naca 0012
TRANSCRIPT
� Goal is to validate AcuSolve for external aerodynamics applications
� NACA 0012 has been studied in depth both numerically and experimentally� AcuSolve results will be compared against:
�“Accepted” CFD simulations using other software
�Available test data
� References
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� References� http://turbmodels.larc.nasa.gov/naca0012_val.html
� Ladson, C. L., "Effects of Independent Variation of Mach and Reynolds Numbers on the Low-Speed Aerodynamic Characteristics of the NACA 0012 Airfoil Section," NASA TM 4074, October 1988
� Gregory, N. and O'Reilly, C. L., "Low-Speed Aerodynamic Characteristics of NACA 0012 Aerofoil Sections, including the Effects of Upper-Surface Roughness Simulation Hoar Frost," NASA R&M 3726, Jan 1970
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� Simulation Conditions
� NACA 0012 airfoil geometry
�Blunt trailing edge
� Re = 6.0e6
� Incompressible flow
�For compressible cases, M=.15
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�For compressible cases, M=.15
� Angles of attack investigated:
�-4 - 20
� Fully turbulent
� Results of interest
� Lift coefficient
� Drag coefficient
� Pressure coefficient
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� NACA 0012 airfoil geometry constructed using acuBuildFoil (GONE script)
� Chord = 1.0
� Trailing edge thickness = .002*chord
� Bounding cylinder diameter = 500*chord
� Span = 50*chord (2-d flowfield simulation)
� Surface tagging used to identify faces for boundary
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� Surface tagging used to identify faces for boundary condition application/meshing
� Enables automated set-up of simulations
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Farfield
Symmetry
Airfoil Walls
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Surface Name Boundary Condition Description
Farfield
Arbitrary inflow/outflow
If inflow: velocity = fixed
eddy viscosity = 1x molecular
If outflow: pressure=0.0
Symmetry Symmetry constraints
Airfoil Walls No-slip walls
� All meshing performed using AcuMeshSim
� Extrusion used to produce a single element in the span
� Mesh growth rate = 1.2
� Leading/trailing edge mesh size = .001*chord
� Surface mesh size at center chord: .02*chord
� Boundary layer first layer height = 1.75e-6*chord (Y+~=1.0 for Re=6.0e6,aoa=0)
� Boundary layer growth rate 1.3
� Boundary layer num layers = 28
� Number of elements across trailing edge height = 8
� Far field mesh size = 100,000*chord
Simple Region Of Influence used up to a distance of .5*chord
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Far field mesh size = 100,000*chord
� Simple Region Of Influence used up to a distance of .5*chord� Base size = .02*chord
� Perimeter edges of airfoil meshed using the following expression (assuming airfoil centered at x=0):sizeMin = .001*chord
sizeMax = .02*chord
xMin = -.5*chord
xMax = .5*chord
xCenter = (xMin+xMax)/2
dx = (xMax-xMin)/2
fac = (dx – abs(x-xCenter))/dx
power = ln(sizeMax/sizeMin)
meshSize = sizeMin*exp(power*fac)
� Total number of nodes = 97,635
� Some optimization capability on model size, can be reduced
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� Drag Coefficient vs. Angle of Attack
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� Lift Coefficient vs. Angle of Attack
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� Drag Coefficient vs. Lift Coefficient
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� Pressure Coefficient – alpha = 0
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� Pressure Coefficient – alpha = 10
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� Pressure Coefficient – alpha = 15
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� AcuSolve results agree very well with both
experimental data and other simulations
� AcuSolve simulations run on unstructured meshes
�Extensive mesh sensitivity study performed to show that
AcuSolve SA results are within 2% of complete grid
convergence
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