Flow computations around two-element and three-element configurations are presented and compared to detailed experimental measurements. The k-epsilon-v(exp 2)(bar) model has been applied and the ability of the model to capture streamline curvature effects, wake-boundary layer confluence, and laminar/turbulent transition is discussed. The numerical results are compared to experimental datasets that include mean quantities (velocity and pressure coefficient) and turbulent quantities (Reynolds normal and shear stresses)

Durbin, P. A.

Iaccarino, G.

English

NASA Technical Reports Server

Center Ior Turbulence Research
Proceedings ol Lhe Summer Program I996
23
Application of the k-e-v 2 model
to multi-component airfoils
By G. Iaccarino 1 AND P. A. Durbin _
Flow computations around two-element and three-element configurations_are pre-
sented and compared to detailed experimental measurements. The k - e - v 2 model
has been applied and the ability of the model to capture streamline curvature effects,
wake-boundary layer confluence, and laminar/turbulent transition is discussed. The
numerical results are compared to experimental datasets that include mean quan-
tities (velocity and pressure coefficient) and turbulent quantities (Reynolds normal
and shear stresses).
1. Introduction
An accurate prediction of turbulent flow over a wing is still a challenging prob-
lem. Even a two-dimensional computation over a multi-element airfoil close to
the maximum lift is an unsolved problem due to the complex geometry producing
complicated viscous flow.
Within the aircraft industry the design of high-lift devices is an important topic
which can have a major influence on the overall economy and safety of the aircraft.
Therefore, development and improvements of numerical tools capable of handling
separated viscous flows are of great interest. Computational methods for the de-
sign of high-lift systems are, traditionally, based on the viscous-inviscid interaction
approach with integral methods for boundary layers and wakes.
Today, due to developments in computer technology and improvements in numer-
ical algorithms, there is a renewed interest in the possibility of obtaining Reynolds
averaged Navier-Stokes solutions for high-lift flows. The main open topics in this
field of application are grid generation and turbulence modeling. The first one has
been addressed and partially solved with the introduction of the zonal methods. By
this way, the computational domain is divided into zones and the mesh and solutions
are computed independently; the matching conditions between different regions pro-
vide boundary conditions for the zones. In particular, multiple-zones meshes can
be either patched (pointwise continuous) or chimera (overlapping) grids. The use of
unstructured grids is another interesting answer to this problem and is still under
development for viscous applications.
1 CIRA, Centro Italiano Ricerche Aerospaziali, Italy
2 Stanford University
https://ntrs.nasa.gov/search.jsp?R=19970014653 2020-06-18T00:16:22+00:00Z
24 G. Iaccarino (_ P. A. Durbin
The other crucial point is the handling of the turbulence for such a complicated
flow situation. There is no shortage of numerical methods to take into account tur-
bulent fluctuations when solving Reynolds Averaged Navier-Stokes (RANS) equa-
tions based either on algebraic or differential equations. It is only the effectiveness
of the models that is at issue.
2. Numerical model
2.1 RANS flow solver
The numerical method is based on an extended version of the incompressible
Navier-Stokes two-dimensional (INS2D) code of Rogers and Kwak (Rogers, 1991).
The incompressible, Reynolds Averaged Navier Stokes equations are solved by an
artificial compressibility method. The basic technique is based on cell-vertex finite
differences over structured meshes. The spatial discretization scheme is a third-order
upwind biased for convective contributions and second-order centered for diffusive
terms. The time integration is implicit and the equations are solved in a coupled
way. The implicit matrices are inverted by ADI line relaxations.
2.2, Turbulence modeling
The turbulence model is based in part on the standard k - ¢ equations:
0tk + u. vk = Pk - _ + v. [(_+ _)vk], (1)
OtC. nt- U • _7c = C_lPk - C_2eT + V. [(v + )VE]. (2)
Another transport equation is introduced to model near-wall effects and the aniso-
tropy of the Reynolds stresses. This reads as
Otv-g + U Vt ,--7 k f --e
• = _ _2_ + v. [(. + .,)v;_], (3)
where v 2 can be regarded as the turbulent intensity normal to streamlines and
k f, the production of v 2, accounts for the redistribution of turbulence intensity
from the streamwise component. By using this equation 'wall-echo' effects are
automatically taken into account. The production of v 2 is modeled by means of an
elliptic relaxation equation for f (Durbin, 1991)
[;L2V2f - f = -_(C, - 1) - - C2¥-. (4)
In the previous equations time and length scales are computed as
T = max [_,6(_)1/2], L=CLmax[k3--/e2,C,(U_a)l/4]. (5)
Application of k - ¢ - t,2 to multi-component airfoils 25
The treatment of the wall boundary conditions for the turbulent quantities is
based on the asymptotic behavior of k and v 2. As y _ 0
k=O, k---, y 2 e (6)
v 2 = O, v 2 _ --y4e
The eddy viscosity is given by
ut = C_v2T.
(7)
(s)
The constants of the model are:
C t, = 0.19, ak = 1, at = 1.3,
C_1 = 1.55, C¢2 = 1.9
C1 = 1.4, C2 = 0.3, CL = 0.3, C¢ = 70. (9)
The space discretization of Eqs. (1) to (4) is the same used for the mean flow and
the time integration is based on the same implicit procedure. The equations are
solved as a coupled two-by-two block tridiagonal system (the mean flow is solved
as a coupled three-by-three system).
3. Two-component configuration
3.1 Experimental test conditions
The experimental test was conducted in the 7x10" wind tunnel at NASA Ames
Research Center, Moffett Field, California (Adair, 1989). The airfoil/flap config-
uration includes a NACA 4412 main airfoil section equipped with a NACA 4415
flap airfoil section. The geometric location of the flap was specified by the flap gap
(FG), the flap overlap (FO), and the flap deflection (hi). In this work, we are using
FG = 0.035c, FO = 0.028c and/i.¢ = 21.8 °, where c is the chord length of the main
airfoil. The angle of attack was set to a = 8.2 ° and flow conditions were specified as
Mach number M = 0.09 and Reynolds number Re = 1.8 × 106. Two-dimensionality
of the measurements was ensured by using fences, and the transition to turbulence
was enforced by using trips at the main airfoil leading edge and at the suction side
of the flap close to the flap pressure minimum.
3.2 Numerical test conditions
A two-dimensional model is used for the computations; it represents the midspan
section of the experimental set-up. The airfoil configuration was characterized by
the value of FG, FO, and 6 f indicated previously. The presence of the wind-
tunnel walls was taken into account because of the large blocking effects, as was
recommended by the experimental investigators (Adair, 1989); for simplicity, slip
boundary conditions were imposed on the wind-tunnel walls. The inlet and outlet
sections were set at 5 chords upwind and 15 chords downwind respectively to mini-
mize their effects on the computed flow field. The angle of attack and the Reynolds
26 G. Iaccarino _4 P. A. Durbin
FIGURE 1. View of the computational grid.
/ /
FIGURE 2. Close-up of the grid around the flap.
number were the same as the experiments, while the flow was assumed to be incom-
pressible. Transition trips were not accounted for: the flow is considered to have a
very low turbulence intensity at the wind tunnel inlet, and the model is allowed to
undergo its natural, bypass transition.
The computational grid was generated by FFA (Sweden Research Center) under
the auspices of the GARTEUR Action Group AG-25. A general view of mesh is
reported in Fig. 1, while a close-up of the grid around the flap is given in Fig. 2.
Due to the complexity of the geometry the computational grid was generated via
a multiblock approach; seven zones were created allowing very good resolution of
the mesh close to the airfoils (a C-type grid); about 100,000 total grid points were
used. The square trailing edges of both airfoils were also retained (see Fig. 2) even
though the resolution in the streamwise direction is quite limited.
Application of k -c- v 2 to multi-component airfoils 27
FIGURE 3. Computed streamlines.
-14
-12
-10
-8
-6
-4
-2
o
2
-3
-2.5
-2
-1.5
-1
Cp -o.5
0
0.5
\ .................. 1.5
2
0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 0 0.05 0.1 0.15 0.2 0,25 0,3 0.35 0.4
x/¢
FIGURE 4. Pressure distributions on the airfoil surface. --
0: measured data.
: computed results;
3.3 Results
The characterization of the flow field is reported in Fig. 3 by means of the stream-
lines. Only a portion of the flow domain is shown. The blocking effect of the wind
tunnel wails and the large curvature of the wake downstream of the flap are evident.
A little separation bubble is also present at the flap trailing edge, in accord with
the experimental findings.
3.3.1 Mean flow: pressure
The comparison between computed and measured pressure coefficient is reported
28 G. Iaccarino _ P. A. Durbin
-1 .;
-1
Cp
-0.8
-0.6
-0.4
-0.2
0
0.2
0.4
Wind Tunnel Working Section
, I , , ! , ,
_ Wind Tunnel Roof
Wind Tunnel Floo:
FIGURE 5. Pressure distributions on the wind tunnel walls.
results; 0: measured data.
-- : computed
in Fig. 4. These results can be compared to those published by Rogers et al. (1993).
The agreement is quite satisfactory for both the main airfoil and the flap. The
suction peak is very well captured on the main airfoil although the stagnation point
is completely misplaced. This is probably due to three-dimensional effects in the
experimental test as can be seen from Fig. 3 of (Adair, 1987). Another reasonable
explanation for this discrepancy is a difference between the geometric location of
the airfoil/flap configuration in the experimental and numerical models. It is worth
noting that the numerical results of Rogers (1993) show this same discrepancy in the
location of tile stagnation point. We point out that the geometry definition of the
airfoil/flap configuration (in terms of FG, FO, and 6t) is somewhat confusing and
this could have led to a different shape of the slot in the numerical and experimental
models.
The pressure peak over the flap is overpredicted and, in particular, located up-
stream with respect to the experimental one. The numerical model fails to capture
the correct pressure plateau at the trailing edge of the flap and, therefore, the sep-
aration region. In particular, the separation point is well captured (it is located at
7% upstream of the trailing edge) as is shown in Fig. 4, but the maximum height
of the recirculation bubble is underpredicted.
In Fig. 5, the pressure distribution over the wind tunnel walls is reported and
compared to the experimental findings. On the working section roof, the agreement
is satisfactory even though an overprediction of the pressure level is present. On the
other hand, at the floor, a shift in the pressure distribution is observed. However,
the grid resolution in the region is quite coarse. Note that as the inlet and the
1.6
1,4
1.2
0.8
0.6
0,4
02
_o'.2o
Application of k - e - v 2 to rnuZti-component airfoils 29
0.2 0.4 0.6 0.8 I 1.2 1.4
UFUe
2
1.8
1,6
1.4
1.2
OJ
0.6 l
0.4 i
O2
_ o o_.'
U/Ue
I_ 1.4
2
1,5
1
0,5
"0_02 0 0 04 0.6 0,8 1
U/lle
12 1.4
Rake 1 Rake 2 Rake 3
Rake 3
FIGURE 6. Mean velocity profiles: : computed results; $: measured data.
outlet are approached the pressure levels become constant. This shows that the
computational domain was large enough.
3.3.2 Mean flow: velocity
The mean velocity was measured at three locations using a hot-wire anemometer
and a 3-D laser velocimeter. The comparison between computed and experimental
x-component velocity is reported in Fig. 6.
The agreement is very encouraging even if there is a difference between computed
and measured flap boundary layer thickness. Comparisons with previous results
by Rogers (1993) confirm that the main differencies are related to a different gap
velocity off the surface of the flap leading edge. It is necessary to point out that in
the numerical model no transition trips are mounted on the flap and, therefore, the
development of the turbulent boundary layer is not the same as in the experiments.
3.3.3 Turbulence results
The evolution of the turbulent boundary layer on the flap surface can be analyzed
from Fig. 7, where the tangential skin friction is reported. The model is capable
of capturing the laminar/turbulent transition automatically as it is evident from
the skin friction rise in the leading edge region. In the work by Lien et al. (1996)
30 G. Iaccarino 6f P. A. Durbin
0.03
0.0_
0.02
"0._ I i I i i i i
O.g_, 1 1.05 1.1 1.1S 1.2 1.2'5 1.3
]K/¢
FIGURE 7. Computed tangential skin friction on the flap surface.
F ;i
/f
t /:
/ //
FIGURE 8. Turbulent kinetic energy contours.
the transitional flow in turbomachinery was investigated and the capability of the
k - ¢ - v 2 were stressed in detail.
In Fig. 8 the turbulent kinetic energy contours are reported to show the strong
interaction between the main airfoil wake and the inviscid jet coming from the slot.
It is also clear that on the lower surfaces of the main and flap the boundary layer
is laminar and very thin.
Application of k - _ - v 2 to multi-component airfoil$ 31
Cruise configuration Take-off configuration
FIGURE 9. Three-element airfoil configuration.
4. Three-component configuration
4.1 Experimental test conditions
The three element airfoil configuration of Fig. 9 was investigated in the Farnbor-
ough (UK) wind tunnel by I.R. Moir (private communications) in the frame of the
AGARD Working Group 14.
The geometric location of the flap and the slat with respect to the main airfoil
was prescribed as:
slat/wing overlap: SO = -0.01c
slat/wing gap: SG = 0.02c
slat deflection: (_ = 25 °
wing/flap overlap: FO = 0
wing/flap gap: FG = 0.023c
flap deflection: _I = 20°
A set of angles of attack were investigated, but relevant measurements correspond
to _ = 20 °. The Reynolds number was Re = 3.52 x 106 and a trip was mounted over
the main airfoil to control transition to turbulence since the wind tunnel turbulence
intensity very low.
Experimental data include pressure surface measurements over the airfoil surface
at two spanwise locations on the wind tunnel model to outline the absence of three-
dimensional effects.
32 G. Iaccarino t"4 P. A. Durbi'n
\
i
2__
(a) (b)
FIGURE 10. (a) Close-up of the computational grids around tile slat; (b) Close-up
of the computational grids around the flap.
4.2 Numerical test condition._
The airfoil configuration was defined using the gap and overlap definitions of the
preceding section. In this case, the presence of wind tunnel walls was not taken into
account, but a correction of the angle of attack (as suggested by the experimental
investigators) was adopted: in particular an incidence of (_ = 20.18 ° was used for
the computations. The far field boundaries of the computational domain were set
to a distance of _ 20 chords from the airfoil. The Reynolds number was the same
used in the experiment and the flow was assumed to be incompressible. Laminar
to turbulent transitions were not fixed.
The computational grid was generated by Rogers (private communication), using
three different overlapping zones. Fig. 10 (a) reports the mesh around the slat and
the main airfoil leading edge, while Fig. 10 (b) reports the grid around the main
airfoil trailing edge and the flap.
4.3 Results
The pressure distributions over the surface of the three airfoil elements are re-
ported in Fig 11. Comparison with the experimental findings is very promising: the
distributions over the main wing and the flap are in very good agreement.
The Cp distribution over the slat presents an overprediction of the suction peak
and this is the main discrepancy between experiments and calculations.
;* °
m
Application of k - ¢ - v2 to multi-component airfoils
-8
-6
.4
.2
0
2 0:1
t
x/c
33
x/c
FIGURE 11. Pressure distributions on the airfoil surface.
0: measured data.
: computed results;
REFERENCES
ADAIR, D. & HORNE, W. C. 1989 Turbulent separated flow over and downstream
of a two-element airfoil. Ezperiments in Fluids. 7, 531-541.
DURBIN, P. 1991 Near-wall turbulence closure modeling without 'damping func-
tions'. Theoretical and Computational Fluid Dynamics. 3_ 1-13.
DURBIN, P. 1995 Separated flow computations with the k - ¢ - _2 model. AIAA
Journal. 33, 659-664.
LARSSON, T. 1994 Separated and high-lift flows over single and multi-element
airfoils. Proceedings of ICAS Conference 1994. 2505-2518.
LIEN, F. S., DURBIN, P. 1996 Non-linear k- e- _2 modeling with application
to high-lift. Proceedings of the Summer Program 1996. NASA Ames/Stanford
Univ.
PAPADAKIS, M., LALL, V. &: HOFFMANN K. A. 1994 Performance of Turbulence
models for planar flows: a selected review. AIAA Paper'94-1873.
ROGERS, S. E., KWAK, D. & KXRIS C. 1991 Numerical solution of the incom-
pressible Navier-Stokes equations for steady-state and time-dependent prob-
lems. AIAA Journal. 29, 603-610.
ROGERS, S. E., KWAK, D. & WILTBERGER N. L. 1993 Efficient simulation of
incompressible viscous flow over single and multielement airfoils. Journal of
Aircraft. 30, 736-743.
34 G. laccarino (Y P. A. Durbin
ROGEaS S. E., MENTER F. R., DUaBIN P. A. & MANSOUrt N. N. 1994 A com-
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Application of the k-epsilon-v(exp 2) model to multi-component airfoils

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