NACA 65-210 (naca65210-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 65-210 (naca65210-il) Reynolds number: 500,000 Max Cl/Cd: 72.94 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca65210-il-500000.txt Download as CSV file: xf-naca65210-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 65-210
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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5299 0.09138 0.08908 -0.0312 1.0000 0.0185
-10.000 -0.5324 0.08480 0.08252 -0.0359 1.0000 0.0185
-9.750 -0.5417 0.07705 0.07473 -0.0426 1.0000 0.0185
-6.750 -0.5414 0.02522 0.02030 -0.0449 0.9826 0.0193
-6.500 -0.5129 0.02228 0.01700 -0.0457 0.9772 0.0192
-6.250 -0.4812 0.02139 0.01616 -0.0477 0.9731 0.0212
-6.000 -0.4485 0.01871 0.01313 -0.0489 0.9697 0.0213
-5.750 -0.4204 0.01717 0.01139 -0.0490 0.9621 0.0220
-5.500 -0.3896 0.01583 0.00989 -0.0497 0.9567 0.0228
-5.250 -0.3630 0.01497 0.00892 -0.0494 0.9482 0.0241
-5.000 -0.3355 0.01439 0.00824 -0.0493 0.9410 0.0254
-4.750 -0.3106 0.01390 0.00766 -0.0487 0.9317 0.0260
-4.500 -0.2893 0.01225 0.00591 -0.0475 0.9233 0.0273
-4.250 -0.2666 0.01145 0.00506 -0.0466 0.9148 0.0291
-4.000 -0.2421 0.01103 0.00462 -0.0461 0.9062 0.0316
-3.500 -0.1917 0.01029 0.00373 -0.0451 0.8903 0.0359
-3.250 -0.1669 0.00974 0.00309 -0.0446 0.8834 0.0408
-3.000 -0.1407 0.00946 0.00278 -0.0443 0.8754 0.0476
-2.750 -0.1157 0.00886 0.00239 -0.0440 0.8684 0.1076
-2.500 -0.0987 0.00675 0.00192 -0.0435 0.8600 0.5509
-2.250 -0.0732 0.00652 0.00191 -0.0430 0.8534 0.6358
-2.000 -0.0463 0.00647 0.00189 -0.0427 0.8462 0.6719
-1.750 -0.0199 0.00645 0.00193 -0.0423 0.8400 0.7056
-1.500 0.0063 0.00647 0.00200 -0.0417 0.8326 0.7371
-1.250 0.0333 0.00652 0.00203 -0.0414 0.8268 0.7552
-1.000 0.0604 0.00653 0.00206 -0.0411 0.8196 0.7686
-0.750 0.0878 0.00655 0.00205 -0.0409 0.8137 0.7784
-0.500 0.1158 0.00655 0.00203 -0.0410 0.8068 0.7856
-0.250 0.1436 0.00655 0.00202 -0.0410 0.8007 0.7917
0.000 0.1715 0.00657 0.00202 -0.0410 0.7944 0.7987
0.250 0.1993 0.00657 0.00203 -0.0410 0.7880 0.8051
0.500 0.2271 0.00660 0.00205 -0.0410 0.7822 0.8120
0.750 0.2549 0.00660 0.00208 -0.0411 0.7756 0.8188
1.000 0.2826 0.00664 0.00211 -0.0410 0.7701 0.8257
1.250 0.3103 0.00665 0.00217 -0.0410 0.7632 0.8328
1.500 0.3377 0.00668 0.00221 -0.0409 0.7565 0.8398
1.750 0.3644 0.00666 0.00221 -0.0405 0.7437 0.8472
2.000 0.3904 0.00666 0.00219 -0.0400 0.7276 0.8547
2.250 0.4165 0.00665 0.00221 -0.0395 0.7096 0.8622
2.500 0.4415 0.00668 0.00219 -0.0387 0.6813 0.8701
2.750 0.4666 0.00674 0.00222 -0.0380 0.6549 0.8779
3.000 0.4919 0.00684 0.00228 -0.0374 0.6251 0.8868
3.250 0.5142 0.00705 0.00237 -0.0362 0.5671 0.8949
3.500 0.5272 0.00823 0.00270 -0.0337 0.3637 0.9053
3.750 0.5348 0.01022 0.00347 -0.0309 0.0876 0.9176
4.000 0.5542 0.01084 0.00391 -0.0294 0.0455 0.9294
4.250 0.5746 0.01118 0.00431 -0.0278 0.0381 0.9419
4.500 0.5954 0.01152 0.00471 -0.0264 0.0336 0.9563
4.750 0.6191 0.01233 0.00561 -0.0258 0.0298 0.9720
5.000 0.6527 0.01287 0.00621 -0.0274 0.0282 0.9868
5.250 0.6813 0.01346 0.00683 -0.0280 0.0261 1.0000
5.500 0.7058 0.01405 0.00745 -0.0278 0.0243 1.0000
5.750 0.7289 0.01489 0.00831 -0.0274 0.0231 1.0000
6.000 0.7500 0.01635 0.00980 -0.0265 0.0220 1.0000
6.250 0.7734 0.01856 0.01209 -0.0260 0.0213 1.0000
6.500 0.7993 0.01929 0.01294 -0.0258 0.0208 1.0000
6.750 0.8248 0.02010 0.01386 -0.0255 0.0199 1.0000
7.000 0.8500 0.02139 0.01529 -0.0251 0.0191 1.0000
7.250 0.8744 0.02317 0.01726 -0.0246 0.0186 1.0000
7.500 0.8973 0.02545 0.01980 -0.0238 0.0184 1.0000
7.750 0.9169 0.02886 0.02359 -0.0225 0.0186 1.0000
8.000 0.9327 0.03235 0.02748 -0.0207 0.0188 1.0000
8.250 0.9373 0.03826 0.03385 -0.0182 0.0204 1.0000
11.750 0.6794 0.10742 0.10550 -0.0197 0.0213 1.0000
12.000 0.6678 0.11667 0.11476 -0.0251 0.0213 1.0000
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Polar data table (+)
Polar graphs
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