NACA 63A010 AIRFOIL (n63010a-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 63A010 AIRFOIL (n63010a-il) Reynolds number: 200,000 Max Cl/Cd: 39.2 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n63010a-il-200000-n5.txt Download as CSV file: xf-n63010a-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.7834 0.07830 0.07478 -0.0101 1.0000 0.0137
-11.000 -0.8266 0.06531 0.06151 -0.0198 1.0000 0.0133
-10.750 -0.8488 0.05928 0.05526 -0.0229 1.0000 0.0133
-10.500 -0.8675 0.05474 0.05049 -0.0234 1.0000 0.0133
-10.250 -0.8832 0.05097 0.04649 -0.0221 1.0000 0.0133
-10.000 -0.8920 0.04742 0.04269 -0.0205 1.0000 0.0135
-9.750 -0.8929 0.04433 0.03934 -0.0190 1.0000 0.0137
-9.500 -0.8860 0.04221 0.03701 -0.0179 1.0000 0.0141
-9.250 -0.8754 0.04043 0.03505 -0.0169 1.0000 0.0147
-9.000 -0.8641 0.03836 0.03276 -0.0158 1.0000 0.0154
-8.750 -0.8532 0.03565 0.02969 -0.0144 1.0000 0.0161
-8.500 -0.8408 0.03272 0.02634 -0.0128 1.0000 0.0166
-8.250 -0.8252 0.03015 0.02339 -0.0115 1.0000 0.0170
-8.000 -0.8070 0.02800 0.02088 -0.0103 1.0000 0.0175
-7.750 -0.7869 0.02623 0.01882 -0.0093 1.0000 0.0181
-7.500 -0.7659 0.02483 0.01716 -0.0084 1.0000 0.0189
-7.250 -0.7458 0.02306 0.01528 -0.0075 1.0000 0.0200
-7.000 -0.7243 0.02185 0.01398 -0.0067 1.0000 0.0208
-6.750 -0.7026 0.02072 0.01275 -0.0058 1.0000 0.0215
-6.500 -0.6811 0.01968 0.01163 -0.0049 1.0000 0.0224
-6.250 -0.6599 0.01873 0.01060 -0.0038 1.0000 0.0234
-6.000 -0.6390 0.01786 0.00963 -0.0027 1.0000 0.0245
-5.750 -0.6178 0.01717 0.00886 -0.0016 1.0000 0.0259
-5.500 -0.5984 0.01632 0.00797 -0.0003 1.0000 0.0280
-5.250 -0.5786 0.01561 0.00724 0.0009 1.0000 0.0304
-5.000 -0.5581 0.01502 0.00660 0.0021 1.0000 0.0332
-4.750 -0.5371 0.01453 0.00600 0.0033 1.0000 0.0367
-4.500 -0.5171 0.01394 0.00544 0.0045 1.0000 0.0441
-4.250 -0.4951 0.01346 0.00498 0.0053 0.9993 0.0570
-4.000 -0.4614 0.01271 0.00447 0.0034 0.9923 0.1025
-3.750 -0.4306 0.01145 0.00389 0.0016 0.9847 0.2447
-3.500 -0.4027 0.01027 0.00354 0.0006 0.9753 0.4257
-3.250 -0.3723 0.00982 0.00339 -0.0002 0.9652 0.5137
-3.000 -0.3417 0.00953 0.00332 -0.0008 0.9549 0.5777
-2.750 -0.3109 0.00934 0.00326 -0.0013 0.9447 0.6310
-2.500 -0.2799 0.00922 0.00325 -0.0018 0.9346 0.6730
-2.250 -0.2502 0.00914 0.00320 -0.0019 0.9227 0.6984
-2.000 -0.2207 0.00908 0.00310 -0.0021 0.9106 0.7174
-1.750 -0.1916 0.00902 0.00302 -0.0021 0.8985 0.7320
-1.500 -0.1630 0.00896 0.00294 -0.0021 0.8862 0.7438
-1.250 -0.1350 0.00891 0.00287 -0.0019 0.8741 0.7554
-1.000 -0.1074 0.00888 0.00279 -0.0016 0.8619 0.7669
-0.750 -0.0804 0.00885 0.00274 -0.0013 0.8496 0.7786
-0.500 -0.0536 0.00883 0.00271 -0.0009 0.8372 0.7906
-0.250 -0.0268 0.00882 0.00269 -0.0004 0.8252 0.8021
0.000 0.0000 0.00881 0.00269 0.0000 0.8135 0.8135
0.250 0.0268 0.00882 0.00269 0.0004 0.8021 0.8252
0.500 0.0536 0.00883 0.00271 0.0009 0.7906 0.8372
0.750 0.0804 0.00885 0.00274 0.0013 0.7786 0.8496
1.000 0.1075 0.00888 0.00279 0.0016 0.7669 0.8619
1.250 0.1350 0.00891 0.00287 0.0019 0.7554 0.8741
1.500 0.1630 0.00896 0.00294 0.0021 0.7438 0.8862
1.750 0.1916 0.00902 0.00302 0.0021 0.7319 0.8985
2.000 0.2207 0.00908 0.00310 0.0021 0.7174 0.9106
2.250 0.2502 0.00914 0.00320 0.0019 0.6984 0.9227
2.500 0.2798 0.00922 0.00325 0.0018 0.6730 0.9346
2.750 0.3108 0.00934 0.00326 0.0013 0.6310 0.9447
3.000 0.3417 0.00953 0.00332 0.0008 0.5777 0.9549
3.250 0.3723 0.00981 0.00339 0.0002 0.5138 0.9652
3.500 0.4026 0.01027 0.00353 -0.0005 0.4258 0.9753
3.750 0.4306 0.01144 0.00389 -0.0016 0.2451 0.9847
4.000 0.4613 0.01271 0.00447 -0.0034 0.1026 0.9923
4.250 0.4951 0.01346 0.00498 -0.0053 0.0569 0.9993
4.500 0.5170 0.01394 0.00544 -0.0045 0.0441 1.0000
4.750 0.5370 0.01453 0.00600 -0.0033 0.0367 1.0000
5.000 0.5580 0.01502 0.00660 -0.0021 0.0332 1.0000
5.250 0.5785 0.01560 0.00724 -0.0009 0.0304 1.0000
5.500 0.5983 0.01632 0.00797 0.0003 0.0280 1.0000
5.750 0.6178 0.01717 0.00886 0.0016 0.0259 1.0000
6.000 0.6390 0.01786 0.00963 0.0027 0.0245 1.0000
6.250 0.6599 0.01873 0.01060 0.0038 0.0234 1.0000
6.500 0.6811 0.01968 0.01163 0.0049 0.0224 1.0000
6.750 0.7026 0.02072 0.01275 0.0058 0.0215 1.0000
7.000 0.7243 0.02185 0.01398 0.0067 0.0208 1.0000
7.250 0.7459 0.02306 0.01528 0.0075 0.0200 1.0000
7.500 0.7659 0.02483 0.01715 0.0084 0.0189 1.0000
7.750 0.7870 0.02623 0.01882 0.0093 0.0181 1.0000
8.000 0.8071 0.02800 0.02089 0.0103 0.0175 1.0000
8.250 0.8253 0.03016 0.02339 0.0115 0.0170 1.0000
8.500 0.8409 0.03272 0.02635 0.0128 0.0166 1.0000
8.750 0.8533 0.03565 0.02969 0.0143 0.0161 1.0000
9.000 0.8642 0.03838 0.03278 0.0158 0.0154 1.0000
9.250 0.8754 0.04046 0.03509 0.0169 0.0147 1.0000
9.500 0.8861 0.04223 0.03704 0.0179 0.0141 1.0000
9.750 0.8930 0.04435 0.03937 0.0190 0.0137 1.0000
10.000 0.8922 0.04745 0.04272 0.0204 0.0135 1.0000
10.250 0.8835 0.05099 0.04652 0.0220 0.0133 1.0000
10.500 0.8678 0.05477 0.05053 0.0233 0.0133 1.0000
10.750 0.8491 0.05934 0.05532 0.0227 0.0133 1.0000
11.000 0.8270 0.06537 0.06158 0.0196 0.0133 1.0000
11.250 0.7814 0.07918 0.07568 0.0093 0.0138 1.0000
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Polar data table (+)
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