NACA-M1 AIRFOIL (m1-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA-M1 AIRFOIL (m1-il) Reynolds number: 200,000 Max Cl/Cd: 32.91 at α=2° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m1-il-200000.txt Download as CSV file: xf-m1-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.7096 0.09320 0.08977 0.0154 1.0000 0.0426
-8.500 -0.7131 0.08791 0.08451 0.0111 1.0000 0.0435
-8.250 -0.7187 0.08188 0.07852 0.0048 1.0000 0.0441
-8.000 -0.7199 0.07546 0.07202 -0.0011 1.0000 0.0452
-7.750 -0.7212 0.06876 0.06491 -0.0078 1.0000 0.0478
-7.500 -0.6682 0.05024 0.04657 -0.0131 1.0000 0.0501
-7.250 -0.6579 0.04679 0.04312 -0.0124 1.0000 0.0518
-7.000 -0.6499 0.04281 0.03904 -0.0123 1.0000 0.0544
-6.750 -0.6414 0.03992 0.03531 -0.0122 1.0000 0.0611
-6.500 -0.6745 0.04549 0.04048 -0.0102 1.0000 0.0628
-6.250 -0.6563 0.04243 0.03751 -0.0100 1.0000 0.0658
-6.000 -0.6391 0.03932 0.03384 -0.0096 1.0000 0.0763
-5.750 -0.6198 0.03735 0.03161 -0.0093 1.0000 0.0894
-5.500 -0.5908 0.02774 0.02091 -0.0066 1.0000 0.0462
-5.250 -0.5667 0.02409 0.01680 -0.0057 1.0000 0.0442
-5.000 -0.5408 0.02184 0.01415 -0.0048 1.0000 0.0455
-4.750 -0.5140 0.02080 0.01272 -0.0041 1.0000 0.0481
-4.500 -0.4882 0.01829 0.00997 -0.0035 1.0000 0.0495
-4.250 -0.4628 0.01670 0.00837 -0.0030 1.0000 0.0531
-4.000 -0.4366 0.01584 0.00745 -0.0026 1.0000 0.0577
-3.750 -0.4106 0.01492 0.00646 -0.0020 1.0000 0.0601
-3.500 -0.3858 0.01381 0.00535 -0.0014 1.0000 0.0640
-3.250 -0.3604 0.01311 0.00468 -0.0009 1.0000 0.0694
-3.000 -0.3346 0.01253 0.00406 -0.0004 1.0000 0.0740
-2.750 -0.3092 0.01184 0.00340 0.0000 1.0000 0.0829
-2.500 -0.2843 0.01093 0.00282 0.0005 1.0000 0.1379
-2.250 -0.2717 0.00828 0.00249 0.0027 1.0000 0.6546
-2.000 -0.2547 0.00774 0.00257 0.0062 1.0000 0.8214
-1.750 -0.2266 0.00770 0.00266 0.0074 1.0000 0.9188
-1.500 -0.1685 0.00786 0.00274 0.0018 1.0000 0.9789
-1.250 -0.1058 0.00794 0.00267 -0.0056 1.0000 1.0000
-1.000 -0.0838 0.00783 0.00250 -0.0047 1.0000 1.0000
-0.750 -0.0622 0.00774 0.00237 -0.0037 1.0000 1.0000
-0.500 -0.0409 0.00768 0.00227 -0.0026 1.0000 1.0000
-0.250 -0.0202 0.00764 0.00222 -0.0014 1.0000 1.0000
0.000 0.0000 0.00763 0.00220 0.0000 1.0000 1.0000
0.250 0.0202 0.00764 0.00222 0.0014 1.0000 1.0000
0.500 0.0409 0.00768 0.00227 0.0026 1.0000 1.0000
0.750 0.0622 0.00774 0.00237 0.0037 1.0000 1.0000
1.000 0.0838 0.00782 0.00250 0.0047 1.0000 1.0000
1.250 0.1058 0.00794 0.00267 0.0056 1.0000 1.0000
1.500 0.1685 0.00786 0.00274 -0.0018 0.9789 1.0000
1.750 0.2266 0.00770 0.00266 -0.0074 0.9187 1.0000
2.000 0.2547 0.00774 0.00257 -0.0062 0.8216 1.0000
2.250 0.2717 0.00828 0.00249 -0.0027 0.6540 1.0000
2.500 0.2843 0.01094 0.00282 -0.0005 0.1369 1.0000
2.750 0.3092 0.01184 0.00340 0.0000 0.0829 1.0000
3.000 0.3346 0.01253 0.00406 0.0004 0.0740 1.0000
3.250 0.3604 0.01311 0.00468 0.0009 0.0694 1.0000
3.500 0.3858 0.01381 0.00535 0.0014 0.0640 1.0000
3.750 0.4106 0.01493 0.00646 0.0020 0.0601 1.0000
4.000 0.4367 0.01584 0.00745 0.0026 0.0577 1.0000
4.250 0.4628 0.01670 0.00837 0.0030 0.0531 1.0000
4.500 0.4882 0.01828 0.00996 0.0035 0.0495 1.0000
4.750 0.5140 0.02081 0.01274 0.0041 0.0481 1.0000
5.000 0.5408 0.02184 0.01415 0.0048 0.0455 1.0000
5.250 0.5667 0.02410 0.01680 0.0056 0.0442 1.0000
5.500 0.5908 0.02774 0.02091 0.0066 0.0462 1.0000
8.750 0.7101 0.09324 0.08981 -0.0156 0.0426 1.0000
9.000 0.7086 0.09804 0.09460 -0.0187 0.0416 1.0000
9.250 0.7088 0.10241 0.09895 -0.0210 0.0404 1.0000
9.500 0.7119 0.10604 0.10257 -0.0219 0.0389 1.0000
12.250 0.5762 0.14251 0.13907 -0.0210 0.0355 1.0000
12.500 0.5711 0.14527 0.14182 -0.0234 0.0349 1.0000
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Polar data table (+)
Polar graphs
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