NACA 0010 (naca0010-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA 0010 (naca0010-il) Reynolds number: 200,000 Max Cl/Cd: 44.41 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca0010-il-200000.txt Download as CSV file: xf-naca0010-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 0010
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
------ -------- --------- --------- -------- -------- --------
-11.750 -0.7617 0.10155 0.09799 0.0000 1.0000 0.0494
-11.500 -0.8675 0.07010 0.06617 -0.0236 1.0000 0.0379
-11.250 -0.8918 0.06396 0.05989 -0.0267 1.0000 0.0375
-11.000 -0.9154 0.05906 0.05480 -0.0272 1.0000 0.0373
-10.750 -0.9373 0.05491 0.05041 -0.0254 1.0000 0.0371
-10.500 -0.9524 0.05047 0.04564 -0.0236 1.0000 0.0371
-10.250 -0.9607 0.04627 0.04103 -0.0215 1.0000 0.0374
-10.000 -0.9614 0.04295 0.03725 -0.0193 1.0000 0.0380
-9.750 -0.9624 0.03858 0.03237 -0.0171 1.0000 0.0392
-9.500 -0.9494 0.03597 0.02968 -0.0162 1.0000 0.0410
-9.250 -0.9324 0.03436 0.02796 -0.0152 1.0000 0.0428
-9.000 -0.9169 0.03218 0.02546 -0.0138 1.0000 0.0446
-8.750 -0.8995 0.03048 0.02341 -0.0124 1.0000 0.0470
-8.500 -0.8824 0.02830 0.02084 -0.0109 1.0000 0.0491
-8.250 -0.8622 0.02622 0.01871 -0.0102 1.0000 0.0516
-8.000 -0.8405 0.02518 0.01757 -0.0094 1.0000 0.0545
-7.750 -0.8182 0.02414 0.01635 -0.0084 1.0000 0.0575
-7.500 -0.7953 0.02328 0.01523 -0.0075 1.0000 0.0595
-7.250 -0.7736 0.02129 0.01323 -0.0067 1.0000 0.0629
-7.000 -0.7508 0.02041 0.01233 -0.0059 1.0000 0.0660
-6.750 -0.7276 0.01954 0.01137 -0.0050 1.0000 0.0689
-6.500 -0.7041 0.01891 0.01062 -0.0041 1.0000 0.0718
-6.250 -0.6833 0.01759 0.00935 -0.0030 1.0000 0.0755
-6.000 -0.6614 0.01684 0.00861 -0.0019 1.0000 0.0795
-5.750 -0.6391 0.01620 0.00790 -0.0008 1.0000 0.0843
-5.500 -0.6189 0.01532 0.00712 0.0006 1.0000 0.0921
-5.250 -0.5982 0.01456 0.00641 0.0018 1.0000 0.1048
-5.000 -0.5782 0.01367 0.00571 0.0032 1.0000 0.1334
-4.750 -0.5583 0.01280 0.00516 0.0044 1.0000 0.1879
-4.500 -0.5375 0.01214 0.00480 0.0054 1.0000 0.2474
-4.250 -0.5160 0.01162 0.00450 0.0065 1.0000 0.3041
-4.000 -0.4944 0.01115 0.00428 0.0075 1.0000 0.3600
-3.750 -0.4725 0.01075 0.00411 0.0085 1.0000 0.4151
-3.500 -0.4503 0.01041 0.00397 0.0095 1.0000 0.4676
-3.250 -0.4281 0.01008 0.00386 0.0106 1.0000 0.5230
-3.000 -0.4060 0.00979 0.00381 0.0118 1.0000 0.5809
-2.750 -0.3834 0.00955 0.00379 0.0129 1.0000 0.6323
-2.500 -0.3604 0.00937 0.00378 0.0141 1.0000 0.6777
-2.250 -0.3373 0.00923 0.00381 0.0152 1.0000 0.7208
-2.000 -0.3143 0.00915 0.00387 0.0165 1.0000 0.7644
-1.750 -0.2921 0.00912 0.00399 0.0180 1.0000 0.8068
-1.500 -0.2705 0.00916 0.00416 0.0197 1.0000 0.8465
-1.250 -0.2491 0.00928 0.00435 0.0214 1.0000 0.8838
-1.000 -0.2245 0.00946 0.00457 0.0224 1.0000 0.9182
-0.750 -0.1738 0.00969 0.00478 0.0180 0.9943 0.9431
-0.500 -0.1201 0.00986 0.00492 0.0129 0.9881 0.9610
-0.250 -0.0607 0.01000 0.00503 0.0066 0.9838 0.9715
0.000 0.0000 0.01004 0.00506 0.0000 0.9778 0.9778
0.250 0.0607 0.01000 0.00503 -0.0066 0.9715 0.9838
0.500 0.1201 0.00986 0.00492 -0.0130 0.9609 0.9881
0.750 0.1738 0.00969 0.00478 -0.0180 0.9431 0.9943
1.000 0.2245 0.00946 0.00457 -0.0224 0.9181 1.0000
1.250 0.2491 0.00928 0.00435 -0.0214 0.8838 1.0000
1.500 0.2705 0.00916 0.00416 -0.0197 0.8465 1.0000
1.750 0.2921 0.00912 0.00399 -0.0180 0.8068 1.0000
2.000 0.3143 0.00915 0.00387 -0.0164 0.7643 1.0000
2.250 0.3373 0.00923 0.00380 -0.0152 0.7208 1.0000
2.500 0.3604 0.00936 0.00378 -0.0141 0.6778 1.0000
2.750 0.3834 0.00955 0.00379 -0.0129 0.6324 1.0000
3.000 0.4060 0.00979 0.00381 -0.0118 0.5811 1.0000
3.250 0.4281 0.01008 0.00386 -0.0106 0.5230 1.0000
3.500 0.4503 0.01041 0.00397 -0.0095 0.4676 1.0000
3.750 0.4725 0.01075 0.00411 -0.0085 0.4152 1.0000
4.000 0.4944 0.01115 0.00428 -0.0075 0.3601 1.0000
4.250 0.5160 0.01162 0.00450 -0.0065 0.3041 1.0000
4.500 0.5375 0.01214 0.00479 -0.0054 0.2474 1.0000
4.750 0.5583 0.01280 0.00516 -0.0044 0.1879 1.0000
5.000 0.5782 0.01367 0.00571 -0.0032 0.1333 1.0000
5.250 0.5982 0.01456 0.00641 -0.0018 0.1048 1.0000
5.500 0.6189 0.01532 0.00712 -0.0006 0.0922 1.0000
5.750 0.6391 0.01620 0.00790 0.0007 0.0843 1.0000
6.000 0.6614 0.01683 0.00861 0.0019 0.0795 1.0000
6.250 0.6833 0.01759 0.00935 0.0030 0.0755 1.0000
6.500 0.7041 0.01891 0.01062 0.0041 0.0718 1.0000
6.750 0.7276 0.01954 0.01137 0.0050 0.0689 1.0000
7.000 0.7508 0.02041 0.01233 0.0059 0.0660 1.0000
7.250 0.7736 0.02129 0.01323 0.0067 0.0629 1.0000
7.500 0.7953 0.02328 0.01523 0.0075 0.0595 1.0000
7.750 0.8183 0.02413 0.01634 0.0084 0.0575 1.0000
8.000 0.8405 0.02518 0.01757 0.0094 0.0545 1.0000
8.250 0.8623 0.02622 0.01871 0.0102 0.0516 1.0000
8.500 0.8825 0.02831 0.02085 0.0109 0.0491 1.0000
8.750 0.8996 0.03049 0.02342 0.0123 0.0470 1.0000
9.000 0.9169 0.03218 0.02546 0.0138 0.0446 1.0000
9.250 0.9325 0.03436 0.02796 0.0152 0.0428 1.0000
9.500 0.9495 0.03595 0.02966 0.0162 0.0409 1.0000
9.750 0.9624 0.03860 0.03240 0.0171 0.0391 1.0000
10.000 0.9615 0.04296 0.03725 0.0193 0.0380 1.0000
10.250 0.9609 0.04628 0.04104 0.0214 0.0374 1.0000
10.500 0.9526 0.05049 0.04566 0.0235 0.0371 1.0000
10.750 0.9374 0.05494 0.05045 0.0253 0.0371 1.0000
11.000 0.9156 0.05910 0.05485 0.0271 0.0372 1.0000
11.250 0.8920 0.06403 0.05996 0.0265 0.0375 1.0000
11.500 0.8678 0.07018 0.06626 0.0234 0.0379 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 0010 (naca0010-il)