NACA 63A010 AIRFOIL (n63010a-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 63A010 AIRFOIL (n63010a-il) Reynolds number: 100,000 Max Cl/Cd: 34.85 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n63010a-il-100000-n5.txt Download as CSV file: xf-n63010a-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.7288 0.08578 0.08081 -0.0110 1.0000 0.0257
-10.500 -0.7457 0.07845 0.07340 -0.0165 1.0000 0.0255
-10.250 -0.7640 0.07232 0.06717 -0.0206 1.0000 0.0253
-10.000 -0.7821 0.06732 0.06202 -0.0227 1.0000 0.0252
-9.750 -0.7988 0.06311 0.05764 -0.0227 1.0000 0.0251
-9.500 -0.8113 0.05904 0.05333 -0.0218 1.0000 0.0250
-9.250 -0.8185 0.05488 0.04887 -0.0207 1.0000 0.0250
-9.000 -0.8210 0.05084 0.04445 -0.0194 1.0000 0.0251
-8.750 -0.8190 0.04699 0.04018 -0.0179 1.0000 0.0253
-8.500 -0.8125 0.04365 0.03634 -0.0162 1.0000 0.0260
-8.250 -0.8035 0.04048 0.03270 -0.0147 1.0000 0.0270
-8.000 -0.7904 0.03759 0.02958 -0.0137 1.0000 0.0281
-7.750 -0.7740 0.03512 0.02677 -0.0126 1.0000 0.0288
-7.500 -0.7556 0.03276 0.02409 -0.0115 1.0000 0.0294
-7.250 -0.7354 0.03064 0.02168 -0.0105 1.0000 0.0302
-7.000 -0.7138 0.02871 0.01951 -0.0096 1.0000 0.0312
-6.750 -0.6915 0.02701 0.01760 -0.0087 1.0000 0.0326
-6.500 -0.6688 0.02571 0.01609 -0.0078 1.0000 0.0351
-6.250 -0.6459 0.02447 0.01464 -0.0069 1.0000 0.0370
-6.000 -0.6257 0.02278 0.01297 -0.0058 1.0000 0.0388
-5.750 -0.6056 0.02163 0.01181 -0.0046 1.0000 0.0411
-5.500 -0.5858 0.02062 0.01073 -0.0033 1.0000 0.0442
-5.250 -0.5658 0.01977 0.00975 -0.0020 1.0000 0.0483
-5.000 -0.5473 0.01883 0.00885 -0.0006 1.0000 0.0560
-4.750 -0.5282 0.01797 0.00796 0.0008 1.0000 0.0663
-4.500 -0.5094 0.01711 0.00719 0.0022 1.0000 0.0882
-4.250 -0.4928 0.01591 0.00643 0.0036 1.0000 0.1523
-4.000 -0.4810 0.01421 0.00579 0.0055 1.0000 0.3417
-3.750 -0.4657 0.01336 0.00564 0.0077 1.0000 0.4921
-3.500 -0.4473 0.01303 0.00559 0.0097 1.0000 0.5786
-3.250 -0.4282 0.01289 0.00562 0.0117 1.0000 0.6391
-3.000 -0.4094 0.01286 0.00569 0.0139 1.0000 0.6896
-2.750 -0.3901 0.01288 0.00578 0.0159 0.9997 0.7293
-2.500 -0.3548 0.01293 0.00577 0.0147 0.9907 0.7590
-2.250 -0.3183 0.01294 0.00571 0.0130 0.9819 0.7781
-2.000 -0.2807 0.01294 0.00562 0.0110 0.9735 0.7931
-1.750 -0.2437 0.01292 0.00551 0.0091 0.9644 0.8069
-1.500 -0.2084 0.01290 0.00542 0.0076 0.9544 0.8203
-1.250 -0.1727 0.01288 0.00536 0.0060 0.9449 0.8333
-1.000 -0.1363 0.01287 0.00531 0.0044 0.9356 0.8451
-0.750 -0.1024 0.01286 0.00528 0.0033 0.9246 0.8570
-0.500 -0.0681 0.01285 0.00525 0.0022 0.9138 0.8688
-0.250 -0.0337 0.01284 0.00523 0.0010 0.9033 0.8805
0.000 0.0000 0.01284 0.00521 0.0000 0.8925 0.8924
0.250 0.0337 0.01284 0.00522 -0.0010 0.8805 0.9033
0.500 0.0681 0.01285 0.00525 -0.0022 0.8688 0.9138
0.750 0.1024 0.01286 0.00528 -0.0033 0.8570 0.9246
1.000 0.1363 0.01287 0.00531 -0.0044 0.8451 0.9356
1.250 0.1726 0.01288 0.00536 -0.0060 0.8333 0.9449
1.500 0.2084 0.01290 0.00542 -0.0076 0.8203 0.9545
1.750 0.2437 0.01292 0.00551 -0.0091 0.8069 0.9644
2.000 0.2807 0.01294 0.00562 -0.0110 0.7931 0.9735
2.250 0.3183 0.01294 0.00571 -0.0130 0.7781 0.9819
2.500 0.3547 0.01293 0.00577 -0.0147 0.7590 0.9907
2.750 0.3900 0.01288 0.00578 -0.0159 0.7293 0.9997
3.000 0.4093 0.01286 0.00569 -0.0139 0.6897 1.0000
3.250 0.4281 0.01289 0.00562 -0.0117 0.6392 1.0000
3.500 0.4472 0.01303 0.00559 -0.0097 0.5787 1.0000
3.750 0.4656 0.01336 0.00564 -0.0077 0.4924 1.0000
4.000 0.4809 0.01420 0.00579 -0.0055 0.3421 1.0000
4.250 0.4928 0.01591 0.00643 -0.0036 0.1524 1.0000
4.500 0.5093 0.01711 0.00719 -0.0021 0.0882 1.0000
4.750 0.5282 0.01797 0.00796 -0.0008 0.0663 1.0000
5.000 0.5472 0.01882 0.00885 0.0006 0.0560 1.0000
5.250 0.5658 0.01977 0.00975 0.0020 0.0484 1.0000
5.500 0.5858 0.02062 0.01073 0.0033 0.0442 1.0000
5.750 0.6056 0.02163 0.01181 0.0046 0.0411 1.0000
6.000 0.6257 0.02278 0.01297 0.0058 0.0388 1.0000
6.250 0.6459 0.02447 0.01464 0.0069 0.0370 1.0000
6.500 0.6688 0.02571 0.01609 0.0078 0.0351 1.0000
6.750 0.6915 0.02701 0.01760 0.0087 0.0326 1.0000
7.000 0.7139 0.02871 0.01951 0.0096 0.0312 1.0000
7.250 0.7354 0.03064 0.02168 0.0105 0.0302 1.0000
7.500 0.7557 0.03276 0.02410 0.0115 0.0294 1.0000
7.750 0.7741 0.03512 0.02677 0.0126 0.0288 1.0000
8.000 0.7904 0.03759 0.02959 0.0137 0.0281 1.0000
8.250 0.8036 0.04048 0.03269 0.0147 0.0270 1.0000
8.500 0.8125 0.04365 0.03634 0.0162 0.0260 1.0000
8.750 0.8191 0.04699 0.04018 0.0178 0.0253 1.0000
9.000 0.8211 0.05084 0.04446 0.0193 0.0251 1.0000
9.250 0.8187 0.05489 0.04888 0.0207 0.0250 1.0000
9.500 0.8115 0.05905 0.05334 0.0217 0.0250 1.0000
9.750 0.7991 0.06313 0.05766 0.0226 0.0251 1.0000
10.000 0.7824 0.06734 0.06205 0.0226 0.0252 1.0000
10.250 0.7643 0.07237 0.06722 0.0205 0.0253 1.0000
10.500 0.7461 0.07851 0.07346 0.0164 0.0255 1.0000
10.750 0.7293 0.08586 0.08088 0.0108 0.0257 1.0000
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