NACA 63-210 AIRFOIL (n63210-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 63-210 AIRFOIL (n63210-il) Reynolds number: 500,000 Max Cl/Cd: 67.69 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63210-il-500000.txt Download as CSV file: xf-n63210-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63-210 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.6413 0.06088 0.05855 -0.0411 1.0000 0.0149
-9.500 -0.6635 0.05575 0.05331 -0.0438 1.0000 0.0150
-9.250 -0.6833 0.05158 0.04900 -0.0437 1.0000 0.0148
-9.000 -0.6996 0.04596 0.04312 -0.0435 1.0000 0.0141
-8.750 -0.7056 0.04060 0.03741 -0.0427 1.0000 0.0145
-8.500 -0.7083 0.03497 0.03132 -0.0414 1.0000 0.0147
-8.250 -0.7045 0.03008 0.02589 -0.0395 1.0000 0.0153
-8.000 -0.6898 0.02839 0.02378 -0.0380 1.0000 0.0162
-7.750 -0.6830 0.02383 0.01889 -0.0364 1.0000 0.0174
-7.500 -0.6637 0.02328 0.01833 -0.0355 1.0000 0.0185
-7.250 -0.6457 0.02222 0.01710 -0.0341 1.0000 0.0201
-7.000 -0.6268 0.02176 0.01646 -0.0325 1.0000 0.0216
-6.750 -0.6108 0.01907 0.01352 -0.0314 0.9995 0.0237
-6.500 -0.5745 0.01857 0.01301 -0.0339 0.9961 0.0259
-6.250 -0.5383 0.01760 0.01191 -0.0359 0.9926 0.0276
-6.000 -0.5017 0.01730 0.01147 -0.0379 0.9884 0.0293
-5.750 -0.4661 0.01569 0.00970 -0.0399 0.9850 0.0303
-5.500 -0.4303 0.01443 0.00837 -0.0418 0.9816 0.0303
-5.250 -0.3972 0.01342 0.00731 -0.0431 0.9749 0.0304
-5.000 -0.3624 0.01253 0.00636 -0.0448 0.9698 0.0305
-4.750 -0.3335 0.01175 0.00554 -0.0452 0.9590 0.0311
-4.500 -0.3060 0.01115 0.00488 -0.0453 0.9471 0.0315
-4.250 -0.2799 0.01067 0.00434 -0.0449 0.9336 0.0318
-4.000 -0.2545 0.01024 0.00383 -0.0444 0.9193 0.0324
-3.750 -0.2289 0.00989 0.00341 -0.0439 0.9051 0.0332
-3.500 -0.2031 0.00960 0.00305 -0.0433 0.8921 0.0347
-3.250 -0.1768 0.00939 0.00274 -0.0429 0.8804 0.0368
-3.000 -0.1502 0.00921 0.00247 -0.0425 0.8699 0.0413
-2.750 -0.1241 0.00863 0.00216 -0.0424 0.8594 0.1161
-2.500 -0.0998 0.00731 0.00179 -0.0428 0.8494 0.3751
-2.250 -0.0739 0.00678 0.00171 -0.0428 0.8405 0.5113
-2.000 -0.0464 0.00671 0.00175 -0.0427 0.8310 0.5724
-1.750 -0.0186 0.00665 0.00172 -0.0426 0.8221 0.6050
-1.500 0.0084 0.00654 0.00168 -0.0423 0.8137 0.6377
-1.250 0.0359 0.00646 0.00171 -0.0421 0.8043 0.6655
-1.000 0.0636 0.00646 0.00171 -0.0420 0.7957 0.6844
-0.750 0.0916 0.00645 0.00170 -0.0419 0.7868 0.6990
-0.500 0.1194 0.00645 0.00171 -0.0418 0.7776 0.7153
-0.250 0.1469 0.00646 0.00172 -0.0416 0.7693 0.7318
0.000 0.1751 0.00646 0.00174 -0.0415 0.7597 0.7434
0.250 0.2032 0.00648 0.00174 -0.0415 0.7510 0.7541
0.500 0.2309 0.00649 0.00176 -0.0414 0.7420 0.7643
0.750 0.2588 0.00649 0.00180 -0.0413 0.7319 0.7751
1.000 0.2863 0.00650 0.00180 -0.0411 0.7178 0.7868
1.250 0.3137 0.00653 0.00182 -0.0409 0.7016 0.7990
1.500 0.3408 0.00655 0.00185 -0.0406 0.6869 0.8107
1.750 0.3675 0.00659 0.00187 -0.0402 0.6682 0.8225
2.000 0.3941 0.00663 0.00191 -0.0398 0.6451 0.8346
2.250 0.4202 0.00671 0.00194 -0.0393 0.6163 0.8467
2.500 0.4460 0.00684 0.00199 -0.0388 0.5823 0.8591
2.750 0.4714 0.00700 0.00207 -0.0382 0.5432 0.8717
3.000 0.4948 0.00731 0.00219 -0.0373 0.4795 0.8841
3.250 0.5152 0.00803 0.00243 -0.0360 0.3570 0.8974
3.500 0.5309 0.00948 0.00299 -0.0345 0.1533 0.9124
3.750 0.5490 0.01046 0.00348 -0.0329 0.0434 0.9285
4.000 0.5704 0.01072 0.00375 -0.0315 0.0344 0.9460
4.250 0.5954 0.01094 0.00404 -0.0308 0.0318 0.9672
4.500 0.6295 0.01126 0.00441 -0.0323 0.0304 1.0000
4.750 0.6572 0.01167 0.00485 -0.0325 0.0293 1.0000
5.000 0.6841 0.01214 0.00537 -0.0326 0.0283 1.0000
5.250 0.7103 0.01267 0.00593 -0.0326 0.0277 1.0000
5.500 0.7357 0.01328 0.00658 -0.0324 0.0272 1.0000
5.750 0.7604 0.01395 0.00730 -0.0320 0.0269 1.0000
6.000 0.7846 0.01471 0.00812 -0.0316 0.0268 1.0000
6.250 0.8085 0.01550 0.00896 -0.0311 0.0265 1.0000
6.500 0.8321 0.01637 0.00986 -0.0306 0.0261 1.0000
6.750 0.8559 0.01733 0.01088 -0.0301 0.0260 1.0000
7.000 0.8799 0.01840 0.01202 -0.0296 0.0260 1.0000
7.250 0.9042 0.01961 0.01332 -0.0292 0.0260 1.0000
7.500 0.9286 0.02095 0.01475 -0.0287 0.0258 1.0000
7.750 0.9516 0.02325 0.01714 -0.0284 0.0247 1.0000
8.000 0.9733 0.02522 0.01936 -0.0278 0.0236 1.0000
8.250 0.9963 0.02626 0.02058 -0.0272 0.0229 1.0000
8.500 1.0180 0.02761 0.02214 -0.0265 0.0220 1.0000
8.750 1.0395 0.02843 0.02307 -0.0259 0.0204 1.0000
9.000 1.0607 0.02898 0.02362 -0.0256 0.0192 1.0000
9.250 1.0686 0.03425 0.02922 -0.0243 0.0177 1.0000
9.500 1.0849 0.03457 0.02983 -0.0229 0.0167 1.0000
9.750 1.0979 0.03656 0.03211 -0.0214 0.0155 1.0000
10.000 1.1087 0.03849 0.03426 -0.0200 0.0146 1.0000
10.250 1.1214 0.03969 0.03555 -0.0189 0.0140 1.0000
10.500 1.1359 0.04038 0.03621 -0.0182 0.0134 1.0000
10.750 1.1331 0.04376 0.03978 -0.0162 0.0130 1.0000
11.000 1.1097 0.04858 0.04493 -0.0124 0.0128 1.0000
11.250 1.0798 0.05343 0.05006 -0.0094 0.0127 1.0000
11.500 1.0602 0.05744 0.05428 -0.0084 0.0127 1.0000
11.750 1.0377 0.06231 0.05935 -0.0088 0.0127 1.0000
12.000 1.0131 0.06813 0.06537 -0.0108 0.0127 1.0000
12.250 1.0034 0.07239 0.06975 -0.0130 0.0127 1.0000
12.500 0.9846 0.07874 0.07626 -0.0169 0.0128 1.0000
12.750 0.9712 0.08503 0.08268 -0.0214 0.0129 1.0000
13.000 0.8258 0.14073 0.13865 -0.0537 0.0195 1.0000
13.250 0.8288 0.14514 0.14306 -0.0560 0.0190 1.0000
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Polar data table (+)
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