The current study expands the application of computational fluid dynamics to three-dimensional multi-element high-lift systems by investigating the flow dynamics created by a slat edge. Flow is computed over a three-element high-lift configuration using an incompressible Navier-Stokes solver with structured, overset grids processed assuming full turbulence with the one-equation Baldwin-Barth turbulence model. The geometry consists of an unswept wing, which spans the wind tunnel test section, a single element half-span Fowler flap, and a three-quarter span slat. Results are presented for the wing configured for landing with a chord based Reynolds number of 3.7 million. Results for the three-quarter span slat case are compared to the full-span slat and two-dimensional investigations

Cummings, Russell M.

English

NASA Technical Reports Server

NASA-CR-204349
California Polytechnic State University
San Luis Obispo, CA 93407
Computational Investigation of Flap-Edges
Final Report
April 1997
NASA Award NCC-2-857
Russell M. Cummings
Aeronautical Engineering Department
https://ntrs.nasa.gov/search.jsp?R=19970017402 2020-06-16T01:56:45+00:00Z
COMPUTATIONAL INVESTIGATION OF A PART-SPAN SLAT
Sandra L. Grabhcrn
Aeronautical Engineering Department
California Polytechnic State University, San Luis Oblspo, CA 93407
Russell M. Cummings
Aeronautical Engineering Department
California Polytechnic State University, san Luis Obispo, CA 93407
ABSTRACT
The current study expands thQ application of computational fluid dynamics to
three-dimensional multi-element high-llft systems by investigating the flow
dynamics created by a slat edge. Flow is computed over a three-element
high-lift configuration using an incompressible Navier-Stokes solver with
structured, overset grids processed assuming full turbulence with the
one-equation Baldwin-Barth turbulence model. The geometry consists of an
unswept wing, which spans the wind tunnel test section, a single element
half-span Fowler flap, and a three-quarter span slat. Results are presented
the wing configured for landing with a chord based Reynolds number of
3.7 million. Results for the three-quarter span slat case are compared to the
full-span slat and two-dimensional investigations.
INTRODUCTION:
h_gh-lift aerodynamics continues to be and important area of research due to the
powerful effect high-lift systems have on the overall design of an aircraft
In order to continue the advancement of the performance of hlgh-lift systems, a
better understanding of the fundamental flow physics is needed. A number of
authors have identified areas important to high-llft flows, such as viscous wake
interaction, confluent wakes and boundary layers, and separated flows.
Computational Fluid Dynamics (CFD) approaches have demonstrated an ability to
capture many of these effects in two dimensions (Refs. I, 2). This ability,
coupled with the detailed flow information provided by a CFD solution, suggested
a computational approach for this work. This study utilizes an incompressible
Navier-Stokes solver in an attempt to further the understanding of high-lift
flows with particular emphasis on the flow affected by the slat edge.
CFD has been used to analyze high-lift systems in many instances, but the vast
majority of this work has been confined to two dimensions. Such work has
contributed greatly to the understanding of the nature of high-lilt'flows,
and while further two dimensional studies still have much to offer, the study
of three-dimensional realistic high-lift flows is of great importance to
aircraft designers. A thorough three-dimensional study should investigate how
two-dimensional results can bs related to the three-dimensional flow phenomena.
This is a relatively young field of study, and the development and assessment of
the software tools and techD, iques required to model 3-D flows is vital to the
growth of the field, A 3-D computational analysis would enhance the
investigation of three-dlmensional flow physics, especially in conjunction with
• ilar experimental investigations.
The goal of this study was to apply the current computational tools to a Dasic
three-dlmenslonal high-lift system consisting of a main element with leading
slat and trailing flap. In this paper, a description of the investigated
geometry, flow solver, turbuleace model, and boundary conditions will be
presented. The grid strategy developed to resolve flow about a 3-D high-lift
system in an efficient manner is discussed. Computations of flow over a
three-element wing between wind tunnel walls will be compared with the results
of a proven 2-D flow solver to verify the computational approach and results.
The results for a 3/4-Span slatted wing (also with a half-span flap) will also
be presented. Conclusions about the computational approach and the flow physics
will be made.
GEOMETRY:
The geometries studied were all based on the NACA 63-215 Mod. B airfoil section,
equipped with 30% chord Fowler flap and convex-shaped slat. Figure (I) shows
the airfoil section with both slat and flap elements deployed.
The full-span slat case, shown in Figures (5-7), is a simple extrusion of the
slat across the leading edge of the wing, so all the spanwise slat stations are
identical. The wing fully spans a modeled wind tunnel test section, which is
modeled after the NASA Ames 7XI0 Foot Wind Tunnel. This full-span slat was
_ded to the three-dimenslonal grid presented in References (3-5), which
_scribe the full-spa_ wing with half-span flap geometry. The three-quarter
span slat case consists of a flapped airfoil over half of the tunnel span, and
an unflapped section comprising the second half of the span, with the slat
extending 3/4 of the tunnel span from the same wall as the flap extends. This
geometry is shown in Figure (8) .
NUMERICALINVESTIGATION:
FLCW SOLVER:
THE INS3D-UP (Incompressible Navler-Stokes Upwind) flow solver was used for this
work (Ref. 6). INS3D-UP is a robust and efficient code that has shown promise
in its ability to model three-dimenslonal high-ilft flows in the past (Ref. !).
The two-dimensional version of the code, INS2D-UP (Reg. 7) has demonstrated the
ability to accurately predict two-dlmensional high-lift flows (Refs. I, 2).
For the current three-element geometry in a landing configuration (i.e. low
speed), the exclusion of compressible effects was felt to be a minor sacrifice
for the efficiency gained by using an incompressible code. However, if time
permits, compressible calculations will be performed and compared to
incompressible results, noting any significant changes in predicted flow
patterns. INS3D utilizes the m@thod of artificial compressibility to directly
couple the flow velocities and pressures by adding a pressure fluctuation term
to the continuity equation. This scheme also converts the incompressible
Navier-Stokes equations into a hyperbolic form which can be best utilized
through the use of an upwind differencing scheme. In this case, the convective
flux terms are differenced with a flux-dlfference splitting scheme based on the
method of Roe. The viscous fluxes are discretized using a standard central-
difference formulation. The system of equations is solved with a Gauss-Seidel
type implicit line relaxation scheme. Convergence for the cases in this study
will require approximately on the order of 400 iterations for a grid with 2.25
million grid points. Typical run times are expected to be about 15 Cray C90
_rs.
TURBULENCE MODEL:
The Baldwln-Barth turbulence model was used for all of the cases discussed in
this report. It has been shown by a number of researchers that this model does
a reasonable job in predicting multi-element airfoil flows, and it was this
recommendation, along with its low computational requirements, that strongly
suggested the use of this model for this study. The current implementation
of turbulence models in INS3D-UP requires that the code be run in either a
fully laminar or turbulent mode, therefore, all of the cases described herein
will be run in a fully turbulent mode, with no modeling of transition.
COMPUTATIONAL GRIDS:
Structured, overset grids were used throughout this study due to the versatility
of the Chimera scheme. Individual grid zones were generated independently and
later combined into one multi-grid using the Chimera based PEGSUS (Ref. 8) code.
PEGSUS merges the grids into a single file and establishes interpolation links
between the grids. This code enables the user to create holes in overlapping
meshes and remove (blank) points which would fall outside of the computational
domain (such as inside a solid surface).
e slat grid was generated as a 2-D plane, using a C-grid topology, and added
-J grids previously generated by other researchers. Grid details can be seen
in Figures (5,6) and are fully described in References (3-5). The wall spacing
is approximately 10E-5 chords, while the circumferential spacing is 10E-3 chords
at the trailing edge of each element. The size of the 2-D slat grid was 121x60,
which was expanded into a 3-D surface by copying spanwise stacks to fill in zhe
third dimension. For the full-span slat case, 4S identical spanwlse grid plane_
were used. Seven zones were requ£red for this simulation: flapped main element,
flap, flap-edge, unflapped main element, patch for exposed surface between main
e3ement sections, slat, and wind tunnel. The span covers a two chord distance.
nwise grid density for the slat should have negligible effects on the flow
solution, so no further full-span refinements will be attempted. The final full
-span slat case grid contained 2.25 mill_on grid points.
Another zone will be added for the part-span slat calculation: the slat-edge
grid, which will increase the total grid points to approximately 2.5 million
grid points. It is likely that the existing main element grids, which currently
contain grid planes clustered about the tunnel mid-span, will need to be
modZfied to include more points in the slat-wake region in order to clarify flow
characteristics, which will further increase total grid points and calculation
time.
Small scale flow features such as secondary separation lines are expected to be
visible in the solutions, with adequate grid resolution, variations of grid
resolution wall be performed in order to develop an adequately resolved 3/4-span
slat solution.
BOUNDARY CONDITIONS:
Most solid surface points were modeled with a no-slip boundary condition, which
specifies zero velocity and a zero normal pressure gradient at the surface.
Other solid surfaces were specified as Chimera boundary points and were allowed
to be updated through interpolation (Ref. 4). Wake-cut boundary conditions
were needed for elements with C-grid topology, and those points were updated by
_irst order averaging cf points on either side of the wake cut. The wind
tunnel walls were represented by slip-walls which impose a zero normal gradient
for all flow variables. The tunnel inflow condition was prescribed with free
stream flow, while the tunnel outflow consisted of constant static pressure and
extrapolated velocity boundary conditions. The outer boundaries of all grids
inside the wind tunnel grid were Chimera boundaries.
RESULTS AND DISCUSSION:
The following section highlights some of the areas to be discussed in the paper.
All of the accompanying figures are draft copies only.
TW0-DIMENSIONA5 MULTI-ELEMENT AIRFOIL:
A two-dlmenslonal grid study concerning the effects of slat location was
performed. Figure (I) shows the layout and overset of the three element grids.
The outer boundary of the main element grid extends approximately 20 chord
lengths in all directions, and solutions were found for free stream flow
conditions. Figure (2) shows the pressure distribution about the airfoil,
contours colored by pressure. The pressure distribution is plotted in
Figure (3) for all three elements. The slat gap (the vertical distance between
increased in increments of 0.002 inches, and the resulting pressure
stributions are shown in Figure (4), where it can be seen that changing the
_p size resulted in nominal differences in pressure over the elements. The
numerical values of the lift coefficient did not vary significantly either, thus
it was decided to leave the slat in its original position. These solutions were
calculated at a Reynolds number of 3.7 million and an angle of attack of i0
degrees.
FULL-SPAN SLAT:
' full-span results will be presented as a validation case. The computational
results are for a free stream Reynolds number of 3.7 million and an angle of
attack of 10 degrees. The computational pressures and forces will be analyzed,
and small three-dimenslonal effects seen in the computations will be discussed.
The flows computed with INS2D-UP and INS3D-UP will be compared. A cross-
section of the unflapped main element with slat is shown in Figure (5), a
cross-section of the three-element slat-main-flap combination is shown in
Fl_dre (6), and _ three-dimensional representation of the multi-element airfoil
is given in Figure (7).
3/4-SPAN SLAT:
The majority of the results in the proposed paper will be focused on the three-
quarter span slat computations, which are also performed for a Reynolds number
of 3.7 million and an angle of attack of 10 degrees. A grid study will be
performed to determine how best to resolve slat-wake flow over the main element
and half-span flap. The llft distribution will be computed, and compared to the
full-span slat case results to determine the effect that the finite span slat
has on the total lift. The lift distribution results for several cross-sections
of this case will be compared to the two-dimenslonal results to better
illustrate the complexity of the flow field. A representation of the 3/4-span
sl_t, main element, half-span flap combination is given in Figure (8).
In addition, several types of flow visualization will be implemented to obtain
_alitatlve understanding of the flow features (such as streamline curvature,
s_paration lines, tip vortex formation, and wake position). The nature of the
flow will be clearly seen through the use of flow visualization.
Experimental analyses of slat-edge and slat-wake effects on an airfoil with a
half-span flap are scheduled in the NASA Ames 7xl0 Wlnd Tunnel for spring 1996
for configurations similar to those generated in this study.
REFERENCES:
I Q
,
o
Rogers, S.,
"Efficient Simulation of Incompressible Viscous Flow Over Single and
Multi-Element Airfoils',
AIAA Paper 92-0405, Jan. 1992.
Rogers, S.,
"Progress in High-Lift Aerodynamic Calculations",
AIAA Paper 93-0194, Jan. 1993,
Mathias, D.,
"Navler-Stokes Analysis of the Flow About a Plap Edge",
Master of Science Thesis, California Polytechnic State University,
San Luis Obispo, CA, 1994.
Mathias, D., Roth, K., Ross, J., Rogers, S., and Cummings, R.,
"Navier-Stckes Analysis of the Flow About a Flap Edge",
AIAA Paper 95-0185, AIAA 33rd Aerospace Sciences Meeting and Exhibit,
Reno, NV, Jan. 1995.
6_
7 ,
,
Mathlas, D., Roth, K., Ross, J., Rogers, S., and Cummings, R.,
"Computational Investigation of a Semi-Span Flap",
6th International Symposium on Computational Fluid Dynamics,
Lake Tahoe, NV, Sept. 1995.
Rogers, S.,
"Numerical Solution of the Incompressible _avier-Stokes Equations",
NASA TM i02!99, Nov. 1990.
Rogers, S., and Kwak, D.,
"Upwind Differencing Scheme for the Time-Accurate Incompressible
Navier-Stokes Equations",
AIAA Journal, Vol. 28, No. 2, Feb. 1990.
Tramel, T, and Subs, 3.,
"PEGSUS 4.0 User's Manual", AEDC TR-gl-8, June 1991.
_igure I: Two-dimensionalmulti-element grid
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, I
--_s f i1'
t
f/f J
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Figure 3: _ressure Distribution over Two-Dimensional Multi-elemerl _,r#uil
I I ! !
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Figure 4: Comparison of Pressure Distributions for variation of Slat Gap 12-D)
FiQure 5: UnflaDped Airfoil Grid Cross SectioF_
Fiqure 6: Flaoped Airfoil Grid Cross Section
Figure 7: 3-D, Full-spa_ Slat Grid Surface Representation



Computational Investigation of Flap-Edges

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