NASA/LANGLEY 64-012 AIRFOIL (n64012-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY 64-012 AIRFOIL (n64012-il) Reynolds number: 200,000 Max Cl/Cd: 39.38 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n64012-il-200000-n5.txt Download as CSV file: xf-n64012-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.750 -0.7959 0.08606 0.08235 -0.0117 1.0000 0.0184
-12.500 -0.8464 0.07280 0.06884 -0.0207 1.0000 0.0179
-12.250 -0.8798 0.06477 0.06056 -0.0250 1.0000 0.0177
-12.000 -0.9028 0.05918 0.05474 -0.0270 1.0000 0.0177
-11.750 -0.9208 0.05474 0.05009 -0.0277 1.0000 0.0178
-11.500 -0.9311 0.05158 0.04675 -0.0274 1.0000 0.0180
-11.250 -0.9433 0.04830 0.04323 -0.0264 1.0000 0.0182
-11.000 -0.9522 0.04544 0.04013 -0.0246 1.0000 0.0183
-10.750 -0.9586 0.04280 0.03724 -0.0222 1.0000 0.0185
-10.500 -0.9587 0.04017 0.03433 -0.0202 1.0000 0.0188
-10.250 -0.9535 0.03763 0.03148 -0.0186 1.0000 0.0191
-10.000 -0.9440 0.03532 0.02885 -0.0172 1.0000 0.0195
-9.750 -0.9311 0.03320 0.02642 -0.0160 1.0000 0.0200
-9.500 -0.9155 0.03147 0.02441 -0.0149 1.0000 0.0207
-9.250 -0.8979 0.03005 0.02269 -0.0140 1.0000 0.0213
-9.000 -0.8790 0.02806 0.02060 -0.0134 1.0000 0.0219
-8.750 -0.8588 0.02668 0.01915 -0.0128 1.0000 0.0225
-8.500 -0.8381 0.02548 0.01785 -0.0122 1.0000 0.0231
-8.250 -0.8169 0.02434 0.01663 -0.0116 1.0000 0.0238
-8.000 -0.7955 0.02326 0.01546 -0.0110 1.0000 0.0246
-7.750 -0.7742 0.02224 0.01436 -0.0103 1.0000 0.0255
-7.500 -0.7530 0.02130 0.01335 -0.0096 1.0000 0.0265
-7.250 -0.7311 0.02059 0.01252 -0.0090 1.0000 0.0277
-7.000 -0.7129 0.01948 0.01143 -0.0080 1.0000 0.0292
-6.750 -0.6930 0.01872 0.01067 -0.0072 0.9987 0.0308
-6.250 -0.6300 0.01717 0.00896 -0.0101 0.9585 0.0352
-6.000 -0.6020 0.01641 0.00815 -0.0107 0.9433 0.0383
-5.750 -0.5760 0.01587 0.00756 -0.0108 0.9288 0.0429
-5.500 -0.5516 0.01539 0.00700 -0.0104 0.9149 0.0476
-5.250 -0.5287 0.01489 0.00647 -0.0097 0.9019 0.0553
-5.000 -0.5060 0.01439 0.00599 -0.0089 0.8897 0.0688
-4.750 -0.4843 0.01378 0.00550 -0.0081 0.8778 0.0987
-4.500 -0.4647 0.01292 0.00498 -0.0071 0.8666 0.1722
-4.250 -0.4472 0.01181 0.00440 -0.0060 0.8561 0.2959
-4.000 -0.4292 0.01090 0.00410 -0.0046 0.8459 0.4378
-3.750 -0.4054 0.01063 0.00401 -0.0038 0.8361 0.5026
-3.500 -0.3802 0.01051 0.00390 -0.0032 0.8275 0.5383
-3.250 -0.3549 0.01042 0.00387 -0.0026 0.8182 0.5749
-3.000 -0.3299 0.01041 0.00392 -0.0017 0.8098 0.6127
-2.750 -0.3042 0.01042 0.00396 -0.0010 0.8011 0.6375
-2.500 -0.2771 0.01039 0.00385 -0.0007 0.7926 0.6489
-2.250 -0.2499 0.01035 0.00374 -0.0005 0.7847 0.6567
-2.000 -0.2222 0.01031 0.00365 -0.0004 0.7766 0.6648
-1.750 -0.1948 0.01027 0.00355 -0.0003 0.7692 0.6721
-1.500 -0.1669 0.01023 0.00346 -0.0003 0.7610 0.6783
-1.250 -0.1394 0.01020 0.00339 -0.0001 0.7536 0.6844
-0.750 -0.0837 0.01015 0.00327 0.0000 0.7387 0.6971
-0.500 -0.0556 0.01014 0.00324 -0.0001 0.7314 0.7039
-0.250 -0.0278 0.01013 0.00321 0.0000 0.7246 0.7104
0.000 0.0000 0.01012 0.00321 0.0000 0.7175 0.7175
0.250 0.0279 0.01013 0.00321 0.0000 0.7104 0.7245
0.500 0.0557 0.01014 0.00324 0.0001 0.7039 0.7313
0.750 0.0837 0.01015 0.00327 0.0000 0.6972 0.7387
1.250 0.1394 0.01020 0.00339 0.0001 0.6844 0.7536
1.500 0.1669 0.01023 0.00346 0.0003 0.6783 0.7610
1.750 0.1948 0.01027 0.00355 0.0002 0.6721 0.7692
2.000 0.2223 0.01031 0.00365 0.0004 0.6647 0.7765
2.250 0.2499 0.01035 0.00374 0.0005 0.6567 0.7847
2.500 0.2771 0.01039 0.00385 0.0007 0.6489 0.7926
2.750 0.3043 0.01042 0.00396 0.0010 0.6375 0.8011
3.000 0.3299 0.01041 0.00392 0.0017 0.6127 0.8098
3.250 0.3549 0.01042 0.00387 0.0025 0.5751 0.8182
3.500 0.3802 0.01051 0.00390 0.0032 0.5383 0.8275
3.750 0.4054 0.01063 0.00401 0.0038 0.5027 0.8361
4.000 0.4292 0.01090 0.00411 0.0046 0.4386 0.8459
4.250 0.4472 0.01181 0.00440 0.0060 0.2956 0.8561
4.500 0.4647 0.01292 0.00498 0.0071 0.1721 0.8666
4.750 0.4843 0.01379 0.00550 0.0081 0.0986 0.8778
5.000 0.5061 0.01439 0.00599 0.0089 0.0688 0.8897
5.250 0.5287 0.01489 0.00647 0.0097 0.0554 0.9019
5.500 0.5517 0.01539 0.00700 0.0104 0.0476 0.9148
5.750 0.5760 0.01586 0.00756 0.0108 0.0428 0.9288
6.000 0.6021 0.01642 0.00815 0.0107 0.0383 0.9434
6.250 0.6301 0.01717 0.00896 0.0101 0.0353 0.9585
6.750 0.6930 0.01872 0.01068 0.0072 0.0308 0.9987
7.000 0.7130 0.01948 0.01143 0.0080 0.0292 1.0000
7.250 0.7313 0.02059 0.01252 0.0090 0.0277 1.0000
7.500 0.7531 0.02131 0.01335 0.0096 0.0264 1.0000
7.750 0.7743 0.02225 0.01437 0.0103 0.0254 1.0000
8.000 0.7956 0.02327 0.01547 0.0110 0.0246 1.0000
8.250 0.8170 0.02435 0.01663 0.0116 0.0238 1.0000
8.500 0.8382 0.02547 0.01783 0.0122 0.0230 1.0000
8.750 0.8590 0.02668 0.01915 0.0128 0.0225 1.0000
9.000 0.8792 0.02806 0.02061 0.0134 0.0220 1.0000
9.250 0.8980 0.03009 0.02274 0.0139 0.0213 1.0000
9.500 0.9156 0.03150 0.02443 0.0149 0.0207 1.0000
9.750 0.9313 0.03316 0.02639 0.0159 0.0199 1.0000
10.000 0.9441 0.03535 0.02888 0.0171 0.0195 1.0000
10.250 0.9534 0.03770 0.03155 0.0186 0.0191 1.0000
10.500 0.9590 0.04016 0.03432 0.0202 0.0188 1.0000
10.750 0.9589 0.04278 0.03723 0.0221 0.0185 1.0000
11.000 0.9522 0.04552 0.04022 0.0245 0.0183 1.0000
11.250 0.9429 0.04841 0.04335 0.0263 0.0182 1.0000
11.500 0.9316 0.05158 0.04675 0.0273 0.0180 1.0000
11.750 0.9163 0.05542 0.05081 0.0275 0.0179 1.0000
12.000 0.9013 0.05949 0.05507 0.0268 0.0178 1.0000
12.250 0.8796 0.06492 0.06072 0.0248 0.0177 1.0000
12.500 0.8435 0.07348 0.06954 0.0201 0.0180 1.0000
12.750 0.7936 0.08690 0.08320 0.0110 0.0185 1.0000
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Polar data table (+)
Polar graphs
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