NASA/LANGLEY 64-012 AIRFOIL (n64012-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY 64-012 AIRFOIL (n64012-il) Reynolds number: 50,000 Max Cl/Cd: 27.12 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n64012-il-50000-n5.txt Download as CSV file: xf-n64012-il-50000-n5.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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.6974 0.09710 0.08992 -0.0133 1.0000 0.0538
-11.250 -0.7193 0.08896 0.08175 -0.0189 1.0000 0.0530
-11.000 -0.7439 0.08174 0.07445 -0.0234 1.0000 0.0516
-10.750 -0.7686 0.07581 0.06839 -0.0264 1.0000 0.0512
-10.500 -0.7902 0.07104 0.06345 -0.0277 1.0000 0.0510
-10.250 -0.8083 0.06711 0.05938 -0.0275 1.0000 0.0510
-10.000 -0.8218 0.06365 0.05571 -0.0263 1.0000 0.0512
-9.750 -0.8291 0.06020 0.05196 -0.0252 1.0000 0.0518
-9.500 -0.8316 0.05685 0.04830 -0.0240 1.0000 0.0525
-9.250 -0.8301 0.05351 0.04460 -0.0228 1.0000 0.0532
-9.000 -0.8241 0.05023 0.04092 -0.0215 1.0000 0.0538
-8.750 -0.8137 0.04711 0.03739 -0.0204 1.0000 0.0545
-8.500 -0.7992 0.04420 0.03405 -0.0194 1.0000 0.0554
-8.250 -0.7815 0.04162 0.03100 -0.0184 1.0000 0.0570
-8.000 -0.7609 0.03926 0.02852 -0.0179 1.0000 0.0598
-7.750 -0.7384 0.03737 0.02648 -0.0175 1.0000 0.0630
-7.500 -0.7118 0.03543 0.02433 -0.0171 1.0000 0.0661
-7.250 -0.6829 0.03377 0.02237 -0.0167 1.0000 0.0699
-7.000 -0.6578 0.03225 0.02098 -0.0164 1.0000 0.0755
-6.750 -0.6325 0.03100 0.01954 -0.0156 1.0000 0.0817
-6.500 -0.6125 0.02967 0.01824 -0.0145 1.0000 0.0877
-6.250 -0.5954 0.02854 0.01705 -0.0131 1.0000 0.0977
-6.000 -0.5822 0.02726 0.01584 -0.0113 1.0000 0.1087
-5.750 -0.5710 0.02596 0.01464 -0.0094 1.0000 0.1239
-5.250 -0.5628 0.02261 0.01223 -0.0040 1.0000 0.2232
-5.000 -0.5677 0.02094 0.01195 0.0008 1.0000 0.3965
-4.750 -0.5612 0.02086 0.01224 0.0052 1.0000 0.5260
-4.500 -0.5530 0.02107 0.01249 0.0094 1.0000 0.5861
-4.250 -0.5444 0.02156 0.01301 0.0139 1.0000 0.6331
-4.000 -0.5367 0.02240 0.01390 0.0194 1.0000 0.6777
-3.750 -0.5224 0.02340 0.01486 0.0243 0.9987 0.7126
-3.500 -0.4882 0.02399 0.01524 0.0242 0.9879 0.7386
-3.250 -0.4513 0.02403 0.01502 0.0223 0.9778 0.7521
-3.000 -0.4117 0.02407 0.01482 0.0201 0.9688 0.7609
-2.750 -0.3734 0.02395 0.01449 0.0176 0.9596 0.7711
-2.500 -0.3351 0.02385 0.01420 0.0151 0.9505 0.7807
-2.250 -0.2973 0.02378 0.01394 0.0128 0.9413 0.7885
-2.000 -0.2599 0.02360 0.01361 0.0102 0.9329 0.7977
-1.750 -0.2235 0.02355 0.01344 0.0084 0.9235 0.8042
-1.500 -0.1925 0.02338 0.01316 0.0070 0.9135 0.8134
-1.250 -0.1532 0.02335 0.01304 0.0046 0.9058 0.8194
-1.000 -0.1269 0.02323 0.01285 0.0042 0.8952 0.8290
-0.750 -0.0887 0.02321 0.01278 0.0022 0.8877 0.8352
-0.500 -0.0630 0.02314 0.01267 0.0020 0.8776 0.8448
-0.250 -0.0298 0.02314 0.01266 0.0007 0.8692 0.8518
0.000 0.0000 0.02311 0.01261 0.0000 0.8609 0.8609
0.250 0.0298 0.02314 0.01266 -0.0007 0.8518 0.8692
0.500 0.0630 0.02313 0.01267 -0.0020 0.8447 0.8776
0.750 0.0888 0.02321 0.01278 -0.0022 0.8351 0.8878
1.000 0.1269 0.02323 0.01285 -0.0042 0.8290 0.8951
1.250 0.1532 0.02334 0.01304 -0.0046 0.8194 0.9058
1.500 0.1926 0.02338 0.01316 -0.0070 0.8134 0.9135
1.750 0.2235 0.02354 0.01343 -0.0084 0.8042 0.9235
2.000 0.2599 0.02360 0.01361 -0.0102 0.7977 0.9329
2.250 0.2973 0.02378 0.01395 -0.0128 0.7885 0.9413
2.500 0.3352 0.02385 0.01420 -0.0151 0.7807 0.9505
2.750 0.3734 0.02395 0.01448 -0.0176 0.7711 0.9596
3.000 0.4116 0.02406 0.01482 -0.0201 0.7610 0.9687
3.250 0.4513 0.02402 0.01502 -0.0223 0.7521 0.9778
3.500 0.4882 0.02399 0.01524 -0.0242 0.7386 0.9879
3.750 0.5224 0.02339 0.01485 -0.0242 0.7124 0.9986
4.000 0.5367 0.02240 0.01387 -0.0194 0.6778 1.0000
4.250 0.5444 0.02156 0.01301 -0.0139 0.6331 1.0000
4.500 0.5530 0.02106 0.01248 -0.0093 0.5858 1.0000
4.750 0.5612 0.02086 0.01224 -0.0052 0.5257 1.0000
5.000 0.5678 0.02094 0.01195 -0.0008 0.3969 1.0000
5.250 0.5628 0.02261 0.01223 0.0040 0.2230 1.0000
5.500 0.5638 0.02446 0.01343 0.0070 0.1540 1.0000
5.750 0.5711 0.02596 0.01464 0.0094 0.1239 1.0000
6.000 0.5822 0.02726 0.01585 0.0113 0.1085 1.0000
6.250 0.5955 0.02853 0.01704 0.0130 0.0977 1.0000
6.500 0.6126 0.02967 0.01824 0.0145 0.0877 1.0000
6.750 0.6326 0.03099 0.01954 0.0156 0.0817 1.0000
7.000 0.6578 0.03224 0.02097 0.0164 0.0755 1.0000
7.250 0.6831 0.03378 0.02238 0.0167 0.0698 1.0000
7.500 0.7119 0.03544 0.02435 0.0171 0.0660 1.0000
7.750 0.7384 0.03737 0.02649 0.0175 0.0630 1.0000
8.000 0.7609 0.03927 0.02851 0.0179 0.0597 1.0000
8.250 0.7817 0.04163 0.03101 0.0184 0.0570 1.0000
8.500 0.7992 0.04422 0.03408 0.0193 0.0554 1.0000
8.750 0.8138 0.04711 0.03739 0.0204 0.0545 1.0000
9.000 0.8243 0.05023 0.04091 0.0215 0.0538 1.0000
9.250 0.8303 0.05353 0.04461 0.0227 0.0532 1.0000
9.500 0.8319 0.05687 0.04831 0.0240 0.0525 1.0000
9.750 0.8292 0.06025 0.05201 0.0252 0.0518 1.0000
10.000 0.8221 0.06370 0.05576 0.0262 0.0512 1.0000
10.250 0.8088 0.06711 0.05938 0.0274 0.0508 1.0000
10.500 0.7906 0.07109 0.06351 0.0276 0.0510 1.0000
10.750 0.7686 0.07594 0.06852 0.0262 0.0513 1.0000
11.000 0.7442 0.08188 0.07459 0.0232 0.0519 1.0000
11.250 0.7193 0.08911 0.08190 0.0186 0.0527 1.0000
11.500 0.6975 0.09734 0.09015 0.0130 0.0537 1.0000
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Polar data table (+)
Polar graphs
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