NACA 64-012A AIRFOIL (n64012a-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 64-012A AIRFOIL (n64012a-il) Reynolds number: 50,000 Max Cl/Cd: 27.48 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64012a-il-50000.txt Download as CSV file: xf-n64012a-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5699 0.10802 0.10086 0.0156 1.0000 0.3613
-9.250 -0.7629 0.07345 0.06642 -0.0243 1.0000 0.1445
-9.000 -0.7976 0.06753 0.06009 -0.0241 1.0000 0.1343
-8.750 -0.7971 0.06288 0.05525 -0.0233 1.0000 0.1313
-8.500 -0.8014 0.05833 0.05030 -0.0218 1.0000 0.1278
-8.250 -0.8145 0.05473 0.04571 -0.0186 1.0000 0.1222
-8.000 -0.8031 0.05086 0.04151 -0.0171 1.0000 0.1217
-7.750 -0.7894 0.04719 0.03751 -0.0156 1.0000 0.1210
-7.500 -0.7738 0.04384 0.03378 -0.0140 1.0000 0.1204
-7.250 -0.7566 0.04085 0.03036 -0.0124 1.0000 0.1209
-7.000 -0.7372 0.03799 0.02721 -0.0111 1.0000 0.1243
-6.750 -0.7142 0.03560 0.02475 -0.0102 1.0000 0.1303
-6.500 -0.6898 0.03339 0.02211 -0.0090 1.0000 0.1353
-6.250 -0.6609 0.03108 0.01987 -0.0087 1.0000 0.1452
-6.000 -0.6315 0.02910 0.01793 -0.0083 1.0000 0.1603
-5.750 -0.6035 0.02720 0.01619 -0.0075 1.0000 0.1831
-5.500 -0.5855 0.02526 0.01458 -0.0054 1.0000 0.2191
-5.250 -0.5835 0.02258 0.01291 -0.0015 1.0000 0.2998
-5.000 -0.6009 0.02187 0.01440 0.0108 1.0000 0.6197
-4.750 -0.5839 0.02500 0.01744 0.0211 1.0000 0.7174
-4.500 -0.4156 0.03320 0.02441 0.0170 1.0000 0.8314
-4.250 -0.3441 0.03320 0.02390 0.0113 1.0000 0.8678
-4.000 -0.2969 0.03241 0.02282 0.0075 1.0000 0.8903
-3.750 -0.2622 0.03154 0.02174 0.0052 1.0000 0.9087
-3.500 -0.2255 0.03061 0.02061 0.0022 1.0000 0.9249
-3.250 -0.1849 0.02960 0.01943 -0.0016 1.0000 0.9393
-3.000 -0.1436 0.02855 0.01823 -0.0057 1.0000 0.9523
-2.750 -0.1039 0.02753 0.01711 -0.0096 1.0000 0.9644
-2.500 -0.0654 0.02656 0.01606 -0.0135 1.0000 0.9757
-2.250 -0.0267 0.02562 0.01505 -0.0175 1.0000 0.9865
-2.000 0.0145 0.02467 0.01406 -0.0221 1.0000 0.9967
-1.750 0.0367 0.02411 0.01351 -0.0232 1.0000 1.0000
-1.500 0.0460 0.02382 0.01327 -0.0220 1.0000 1.0000
-1.250 0.0495 0.02368 0.01318 -0.0199 1.0000 1.0000
-1.000 0.0460 0.02369 0.01324 -0.0167 1.0000 1.0000
-0.750 0.0372 0.02380 0.01338 -0.0128 1.0000 1.0000
-0.500 0.0255 0.02392 0.01352 -0.0085 1.0000 1.0000
-0.250 0.0129 0.02401 0.01361 -0.0043 1.0000 1.0000
0.000 0.0000 0.02404 0.01365 0.0000 1.0000 1.0000
0.250 -0.0129 0.02401 0.01361 0.0043 1.0000 1.0000
0.500 -0.0255 0.02392 0.01352 0.0086 1.0000 1.0000
0.750 -0.0372 0.02380 0.01338 0.0128 1.0000 1.0000
1.000 -0.0460 0.02369 0.01324 0.0167 1.0000 1.0000
1.250 -0.0495 0.02367 0.01318 0.0199 1.0000 1.0000
1.500 -0.0460 0.02381 0.01326 0.0220 1.0000 1.0000
1.750 -0.0367 0.02410 0.01350 0.0232 1.0000 1.0000
2.000 -0.0144 0.02466 0.01405 0.0221 0.9967 1.0000
2.250 0.0267 0.02561 0.01504 0.0175 0.9865 1.0000
2.500 0.0655 0.02655 0.01605 0.0135 0.9757 1.0000
2.750 0.1040 0.02752 0.01710 0.0096 0.9644 1.0000
3.000 0.1438 0.02854 0.01822 0.0056 0.9524 1.0000
3.250 0.1851 0.02959 0.01942 0.0015 0.9394 1.0000
3.500 0.2258 0.03060 0.02060 -0.0023 0.9249 1.0000
3.750 0.2625 0.03153 0.02173 -0.0052 0.9088 1.0000
4.000 0.2973 0.03240 0.02280 -0.0075 0.8903 1.0000
4.250 0.3445 0.03319 0.02389 -0.0114 0.8678 1.0000
4.500 0.4159 0.03318 0.02439 -0.0170 0.8313 1.0000
4.750 0.5836 0.02501 0.01746 -0.0211 0.7176 1.0000
5.000 0.6009 0.02187 0.01440 -0.0108 0.6198 1.0000
5.250 0.5834 0.02258 0.01291 0.0015 0.3000 1.0000
5.500 0.5855 0.02525 0.01457 0.0055 0.2192 1.0000
5.750 0.6035 0.02720 0.01619 0.0075 0.1831 1.0000
6.000 0.6314 0.02910 0.01793 0.0083 0.1603 1.0000
6.250 0.6609 0.03108 0.01987 0.0087 0.1452 1.0000
6.500 0.6898 0.03339 0.02211 0.0090 0.1353 1.0000
6.750 0.7141 0.03560 0.02475 0.0102 0.1303 1.0000
7.000 0.7372 0.03799 0.02721 0.0111 0.1243 1.0000
7.250 0.7566 0.04085 0.03036 0.0124 0.1209 1.0000
7.500 0.7738 0.04384 0.03379 0.0140 0.1204 1.0000
7.750 0.7894 0.04719 0.03751 0.0156 0.1210 1.0000
8.000 0.8031 0.05087 0.04152 0.0171 0.1217 1.0000
8.250 0.8145 0.05474 0.04572 0.0185 0.1222 1.0000
8.500 0.8016 0.05834 0.05032 0.0218 0.1278 1.0000
8.750 0.7973 0.06290 0.05527 0.0232 0.1313 1.0000
9.000 0.7979 0.06756 0.06011 0.0241 0.1344 1.0000
9.250 0.7631 0.07350 0.06647 0.0242 0.1445 1.0000
9.500 0.6527 0.06888 0.06224 0.0312 0.1485 1.0000
9.750 0.5807 0.07881 0.07219 0.0254 0.1570 1.0000
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Polar data table (+)
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