FloEFD Aerospace Validation and Test- Cases€¦ · — Most CFD comparisons are run fully...

FloEFD Aerospace Validation and Test-Cases

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© 2010 Mentor Graphics Corp.

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Contents

1. Introduction a. Background b. Relevance of the Test-Cases c. Objective

2. Sources for Aerospace Test-Cases

3. Guidelines for Model Setup

4. Test-Cases: a. Sabjen Diffuser b. NACA 0012 Aerofoil c. RAE 2822 Aerofoil d. NLR 7301 Aerofoil with Trailing Edge Flap (work in progress) e. ONERA M6 Wing f. Future Work

5. Recommendations

6. References

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Introduction – Objective

Determine a set of “best-practice guidelines” for solving these types of problem using FloEFD

Ideally we want to use FloEFD “out of the box”

— Retain the benefits of ease of use & fast model setup

— Use automated settings as much as possible

— Avoid dependence on advanced functionality

– Needs to be accessible to the average user

A basic working knowledge of aerodynamics is assumed, i.e.

— Formulation of aerodynamic coefficients etc.

— We have some idea of the kind of solution we expect.....

– E.g. Shock waves, flow separation

— .....but no detailed a priori knowledge of the answer

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Sources for Aerospace Test-Cases

NPARC Alliance CFD Verification and Validation Archive — http://www.grc.nasa.gov/WWW/wind/valid/archive.html

AGARD AR-138, “Experimental Data Base for Computer Program Assessment”, May 1979

AGARD AR-303, “A Selection of Experimental Test Cases for the Validation of CFD Codes”, August 1994.

AIAA CFD Drag Prediction Workshop — http://aaac.larc.nasa.gov/tsab/cfdlarc/aiaa-dpw/

AIAA CFD High Lift Prediction Workshop — http://hiliftpw.larc.nasa.gov/

A large list of test-cases can also be found here: — http://www.dept.aoe.vt.edu/~mason/Mason_f/CAtxtAppB.pdf

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Used here

future work!

http://www.grc.nasa.gov/WWW/wind/valid/archive.html

http://aaac.larc.nasa.gov/tsab/cfdlarc/aiaa-dpw/



http://hiliftpw.larc.nasa.gov/

http://www.dept.aoe.vt.edu/~mason/Mason_f/CAtxtAppB.pdf


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Guidelines for Model Setup (1)

Wings and aerofoils should normally be modelled as external flow as wind-tunnel data usually include wall-corrections

— Ducts, nozzles etc. are obviously internal flow cases!

Laminar flow, fully turbulent or both? — Most wind-tunnel models use a trip near the leading edge to ensure boundary layer transition

– (to reduce scale effects between model scale and full scale) – Occasionally wind-tunnel tests are conducted with both free and fixed transition

— Most CFD comparisons are run fully turbulent — In most cases we should use the turbulent flow only option

– The NLR7301 + Flap case shown here is one exception to this rule!

Ensure calculations are run at the correct Reynolds number and Mach Number

— Specify initial conditions in terms of: – Temperature and Density

– Assume a fixed temperature (say, 300K) for which we know the absolute viscosity, m – Density, r, can then be determined from Reynolds number, Re=rVl/m

– Mach number – X & Z components can be calculated from the angle of incidence, a

N.B. All calculations were run using the low Mach number option

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Goals (1)

Aerodynamics is primarily about getting the pressure distribution right, from which the forces & moments are then determined

— So we need to select appropriate goals

Most of these are Point and Surface Goals which are not used for convergence control, but instead used to define Equation Goals

— GG Mmax – ensures any shocks are stable

— PG’s for (p, T, r, V, u, w, M)∞ - used to calculate coefficients and

for background information

— Define these at a point on the upstream boundary, (x, y, z)=(xmin, 0.0, 0.0)

— SG’s for X & Z force components – used to determine force coefficients

— SG for minimum static pressure – used to measure the suction peak

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Goals (2)

EG’s:

— Normal Force Coefficient, CN = SG Force (Z)/(qS)

— Axial Force Coefficient, CA = SG Force (X)/(qS)

– q = ½rV2 = ½gp∞M2; S=reference area

— Lift Coefficient, CL = CNcosa – CAsina = CN.u/V – CA.w/V

— Drag Coefficient, CD = CAcosa + CNsina = CA.u/V + CN.w/V

— Negative of Minimum Pressure Coefficient,

-Cpmin = - (SG Min Static Pressure - p∞)/q

– The negative sign is just to give the goal a positive value I just think it makes goal plotting a bit neater!

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N.B. This is a neat trick using basic trigonometry! V = Velocity (Hypotenuse) u = x-velocity (Adjacent) w = z-velocity (Opposite) Defining the Equation Goal in this way means we can have a single goal that will work for any a, instead of having to edit the goal every time!

sina = O/H = w/V cosa = A/H = u/V


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Calculation Control Options

— We need to give the flow time to develop fully so increase minimum refinement number to 10

— However, we also want to avoid excessively long run times so set maximum travels to Manual = 15

— The default automatic convergence criteria are not adequate (they are far to loose) so we can use manual settings

– Although we don’t have a detailed knowledge of the solution in advance, we can use our background knowledge of aerodynamics to make an intelligent guess, hence the following suggestions:

– GG Max Mach Number = 0.001

– M∞ ≈ 0.1 - 1.0 for low speed & transonic flows

– CL = 0.001

– CL ≈ 0.1 - 1.0 is not uncommon, may be higher for high lift configurations

– -Cpmin = 0.1

– Suction peaks of Cp ≈ -0.1 to -1.0 are not uncommon

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Computational Domain

Recommended dimensions:

— -16.0m < x < 20.0m (+x points downstream)

— 0.0m < y < 10.0m (+y points spanwise in starboard direction)

– Symmetry boundary condition imposed at ymin

— -10.0m < z < 10.0m (+z points vertically upwards)

– Based on chord length, c=1.0m in all 2D cases and root chord, cr=0.8059m for ONERA M6 wing (3D)

– These dimensions can be scaled and used as a guideline for different length scales

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Mesh Settings

General Guidelines — Rule of thumb is to have ≈100 cells along the root chord — Switch on Solution Adaptive Meshing to ensure salient flow

features are well resolved – Switch from Tabular to Periodic Refinement, but otherwise the default

settings are adequate – For a fine grid solution increase Refinement Level from 2 to 3, but

avoid increasing it further to prevent very long run times

2D — Mesh Level 8 seems to work consistently well for 2D aerofoils — Away from the geometry the basic mesh is probably finer than

strictly necessary – Penalty for this (increased cell count, longer run times) is not very big

for 2D cases

— Model setup is simpler if we stick to automatic settings — Fairly easy to modify basic mesh if we want to reduce run times

even further

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Mesh Settings

3D — Automatic settings yield high number of cells even in the

basic mesh – => Long run times and high memory requirements!

— To avoid this we must manually define the basic mesh and use control planes to concentrate the grid around the model

— We can also define 1-2 nested regions of local initial mesh to achieve ≈100 cells along the root chord and to ensure a reasonable level of refinement in the near-field

– It is not strictly necessary to have cells of unit aspect ratio around wings

– We can stretch the cells in the spanwise direction as flow gradients in this direction are generally much smaller than in the chordwise direction

Nested mesh regions around

ONERA M6 Wing


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Sajben Diffuser

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Mach no. Contours, Weak Shock Case

Mach no. Contours, Strong Shock Case


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NACA 0012 Aerofoil

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Vladimir Gavriliouk & EvgenyOdintsov: “Conclusion: New Solution Adaptive Refinement (SAR 9.0) technique allows much more accurate NACA 0012 simulation in comparison with standard FloEFD.”

Experimental results taken from Ref.1 (N.B. read from graph as tabulated results were unavailable!) CFD flow conditions taken from Ref.2, which include wind-tunnel wall corrections. For comparison, additional CFD results are shown, taken from Ref.3

a=2.26°; M=0.799; Re=9.0x106


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NACA 0012 Aerofoil

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FloEFD v11 ER NACA 0012 Aerofoil a=2.26°; M=0.799; Re=9.0x106


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RAE 2822 Aerofoil

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FloEFD v11 ER RAE 2822 Case 6 a=2.44°; M=0.725; Re=6.5x106


Experiment FloEFD WIND


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RAE 2822 Aerofoil

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Baldwin & Barth – Ref.3


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NLR 7301 Aerofoil with Trailing Edge Flap

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FloEFD v11 ER NLR 7301 + Flap a=6.0°; M=0.185; Re=2.51x106

Reasonable agreement in CP for attached flow case. Suction peak is under-predicted which may explain some of the discrepancy in drag. Still working on extracting experimental data points for accurate plotting. FloEFD work still in progress!


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NLR 7301 Aerofoil with Trailing Edge Flap

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Good agreement in CL for attached flow

Drag is hugely over-predicted, even for attached flow case!!!

Onset of flow separation and stall is much earlier than seen in experiment


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ONERA M6 Wing

ONERA is the national aerospace research centre of France:

— Office National d'Études et de Recherches Aérospatiales

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“The Onera M6 wing is a classic CFD validation case for external

flows because of its simple geometry combined with complexities

of transonic flow (i.e. local supersonic flow, shocks, and turbulent

boundary layers separation). It has almost become a standard for

CFD codes because of its inclusion as a validation case in

numerous CFD papers over the years. In the proceedings of a

single conference, the 14th AIAA CFD Conference (1999), the

Onera M6 wing was included in 10 of the approximately 130

papers!”

(J W Slater, Ref.8)


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ONERA M6 Wing

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FloEFD v11 ER Medium Mesh

FloEFD v11 ER Fine Mesh

TFLO2000 by Medic et al – Ref.7

ONERA M6 a=3.06°; M= 0.8395; Re=11.7x106

Weak leading edge shock and strong mid-chord shock coalesce near the tip to form a l-shock (Greek “lambda”)

The main flow features are well captured by FloEFD although shock at mid-chord could be crisper


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ONERA M6 Wing

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Generally good agreement with experiment although pressure recovery after mid-chord shock could be better

Resolution of shock near the tip is improved with fine mesh, but further improvement is still required


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ONERA M6 Wing

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FloEFD v11 ER Medium & Fine Meshes

TFLO2000 by Medic et al – Ref.7


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Recommendations

Better communication and “cross pollination” between Sales, AE’s and Development! — Let’s make full use of all the expertise we have in-house!

For ducts and nozzles use a mass flow inlet boundary condition — Adjust mass flow to get correct pressure ratio between inlet and outlet

Use Low Mach Number option to ensure flow separations are modelled correctly

Use Fully Turbulent option in almost all cases — Only exception is if you are comparing to an experiment with free-transition

Use Solution Adaptive Meshing to ensure flow features are well resolved

Set appropriate goals for the aerodynamic features of interest — E.g. Force & moment coefficients, shock waves, suction peaks — Set appropriate Manual convergence criteria on goals if the Automatic settings are too loose

For 3D cases, avoid very large meshes & long run times by manually defining the basic mesh

— Use Control Planes to cluster grid around the model and expand smoothly away — Use nested regions of Local Initial Mesh near the model

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References

1. Harris, C D, “Two-Dimensional Aerodynamics Characteristics of the NACA 0012 Airfoil in the Langley 8-Foot Transonic Pressure Tunnel”, NASA Technical Memorandum 81927, April 1981

2. Maksymiuk, C M and Pulliam T H, “Viscous Transonic Airfoil Workshop Results Using ARC2D”, AIAA-87-0415, January 1987

3. Baldwin, B S and Barth, T J, “A One-Equation Turbulence Transport Model for High Reynolds Number Wall-Bounded Flows”, NASA Technical Memorandum 102847, August 1990

4. Cook, P.H., M.A. McDonald, M.C.P. Firmin, "Aerofoil RAE 2822 - Pressure Distributions, and Boundary Layer and Wake Measurements," Experimental Data Base for Computer Program Assessment, AGARD Report AR 138, 1979

5. Slater, J W, “NPARC Alliance Validation Archive - RAE 2822 Transonic Airfoil: Study #1”, http://www.grc.nasa.gov/WWW/wind/valid/raetaf/raetaf01/raetaf01.html, December 2002

6. Schmitt, V. and F. Charpin, "Pressure Distributions on the ONERA-M6-Wing at Transonic Mach Numbers," Experimental Data Base for Computer Program Assessment. Report of the Fluid Dynamics Panel Working Group 04, AGARD AR 138, May 1979

7. Medic, G, Kalitzin, G, Iaccarino, G and van der Weide, E, “Adaptive Wall Functions with Applications”, Unpublished

8. Slater, J W, “NPARC Alliance Validation Archive – ONERA M6 Wing”, http://www.grc.nasa.gov/WWW/wind/valid/m6wing/m6wing.html, July 2008

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