NACA 65-209 (naca65209-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 65-209 (naca65209-il) Reynolds number: 500,000 Max Cl/Cd: 71.32 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca65209-il-500000.txt Download as CSV file: xf-naca65209-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 65-209
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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.5860 0.05117 0.04835 -0.0420 1.0000 0.0168
-7.500 -0.5937 0.04627 0.04330 -0.0404 1.0000 0.0173
-7.250 -0.5904 0.04421 0.04121 -0.0385 1.0000 0.0178
-7.000 -0.5908 0.04230 0.03922 -0.0354 1.0000 0.0181
-6.750 -0.5696 0.03963 0.03640 -0.0369 0.9973 0.0192
-6.500 -0.5397 0.03653 0.03304 -0.0394 0.9938 0.0215
-6.250 -0.5018 0.03665 0.03277 -0.0402 0.9898 0.0243
-5.500 -0.4199 0.01976 0.01447 -0.0446 0.9776 0.0200
-5.250 -0.3869 0.01737 0.01177 -0.0457 0.9740 0.0202
-5.000 -0.3525 0.01616 0.01037 -0.0471 0.9707 0.0219
-4.750 -0.3214 0.01467 0.00871 -0.0477 0.9653 0.0226
-4.500 -0.2907 0.01369 0.00762 -0.0483 0.9592 0.0235
-4.250 -0.2601 0.01317 0.00701 -0.0488 0.9532 0.0243
-4.000 -0.2369 0.01141 0.00519 -0.0481 0.9445 0.0268
-3.750 -0.2117 0.01077 0.00449 -0.0477 0.9359 0.0287
-3.500 -0.1861 0.01029 0.00395 -0.0472 0.9281 0.0309
-3.250 -0.1610 0.00989 0.00349 -0.0467 0.9191 0.0335
-3.000 -0.1355 0.00949 0.00300 -0.0462 0.9114 0.0373
-2.750 -0.1100 0.00910 0.00257 -0.0458 0.9027 0.0449
-2.500 -0.0843 0.00866 0.00221 -0.0454 0.8947 0.0778
-2.250 -0.0669 0.00645 0.00178 -0.0449 0.8865 0.5712
-2.000 -0.0418 0.00621 0.00179 -0.0443 0.8783 0.6605
-1.500 0.0099 0.00611 0.00184 -0.0431 0.8631 0.7437
-1.250 0.0359 0.00613 0.00187 -0.0424 0.8565 0.7687
-1.000 0.0624 0.00615 0.00191 -0.0420 0.8484 0.7880
-0.750 0.0890 0.00618 0.00192 -0.0416 0.8417 0.8024
-0.500 0.1161 0.00617 0.00191 -0.0414 0.8340 0.8102
-0.250 0.1438 0.00619 0.00188 -0.0413 0.8274 0.8184
0.000 0.1710 0.00618 0.00189 -0.0411 0.8197 0.8257
0.250 0.1985 0.00620 0.00189 -0.0410 0.8131 0.8336
0.500 0.2257 0.00621 0.00192 -0.0409 0.8055 0.8414
0.750 0.2529 0.00623 0.00194 -0.0407 0.7989 0.8491
1.000 0.2801 0.00624 0.00198 -0.0406 0.7913 0.8574
1.250 0.3069 0.00627 0.00203 -0.0403 0.7840 0.8651
1.750 0.3588 0.00628 0.00204 -0.0392 0.7574 0.8818
2.000 0.3840 0.00627 0.00204 -0.0384 0.7374 0.8907
2.250 0.4081 0.00630 0.00200 -0.0374 0.7067 0.9001
2.500 0.4317 0.00637 0.00203 -0.0363 0.6756 0.9090
2.750 0.4554 0.00646 0.00207 -0.0353 0.6401 0.9190
3.000 0.4771 0.00669 0.00212 -0.0339 0.5797 0.9302
3.250 0.4871 0.00797 0.00242 -0.0307 0.3425 0.9444
3.500 0.4967 0.00986 0.00311 -0.0282 0.0604 0.9613
3.750 0.5256 0.01041 0.00360 -0.0286 0.0388 0.9749
4.000 0.5599 0.01087 0.00411 -0.0303 0.0328 0.9876
4.250 0.5877 0.01176 0.00506 -0.0308 0.0282 1.0000
4.500 0.6131 0.01221 0.00554 -0.0306 0.0266 1.0000
4.750 0.6376 0.01285 0.00624 -0.0302 0.0251 1.0000
5.000 0.6617 0.01357 0.00701 -0.0298 0.0237 1.0000
5.250 0.6856 0.01442 0.00790 -0.0292 0.0226 1.0000
5.500 0.7096 0.01532 0.00881 -0.0288 0.0213 1.0000
5.750 0.7321 0.01730 0.01084 -0.0281 0.0198 1.0000
6.000 0.7570 0.01971 0.01343 -0.0276 0.0192 1.0000
6.250 0.7826 0.02074 0.01461 -0.0272 0.0188 1.0000
6.500 0.8073 0.02255 0.01664 -0.0266 0.0184 1.0000
6.750 0.8312 0.02433 0.01866 -0.0259 0.0178 1.0000
7.000 0.8539 0.02631 0.02090 -0.0250 0.0170 1.0000
7.250 0.8708 0.03037 0.02540 -0.0232 0.0177 1.0000
11.000 0.6775 0.10293 0.10101 -0.0217 0.0193 1.0000
11.250 0.6695 0.11103 0.10912 -0.0262 0.0193 1.0000
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Polar data table (+)
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