NACA 63-215 AIRFOIL (n63215-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 63-215 AIRFOIL (n63215-il) Reynolds number: 1,000,000 Max Cl/Cd: 80.8 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n63215-il-1000000-n5.txt Download as CSV file: xf-n63215-il-1000000-n5.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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-19.750 -1.1219 0.10543 0.10235 -0.0207 1.0000 0.0098
-19.500 -1.1409 0.09801 0.09480 -0.0249 1.0000 0.0099
-19.250 -1.1577 0.09118 0.08785 -0.0287 1.0000 0.0099
-19.000 -1.1720 0.08502 0.08158 -0.0321 1.0000 0.0100
-18.750 -1.1844 0.07932 0.07576 -0.0351 1.0000 0.0100
-18.500 -1.1950 0.07408 0.07042 -0.0378 1.0000 0.0101
-18.250 -1.2029 0.06942 0.06566 -0.0401 1.0000 0.0101
-18.000 -1.2088 0.06517 0.06131 -0.0421 1.0000 0.0102
-17.750 -1.2120 0.06142 0.05747 -0.0437 1.0000 0.0104
-17.500 -1.2175 0.05747 0.05342 -0.0454 1.0000 0.0104
-17.250 -1.2173 0.05436 0.05022 -0.0466 1.0000 0.0105
-17.000 -1.2180 0.05124 0.04703 -0.0477 1.0000 0.0107
-16.750 -1.2188 0.04820 0.04390 -0.0487 1.0000 0.0107
-16.500 -1.2175 0.04551 0.04112 -0.0494 1.0000 0.0109
-16.250 -1.2161 0.04295 0.03848 -0.0500 1.0000 0.0110
-16.000 -1.2114 0.04083 0.03628 -0.0503 1.0000 0.0111
-15.750 -1.2075 0.03864 0.03402 -0.0506 1.0000 0.0112
-15.500 -1.2028 0.03660 0.03190 -0.0508 1.0000 0.0113
-15.250 -1.1989 0.03454 0.02977 -0.0509 1.0000 0.0114
-15.000 -1.1956 0.03248 0.02764 -0.0508 1.0000 0.0116
-14.750 -1.1902 0.03067 0.02576 -0.0506 1.0000 0.0118
-14.500 -1.1831 0.02907 0.02412 -0.0503 1.0000 0.0120
-14.250 -1.1747 0.02762 0.02261 -0.0499 1.0000 0.0121
-14.000 -1.1655 0.02629 0.02123 -0.0494 1.0000 0.0124
-13.750 -1.1557 0.02504 0.01993 -0.0488 1.0000 0.0125
-13.500 -1.1452 0.02389 0.01874 -0.0481 1.0000 0.0128
-13.250 -1.1338 0.02285 0.01765 -0.0473 1.0000 0.0130
-13.000 -1.1225 0.02185 0.01660 -0.0465 1.0000 0.0132
-12.750 -1.1106 0.02094 0.01565 -0.0455 1.0000 0.0134
-12.500 -1.0991 0.02008 0.01474 -0.0444 1.0000 0.0136
-12.250 -1.0864 0.01936 0.01397 -0.0431 1.0000 0.0140
-12.000 -1.0510 0.01844 0.01298 -0.0466 0.9484 0.0144
-11.750 -1.0357 0.01781 0.01221 -0.0455 0.9096 0.0148
-11.500 -1.0269 0.01735 0.01164 -0.0429 0.8858 0.0151
-11.250 -1.0116 0.01689 0.01110 -0.0414 0.8700 0.0156
-11.000 -0.9938 0.01646 0.01060 -0.0402 0.8568 0.0160
-10.750 -0.9749 0.01604 0.01010 -0.0392 0.8455 0.0165
-10.500 -0.9544 0.01565 0.00963 -0.0384 0.8348 0.0170
-10.250 -0.9334 0.01525 0.00917 -0.0376 0.8256 0.0176
-10.000 -0.9120 0.01485 0.00874 -0.0369 0.8161 0.0185
-9.750 -0.8893 0.01449 0.00834 -0.0364 0.8079 0.0192
-9.500 -0.8661 0.01416 0.00795 -0.0359 0.7995 0.0201
-9.250 -0.8420 0.01385 0.00758 -0.0356 0.7920 0.0208
-9.000 -0.8181 0.01351 0.00720 -0.0352 0.7843 0.0217
-8.750 -0.7937 0.01319 0.00686 -0.0349 0.7773 0.0228
-8.500 -0.7687 0.01290 0.00653 -0.0347 0.7696 0.0238
-8.250 -0.7433 0.01265 0.00622 -0.0345 0.7628 0.0248
-8.000 -0.7175 0.01238 0.00592 -0.0343 0.7558 0.0258
-7.750 -0.6921 0.01209 0.00560 -0.0342 0.7487 0.0273
-7.500 -0.6660 0.01183 0.00531 -0.0341 0.7422 0.0286
-7.250 -0.6396 0.01160 0.00505 -0.0340 0.7351 0.0298
-7.000 -0.6129 0.01139 0.00480 -0.0340 0.7287 0.0308
-6.750 -0.5862 0.01113 0.00452 -0.0339 0.7221 0.0328
-6.500 -0.5594 0.01091 0.00427 -0.0339 0.7155 0.0348
-6.250 -0.5321 0.01071 0.00405 -0.0340 0.7093 0.0366
-6.000 -0.5051 0.01048 0.00382 -0.0340 0.7025 0.0403
-5.750 -0.4779 0.01027 0.00360 -0.0341 0.6965 0.0456
-5.500 -0.4506 0.01001 0.00338 -0.0342 0.6903 0.0545
-5.250 -0.4234 0.00977 0.00316 -0.0342 0.6836 0.0653
-5.000 -0.3960 0.00951 0.00295 -0.0343 0.6777 0.0800
-4.750 -0.3686 0.00923 0.00273 -0.0345 0.6715 0.1009
-4.500 -0.3414 0.00893 0.00252 -0.0346 0.6653 0.1286
-4.250 -0.3138 0.00863 0.00233 -0.0348 0.6594 0.1574
-4.000 -0.2863 0.00831 0.00213 -0.0350 0.6530 0.1923
-3.750 -0.2589 0.00797 0.00193 -0.0352 0.6473 0.2361
-3.500 -0.2314 0.00757 0.00173 -0.0355 0.6418 0.2898
-3.250 -0.2037 0.00724 0.00156 -0.0357 0.6355 0.3398
-3.000 -0.1757 0.00696 0.00144 -0.0360 0.6297 0.3898
-2.750 -0.1471 0.00678 0.00136 -0.0363 0.6240 0.4260
-2.500 -0.1184 0.00664 0.00130 -0.0365 0.6183 0.4582
-2.250 -0.0895 0.00656 0.00126 -0.0368 0.6130 0.4776
-2.000 -0.0602 0.00652 0.00123 -0.0371 0.6070 0.4883
-1.750 -0.0311 0.00648 0.00120 -0.0374 0.6011 0.5023
-1.500 -0.0019 0.00645 0.00119 -0.0376 0.5962 0.5152
-1.000 0.0567 0.00643 0.00115 -0.0382 0.5849 0.5301
-0.750 0.0860 0.00644 0.00115 -0.0385 0.5797 0.5369
-0.500 0.1154 0.00643 0.00115 -0.0388 0.5743 0.5433
-0.250 0.1446 0.00644 0.00115 -0.0390 0.5687 0.5502
0.000 0.1739 0.00645 0.00116 -0.0393 0.5634 0.5564
0.250 0.2032 0.00645 0.00117 -0.0396 0.5576 0.5625
0.500 0.2323 0.00648 0.00119 -0.0398 0.5514 0.5687
0.750 0.2616 0.00650 0.00121 -0.0401 0.5456 0.5740
1.000 0.2908 0.00652 0.00124 -0.0404 0.5390 0.5797
1.250 0.3199 0.00656 0.00127 -0.0406 0.5330 0.5854
1.500 0.3492 0.00658 0.00130 -0.0409 0.5270 0.5909
1.750 0.3782 0.00662 0.00135 -0.0412 0.5204 0.5969
2.000 0.4073 0.00666 0.00139 -0.0414 0.5138 0.6025
2.250 0.4361 0.00673 0.00145 -0.0416 0.5036 0.6077
2.500 0.4648 0.00679 0.00151 -0.0418 0.4905 0.6140
2.750 0.4936 0.00687 0.00158 -0.0420 0.4798 0.6201
3.000 0.5221 0.00696 0.00165 -0.0422 0.4691 0.6257
3.250 0.5503 0.00708 0.00174 -0.0423 0.4514 0.6317
3.500 0.5778 0.00729 0.00185 -0.0424 0.4229 0.6377
3.750 0.6052 0.00749 0.00198 -0.0424 0.3958 0.6439
4.000 0.6309 0.00787 0.00218 -0.0422 0.3483 0.6505
4.250 0.6561 0.00832 0.00243 -0.0420 0.3013 0.6561
4.500 0.6811 0.00876 0.00271 -0.0418 0.2578 0.6624
4.750 0.7057 0.00924 0.00301 -0.0415 0.2143 0.6692
5.000 0.7307 0.00966 0.00330 -0.0413 0.1802 0.6754
5.250 0.7558 0.01005 0.00358 -0.0410 0.1517 0.6821
5.500 0.7806 0.01047 0.00388 -0.0407 0.1247 0.6884
5.750 0.8047 0.01092 0.00422 -0.0403 0.0969 0.6951
6.000 0.8288 0.01135 0.00455 -0.0400 0.0753 0.7022
6.250 0.8536 0.01169 0.00485 -0.0397 0.0620 0.7084
6.500 0.8782 0.01204 0.00517 -0.0393 0.0513 0.7156
6.750 0.9028 0.01238 0.00548 -0.0390 0.0432 0.7224
7.000 0.9272 0.01270 0.00580 -0.0386 0.0376 0.7293
7.250 0.9516 0.01302 0.00612 -0.0382 0.0341 0.7365
7.500 0.9759 0.01332 0.00645 -0.0378 0.0315 0.7433
7.750 1.0001 0.01361 0.00677 -0.0374 0.0297 0.7510
8.000 1.0234 0.01395 0.00712 -0.0368 0.0279 0.7579
8.250 1.0465 0.01428 0.00749 -0.0363 0.0264 0.7659
8.750 1.0919 0.01493 0.00822 -0.0350 0.0241 0.7815
9.250 1.1335 0.01570 0.00906 -0.0332 0.0217 0.7987
9.500 1.1541 0.01605 0.00946 -0.0323 0.0208 0.8075
9.750 1.1731 0.01644 0.00991 -0.0311 0.0199 0.8176
10.000 1.1884 0.01685 0.01036 -0.0293 0.0188 0.8281
10.250 1.2015 0.01732 0.01088 -0.0272 0.0178 0.8401
10.500 1.2149 0.01777 0.01142 -0.0252 0.0172 0.8540
10.750 1.2274 0.01829 0.01201 -0.0232 0.0165 0.8705
11.000 1.2381 0.01885 0.01267 -0.0210 0.0157 0.8925
11.250 1.2462 0.01943 0.01337 -0.0184 0.0151 0.9333
11.750 1.2757 0.02094 0.01501 -0.0165 0.0140 1.0000
12.000 1.2887 0.02181 0.01591 -0.0154 0.0134 1.0000
12.250 1.3010 0.02274 0.01687 -0.0143 0.0129 1.0000
12.500 1.3124 0.02377 0.01793 -0.0132 0.0125 1.0000
12.750 1.3233 0.02489 0.01908 -0.0122 0.0122 1.0000
13.000 1.3329 0.02612 0.02036 -0.0112 0.0119 1.0000
13.250 1.3430 0.02737 0.02166 -0.0103 0.0116 1.0000
13.500 1.3528 0.02868 0.02302 -0.0095 0.0115 1.0000
13.750 1.3620 0.03007 0.02447 -0.0088 0.0113 1.0000
14.000 1.3706 0.03155 0.02600 -0.0081 0.0112 1.0000
14.250 1.3787 0.03311 0.02762 -0.0076 0.0110 1.0000
14.500 1.3865 0.03474 0.02931 -0.0070 0.0108 1.0000
14.750 1.3932 0.03651 0.03114 -0.0066 0.0106 1.0000
15.000 1.3990 0.03841 0.03311 -0.0062 0.0105 1.0000
15.250 1.4043 0.04040 0.03516 -0.0060 0.0104 1.0000
15.500 1.4091 0.04252 0.03734 -0.0058 0.0102 1.0000
15.750 1.4131 0.04478 0.03967 -0.0058 0.0102 1.0000
16.000 1.4167 0.04717 0.04212 -0.0059 0.0100 1.0000
16.250 1.4191 0.04977 0.04479 -0.0062 0.0098 1.0000
16.500 1.4200 0.05261 0.04771 -0.0065 0.0097 1.0000
16.750 1.4203 0.05563 0.05080 -0.0071 0.0096 1.0000
17.000 1.4199 0.05882 0.05408 -0.0078 0.0095 1.0000
17.250 1.4174 0.06241 0.05776 -0.0088 0.0094 1.0000
17.500 1.4156 0.06596 0.06140 -0.0098 0.0093 1.0000
17.750 1.4156 0.06937 0.06490 -0.0109 0.0092 1.0000
18.000 1.4136 0.07317 0.06879 -0.0122 0.0092 1.0000
18.250 1.4108 0.07718 0.07290 -0.0138 0.0091 1.0000
18.500 1.4066 0.08147 0.07729 -0.0155 0.0091 1.0000
18.750 1.4021 0.08594 0.08187 -0.0174 0.0090 1.0000
19.000 1.3956 0.09085 0.08688 -0.0196 0.0090 1.0000
19.250 1.3881 0.09603 0.09217 -0.0221 0.0090 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 63-215 AIRFOIL (n63215-il)