NACA 64-108 AIRFOIL (n64108-il) Xfoil prediction polar at RE=500,000 Ncrit=5
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Airfoil: NACA 64-108 AIRFOIL (n64108-il) Reynolds number: 500,000 Max Cl/Cd: 50.28 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n64108-il-500000-n5.txt Download as CSV file: xf-n64108-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-108 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.6546 0.08420 0.08203 -0.0052 1.0000 0.0052
-9.500 -0.6629 0.07773 0.07559 -0.0098 1.0000 0.0052
-9.250 -0.6773 0.06924 0.06712 -0.0183 1.0000 0.0051
-9.000 -0.6969 0.06346 0.06128 -0.0222 1.0000 0.0051
-8.750 -0.7121 0.05696 0.05464 -0.0244 1.0000 0.0051
-8.500 -0.7217 0.05013 0.04757 -0.0255 1.0000 0.0051
-8.250 -0.7288 0.04216 0.03923 -0.0255 1.0000 0.0051
-8.000 -0.7505 0.02634 0.02218 -0.0231 1.0000 0.0054
-7.750 -0.7332 0.02365 0.01914 -0.0225 1.0000 0.0057
-7.500 -0.7113 0.02241 0.01771 -0.0221 1.0000 0.0059
-7.250 -0.6886 0.02127 0.01640 -0.0218 1.0000 0.0063
-7.000 -0.6662 0.01969 0.01458 -0.0213 1.0000 0.0068
-6.750 -0.6434 0.01816 0.01280 -0.0207 1.0000 0.0074
-6.500 -0.6188 0.01744 0.01194 -0.0203 1.0000 0.0083
-6.250 -0.5942 0.01673 0.01105 -0.0199 1.0000 0.0088
-6.000 -0.5729 0.01472 0.00881 -0.0190 1.0000 0.0097
-5.750 -0.5490 0.01407 0.00810 -0.0186 1.0000 0.0105
-5.500 -0.5230 0.01348 0.00746 -0.0185 0.9979 0.0113
-5.250 -0.4906 0.01286 0.00675 -0.0197 0.9881 0.0124
-5.000 -0.4578 0.01244 0.00626 -0.0211 0.9786 0.0138
-4.750 -0.4254 0.01189 0.00560 -0.0223 0.9684 0.0144
-4.500 -0.3946 0.01106 0.00467 -0.0232 0.9571 0.0153
-4.250 -0.3653 0.01039 0.00392 -0.0237 0.9449 0.0169
-4.000 -0.3372 0.01004 0.00352 -0.0239 0.9321 0.0187
-3.750 -0.3104 0.00971 0.00312 -0.0237 0.9191 0.0206
-3.500 -0.2839 0.00946 0.00276 -0.0234 0.9066 0.0229
-3.250 -0.2576 0.00921 0.00244 -0.0230 0.8943 0.0265
-3.000 -0.2313 0.00899 0.00218 -0.0227 0.8825 0.0329
-2.750 -0.2047 0.00881 0.00197 -0.0225 0.8708 0.0418
-2.500 -0.1783 0.00851 0.00175 -0.0223 0.8595 0.0768
-2.250 -0.1535 0.00767 0.00147 -0.0223 0.8483 0.2344
-2.000 -0.1293 0.00676 0.00122 -0.0222 0.8375 0.4316
-1.750 -0.1030 0.00646 0.00113 -0.0220 0.8273 0.5121
-1.500 -0.0768 0.00619 0.00108 -0.0218 0.8165 0.5899
-1.250 -0.0512 0.00595 0.00111 -0.0213 0.8059 0.6693
-1.000 -0.0246 0.00586 0.00111 -0.0209 0.7955 0.7068
-0.500 0.0303 0.00583 0.00106 -0.0206 0.7752 0.7350
-0.250 0.0581 0.00584 0.00104 -0.0206 0.7648 0.7454
0.000 0.0857 0.00584 0.00104 -0.0205 0.7545 0.7558
0.250 0.1132 0.00585 0.00105 -0.0204 0.7446 0.7664
0.500 0.1409 0.00586 0.00107 -0.0203 0.7346 0.7771
0.750 0.1685 0.00588 0.00110 -0.0202 0.7243 0.7877
1.000 0.1956 0.00593 0.00112 -0.0200 0.7055 0.7985
1.250 0.2222 0.00600 0.00115 -0.0196 0.6793 0.8095
1.500 0.2488 0.00608 0.00120 -0.0193 0.6510 0.8206
1.750 0.2743 0.00629 0.00123 -0.0187 0.5955 0.8317
2.000 0.2993 0.00659 0.00129 -0.0182 0.5238 0.8430
2.250 0.3246 0.00690 0.00140 -0.0178 0.4552 0.8550
2.750 0.3683 0.00888 0.00208 -0.0166 0.1017 0.8802
3.000 0.3929 0.00928 0.00233 -0.0161 0.0520 0.8935
3.250 0.4180 0.00951 0.00255 -0.0156 0.0377 0.9077
3.500 0.4432 0.00972 0.00281 -0.0150 0.0310 0.9229
3.750 0.4686 0.00997 0.00311 -0.0145 0.0257 0.9393
4.000 0.4966 0.01020 0.00341 -0.0146 0.0222 0.9568
4.500 0.5580 0.01122 0.00453 -0.0163 0.0165 1.0000
4.750 0.5841 0.01164 0.00502 -0.0161 0.0159 1.0000
5.000 0.6099 0.01213 0.00558 -0.0159 0.0152 1.0000
5.250 0.6352 0.01273 0.00625 -0.0156 0.0144 1.0000
5.500 0.6609 0.01324 0.00682 -0.0154 0.0132 1.0000
5.750 0.6865 0.01374 0.00735 -0.0152 0.0122 1.0000
6.000 0.7116 0.01436 0.00803 -0.0149 0.0114 1.0000
6.250 0.7355 0.01530 0.00907 -0.0145 0.0107 1.0000
6.500 0.7575 0.01699 0.01095 -0.0137 0.0098 1.0000
6.750 0.7832 0.01744 0.01149 -0.0135 0.0092 1.0000
7.000 0.8080 0.01821 0.01238 -0.0132 0.0084 1.0000
7.250 0.8317 0.01936 0.01371 -0.0127 0.0077 1.0000
7.500 0.8554 0.02037 0.01487 -0.0123 0.0071 1.0000
7.750 0.8789 0.02126 0.01588 -0.0120 0.0067 1.0000
8.000 0.9021 0.02209 0.01686 -0.0117 0.0063 1.0000
8.250 0.9181 0.02516 0.02032 -0.0106 0.0058 1.0000
8.500 0.9353 0.02781 0.02337 -0.0095 0.0056 1.0000
8.750 0.9433 0.03294 0.02912 -0.0077 0.0051 1.0000
9.000 0.9174 0.04657 0.04367 -0.0040 0.0046 1.0000
9.250 0.9014 0.05457 0.05206 -0.0025 0.0044 1.0000
9.500 0.8840 0.06089 0.05860 -0.0019 0.0044 1.0000
9.750 0.8581 0.06601 0.06383 -0.0013 0.0045 1.0000
10.000 0.8347 0.07196 0.06987 -0.0043 0.0045 1.0000
10.250 0.8111 0.08123 0.07922 -0.0123 0.0049 1.0000
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Polar data table (+)
Polar graphs
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