NACA 65(1)-212 a=0.6 (naca651212a06-il) Xfoil prediction polar at RE=200,000 Ncrit=9
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Airfoil: NACA 65(1)-212 a=0.6 (naca651212a06-il) Reynolds number: 200,000 Max Cl/Cd: 66.71 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca651212a06-il-200000.txt Download as CSV file: xf-naca651212a06-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5139 0.08465 0.08128 -0.0444 1.0000 0.0579
-9.750 -0.5304 0.07783 0.07446 -0.0506 1.0000 0.0579
-9.500 -0.5444 0.07287 0.06949 -0.0537 1.0000 0.0576
-9.250 -0.5761 0.06959 0.06612 -0.0546 1.0000 0.0584
-9.000 -0.6010 0.06766 0.06408 -0.0522 1.0000 0.0587
-8.750 -0.6233 0.06678 0.06300 -0.0493 0.9946 0.0592
-8.500 -0.6171 0.05839 0.05451 -0.0531 0.9874 0.0612
-8.250 -0.5879 0.05519 0.05143 -0.0558 0.9843 0.0645
-8.000 -0.5695 0.05164 0.04755 -0.0587 0.9767 0.0700
-7.750 -0.5556 0.04664 0.04211 -0.0617 0.9701 0.0756
-7.500 -0.5322 0.04383 0.03928 -0.0630 0.9630 0.0794
-7.250 -0.5159 0.04082 0.03578 -0.0639 0.9555 0.0896
-7.000 -0.4942 0.03840 0.03335 -0.0642 0.9477 0.0942
-6.750 -0.4725 0.03291 0.02651 -0.0605 0.9394 0.0588
-6.500 -0.4526 0.02839 0.02126 -0.0584 0.9320 0.0477
-6.250 -0.4280 0.02584 0.01850 -0.0579 0.9262 0.0463
-6.000 -0.4050 0.02392 0.01633 -0.0569 0.9191 0.0461
-5.750 -0.3801 0.02271 0.01489 -0.0561 0.9140 0.0468
-5.500 -0.3562 0.02084 0.01289 -0.0555 0.9074 0.0487
-5.250 -0.3315 0.01954 0.01152 -0.0548 0.9020 0.0498
-5.000 -0.3080 0.01858 0.01054 -0.0540 0.8968 0.0515
-4.750 -0.2852 0.01782 0.00977 -0.0531 0.8904 0.0536
-4.500 -0.2625 0.01722 0.00912 -0.0520 0.8856 0.0570
-4.250 -0.2405 0.01663 0.00848 -0.0511 0.8801 0.0604
-4.000 -0.2207 0.01587 0.00774 -0.0498 0.8742 0.0654
-3.750 -0.1984 0.01541 0.00720 -0.0487 0.8699 0.0722
-3.500 -0.1772 0.01482 0.00668 -0.0477 0.8641 0.0905
-3.250 -0.1761 0.01230 0.00607 -0.0441 0.8580 0.4927
-3.000 -0.1587 0.01209 0.00631 -0.0413 0.8544 0.6461
-2.750 -0.1361 0.01225 0.00652 -0.0400 0.8486 0.6901
-2.500 -0.1125 0.01240 0.00666 -0.0388 0.8438 0.7190
-2.250 -0.0907 0.01268 0.00697 -0.0366 0.8404 0.7541
-2.000 -0.0706 0.01279 0.00711 -0.0343 0.8360 0.7878
-1.750 -0.0485 0.01265 0.00703 -0.0328 0.8308 0.8106
-1.500 -0.0229 0.01255 0.00694 -0.0320 0.8270 0.8234
-1.250 0.0043 0.01252 0.00685 -0.0317 0.8241 0.8317
-1.000 0.0307 0.01270 0.00705 -0.0314 0.8191 0.8379
-0.750 0.0570 0.01282 0.00716 -0.0311 0.8145 0.8454
-0.500 0.0843 0.01296 0.00728 -0.0306 0.8111 0.8510
-0.250 0.1114 0.01309 0.00737 -0.0302 0.8085 0.8586
0.000 0.1363 0.01340 0.00770 -0.0298 0.8027 0.8664
0.250 0.1619 0.01355 0.00783 -0.0295 0.7984 0.8776
0.500 0.1860 0.01363 0.00786 -0.0289 0.7951 0.8923
0.750 0.2094 0.01373 0.00796 -0.0281 0.7911 0.9052
1.000 0.2329 0.01383 0.00809 -0.0275 0.7857 0.9170
1.250 0.2598 0.01380 0.00808 -0.0273 0.7820 0.9282
1.500 0.2942 0.01374 0.00804 -0.0285 0.7793 0.9358
1.750 0.3283 0.01386 0.00820 -0.0300 0.7752 0.9449
2.000 0.3656 0.01399 0.00840 -0.0323 0.7704 0.9532
2.250 0.4074 0.01401 0.00846 -0.0354 0.7670 0.9590
2.500 0.4512 0.01397 0.00846 -0.0386 0.7641 0.9644
2.750 0.4922 0.01407 0.00866 -0.0417 0.7581 0.9713
3.000 0.5364 0.01398 0.00865 -0.0452 0.7525 0.9766
3.250 0.5767 0.01363 0.00833 -0.0473 0.7442 0.9838
3.500 0.6185 0.01256 0.00724 -0.0487 0.7231 0.9901
3.750 0.6590 0.01152 0.00612 -0.0500 0.6932 0.9980
4.000 0.6834 0.01116 0.00575 -0.0492 0.6739 1.0000
4.250 0.7012 0.01098 0.00562 -0.0472 0.6511 1.0000
4.500 0.7192 0.01080 0.00540 -0.0450 0.6147 1.0000
4.750 0.7311 0.01096 0.00517 -0.0415 0.5070 1.0000
5.000 0.7135 0.01337 0.00602 -0.0343 0.2331 1.0000
5.250 0.7029 0.01551 0.00718 -0.0286 0.0856 1.0000
5.500 0.7114 0.01638 0.00798 -0.0253 0.0677 1.0000
5.750 0.7194 0.01726 0.00884 -0.0221 0.0594 1.0000
6.000 0.7310 0.01799 0.00963 -0.0195 0.0552 1.0000
6.250 0.7432 0.01879 0.01045 -0.0170 0.0518 1.0000
6.500 0.7544 0.01984 0.01145 -0.0145 0.0485 1.0000
6.750 0.7692 0.02106 0.01267 -0.0126 0.0460 1.0000
7.000 0.7891 0.02195 0.01363 -0.0114 0.0439 1.0000
7.250 0.8124 0.02312 0.01483 -0.0107 0.0422 1.0000
7.500 0.8388 0.02445 0.01621 -0.0104 0.0410 1.0000
7.750 0.8673 0.02602 0.01785 -0.0105 0.0400 1.0000
8.000 0.8943 0.02766 0.01957 -0.0106 0.0386 1.0000
8.250 0.9251 0.03074 0.02273 -0.0114 0.0371 1.0000
8.500 0.9496 0.03309 0.02530 -0.0110 0.0371 1.0000
8.750 0.9709 0.03523 0.02771 -0.0100 0.0373 1.0000
12.500 0.6399 0.11431 0.11112 -0.0024 0.0744 1.0000
12.750 0.7459 0.12304 0.11955 -0.0097 0.0641 1.0000
13.000 0.7511 0.12628 0.12280 -0.0104 0.0619 1.0000
13.250 0.7705 0.12586 0.12242 -0.0070 0.0600 1.0000
13.500 0.5931 0.13556 0.13234 -0.0150 0.0666 1.0000
13.750 0.5905 0.13889 0.13568 -0.0169 0.0639 1.0000
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Polar data table (+)
Polar graphs
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