V Conferência Nacional de Mecânica dos Fluidos, Termodinâmica e Energia

V Conferência Nacional de Mecânica dos Fluidos, Termodinâmica e Energia
MEFTE 2014, 11–12 Setembro 2014, Porto, Portugal
© APMTAC, 2014
Ground-Effect Simulations based on an Efficient Boundary Element
Method
G. Farinha1, P. Quental1, L. Eça1, D. Matos Chaves2
1
Departmento de Engenharia Mecânica, Instituto Superior Técnico, Universidade de Lisboa, Av. Rovisco Pais 1,
1049-001 Lisboa, Portugal
email: [email protected]; [email protected]; [email protected]
2
Technip UK Limited, UK
email: [email protected]
ABSTRACT: This paper presents the application of a boundary element method (commonly known as a
panel method), which is able to solve the steady, two-dimensional, potential flow around airfoils, with an
arbitrary number of components, to the calculation of airfoils in the vicinity of a wall, i.e. including
ground-effect. The method requires modest computer requirements and extremely small calculation times
even for discretizations that lead to negligible numerical uncertainties. At small angles of attack, the
results obtained from this simple approach are qualitatively similar to those obtained experimentally.
KEY-WORDS: Boundary element method, airfoils, ground-effect, numerical uncertainty.
1
INTRODUCTION
The fast development of computer hardware has replaced mathematical models based on potential
flow theory by viscous flow solvers based on the Reynolds-Averaged Navier-Stokes (RANS) equations in
many practical applications including the calculation of the flow around airfoils, see for example [1].
However, there are specific problems, as for example optimization and/or uncertainty quantification,
where the number of solutions required makes the use of RANS solvers extremely time consuming.
Therefore, fast and robust potential flow solvers can still play a useful role in engineering applications.
In this paper, we address the calculation of the lift coefficient in single and two-component airfoils in
the vicinity of a wall, i.e. including the presence of ground effect. To this end we apply a boundary
element method (commonly known as a panel method) that is able to solve the steady, two-dimensional,
potential flow around airfoils with an arbitrary number of components with modest computer
requirements and an extremely small calculation time. The presence of the wall is modeled by the method
of images, which means that we will double the number of components of each test case. The goal of the
present work is twofold:
1.
Assess the grid density required to obtain negligible numerical uncertainties in the predicted lift
coefficient.
2.
Check the ability of this simple approach to obtain reasonable predictions of the ground effect in
the lift coefficient.
The first topic is addressed with grid refinement studies performed to a single element airfoil in an
unbounded domain and for an airfoil with a high-lift device in the vicinity of a wall. This type of studies
requires a careful definition of the airfoil geometry to allow the generation of geometrically similar grids
with an arbitrary number of panels.
Naturally, we cannot expect to make quantitative predictions of the lift coefficient based on potential
flow theory. Therefore, our aim is to compute the ratio between the lift coefficients obtained with and
without ground effect for single and double element airfoils and compare it with values obtained from
experiments.
The paper is organized in the following way: section 2 gives a brief description of the boundary
element method (panel method); the interpolation of airfoil geometries and grid generation is described in
section 3 and the results of the application to single and double element airfoils is presented and discussed
in section 4; Finally, section 5 summarizes the main conclusions of this study.
2
2.1
BOUNDARY ELEMENT METHOD
Mathematical formulation
The flow of an ideal fluid around an arbitrary number of lifting bodies (airfoils) is assumed to be
steady, two-dimensional, incompressible and irrotational. In such conditions, the velocity potential 

 

satisfies the Laplace equation     0 . The boundary conditions are obtained from the


impermeability condition on the airfoils surface   n  0 and the velocity vector at infinity is equal to


the undisturbed velocity   U  .
An efficient way to obtain  is to explore the linearity of the problem, see for example [2], and to
decompose the velocity potential into      , where   is the known velocity potential of the
undisturbed flow and  is the perturbation potential represented by line sources/sinks distributed on the
surface of the foils. All these velocity potentials satisfy the Laplace equation and so the application of the
boundary condition at each point P on the surface of the N c components of the airfoil determines the
strength  of the sources/sinks.
  P  N
2
c
 q 

ln  r  P, q   dS     nP  0
  
Si 2
nP
 i 1

(1)
The velocity potential defined by      gives a flow with no circulation, which does not
satisfy the Kutta condition, i.e. finite velocity at the trailing edge of all airfoil components. In the present
method, vortices are distributed along the mean line of each airfoil component (with a velocity potencial
defined by  i  i to guarantee the satisfaction of the Kutta condition.  i are
N airfoils
constants to be
determined by the application of the Kutta condition to each airfoil component. This leads to
  P  N
2
2.2
c
 q 
 Nc
ln  r  P, q   dS     i   nP    nP  0 (2)
  
S
i
 i 1 i 2 nP
 i 1
  
Discretization
There are three different approximations required for the transformation of equation (2) into a
system of algebraic equations: the geometric description of the airfoil components; the function that
defines the strength of the source/sink distributions and the distribution of the vortices along the mean
lines of the airfoil components.
Each airfoil component is discretized with NCPi flat panels (half on each side of the airfoil
component) which are defined by NBPi boundary points, which is the simplest choice available for the
geometric discretization. Nevertheless, it is a consistent approach, i.e. the discretization error tends to zero
when the number of panels/elements goes to infinity. In each linear element/flat panel, the source/sink
strength
 q
is assumed to be constant. This means that will have NCPi unknown source/sinks
strengths for each airfoil element.
The vortices that guarantee the satisfaction of the Kutta condition are located on mean line panels
that are obtained from the average of the surface boundary points. The strength of the vortex distribution
is piecewise constant and is proportional to sm0.4 [2], where sm is the distance of the mid-point of the
i
i
panel to the trailing edge.
The present discretization leads to
NTeq 
N airfoils
  NCP  1 unknowns that are determined by
i 1
i
satisfying the boundary condition of zero normal velocity (equation (2)) at the midpoint (control points)
Nc
of each surface panel (
 NCP equations) and satisfying the Kutta condition at the trailing edge of all
i 1
i
the airfoil components ( N c equations). Following the results presented in [2], the Kutta condition is
numerically satisfied by the equality of the tangential velocity components at the control points of the two
panels that contain the trailing edge. Therefore, we obtain a system of NTeq algebraic equations that has
the following structure for a single component airfoil:
NCP
A
j 1
ij
j
  Ai ( NCP 1)    ni for i  1 to NCP
NCP
A
j 1
where
(3)
 j   A( NCP 1)( NCP 1)    t1    tNCP
( NCP 1) j




Ai ( NCP 1)    nP , A( NCP1) j  B1 j  BNCPj and A( NCP 1)( NCP 1)    t1    tNCP .
Aij and Bij are the matrices of influence coefficients [2] and ti are the unit vectors tangent to the panels.
The system of
 NCP  1   NCP  1
linear equations can be designated by Aq  b . For N airfoils ,
the system of NTeq equations has a similar structure to that defined above Cij Q j  Bi , where each entry
of matrix Cij and vectors Q j and
Bi correspond to A , q and b defined above.
C11 C12 ... C1Nc  Q1   B1 


  
C21 C22 ... C2 Nc  Q2   B2 
 . . ... .
 .   . 


  
CNc 1 CNc 2 ... CNc Nc  QNc   BNc 
For example,
(4)
C12 contains the influence of the singularity distributions on the second component of the
airfoil in the first component.
With the present method it is straightforward to study ground effect using the method of images.
The ground is assumed to be a mirror and so we just have to include the mirrored airfoil components to
guarantee that the ground is a streamline.
2.3
Solution procedure
The method has been implemented in MATLAB© [3] with a modular structure. There are four
basic routines required: geometry definition that includes the determination of control points, panel length
and normal and tangential vectors; calculation of the influence coefficients and post-processing for the
determination of the pressure distribution on the surface of the airfoil components, which allows the
determination of the force coefficients by integration.
3
GRID GENERATION
Although grid generation for the present method seems to be a trivial problem, the fact that many
airfoil configurations are defined by tables of points leads to the need to use interpolation in the grid
generation procedure. Furthermore, typical grid generation procedures use cosine distributions along the
chord that must be done carefully to avoid geometric discretization errors that do not vanish with grid
refinement.
The present grid generation procedure relies on an interpolation procedure that splits any airfoil
component into three regions: leading edge, upper surface and lower surface. The generation of a grid for
an airfoil defined by a table of points that has the x coordinate aligned with the chord and the y coordinate
perpendicular to it includes the following steps:
1. Determine a 4th order polynomial of x=f(y) based on the leading edge (replaced by the point with
minimum x if the leading edge is not included) and the two surrounding points of the upper and
lower surfaces (5 points in total).
2.
Determine the minimum x coordinate defined by x=f(y), designated by xmin.
3.
Determine two 3rd order splines for the upper and lower surfaces of the airfoil including all
points available on the table except the leading edge and the nearest point to it. The firstderivatives on the left side (closest to the leading edge) are obtained from x=f(y) and the second
derivative at the trailing edge (xmax) is assumed to be zero.
4.
Make a cosine distribution along the line defined by xmin and xmax and determine the y
coordinates from the 4th order polynomial or the 3rd order spline of the upper and lower surfaces.
An example of the interpolation procedure is given figure 1 that presents a 160 panels grid of the
NACA 4412 airfoil with an enlarged view of the leading edge region.
81 Points
161 Interpolated
NACA 4412
81 Points
161 Interpolated
NACA 4412
Leading edge
xmin
Figure 1: Interpolation of a 161 points grid from a table of 81 points of the NACA 4412 airfoil.
Comparison with the analytical definition of the airfoil.
4
4.1
RESULTS
Test cases
We have studied ground effect for two different airfoil configurations: a single element airfoil [4]
and a two component airfoil [5,6] in two different configurations. Figure 2 illustrates the geometries of
the selected test cases. The plots include the images used to simulate the ground effect and the definition
of the distance to the ground h.
4.2
Numerical uncertainty
The first step of our study is the estimation of the numerical uncertainty of the calculated lift
coefficient Cl of the single and double element airfoils for different values of h/c where c is the chord for
the single element or the distance between the leading edge of the main element and the average location
of the trailing edge for the double element. c is also used as the reference length for the determination of
Cl. To this end we have made 13 grids for each case that range from 160 to 1280 panels per airfoil
element, which correspond to a grid refinement ratio of 8. The numerical uncertainty is determined with
the procedure proposed in [7] based on the data of the 5 finest grids.
Figure 3 presents the convergence of the lift coefficient Cl with the grid refinement as a function of
the typical grid size hi/h1 for the three geometries tested and different values of h/c. The results suggest
the following remarks:
h
h
Main element
Low flap
High flap
Figure 2: Illustration of the single component airfoil and of the two configuration of the double
component airfoil.
2.5
4.5
6.5
4.25
2.25
2
h/c=
1h+h2
h/c=0.45
1h+h2
h/c=0.67
1h+h2
h/c=0.31
1h+h2
h/c=0.31
1h+h2
6
4
5.5
3.75
3.5
5
1.75
Cl
Cl
Cl
3.25
3
4
2.75
1.5
2.5
2.25
1.25
2
h/c=
1h+h2
h/c=0.262
1h+h2
3.5
h/c=0.4
1h+h2
h/c=0.211
1h+h2
3
1
2
3
4
hi/h1
5
Single element
6
7
8
1.5
0
h/c=
1h+h2
h/c=0.34
1h+h2
h/c=0.247
1h+h2
h/c=0.4
1h+h2
h/c=0.284
1h+h2
h/c=0.211
1h+h2
2.5
1.75
1
0
4.5
1
2
3
4
hi/h1
5
Low flap
6
7
8
2
0
1
2
3
4
hi/h1
5
6
7
8
High flap
Figure 3: Convergence of the lift coefficient Cl of the three configurations (single and double element
airfoils) tested with the typical grid size hi/h1.
4.3

The convergence properties are similar for all the geometries tested. The best fit is obtained with
a two terms power series expansion, but the results exhibit a nearly linear behaviour with hi/h1
that changes slope for the three finest grids (the scale of the plots does not allow a clear
identification of this data behaviour).

The numerical uncertainty of the lift coefficient determination is smaller than that obtained from
the size of the symbols in the plots. For the present level of grid refinement, the largest estimated
uncertainty of all the cases plotted in figure 3 is smaller than 0.4%.
Ground effect
Naturally, we cannot expect to make realistic predictions of the lift coefficient of such
configurations with a model based on potential (ideal) fluid theory. Our aim is to explore the ability to
determine in a fast and efficient way the qualitative consequences of ground effect. Therefore, our
quantity of interest is the ratio between the lift coefficients obtained with and without ground effect.
Unfortunately, we do not have the experimental lift coefficient for the double element airfoils without
ground effect and so for those cases we have assumed that Cl∞ corresponds to the largest value of h/c.
Figure 4 presents the results obtained for the three geometries tested. The data suggest the following
remarks:

The qualitative effect of the ground effect is reasonably captured by the present potential flow
model, which was the main goal of the present exercise.

The numerical uncertainty of the calculations is much smaller than the difference between
predictions and experiments confirming the quantitative limitations of potential flow models,
especially for the smallest values of h/c tested.
Panel
2.5
Panel
1.6
Experimental, [4]
Panel
1.5
Experimental, [5]
Experimental, [5]
1.5
2.25
1.4
1.4
2
1.3
1.75
Cl/Cl
Cl/Cl
Cl/Cl
1.3
1.2
1.2
1.1
1.1
1.5
1
1
1.25
0.9
0.9
1
0.1
0.2
0.3
0.4
h/c
0.5
Single element
0.6
0.7
0.8
0.1
0.2
0.3
h/c
Low flap
0.4
0.5
0.8
0.2
0.3
h/c
0.4
High flap
Figure 4: Ratio of lift coefficient Cl without and with ground effect as a function of the non-dimensional
distance to the ground h/c.
5
CONCLUSIONS
The present paper presents a simple potential flow model for the calculation of the flow around
airfoils with multi-components, which allows a fast and efficient evaluation of the ground effect.
Naturally, the method is limited to the evaluation of lift coefficients. Nevertheless, the qualitative trends
of ground effect are reasonably captured for three practical test geometries with very modest computer
requirements (at the level of a cell phone in nowadays!).
REFERENCES
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Interaction Method, International Shipbuilding Progress, 40(422), 137-163.
[3] MATLAB version 7.10.0. Natick, Massachusetts: The MathWorks Inc., 2010.
[4] Zerihan, J., & Zhang, X. (2000). Aerodynamics of a Single Element Wing in Ground Effect. Journal of Aircraft.
37, 1058-1064.
[5] Zhang, X., & Zerihan, J. (2003). Aerodynamics of a Double-Element Wing in Ground Effect. AAIA Journal. 41,
1007-1016.
[6] Zhang, X., Toet, W., & Zerihan, J. (2006). Ground Effect Aerodynamics of Race Cars. ASME Appl. Mech.
Rev. 59(1), 33-49
[7] Eça L. Hoekstra M. (2014). A procedure for the estimation of the numerical uncertainty of CFD calculations
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