NACA 63-210 AIRFOIL (n63210-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 63-210 AIRFOIL (n63210-il) Reynolds number: 200,000 Max Cl/Cd: 48.69 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n63210-il-200000-n5.txt Download as CSV file: xf-n63210-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.6022 0.08750 0.08397 -0.0195 1.0000 0.0142
-10.250 -0.6864 0.05894 0.05523 -0.0398 1.0000 0.0127
-10.000 -0.7079 0.05365 0.04979 -0.0429 1.0000 0.0128
-9.750 -0.7270 0.04971 0.04568 -0.0430 1.0000 0.0128
-9.500 -0.7361 0.04603 0.04177 -0.0428 1.0000 0.0129
-9.250 -0.7344 0.04340 0.03897 -0.0424 1.0000 0.0132
-9.000 -0.7295 0.04075 0.03610 -0.0419 1.0000 0.0136
-8.750 -0.7210 0.03824 0.03336 -0.0413 1.0000 0.0141
-8.500 -0.7108 0.03562 0.03045 -0.0405 1.0000 0.0149
-8.250 -0.6988 0.03291 0.02736 -0.0396 1.0000 0.0164
-8.000 -0.6839 0.03061 0.02455 -0.0385 1.0000 0.0181
-7.750 -0.6710 0.02757 0.02114 -0.0375 1.0000 0.0197
-7.500 -0.6527 0.02619 0.01961 -0.0367 1.0000 0.0210
-7.250 -0.6337 0.02487 0.01809 -0.0357 1.0000 0.0224
-7.000 -0.6141 0.02380 0.01681 -0.0347 1.0000 0.0241
-6.750 -0.5938 0.02312 0.01590 -0.0336 1.0000 0.0260
-6.500 -0.5754 0.02167 0.01429 -0.0325 1.0000 0.0276
-6.250 -0.5490 0.02063 0.01320 -0.0329 0.9968 0.0291
-6.000 -0.5153 0.01943 0.01182 -0.0345 0.9906 0.0294
-5.750 -0.4807 0.01826 0.01052 -0.0362 0.9852 0.0294
-5.500 -0.4478 0.01727 0.00944 -0.0375 0.9780 0.0293
-5.250 -0.4148 0.01639 0.00849 -0.0389 0.9707 0.0294
-5.000 -0.3817 0.01562 0.00763 -0.0404 0.9632 0.0295
-4.750 -0.3508 0.01495 0.00690 -0.0413 0.9533 0.0296
-4.500 -0.3202 0.01437 0.00626 -0.0422 0.9430 0.0299
-4.250 -0.2901 0.01387 0.00567 -0.0429 0.9320 0.0303
-4.000 -0.2605 0.01344 0.00514 -0.0433 0.9202 0.0308
-3.750 -0.2317 0.01308 0.00469 -0.0436 0.9076 0.0316
-3.500 -0.2030 0.01275 0.00427 -0.0438 0.8958 0.0327
-3.250 -0.1745 0.01247 0.00390 -0.0439 0.8850 0.0345
-3.000 -0.1466 0.01219 0.00356 -0.0438 0.8750 0.0393
-2.750 -0.1194 0.01183 0.00325 -0.0437 0.8655 0.0643
-2.500 -0.0962 0.01039 0.00281 -0.0440 0.8553 0.3047
-2.250 -0.0730 0.00951 0.00282 -0.0436 0.8456 0.5222
-2.000 -0.0463 0.00945 0.00280 -0.0432 0.8366 0.5856
-1.750 -0.0194 0.00937 0.00274 -0.0428 0.8267 0.6104
-1.500 0.0065 0.00927 0.00277 -0.0422 0.8173 0.6426
-1.250 0.0318 0.00926 0.00287 -0.0413 0.8084 0.6873
-1.000 0.0580 0.00926 0.00293 -0.0407 0.7987 0.7135
-0.750 0.0854 0.00926 0.00289 -0.0405 0.7896 0.7233
-0.500 0.1128 0.00926 0.00285 -0.0403 0.7810 0.7335
-0.250 0.1401 0.00925 0.00285 -0.0400 0.7714 0.7435
0.000 0.1673 0.00926 0.00284 -0.0398 0.7629 0.7535
0.250 0.1945 0.00927 0.00285 -0.0395 0.7538 0.7642
0.500 0.2217 0.00929 0.00288 -0.0393 0.7447 0.7752
0.750 0.2485 0.00931 0.00291 -0.0388 0.7363 0.7859
1.000 0.2754 0.00933 0.00297 -0.0385 0.7266 0.7971
1.250 0.3020 0.00936 0.00303 -0.0381 0.7160 0.8087
1.500 0.3274 0.00938 0.00299 -0.0373 0.6936 0.8207
1.750 0.3522 0.00940 0.00297 -0.0363 0.6665 0.8323
2.000 0.3776 0.00945 0.00304 -0.0356 0.6498 0.8441
2.250 0.4031 0.00950 0.00311 -0.0349 0.6348 0.8561
2.500 0.4282 0.00955 0.00320 -0.0342 0.6167 0.8685
2.750 0.4523 0.00965 0.00324 -0.0331 0.5864 0.8816
3.000 0.4749 0.00984 0.00328 -0.0318 0.5337 0.8956
3.250 0.4961 0.01019 0.00334 -0.0304 0.4560 0.9113
3.500 0.5131 0.01117 0.00358 -0.0288 0.2977 0.9308
3.750 0.5320 0.01259 0.00416 -0.0281 0.1219 0.9537
4.000 0.5614 0.01346 0.00470 -0.0291 0.0494 0.9803
4.250 0.5895 0.01394 0.00513 -0.0296 0.0369 1.0000
4.500 0.6156 0.01437 0.00557 -0.0295 0.0325 1.0000
4.750 0.6412 0.01487 0.00607 -0.0294 0.0298 1.0000
5.000 0.6666 0.01539 0.00667 -0.0293 0.0287 1.0000
5.250 0.6918 0.01591 0.00728 -0.0291 0.0281 1.0000
5.500 0.7163 0.01651 0.00795 -0.0288 0.0275 1.0000
5.750 0.7402 0.01717 0.00868 -0.0284 0.0270 1.0000
6.000 0.7635 0.01791 0.00949 -0.0279 0.0265 1.0000
6.250 0.7865 0.01871 0.01038 -0.0274 0.0261 1.0000
6.500 0.8091 0.01960 0.01133 -0.0268 0.0257 1.0000
6.750 0.8317 0.02055 0.01235 -0.0262 0.0254 1.0000
7.000 0.8544 0.02159 0.01347 -0.0256 0.0252 1.0000
7.250 0.8774 0.02272 0.01469 -0.0251 0.0250 1.0000
7.500 0.9008 0.02397 0.01604 -0.0246 0.0248 1.0000
7.750 0.9242 0.02534 0.01758 -0.0241 0.0247 1.0000
8.000 0.9476 0.02688 0.01928 -0.0236 0.0246 1.0000
8.250 0.9702 0.02850 0.02111 -0.0231 0.0244 1.0000
8.500 0.9914 0.02976 0.02253 -0.0225 0.0234 1.0000
8.750 1.0113 0.03094 0.02378 -0.0221 0.0221 1.0000
9.000 1.0274 0.03391 0.02698 -0.0214 0.0209 1.0000
9.250 1.0438 0.03554 0.02895 -0.0203 0.0203 1.0000
9.500 1.0589 0.03709 0.03085 -0.0191 0.0192 1.0000
9.750 1.0708 0.03915 0.03326 -0.0178 0.0179 1.0000
10.000 1.0789 0.04166 0.03609 -0.0162 0.0170 1.0000
10.250 1.0845 0.04403 0.03872 -0.0147 0.0162 1.0000
10.500 1.0862 0.04644 0.04137 -0.0129 0.0156 1.0000
10.750 1.0828 0.04864 0.04376 -0.0107 0.0152 1.0000
11.000 1.0770 0.05089 0.04615 -0.0088 0.0148 1.0000
11.250 1.0668 0.05389 0.04933 -0.0074 0.0145 1.0000
11.500 1.0526 0.05763 0.05324 -0.0070 0.0142 1.0000
11.750 1.0327 0.06250 0.05833 -0.0076 0.0140 1.0000
12.000 1.0107 0.06820 0.06424 -0.0096 0.0139 1.0000
12.250 0.9913 0.07423 0.07048 -0.0127 0.0140 1.0000
12.500 0.9716 0.08115 0.07760 -0.0171 0.0141 1.0000
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Polar data table (+)
Polar graphs
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