NACA 64-012A AIRFOIL (n64012a-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 64-012A AIRFOIL (n64012a-il) Reynolds number: 100,000 Max Cl/Cd: 35.53 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n64012a-il-100000-n5.txt Download as CSV file: xf-n64012a-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.7268 0.10113 0.09592 -0.0123 1.0000 0.0288
-12.250 -0.7458 0.09231 0.08705 -0.0181 1.0000 0.0285
-12.000 -0.7676 0.08429 0.07894 -0.0235 1.0000 0.0283
-11.750 -0.7898 0.07741 0.07193 -0.0277 1.0000 0.0281
-11.500 -0.8112 0.07158 0.06595 -0.0306 1.0000 0.0280
-11.250 -0.8315 0.06663 0.06080 -0.0320 1.0000 0.0279
-11.000 -0.8500 0.06239 0.05635 -0.0321 1.0000 0.0279
-10.750 -0.8668 0.05869 0.05241 -0.0309 1.0000 0.0279
-10.500 -0.8813 0.05541 0.04885 -0.0285 1.0000 0.0281
-10.250 -0.8906 0.05215 0.04521 -0.0261 1.0000 0.0285
-9.750 -0.8905 0.04606 0.03850 -0.0222 1.0000 0.0297
-9.500 -0.8817 0.04367 0.03592 -0.0208 1.0000 0.0304
-9.250 -0.8714 0.04122 0.03320 -0.0193 1.0000 0.0309
-9.000 -0.8587 0.03882 0.03048 -0.0179 1.0000 0.0314
-8.750 -0.8436 0.03660 0.02797 -0.0165 1.0000 0.0320
-8.500 -0.8263 0.03453 0.02564 -0.0154 1.0000 0.0328
-8.250 -0.8071 0.03261 0.02348 -0.0143 1.0000 0.0337
-8.000 -0.7868 0.03091 0.02156 -0.0133 1.0000 0.0350
-7.750 -0.7660 0.02953 0.01993 -0.0123 1.0000 0.0369
-7.500 -0.7449 0.02794 0.01821 -0.0114 1.0000 0.0386
-7.250 -0.7249 0.02657 0.01685 -0.0105 1.0000 0.0402
-7.000 -0.7054 0.02540 0.01565 -0.0093 1.0000 0.0420
-6.750 -0.6868 0.02432 0.01452 -0.0079 1.0000 0.0441
-6.500 -0.6690 0.02340 0.01350 -0.0063 1.0000 0.0471
-6.250 -0.6541 0.02240 0.01252 -0.0044 1.0000 0.0507
-6.000 -0.6389 0.02161 0.01172 -0.0025 1.0000 0.0552
-5.750 -0.6238 0.02087 0.01090 -0.0003 1.0000 0.0598
-5.500 -0.6105 0.02005 0.01014 0.0018 1.0000 0.0673
-5.250 -0.5969 0.01932 0.00944 0.0040 1.0000 0.0776
-5.000 -0.5837 0.01859 0.00877 0.0062 1.0000 0.0941
-4.500 -0.5468 0.01602 0.00727 0.0073 0.9920 0.2739
-4.250 -0.5220 0.01469 0.00707 0.0070 0.9826 0.4919
-4.000 -0.4888 0.01451 0.00707 0.0060 0.9743 0.5736
-3.750 -0.4561 0.01449 0.00709 0.0052 0.9650 0.6206
-3.500 -0.4238 0.01458 0.00721 0.0047 0.9563 0.6624
-3.250 -0.3912 0.01479 0.00744 0.0045 0.9482 0.6989
-3.000 -0.3610 0.01483 0.00741 0.0043 0.9379 0.7205
-2.750 -0.3271 0.01481 0.00731 0.0033 0.9298 0.7322
-2.500 -0.2949 0.01475 0.00715 0.0025 0.9203 0.7419
-2.250 -0.2643 0.01466 0.00696 0.0019 0.9103 0.7518
-2.000 -0.2309 0.01458 0.00682 0.0009 0.9024 0.7594
-1.750 -0.2025 0.01450 0.00665 0.0008 0.8915 0.7681
-1.500 -0.1725 0.01444 0.00654 0.0005 0.8821 0.7761
-1.250 -0.1422 0.01437 0.00641 0.0001 0.8733 0.7842
-1.000 -0.1142 0.01433 0.00633 0.0001 0.8630 0.7928
-0.750 -0.0847 0.01429 0.00626 -0.0001 0.8545 0.8008
-0.500 -0.0570 0.01425 0.00619 0.0000 0.8449 0.8101
-0.250 -0.0283 0.01425 0.00619 0.0000 0.8357 0.8179
0.000 0.0000 0.01422 0.00613 0.0000 0.8275 0.8275
0.250 0.0283 0.01425 0.00619 0.0000 0.8179 0.8357
0.500 0.0570 0.01425 0.00619 0.0000 0.8101 0.8448
0.750 0.0847 0.01429 0.00626 0.0001 0.8008 0.8545
1.000 0.1143 0.01433 0.00633 -0.0001 0.7928 0.8630
1.250 0.1422 0.01437 0.00641 -0.0001 0.7842 0.8733
1.500 0.1725 0.01444 0.00654 -0.0005 0.7761 0.8821
1.750 0.2025 0.01450 0.00665 -0.0008 0.7681 0.8915
2.000 0.2309 0.01458 0.00682 -0.0009 0.7594 0.9024
2.250 0.2643 0.01466 0.00696 -0.0019 0.7518 0.9103
2.500 0.2949 0.01475 0.00715 -0.0025 0.7419 0.9203
2.750 0.3271 0.01481 0.00731 -0.0033 0.7322 0.9298
3.000 0.3610 0.01483 0.00741 -0.0043 0.7205 0.9379
3.250 0.3912 0.01479 0.00744 -0.0045 0.6989 0.9482
3.500 0.4238 0.01458 0.00721 -0.0047 0.6624 0.9563
3.750 0.4561 0.01449 0.00709 -0.0052 0.6205 0.9650
4.000 0.4888 0.01451 0.00707 -0.0060 0.5737 0.9743
4.250 0.5220 0.01469 0.00707 -0.0070 0.4920 0.9826
4.500 0.5468 0.01602 0.00727 -0.0073 0.2740 0.9920
5.000 0.5837 0.01859 0.00877 -0.0062 0.0941 1.0000
5.250 0.5969 0.01932 0.00944 -0.0040 0.0776 1.0000
5.500 0.6105 0.02005 0.01014 -0.0018 0.0673 1.0000
5.750 0.6238 0.02086 0.01089 0.0003 0.0598 1.0000
6.000 0.6389 0.02160 0.01172 0.0025 0.0552 1.0000
6.250 0.6541 0.02240 0.01252 0.0044 0.0507 1.0000
6.500 0.6690 0.02340 0.01350 0.0063 0.0471 1.0000
6.750 0.6868 0.02431 0.01452 0.0079 0.0441 1.0000
7.000 0.7054 0.02540 0.01565 0.0093 0.0420 1.0000
7.250 0.7249 0.02657 0.01685 0.0105 0.0402 1.0000
7.500 0.7449 0.02794 0.01821 0.0114 0.0386 1.0000
7.750 0.7660 0.02953 0.01993 0.0123 0.0369 1.0000
8.000 0.7869 0.03091 0.02156 0.0133 0.0350 1.0000
8.250 0.8072 0.03261 0.02348 0.0143 0.0337 1.0000
8.500 0.8264 0.03453 0.02565 0.0153 0.0328 1.0000
8.750 0.8437 0.03660 0.02798 0.0165 0.0320 1.0000
9.000 0.8588 0.03883 0.03049 0.0178 0.0314 1.0000
9.250 0.8715 0.04123 0.03320 0.0193 0.0309 1.0000
9.500 0.8819 0.04368 0.03592 0.0208 0.0304 1.0000
9.750 0.8907 0.04606 0.03850 0.0222 0.0297 1.0000
10.250 0.8908 0.05217 0.04523 0.0260 0.0285 1.0000
10.500 0.8815 0.05543 0.04887 0.0284 0.0281 1.0000
10.750 0.8670 0.05872 0.05245 0.0308 0.0279 1.0000
11.000 0.8503 0.06243 0.05639 0.0320 0.0279 1.0000
11.250 0.8318 0.06669 0.06087 0.0319 0.0279 1.0000
11.500 0.8116 0.07166 0.06602 0.0304 0.0280 1.0000
11.750 0.7903 0.07749 0.07202 0.0275 0.0281 1.0000
12.000 0.7681 0.08440 0.07905 0.0233 0.0283 1.0000
12.250 0.7463 0.09246 0.08720 0.0178 0.0285 1.0000
12.500 0.7273 0.10136 0.09615 0.0120 0.0288 1.0000
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Polar data table (+)
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