NACA 65(1)-212 a=0.6 (naca651212a06-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 65(1)-212 a=0.6 (naca651212a06-il) Reynolds number: 50,000 Max Cl/Cd: 29.47 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca651212a06-il-50000-n5.txt Download as CSV file: xf-naca651212a06-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 65(1)-212 a=0.6
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.5124 0.12761 0.12036 -0.0243 1.0000 0.0718
-11.750 -0.4992 0.12331 0.11603 -0.0239 1.0000 0.0670
-11.500 -0.5176 0.11530 0.10812 -0.0316 1.0000 0.0592
-11.250 -0.5105 0.11123 0.10406 -0.0321 1.0000 0.0584
-11.000 -0.5097 0.10653 0.09940 -0.0342 1.0000 0.0581
-10.750 -0.5108 0.10153 0.09444 -0.0368 1.0000 0.0577
-10.500 -0.5136 0.09620 0.08916 -0.0399 1.0000 0.0574
-10.250 -0.5177 0.09061 0.08362 -0.0434 1.0000 0.0569
-10.000 -0.5260 0.08495 0.07799 -0.0472 1.0000 0.0563
-9.750 -0.5391 0.07954 0.07260 -0.0507 1.0000 0.0557
-9.500 -0.5545 0.07495 0.06800 -0.0528 1.0000 0.0550
-9.250 -0.5720 0.07119 0.06421 -0.0531 1.0000 0.0545
-9.000 -0.5879 0.06774 0.06070 -0.0524 1.0000 0.0542
-8.750 -0.6012 0.06450 0.05735 -0.0510 1.0000 0.0541
-8.500 -0.6138 0.06166 0.05438 -0.0486 1.0000 0.0539
-8.250 -0.6268 0.05924 0.05182 -0.0451 1.0000 0.0537
-8.000 -0.6391 0.05697 0.04937 -0.0413 1.0000 0.0535
-7.750 -0.6488 0.05469 0.04687 -0.0375 1.0000 0.0532
-7.500 -0.6552 0.05231 0.04423 -0.0339 1.0000 0.0532
-7.250 -0.6576 0.04995 0.04158 -0.0307 1.0000 0.0530
-7.000 -0.6504 0.04715 0.03839 -0.0291 0.9974 0.0532
-6.750 -0.6275 0.04390 0.03461 -0.0300 0.9902 0.0536
-6.500 -0.6016 0.04115 0.03132 -0.0310 0.9838 0.0557
-6.250 -0.5748 0.03881 0.02838 -0.0316 0.9773 0.0582
-6.000 -0.5452 0.03669 0.02569 -0.0323 0.9719 0.0594
-5.750 -0.5150 0.03442 0.02329 -0.0333 0.9668 0.0610
-5.500 -0.4806 0.03270 0.02143 -0.0347 0.9628 0.0636
-5.250 -0.4515 0.03149 0.02006 -0.0351 0.9572 0.0683
-5.000 -0.4178 0.03040 0.01868 -0.0359 0.9527 0.0728
-4.750 -0.3844 0.02911 0.01739 -0.0371 0.9491 0.0772
-4.500 -0.3615 0.02830 0.01645 -0.0365 0.9423 0.0835
-4.250 -0.3336 0.02736 0.01545 -0.0372 0.9372 0.0942
-4.000 -0.3092 0.02646 0.01447 -0.0372 0.9316 0.1065
-3.750 -0.2884 0.02539 0.01358 -0.0368 0.9252 0.1369
-3.500 -0.2790 0.02269 0.01311 -0.0349 0.9202 0.4689
-3.250 -0.2700 0.02287 0.01375 -0.0297 0.9129 0.6551
-3.000 -0.2557 0.02365 0.01470 -0.0239 0.9076 0.7629
-2.750 -0.1996 0.02561 0.01657 -0.0221 0.9081 0.8450
-2.500 -0.1560 0.02573 0.01636 -0.0251 0.9050 0.8556
-2.250 -0.1296 0.02581 0.01620 -0.0254 0.8990 0.8671
-2.000 -0.1017 0.02584 0.01602 -0.0261 0.8935 0.8788
-1.500 -0.0532 0.02582 0.01567 -0.0264 0.8823 0.9032
-1.250 -0.0200 0.02579 0.01547 -0.0282 0.8778 0.9132
-1.000 0.0194 0.02576 0.01529 -0.0311 0.8745 0.9220
-0.750 0.0485 0.02581 0.01524 -0.0323 0.8687 0.9323
-0.500 0.0921 0.02584 0.01517 -0.0362 0.8647 0.9391
-0.250 0.1339 0.02586 0.01511 -0.0398 0.8611 0.9470
0.000 0.1786 0.02590 0.01508 -0.0439 0.8578 0.9541
0.250 0.2119 0.02605 0.01522 -0.0462 0.8519 0.9641
0.500 0.2512 0.02614 0.01531 -0.0495 0.8476 0.9734
0.750 0.2955 0.02618 0.01536 -0.0535 0.8443 0.9817
1.000 0.3302 0.02639 0.01561 -0.0563 0.8386 0.9932
1.250 0.3574 0.02660 0.01585 -0.0575 0.8328 1.0000
1.500 0.3780 0.02674 0.01603 -0.0570 0.8280 1.0000
1.750 0.3707 0.02725 0.01654 -0.0521 0.8182 1.0000
2.000 0.3918 0.02746 0.01678 -0.0515 0.8132 1.0000
2.250 0.3885 0.02803 0.01734 -0.0471 0.8039 1.0000
2.500 0.4100 0.02833 0.01768 -0.0464 0.7986 1.0000
2.750 0.4135 0.02891 0.01828 -0.0431 0.7900 1.0000
3.000 0.4346 0.02927 0.01873 -0.0423 0.7840 1.0000
3.250 0.4445 0.02983 0.01933 -0.0398 0.7757 1.0000
3.500 0.4665 0.03021 0.01981 -0.0391 0.7689 1.0000
3.750 0.4801 0.03074 0.02043 -0.0372 0.7601 1.0000
4.000 0.5058 0.03105 0.02092 -0.0369 0.7530 1.0000
4.250 0.5192 0.03162 0.02160 -0.0350 0.7430 1.0000
4.500 0.5516 0.03174 0.02193 -0.0354 0.7364 1.0000
4.750 0.5645 0.03229 0.02263 -0.0333 0.7246 1.0000
5.000 0.5864 0.03254 0.02312 -0.0322 0.7133 1.0000
5.250 0.6183 0.03226 0.02313 -0.0317 0.7019 1.0000
5.500 0.6733 0.02698 0.01811 -0.0272 0.6416 1.0000
5.750 0.6804 0.02550 0.01671 -0.0210 0.5752 1.0000
6.000 0.7049 0.02392 0.01458 -0.0163 0.4011 1.0000
6.250 0.6936 0.02600 0.01512 -0.0105 0.2138 1.0000
6.500 0.6849 0.02820 0.01662 -0.0064 0.1407 1.0000
6.750 0.6849 0.03003 0.01811 -0.0033 0.1103 1.0000
7.000 0.6908 0.03160 0.01957 -0.0010 0.0958 1.0000
7.250 0.6998 0.03309 0.02103 0.0010 0.0869 1.0000
7.500 0.7119 0.03447 0.02244 0.0027 0.0787 1.0000
7.750 0.7291 0.03575 0.02380 0.0041 0.0720 1.0000
8.000 0.7549 0.03695 0.02510 0.0050 0.0673 1.0000
8.250 0.7988 0.03835 0.02656 0.0041 0.0617 1.0000
8.500 0.8421 0.04011 0.02863 0.0026 0.0562 1.0000
8.750 0.8813 0.04240 0.03109 0.0014 0.0539 1.0000
9.000 0.9138 0.04518 0.03402 0.0007 0.0521 1.0000
9.250 0.9350 0.04796 0.03714 0.0013 0.0507 1.0000
9.500 0.9474 0.05053 0.04022 0.0028 0.0491 1.0000
9.750 0.9560 0.05330 0.04342 0.0046 0.0476 1.0000
10.000 0.9612 0.05627 0.04677 0.0064 0.0469 1.0000
10.250 0.9618 0.05943 0.05029 0.0084 0.0466 1.0000
10.500 0.9569 0.06255 0.05373 0.0108 0.0465 1.0000
10.750 0.9473 0.06575 0.05722 0.0131 0.0465 1.0000
11.000 0.9340 0.06921 0.06093 0.0150 0.0465 1.0000
11.250 0.9187 0.07291 0.06488 0.0162 0.0466 1.0000
11.500 0.9005 0.07707 0.06925 0.0167 0.0468 1.0000
11.750 0.8806 0.08170 0.07407 0.0163 0.0471 1.0000
12.000 0.8588 0.08701 0.07955 0.0148 0.0474 1.0000
12.250 0.8359 0.09312 0.08579 0.0123 0.0478 1.0000
12.500 0.8131 0.10008 0.09285 0.0086 0.0482 1.0000
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Polar data table (+)
Polar graphs
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