NACA M11 AIRFOIL (m11-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M11 AIRFOIL (m11-il) Reynolds number: 100,000 Max Cl/Cd: 50.43 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m11-il-100000.txt Download as CSV file: xf-m11-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M11 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4153 0.10316 0.09830 -0.0128 1.0000 0.0996
-9.500 -0.5001 0.10550 0.10032 -0.0096 1.0000 0.0925
-9.250 -0.4975 0.10232 0.09717 -0.0107 1.0000 0.0969
-9.000 -0.4101 0.09113 0.08635 -0.0149 1.0000 0.1071
-8.750 -0.4104 0.08750 0.08275 -0.0158 1.0000 0.1113
-8.500 -0.4256 0.08405 0.07937 -0.0192 1.0000 0.1146
-8.250 -0.4524 0.08059 0.07601 -0.0243 1.0000 0.1156
-8.000 -0.4296 0.07530 0.07071 -0.0207 1.0000 0.1192
-7.750 -0.4238 0.07181 0.06723 -0.0201 1.0000 0.1240
-7.500 -0.4390 0.06810 0.06360 -0.0222 1.0000 0.1272
-7.250 -0.4720 0.06444 0.05987 -0.0268 1.0000 0.1302
-7.000 -0.4561 0.05955 0.05509 -0.0242 1.0000 0.1340
-6.750 -0.4537 0.05627 0.05179 -0.0238 1.0000 0.1422
-6.500 -0.5044 0.06396 0.05905 -0.0231 1.0000 0.1407
-6.250 -0.5030 0.06048 0.05549 -0.0231 1.0000 0.1489
-6.000 -0.5022 0.05818 0.05295 -0.0228 1.0000 0.1614
-5.750 -0.4881 0.05477 0.04969 -0.0205 1.0000 0.1667
-5.500 -0.4830 0.05198 0.04679 -0.0193 1.0000 0.1788
-5.250 -0.4766 0.04957 0.04427 -0.0176 1.0000 0.1925
-5.000 -0.4690 0.04728 0.04192 -0.0155 1.0000 0.2071
-4.750 -0.4609 0.04494 0.03958 -0.0131 1.0000 0.2228
-4.500 -0.4526 0.04286 0.03758 -0.0102 1.0000 0.2414
-4.250 -0.4478 0.04105 0.03576 -0.0071 1.0000 0.2684
-4.000 -0.4115 0.03325 0.02561 -0.0073 1.0000 0.1157
-3.750 -0.3946 0.02968 0.02160 -0.0049 1.0000 0.1005
-3.500 -0.3786 0.02735 0.01911 -0.0030 1.0000 0.0972
-3.250 -0.3617 0.02570 0.01716 -0.0010 1.0000 0.0960
-3.000 -0.3440 0.02441 0.01559 0.0009 1.0000 0.0965
-2.750 -0.3250 0.02309 0.01400 0.0025 1.0000 0.0952
-2.500 -0.3050 0.02198 0.01266 0.0040 1.0000 0.0943
-2.250 -0.2848 0.02108 0.01159 0.0054 1.0000 0.0944
-2.000 -0.2650 0.02034 0.01073 0.0066 1.0000 0.0953
-1.750 -0.2326 0.01963 0.00993 0.0055 0.9961 0.0977
-1.500 -0.1876 0.01888 0.00921 0.0018 0.9872 0.1046
-1.250 -0.1444 0.01823 0.00864 -0.0015 0.9779 0.1124
-1.000 -0.0998 0.01761 0.00812 -0.0049 0.9689 0.1264
-0.750 0.0185 0.01424 0.00757 -0.0218 0.9847 1.0000
-0.500 0.0720 0.01439 0.00749 -0.0274 0.9730 1.0000
-0.250 0.1244 0.01447 0.00742 -0.0327 0.9610 1.0000
0.000 0.1778 0.01449 0.00731 -0.0381 0.9494 1.0000
0.250 0.2383 0.01437 0.00712 -0.0446 0.9402 1.0000
0.500 0.2891 0.01421 0.00693 -0.0492 0.9272 1.0000
0.750 0.3351 0.01404 0.00674 -0.0526 0.9128 1.0000
1.000 0.3760 0.01388 0.00657 -0.0548 0.8973 1.0000
1.250 0.4083 0.01380 0.00648 -0.0552 0.8795 1.0000
1.500 0.4359 0.01380 0.00647 -0.0547 0.8608 1.0000
1.750 0.4619 0.01381 0.00647 -0.0537 0.8426 1.0000
2.000 0.4864 0.01386 0.00651 -0.0524 0.8249 1.0000
2.250 0.5095 0.01394 0.00658 -0.0508 0.8072 1.0000
2.500 0.5309 0.01411 0.00675 -0.0491 0.7878 1.0000
2.750 0.5530 0.01425 0.00691 -0.0474 0.7693 1.0000
3.000 0.5755 0.01439 0.00704 -0.0457 0.7512 1.0000
3.250 0.5973 0.01454 0.00719 -0.0439 0.7316 1.0000
3.500 0.6186 0.01467 0.00733 -0.0420 0.7097 1.0000
3.750 0.6405 0.01475 0.00740 -0.0400 0.6878 1.0000
4.000 0.6615 0.01481 0.00745 -0.0379 0.6626 1.0000
4.250 0.6821 0.01484 0.00746 -0.0357 0.6344 1.0000
4.500 0.7026 0.01487 0.00745 -0.0335 0.6044 1.0000
4.750 0.7231 0.01494 0.00752 -0.0314 0.5736 1.0000
5.000 0.7431 0.01507 0.00764 -0.0293 0.5398 1.0000
5.250 0.7625 0.01524 0.00778 -0.0272 0.5014 1.0000
5.500 0.7807 0.01548 0.00797 -0.0249 0.4537 1.0000
5.750 0.7958 0.01593 0.00821 -0.0221 0.3780 1.0000
6.000 0.8004 0.01761 0.00893 -0.0182 0.2221 1.0000
6.250 0.8074 0.01972 0.01033 -0.0150 0.1510 1.0000
6.500 0.8206 0.02128 0.01164 -0.0125 0.1240 1.0000
6.750 0.8370 0.02267 0.01297 -0.0105 0.1060 1.0000
7.000 0.8551 0.02426 0.01447 -0.0088 0.0941 1.0000
7.250 0.8749 0.02603 0.01615 -0.0075 0.0838 1.0000
7.500 0.8972 0.02768 0.01795 -0.0061 0.0772 1.0000
7.750 0.9194 0.03007 0.02031 -0.0053 0.0712 1.0000
8.000 0.9396 0.03185 0.02243 -0.0037 0.0668 1.0000
8.250 0.9600 0.03419 0.02499 -0.0024 0.0640 1.0000
8.500 0.9781 0.03680 0.02780 -0.0010 0.0617 1.0000
8.750 0.9912 0.04074 0.03202 0.0005 0.0596 1.0000
9.000 0.9995 0.04300 0.03484 0.0034 0.0581 1.0000
9.250 1.0066 0.04650 0.03877 0.0058 0.0581 1.0000
9.500 1.0101 0.05042 0.04308 0.0082 0.0585 1.0000
9.750 1.0118 0.05490 0.04785 0.0104 0.0592 1.0000
10.000 1.0168 0.05866 0.05191 0.0124 0.0607 1.0000
10.250 0.9703 0.06356 0.05770 0.0179 0.0656 1.0000
10.500 0.9452 0.06778 0.06213 0.0207 0.0676 1.0000
10.750 0.9233 0.07214 0.06662 0.0218 0.0691 1.0000
11.000 0.8085 0.06777 0.06260 0.0248 0.0670 1.0000
11.250 0.7737 0.07516 0.07007 0.0216 0.0682 1.0000
11.500 0.7427 0.08360 0.07857 0.0174 0.0696 1.0000
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Polar data table (+)
Polar graphs
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