NACA M14 AIRFOIL (m14-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M14 AIRFOIL (m14-il) Reynolds number: 500,000 Max Cl/Cd: 102.93 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m14-il-500000.txt Download as CSV file: xf-m14-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M14 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4524 0.11936 0.11710 -0.0022 1.0000 0.0190
-9.750 -0.4505 0.11581 0.11356 -0.0052 1.0000 0.0191
-9.500 -0.4472 0.11203 0.10980 -0.0077 1.0000 0.0192
-9.250 -0.4431 0.10816 0.10595 -0.0101 1.0000 0.0192
-9.000 -0.4387 0.10430 0.10210 -0.0124 1.0000 0.0192
-8.750 -0.4385 0.09910 0.09693 -0.0135 1.0000 0.0195
-8.500 -0.4298 0.09597 0.09380 -0.0127 1.0000 0.0198
-8.250 -0.4227 0.09301 0.09085 -0.0134 1.0000 0.0201
-8.000 -0.4169 0.09000 0.08786 -0.0147 1.0000 0.0204
-7.750 -0.4121 0.08697 0.08485 -0.0164 1.0000 0.0208
-7.500 -0.4073 0.08380 0.08171 -0.0189 1.0000 0.0215
-7.250 -0.3985 0.08010 0.07802 -0.0227 1.0000 0.0225
-7.000 -0.3836 0.07531 0.07320 -0.0309 1.0000 0.0243
-6.750 -0.3672 0.07066 0.06850 -0.0369 1.0000 0.0247
-6.500 -0.3555 0.06670 0.06450 -0.0393 0.9998 0.0248
-6.250 -0.3240 0.06078 0.05845 -0.0464 0.9903 0.0248
-6.000 -0.3080 0.05405 0.05166 -0.0507 0.9779 0.0257
-5.750 -0.2831 0.05162 0.04917 -0.0530 0.9629 0.0265
-5.500 -0.2587 0.04873 0.04618 -0.0552 0.9432 0.0276
-4.750 -0.1864 0.03520 0.03168 -0.0590 0.8824 0.0329
-4.500 -0.1688 0.03290 0.02927 -0.0589 0.8647 0.0339
-4.250 -0.1473 0.03148 0.02772 -0.0587 0.8479 0.0349
-4.000 -0.1240 0.02972 0.02578 -0.0586 0.8325 0.0364
-3.750 -0.0988 0.02758 0.02337 -0.0585 0.8183 0.0391
-3.500 -0.0677 0.02726 0.02258 -0.0576 0.8049 0.0422
-3.000 -0.0230 0.02163 0.01660 -0.0579 0.7800 0.0469
-2.750 0.0045 0.02081 0.01553 -0.0576 0.7681 0.0519
-2.500 0.0302 0.01882 0.01323 -0.0575 0.7570 0.0568
-2.250 0.0565 0.01805 0.01238 -0.0574 0.7461 0.0596
-2.000 0.0862 0.01944 0.01354 -0.0569 0.7346 0.0667
-1.750 0.1108 0.01631 0.01028 -0.0572 0.7242 0.0716
-1.500 0.1423 0.01314 0.00634 -0.0557 0.7154 0.0433
-1.250 0.1702 0.01264 0.00575 -0.0555 0.7052 0.0416
-1.000 0.1979 0.01170 0.00470 -0.0553 0.6952 0.0410
-0.750 0.2254 0.01115 0.00406 -0.0551 0.6859 0.0410
-0.500 0.2529 0.01073 0.00357 -0.0549 0.6764 0.0412
-0.250 0.2806 0.01038 0.00319 -0.0548 0.6670 0.0415
0.000 0.3081 0.01011 0.00286 -0.0547 0.6582 0.0417
0.250 0.3358 0.00989 0.00261 -0.0546 0.6487 0.0424
0.500 0.3633 0.00960 0.00228 -0.0545 0.6399 0.0436
0.750 0.3910 0.00945 0.00207 -0.0545 0.6313 0.0446
1.000 0.4191 0.00932 0.00194 -0.0545 0.6220 0.0467
1.250 0.4470 0.00927 0.00185 -0.0544 0.6134 0.0504
1.500 0.4750 0.00920 0.00180 -0.0544 0.6045 0.0642
1.750 0.4992 0.00799 0.00189 -0.0545 0.5963 0.5724
2.250 0.5707 0.00712 0.00198 -0.0576 0.5779 1.0000
2.500 0.5972 0.00723 0.00202 -0.0573 0.5698 1.0000
2.750 0.6238 0.00732 0.00207 -0.0571 0.5607 1.0000
3.000 0.6506 0.00742 0.00214 -0.0569 0.5518 1.0000
3.250 0.6771 0.00755 0.00220 -0.0566 0.5434 1.0000
3.500 0.7041 0.00763 0.00230 -0.0565 0.5342 1.0000
3.750 0.7309 0.00776 0.00240 -0.0563 0.5254 1.0000
4.000 0.7577 0.00789 0.00250 -0.0562 0.5159 1.0000
4.250 0.7847 0.00801 0.00264 -0.0560 0.5066 1.0000
4.500 0.8114 0.00815 0.00276 -0.0559 0.4953 1.0000
4.750 0.8379 0.00828 0.00287 -0.0557 0.4781 1.0000
5.000 0.8639 0.00847 0.00297 -0.0554 0.4562 1.0000
5.250 0.8903 0.00865 0.00313 -0.0552 0.4363 1.0000
5.500 0.9160 0.00890 0.00329 -0.0550 0.4102 1.0000
5.750 0.9416 0.00918 0.00349 -0.0547 0.3844 1.0000
6.000 0.9672 0.00948 0.00373 -0.0545 0.3591 1.0000
6.250 0.9917 0.00991 0.00404 -0.0541 0.3247 1.0000
6.500 1.0137 0.01067 0.00448 -0.0535 0.2573 1.0000
6.750 1.0246 0.01292 0.00575 -0.0517 0.0877 1.0000
7.000 1.0416 0.01433 0.00683 -0.0504 0.0314 1.0000
7.250 1.0628 0.01516 0.00769 -0.0494 0.0254 1.0000
7.500 1.0859 0.01568 0.00827 -0.0488 0.0232 1.0000
7.750 1.1076 0.01633 0.00896 -0.0480 0.0212 1.0000
8.000 1.1258 0.01735 0.01005 -0.0467 0.0195 1.0000
8.250 1.1424 0.01847 0.01128 -0.0452 0.0186 1.0000
8.500 1.1610 0.01930 0.01221 -0.0439 0.0179 1.0000
8.750 1.1783 0.02021 0.01320 -0.0426 0.0171 1.0000
9.000 1.1952 0.02110 0.01415 -0.0412 0.0162 1.0000
9.250 1.2109 0.02205 0.01514 -0.0397 0.0154 1.0000
9.500 1.2225 0.02331 0.01645 -0.0377 0.0148 1.0000
9.750 1.2287 0.02512 0.01834 -0.0350 0.0142 1.0000
10.000 1.2347 0.02735 0.02066 -0.0322 0.0139 1.0000
10.250 1.2474 0.02839 0.02181 -0.0302 0.0137 1.0000
10.500 1.2598 0.02969 0.02323 -0.0284 0.0134 1.0000
10.750 1.2721 0.03120 0.02486 -0.0267 0.0131 1.0000
11.000 1.2845 0.03294 0.02675 -0.0250 0.0129 1.0000
11.250 1.2959 0.03483 0.02880 -0.0234 0.0126 1.0000
11.500 1.3050 0.03671 0.03084 -0.0217 0.0123 1.0000
11.750 1.3121 0.03858 0.03288 -0.0200 0.0119 1.0000
12.000 1.3177 0.04049 0.03492 -0.0184 0.0116 1.0000
12.250 1.3217 0.04270 0.03729 -0.0170 0.0114 1.0000
12.500 1.3223 0.04554 0.04034 -0.0154 0.0113 1.0000
12.750 1.3178 0.04903 0.04409 -0.0139 0.0114 1.0000
13.000 1.3061 0.05362 0.04901 -0.0124 0.0116 1.0000
13.250 1.2889 0.05884 0.05455 -0.0115 0.0119 1.0000
13.500 1.2694 0.06435 0.06034 -0.0114 0.0122 1.0000
13.750 1.2496 0.06999 0.06621 -0.0121 0.0125 1.0000
14.000 1.2296 0.07583 0.07226 -0.0135 0.0127 1.0000
14.250 1.2088 0.08208 0.07869 -0.0156 0.0129 1.0000
14.500 1.1874 0.08871 0.08548 -0.0183 0.0131 1.0000
14.750 1.1652 0.09591 0.09283 -0.0215 0.0133 1.0000
15.000 1.0112 0.09304 0.09005 -0.0079 0.0128 1.0000
15.250 0.9844 0.09952 0.09667 -0.0106 0.0129 1.0000
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Polar data table (+)
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