NACA-M1 AIRFOIL (m1-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA-M1 AIRFOIL (m1-il) Reynolds number: 50,000 Max Cl/Cd: 20.63 at α=3° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m1-il-50000-n5.txt Download as CSV file: xf-m1-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA-M1 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5665 0.11358 0.10688 0.0100 1.0000 0.0584
-10.000 -0.5684 0.10914 0.10248 0.0088 1.0000 0.0574
-9.750 -0.7074 0.11293 0.10608 0.0153 1.0000 0.0482
-9.500 -0.7026 0.10855 0.10171 0.0154 1.0000 0.0472
-9.250 -0.7022 0.10370 0.09690 0.0133 1.0000 0.0466
-9.000 -0.7032 0.09859 0.09184 0.0103 1.0000 0.0463
-8.750 -0.7053 0.09318 0.08647 0.0065 1.0000 0.0459
-8.500 -0.7091 0.08768 0.08101 0.0024 1.0000 0.0454
-8.250 -0.7118 0.08220 0.07549 -0.0010 1.0000 0.0451
-8.000 -0.7132 0.07675 0.06996 -0.0039 1.0000 0.0449
-7.750 -0.7130 0.07138 0.06444 -0.0062 1.0000 0.0450
-7.500 -0.7103 0.06609 0.05892 -0.0080 1.0000 0.0452
-7.250 -0.7047 0.06091 0.05336 -0.0092 1.0000 0.0454
-7.000 -0.6960 0.05588 0.04793 -0.0098 1.0000 0.0455
-6.750 -0.6841 0.05106 0.04263 -0.0100 1.0000 0.0456
-6.500 -0.6690 0.04654 0.03756 -0.0098 1.0000 0.0461
-6.250 -0.6508 0.04261 0.03280 -0.0093 1.0000 0.0484
-6.000 -0.6323 0.03927 0.02932 -0.0091 1.0000 0.0521
-5.750 -0.6102 0.03613 0.02572 -0.0085 1.0000 0.0547
-5.500 -0.5860 0.03327 0.02228 -0.0078 1.0000 0.0586
-5.250 -0.5616 0.03084 0.01941 -0.0072 1.0000 0.0650
-5.000 -0.5359 0.02865 0.01689 -0.0065 1.0000 0.0693
-4.750 -0.5094 0.02696 0.01479 -0.0058 1.0000 0.0768
-4.500 -0.4842 0.02536 0.01315 -0.0052 1.0000 0.0845
-4.250 -0.4577 0.02395 0.01147 -0.0043 1.0000 0.0886
-4.000 -0.4324 0.02265 0.01006 -0.0036 1.0000 0.0937
-3.750 -0.4067 0.02156 0.00875 -0.0030 1.0000 0.1024
-3.500 -0.3819 0.02043 0.00762 -0.0025 1.0000 0.1187
-3.250 -0.3586 0.01896 0.00654 -0.0020 1.0000 0.1788
-3.000 -0.3437 0.01666 0.00593 -0.0001 1.0000 0.4845
-2.750 -0.3203 0.01563 0.00623 0.0047 1.0000 0.8387
-2.500 -0.2117 0.01581 0.00581 -0.0090 1.0000 1.0000
-2.250 -0.1909 0.01552 0.00526 -0.0084 1.0000 1.0000
-2.000 -0.1700 0.01529 0.00482 -0.0076 1.0000 1.0000
-1.750 -0.1489 0.01510 0.00445 -0.0068 1.0000 1.0000
-1.500 -0.1277 0.01495 0.00414 -0.0059 1.0000 1.0000
-1.250 -0.1064 0.01483 0.00387 -0.0050 1.0000 1.0000
-1.000 -0.0851 0.01473 0.00367 -0.0041 1.0000 1.0000
-0.750 -0.0639 0.01466 0.00352 -0.0031 1.0000 1.0000
-0.500 -0.0426 0.01461 0.00341 -0.0021 1.0000 1.0000
-0.250 -0.0213 0.01458 0.00334 -0.0010 1.0000 1.0000
0.000 0.0000 0.01457 0.00331 0.0000 1.0000 1.0000
0.250 0.0213 0.01458 0.00334 0.0010 1.0000 1.0000
0.500 0.0426 0.01461 0.00341 0.0021 1.0000 1.0000
0.750 0.0639 0.01466 0.00351 0.0031 1.0000 1.0000
1.000 0.0851 0.01473 0.00367 0.0041 1.0000 1.0000
1.250 0.1064 0.01482 0.00387 0.0050 1.0000 1.0000
1.500 0.1277 0.01495 0.00414 0.0059 1.0000 1.0000
1.750 0.1489 0.01510 0.00445 0.0068 1.0000 1.0000
2.000 0.1700 0.01529 0.00482 0.0076 1.0000 1.0000
2.250 0.1910 0.01552 0.00526 0.0084 1.0000 1.0000
2.500 0.2118 0.01580 0.00581 0.0090 1.0000 1.0000
2.750 0.3204 0.01563 0.00623 -0.0047 0.8381 1.0000
3.000 0.3437 0.01666 0.00593 0.0001 0.4837 1.0000
3.250 0.3586 0.01897 0.00654 0.0020 0.1783 1.0000
3.500 0.3819 0.02043 0.00762 0.0025 0.1187 1.0000
3.750 0.4068 0.02156 0.00875 0.0030 0.1024 1.0000
4.000 0.4324 0.02265 0.01006 0.0036 0.0937 1.0000
4.250 0.4578 0.02395 0.01147 0.0043 0.0886 1.0000
4.500 0.4842 0.02536 0.01315 0.0052 0.0845 1.0000
4.750 0.5094 0.02695 0.01479 0.0058 0.0768 1.0000
5.000 0.5359 0.02865 0.01690 0.0065 0.0693 1.0000
5.250 0.5617 0.03084 0.01941 0.0072 0.0650 1.0000
5.500 0.5860 0.03327 0.02228 0.0078 0.0586 1.0000
5.750 0.6102 0.03613 0.02571 0.0085 0.0547 1.0000
6.000 0.6323 0.03927 0.02932 0.0091 0.0521 1.0000
6.250 0.6508 0.04261 0.03280 0.0093 0.0483 1.0000
6.500 0.6691 0.04654 0.03756 0.0098 0.0461 1.0000
6.750 0.6841 0.05106 0.04264 0.0100 0.0456 1.0000
7.000 0.6960 0.05589 0.04794 0.0098 0.0455 1.0000
7.250 0.7048 0.06092 0.05336 0.0092 0.0454 1.0000
7.500 0.7104 0.06610 0.05893 0.0079 0.0451 1.0000
7.750 0.7130 0.07139 0.06445 0.0062 0.0450 1.0000
8.000 0.7134 0.07677 0.06998 0.0038 0.0449 1.0000
8.250 0.7120 0.08221 0.07551 0.0009 0.0451 1.0000
8.500 0.7094 0.08771 0.08103 -0.0025 0.0454 1.0000
8.750 0.7056 0.09322 0.08651 -0.0066 0.0459 1.0000
9.000 0.7036 0.09863 0.09188 -0.0104 0.0463 1.0000
9.250 0.7027 0.10373 0.09693 -0.0134 0.0466 1.0000
9.500 0.7031 0.10858 0.10175 -0.0155 0.0472 1.0000
9.750 0.7081 0.11295 0.10611 -0.0153 0.0482 1.0000
10.000 0.5678 0.10895 0.10229 -0.0087 0.0575 1.0000
10.250 0.5660 0.11337 0.10668 -0.0099 0.0585 1.0000
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Polar data table (+)
Polar graphs
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