FX 66-H-60 (fx66h60-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: FX 66-H-60 (fx66h60-il) Reynolds number: 500,000 Max Cl/Cd: 67.84 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx66h60-il-500000-n5.txt Download as CSV file: xf-fx66h60-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: FX 66-H-60
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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.6103 0.09195 0.08985 0.0275 1.0000 0.0038
-7.750 -0.6108 0.08797 0.08591 0.0240 1.0000 0.0042
-7.500 -0.6140 0.08395 0.08190 0.0214 1.0000 0.0042
-6.250 -0.5236 0.05620 0.05417 0.0105 1.0000 0.0093
-6.000 -0.5144 0.05149 0.04938 0.0094 1.0000 0.0093
-5.750 -0.5037 0.04681 0.04460 0.0087 1.0000 0.0093
-5.500 -0.4911 0.04221 0.03989 0.0083 1.0000 0.0093
-4.500 -0.4549 0.03635 0.03307 0.0096 1.0000 0.0076
-4.250 -0.4341 0.03241 0.02888 0.0107 1.0000 0.0068
-4.000 -0.3905 0.02731 0.02293 0.0079 0.7929 0.0063
-3.750 -0.3694 0.02419 0.01926 0.0100 0.7302 0.0064
-3.500 -0.3461 0.02144 0.01600 0.0118 0.6824 0.0070
-3.250 -0.3216 0.01926 0.01332 0.0132 0.6405 0.0077
-3.000 -0.2956 0.01857 0.01233 0.0140 0.6009 0.0090
-2.500 -0.2443 0.01560 0.00865 0.0157 0.5413 0.0091
-2.250 -0.2184 0.01410 0.00685 0.0165 0.5158 0.0089
-2.000 -0.1925 0.01296 0.00548 0.0173 0.4927 0.0087
-1.750 -0.1670 0.01206 0.00439 0.0180 0.4731 0.0086
-1.500 -0.1417 0.01134 0.00351 0.0187 0.4545 0.0086
-1.250 -0.1164 0.01077 0.00276 0.0194 0.4370 0.0086
-1.000 -0.0907 0.01037 0.00222 0.0200 0.4206 0.0087
-0.750 -0.0644 0.01008 0.00179 0.0205 0.4064 0.0091
-0.500 -0.0379 0.00987 0.00142 0.0209 0.3925 0.0106
-0.250 -0.0112 0.00971 0.00123 0.0213 0.3789 0.0239
0.250 0.0146 0.00678 0.00117 0.0273 0.3599 0.9046
0.500 0.0398 0.00687 0.00126 0.0287 0.3501 0.9520
0.750 0.0811 0.00707 0.00134 0.0263 0.3380 0.9710
1.000 0.1173 0.00717 0.00134 0.0245 0.3268 0.9744
1.250 0.1476 0.00723 0.00132 0.0240 0.3178 0.9760
1.500 0.1771 0.00730 0.00131 0.0236 0.3098 0.9777
1.750 0.2060 0.00737 0.00132 0.0233 0.3017 0.9796
2.000 0.2344 0.00744 0.00134 0.0231 0.2936 0.9817
2.500 0.2962 0.00759 0.00141 0.0217 0.2791 0.9837
3.000 0.3574 0.00776 0.00155 0.0203 0.2666 0.9865
3.250 0.3873 0.00787 0.00164 0.0198 0.2600 0.9881
3.500 0.4169 0.00796 0.00174 0.0193 0.2540 0.9898
3.750 0.4461 0.00807 0.00186 0.0189 0.2484 0.9916
4.000 0.4752 0.00818 0.00203 0.0186 0.2436 0.9935
4.250 0.5053 0.00829 0.00218 0.0180 0.2380 0.9951
4.500 0.5355 0.00844 0.00235 0.0173 0.2325 0.9969
4.750 0.5654 0.00857 0.00249 0.0167 0.2179 0.9986
5.000 0.5943 0.00877 0.00264 0.0163 0.1948 1.0000
5.250 0.6167 0.00909 0.00284 0.0172 0.1584 1.0000
5.500 0.6335 0.01062 0.00377 0.0184 0.0283 1.0000
5.750 0.6553 0.01120 0.00433 0.0194 0.0118 1.0000
6.000 0.6780 0.01167 0.00493 0.0204 0.0085 1.0000
6.250 0.7008 0.01222 0.00560 0.0213 0.0065 1.0000
6.500 0.7227 0.01303 0.00655 0.0222 0.0053 1.0000
6.750 0.7433 0.01416 0.00787 0.0232 0.0043 1.0000
7.000 0.7672 0.01454 0.00830 0.0238 0.0037 1.0000
7.250 0.7891 0.01545 0.00935 0.0245 0.0033 1.0000
7.500 0.8106 0.01645 0.01049 0.0253 0.0030 1.0000
7.750 0.8322 0.01745 0.01169 0.0260 0.0026 1.0000
8.000 0.8533 0.01853 0.01289 0.0266 0.0024 1.0000
8.250 0.8740 0.01966 0.01415 0.0272 0.0021 1.0000
8.500 0.8911 0.02161 0.01630 0.0280 0.0019 1.0000
8.750 0.9041 0.02453 0.01954 0.0292 0.0018 1.0000
9.000 0.9132 0.02829 0.02371 0.0303 0.0017 1.0000
9.500 0.9245 0.03559 0.03171 0.0315 0.0017 1.0000
9.750 0.9231 0.03938 0.03582 0.0315 0.0017 1.0000
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Polar data table (+)
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