NACA 65-209 (naca65209-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 65-209 (naca65209-il) Reynolds number: 200,000 Max Cl/Cd: 59.33 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca65209-il-200000.txt Download as CSV file: xf-naca65209-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5310 0.08901 0.08548 -0.0262 1.0000 0.0389
-9.000 -0.5316 0.08489 0.08140 -0.0284 1.0000 0.0401
-8.750 -0.5350 0.08024 0.07679 -0.0317 1.0000 0.0405
-8.500 -0.5446 0.07489 0.07148 -0.0366 1.0000 0.0407
-8.250 -0.5571 0.07062 0.06719 -0.0390 1.0000 0.0412
-8.000 -0.5631 0.06648 0.06299 -0.0405 1.0000 0.0418
-7.750 -0.5662 0.06256 0.05898 -0.0414 1.0000 0.0430
-7.500 -0.5675 0.05877 0.05501 -0.0416 1.0000 0.0451
-7.250 -0.5689 0.05816 0.05384 -0.0400 1.0000 0.0476
-7.000 -0.5734 0.05310 0.04855 -0.0385 1.0000 0.0483
-6.750 -0.5682 0.04784 0.04346 -0.0376 1.0000 0.0500
-6.500 -0.5620 0.04538 0.04100 -0.0356 1.0000 0.0516
-6.250 -0.5569 0.04318 0.03870 -0.0333 1.0000 0.0540
-6.000 -0.5494 0.04502 0.03982 -0.0292 1.0000 0.0609
-5.750 -0.5470 0.03827 0.03317 -0.0283 1.0000 0.0634
-5.500 -0.5367 0.03610 0.03101 -0.0267 1.0000 0.0667
-5.250 -0.5265 0.03430 0.02877 -0.0247 1.0000 0.0766
-5.000 -0.4952 0.03156 0.02592 -0.0274 0.9962 0.0914
-4.500 -0.4162 0.02271 0.01556 -0.0277 0.9897 0.0443
-4.250 -0.3801 0.02027 0.01277 -0.0290 0.9863 0.0426
-4.000 -0.3415 0.01897 0.01123 -0.0310 0.9832 0.0454
-3.750 -0.3079 0.01752 0.00962 -0.0319 0.9784 0.0460
-3.500 -0.2721 0.01645 0.00844 -0.0335 0.9740 0.0474
-3.250 -0.2364 0.01496 0.00702 -0.0355 0.9707 0.0528
-3.000 -0.2027 0.01424 0.00628 -0.0370 0.9652 0.0589
-2.750 -0.1673 0.01340 0.00543 -0.0389 0.9600 0.0686
-2.500 -0.1355 0.01089 0.00451 -0.0415 0.9566 0.4377
-2.250 -0.1145 0.01027 0.00494 -0.0393 0.9496 0.7298
-2.000 -0.0854 0.01047 0.00530 -0.0383 0.9445 0.7957
-1.750 -0.0604 0.01076 0.00563 -0.0364 0.9388 0.8353
-1.500 -0.0387 0.01099 0.00585 -0.0338 0.9317 0.8624
-1.250 -0.0098 0.01110 0.00594 -0.0330 0.9277 0.8805
-1.000 0.0146 0.01107 0.00586 -0.0323 0.9191 0.8902
-0.750 0.0473 0.01101 0.00574 -0.0331 0.9146 0.8981
-0.500 0.0723 0.01099 0.00568 -0.0324 0.9068 0.9070
-0.250 0.1023 0.01092 0.00556 -0.0328 0.9014 0.9162
0.000 0.1295 0.01089 0.00553 -0.0326 0.8944 0.9246
0.250 0.1595 0.01083 0.00545 -0.0329 0.8885 0.9332
0.500 0.1889 0.01079 0.00541 -0.0332 0.8824 0.9428
0.750 0.2239 0.01075 0.00538 -0.0347 0.8761 0.9497
1.000 0.2591 0.01068 0.00531 -0.0361 0.8716 0.9576
1.250 0.2977 0.01068 0.00537 -0.0386 0.8644 0.9638
1.500 0.3355 0.01062 0.00535 -0.0407 0.8590 0.9708
1.750 0.3777 0.01060 0.00541 -0.0440 0.8522 0.9756
2.000 0.4170 0.01048 0.00535 -0.0463 0.8446 0.9819
2.250 0.4553 0.01015 0.00506 -0.0478 0.8280 0.9874
2.500 0.4887 0.00964 0.00453 -0.0478 0.7963 0.9950
2.750 0.5146 0.00924 0.00403 -0.0465 0.7525 1.0000
3.000 0.5299 0.00915 0.00389 -0.0438 0.7190 1.0000
3.250 0.5429 0.00915 0.00376 -0.0406 0.6621 1.0000
3.500 0.5440 0.00992 0.00361 -0.0351 0.4522 1.0000
3.750 0.5348 0.01294 0.00467 -0.0299 0.0841 1.0000
4.000 0.5548 0.01384 0.00547 -0.0287 0.0614 1.0000
4.250 0.5751 0.01479 0.00641 -0.0276 0.0532 1.0000
4.500 0.5964 0.01575 0.00741 -0.0266 0.0496 1.0000
4.750 0.6189 0.01675 0.00847 -0.0257 0.0466 1.0000
5.000 0.6421 0.01778 0.00946 -0.0251 0.0427 1.0000
5.250 0.6661 0.01997 0.01161 -0.0245 0.0402 1.0000
5.500 0.6933 0.02178 0.01354 -0.0242 0.0395 1.0000
5.750 0.7207 0.02380 0.01574 -0.0238 0.0393 1.0000
6.000 0.7478 0.02711 0.01927 -0.0235 0.0400 1.0000
10.000 0.7142 0.08323 0.08019 -0.0054 0.0505 1.0000
10.250 0.6966 0.09044 0.08744 -0.0093 0.0508 1.0000
10.500 0.6805 0.09873 0.09573 -0.0147 0.0509 1.0000
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Polar data table (+)
Polar graphs
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