NACA 63012A AIRFOIL (n63012a-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: NACA 63012A AIRFOIL (n63012a-il) Reynolds number: 100,000 Max Cl/Cd: 41.25 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63012a-il-100000.txt Download as CSV file: xf-n63012a-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.6451 0.10393 0.09898 0.0008 1.0000 0.1695
-10.000 -0.7116 0.06091 0.05593 -0.0347 1.0000 0.0882
-9.750 -0.7390 0.05556 0.05044 -0.0336 1.0000 0.0844
-9.500 -0.7820 0.04933 0.04370 -0.0300 1.0000 0.0758
-9.250 -0.8475 0.05545 0.04925 -0.0237 1.0000 0.0739
-9.000 -0.8659 0.05060 0.04324 -0.0191 1.0000 0.0659
-8.750 -0.8555 0.04768 0.03993 -0.0173 1.0000 0.0656
-8.500 -0.8453 0.04389 0.03581 -0.0156 1.0000 0.0658
-8.250 -0.8314 0.04062 0.03218 -0.0140 1.0000 0.0659
-8.000 -0.8145 0.03794 0.02913 -0.0124 1.0000 0.0656
-7.750 -0.7959 0.03474 0.02560 -0.0112 1.0000 0.0659
-7.500 -0.7740 0.03158 0.02225 -0.0104 1.0000 0.0670
-7.250 -0.7514 0.02947 0.02008 -0.0097 1.0000 0.0698
-7.000 -0.7290 0.02793 0.01843 -0.0087 1.0000 0.0740
-6.750 -0.7053 0.02642 0.01669 -0.0076 1.0000 0.0775
-6.500 -0.6811 0.02436 0.01475 -0.0069 1.0000 0.0822
-6.250 -0.6602 0.02319 0.01353 -0.0056 1.0000 0.0900
-6.000 -0.6425 0.02176 0.01228 -0.0039 1.0000 0.1002
-5.750 -0.6277 0.02048 0.01113 -0.0017 1.0000 0.1139
-5.500 -0.6162 0.01916 0.01003 0.0010 1.0000 0.1381
-5.250 -0.6115 0.01722 0.00876 0.0045 1.0000 0.2086
-5.000 -0.6175 0.01497 0.00823 0.0099 1.0000 0.4738
-4.750 -0.6054 0.01487 0.00841 0.0135 1.0000 0.5696
-4.500 -0.5896 0.01502 0.00858 0.0163 1.0000 0.6188
-4.250 -0.5728 0.01525 0.00882 0.0191 1.0000 0.6545
-4.000 -0.5558 0.01551 0.00907 0.0217 1.0000 0.6846
-3.750 -0.5391 0.01574 0.00925 0.0243 1.0000 0.7119
-3.500 -0.5221 0.01607 0.00960 0.0272 1.0000 0.7353
-3.250 -0.5062 0.01632 0.00983 0.0300 1.0000 0.7594
-3.000 -0.4900 0.01662 0.01013 0.0330 1.0000 0.7808
-2.750 -0.4743 0.01676 0.01022 0.0355 1.0000 0.8022
-2.500 -0.4580 0.01691 0.01033 0.0379 1.0000 0.8201
-2.250 -0.4413 0.01701 0.01040 0.0400 1.0000 0.8368
-2.000 -0.4049 0.01724 0.01054 0.0382 0.9942 0.8522
-1.750 -0.3622 0.01744 0.01062 0.0350 0.9854 0.8660
-1.500 -0.3200 0.01761 0.01071 0.0318 0.9764 0.8791
-1.250 -0.2686 0.01791 0.01091 0.0271 0.9700 0.8891
-1.000 -0.2248 0.01805 0.01099 0.0237 0.9607 0.9001
-0.750 -0.1739 0.01821 0.01108 0.0187 0.9538 0.9110
-0.500 -0.1188 0.01839 0.01121 0.0132 0.9462 0.9188
-0.250 -0.0581 0.01847 0.01126 0.0063 0.9411 0.9269
0.000 0.0000 0.01853 0.01132 0.0000 0.9333 0.9333
0.250 0.0580 0.01847 0.01126 -0.0063 0.9269 0.9411
0.500 0.1188 0.01838 0.01121 -0.0132 0.9188 0.9462
0.750 0.1738 0.01821 0.01107 -0.0187 0.9110 0.9538
1.000 0.2248 0.01805 0.01099 -0.0237 0.9001 0.9607
1.250 0.2686 0.01790 0.01091 -0.0271 0.8891 0.9700
1.500 0.3200 0.01761 0.01070 -0.0318 0.8791 0.9765
1.750 0.3621 0.01743 0.01062 -0.0350 0.8660 0.9855
2.000 0.4048 0.01724 0.01054 -0.0382 0.8522 0.9942
2.250 0.4409 0.01701 0.01039 -0.0400 0.8368 1.0000
2.500 0.4576 0.01691 0.01033 -0.0379 0.8201 1.0000
2.750 0.4739 0.01676 0.01021 -0.0354 0.8023 1.0000
3.000 0.4896 0.01661 0.01013 -0.0329 0.7809 1.0000
3.250 0.5058 0.01632 0.00983 -0.0300 0.7595 1.0000
3.500 0.5217 0.01607 0.00960 -0.0272 0.7355 1.0000
3.750 0.5386 0.01574 0.00925 -0.0243 0.7120 1.0000
4.000 0.5553 0.01551 0.00907 -0.0217 0.6848 1.0000
4.250 0.5723 0.01525 0.00882 -0.0190 0.6548 1.0000
4.500 0.5891 0.01502 0.00857 -0.0162 0.6192 1.0000
4.750 0.6049 0.01487 0.00841 -0.0134 0.5702 1.0000
5.000 0.6171 0.01496 0.00823 -0.0098 0.4753 1.0000
5.250 0.6113 0.01719 0.00875 -0.0044 0.2099 1.0000
5.500 0.6160 0.01915 0.01002 -0.0009 0.1383 1.0000
5.750 0.6275 0.02047 0.01112 0.0017 0.1140 1.0000
6.000 0.6423 0.02175 0.01227 0.0040 0.1002 1.0000
6.250 0.6600 0.02318 0.01352 0.0056 0.0901 1.0000
6.500 0.6809 0.02436 0.01475 0.0070 0.0822 1.0000
6.750 0.7052 0.02641 0.01668 0.0077 0.0776 1.0000
7.000 0.7289 0.02793 0.01843 0.0088 0.0740 1.0000
7.250 0.7514 0.02946 0.02008 0.0097 0.0698 1.0000
7.500 0.7740 0.03158 0.02225 0.0104 0.0670 1.0000
7.750 0.7959 0.03474 0.02561 0.0112 0.0659 1.0000
8.000 0.8146 0.03795 0.02912 0.0124 0.0656 1.0000
8.250 0.8316 0.04063 0.03219 0.0140 0.0659 1.0000
8.500 0.8455 0.04389 0.03581 0.0156 0.0658 1.0000
8.750 0.8558 0.04771 0.03997 0.0172 0.0656 1.0000
9.000 0.8663 0.05061 0.04326 0.0190 0.0659 1.0000
9.250 0.8480 0.05552 0.04932 0.0236 0.0739 1.0000
9.500 0.8552 0.06512 0.05914 0.0239 0.0912 1.0000
9.750 0.7387 0.05553 0.05042 0.0335 0.0845 1.0000
10.000 0.7115 0.06090 0.05592 0.0346 0.0882 1.0000
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Polar data table (+)
Polar graphs
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