NACA 64-008A AIRFOIL (n64008a-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 64-008A AIRFOIL (n64008a-il) Reynolds number: 1,000,000 Max Cl/Cd: 59.37 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n64008a-il-1000000-n5.txt Download as CSV file: xf-n64008a-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-008A AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -1.0145 0.02567 0.02220 -0.0084 1.0000 0.0027
-9.750 -1.0004 0.02299 0.01916 -0.0073 1.0000 0.0028
-9.500 -0.9824 0.02109 0.01699 -0.0064 1.0000 0.0029
-9.250 -0.9625 0.01960 0.01527 -0.0056 1.0000 0.0030
-9.000 -0.9412 0.01839 0.01388 -0.0050 1.0000 0.0031
-8.750 -0.9189 0.01740 0.01272 -0.0044 1.0000 0.0032
-8.500 -0.8987 0.01587 0.01096 -0.0035 1.0000 0.0035
-8.250 -0.8743 0.01536 0.01039 -0.0032 1.0000 0.0039
-8.000 -0.8490 0.01509 0.01011 -0.0030 1.0000 0.0042
-7.750 -0.8240 0.01472 0.00970 -0.0028 1.0000 0.0046
-7.500 -0.7995 0.01423 0.00914 -0.0024 1.0000 0.0050
-7.250 -0.7754 0.01364 0.00847 -0.0019 1.0000 0.0054
-7.000 -0.7511 0.01312 0.00787 -0.0014 1.0000 0.0056
-6.750 -0.7265 0.01270 0.00738 -0.0010 1.0000 0.0058
-6.500 -0.7040 0.01186 0.00647 -0.0002 1.0000 0.0066
-6.250 -0.6792 0.01157 0.00615 0.0002 1.0000 0.0071
-6.000 -0.6545 0.01131 0.00587 0.0006 1.0000 0.0077
-5.750 -0.6299 0.01106 0.00560 0.0011 1.0000 0.0085
-5.500 -0.6045 0.01074 0.00526 0.0014 0.9993 0.0091
-5.250 -0.5726 0.01037 0.00483 0.0003 0.9953 0.0096
-5.000 -0.5413 0.01005 0.00449 -0.0007 0.9899 0.0099
-4.750 -0.5107 0.00943 0.00375 -0.0016 0.9831 0.0109
-4.500 -0.4793 0.00899 0.00325 -0.0026 0.9752 0.0124
-4.250 -0.4482 0.00869 0.00292 -0.0034 0.9649 0.0131
-4.000 -0.4184 0.00844 0.00262 -0.0040 0.9519 0.0138
-3.750 -0.3904 0.00823 0.00236 -0.0040 0.9363 0.0145
-3.500 -0.3638 0.00807 0.00212 -0.0038 0.9188 0.0155
-3.250 -0.3375 0.00793 0.00192 -0.0034 0.9015 0.0168
-3.000 -0.3113 0.00779 0.00172 -0.0031 0.8848 0.0218
-2.750 -0.2849 0.00765 0.00156 -0.0028 0.8691 0.0317
-2.500 -0.2585 0.00748 0.00141 -0.0026 0.8541 0.0498
-2.250 -0.2321 0.00729 0.00126 -0.0024 0.8401 0.0790
-2.000 -0.2062 0.00696 0.00111 -0.0022 0.8270 0.1426
-1.500 -0.1552 0.00615 0.00082 -0.0017 0.8017 0.3210
-1.250 -0.1321 0.00537 0.00066 -0.0013 0.7901 0.5085
-1.000 -0.1079 0.00493 0.00063 -0.0007 0.7786 0.6277
-0.750 -0.0816 0.00480 0.00061 -0.0004 0.7672 0.6738
-0.500 -0.0547 0.00474 0.00060 -0.0002 0.7564 0.7006
-0.250 -0.0275 0.00470 0.00059 -0.0001 0.7452 0.7205
0.000 0.0000 0.00469 0.00059 0.0000 0.7339 0.7339
0.250 0.0275 0.00470 0.00059 0.0001 0.7205 0.7451
0.500 0.0547 0.00474 0.00060 0.0002 0.7006 0.7564
0.750 0.0816 0.00480 0.00061 0.0004 0.6736 0.7673
1.000 0.1079 0.00493 0.00063 0.0007 0.6277 0.7786
1.250 0.1321 0.00536 0.00066 0.0013 0.5091 0.7901
1.500 0.1551 0.00615 0.00082 0.0018 0.3196 0.8017
2.000 0.2060 0.00697 0.00111 0.0022 0.1396 0.8270
2.250 0.2319 0.00729 0.00126 0.0024 0.0793 0.8402
2.500 0.2583 0.00748 0.00141 0.0026 0.0494 0.8542
2.750 0.2847 0.00765 0.00156 0.0028 0.0314 0.8692
3.000 0.3110 0.00779 0.00172 0.0031 0.0220 0.8848
3.250 0.3373 0.00793 0.00192 0.0035 0.0168 0.9015
3.500 0.3636 0.00806 0.00212 0.0038 0.0155 0.9188
3.750 0.3903 0.00823 0.00236 0.0041 0.0145 0.9364
4.000 0.4183 0.00844 0.00262 0.0040 0.0138 0.9520
4.250 0.4482 0.00869 0.00292 0.0034 0.0131 0.9651
4.500 0.4793 0.00899 0.00325 0.0026 0.0124 0.9753
4.750 0.5109 0.00943 0.00375 0.0016 0.0109 0.9832
5.000 0.5415 0.01005 0.00449 0.0007 0.0099 0.9899
5.250 0.5729 0.01036 0.00483 -0.0003 0.0096 0.9953
5.500 0.6047 0.01074 0.00525 -0.0014 0.0091 0.9994
5.750 0.6300 0.01106 0.00560 -0.0011 0.0084 1.0000
6.000 0.6546 0.01130 0.00587 -0.0006 0.0077 1.0000
6.250 0.6793 0.01157 0.00615 -0.0002 0.0071 1.0000
6.500 0.7041 0.01186 0.00646 0.0002 0.0066 1.0000
6.750 0.7265 0.01270 0.00738 0.0010 0.0058 1.0000
7.000 0.7510 0.01312 0.00787 0.0015 0.0056 1.0000
7.250 0.7753 0.01364 0.00847 0.0019 0.0054 1.0000
7.500 0.7994 0.01421 0.00912 0.0024 0.0050 1.0000
7.750 0.8237 0.01474 0.00972 0.0028 0.0046 1.0000
8.000 0.8487 0.01509 0.01010 0.0031 0.0042 1.0000
8.250 0.8739 0.01537 0.01040 0.0033 0.0039 1.0000
8.500 0.8982 0.01587 0.01097 0.0036 0.0035 1.0000
8.750 0.9183 0.01739 0.01271 0.0045 0.0032 1.0000
9.000 0.9405 0.01838 0.01387 0.0051 0.0031 1.0000
9.250 0.9617 0.01961 0.01528 0.0058 0.0030 1.0000
9.500 0.9816 0.02110 0.01700 0.0065 0.0029 1.0000
9.750 0.9994 0.02300 0.01918 0.0075 0.0028 1.0000
10.000 1.0135 0.02567 0.02221 0.0087 0.0027 1.0000
12.500 0.7208 0.14303 0.14142 -0.0359 0.0023 1.0000
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Polar data table (+)
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