NACA 64-110 AIRFOIL (n64110-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 64-110 AIRFOIL (n64110-il) Reynolds number: 500,000 Max Cl/Cd: 54.09 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n64110-il-500000-n5.txt Download as CSV file: xf-n64110-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-110 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
------ -------- --------- --------- -------- -------- --------
-11.000 -0.8806 0.04605 0.04325 -0.0344 1.0000 0.0075
-10.750 -0.9247 0.03920 0.03596 -0.0313 1.0000 0.0076
-10.500 -0.9415 0.03357 0.02978 -0.0289 1.0000 0.0079
-10.250 -0.9386 0.03013 0.02591 -0.0273 1.0000 0.0082
-10.000 -0.9273 0.02780 0.02325 -0.0261 1.0000 0.0084
-9.750 -0.9161 0.02523 0.02037 -0.0249 1.0000 0.0088
-9.500 -0.8973 0.02410 0.01911 -0.0243 1.0000 0.0091
-9.250 -0.8767 0.02322 0.01813 -0.0238 1.0000 0.0094
-9.000 -0.8554 0.02236 0.01716 -0.0234 1.0000 0.0097
-8.750 -0.8339 0.02142 0.01608 -0.0229 1.0000 0.0101
-8.500 -0.8117 0.02057 0.01509 -0.0224 1.0000 0.0107
-8.250 -0.7893 0.01969 0.01407 -0.0219 1.0000 0.0113
-8.000 -0.7667 0.01880 0.01304 -0.0214 1.0000 0.0118
-7.750 -0.7439 0.01797 0.01207 -0.0208 1.0000 0.0120
-7.500 -0.7237 0.01660 0.01058 -0.0199 1.0000 0.0127
-7.250 -0.6984 0.01584 0.00977 -0.0200 0.9950 0.0132
-7.000 -0.6660 0.01530 0.00918 -0.0214 0.9807 0.0139
-6.750 -0.6334 0.01480 0.00862 -0.0229 0.9683 0.0148
-6.500 -0.6025 0.01416 0.00790 -0.0239 0.9556 0.0155
-6.250 -0.5742 0.01357 0.00721 -0.0243 0.9421 0.0161
-6.000 -0.5484 0.01306 0.00659 -0.0241 0.9282 0.0166
-5.750 -0.5234 0.01265 0.00608 -0.0236 0.9148 0.0170
-5.500 -0.5001 0.01199 0.00535 -0.0230 0.9019 0.0184
-5.250 -0.4752 0.01163 0.00494 -0.0225 0.8900 0.0198
-5.000 -0.4499 0.01130 0.00453 -0.0221 0.8786 0.0210
-4.750 -0.4242 0.01100 0.00415 -0.0218 0.8674 0.0226
-4.500 -0.3981 0.01075 0.00382 -0.0215 0.8570 0.0239
-4.250 -0.3723 0.01040 0.00341 -0.0212 0.8469 0.0272
-4.000 -0.3458 0.01016 0.00312 -0.0210 0.8367 0.0307
-3.750 -0.3189 0.00998 0.00288 -0.0209 0.8270 0.0339
-3.500 -0.2923 0.00973 0.00262 -0.0207 0.8176 0.0430
-3.250 -0.2655 0.00946 0.00239 -0.0206 0.8082 0.0614
-3.000 -0.2395 0.00901 0.00213 -0.0206 0.7991 0.1176
-2.750 -0.2142 0.00839 0.00185 -0.0205 0.7900 0.2186
-2.500 -0.1900 0.00747 0.00153 -0.0205 0.7804 0.3848
-2.250 -0.1639 0.00706 0.00142 -0.0205 0.7715 0.4809
-2.000 -0.1367 0.00689 0.00134 -0.0204 0.7628 0.5256
-1.750 -0.1098 0.00670 0.00132 -0.0203 0.7543 0.5809
-1.500 -0.0831 0.00657 0.00133 -0.0200 0.7459 0.6326
-1.250 -0.0553 0.00652 0.00131 -0.0199 0.7368 0.6524
-1.000 -0.0272 0.00651 0.00127 -0.0200 0.7283 0.6629
-0.750 0.0009 0.00651 0.00124 -0.0200 0.7199 0.6719
-0.500 0.0291 0.00650 0.00122 -0.0201 0.7121 0.6793
-0.250 0.0574 0.00651 0.00120 -0.0201 0.7042 0.6875
0.000 0.0856 0.00650 0.00121 -0.0202 0.6962 0.6955
0.500 0.1419 0.00653 0.00124 -0.0203 0.6802 0.7121
0.750 0.1701 0.00655 0.00126 -0.0203 0.6732 0.7203
1.000 0.1983 0.00656 0.00130 -0.0204 0.6653 0.7289
1.250 0.2263 0.00659 0.00135 -0.0204 0.6561 0.7376
1.500 0.2541 0.00664 0.00139 -0.0204 0.6432 0.7470
1.750 0.2815 0.00669 0.00144 -0.0203 0.6230 0.7557
2.000 0.3082 0.00683 0.00147 -0.0200 0.5864 0.7644
2.250 0.3344 0.00702 0.00153 -0.0197 0.5380 0.7735
2.500 0.3609 0.00721 0.00162 -0.0195 0.4964 0.7826
2.750 0.3864 0.00757 0.00174 -0.0192 0.4256 0.7924
3.000 0.4090 0.00839 0.00207 -0.0187 0.2887 0.8022
3.250 0.4323 0.00916 0.00242 -0.0183 0.1785 0.8117
3.500 0.4561 0.00984 0.00277 -0.0180 0.0920 0.8218
3.750 0.4810 0.01029 0.00306 -0.0176 0.0511 0.8321
4.000 0.5068 0.01056 0.00334 -0.0173 0.0396 0.8425
4.250 0.5327 0.01079 0.00361 -0.0170 0.0341 0.8532
4.750 0.5830 0.01138 0.00429 -0.0161 0.0257 0.8765
5.000 0.6079 0.01160 0.00460 -0.0156 0.0234 0.8891
5.250 0.6321 0.01188 0.00493 -0.0149 0.0211 0.9026
5.500 0.6552 0.01227 0.00537 -0.0141 0.0191 0.9176
5.750 0.6781 0.01270 0.00588 -0.0132 0.0179 0.9343
6.000 0.7033 0.01308 0.00637 -0.0128 0.0172 0.9533
6.250 0.7340 0.01357 0.00693 -0.0137 0.0163 0.9765
6.500 0.7617 0.01411 0.00752 -0.0141 0.0155 1.0000
6.750 0.7860 0.01467 0.00813 -0.0138 0.0150 1.0000
7.000 0.8105 0.01520 0.00869 -0.0135 0.0143 1.0000
7.250 0.8338 0.01587 0.00940 -0.0131 0.0135 1.0000
7.500 0.8549 0.01696 0.01057 -0.0123 0.0129 1.0000
7.750 0.8780 0.01771 0.01141 -0.0118 0.0126 1.0000
8.000 0.9007 0.01856 0.01235 -0.0113 0.0124 1.0000
8.250 0.9232 0.01942 0.01333 -0.0107 0.0120 1.0000
8.500 0.9456 0.02032 0.01434 -0.0101 0.0115 1.0000
8.750 0.9680 0.02108 0.01522 -0.0097 0.0110 1.0000
9.000 0.9900 0.02185 0.01608 -0.0092 0.0106 1.0000
9.250 1.0113 0.02267 0.01700 -0.0086 0.0102 1.0000
9.500 1.0317 0.02359 0.01801 -0.0079 0.0099 1.0000
9.750 1.0511 0.02457 0.01909 -0.0072 0.0095 1.0000
10.000 1.0657 0.02648 0.02119 -0.0061 0.0091 1.0000
10.250 1.0796 0.02837 0.02333 -0.0048 0.0089 1.0000
10.500 1.0938 0.02996 0.02519 -0.0036 0.0087 1.0000
10.750 1.1052 0.03177 0.02724 -0.0022 0.0083 1.0000
11.000 1.1147 0.03352 0.02922 -0.0006 0.0080 1.0000
11.250 1.1169 0.03572 0.03168 0.0016 0.0077 1.0000
11.500 1.1135 0.03793 0.03411 0.0043 0.0075 1.0000
11.750 1.1112 0.03993 0.03628 0.0063 0.0073 1.0000
12.000 1.0995 0.04316 0.03975 0.0080 0.0072 1.0000
12.250 1.0815 0.04736 0.04420 0.0088 0.0071 1.0000
12.500 1.0718 0.05083 0.04784 0.0086 0.0070 1.0000
12.750 1.0423 0.05756 0.05484 0.0065 0.0070 1.0000
13.000 1.0093 0.06592 0.06345 0.0022 0.0070 1.0000
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Polar data table (+)
Polar graphs
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