NACA M20 AIRFOIL (m20-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M20 AIRFOIL (m20-il) Reynolds number: 1,000,000 Max Cl/Cd: 125.27 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m20-il-1000000-n5.txt Download as CSV file: xf-m20-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M20 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3799 0.09528 0.09260 0.0016 0.6219 0.0062
-7.750 -0.3763 0.09170 0.08902 -0.0010 0.6175 0.0065
-7.500 -0.3717 0.08679 0.08411 -0.0054 0.6133 0.0070
-7.250 -0.3637 0.08103 0.07831 -0.0108 0.6092 0.0075
-7.000 -0.3479 0.07822 0.07547 -0.0137 0.6046 0.0076
-6.750 -0.3310 0.07531 0.07252 -0.0166 0.5995 0.0079
-6.500 -0.3132 0.07222 0.06938 -0.0197 0.5940 0.0082
-6.250 -0.2943 0.06871 0.06581 -0.0231 0.5893 0.0088
-6.000 -0.2748 0.06188 0.05889 -0.0290 0.5854 0.0101
-5.750 -0.2526 0.05960 0.05654 -0.0311 0.5800 0.0103
-5.500 -0.2296 0.05722 0.05409 -0.0332 0.5753 0.0106
-5.250 -0.2059 0.05454 0.05134 -0.0355 0.5704 0.0111
-5.000 -0.1790 0.04738 0.04399 -0.0402 0.5663 0.0132
-4.750 -0.1542 0.04570 0.04222 -0.0414 0.5608 0.0135
-4.500 -0.1289 0.04399 0.04044 -0.0426 0.5555 0.0138
-4.250 -0.1032 0.04215 0.03851 -0.0437 0.5502 0.0142
-4.000 -0.0767 0.03989 0.03614 -0.0449 0.5457 0.0151
-3.750 -0.0458 0.03272 0.02863 -0.0470 0.5423 0.0174
-3.500 -0.0196 0.03159 0.02740 -0.0475 0.5367 0.0177
-3.250 0.0067 0.03057 0.02630 -0.0480 0.5313 0.0180
-3.000 0.0336 0.02906 0.02469 -0.0484 0.5263 0.0183
-2.750 0.0609 0.02722 0.02270 -0.0488 0.5214 0.0188
-2.500 0.0887 0.02499 0.02028 -0.0490 0.5169 0.0193
-2.250 0.1170 0.02233 0.01740 -0.0490 0.5129 0.0203
-2.000 0.1452 0.01823 0.01291 -0.0486 0.5085 0.0215
-1.750 0.1725 0.01557 0.00989 -0.0483 0.5037 0.0220
-1.500 0.2002 0.01430 0.00839 -0.0483 0.4991 0.0221
-1.250 0.2282 0.01345 0.00735 -0.0483 0.4942 0.0222
-1.000 0.2564 0.01313 0.00692 -0.0484 0.4895 0.0224
-0.750 0.2846 0.01251 0.00616 -0.0485 0.4853 0.0224
-0.500 0.3127 0.01175 0.00525 -0.0486 0.4805 0.0225
-0.250 0.3402 0.01049 0.00377 -0.0486 0.4755 0.0231
0.000 0.3682 0.01008 0.00330 -0.0486 0.4712 0.0238
0.250 0.3965 0.00985 0.00304 -0.0488 0.4671 0.0242
0.500 0.4247 0.00968 0.00285 -0.0489 0.4625 0.0248
1.000 0.4810 0.00940 0.00250 -0.0492 0.4539 0.0257
1.250 0.5092 0.00926 0.00235 -0.0493 0.4495 0.0261
1.500 0.5372 0.00916 0.00222 -0.0494 0.4454 0.0264
1.750 0.5654 0.00906 0.00210 -0.0495 0.4417 0.0267
2.000 0.5935 0.00897 0.00200 -0.0497 0.4378 0.0269
2.250 0.6217 0.00892 0.00194 -0.0499 0.4332 0.0274
2.500 0.6498 0.00890 0.00191 -0.0500 0.4283 0.0279
2.750 0.6781 0.00885 0.00186 -0.0502 0.4240 0.0282
3.000 0.7062 0.00885 0.00183 -0.0504 0.4179 0.0284
3.250 0.7343 0.00887 0.00183 -0.0506 0.4112 0.0287
3.500 0.7625 0.00889 0.00185 -0.0508 0.4042 0.0293
3.750 0.7906 0.00894 0.00188 -0.0510 0.3997 0.0300
4.000 0.8189 0.00896 0.00193 -0.0513 0.3965 0.0318
4.250 0.8471 0.00899 0.00198 -0.0515 0.3925 0.0381
4.500 0.8751 0.00901 0.00209 -0.0517 0.3879 0.0822
4.750 0.9029 0.00904 0.00222 -0.0520 0.3833 0.1319
5.000 0.9217 0.00757 0.00249 -0.0505 0.3742 0.9747
5.250 0.9815 0.00786 0.00271 -0.0581 0.3564 1.0000
5.500 1.0072 0.00804 0.00285 -0.0580 0.3431 1.0000
5.750 1.0323 0.00837 0.00305 -0.0578 0.3175 1.0000
6.000 1.0563 0.00889 0.00337 -0.0576 0.2793 1.0000
6.250 1.0757 0.01019 0.00416 -0.0570 0.1885 1.0000
6.500 1.0932 0.01166 0.00514 -0.0563 0.0958 1.0000
6.750 1.1112 0.01287 0.00604 -0.0553 0.0268 1.0000
7.000 1.1345 0.01327 0.00642 -0.0549 0.0186 1.0000
7.250 1.1578 0.01363 0.00678 -0.0545 0.0156 1.0000
7.500 1.1809 0.01404 0.00720 -0.0540 0.0127 1.0000
7.750 1.2043 0.01438 0.00756 -0.0536 0.0118 1.0000
8.000 1.2272 0.01477 0.00798 -0.0532 0.0108 1.0000
8.250 1.2494 0.01523 0.00844 -0.0527 0.0097 1.0000
8.500 1.2706 0.01578 0.00904 -0.0521 0.0086 1.0000
8.750 1.2926 0.01621 0.00949 -0.0516 0.0081 1.0000
9.000 1.3138 0.01670 0.01001 -0.0511 0.0075 1.0000
9.250 1.3341 0.01724 0.01058 -0.0504 0.0070 1.0000
9.500 1.3533 0.01785 0.01121 -0.0497 0.0066 1.0000
9.750 1.3696 0.01865 0.01207 -0.0486 0.0061 1.0000
10.000 1.3854 0.01941 0.01290 -0.0475 0.0058 1.0000
10.250 1.3995 0.02020 0.01375 -0.0462 0.0056 1.0000
10.500 1.4080 0.02118 0.01479 -0.0443 0.0054 1.0000
10.750 1.4166 0.02240 0.01608 -0.0431 0.0052 1.0000
11.000 1.4264 0.02376 0.01750 -0.0423 0.0050 1.0000
11.250 1.4362 0.02524 0.01904 -0.0418 0.0047 1.0000
11.500 1.4456 0.02684 0.02071 -0.0414 0.0045 1.0000
11.750 1.4531 0.02869 0.02261 -0.0411 0.0044 1.0000
12.000 1.4577 0.03087 0.02487 -0.0407 0.0042 1.0000
12.250 1.4583 0.03351 0.02760 -0.0403 0.0040 1.0000
12.500 1.4571 0.03636 0.03054 -0.0399 0.0039 1.0000
12.750 1.4597 0.03882 0.03308 -0.0396 0.0039 1.0000
13.000 1.4604 0.04147 0.03582 -0.0393 0.0038 1.0000
13.250 1.4600 0.04428 0.03872 -0.0391 0.0038 1.0000
13.500 1.4586 0.04722 0.04175 -0.0388 0.0037 1.0000
13.750 1.4563 0.05027 0.04489 -0.0386 0.0036 1.0000
14.000 1.4537 0.05342 0.04813 -0.0385 0.0036 1.0000
14.250 1.4516 0.05663 0.05143 -0.0385 0.0035 1.0000
14.500 1.4492 0.05994 0.05483 -0.0386 0.0034 1.0000
14.750 1.4468 0.06332 0.05830 -0.0388 0.0034 1.0000
15.000 1.4448 0.06668 0.06174 -0.0390 0.0033 1.0000
15.250 1.4423 0.07013 0.06527 -0.0393 0.0032 1.0000
15.500 1.4397 0.07366 0.06888 -0.0396 0.0032 1.0000
15.750 1.4376 0.07718 0.07248 -0.0400 0.0031 1.0000
16.000 1.4356 0.08065 0.07603 -0.0405 0.0031 1.0000
16.250 1.4334 0.08421 0.07967 -0.0410 0.0030 1.0000
16.500 1.4310 0.08786 0.08340 -0.0415 0.0030 1.0000
16.750 1.4290 0.09146 0.08707 -0.0422 0.0029 1.0000
17.000 1.4267 0.09514 0.09083 -0.0429 0.0029 1.0000
17.250 1.4244 0.09883 0.09460 -0.0436 0.0028 1.0000
17.500 1.4219 0.10260 0.09845 -0.0444 0.0028 1.0000
17.750 1.4187 0.10653 0.10246 -0.0454 0.0027 1.0000
18.000 1.4155 0.11045 0.10646 -0.0463 0.0027 1.0000
18.250 1.4107 0.11466 0.11075 -0.0475 0.0026 1.0000
18.500 1.4055 0.11890 0.11509 -0.0486 0.0026 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M20 AIRFOIL (m20-il)