NACA 65-206 (naca65206-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 65-206 (naca65206-il) Reynolds number: 200,000 Max Cl/Cd: 54.16 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca65206-il-200000.txt Download as CSV file: xf-naca65206-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 65-206
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.5406 0.08529 0.08182 -0.0136 1.0000 0.0236
-7.750 -0.5417 0.08158 0.07815 -0.0159 1.0000 0.0240
-7.500 -0.5416 0.07738 0.07399 -0.0198 1.0000 0.0245
-7.250 -0.5376 0.07287 0.06947 -0.0239 1.0000 0.0252
-7.000 -0.5316 0.06842 0.06498 -0.0273 1.0000 0.0258
-6.750 -0.5234 0.06405 0.06053 -0.0303 1.0000 0.0272
-6.500 -0.5086 0.05993 0.05622 -0.0336 1.0000 0.0304
-6.250 -0.4876 0.05824 0.05408 -0.0354 1.0000 0.0320
-6.000 -0.4833 0.05079 0.04653 -0.0367 1.0000 0.0332
-5.750 -0.4726 0.04677 0.04255 -0.0366 1.0000 0.0348
-5.500 -0.4581 0.04359 0.03927 -0.0365 1.0000 0.0370
-5.250 -0.4408 0.04049 0.03596 -0.0363 1.0000 0.0399
-5.000 -0.4136 0.04082 0.03556 -0.0350 1.0000 0.0451
-4.750 -0.4020 0.03409 0.02883 -0.0354 1.0000 0.0475
-4.500 -0.3842 0.03150 0.02619 -0.0349 1.0000 0.0511
-4.250 -0.3630 0.02930 0.02351 -0.0340 1.0000 0.0604
-4.000 -0.3438 0.02701 0.02111 -0.0333 1.0000 0.0670
-3.750 -0.3237 0.02496 0.01885 -0.0325 1.0000 0.0798
-3.000 -0.2441 0.01766 0.01021 -0.0274 1.0000 0.0423
-2.750 -0.2195 0.01576 0.00808 -0.0262 1.0000 0.0390
-2.500 -0.1955 0.01441 0.00660 -0.0251 1.0000 0.0385
-2.250 -0.1722 0.01338 0.00553 -0.0241 1.0000 0.0411
-2.000 -0.1491 0.01253 0.00465 -0.0234 1.0000 0.0480
-1.750 -0.1253 0.01184 0.00391 -0.0229 1.0000 0.0521
-1.500 -0.1013 0.01133 0.00331 -0.0224 1.0000 0.0625
-1.250 -0.0905 0.00833 0.00331 -0.0192 1.0000 0.8133
-1.000 -0.0701 0.00809 0.00328 -0.0163 1.0000 1.0000
-0.750 -0.0366 0.00822 0.00323 -0.0182 0.9970 1.0000
-0.500 0.0045 0.00838 0.00324 -0.0216 0.9914 1.0000
-0.250 0.0452 0.00853 0.00326 -0.0248 0.9854 1.0000
0.000 0.0871 0.00868 0.00334 -0.0282 0.9802 1.0000
0.250 0.1255 0.00877 0.00339 -0.0308 0.9732 1.0000
0.500 0.1689 0.00888 0.00349 -0.0345 0.9687 1.0000
0.750 0.2056 0.00895 0.00358 -0.0366 0.9607 1.0000
1.000 0.2496 0.00899 0.00367 -0.0403 0.9560 1.0000
1.250 0.2875 0.00903 0.00377 -0.0426 0.9483 1.0000
1.500 0.3306 0.00901 0.00386 -0.0458 0.9428 1.0000
1.750 0.3740 0.00866 0.00363 -0.0480 0.9252 1.0000
2.000 0.4033 0.00818 0.00327 -0.0463 0.8904 1.0000
2.250 0.4205 0.00787 0.00283 -0.0419 0.8240 1.0000
2.500 0.4365 0.00806 0.00256 -0.0377 0.6917 1.0000
2.750 0.4347 0.01172 0.00328 -0.0327 0.0713 1.0000
3.000 0.4586 0.01261 0.00415 -0.0320 0.0506 1.0000
3.250 0.4823 0.01355 0.00517 -0.0312 0.0436 1.0000
3.500 0.5060 0.01464 0.00633 -0.0304 0.0397 1.0000
3.750 0.5300 0.01611 0.00779 -0.0295 0.0379 1.0000
4.000 0.5557 0.01804 0.00977 -0.0288 0.0369 1.0000
4.250 0.5824 0.01967 0.01165 -0.0281 0.0332 1.0000
4.500 0.6096 0.02221 0.01447 -0.0272 0.0338 1.0000
5.250 0.6717 0.01904 0.01343 -0.0196 0.0704 1.0000
5.500 0.6895 0.02148 0.01607 -0.0191 0.0532 1.0000
6.500 0.7786 0.05047 0.04545 -0.0190 0.0364 1.0000
6.750 0.7832 0.05721 0.05256 -0.0184 0.0355 1.0000
7.000 0.7996 0.05791 0.05389 -0.0167 0.0330 1.0000
7.250 0.8084 0.06211 0.05836 -0.0165 0.0310 1.0000
7.500 0.8132 0.06655 0.06300 -0.0167 0.0298 1.0000
7.750 0.8148 0.07103 0.06764 -0.0173 0.0290 1.0000
8.000 0.8132 0.07558 0.07230 -0.0183 0.0284 1.0000
8.250 0.8077 0.08018 0.07698 -0.0197 0.0280 1.0000
8.500 0.7969 0.08461 0.08146 -0.0215 0.0280 1.0000
8.750 0.7860 0.09028 0.08714 -0.0264 0.0280 1.0000
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Polar data table (+)
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