 0.9956
6.250 0.9722 0.01071 0.00440 -0.0249 0.3085 0.9968
6.500 1.0027 0.01081 0.00455 -0.0257 0.3047 0.9980
6.750 1.0328 0.01095 0.00470 -0.0264 0.3002 0.9992
7.000 1.0608 0.01112 0.00487 -0.0268 0.2961 1.0000
7.250 1.0841 0.01124 0.00505 -0.0260 0.2915 1.0000
7.500 1.1067 0.01140 0.00524 -0.0252 0.2863 1.0000
7.750 1.1287 0.01160 0.00544 -0.0243 0.2813 1.0000
8.000 1.1514 0.01175 0.00566 -0.0235 0.2765 1.0000
8.250 1.1734 0.01195 0.00588 -0.0226 0.2712 1.0000
8.500 1.1949 0.01217 0.00613 -0.0216 0.2663 1.0000
8.750 1.2168 0.01237 0.00638 -0.0207 0.2600 1.0000
9.000 1.2372 0.01265 0.00666 -0.0195 0.2534 1.0000
9.250 1.2584 0.01288 0.00694 -0.0185 0.2455 1.0000
9.500 1.2782 0.01319 0.00726 -0.0173 0.2376 1.0000
9.750 1.2979 0.01350 0.00760 -0.0161 0.2286 1.0000
10.000 1.3167 0.01385 0.00797 -0.0148 0.2189 1.0000
10.250 1.3337 0.01430 0.00840 -0.0132 0.2060 1.0000
10.500 1.3482 0.01488 0.00893 -0.0113 0.1888 1.0000
10.750 1.3605 0.01556 0.00955 -0.0092 0.1714 1.0000
11.000 1.3719 0.01626 0.01020 -0.0068 0.1554 1.0000
11.250 1.3791 0.01712 0.01099 -0.0040 0.1374 1.0000
11.500 1.3839 0.01803 0.01184 -0.0008 0.1223 1.0000
11.750 1.3838 0.01906 0.01281 0.0031 0.1073 1.0000
12.000 1.3786 0.01997 0.01370 0.0078 0.0975 1.0000
12.250 1.3751 0.02104 0.01478 0.0117 0.0881 1.0000
12.500 1.3730 0.02235 0.01612 0.0145 0.0811 1.0000
12.750 1.3700 0.02408 0.01788 0.0165 0.0746 1.0000
13.000 1.3691 0.02603 0.01988 0.0177 0.0686 1.0000
13.250 1.3657 0.02849 0.02238 0.0183 0.0630 1.0000
13.500 1.3620 0.03121 0.02516 0.0185 0.0575 1.0000
13.750 1.3570 0.03422 0.02822 0.0184 0.0535 1.0000
14.000 1.3521 0.03730 0.03139 0.0182 0.0498 1.0000
14.500 1.3364 0.04430 0.03853 0.0175 0.0438 1.0000
14.750 1.3258 0.04817 0.04246 0.0170 0.0403 1.0000
15.000 1.3165 0.05195 0.04632 0.0164 0.0387 1.0000
15.250 1.3060 0.05591 0.05035 0.0157 0.0359 1.0000
15.500 1.2948 0.06008 0.05459 0.0149 0.0330 1.0000
15.750 1.2842 0.06431 0.05887 0.0140 0.0304 1.0000
16.000 1.2776 0.06812 0.06277 0.0131 0.0290 1.0000
16.250 1.2657 0.07271 0.06740 0.0119 0.0260 1.0000
16.500 1.2573 0.07689 0.07164 0.0107 0.0237 1.0000
16.750 1.2493 0.08108 0.07590 0.0096 0.0218 1.0000
17.000 1.2415 0.08532 0.08019 0.0084 0.0198 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 637 AIRFOIL (e637-il)