NASA/LANGLEY 64-012 AIRFOIL (n64012-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY 64-012 AIRFOIL (n64012-il) Reynolds number: 500,000 Max Cl/Cd: 56.48 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64012-il-500000.txt Download as CSV file: xf-n64012-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY 64-012 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.8334 0.07142 0.06878 -0.0225 1.0000 0.0202
-12.000 -0.8534 0.06580 0.06307 -0.0253 1.0000 0.0197
-11.750 -0.8802 0.05994 0.05706 -0.0274 1.0000 0.0192
-11.500 -0.9031 0.05528 0.05222 -0.0280 1.0000 0.0192
-11.250 -0.9290 0.05065 0.04737 -0.0271 1.0000 0.0188
-11.000 -0.9453 0.04756 0.04409 -0.0250 1.0000 0.0190
-10.750 -0.9657 0.04385 0.04012 -0.0214 1.0000 0.0187
-10.500 -0.9746 0.03998 0.03592 -0.0187 1.0000 0.0188
-10.250 -0.9780 0.03590 0.03144 -0.0163 1.0000 0.0186
-10.000 -0.9717 0.03283 0.02804 -0.0145 1.0000 0.0187
-9.750 -0.9601 0.03013 0.02503 -0.0130 1.0000 0.0189
-9.500 -0.9445 0.02801 0.02267 -0.0119 1.0000 0.0192
-9.250 -0.9264 0.02613 0.02055 -0.0109 1.0000 0.0196
-9.000 -0.9062 0.02476 0.01899 -0.0102 1.0000 0.0202
-8.750 -0.8846 0.02414 0.01821 -0.0095 1.0000 0.0207
-8.500 -0.8637 0.02171 0.01560 -0.0089 1.0000 0.0213
-8.250 -0.8420 0.02010 0.01394 -0.0084 1.0000 0.0219
-8.000 -0.8197 0.01913 0.01295 -0.0079 1.0000 0.0225
-7.750 -0.7969 0.01831 0.01209 -0.0075 1.0000 0.0233
-7.500 -0.7741 0.01750 0.01125 -0.0070 1.0000 0.0240
-7.250 -0.7515 0.01677 0.01047 -0.0065 1.0000 0.0249
-7.000 -0.7258 0.01621 0.00987 -0.0065 0.9948 0.0261
-6.750 -0.6924 0.01511 0.00869 -0.0084 0.9776 0.0274
-6.500 -0.6619 0.01413 0.00768 -0.0097 0.9622 0.0292
-6.250 -0.6362 0.01361 0.00711 -0.0095 0.9458 0.0309
-6.000 -0.6146 0.01322 0.00664 -0.0084 0.9302 0.0328
-5.750 -0.5937 0.01276 0.00609 -0.0071 0.9162 0.0348
-5.500 -0.5728 0.01220 0.00550 -0.0059 0.9034 0.0388
-5.250 -0.5491 0.01190 0.00513 -0.0052 0.8921 0.0426
-5.000 -0.5265 0.01141 0.00460 -0.0042 0.8815 0.0494
-4.750 -0.5026 0.01100 0.00419 -0.0036 0.8709 0.0608
-4.500 -0.4801 0.01030 0.00373 -0.0029 0.8608 0.1110
-4.250 -0.4612 0.00912 0.00315 -0.0021 0.8513 0.2521
-4.000 -0.4429 0.00784 0.00267 -0.0011 0.8412 0.4423
-3.750 -0.4177 0.00755 0.00256 -0.0007 0.8320 0.5067
-3.500 -0.3914 0.00743 0.00245 -0.0003 0.8236 0.5399
-3.250 -0.3639 0.00735 0.00237 -0.0003 0.8144 0.5626
-3.000 -0.3365 0.00731 0.00229 -0.0001 0.8066 0.5787
-2.750 -0.3087 0.00726 0.00223 -0.0001 0.7977 0.5943
-2.500 -0.2813 0.00723 0.00218 0.0001 0.7898 0.6119
-2.250 -0.2540 0.00721 0.00220 0.0004 0.7811 0.6334
-2.000 -0.2266 0.00724 0.00222 0.0006 0.7730 0.6524
-1.750 -0.1986 0.00721 0.00217 0.0006 0.7647 0.6616
-1.500 -0.1702 0.00721 0.00212 0.0005 0.7569 0.6695
-1.250 -0.1421 0.00717 0.00208 0.0005 0.7492 0.6768
-1.000 -0.1136 0.00718 0.00204 0.0004 0.7415 0.6847
-0.750 -0.0854 0.00714 0.00201 0.0003 0.7340 0.6910
-0.500 -0.0569 0.00714 0.00199 0.0002 0.7265 0.6983
-0.250 -0.0285 0.00712 0.00196 0.0001 0.7189 0.7047
0.000 -0.0001 0.00713 0.00196 0.0000 0.7118 0.7118
0.250 0.0284 0.00712 0.00196 -0.0001 0.7047 0.7189
0.500 0.0567 0.00714 0.00199 -0.0002 0.6983 0.7265
0.750 0.0852 0.00714 0.00201 -0.0003 0.6910 0.7340
1.000 0.1135 0.00718 0.00204 -0.0003 0.6847 0.7415
1.250 0.1420 0.00717 0.00208 -0.0005 0.6768 0.7492
1.500 0.1700 0.00721 0.00212 -0.0005 0.6695 0.7569
1.750 0.1985 0.00721 0.00217 -0.0006 0.6615 0.7647
2.000 0.2264 0.00724 0.00222 -0.0005 0.6524 0.7730
2.250 0.2539 0.00721 0.00220 -0.0003 0.6336 0.7811
2.500 0.2811 0.00723 0.00218 -0.0001 0.6120 0.7898
2.750 0.3085 0.00726 0.00223 0.0001 0.5944 0.7977
3.000 0.3363 0.00731 0.00229 0.0001 0.5787 0.8066
3.250 0.3637 0.00735 0.00237 0.0003 0.5626 0.8145
3.500 0.3912 0.00743 0.00245 0.0004 0.5398 0.8236
3.750 0.4176 0.00755 0.00256 0.0007 0.5069 0.8321
4.000 0.4428 0.00784 0.00267 0.0011 0.4429 0.8413
4.250 0.4611 0.00911 0.00315 0.0021 0.2531 0.8513
4.500 0.4800 0.01029 0.00373 0.0030 0.1120 0.8608
4.750 0.5025 0.01100 0.00419 0.0036 0.0609 0.8709
5.000 0.5264 0.01140 0.00460 0.0043 0.0494 0.8815
5.250 0.5490 0.01190 0.00513 0.0052 0.0427 0.8922
5.500 0.5726 0.01220 0.00550 0.0060 0.0390 0.9035
5.750 0.5935 0.01276 0.00610 0.0071 0.0348 0.9163
6.000 0.6144 0.01322 0.00664 0.0084 0.0329 0.9303
6.250 0.6360 0.01361 0.00711 0.0096 0.0310 0.9459
6.500 0.6618 0.01412 0.00768 0.0097 0.0292 0.9623
6.750 0.6923 0.01515 0.00873 0.0084 0.0273 0.9777
7.000 0.7260 0.01619 0.00985 0.0065 0.0261 0.9949
7.250 0.7515 0.01676 0.01046 0.0065 0.0250 1.0000
7.500 0.7741 0.01750 0.01124 0.0070 0.0240 1.0000
7.750 0.7969 0.01827 0.01205 0.0075 0.0232 1.0000
8.000 0.8195 0.01911 0.01292 0.0080 0.0225 1.0000
8.250 0.8419 0.02009 0.01393 0.0084 0.0219 1.0000
8.500 0.8635 0.02170 0.01559 0.0090 0.0213 1.0000
8.750 0.8844 0.02400 0.01807 0.0096 0.0207 1.0000
9.000 0.9059 0.02482 0.01906 0.0102 0.0202 1.0000
9.250 0.9262 0.02613 0.02055 0.0110 0.0196 1.0000
9.500 0.9443 0.02800 0.02266 0.0119 0.0193 1.0000
9.750 0.9598 0.03016 0.02507 0.0131 0.0189 1.0000
10.000 0.9716 0.03276 0.02797 0.0145 0.0187 1.0000
10.250 0.9765 0.03618 0.03174 0.0165 0.0187 1.0000
10.500 0.9742 0.04000 0.03593 0.0188 0.0187 1.0000
10.750 0.9633 0.04420 0.04049 0.0216 0.0190 1.0000
11.000 0.9435 0.04777 0.04432 0.0251 0.0191 1.0000
11.250 0.9236 0.05138 0.04812 0.0272 0.0193 1.0000
11.500 0.9025 0.05535 0.05229 0.0280 0.0192 1.0000
11.750 0.8802 0.06004 0.05713 0.0274 0.0198 1.0000
12.000 0.8555 0.06554 0.06278 0.0255 0.0199 1.0000
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Polar data table (+)
Polar graphs
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