NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA M10 AIRFOIL (m10-il) Reynolds number: 100,000 Max Cl/Cd: 46.5 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m10-il-100000.txt Download as CSV file: xf-m10-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M10 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5714 0.10068 0.09588 0.0092 1.0000 0.0737
-8.000 -0.5728 0.09725 0.09250 0.0047 1.0000 0.0761
-7.750 -0.5734 0.09376 0.08900 -0.0068 1.0000 0.0777
-7.500 -0.5687 0.08895 0.08412 -0.0133 1.0000 0.0786
-7.250 -0.5580 0.08390 0.07920 -0.0055 1.0000 0.0814
-7.000 -0.5481 0.08012 0.07541 -0.0064 1.0000 0.0852
-6.750 -0.5346 0.07619 0.07125 -0.0166 1.0000 0.0916
-6.500 -0.5249 0.07068 0.06575 -0.0178 1.0000 0.0938
-6.250 -0.5128 0.06714 0.06227 -0.0158 1.0000 0.0984
-6.000 -0.4958 0.06272 0.05757 -0.0214 1.0000 0.1079
-5.750 -0.4812 0.05948 0.05434 -0.0203 1.0000 0.1169
-5.500 -0.4647 0.05533 0.05008 -0.0219 1.0000 0.1249
-5.250 -0.4454 0.05169 0.04623 -0.0238 1.0000 0.1372
-5.000 -0.4262 0.04840 0.04278 -0.0246 1.0000 0.1512
-4.750 -0.4074 0.04522 0.03952 -0.0246 1.0000 0.1663
-4.500 -0.3889 0.04262 0.03681 -0.0246 1.0000 0.1937
-4.250 -0.3709 0.03996 0.03412 -0.0239 1.0000 0.2223
-4.000 -0.3530 0.03730 0.03146 -0.0229 1.0000 0.2525
-3.750 -0.3363 0.03491 0.02910 -0.0214 1.0000 0.2943
-3.500 -0.2649 0.02685 0.01846 -0.0262 1.0000 0.1050
-3.250 -0.2389 0.02406 0.01540 -0.0256 1.0000 0.0997
-3.000 -0.2119 0.02195 0.01283 -0.0248 1.0000 0.0963
-2.750 -0.1867 0.02063 0.01132 -0.0241 1.0000 0.0999
-2.500 -0.1613 0.01941 0.00989 -0.0232 1.0000 0.1026
-2.250 -0.1362 0.01824 0.00853 -0.0224 1.0000 0.1036
-2.000 -0.1115 0.01732 0.00748 -0.0215 1.0000 0.1058
-1.750 -0.0878 0.01622 0.00649 -0.0207 1.0000 0.1101
-1.500 -0.0646 0.01554 0.00587 -0.0198 1.0000 0.1179
-1.250 -0.0422 0.01489 0.00536 -0.0190 1.0000 0.1320
-1.000 -0.0196 0.01420 0.00480 -0.0182 1.0000 0.1530
-0.750 0.0270 0.01098 0.00413 -0.0207 1.0000 1.0000
-0.500 0.0453 0.01116 0.00412 -0.0193 1.0000 1.0000
-0.250 0.0634 0.01141 0.00423 -0.0182 1.0000 1.0000
0.000 0.0817 0.01172 0.00441 -0.0173 1.0000 1.0000
0.250 0.1001 0.01210 0.00470 -0.0167 1.0000 1.0000
0.500 0.1432 0.01238 0.00490 -0.0208 0.9909 1.0000
0.750 0.1975 0.01255 0.00503 -0.0267 0.9765 1.0000
1.000 0.2521 0.01263 0.00511 -0.0326 0.9620 1.0000
1.250 0.3046 0.01265 0.00516 -0.0377 0.9468 1.0000
1.500 0.3518 0.01264 0.00522 -0.0415 0.9299 1.0000
1.750 0.3914 0.01265 0.00528 -0.0435 0.9108 1.0000
2.000 0.4229 0.01272 0.00540 -0.0437 0.8892 1.0000
2.250 0.4504 0.01282 0.00554 -0.0429 0.8681 1.0000
2.500 0.4747 0.01296 0.00576 -0.0414 0.8458 1.0000
2.750 0.4979 0.01313 0.00596 -0.0397 0.8236 1.0000
3.000 0.5206 0.01329 0.00615 -0.0378 0.8012 1.0000
3.250 0.5414 0.01336 0.00622 -0.0351 0.7724 1.0000
3.500 0.5610 0.01333 0.00619 -0.0320 0.7343 1.0000
3.750 0.5813 0.01331 0.00611 -0.0292 0.6926 1.0000
4.000 0.6025 0.01333 0.00606 -0.0266 0.6459 1.0000
4.250 0.6236 0.01341 0.00603 -0.0243 0.5803 1.0000
4.500 0.6426 0.01384 0.00604 -0.0216 0.4503 1.0000
4.750 0.6525 0.01723 0.00732 -0.0193 0.1204 1.0000
5.000 0.6743 0.01870 0.00867 -0.0182 0.0975 1.0000
5.250 0.6959 0.02015 0.01008 -0.0170 0.0848 1.0000
5.500 0.7176 0.02205 0.01178 -0.0160 0.0761 1.0000
5.750 0.7433 0.02364 0.01351 -0.0149 0.0715 1.0000
6.000 0.7678 0.02569 0.01549 -0.0143 0.0648 1.0000
6.250 0.7938 0.02827 0.01832 -0.0135 0.0624 1.0000
6.500 0.8195 0.03085 0.02127 -0.0125 0.0618 1.0000
6.750 0.8436 0.03386 0.02472 -0.0115 0.0619 1.0000
7.000 0.8654 0.03662 0.02802 -0.0102 0.0604 1.0000
7.250 0.8850 0.04056 0.03241 -0.0091 0.0617 1.0000
7.500 0.9031 0.04570 0.03847 -0.0071 0.0732 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M10 AIRFOIL (m10-il)