NASA/LANGLEY LS(1)-0013 AIRFOIL (ls013-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: NASA/LANGLEY LS(1)-0013 AIRFOIL (ls013-il) Reynolds number: 50,000 Max Cl/Cd: 26.38 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ls013-il-50000.txt Download as CSV file: xf-ls013-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY LS(1)-0013 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.7852 0.08283 0.07574 -0.0152 1.0000 0.1690
-9.750 -0.7993 0.07697 0.06983 -0.0163 1.0000 0.1659
-9.500 -0.8684 0.06822 0.06047 -0.0175 1.0000 0.1546
-9.250 -0.8710 0.06337 0.05535 -0.0171 1.0000 0.1542
-9.000 -0.8723 0.05881 0.05039 -0.0164 1.0000 0.1542
-8.750 -0.8697 0.05457 0.04565 -0.0155 1.0000 0.1546
-8.500 -0.8613 0.05065 0.04122 -0.0146 1.0000 0.1552
-8.250 -0.8418 0.04710 0.03764 -0.0144 1.0000 0.1585
-8.000 -0.8233 0.04438 0.03473 -0.0139 1.0000 0.1637
-7.750 -0.8066 0.04146 0.03133 -0.0132 1.0000 0.1678
-7.500 -0.7870 0.03869 0.02819 -0.0126 1.0000 0.1729
-7.250 -0.7636 0.03659 0.02602 -0.0122 1.0000 0.1812
-7.000 -0.7410 0.03435 0.02347 -0.0116 1.0000 0.1886
-6.750 -0.7177 0.03260 0.02165 -0.0110 1.0000 0.1999
-6.500 -0.6935 0.03085 0.01994 -0.0104 1.0000 0.2122
-6.250 -0.6701 0.02931 0.01845 -0.0095 1.0000 0.2280
-6.000 -0.6489 0.02794 0.01710 -0.0081 1.0000 0.2494
-5.750 -0.6289 0.02655 0.01604 -0.0065 1.0000 0.2782
-5.500 -0.6145 0.02508 0.01500 -0.0042 1.0000 0.3229
-5.250 -0.6088 0.02308 0.01395 -0.0004 1.0000 0.4127
-5.000 -0.6045 0.02356 0.01601 0.0087 1.0000 0.6064
-4.750 -0.5926 0.02593 0.01833 0.0167 1.0000 0.6891
-4.500 -0.5725 0.02813 0.02034 0.0231 1.0000 0.7367
-4.250 -0.5406 0.03018 0.02213 0.0275 1.0000 0.7775
-4.000 -0.4587 0.03268 0.02419 0.0251 1.0000 0.8247
-3.750 -0.2979 0.03362 0.02445 0.0076 1.0000 0.8838
-3.500 -0.2261 0.03268 0.02322 -0.0001 1.0000 0.9183
-3.250 -0.1703 0.03151 0.02183 -0.0064 1.0000 0.9447
-3.000 -0.1146 0.03027 0.02042 -0.0132 1.0000 0.9680
-2.750 -0.0516 0.02879 0.01879 -0.0217 1.0000 0.9882
-2.500 -0.0080 0.02776 0.01767 -0.0271 1.0000 1.0000
-2.250 -0.0017 0.02757 0.01744 -0.0256 1.0000 1.0000
-2.000 -0.0001 0.02753 0.01737 -0.0231 1.0000 1.0000
-1.750 0.0000 0.02755 0.01736 -0.0204 1.0000 1.0000
-1.500 -0.0003 0.02758 0.01737 -0.0175 1.0000 1.0000
-1.250 -0.0006 0.02761 0.01738 -0.0146 1.0000 1.0000
-1.000 -0.0008 0.02764 0.01739 -0.0117 1.0000 1.0000
-0.750 -0.0008 0.02766 0.01740 -0.0087 1.0000 1.0000
-0.500 -0.0006 0.02767 0.01740 -0.0058 1.0000 1.0000
-0.250 -0.0003 0.02768 0.01741 -0.0029 1.0000 1.0000
0.000 0.0000 0.02768 0.01741 0.0000 1.0000 1.0000
0.250 0.0003 0.02768 0.01741 0.0029 1.0000 1.0000
0.500 0.0006 0.02767 0.01740 0.0058 1.0000 1.0000
0.750 0.0008 0.02765 0.01739 0.0087 1.0000 1.0000
1.000 0.0008 0.02763 0.01739 0.0117 1.0000 1.0000
1.250 0.0006 0.02760 0.01737 0.0146 1.0000 1.0000
1.500 0.0002 0.02757 0.01736 0.0175 1.0000 1.0000
1.750 0.0000 0.02753 0.01735 0.0204 1.0000 1.0000
2.000 0.0001 0.02752 0.01736 0.0231 1.0000 1.0000
2.250 0.0017 0.02755 0.01743 0.0256 1.0000 1.0000
2.500 0.0081 0.02775 0.01766 0.0270 1.0000 1.0000
2.750 0.0516 0.02877 0.01878 0.0217 0.9882 1.0000
3.000 0.1147 0.03025 0.02041 0.0132 0.9681 1.0000
3.250 0.1703 0.03149 0.02182 0.0064 0.9447 1.0000
3.500 0.2262 0.03267 0.02320 0.0001 0.9184 1.0000
3.750 0.2978 0.03361 0.02443 -0.0075 0.8839 1.0000
4.000 0.4587 0.03267 0.02418 -0.0251 0.8247 1.0000
4.250 0.5405 0.03017 0.02212 -0.0275 0.7775 1.0000
4.500 0.5723 0.02813 0.02034 -0.0231 0.7368 1.0000
4.750 0.5924 0.02593 0.01833 -0.0167 0.6892 1.0000
5.000 0.6044 0.02357 0.01602 -0.0087 0.6068 1.0000
5.250 0.6087 0.02307 0.01395 0.0004 0.4129 1.0000
5.500 0.6144 0.02507 0.01499 0.0042 0.3230 1.0000
5.750 0.6288 0.02655 0.01604 0.0066 0.2783 1.0000
6.000 0.6488 0.02794 0.01709 0.0082 0.2495 1.0000
6.250 0.6700 0.02930 0.01845 0.0095 0.2280 1.0000
6.500 0.6934 0.03085 0.01994 0.0104 0.2122 1.0000
6.750 0.7176 0.03260 0.02165 0.0110 0.1998 1.0000
7.000 0.7410 0.03435 0.02347 0.0116 0.1886 1.0000
7.250 0.7636 0.03659 0.02602 0.0122 0.1812 1.0000
7.500 0.7870 0.03869 0.02817 0.0126 0.1729 1.0000
7.750 0.8067 0.04146 0.03133 0.0132 0.1678 1.0000
8.000 0.8234 0.04438 0.03473 0.0139 0.1637 1.0000
8.250 0.8419 0.04711 0.03764 0.0144 0.1585 1.0000
8.500 0.8614 0.05066 0.04122 0.0146 0.1552 1.0000
8.750 0.8698 0.05457 0.04566 0.0155 0.1546 1.0000
9.000 0.8724 0.05882 0.05040 0.0163 0.1542 1.0000
9.250 0.8712 0.06338 0.05536 0.0170 0.1542 1.0000
9.500 0.8687 0.06824 0.06049 0.0174 0.1547 1.0000
9.750 0.7997 0.07701 0.06987 0.0163 0.1659 1.0000
10.000 0.7858 0.08287 0.07578 0.0151 0.1690 1.0000
10.250 0.7145 0.10049 0.09344 0.0020 0.2016 1.0000
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Polar data table (+)
Polar graphs
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