Naca 0012

Validation of AcuSolve for

Aerodynamics of Lifting

Bodies

� 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

Copyright 2010 ACUSIM Software, Inc. All rights reserved. www.acusim.com

� 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


�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


� 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

Copyright 2010 ACUSIM Software, Inc. All rights reserved. www.acusim.com5

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


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


� Drag Coefficient vs. Angle of Attack


� Lift Coefficient vs. Angle of Attack


� Drag Coefficient vs. Lift Coefficient


� Pressure Coefficient – alpha = 0


� 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


Naca 0012

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