NACA 66(1)-212 (naca661212-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 66(1)-212 (naca661212-il) Reynolds number: 200,000 Max Cl/Cd: 47.64 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca661212-il-200000.txt Download as CSV file: xf-naca661212-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 66(1)-212
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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.3907 0.09039 0.08709 -0.0502 1.0000 0.0582
-10.500 -0.4024 0.08609 0.08284 -0.0511 1.0000 0.0597
-10.250 -0.4226 0.08243 0.07926 -0.0506 1.0000 0.0606
-10.000 -0.4454 0.07998 0.07688 -0.0481 1.0000 0.0603
-9.750 -0.4687 0.07677 0.07373 -0.0466 1.0000 0.0596
-9.500 -0.6619 0.08048 0.07692 -0.0456 0.9993 0.0507
-9.250 -0.6671 0.07642 0.07253 -0.0485 0.9920 0.0509
-9.000 -0.6584 0.06708 0.06338 -0.0509 0.9902 0.0527
-8.750 -0.6413 0.06379 0.06010 -0.0524 0.9869 0.0545
-8.500 -0.6284 0.06015 0.05632 -0.0546 0.9834 0.0570
-8.250 -0.6253 0.05694 0.05281 -0.0546 0.9763 0.0606
-8.000 -0.6230 0.05221 0.04752 -0.0555 0.9711 0.0654
-7.750 -0.6058 0.04964 0.04509 -0.0557 0.9668 0.0686
-7.250 -0.5772 0.04382 0.03870 -0.0559 0.9586 0.0810
-7.000 -0.5586 0.04186 0.03659 -0.0557 0.9551 0.0869
-6.750 -0.5524 0.03959 0.03406 -0.0535 0.9490 0.0950
-6.500 -0.5323 0.03768 0.03186 -0.0537 0.9459 0.1080
-6.250 -0.4928 0.03087 0.02349 -0.0506 0.9441 0.0513
-6.000 -0.4594 0.02705 0.01947 -0.0520 0.9432 0.0485
-5.750 -0.4500 0.02570 0.01793 -0.0485 0.9371 0.0478
-5.500 -0.4222 0.02412 0.01614 -0.0485 0.9347 0.0475
-5.250 -0.3907 0.02288 0.01473 -0.0492 0.9329 0.0487
-5.000 -0.3579 0.02227 0.01393 -0.0502 0.9313 0.0508
-4.750 -0.3247 0.02051 0.01218 -0.0514 0.9303 0.0524
-4.500 -0.2946 0.01944 0.01117 -0.0522 0.9290 0.0553
-4.250 -0.2939 0.01930 0.01102 -0.0473 0.9220 0.0569
-4.000 -0.2712 0.01889 0.01057 -0.0466 0.9190 0.0616
-3.750 -0.2466 0.01821 0.00989 -0.0465 0.9169 0.0664
-3.500 -0.2189 0.01770 0.00937 -0.0469 0.9153 0.0744
-3.250 -0.2116 0.01756 0.00925 -0.0433 0.9107 0.0859
-3.000 -0.2069 0.01686 0.00899 -0.0395 0.9054 0.1654
-2.750 -0.2151 0.01475 0.00927 -0.0330 0.9019 0.7112
-2.500 -0.1909 0.01543 0.01003 -0.0309 0.9004 0.7774
-2.000 -0.1830 0.01689 0.01151 -0.0197 0.8913 0.8291
-1.750 -0.1687 0.01783 0.01249 -0.0148 0.8886 0.8523
-1.500 -0.1467 0.01848 0.01310 -0.0121 0.8866 0.8714
-1.250 -0.1165 0.01906 0.01363 -0.0112 0.8854 0.8861
-1.000 -0.1126 0.01945 0.01401 -0.0062 0.8802 0.8970
-0.750 -0.0940 0.01968 0.01419 -0.0043 0.8763 0.9069
-0.500 -0.0623 0.01976 0.01421 -0.0053 0.8740 0.9115
-0.250 -0.0346 0.01970 0.01409 -0.0058 0.8717 0.9163
0.000 0.0047 0.01971 0.01405 -0.0083 0.8703 0.9185
0.250 0.0048 0.01990 0.01424 -0.0037 0.8634 0.9247
0.500 0.0269 0.01986 0.01417 -0.0032 0.8595 0.9286
0.750 0.0586 0.01977 0.01406 -0.0043 0.8570 0.9315
1.000 0.1020 0.01975 0.01403 -0.0077 0.8555 0.9323
1.250 0.1057 0.01995 0.01425 -0.0039 0.8476 0.9375
1.500 0.1379 0.01994 0.01425 -0.0052 0.8442 0.9401
1.750 0.1780 0.01988 0.01421 -0.0078 0.8420 0.9418
2.000 0.2228 0.01979 0.01416 -0.0113 0.8406 0.9427
2.250 0.2252 0.02000 0.01440 -0.0073 0.8309 0.9487
2.500 0.2700 0.01987 0.01433 -0.0107 0.8283 0.9493
2.750 0.3222 0.01957 0.01412 -0.0152 0.8265 0.9489
3.000 0.3512 0.01935 0.01398 -0.0155 0.8170 0.9518
3.250 0.4393 0.01700 0.01172 -0.0235 0.8110 0.9463
3.500 0.4880 0.01524 0.01001 -0.0250 0.7926 0.9461
3.750 0.5329 0.01355 0.00831 -0.0259 0.7657 0.9464
4.000 0.5639 0.01282 0.00763 -0.0255 0.7368 0.9486
4.250 0.5841 0.01226 0.00705 -0.0232 0.6716 0.9526
4.500 0.5715 0.01312 0.00652 -0.0148 0.4091 0.9602
4.750 0.5550 0.01559 0.00741 -0.0084 0.1177 0.9691
5.000 0.5747 0.01673 0.00824 -0.0078 0.0727 0.9737
5.250 0.6036 0.01753 0.00909 -0.0088 0.0633 0.9771
5.500 0.6248 0.01859 0.01012 -0.0086 0.0570 0.9821
5.750 0.6545 0.01933 0.01092 -0.0096 0.0521 0.9854
6.000 0.6849 0.02036 0.01196 -0.0109 0.0488 0.9883
6.250 0.7168 0.02193 0.01347 -0.0122 0.0464 0.9904
6.500 0.7517 0.02333 0.01496 -0.0138 0.0439 0.9922
6.750 0.7869 0.02467 0.01643 -0.0154 0.0421 0.9943
7.000 0.8245 0.02651 0.01846 -0.0172 0.0415 0.9957
7.250 0.8608 0.02887 0.02110 -0.0187 0.0416 0.9976
7.500 0.8941 0.03175 0.02432 -0.0199 0.0425 1.0000
7.750 0.9015 0.03355 0.02636 -0.0164 0.0426 1.0000
8.000 0.9036 0.03527 0.02830 -0.0122 0.0427 1.0000
8.250 0.9203 0.03875 0.03188 -0.0105 0.0464 1.0000
10.000 0.8596 0.06053 0.05637 0.0153 0.0697 1.0000
10.250 0.8317 0.06189 0.05799 0.0204 0.0691 1.0000
10.500 0.8002 0.06447 0.06076 0.0242 0.0688 1.0000
10.750 0.7708 0.06825 0.06469 0.0261 0.0686 1.0000
11.000 0.7433 0.07302 0.06959 0.0266 0.0686 1.0000
11.250 0.7169 0.07854 0.07521 0.0260 0.0688 1.0000
11.500 0.6909 0.08473 0.08148 0.0242 0.0689 1.0000
11.750 0.6641 0.09170 0.08851 0.0212 0.0690 1.0000
12.000 0.6354 0.09990 0.09677 0.0162 0.0691 1.0000
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