NACA 64-008A AIRFOIL (n64008a-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 64-008A AIRFOIL (n64008a-il) Reynolds number: 50,000 Max Cl/Cd: 23.28 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n64008a-il-50000-n5.txt Download as CSV file: xf-n64008a-il-50000-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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.6780 0.09403 0.08726 -0.0016 1.0000 0.0473
-9.000 -0.6841 0.08781 0.08107 -0.0066 1.0000 0.0461
-8.750 -0.6945 0.08215 0.07540 -0.0105 1.0000 0.0449
-8.500 -0.7045 0.07691 0.07008 -0.0130 1.0000 0.0438
-8.250 -0.7106 0.07194 0.06495 -0.0146 1.0000 0.0433
-8.000 -0.7119 0.06745 0.06027 -0.0154 1.0000 0.0436
-7.750 -0.7104 0.06305 0.05561 -0.0157 1.0000 0.0438
-7.500 -0.7061 0.05862 0.05086 -0.0157 1.0000 0.0438
-7.250 -0.6989 0.05417 0.04599 -0.0153 1.0000 0.0432
-7.000 -0.6884 0.04996 0.04129 -0.0145 1.0000 0.0427
-6.750 -0.6748 0.04610 0.03694 -0.0136 1.0000 0.0426
-6.500 -0.6585 0.04258 0.03294 -0.0126 1.0000 0.0429
-6.250 -0.6397 0.03938 0.02928 -0.0115 1.0000 0.0436
-6.000 -0.6192 0.03670 0.02622 -0.0107 1.0000 0.0459
-5.750 -0.5970 0.03440 0.02343 -0.0097 1.0000 0.0502
-5.500 -0.5721 0.03208 0.02059 -0.0087 1.0000 0.0523
-5.250 -0.5460 0.02965 0.01784 -0.0079 1.0000 0.0540
-5.000 -0.5204 0.02766 0.01577 -0.0071 1.0000 0.0568
-4.750 -0.4952 0.02608 0.01404 -0.0060 1.0000 0.0607
-4.500 -0.4719 0.02475 0.01254 -0.0049 1.0000 0.0685
-4.250 -0.4497 0.02355 0.01122 -0.0039 1.0000 0.0800
-4.000 -0.4288 0.02218 0.00978 -0.0027 1.0000 0.0927
-3.750 -0.4088 0.02055 0.00845 -0.0018 1.0000 0.1310
-3.500 -0.4043 0.01736 0.00744 0.0014 1.0000 0.4751
-3.250 -0.3938 0.01692 0.00785 0.0077 1.0000 0.7151
-3.000 -0.3651 0.01801 0.00912 0.0145 1.0000 0.8731
-2.750 -0.3184 0.01819 0.00884 0.0125 1.0000 0.9111
-2.500 -0.2762 0.01795 0.00819 0.0095 1.0000 0.9268
-2.250 -0.2344 0.01766 0.00760 0.0063 1.0000 0.9408
-2.000 -0.1929 0.01737 0.00705 0.0029 1.0000 0.9541
-1.750 -0.1514 0.01706 0.00654 -0.0005 1.0000 0.9673
-1.500 -0.1098 0.01675 0.00604 -0.0042 1.0000 0.9803
-1.250 -0.0681 0.01645 0.00561 -0.0079 1.0000 0.9934
-1.000 -0.0392 0.01617 0.00529 -0.0093 1.0000 1.0000
-0.750 -0.0261 0.01598 0.00508 -0.0076 1.0000 1.0000
-0.500 -0.0154 0.01583 0.00494 -0.0055 1.0000 1.0000
-0.250 -0.0071 0.01574 0.00486 -0.0029 1.0000 1.0000
0.000 0.0000 0.01570 0.00484 0.0000 1.0000 1.0000
0.250 0.0071 0.01574 0.00486 0.0029 1.0000 1.0000
0.500 0.0154 0.01583 0.00494 0.0055 1.0000 1.0000
0.750 0.0261 0.01598 0.00508 0.0076 1.0000 1.0000
1.000 0.0392 0.01617 0.00529 0.0093 1.0000 1.0000
1.250 0.0681 0.01644 0.00561 0.0079 0.9934 1.0000
1.500 0.1098 0.01675 0.00604 0.0042 0.9803 1.0000
1.750 0.1514 0.01706 0.00654 0.0005 0.9673 1.0000
2.000 0.1928 0.01736 0.00705 -0.0029 0.9542 1.0000
2.250 0.2344 0.01766 0.00759 -0.0063 0.9408 1.0000
2.500 0.2762 0.01795 0.00818 -0.0095 0.9269 1.0000
2.750 0.3183 0.01818 0.00884 -0.0125 0.9111 1.0000
3.000 0.3651 0.01801 0.00912 -0.0145 0.8732 1.0000
3.250 0.3939 0.01692 0.00785 -0.0077 0.7151 1.0000
3.500 0.4043 0.01737 0.00744 -0.0014 0.4745 1.0000
3.750 0.4089 0.02055 0.00845 0.0018 0.1310 1.0000
4.000 0.4288 0.02218 0.00979 0.0027 0.0926 1.0000
4.250 0.4497 0.02355 0.01122 0.0039 0.0800 1.0000
4.500 0.4719 0.02475 0.01254 0.0049 0.0685 1.0000
4.750 0.4952 0.02608 0.01404 0.0060 0.0607 1.0000
5.000 0.5204 0.02766 0.01577 0.0071 0.0568 1.0000
5.250 0.5461 0.02965 0.01784 0.0079 0.0540 1.0000
5.500 0.5721 0.03208 0.02059 0.0087 0.0523 1.0000
5.750 0.5970 0.03440 0.02343 0.0097 0.0502 1.0000
6.000 0.6192 0.03670 0.02623 0.0107 0.0459 1.0000
6.250 0.6398 0.03938 0.02928 0.0116 0.0436 1.0000
6.500 0.6585 0.04258 0.03294 0.0126 0.0429 1.0000
6.750 0.6748 0.04610 0.03694 0.0136 0.0426 1.0000
7.000 0.6884 0.04996 0.04129 0.0145 0.0427 1.0000
7.250 0.6989 0.05417 0.04599 0.0153 0.0432 1.0000
7.500 0.7061 0.05862 0.05086 0.0157 0.0438 1.0000
7.750 0.7104 0.06304 0.05561 0.0157 0.0438 1.0000
8.000 0.7120 0.06744 0.06026 0.0154 0.0436 1.0000
8.250 0.7107 0.07193 0.06494 0.0146 0.0433 1.0000
8.500 0.7047 0.07689 0.07006 0.0130 0.0438 1.0000
8.750 0.6947 0.08215 0.07540 0.0105 0.0448 1.0000
9.000 0.6844 0.08783 0.08109 0.0066 0.0461 1.0000
9.250 0.6784 0.09405 0.08728 0.0015 0.0473 1.0000
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Polar data table (+)
Polar graphs
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