NACA 66-206 (naca66206-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 66-206 (naca66206-il) Reynolds number: 200,000 Max Cl/Cd: 50.68 at α=2° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca66206-il-200000-n5.txt Download as CSV file: xf-naca66206-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 66-206
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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.5140 0.08558 0.08206 -0.0244 1.0000 0.0107
-7.750 -0.5142 0.08151 0.07800 -0.0275 1.0000 0.0107
-7.000 -0.5149 0.06750 0.06396 -0.0323 1.0000 0.0097
-6.750 -0.5123 0.06334 0.05975 -0.0333 1.0000 0.0092
-6.500 -0.5073 0.05914 0.05548 -0.0340 1.0000 0.0088
-6.250 -0.5007 0.05521 0.05145 -0.0342 1.0000 0.0084
-6.000 -0.4926 0.05129 0.04739 -0.0340 1.0000 0.0081
-5.750 -0.4837 0.04754 0.04348 -0.0332 1.0000 0.0079
-5.500 -0.4737 0.04385 0.03959 -0.0320 1.0000 0.0077
-5.250 -0.4622 0.04033 0.03585 -0.0305 1.0000 0.0076
-5.000 -0.4397 0.03653 0.03172 -0.0307 0.9982 0.0084
-4.250 -0.3435 0.02703 0.02097 -0.0337 0.9881 0.0107
-4.000 -0.3135 0.02372 0.01722 -0.0343 0.9848 0.0108
-3.750 -0.2836 0.02028 0.01324 -0.0348 0.9816 0.0110
-3.500 -0.2531 0.01719 0.00972 -0.0354 0.9794 0.0124
-3.250 -0.2214 0.01637 0.00869 -0.0365 0.9767 0.0172
-3.000 -0.1921 0.01492 0.00704 -0.0365 0.9732 0.0190
-2.750 -0.1610 0.01392 0.00590 -0.0371 0.9703 0.0215
-2.500 -0.1296 0.01287 0.00482 -0.0383 0.9680 0.0268
-2.250 -0.0998 0.01219 0.00404 -0.0389 0.9642 0.0294
-2.000 -0.0694 0.01174 0.00344 -0.0396 0.9605 0.0345
-1.750 -0.0376 0.01128 0.00291 -0.0407 0.9575 0.0513
-1.500 -0.0237 0.00855 0.00294 -0.0388 0.9538 0.7640
-1.250 -0.0077 0.00852 0.00315 -0.0350 0.9478 0.8692
-1.000 0.0136 0.00858 0.00323 -0.0324 0.9440 0.9243
-0.750 0.0486 0.00861 0.00319 -0.0333 0.9428 0.9565
-0.500 0.0846 0.00859 0.00307 -0.0352 0.9400 0.9625
-0.250 0.1178 0.00858 0.00300 -0.0367 0.9361 0.9681
0.000 0.1542 0.00856 0.00295 -0.0387 0.9332 0.9720
0.250 0.1900 0.00854 0.00292 -0.0407 0.9305 0.9767
0.500 0.2262 0.00853 0.00293 -0.0428 0.9276 0.9806
0.750 0.2600 0.00853 0.00297 -0.0444 0.9229 0.9860
1.000 0.2964 0.00849 0.00298 -0.0463 0.9180 0.9898
1.250 0.3303 0.00840 0.00296 -0.0475 0.9076 0.9946
1.500 0.3613 0.00816 0.00280 -0.0475 0.8830 0.9997
1.750 0.3828 0.00804 0.00271 -0.0457 0.8574 1.0000
2.000 0.4019 0.00793 0.00251 -0.0430 0.8077 1.0000
2.250 0.4087 0.00863 0.00209 -0.0374 0.5437 1.0000
2.500 0.4100 0.01110 0.00267 -0.0335 0.1231 1.0000
2.750 0.4307 0.01196 0.00317 -0.0325 0.0450 1.0000
3.000 0.4544 0.01250 0.00379 -0.0318 0.0362 1.0000
3.250 0.4781 0.01308 0.00452 -0.0312 0.0299 1.0000
3.500 0.5009 0.01385 0.00537 -0.0304 0.0253 1.0000
3.750 0.5241 0.01470 0.00633 -0.0296 0.0228 1.0000
4.000 0.5471 0.01585 0.00758 -0.0286 0.0207 1.0000
4.250 0.5713 0.01692 0.00873 -0.0279 0.0177 1.0000
4.500 0.5947 0.01915 0.01108 -0.0272 0.0144 1.0000
4.750 0.6209 0.02076 0.01296 -0.0264 0.0132 1.0000
5.000 0.6469 0.02280 0.01537 -0.0255 0.0114 1.0000
5.250 0.6715 0.02412 0.01703 -0.0249 0.0086 1.0000
5.500 0.6944 0.02665 0.01995 -0.0237 0.0079 1.0000
5.750 0.7154 0.02974 0.02347 -0.0224 0.0076 1.0000
6.000 0.7342 0.03334 0.02752 -0.0208 0.0075 1.0000
6.250 0.7511 0.03743 0.03205 -0.0190 0.0075 1.0000
6.500 0.7658 0.04228 0.03735 -0.0170 0.0079 1.0000
6.750 0.7778 0.04755 0.04303 -0.0152 0.0084 1.0000
7.000 0.7871 0.05273 0.04855 -0.0138 0.0089 1.0000
7.250 0.7938 0.05781 0.05390 -0.0128 0.0095 1.0000
7.500 0.7975 0.06283 0.05914 -0.0123 0.0100 1.0000
7.750 0.7982 0.06775 0.06424 -0.0122 0.0106 1.0000
8.000 0.7961 0.07247 0.06909 -0.0126 0.0111 1.0000
8.250 0.7907 0.07716 0.07387 -0.0134 0.0115 1.0000
8.500 0.7799 0.08157 0.07833 -0.0142 0.0119 1.0000
8.750 0.7702 0.08642 0.08321 -0.0170 0.0120 1.0000
9.000 0.7624 0.09226 0.08904 -0.0220 0.0123 1.0000
9.250 0.7575 0.09782 0.09457 -0.0257 0.0128 1.0000
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Polar data table (+)
Polar graphs
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