NACA 63A010 AIRFOIL (n63010a-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 63A010 AIRFOIL (n63010a-il) Reynolds number: 1,000,000 Max Cl/Cd: 59.73 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63010a-il-1000000.txt Download as CSV file: xf-n63010a-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63A010 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.8763 0.06182 0.06001 -0.0175 1.0000 0.0082
-11.250 -0.9108 0.05361 0.05159 -0.0222 1.0000 0.0081
-11.000 -0.9484 0.04667 0.04437 -0.0227 1.0000 0.0080
-10.750 -0.9842 0.04040 0.03771 -0.0194 1.0000 0.0080
-10.500 -1.0000 0.03490 0.03176 -0.0166 1.0000 0.0081
-10.250 -0.9999 0.03110 0.02760 -0.0144 1.0000 0.0083
-10.000 -0.9912 0.02835 0.02453 -0.0127 1.0000 0.0085
-9.750 -0.9768 0.02643 0.02237 -0.0114 1.0000 0.0087
-9.500 -0.9582 0.02523 0.02099 -0.0105 1.0000 0.0090
-9.250 -0.9446 0.02271 0.01813 -0.0090 1.0000 0.0093
-9.000 -0.9293 0.02043 0.01567 -0.0078 1.0000 0.0099
-8.750 -0.9073 0.01975 0.01493 -0.0072 1.0000 0.0102
-8.500 -0.8854 0.01891 0.01400 -0.0065 1.0000 0.0106
-8.250 -0.8634 0.01800 0.01299 -0.0058 1.0000 0.0109
-8.000 -0.8414 0.01709 0.01199 -0.0050 1.0000 0.0112
-7.750 -0.8194 0.01620 0.01101 -0.0042 1.0000 0.0116
-7.500 -0.7973 0.01541 0.01013 -0.0033 1.0000 0.0120
-7.250 -0.7750 0.01474 0.00939 -0.0025 1.0000 0.0124
-7.000 -0.7520 0.01430 0.00890 -0.0017 1.0000 0.0127
-6.750 -0.7339 0.01315 0.00764 0.0000 1.0000 0.0133
-6.500 -0.7153 0.01227 0.00669 0.0015 1.0000 0.0140
-6.250 -0.6945 0.01175 0.00614 0.0028 1.0000 0.0146
-6.000 -0.6736 0.01131 0.00568 0.0040 1.0000 0.0151
-5.750 -0.6518 0.01091 0.00524 0.0050 0.9998 0.0157
-5.500 -0.6173 0.01046 0.00475 0.0033 0.9972 0.0166
-5.250 -0.5821 0.01009 0.00435 0.0015 0.9946 0.0175
-5.000 -0.5492 0.00956 0.00375 0.0002 0.9905 0.0194
-4.750 -0.5148 0.00919 0.00337 -0.0015 0.9864 0.0220
-4.500 -0.4804 0.00893 0.00310 -0.0030 0.9815 0.0245
-4.250 -0.4491 0.00849 0.00271 -0.0039 0.9728 0.0373
-4.000 -0.4185 0.00813 0.00244 -0.0047 0.9625 0.0599
-3.750 -0.3908 0.00766 0.00217 -0.0049 0.9491 0.1098
-3.500 -0.3665 0.00704 0.00188 -0.0045 0.9327 0.1983
-3.250 -0.3439 0.00636 0.00160 -0.0037 0.9152 0.3112
-3.000 -0.3208 0.00582 0.00138 -0.0030 0.8983 0.4143
-2.750 -0.2964 0.00548 0.00125 -0.0024 0.8822 0.4882
-2.500 -0.2711 0.00526 0.00117 -0.0019 0.8670 0.5456
-2.250 -0.2446 0.00517 0.00110 -0.0016 0.8527 0.5729
-2.000 -0.2180 0.00509 0.00104 -0.0013 0.8391 0.5982
-1.750 -0.1915 0.00500 0.00100 -0.0010 0.8261 0.6326
-1.500 -0.1647 0.00493 0.00098 -0.0007 0.8138 0.6595
-1.000 -0.1102 0.00488 0.00093 -0.0004 0.7907 0.6961
-0.750 -0.0829 0.00487 0.00092 -0.0002 0.7797 0.7111
-0.500 -0.0554 0.00486 0.00091 -0.0001 0.7687 0.7251
-0.250 -0.0276 0.00486 0.00090 -0.0001 0.7580 0.7370
0.000 0.0000 0.00485 0.00090 0.0000 0.7476 0.7476
0.250 0.0276 0.00486 0.00090 0.0001 0.7370 0.7581
0.500 0.0553 0.00486 0.00091 0.0001 0.7251 0.7687
0.750 0.0829 0.00487 0.00092 0.0002 0.7111 0.7797
1.000 0.1102 0.00488 0.00093 0.0004 0.6962 0.7907
1.250 0.1374 0.00491 0.00095 0.0006 0.6794 0.8020
1.500 0.1647 0.00493 0.00098 0.0007 0.6595 0.8138
1.750 0.1915 0.00500 0.00100 0.0010 0.6324 0.8261
2.000 0.2180 0.00509 0.00104 0.0013 0.5982 0.8391
2.250 0.2446 0.00517 0.00110 0.0016 0.5728 0.8527
2.500 0.2711 0.00526 0.00117 0.0019 0.5445 0.8670
2.750 0.2964 0.00548 0.00125 0.0024 0.4888 0.8822
3.000 0.3209 0.00582 0.00138 0.0030 0.4144 0.8983
3.250 0.3440 0.00635 0.00159 0.0037 0.3123 0.9151
3.500 0.3665 0.00704 0.00188 0.0045 0.1975 0.9326
3.750 0.3908 0.00766 0.00217 0.0049 0.1094 0.9491
4.000 0.4186 0.00813 0.00244 0.0047 0.0600 0.9625
4.250 0.4491 0.00849 0.00270 0.0039 0.0374 0.9728
4.500 0.4803 0.00893 0.00310 0.0031 0.0245 0.9815
4.750 0.5147 0.00919 0.00337 0.0015 0.0220 0.9864
5.000 0.5491 0.00956 0.00375 -0.0001 0.0194 0.9905
5.250 0.5821 0.01009 0.00434 -0.0015 0.0175 0.9946
5.500 0.6172 0.01046 0.00475 -0.0033 0.0166 0.9973
5.750 0.6517 0.01091 0.00523 -0.0050 0.0157 0.9999
6.000 0.6733 0.01131 0.00568 -0.0039 0.0151 1.0000
6.250 0.6943 0.01175 0.00614 -0.0027 0.0145 1.0000
6.500 0.7152 0.01227 0.00670 -0.0015 0.0140 1.0000
6.750 0.7338 0.01316 0.00764 0.0001 0.0133 1.0000
7.000 0.7518 0.01431 0.00890 0.0017 0.0127 1.0000
7.250 0.7750 0.01473 0.00939 0.0025 0.0124 1.0000
7.500 0.7973 0.01541 0.01013 0.0033 0.0120 1.0000
7.750 0.8194 0.01621 0.01102 0.0042 0.0116 1.0000
8.000 0.8415 0.01710 0.01199 0.0050 0.0112 1.0000
8.250 0.8635 0.01801 0.01301 0.0058 0.0109 1.0000
8.500 0.8854 0.01893 0.01402 0.0065 0.0106 1.0000
8.750 0.9074 0.01976 0.01495 0.0072 0.0102 1.0000
9.000 0.9294 0.02045 0.01569 0.0077 0.0099 1.0000
9.250 0.9448 0.02272 0.01815 0.0090 0.0093 1.0000
9.500 0.9584 0.02524 0.02101 0.0105 0.0090 1.0000
9.750 0.9769 0.02645 0.02239 0.0114 0.0087 1.0000
10.000 0.9915 0.02836 0.02454 0.0127 0.0085 1.0000
10.250 1.0001 0.03112 0.02761 0.0143 0.0083 1.0000
10.500 1.0001 0.03495 0.03182 0.0165 0.0081 1.0000
10.750 0.9844 0.04043 0.03775 0.0193 0.0080 1.0000
11.000 0.9490 0.04671 0.04441 0.0226 0.0080 1.0000
11.250 0.9109 0.05372 0.05171 0.0220 0.0081 1.0000
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Polar data table (+)
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