NACA M14 AIRFOIL (m14-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M14 AIRFOIL (m14-il) Reynolds number: 1,000,000 Max Cl/Cd: 121.88 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m14-il-1000000.txt Download as CSV file: xf-m14-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4335 0.09011 0.08860 -0.0156 1.0000 0.0149
-8.000 -0.4318 0.08605 0.08455 -0.0187 1.0000 0.0149
-7.750 -0.4322 0.08031 0.07883 -0.0228 1.0000 0.0153
-7.500 -0.4209 0.07779 0.07631 -0.0242 1.0000 0.0155
-7.250 -0.4089 0.07500 0.07351 -0.0265 1.0000 0.0157
-7.000 -0.3931 0.07173 0.07023 -0.0300 0.9950 0.0162
-6.750 -0.3678 0.06763 0.06608 -0.0362 0.9710 0.0174
-6.500 -0.3393 0.06088 0.05915 -0.0450 0.9362 0.0191
-6.250 -0.3230 0.05623 0.05432 -0.0476 0.9078 0.0192
-6.000 -0.3048 0.05129 0.04919 -0.0500 0.8845 0.0192
-5.750 -0.2937 0.04414 0.04180 -0.0528 0.8647 0.0199
-5.500 -0.2736 0.04278 0.04033 -0.0531 0.8456 0.0203
-5.250 -0.2521 0.04132 0.03874 -0.0535 0.8278 0.0210
-5.000 -0.2287 0.03901 0.03626 -0.0543 0.8117 0.0225
-4.750 -0.1964 0.03519 0.03215 -0.0550 0.7976 0.0247
-4.500 -0.1728 0.03060 0.02722 -0.0553 0.7848 0.0248
-4.250 -0.1569 0.02303 0.01909 -0.0556 0.7738 0.0260
-4.000 -0.1319 0.02212 0.01807 -0.0556 0.7611 0.0265
-3.500 -0.0803 0.01985 0.01546 -0.0555 0.7378 0.0283
-3.250 -0.0534 0.01864 0.01405 -0.0553 0.7273 0.0302
-3.000 -0.0239 0.01940 0.01467 -0.0551 0.7165 0.0322
-2.750 0.0011 0.01605 0.01081 -0.0547 0.7069 0.0329
-2.500 -0.5844 0.45761 0.45506 0.1219 0.7866 0.0262
-1.750 0.1109 0.01153 0.00563 -0.0542 0.6682 0.0310
-1.500 0.1387 0.01079 0.00475 -0.0541 0.6592 0.0308
-1.250 0.1666 0.01015 0.00402 -0.0540 0.6501 0.0306
-1.000 0.1944 0.00963 0.00342 -0.0540 0.6414 0.0306
-0.750 0.2223 0.00929 0.00302 -0.0539 0.6323 0.0310
-0.500 0.2503 0.00902 0.00271 -0.0539 0.6234 0.0315
-0.250 0.2782 0.00883 0.00247 -0.0539 0.6149 0.0318
0.000 0.3063 0.00865 0.00226 -0.0539 0.6059 0.0321
0.250 0.3343 0.00849 0.00207 -0.0539 0.5974 0.0323
0.500 0.3623 0.00833 0.00188 -0.0540 0.5888 0.0324
0.750 0.3903 0.00811 0.00162 -0.0540 0.5807 0.0330
1.250 0.4467 0.00790 0.00133 -0.0541 0.5637 0.0351
1.500 0.4749 0.00786 0.00126 -0.0542 0.5558 0.0376
1.750 0.5032 0.00783 0.00122 -0.0543 0.5476 0.0404
2.000 0.5314 0.00782 0.00122 -0.0543 0.5396 0.0490
2.250 0.5590 0.00753 0.00130 -0.0545 0.5311 0.2191
2.500 0.5979 0.00576 0.00147 -0.0575 0.5230 1.0000
2.750 0.6244 0.00585 0.00150 -0.0572 0.5148 1.0000
3.000 0.6512 0.00592 0.00155 -0.0570 0.5066 1.0000
3.250 0.6779 0.00602 0.00161 -0.0568 0.4981 1.0000
3.500 0.7049 0.00610 0.00168 -0.0567 0.4896 1.0000
3.750 0.7318 0.00621 0.00175 -0.0565 0.4798 1.0000
4.000 0.7587 0.00633 0.00183 -0.0564 0.4674 1.0000
4.250 0.7855 0.00647 0.00192 -0.0563 0.4490 1.0000
4.500 0.8119 0.00668 0.00202 -0.0561 0.4255 1.0000
4.750 0.8385 0.00688 0.00214 -0.0560 0.4024 1.0000
5.000 0.8650 0.00711 0.00229 -0.0559 0.3815 1.0000
5.250 0.8910 0.00740 0.00247 -0.0557 0.3540 1.0000
5.500 0.9172 0.00767 0.00266 -0.0556 0.3294 1.0000
5.750 0.9426 0.00805 0.00289 -0.0554 0.2971 1.0000
6.000 0.9672 0.00856 0.00319 -0.0551 0.2507 1.0000
6.250 0.9862 0.00984 0.00390 -0.0542 0.1427 1.0000
6.500 1.0039 0.01129 0.00480 -0.0531 0.0431 1.0000
6.750 1.0271 0.01194 0.00532 -0.0525 0.0224 1.0000
7.000 1.0515 0.01240 0.00579 -0.0520 0.0183 1.0000
7.250 1.0762 0.01278 0.00622 -0.0517 0.0168 1.0000
7.500 1.1001 0.01327 0.00672 -0.0512 0.0153 1.0000
7.750 1.1218 0.01401 0.00755 -0.0504 0.0138 1.0000
8.000 1.1457 0.01443 0.00801 -0.0499 0.0133 1.0000
8.250 1.1690 0.01490 0.00852 -0.0494 0.0126 1.0000
8.500 1.1917 0.01541 0.00907 -0.0488 0.0118 1.0000
8.750 1.2132 0.01602 0.00972 -0.0480 0.0112 1.0000
9.000 1.2307 0.01701 0.01079 -0.0467 0.0105 1.0000
9.250 1.2446 0.01829 0.01217 -0.0449 0.0100 1.0000
9.500 1.2646 0.01890 0.01285 -0.0439 0.0098 1.0000
9.750 1.2827 0.01965 0.01367 -0.0427 0.0095 1.0000
10.000 1.2991 0.02050 0.01459 -0.0413 0.0092 1.0000
10.250 1.3141 0.02141 0.01557 -0.0397 0.0089 1.0000
10.500 1.3280 0.02231 0.01654 -0.0380 0.0086 1.0000
10.750 1.3400 0.02313 0.01741 -0.0360 0.0083 1.0000
11.000 1.3504 0.02400 0.01834 -0.0339 0.0080 1.0000
11.250 1.3593 0.02507 0.01945 -0.0318 0.0078 1.0000
11.500 1.3621 0.02674 0.02121 -0.0294 0.0075 1.0000
11.750 1.3536 0.02981 0.02442 -0.0261 0.0072 1.0000
12.000 1.3608 0.03142 0.02615 -0.0247 0.0071 1.0000
12.250 1.3695 0.03291 0.02774 -0.0237 0.0070 1.0000
12.500 1.3764 0.03466 0.02960 -0.0226 0.0069 1.0000
12.750 1.3817 0.03665 0.03170 -0.0216 0.0068 1.0000
13.000 1.3856 0.03884 0.03402 -0.0207 0.0067 1.0000
13.250 1.3884 0.04120 0.03652 -0.0199 0.0066 1.0000
13.500 1.3896 0.04379 0.03924 -0.0192 0.0065 1.0000
13.750 1.3893 0.04661 0.04220 -0.0187 0.0064 1.0000
14.000 1.3876 0.04966 0.04540 -0.0183 0.0063 1.0000
14.250 1.3844 0.05294 0.04882 -0.0182 0.0062 1.0000
14.500 1.3796 0.05651 0.05254 -0.0183 0.0061 1.0000
14.750 1.3729 0.06039 0.05657 -0.0186 0.0061 1.0000
15.000 1.3644 0.06464 0.06098 -0.0193 0.0060 1.0000
15.250 1.3550 0.06919 0.06567 -0.0203 0.0060 1.0000
15.500 1.3444 0.07408 0.07071 -0.0216 0.0059 1.0000
15.750 1.3308 0.07971 0.07651 -0.0234 0.0059 1.0000
16.000 1.3163 0.08572 0.08268 -0.0257 0.0059 1.0000
16.250 1.3017 0.09197 0.08908 -0.0283 0.0058 1.0000
16.500 1.2833 0.09923 0.09651 -0.0315 0.0058 1.0000
16.750 1.2622 0.10735 0.10480 -0.0353 0.0058 1.0000
17.000 1.2410 0.11583 0.11345 -0.0397 0.0058 1.0000
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Polar data table (+)
Polar graphs
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