NACA 64A210 (naca64a210-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 64A210 (naca64a210-il) Reynolds number: 200,000 Max Cl/Cd: 59.25 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca64a210-il-200000-n5.txt Download as CSV file: xf-naca64a210-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64A210
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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5863 0.08287 0.07933 -0.0283 1.0000 0.0155
-9.750 -0.5947 0.07468 0.07118 -0.0349 1.0000 0.0154
-9.500 -0.6083 0.06709 0.06354 -0.0419 1.0000 0.0151
-9.250 -0.6251 0.06154 0.05789 -0.0455 1.0000 0.0150
-9.000 -0.6392 0.05640 0.05258 -0.0465 1.0000 0.0150
-8.750 -0.6491 0.05122 0.04716 -0.0465 1.0000 0.0150
-8.500 -0.6559 0.04631 0.04196 -0.0453 1.0000 0.0151
-8.250 -0.6604 0.04177 0.03706 -0.0432 1.0000 0.0151
-8.000 -0.6619 0.03774 0.03261 -0.0405 1.0000 0.0153
-7.750 -0.6592 0.03440 0.02881 -0.0377 1.0000 0.0155
-7.500 -0.6549 0.03120 0.02522 -0.0349 1.0000 0.0159
-7.250 -0.6456 0.02898 0.02275 -0.0327 1.0000 0.0163
-7.000 -0.6196 0.02770 0.02132 -0.0337 0.9971 0.0174
-6.750 -0.5891 0.02625 0.01963 -0.0352 0.9934 0.0188
-6.500 -0.5595 0.02407 0.01706 -0.0361 0.9893 0.0198
-6.250 -0.5278 0.02213 0.01475 -0.0373 0.9861 0.0207
-6.000 -0.4972 0.02068 0.01303 -0.0382 0.9818 0.0215
-5.750 -0.4660 0.01921 0.01139 -0.0393 0.9781 0.0232
-5.500 -0.4329 0.01810 0.01023 -0.0409 0.9751 0.0247
-5.250 -0.4036 0.01715 0.00918 -0.0415 0.9695 0.0259
-5.000 -0.3717 0.01626 0.00821 -0.0426 0.9652 0.0275
-4.750 -0.3419 0.01550 0.00737 -0.0433 0.9598 0.0292
-4.500 -0.3119 0.01483 0.00663 -0.0440 0.9540 0.0319
-4.250 -0.2805 0.01415 0.00591 -0.0451 0.9494 0.0363
-4.000 -0.2529 0.01367 0.00535 -0.0452 0.9418 0.0407
-3.750 -0.2220 0.01312 0.00478 -0.0460 0.9366 0.0495
-3.500 -0.1949 0.01268 0.00437 -0.0461 0.9288 0.0685
-3.250 -0.1668 0.01198 0.00394 -0.0465 0.9227 0.1312
-3.000 -0.1452 0.01079 0.00353 -0.0462 0.9140 0.3216
-2.750 -0.1220 0.00992 0.00343 -0.0456 0.9074 0.5198
-2.500 -0.0974 0.00970 0.00339 -0.0448 0.8987 0.5881
-2.250 -0.0706 0.00955 0.00332 -0.0443 0.8921 0.6377
-2.000 -0.0465 0.00945 0.00337 -0.0431 0.8832 0.6912
-1.750 -0.0211 0.00939 0.00338 -0.0422 0.8759 0.7289
-1.500 0.0057 0.00933 0.00328 -0.0418 0.8675 0.7448
-1.250 0.0328 0.00927 0.00319 -0.0415 0.8594 0.7564
-1.000 0.0602 0.00921 0.00310 -0.0413 0.8514 0.7670
-0.750 0.0871 0.00917 0.00303 -0.0410 0.8425 0.7781
-0.500 0.1146 0.00912 0.00294 -0.0408 0.8349 0.7894
-0.250 0.1412 0.00907 0.00291 -0.0404 0.8252 0.8006
0.000 0.1681 0.00903 0.00287 -0.0401 0.8166 0.8125
0.250 0.1951 0.00899 0.00285 -0.0397 0.8080 0.8250
0.500 0.2220 0.00896 0.00285 -0.0394 0.7986 0.8387
0.750 0.2494 0.00894 0.00285 -0.0391 0.7902 0.8538
1.000 0.2773 0.00893 0.00289 -0.0389 0.7811 0.8704
1.250 0.3063 0.00894 0.00293 -0.0390 0.7722 0.8890
1.500 0.3374 0.00896 0.00297 -0.0395 0.7639 0.9086
1.750 0.3708 0.00899 0.00305 -0.0407 0.7544 0.9288
2.000 0.4054 0.00903 0.00314 -0.0421 0.7455 0.9493
2.250 0.4412 0.00908 0.00321 -0.0438 0.7355 0.9680
2.500 0.4766 0.00913 0.00328 -0.0454 0.7215 0.9860
2.750 0.5106 0.00919 0.00331 -0.0467 0.6954 1.0000
3.000 0.5324 0.00931 0.00328 -0.0453 0.6558 1.0000
3.250 0.5544 0.00950 0.00331 -0.0440 0.6134 1.0000
3.500 0.5766 0.00975 0.00343 -0.0428 0.5698 1.0000
3.750 0.5978 0.01009 0.00356 -0.0415 0.5115 1.0000
4.000 0.6143 0.01083 0.00378 -0.0395 0.3948 1.0000
4.250 0.6260 0.01225 0.00436 -0.0372 0.2300 1.0000
4.500 0.6414 0.01350 0.00499 -0.0357 0.1134 1.0000
4.750 0.6614 0.01432 0.00557 -0.0346 0.0672 1.0000
5.000 0.6833 0.01494 0.00612 -0.0338 0.0494 1.0000
5.250 0.7052 0.01558 0.00673 -0.0329 0.0393 1.0000
5.500 0.7277 0.01613 0.00733 -0.0322 0.0333 1.0000
5.750 0.7480 0.01697 0.00818 -0.0311 0.0295 1.0000
6.000 0.7697 0.01765 0.00898 -0.0301 0.0276 1.0000
6.250 0.7906 0.01843 0.00985 -0.0291 0.0260 1.0000
6.500 0.8112 0.01928 0.01077 -0.0280 0.0246 1.0000
6.750 0.8317 0.02021 0.01175 -0.0270 0.0235 1.0000
7.000 0.8520 0.02124 0.01283 -0.0260 0.0225 1.0000
7.250 0.8708 0.02284 0.01448 -0.0248 0.0212 1.0000
7.500 0.8932 0.02376 0.01554 -0.0241 0.0203 1.0000
7.750 0.9151 0.02504 0.01697 -0.0233 0.0197 1.0000
8.000 0.9370 0.02652 0.01863 -0.0224 0.0190 1.0000
8.250 0.9583 0.02819 0.02052 -0.0216 0.0184 1.0000
8.500 0.9785 0.03005 0.02262 -0.0206 0.0179 1.0000
8.750 0.9971 0.03169 0.02452 -0.0195 0.0171 1.0000
9.000 1.0144 0.03301 0.02601 -0.0185 0.0162 1.0000
9.250 1.0303 0.03441 0.02753 -0.0174 0.0155 1.0000
9.500 1.0422 0.03697 0.03032 -0.0160 0.0150 1.0000
9.750 1.0483 0.04035 0.03408 -0.0138 0.0147 1.0000
10.000 1.0520 0.04330 0.03746 -0.0115 0.0145 1.0000
10.250 1.0499 0.04668 0.04126 -0.0088 0.0143 1.0000
10.500 1.0408 0.05018 0.04511 -0.0056 0.0141 1.0000
10.750 1.0266 0.05379 0.04903 -0.0026 0.0140 1.0000
11.000 1.0092 0.05775 0.05327 -0.0005 0.0139 1.0000
11.250 0.9898 0.06219 0.05796 0.0004 0.0139 1.0000
11.500 0.9690 0.06729 0.06328 -0.0002 0.0139 1.0000
11.750 0.9471 0.07319 0.06937 -0.0023 0.0139 1.0000
12.000 0.9247 0.08008 0.07642 -0.0061 0.0140 1.0000
12.250 0.9020 0.08827 0.08474 -0.0115 0.0142 1.0000
12.500 0.8794 0.09825 0.09482 -0.0185 0.0144 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 64A210 (naca64a210-il)