NACA 63012A AIRFOIL (n63012a-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 63012A AIRFOIL (n63012a-il) Reynolds number: 50,000 Max Cl/Cd: 28.9 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63012a-il-50000.txt Download as CSV file: xf-n63012a-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63012A 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.7108 0.08833 0.08144 -0.0176 1.0000 0.1504
-9.750 -0.7385 0.08012 0.07323 -0.0225 1.0000 0.1400
-9.500 -0.7898 0.07310 0.06598 -0.0245 1.0000 0.1305
-9.250 -0.7991 0.06811 0.06083 -0.0243 1.0000 0.1275
-9.000 -0.8380 0.06259 0.05450 -0.0224 1.0000 0.1197
-8.750 -0.8351 0.05851 0.05005 -0.0210 1.0000 0.1187
-8.500 -0.8299 0.05454 0.04571 -0.0195 1.0000 0.1185
-8.250 -0.8220 0.05079 0.04155 -0.0178 1.0000 0.1184
-8.000 -0.8099 0.04719 0.03755 -0.0162 1.0000 0.1182
-7.750 -0.7953 0.04384 0.03373 -0.0146 1.0000 0.1184
-7.500 -0.7782 0.04073 0.03019 -0.0131 1.0000 0.1197
-7.250 -0.7553 0.03789 0.02739 -0.0125 1.0000 0.1257
-7.000 -0.7346 0.03558 0.02466 -0.0111 1.0000 0.1317
-6.750 -0.7077 0.03296 0.02200 -0.0105 1.0000 0.1388
-6.500 -0.6815 0.03085 0.01982 -0.0098 1.0000 0.1522
-6.250 -0.6538 0.02880 0.01787 -0.0091 1.0000 0.1710
-6.000 -0.6306 0.02680 0.01611 -0.0078 1.0000 0.2021
-5.750 -0.6172 0.02442 0.01438 -0.0053 1.0000 0.2585
-5.500 -0.6256 0.02164 0.01359 0.0015 1.0000 0.4447
-5.250 -0.6252 0.02294 0.01550 0.0114 1.0000 0.6122
-5.000 -0.6070 0.02495 0.01740 0.0185 1.0000 0.6760
-4.750 -0.5790 0.02705 0.01930 0.0243 1.0000 0.7240
-4.500 -0.5300 0.02969 0.02159 0.0280 1.0000 0.7717
-4.250 -0.4384 0.03190 0.02316 0.0236 1.0000 0.8131
-4.000 -0.3768 0.03184 0.02267 0.0190 1.0000 0.8403
-3.750 -0.3409 0.03124 0.02185 0.0173 1.0000 0.8628
-3.500 -0.2942 0.03052 0.02089 0.0135 1.0000 0.8822
-3.250 -0.2501 0.02971 0.01986 0.0097 1.0000 0.9000
-3.000 -0.2103 0.02885 0.01885 0.0063 1.0000 0.9164
-2.750 -0.1710 0.02797 0.01785 0.0028 1.0000 0.9317
-2.500 -0.1314 0.02709 0.01685 -0.0010 1.0000 0.9462
-2.250 -0.0910 0.02620 0.01589 -0.0050 1.0000 0.9601
-2.000 -0.0487 0.02531 0.01493 -0.0096 1.0000 0.9737
-1.750 -0.0037 0.02439 0.01397 -0.0149 1.0000 0.9869
-1.500 0.0444 0.02345 0.01302 -0.0209 1.0000 0.9997
-1.250 0.0520 0.02317 0.01279 -0.0196 1.0000 1.0000
-1.000 0.0511 0.02308 0.01275 -0.0170 1.0000 1.0000
-0.750 0.0430 0.02312 0.01283 -0.0133 1.0000 1.0000
-0.500 0.0301 0.02323 0.01296 -0.0090 1.0000 1.0000
-0.250 0.0153 0.02331 0.01305 -0.0045 1.0000 1.0000
0.000 0.0000 0.02334 0.01308 0.0000 1.0000 1.0000
0.250 -0.0153 0.02331 0.01305 0.0045 1.0000 1.0000
0.500 -0.0301 0.02322 0.01295 0.0090 1.0000 1.0000
0.750 -0.0431 0.02312 0.01283 0.0133 1.0000 1.0000
1.000 -0.0511 0.02307 0.01274 0.0170 1.0000 1.0000
1.250 -0.0520 0.02317 0.01278 0.0196 1.0000 1.0000
1.500 -0.0444 0.02345 0.01301 0.0209 0.9997 1.0000
1.750 0.0037 0.02438 0.01397 0.0149 0.9869 1.0000
2.000 0.0487 0.02530 0.01492 0.0096 0.9738 1.0000
2.250 0.0910 0.02619 0.01588 0.0050 0.9602 1.0000
2.500 0.1315 0.02708 0.01684 0.0009 0.9462 1.0000
2.750 0.1711 0.02796 0.01784 -0.0028 0.9317 1.0000
3.000 0.2105 0.02884 0.01884 -0.0064 0.9164 1.0000
3.250 0.2504 0.02970 0.01985 -0.0098 0.9000 1.0000
3.500 0.2944 0.03051 0.02088 -0.0136 0.8823 1.0000
3.750 0.3413 0.03122 0.02183 -0.0174 0.8628 1.0000
4.000 0.3772 0.03183 0.02266 -0.0191 0.8404 1.0000
4.250 0.4388 0.03189 0.02314 -0.0237 0.8131 1.0000
4.500 0.5301 0.02967 0.02157 -0.0280 0.7718 1.0000
4.750 0.5789 0.02705 0.01929 -0.0243 0.7242 1.0000
5.000 0.6068 0.02495 0.01740 -0.0185 0.6762 1.0000
5.250 0.6249 0.02296 0.01552 -0.0114 0.6127 1.0000
5.500 0.6255 0.02164 0.01360 -0.0015 0.4462 1.0000
5.750 0.6172 0.02441 0.01437 0.0053 0.2588 1.0000
6.000 0.6306 0.02679 0.01610 0.0078 0.2022 1.0000
6.250 0.6537 0.02879 0.01787 0.0091 0.1710 1.0000
6.500 0.6815 0.03084 0.01982 0.0098 0.1522 1.0000
6.750 0.7077 0.03296 0.02199 0.0105 0.1388 1.0000
7.000 0.7346 0.03557 0.02465 0.0111 0.1317 1.0000
7.250 0.7553 0.03789 0.02739 0.0125 0.1257 1.0000
7.500 0.7782 0.04074 0.03019 0.0131 0.1197 1.0000
7.750 0.7954 0.04384 0.03373 0.0146 0.1184 1.0000
8.000 0.8101 0.04720 0.03756 0.0162 0.1182 1.0000
8.250 0.8222 0.05080 0.04157 0.0178 0.1184 1.0000
8.500 0.8301 0.05455 0.04573 0.0194 0.1185 1.0000
8.750 0.8354 0.05853 0.05007 0.0209 0.1187 1.0000
9.000 0.8383 0.06263 0.05454 0.0223 0.1197 1.0000
9.250 0.7244 0.05840 0.05151 0.0297 0.1300 1.0000
9.500 0.7132 0.06341 0.05662 0.0307 0.1319 1.0000
9.750 0.7121 0.06851 0.06178 0.0312 0.1336 1.0000
10.000 0.6866 0.09223 0.08529 0.0124 0.1563 1.0000
10.250 0.5829 0.11957 0.11234 -0.0190 0.3288 1.0000
10.500 0.6025 0.12470 0.11751 -0.0191 0.3240 1.0000
10.750 0.5884 0.12726 0.11998 -0.0201 0.3141 1.0000
11.000 0.6144 0.13308 0.12587 -0.0199 0.3088 1.0000
11.250 0.4529 0.12613 0.11924 -0.0042 0.3073 1.0000
11.500 0.4511 0.12857 0.12167 -0.0041 0.2961 1.0000
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Polar data table (+)
Polar graphs
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