NACA 63-215 AIRFOIL (n63215-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 63-215 AIRFOIL (n63215-il) Reynolds number: 50,000 Max Cl/Cd: 29.72 at α=8.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63215-il-50000.txt Download as CSV file: xf-n63215-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63-215 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.6764 0.07608 0.06892 -0.0491 1.0000 0.1409
-10.500 -0.6958 0.07079 0.06360 -0.0504 1.0000 0.1393
-10.250 -0.7220 0.06624 0.05898 -0.0502 1.0000 0.1377
-10.000 -0.7458 0.06196 0.05453 -0.0494 1.0000 0.1362
-9.750 -0.7623 0.05803 0.05037 -0.0481 1.0000 0.1355
-9.500 -0.7718 0.05449 0.04656 -0.0464 1.0000 0.1359
-9.250 -0.7769 0.05132 0.04310 -0.0445 1.0000 0.1371
-9.000 -0.7798 0.04847 0.03992 -0.0421 1.0000 0.1386
-8.750 -0.7810 0.04589 0.03698 -0.0393 1.0000 0.1401
-8.500 -0.7802 0.04357 0.03426 -0.0364 1.0000 0.1415
-8.250 -0.7696 0.04115 0.03183 -0.0344 1.0000 0.1453
-8.000 -0.7596 0.03940 0.03001 -0.0321 1.0000 0.1508
-7.750 -0.7529 0.03765 0.02792 -0.0294 1.0000 0.1556
-7.500 -0.7385 0.03574 0.02601 -0.0276 1.0000 0.1615
-7.250 -0.7261 0.03431 0.02443 -0.0254 1.0000 0.1702
-7.000 -0.7101 0.03279 0.02302 -0.0236 1.0000 0.1807
-6.750 -0.6945 0.03137 0.02166 -0.0216 1.0000 0.1940
-6.500 -0.6814 0.03007 0.02050 -0.0194 1.0000 0.2130
-6.250 -0.6704 0.02865 0.01935 -0.0169 1.0000 0.2400
-6.000 -0.6645 0.02695 0.01823 -0.0141 1.0000 0.2879
-5.750 -0.6655 0.02554 0.01812 -0.0095 1.0000 0.4085
-5.500 -0.6611 0.02722 0.02027 -0.0027 1.0000 0.5147
-5.250 -0.6515 0.02931 0.02236 0.0036 1.0000 0.5627
-5.000 -0.6409 0.03126 0.02426 0.0096 1.0000 0.5964
-4.750 -0.6300 0.03308 0.02601 0.0156 1.0000 0.6253
-4.500 -0.6210 0.03431 0.02714 0.0209 1.0000 0.6537
-4.250 -0.6070 0.03621 0.02897 0.0273 1.0000 0.6776
-4.000 -0.5991 0.03674 0.02939 0.0319 1.0000 0.7036
-3.750 -0.5820 0.03787 0.03041 0.0366 1.0000 0.7256
-3.500 -0.5690 0.03817 0.03058 0.0401 1.0000 0.7477
-3.250 -0.5618 0.03788 0.03016 0.0432 1.0000 0.7687
-3.000 -0.5468 0.03786 0.03002 0.0458 1.0000 0.7879
-2.750 -0.5293 0.03776 0.02979 0.0475 1.0000 0.8059
-2.500 -0.5135 0.03743 0.02933 0.0488 1.0000 0.8227
-2.250 -0.4709 0.03764 0.02932 0.0452 0.9935 0.8392
-2.000 -0.4292 0.03770 0.02919 0.0411 0.9855 0.8542
-1.750 -0.3906 0.03770 0.02902 0.0374 0.9772 0.8683
-1.500 -0.3322 0.03807 0.02919 0.0304 0.9693 0.8810
-1.250 -0.2830 0.03814 0.02910 0.0248 0.9610 0.8921
-1.000 -0.2326 0.03832 0.02915 0.0187 0.9531 0.9037
-0.750 -0.1975 0.03831 0.02904 0.0152 0.9451 0.9154
-0.500 -0.1327 0.03865 0.02925 0.0064 0.9370 0.9252
-0.250 -0.0826 0.03882 0.02934 0.0001 0.9287 0.9353
0.000 -0.0399 0.03903 0.02949 -0.0049 0.9209 0.9459
0.250 0.0111 0.03929 0.02970 -0.0115 0.9125 0.9562
0.500 0.0622 0.03958 0.02996 -0.0182 0.9043 0.9657
0.750 0.1160 0.03991 0.03028 -0.0253 0.8959 0.9757
1.000 0.1621 0.04026 0.03064 -0.0313 0.8873 0.9859
1.250 0.2211 0.04067 0.03107 -0.0394 0.8789 0.9958
1.500 0.2359 0.04110 0.03152 -0.0401 0.8704 1.0000
1.750 0.2468 0.04156 0.03198 -0.0397 0.8626 1.0000
2.000 0.2367 0.04199 0.03241 -0.0360 0.8550 1.0000
2.250 0.2372 0.04241 0.03283 -0.0340 0.8475 1.0000
2.500 0.2317 0.04283 0.03325 -0.0310 0.8405 1.0000
2.750 0.2146 0.04304 0.03344 -0.0264 0.8340 1.0000
3.000 0.2249 0.04360 0.03400 -0.0256 0.8265 1.0000
3.250 0.1952 0.04357 0.03392 -0.0192 0.8214 1.0000
3.500 0.2081 0.04427 0.03462 -0.0190 0.8135 1.0000
3.750 0.2189 0.04513 0.03545 -0.0188 0.8050 1.0000
4.000 0.2568 0.04644 0.03679 -0.0217 0.7937 1.0000
4.250 0.2618 0.04736 0.03770 -0.0210 0.7846 1.0000
4.500 0.2890 0.04862 0.03900 -0.0225 0.7723 1.0000
4.750 0.3227 0.04994 0.04037 -0.0246 0.7588 1.0000
5.000 0.3525 0.05117 0.04167 -0.0261 0.7449 1.0000
5.250 0.3680 0.05238 0.04293 -0.0261 0.7309 1.0000
5.500 0.3936 0.05361 0.04422 -0.0269 0.7149 1.0000
5.750 0.4232 0.05473 0.04544 -0.0278 0.6976 1.0000
6.000 0.4587 0.05569 0.04654 -0.0289 0.6793 1.0000
6.250 0.4901 0.05658 0.04754 -0.0293 0.6604 1.0000
6.500 0.5065 0.05769 0.04874 -0.0288 0.6408 1.0000
6.750 0.5315 0.05853 0.04973 -0.0285 0.6202 1.0000
7.000 0.5784 0.05824 0.04965 -0.0288 0.5983 1.0000
7.250 0.6028 0.05867 0.05025 -0.0278 0.5754 1.0000
7.500 0.6545 0.05685 0.04871 -0.0268 0.5504 1.0000
7.750 0.6941 0.05501 0.04712 -0.0246 0.5230 1.0000
8.000 0.8742 0.03612 0.02905 -0.0190 0.4769 1.0000
8.250 0.9176 0.03088 0.02361 -0.0139 0.4035 1.0000
8.500 0.9176 0.03106 0.02304 -0.0085 0.3203 1.0000
8.750 0.9170 0.03303 0.02418 -0.0048 0.2602 1.0000
9.000 0.9317 0.03529 0.02581 -0.0031 0.2177 1.0000
9.250 0.9624 0.03765 0.02782 -0.0031 0.1871 1.0000
9.500 0.9925 0.04007 0.03026 -0.0032 0.1684 1.0000
9.750 1.0237 0.04265 0.03283 -0.0036 0.1545 1.0000
10.000 1.0558 0.04544 0.03558 -0.0042 0.1439 1.0000
10.250 1.0664 0.04826 0.03884 -0.0028 0.1384 1.0000
10.500 1.0849 0.05106 0.04179 -0.0022 0.1321 1.0000
10.750 1.1039 0.05478 0.04564 -0.0021 0.1279 1.0000
11.000 1.0988 0.05802 0.04936 0.0002 0.1262 1.0000
11.250 1.0896 0.06147 0.05320 0.0024 0.1248 1.0000
11.500 1.0737 0.06491 0.05696 0.0048 0.1238 1.0000
11.750 1.0535 0.06876 0.06107 0.0067 0.1237 1.0000
12.000 1.0290 0.07323 0.06579 0.0075 0.1242 1.0000
12.250 1.0002 0.07854 0.07131 0.0072 0.1251 1.0000
12.500 0.9705 0.08475 0.07769 0.0056 0.1263 1.0000
12.750 0.9420 0.09184 0.08488 0.0030 0.1276 1.0000
13.000 0.9197 0.09937 0.09248 0.0000 0.1286 1.0000
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Polar data table (+)
Polar graphs
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