NACA 66-021 AIRFOIL (n66021-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 66-021 AIRFOIL (n66021-il) Reynolds number: 1,000,000 Max Cl/Cd: 64.89 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n66021-il-1000000-n5.txt Download as CSV file: xf-n66021-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 66-021 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
------ -------- --------- --------- -------- -------- --------
-19.500 -1.0637 0.06950 0.06423 -0.0757 0.8242 0.0120
-19.250 -1.0617 0.06681 0.06145 -0.0763 0.8201 0.0121
-19.000 -1.0581 0.06440 0.05894 -0.0768 0.8160 0.0122
-18.750 -1.0564 0.06183 0.05628 -0.0771 0.8121 0.0122
-18.500 -1.0531 0.05944 0.05381 -0.0774 0.8090 0.0123
-18.250 -1.0468 0.05743 0.05173 -0.0775 0.8053 0.0123
-18.000 -1.0408 0.05545 0.04966 -0.0776 0.8015 0.0124
-17.750 -1.0362 0.05338 0.04751 -0.0775 0.7975 0.0124
-17.250 -1.0226 0.04977 0.04375 -0.0773 0.7911 0.0126
-17.000 -1.0141 0.04817 0.04208 -0.0771 0.7881 0.0126
-16.750 -1.0050 0.04665 0.04049 -0.0768 0.7851 0.0127
-16.500 -0.9998 0.04485 0.03863 -0.0763 0.7820 0.0130
-16.250 -0.9894 0.04353 0.03725 -0.0760 0.7789 0.0130
-16.000 -0.9827 0.04194 0.03559 -0.0754 0.7760 0.0131
-15.750 -0.9718 0.04069 0.03428 -0.0749 0.7735 0.0132
-15.500 -0.9619 0.03938 0.03294 -0.0744 0.7714 0.0134
-15.250 -0.9507 0.03820 0.03173 -0.0739 0.7690 0.0137
-15.000 -0.9379 0.03716 0.03064 -0.0734 0.7664 0.0139
-14.750 -0.9257 0.03610 0.02955 -0.0728 0.7639 0.0141
-14.500 -0.9135 0.03506 0.02844 -0.0722 0.7615 0.0140
-14.250 -0.9004 0.03411 0.02744 -0.0716 0.7591 0.0142
-14.000 -0.8865 0.03322 0.02651 -0.0710 0.7566 0.0147
-13.750 -0.8728 0.03232 0.02557 -0.0704 0.7548 0.0149
-13.500 -0.8586 0.03145 0.02467 -0.0698 0.7532 0.0151
-13.250 -0.8439 0.03064 0.02382 -0.0692 0.7514 0.0152
-13.000 -0.8287 0.02987 0.02302 -0.0686 0.7495 0.0156
-12.750 -0.8143 0.02907 0.02220 -0.0679 0.7475 0.0159
-12.500 -0.7998 0.02828 0.02139 -0.0672 0.7456 0.0164
-12.250 -0.7841 0.02759 0.02066 -0.0665 0.7436 0.0168
-12.000 -0.7677 0.02695 0.02000 -0.0659 0.7418 0.0173
-11.750 -0.7515 0.02632 0.01934 -0.0652 0.7399 0.0180
-11.250 -0.7192 0.02504 0.01801 -0.0638 0.7369 0.0192
-11.000 -0.7023 0.02444 0.01741 -0.0632 0.7357 0.0199
-10.750 -0.6850 0.02387 0.01683 -0.0625 0.7345 0.0207
-10.500 -0.6672 0.02334 0.01629 -0.0619 0.7331 0.0215
-10.250 -0.6499 0.02280 0.01573 -0.0612 0.7317 0.0224
-10.000 -0.6334 0.02222 0.01516 -0.0604 0.7304 0.0236
-9.750 -0.6156 0.02172 0.01464 -0.0597 0.7290 0.0246
-9.500 -0.5972 0.02126 0.01416 -0.0591 0.7276 0.0258
-9.250 -0.5799 0.02075 0.01365 -0.0583 0.7262 0.0272
-9.000 -0.5622 0.02027 0.01317 -0.0575 0.7248 0.0289
-8.750 -0.5447 0.01981 0.01269 -0.0567 0.7234 0.0307
-8.500 -0.5270 0.01935 0.01223 -0.0558 0.7219 0.0327
-8.250 -0.5095 0.01890 0.01178 -0.0549 0.7205 0.0365
-8.000 -0.4915 0.01846 0.01136 -0.0541 0.7195 0.0402
-7.750 -0.4759 0.01792 0.01088 -0.0530 0.7187 0.0493
-7.500 -0.4617 0.01735 0.01037 -0.0516 0.7179 0.0631
-7.250 -0.4482 0.01676 0.00988 -0.0501 0.7170 0.0813
-7.000 -0.4371 0.01610 0.00934 -0.0482 0.7161 0.1057
-6.750 -0.4268 0.01542 0.00879 -0.0462 0.7151 0.1318
-6.500 -0.4217 0.01458 0.00816 -0.0434 0.7140 0.1734
-6.250 -0.4096 0.01401 0.00767 -0.0415 0.7130 0.1942
-6.000 -0.4122 0.01297 0.00689 -0.0374 0.7119 0.2481
-5.750 -0.4064 0.01225 0.00630 -0.0344 0.7108 0.2795
-5.500 -0.4055 0.01142 0.00563 -0.0306 0.7097 0.3180
-5.250 -0.4094 0.01051 0.00490 -0.0258 0.7086 0.3618
-5.000 -0.4217 0.00949 0.00409 -0.0194 0.7073 0.4123
-4.750 -0.4375 0.00858 0.00337 -0.0119 0.7060 0.4589
-4.500 -0.4565 0.00801 0.00297 -0.0032 0.7047 0.5030
-4.250 -0.4645 0.00737 0.00259 0.0035 0.7034 0.5629
-4.000 -0.4566 0.00704 0.00252 0.0071 0.7024 0.6297
-3.750 -0.4312 0.00703 0.00253 0.0074 0.7019 0.6469
-3.500 -0.4043 0.00704 0.00254 0.0073 0.7015 0.6565
-3.250 -0.3767 0.00706 0.00255 0.0070 0.7010 0.6637
-3.000 -0.3489 0.00708 0.00253 0.0067 0.7005 0.6687
-2.750 -0.3203 0.00708 0.00255 0.0063 0.6999 0.6728
-2.500 -0.2916 0.00710 0.00257 0.0059 0.6993 0.6764
-2.250 -0.2629 0.00712 0.00258 0.0054 0.6986 0.6797
-2.000 -0.2341 0.00714 0.00257 0.0049 0.6979 0.6821
-1.750 -0.2054 0.00716 0.00256 0.0044 0.6973 0.6841
-1.500 -0.1764 0.00716 0.00255 0.0039 0.6967 0.6867
-1.250 -0.1471 0.00716 0.00257 0.0032 0.6961 0.6892
-1.000 -0.1177 0.00716 0.00258 0.0026 0.6955 0.6907
-0.750 -0.0883 0.00716 0.00257 0.0020 0.6949 0.6914
-0.500 -0.0589 0.00715 0.00256 0.0013 0.6943 0.6919
-0.250 -0.0294 0.00715 0.00255 0.0007 0.6937 0.6925
0.000 0.0000 0.00714 0.00255 0.0000 0.6931 0.6931
0.250 0.0294 0.00715 0.00255 -0.0007 0.6925 0.6937
0.500 0.0589 0.00715 0.00256 -0.0013 0.6919 0.6943
0.750 0.0883 0.00716 0.00257 -0.0020 0.6914 0.6949
1.000 0.1177 0.00716 0.00258 -0.0026 0.6906 0.6955
1.250 0.1471 0.00716 0.00257 -0.0032 0.6892 0.6961
1.500 0.1764 0.00716 0.00255 -0.0039 0.6867 0.6966
1.750 0.2054 0.00716 0.00256 -0.0044 0.6842 0.6972
2.000 0.2341 0.00714 0.00257 -0.0049 0.6821 0.6978
2.250 0.2629 0.00712 0.00257 -0.0054 0.6797 0.6986
2.500 0.2916 0.00710 0.00257 -0.0059 0.6764 0.6993
2.750 0.3203 0.00708 0.00255 -0.0063 0.6728 0.6999
3.000 0.3489 0.00708 0.00253 -0.0067 0.6689 0.7005
3.250 0.3766 0.00705 0.00254 -0.0070 0.6634 0.7010
3.500 0.4041 0.00703 0.00253 -0.0072 0.6557 0.7015
3.750 0.4313 0.00703 0.00253 -0.0074 0.6473 0.7019
4.000 0.4568 0.00704 0.00253 -0.0072 0.6312 0.7024
4.250 0.4659 0.00734 0.00258 -0.0037 0.5659 0.7034
4.500 0.4547 0.00804 0.00299 0.0035 0.5001 0.7047
4.750 0.4350 0.00862 0.00340 0.0124 0.4561 0.7061
5.000 0.4216 0.00949 0.00409 0.0194 0.4120 0.7073
5.250 0.4092 0.01052 0.00490 0.0259 0.3613 0.7085
5.500 0.4068 0.01138 0.00560 0.0304 0.3203 0.7097
5.750 0.4080 0.01221 0.00627 0.0342 0.2825 0.7108
6.000 0.4091 0.01308 0.00696 0.0379 0.2412 0.7119
6.250 0.4164 0.01378 0.00752 0.0405 0.2091 0.7129
6.500 0.4219 0.01458 0.00815 0.0433 0.1736 0.7140
6.750 0.4257 0.01546 0.00882 0.0463 0.1287 0.7151
7.000 0.4337 0.01623 0.00942 0.0487 0.0965 0.7161
7.250 0.4473 0.01680 0.00990 0.0502 0.0785 0.7170
7.500 0.4624 0.01732 0.01036 0.0515 0.0642 0.7179
7.750 0.4765 0.01790 0.01086 0.0529 0.0499 0.7188
8.000 0.4925 0.01842 0.01133 0.0540 0.0413 0.7195
8.250 0.5092 0.01892 0.01179 0.0550 0.0352 0.7205
8.500 0.5274 0.01934 0.01222 0.0558 0.0328 0.7219
8.750 0.5448 0.01980 0.01268 0.0566 0.0304 0.7233
9.000 0.5625 0.02026 0.01316 0.0574 0.0286 0.7249
9.250 0.5801 0.02075 0.01365 0.0582 0.0270 0.7262
9.500 0.5977 0.02124 0.01414 0.0590 0.0255 0.7276
9.750 0.6161 0.02170 0.01462 0.0597 0.0246 0.7290
10.000 0.6338 0.02221 0.01514 0.0604 0.0233 0.7303
10.250 0.6506 0.02277 0.01570 0.0611 0.0223 0.7317
10.500 0.6677 0.02332 0.01627 0.0618 0.0215 0.7331
10.750 0.6854 0.02386 0.01682 0.0625 0.0208 0.7345
11.000 0.7028 0.02442 0.01738 0.0631 0.0196 0.7357
11.250 0.7197 0.02501 0.01798 0.0637 0.0190 0.7369
11.750 0.7523 0.02628 0.01929 0.0651 0.0178 0.7399
12.000 0.7686 0.02691 0.01996 0.0658 0.0172 0.7418
12.250 0.7846 0.02757 0.02064 0.0665 0.0165 0.7437
12.500 0.8003 0.02826 0.02136 0.0671 0.0164 0.7455
12.750 0.8154 0.02901 0.02214 0.0678 0.0162 0.7475
13.000 0.8300 0.02980 0.02295 0.0685 0.0157 0.7495
13.250 0.8441 0.03064 0.02382 0.0692 0.0153 0.7513
13.500 0.8594 0.03142 0.02463 0.0697 0.0150 0.7532
13.750 0.8738 0.03226 0.02551 0.0703 0.0149 0.7548
14.000 0.8877 0.03315 0.02644 0.0709 0.0147 0.7566
14.250 0.9012 0.03406 0.02740 0.0715 0.0145 0.7591
14.500 0.9145 0.03500 0.02838 0.0721 0.0141 0.7615
14.750 0.9265 0.03605 0.02949 0.0728 0.0139 0.7640
15.000 0.9390 0.03709 0.03057 0.0733 0.0138 0.7665
15.250 0.9510 0.03819 0.03170 0.0738 0.0135 0.7690
15.500 0.9627 0.03933 0.03290 0.0743 0.0134 0.7714
15.750 0.9740 0.04053 0.03414 0.0748 0.0134 0.7736
16.000 0.9832 0.04190 0.03556 0.0753 0.0132 0.7761
16.250 0.9936 0.04320 0.03691 0.0757 0.0131 0.7791
16.500 0.9976 0.04504 0.03882 0.0764 0.0127 0.7822
16.750 1.0092 0.04630 0.04014 0.0766 0.0126 0.7853
17.000 1.0143 0.04816 0.04207 0.0770 0.0126 0.7882
17.250 1.0225 0.04979 0.04377 0.0772 0.0126 0.7912
17.750 1.0362 0.05339 0.04751 0.0775 0.0125 0.7977
18.000 1.0429 0.05527 0.04948 0.0776 0.0124 0.8018
18.250 1.0489 0.05724 0.05153 0.0775 0.0123 0.8053
18.500 1.0515 0.05961 0.05398 0.0773 0.0123 0.8091
19.000 1.0619 0.06400 0.05854 0.0767 0.0121 0.8165
19.250 1.0638 0.06660 0.06124 0.0763 0.0121 0.8206
|
Polar data table (+)
Polar graphs
<< Back to NACA 66-021 AIRFOIL (n66021-il)