NACA 64-012A AIRFOIL (n64012a-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 64-012A AIRFOIL (n64012a-il) Reynolds number: 1,000,000 Max Cl/Cd: 64.88 at α=8.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n64012a-il-1000000-n5.txt Download as CSV file: xf-n64012a-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-012A 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
------ -------- --------- --------- -------- -------- --------
-18.000 -1.0719 0.12038 0.11827 0.0105 1.0000 0.0068
-17.750 -1.1445 0.10053 0.09818 -0.0001 1.0000 0.0067
-17.500 -1.1872 0.08762 0.08507 -0.0074 1.0000 0.0067
-17.250 -1.2161 0.07821 0.07549 -0.0127 1.0000 0.0066
-17.000 -1.2366 0.07087 0.06799 -0.0166 1.0000 0.0067
-16.750 -1.2513 0.06493 0.06190 -0.0195 1.0000 0.0067
-16.500 -1.2633 0.05976 0.05660 -0.0218 1.0000 0.0067
-16.250 -1.2714 0.05539 0.05209 -0.0235 1.0000 0.0067
-16.000 -1.2771 0.05153 0.04810 -0.0248 1.0000 0.0068
-15.750 -1.2815 0.04797 0.04440 -0.0258 1.0000 0.0068
-15.500 -1.2922 0.04386 0.04014 -0.0266 1.0000 0.0069
-15.250 -1.2969 0.04061 0.03675 -0.0269 1.0000 0.0070
-15.000 -1.2973 0.03796 0.03400 -0.0269 1.0000 0.0072
-14.750 -1.2945 0.03575 0.03168 -0.0267 1.0000 0.0073
-14.500 -1.2895 0.03383 0.02968 -0.0263 1.0000 0.0074
-14.250 -1.2828 0.03215 0.02790 -0.0257 1.0000 0.0075
-14.000 -1.2748 0.03065 0.02632 -0.0250 1.0000 0.0076
-13.750 -1.2658 0.02929 0.02488 -0.0242 1.0000 0.0077
-13.500 -1.2561 0.02804 0.02355 -0.0232 1.0000 0.0078
-13.250 -1.2456 0.02690 0.02231 -0.0221 1.0000 0.0079
-13.000 -1.2345 0.02584 0.02117 -0.0209 1.0000 0.0081
-12.750 -1.2228 0.02485 0.02009 -0.0195 1.0000 0.0082
-12.500 -1.2104 0.02393 0.01909 -0.0181 1.0000 0.0084
-12.250 -1.1971 0.02310 0.01817 -0.0166 1.0000 0.0085
-12.000 -1.1827 0.02235 0.01734 -0.0151 1.0000 0.0087
-11.750 -1.1667 0.02167 0.01659 -0.0138 1.0000 0.0089
-11.500 -1.1488 0.02100 0.01585 -0.0127 1.0000 0.0091
-11.250 -1.1301 0.02037 0.01515 -0.0118 1.0000 0.0092
-11.000 -1.1109 0.01974 0.01443 -0.0108 1.0000 0.0093
-10.750 -1.0913 0.01913 0.01376 -0.0099 1.0000 0.0094
-10.500 -1.0753 0.01813 0.01267 -0.0086 1.0000 0.0097
-10.250 -1.0568 0.01736 0.01184 -0.0074 1.0000 0.0099
-10.000 -1.0368 0.01674 0.01117 -0.0065 1.0000 0.0101
-9.750 -1.0159 0.01621 0.01060 -0.0056 1.0000 0.0104
-9.500 -0.9944 0.01575 0.01010 -0.0049 1.0000 0.0107
-9.250 -0.9729 0.01531 0.00963 -0.0040 1.0000 0.0109
-9.000 -0.9518 0.01486 0.00915 -0.0031 1.0000 0.0112
-8.750 -0.9313 0.01441 0.00866 -0.0020 1.0000 0.0114
-8.500 -0.9040 0.01394 0.00814 -0.0024 0.9962 0.0117
-8.250 -0.8741 0.01347 0.00762 -0.0033 0.9894 0.0120
-8.000 -0.8436 0.01303 0.00714 -0.0043 0.9812 0.0123
-7.750 -0.8130 0.01264 0.00670 -0.0053 0.9712 0.0125
-7.500 -0.7840 0.01221 0.00623 -0.0059 0.9577 0.0130
-7.250 -0.7578 0.01180 0.00577 -0.0058 0.9405 0.0137
-7.000 -0.7331 0.01150 0.00541 -0.0054 0.9214 0.0143
-6.750 -0.7088 0.01125 0.00508 -0.0048 0.9027 0.0149
-6.500 -0.6841 0.01101 0.00477 -0.0043 0.8855 0.0155
-6.250 -0.6589 0.01080 0.00448 -0.0039 0.8693 0.0162
-6.000 -0.6334 0.01059 0.00420 -0.0035 0.8545 0.0170
-5.750 -0.6081 0.01035 0.00391 -0.0032 0.8408 0.0189
-5.500 -0.5821 0.01015 0.00367 -0.0029 0.8280 0.0209
-5.250 -0.5558 0.00997 0.00344 -0.0028 0.8164 0.0231
-5.000 -0.5297 0.00976 0.00322 -0.0026 0.8057 0.0269
-4.750 -0.5032 0.00959 0.00302 -0.0024 0.7955 0.0309
-4.500 -0.4767 0.00940 0.00282 -0.0023 0.7857 0.0371
-4.250 -0.4503 0.00919 0.00263 -0.0022 0.7764 0.0469
-4.000 -0.4238 0.00900 0.00245 -0.0020 0.7670 0.0588
-3.750 -0.3973 0.00878 0.00228 -0.0019 0.7585 0.0755
-3.500 -0.3714 0.00851 0.00209 -0.0017 0.7504 0.1042
-3.250 -0.3465 0.00806 0.00186 -0.0015 0.7419 0.1648
-3.000 -0.3229 0.00750 0.00162 -0.0010 0.7337 0.2523
-2.750 -0.2981 0.00706 0.00142 -0.0007 0.7252 0.3232
-2.500 -0.2742 0.00656 0.00122 -0.0002 0.7176 0.4116
-2.250 -0.2486 0.00625 0.00112 0.0000 0.7099 0.4760
-2.000 -0.2219 0.00609 0.00106 0.0002 0.7028 0.5118
-1.750 -0.1951 0.00594 0.00101 0.0003 0.6950 0.5471
-1.500 -0.1680 0.00584 0.00097 0.0004 0.6875 0.5755
-1.250 -0.1404 0.00578 0.00095 0.0004 0.6795 0.5943
-1.000 -0.1126 0.00574 0.00092 0.0004 0.6727 0.6095
-0.750 -0.0847 0.00570 0.00091 0.0003 0.6655 0.6220
-0.250 -0.0282 0.00569 0.00088 0.0001 0.6513 0.6374
0.000 0.0000 0.00570 0.00088 0.0000 0.6442 0.6442
0.250 0.0283 0.00569 0.00088 -0.0001 0.6374 0.6513
0.750 0.0847 0.00570 0.00091 -0.0003 0.6221 0.6655
1.000 0.1127 0.00574 0.00092 -0.0004 0.6095 0.6727
1.250 0.1404 0.00578 0.00095 -0.0004 0.5942 0.6795
1.500 0.1681 0.00584 0.00097 -0.0004 0.5755 0.6875
1.750 0.1951 0.00594 0.00101 -0.0003 0.5469 0.6949
2.000 0.2219 0.00610 0.00106 -0.0002 0.5115 0.7028
2.250 0.2487 0.00625 0.00112 0.0000 0.4759 0.7099
2.500 0.2741 0.00657 0.00122 0.0002 0.4098 0.7176
2.750 0.2982 0.00706 0.00142 0.0007 0.3233 0.7252
3.000 0.3229 0.00750 0.00162 0.0010 0.2527 0.7338
3.250 0.3465 0.00807 0.00186 0.0015 0.1641 0.7420
3.500 0.3714 0.00851 0.00209 0.0017 0.1044 0.7504
3.750 0.3973 0.00878 0.00228 0.0019 0.0757 0.7585
4.000 0.4239 0.00900 0.00245 0.0020 0.0587 0.7670
4.250 0.4503 0.00919 0.00263 0.0022 0.0469 0.7764
4.500 0.4768 0.00940 0.00282 0.0023 0.0371 0.7857
4.750 0.5033 0.00959 0.00302 0.0024 0.0309 0.7955
5.000 0.5297 0.00976 0.00322 0.0026 0.0269 0.8057
5.250 0.5559 0.00997 0.00344 0.0028 0.0231 0.8165
5.500 0.5821 0.01015 0.00367 0.0029 0.0209 0.8280
5.750 0.6081 0.01035 0.00391 0.0032 0.0189 0.8407
6.000 0.6335 0.01059 0.00420 0.0035 0.0170 0.8544
6.250 0.6590 0.01080 0.00448 0.0039 0.0162 0.8692
6.500 0.6841 0.01101 0.00477 0.0043 0.0155 0.8857
6.750 0.7088 0.01124 0.00508 0.0048 0.0149 0.9029
7.000 0.7332 0.01150 0.00541 0.0054 0.0143 0.9214
7.250 0.7578 0.01180 0.00577 0.0058 0.0137 0.9403
7.500 0.7841 0.01221 0.00623 0.0059 0.0130 0.9576
7.750 0.8130 0.01264 0.00670 0.0053 0.0125 0.9713
8.000 0.8435 0.01303 0.00714 0.0043 0.0123 0.9813
8.250 0.8739 0.01347 0.00762 0.0033 0.0120 0.9894
8.500 0.9038 0.01394 0.00814 0.0024 0.0117 0.9961
8.750 0.9313 0.01441 0.00866 0.0020 0.0114 1.0000
9.000 0.9518 0.01486 0.00914 0.0031 0.0112 1.0000
9.250 0.9729 0.01531 0.00963 0.0040 0.0109 1.0000
9.500 0.9944 0.01575 0.01010 0.0049 0.0107 1.0000
9.750 1.0159 0.01621 0.01060 0.0057 0.0104 1.0000
10.000 1.0367 0.01674 0.01117 0.0065 0.0101 1.0000
10.250 1.0567 0.01736 0.01184 0.0075 0.0099 1.0000
10.500 1.0752 0.01813 0.01267 0.0086 0.0097 1.0000
10.750 1.0913 0.01912 0.01375 0.0100 0.0094 1.0000
11.000 1.1109 0.01973 0.01443 0.0108 0.0093 1.0000
11.250 1.1301 0.02036 0.01514 0.0118 0.0092 1.0000
11.500 1.1488 0.02100 0.01585 0.0127 0.0091 1.0000
11.750 1.1669 0.02165 0.01657 0.0137 0.0089 1.0000
12.000 1.1828 0.02234 0.01734 0.0151 0.0087 1.0000
12.250 1.1973 0.02309 0.01816 0.0165 0.0085 1.0000
12.500 1.2107 0.02393 0.01908 0.0180 0.0084 1.0000
12.750 1.2231 0.02485 0.02009 0.0195 0.0082 1.0000
13.000 1.2349 0.02584 0.02117 0.0208 0.0081 1.0000
13.250 1.2461 0.02689 0.02231 0.0220 0.0079 1.0000
13.500 1.2568 0.02803 0.02353 0.0231 0.0078 1.0000
13.750 1.2664 0.02928 0.02487 0.0241 0.0077 1.0000
14.000 1.2755 0.03064 0.02631 0.0249 0.0076 1.0000
14.250 1.2837 0.03213 0.02789 0.0256 0.0075 1.0000
14.500 1.2905 0.03381 0.02965 0.0261 0.0074 1.0000
14.750 1.2955 0.03573 0.03166 0.0265 0.0072 1.0000
15.000 1.2985 0.03793 0.03397 0.0267 0.0071 1.0000
15.250 1.2978 0.04061 0.03676 0.0267 0.0070 1.0000
15.500 1.2931 0.04388 0.04016 0.0264 0.0069 1.0000
15.750 1.2829 0.04795 0.04438 0.0256 0.0068 1.0000
16.000 1.2790 0.05145 0.04802 0.0246 0.0068 1.0000
16.250 1.2734 0.05530 0.05199 0.0233 0.0067 1.0000
16.500 1.2649 0.05973 0.05657 0.0216 0.0067 1.0000
16.750 1.2538 0.06478 0.06175 0.0193 0.0067 1.0000
17.000 1.2393 0.07068 0.06780 0.0164 0.0066 1.0000
17.250 1.2183 0.07812 0.07540 0.0124 0.0066 1.0000
17.500 1.1893 0.08755 0.08500 0.0071 0.0067 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 64-012A AIRFOIL (n64012a-il)