Download Aerodynamic Influence of a Half-Span Model Installation for High

48th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition
4 - 7 January 2010, Orlando, Florida
AIAA 2010-684
Aerodynamic Influence of a Half-Span Model Installation
for High-Lift Configuration Experiment
Yuzuru YOKOKAWA *, Mitsuhiro MURAYAMA †, Hiroshi UCHIDA ‡, Kentaro TANAKA §,
Takeshi ITO ** and Kazuomi YAMAMOTO ††
Japan Aerospace Exploration Agency, Mitaka, Tokyo, 181-0015, JAPAN
Lowspeed wind tunnel experiment of a high-lift configuration aircraft model JSM (Jaxa
high-lift configuration Standard Model) is implemented at 6.5m by 5.5m low-speed wind
tunnel in JAXA (JAXA-LWT1). The aerodynamic influence of a half-span model installation
is assessed with changing a height of a spacer, which locates between a fuselage and wind
tunnel wall. Variation of the aerodynamic characteristics depending on the spacer height in
the experiment agrees well with computational results which is assumed wind tunnel wall
boundary layer. Aerodynamic coefficients obtained in the experiment become closer to the
result of a free-flight computation when the spacer height becomes lower. Local lift
distribution estimated by the pressure distribution around the wing cross section changes
continuously along spanwise direction when the spacer height increases or decreases. On the
other hand, the local drag changes mainly in the inboard area. The variation of the induced
drag evaluated by the relation between CD and square of CL shows that an effect of the
spacer installation mainly appears as an increment of the effective aspect ratio and resultant
changes in slope of the lift curve and reduction of the induced drag. However, an influence of
the flow interaction near wind tunnel wall is not negligible. An estimation of desirable height
of the spacer for the half-span model experiment is performed by comparing the
computational result for the full-span model in free-flight condition and the half-span model
installed on non-slip wall condition. Based on the effective aspect ratio of free-flight
condition, the spacer height approximately 2 ~ 3 times of the displacement thickness of the
floor boundary layer is a best candidate for the spacer height.
Nomenclature
C
η
Re
Ae
V
α, AoA
CL
CD
CM
cl
cd
Cp
=
=
=
=
=
=
=
=
=
=
=
=
mean aerodynamic chord of stowed configuration
dimensionless spanwise station from the center line of the fuselage
Reynolds number based on freestream velocity and mean aerodynamic chord
Effective aspect ratio
freestream velocity
angle of attack
lift coefficient
drag coefficient
pitching-moment coefficient
lift coefficient at local cross section
drag coefficient at local cross section
static pressure coefficient
*
Associate Senior Researcher, Civil Transport Team, Aviation Program Group, AIAA Member.
Researcher, Civil Transport Team, Aviation Program Group, AIAA Member.
‡
Engineer, IHI Aerospace Engineering Co., Ltd.
§
Engineer, Ryoyu Systems Co., Ltd.
**
Senior Researcher, Wind Tunnel Technology Center, Institute of Aerospace Technology, AIAA Senior Member.
††
Senior Researcher, Civil Transport Team, Aviation Program Group, AIAA Senior Member.
†
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Copyright © 2010 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
I.
Introduction
Ahigh-lift configuration is essential issues for an aerodynamic design because successful improvement of the
balance between aerodynamic efficiency and silent level has a strong impact on the operating cost and environments
ccurate prediction of the maximum lift coefficient, lift/drag ratio and their Reynolds number dependency for
around airports, such as reduction of weight, increase of payload, fuel consumption and aerodynamic noise emission.
Furthermore, shortening of the design process by a structure simplification, reduction of production cost and
maintenance cost are also expected1,2. Owing to the present development of CFD solving Reynolds-averaged
Navier-Stokes (RANS) equations and computer technologies, accurate prediction has been possible for the cruising
phase with the least wind tunnel test3.4. However, regarding the precise prediction of the aerodynamic forces in
landing or take-off phase, it is not easy to achieve the same accuracy compared with the cruising phase due to the
complicated geometry and the flow features. On the other hand, from an experimental point of view, an estimation
of aerodynamic performance in high-lift configuration has been conducted through a wind tunnel experiment with
scale models although data probably contain some uncertainties due to the difficulties in model construction,
influence of Reynolds number, some interferences and model deformation. Based on such background, in the recent
days, an efficient design process combining CFD with a wind tunnel experiment has been strongly required.
Therefore, the development of high-lift system design technology has been extensively performed from both
experimental and computational aspects5-8.
In Civil Transport Team of Japan Aerospace Exploration Agency (JAXA/CTT), a research program to make
design methodology for advanced high-lift system has been performed. In this research program, lowspeed wind
tunnel experiment using a realistic aircraft configuration model with a leading-edge slat, flaps, a fuselage and a
nacelle-pylon was conducted through October 2005 - July 2009. In the experiment, we focused our attention to
acquiring both precise validation data for CFD and detailed flow information for a prediction of lift performance in
actual aircraft configuration9,10. At the same time, computations with Reynolds-averaged Navier-Stokes (RANS)
equations were performed with both unstructured mesh method and multiblock structured mesh method11-13. As a
result of extensive work in the program, large amount of data and knowledge from both experimental and
computational results14-21 such as an interaction between a nacelle and wing, Reynolds number effect, an effect of
nacelle chine, optimum position of a slat and a flap, and so on, could be stored and is now being discussed. However,
in CFD workshop for a highlift configuration in Japan, it was shown that the slope of the lift curve and drag
coefficient almost agrees among many CFD results but it does not agree between CFDs and experiment. As one of
the reason for this disagreement between CFDs and experiment, influence of a half-span model installation in the
experiment is suspected.
In this paper, aerodynamic influence of a half-span model installation for highlift configuration experiment is
assessed and optimum height of a spacer (in other words, a stand-off or a peniche) is discussed. Some activities to
improve a test technique using half-span (semi-span) aircraft model have been reported. In NASA Langley research
center, a series of assessment for an influence of the stand-off height and a shape on the flow over the semi-span
model was performed. Computational and experimental examinations using a semi- and a full-span model revealed
that flow on whole of the upper surface of the wing is accelerated when the semi-span model is used. And at the
same time, an increased cross flow was observed on the inboard area of the upper wing surface22-24. They concluded
that a stand-off height equal to twice of the tunnel-empty wall boundary layer was best. And also, several filleted
and undercut nose shapes were found to eliminate the separation of the sidewall boundary layer which was
generated with the original shape24. In the European research program EUROLIFT25, 10% higher drag and
maximum lift in the free-flight calculations compared to experimental and numerical in-tunnel results were observed
when KH3Y highlift configuration half model was used in ETW. It was explained that the mounting of the half
model in the tunnel causes a redistribution of the velocity field due to cross flow velocity components in the plane of
symmetry of the half model. In ELIRET26 project, comparisons of the numerical and the ETW results were
presented indicating the peniche and wind tunnel influences on the flow around the DLR F11 half model high lift
configuration. The main peniche effect on the model flow was based on its additional flow displacement leading to
an additional flow velocity around the fuselage and the inboard wing compared to a configuration without a peniche.
The variation of the peniche height only changes the strength of the displacement effect, but it could not be avoided
in any case. Murayama et al.17,20 performed CFD to understand an influence of the spacer with three heights of the
spacer in the free-air conditions with/without the floor boundary layer. The results showed that the spacer and floor
boundary layer generated large positive and negative velocity changes in the plane to assume the symmetric
condition and changed the effective angle of attack locally near the fuselage. The changes reduced drag especially at
high angle of attack and generated the difference of a polar curve between the corrected wind tunnel test data and
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free-air computations. Through the investigations, a height of the boundary layer spacer related to the displacement
thickness of the floor boundary layer showed less difference with the results in the free-air conditions.
In this paper, at first, an effect of the height of the spacer on aerodynamic coefficients and pressure coefficient
are presented using experimental results with four heights of the spacer. Following that, in order to understand how
the existence of the spacer changes the performance of the wing, lift and drag is analyzed with the lift line theory
and surface flow pattern in the different spacer conditions are observed. Finally, estimation of a desirable spacer
height for the half-span model experiment is performed by comparing the computational result between a full-span
model in free-flight condition and the half-span model installed on non-slip wall condition.
II.
Model of High-lift Device Testing
A generic high-lift configuration aircraft model “JSM (Jaxa high-lift configuration Standard Model)” which
assumes a typical 100-passenger class regional jet airliner is used in the series of the research. Specification of the
assumed aircraft is as follows. The maximum take-off weight is about 100,000lbs, the cruising speed is M = 0.8, the
maximum speed is M = 0.84, the cruising range is 2300 n.mile, and the landing and the take-off distances are 4400ft
and 5000ft, respectively. The conceptual and basic design of this aircraft was planned by JAXA, on the other hand,
the precise design and construction of the model was carried out by the collaboration between an aircraft
manufacturer in Japan and JAXA. This model is available also in pressurized tunnels up to three times of
atmospheric pressure.
To maximize Reynolds number based on the
mean aerodynamic chord (MAC), the model was
a half-span type. As shown in Fig.1 the model
was equipped with a circular fuselage and a flow
through nacelle-pylon at under the wing and
three Flap Track Fairings (FTF) as well as highlift devices, so that the model can provide
detailed flow fields occurring in actual aircraft.
2.3m
The model geometries are 2.3m in the half
wing span, 4.9m in the fuselage length, 0.5m in
the fuselage diameter and 0.3m in the nacelle
diameter, which is about 17% similarity of the
assumed aircraft. The wing has 9.42 of the
aspect ratio and 33degrees of the leading-edge
sweep angle. The specification of the wing is
listed in Table 1. The extent of the leading-edge
4.9m
slat is 85% of semi-span length and the chord
length is 12% of stowed wing chord length,
which is supported by eight slat tracks. The flap
is 77% of semi-span length and 30% of stowed
wing chord length. There are double-slotted flap
Figure 1. JSM in JAXA-LWT1 test section.
at the inboard (up to 37%) and a single-slotted
flap at the outboard (up to 77%), as shown in Fig.
1. For the landing phase, deflection angles of the high-lift devices are 25deg for the slat, 35deg for the flap and
20deg for the aft-flap. The configurations of the high-lift devices are given in Table 2.
Table 1. Specification of the wing.
Half span length
2.3m
Mean aerodynamic chord
0.5292m
Wing area (half span)
1.1233m2
Sweep angle (leading-edge)
33degrees
Aspect ratio
9.42
Taper ratio
0.333
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Table 2. Configurations of the high-lift devices.
Wing elements
Gap[1/c]
Overlap[1/c]
Slat
Flap
Inboard section
0.015
0.004
Outboard section
0.024
0.006
Inboard section
0.011
0.007
Outboard section
0.015
0.010
0.007
0.004
Aft-flap
Angle[deg]
25
35
20
The model has total of 456 static pressure taps at seven cross sections on the wing, four cross sections on the
fuselage and also on the nacelle-pylon. Figure 3 shows the locations of the cross sections for the pressure
measurement on the wing and the fuselage. The great care was taken for layout of the tubes from static pressure
ports inside the slat or the flaps, because it is often reported7 that they tend to disturb the local flow if they are
exposed. Therefore, they are veiled inside the support tracks or the FTF.
η = 0.89
η = 0.77
η = 0.56
η = 0.41
η = 0.33
η = 0.25
η = 0.16
O-O
N-N
M-M
L-L
Figure 2. Cross sections for pressure tap.
III.
Facility and Experimental Apparatus
Wind tunnel experiment was conducted in 6.5m by 5.5m Lowspeed Wind Tunnel in JAXA (JAXA-LWT1),
which was an atmospheric pressure closed-circuit tunnel with an octagonal cross section. In this experiment with
JSM, Reynolds number based on MAC was up to about 2.5million.
The model was mounted vertically on the turntable for changing an angle of attack. To avoid interference between
the model and the boundary layer of the tunnel wall, a spacer was set below the fuselage. The spacer has the same
cross sectional shape with the fuselage symmetry plane. The symmetry plane of the fuselage is simply extended to
the tunnel wall. The spacer with 150mm height which corresponds to 99%-thickness of the wall boundary layer is
used originally9,10, however at this moment, four kinds of the spacer height are tested to know variation in
aerodynamic performance and to assess an optimum configuration for the half-model experiment as shown in Fig.3.
They were 0mm, 30mm, 80mm and 150mm. Height of 0mm means without spacer configuration. 30mm is nearly
equal to displacement thickness of the wall boundary layer.
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Under the model, five components moment-type balance was installed to measure two elements of forces and
three elements of moments. The accuracy of the balance is about 0.01 and 10 counts in lift and drag coefficient,
respectively, which is sufficient level in the aerodynamic testing for high lift devices27. The surface pressure
distribution was measured by electronic pressure scanners mounted inside the model. Wind tunnel wall interferences
by the model blockage, streamline curvature, downwash were corrected by the theoretical way that was invented by
Alan Pope28. The reliability of the correction was confirmed by comparing with the result given by the method
specially developed for JAXA-LWT1, which was applied the panel method and actually measured tunnel wall static
pressure. Surface flow patterns were visualized by two ways, the tuft method and the oil-flow method. The tuft
method can give overview of separation patterns briefly with changing an angle of attack. On the other hand, the oilflow visualization provides a lot of information for local flow phenomena. The flow pattern images given by each
method were taken by digital cameras and digital video camera during or after the wind tunnel running. Referenced
papers can give an opportunity to view the other measurements applied to this model throughout the series of our
experiment canpaign29-33.
Figure 3. Boundary layer spacers.
(left to right : without spacer/30mm/80mm/150mm)
IV.
Computational Method
As an unstructured mesh generator and flow solver, TAS (Tohoku university Aerodynamic Simulation) codes34,35
are used in this study. TAS_Mesh is a mesh generator with graphical user interface (GUI) tools36,37. It can generate
triangular surface mesh with the advancing front method and tetrahedral volume mesh using Delaunay tetrahedral
meshing38, as well as hybrid volume mesh composed of tetrahedrons, prisms, and pyramids for viscous flows with
high Reynolds number39. The unstructured surface meshing using isotropic triangles is semi-automatic and the
volume mesh generation is fully automated. In TAS_Flow, Navier-Stokes equations are solved on the unstructured
mesh by a cell-vertex finite volume method. HLLEW method40 is used for the numerical flux computations. Secondorder spatial accuracy is realized by a linear reconstruction of the primitive variables. LU-SGS implicit method41 is
used for time integration. A variation of Spalart-Allmaras turbulence model, which reduces the eddy viscosity in the
regions of high vorticity42,43, is used in the present computations to simulate turbulent flows. In this study, a simple
combination using the minimum of the vorticity and strain rate is utilized in the modification43. The modified model
computes turbulent vortical flow without adding much dissipation to the vortex core.
Two kinds of the computations are conducted for a full-span model in free-flight condition and the half-span
model installed on non-slip wall which assumed the configuration of the wind tunnel experiment, however the side
wall and ceiling is removed. Figure 4 shows the unstructured mesh for half-model computation. This mesh is
generated using that of the free-flight computation. Therefore, results from two kinds of computations can be
compared without errors which are caused by differences in computational technique. FTF and brackets to support
the high-lift devices are removed in the computations. The unstructured mesh has about 6~8 million mesh points in
volume mesh. The minimum spacing in the normal direction to the wing surface was 0.02/√Re. Fully turbulent
condition is applied. Computations are carried out on Fujitsu PRIMEPOWER HPC2500 multi-processor, which is
the main machine of Numerical Simulator III system in JAXA44. Details of the computation are shown in Murayama
et al11-17,20.
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Figure 4. Computational meshes to conduct parametric study of the height of the spacer.
V.
Results and Discussions
A. Basic Aerodynamics
In this section, basic aerodynamics obtained in the wind tunnel experiment for each of the spacer heights is
presented. In the measurement, the data are acquired in pitch pause mode as the angle of attack is increasing. All
data are provided after the correction of the tunnel wall and other interference by a method given by Alan Pope28.
The difference in angle of attack between before and after the correction is about 0.55degrees at the point of the
maximum lift coefficient condition for freestream velocity of 60m/s.
1. Forces, Moment
Figure 5 shows the change in lift, drag and pitching-moment coefficients with changing heights of the spacer at
freestream velocity of 60m/s. As shown in Fig.5(a), CL value and its slope increase as the spacer height becomes
higher. And, stall angle of attack and maximum lift value is changed depending on the spacer height. Furthermore,
kink at 4degree which appears on lift curve of 150mm spacer case is not observed in other cases. This indicates a
change in local flow physics on the wing. Figure 6 shows a relation between slope of the lift curve and spacer height.
The curve seems to approximate with a polynomial function. On the other hand, drag in the polar curve shown in
Fig.5(b) decreases as the spacer becomes higher, which indicates variation in induced drag. Pitching-moment shown
in Fig.5(c) reduces as the spacer becomes higher, but the relation between lift does not change in the region of
moderate lift value. This means that an aerodynamic center does not move in case the spacer height changes except
for the lower and higher lift conditions. Difference of the aerodynamic coefficients for the spacer height from
0mm(without spacer) to 150mm are approximately as follows; ∆CL = 0.08, ∆CD = 0.013, ∆CM = 0.03. Variation of
the aerodynamic characteristics with spacer height in the experiment agrees well with the computational result
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which is assumed a wind tunnel wall boundary layer17-20. All of the experimental results become closer to the result
of free-flight computation when the spacer height becomes lower.
3
2.5
2.5
CL
CL
3
2
2
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
1.5
0
5
10
15
20
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
1.5
25
0.1
0.2
0.3
AoA [deg]
0.4
0.5
CD
(a) CL-α
(b) CL-CD
3
0
2.5
CL
-0.4
2
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
-0.6
-0.8
0
5
10
15
20
25
1.5
-0.7
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
-0.6
-0.5
-0.4
AoA [deg]
CM
(c) CM-α
(d) CL-CM
Figure 5. Effect of the spacer height on aerodynamic coefficients.
0.11
0.105
dCL/d α
CM
-0.2
0.1
0.095
0.09
0
30
60
90
120
Height[mm]
150
180
Figure 6. Slope of the lift curves versus spacer heights.
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-0.3
-0.2
-0.1
2. Static Pressure
Figure 7 shows static pressure distributions at the in- and the outboard cross sections η = 0.16 to 0.56 when the
spacer heights are changed. Results of three kinds of angle of attack conditions at freestream velocity of 60m/s are
compared. Horizontal axis is the location from the nose of the fuselage which is normalized by mean aerodynamic
chord. In the case of lower angle of attack AoA=4deg shown in Fig.7(a), difference in the pressure distribution is
little on both the in- and the outboard cross sections for each of the spacer heights. However, in higher angle of
attack condition AoA=10deg shown in Fig.7(b), the pressure on the slat and the leading-edge of the main-wing has
difference depending on a spacer height. And further, that character is visible also in larger angle of attack case at
AoA=13deg shown in Fig.7(c). Although a difference in pressure increases when an angle of attack becomes larger,
the location on the wing where the pressure distribution is affected by a spacer installation seems limited near the
leading edge region. The tendency in this result agrees with that of the computation by Murayama et al.17,20.
-6
-5
-4
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
-5
-4
-3
Cp
Cp
-6
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
-2
-3
-2
-1
-1
0
0
1
1
3.4
3.9
4.4
X/C
4.9
5.4
4.5
5
X/C
5.5
(a) AoA = 4deg
-6
-5
-4
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
-5
-4
-3
Cp
Cp
-6
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
-2
-3
-2
-1
-1
0
0
1
1
3.4
3.9
4.4
X/C
4.9
5.4
4.5
5
X/C
5.5
(b) AoA = 10deg
-6
-4
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
-5
-4
-3
Cp
Cp
-6
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
-5
-2
-3
-2
-1
-1
0
0
1
1
3.4
3.9
4.4
X/C
4.9
5.4
4.5
5
X/C
(c) AoA = 13deg
Figure 7. Static pressure distributions.
(Left: inboard-section at η = 0.16, Right: outboard-section at η =0.56)
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5.5
Spanwise distributions of local lift and drag coefficients with changing the spacer height at freestream velocity
of 60m/s are shown in Fig.8. They are evaluated by integrating static pressure distributions at seven cross sections
on the wing. Geometrical angles of attack of the model are 4deg, 10deg and 13deg. As shown in Fig.8, local lift
increases continuously from the inboard area to the outboard area when the spacer becomes higher. The magnitude
of the increment with spacer height seems to depend on the magnitude of lift coefficient at each of the cross sections.
From this aspect, the influence of the spacer height seems to be explained based on the lifting line theory rather than
fluid dynamic interaction. On the other hand, distributions of local drag coefficients shows the different
characteristics. As shown in right row of Fig.8, the local drag increases mainly in the inboard area. This result is
suggestive of the fluid dynamic interaction near this area, and discussed in detail in the next section.
3.5
0.8
3
0.6
0.4
2
cd
cl
2.5
0.2
1.5
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
1
0.5
0
0.2
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
0
-0.2
0.4
0.6
0.8
0
1
0.2
0.4
0.6
0.8
1
0.6
0.8
1
0.6
0.8
1
η
η
(a) AoA = 4deg
3.5
0.8
3
0.6
0.4
2
cd
cl
2.5
0.2
1.5
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
1
0.5
0
0.2
-0.2
0.4
0.6
0.8
η
1
0
0.2
0.4
η
(b) AoA = 10deg
3.5
0.8
3
0.6
2.5
0.4
2
cd
cl
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
0
0.2
1.5
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
1
0.5
0
0.2
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
0
-0.2
0.4
0.6
0.8
1
0
0.2
η
0.4
η
(c) AoA = 13deg
Figure 8. Local lift and drag distribution along spanwise direction.
(Left: Lift coefficient, Right: Drag coefficient)
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In order to roughly estimate a variation in the effective angle of attack, Figure 9 compares a static pressure
coefficient at two parts of the wing cross sections for AoA=10deg. As the typical case, spacer heights of 150mm and
30mm are selected, which corresponds to 99%-thickness of the wall boundary layer and the latter is nearly equal to
displacement thickness of the wall boundary layer, respectively. It is estimated that the effective angle of attack for
150mm spacer case is 1.5degree and 0.5degree higher than that of 30mm spacer case at the in- and the outboard
sections, respectively. This result almost agrees with computational results17,20.
AoA=8deg, 30mm Spacer (WTT)
AoA=9deg, 30mm Spacer (WTT)
AoA=10deg, 30mm Spacer (WTT)
AoA=11deg, 30mm Spacer (WTT)
AoA=12deg, 30mm Spacer (WTT)
AoA=10deg, 150mm Spacer (WTT)
-5
-4
-2
-4
-3
Cp
Cp
-3
AoA=8deg, 30mm Spacer (WTT)
AoA=9deg, 30mm Spacer (WTT)
AoA=10deg, 30mm Spacer (WTT)
AoA=11deg, 30mm Spacer (WTT)
AoA=12deg, 30mm Spacer (WTT)
AoA=10deg, 150mm Spacer (WTT)
-5
-2
-1
-1
0
0
1
1
3.4
3.9
4.4
X/C
4.9
4.5
5.4
5
X/C
5.5
Figure 9. Variation of the effective angle of attack at AoA=10deg.
(left: η = 0.16, right: η = 0.56)
B. Local aerodynamics and flow field
In this section, further details of lift and drag performance and a surface flow pattern in the different spacer
conditions are discussed.
1. Effect of the aspect ratio
As mentioned in the previous section, aerodynamic coefficients are strongly affected by the installation of the
spacer and it is possibly due to two reasons. They are based on the lifting line theory or fluid dynamic interaction.
Figure 10 shows the variation of the induced drag evaluated by the relation between CD and square of CL. Two
remarkable features are found in the figure. The first one is that drag coefficient at zero-lift angle of attack which is
estimated by the extrapolation is observed almost the same among each of the spacers. And the other is that the
slope of the curve becomes lower when the spacer height increases. Therefore, it can be said that the effect of the
spacer installation is mainly caused by a change in effective aspect ratio.
0.6
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
0.5
CD
0.4
0.3
0.2
0.1
0
1
2
3
4
5
6
7
8
CL2
Figure 10. Variation of the induced drag.
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9
Figure 11 shows relation between the effective aspect ratio and height of the spacer. The effective aspect ratio is
briefly estimated as the slope of CD and square of CL shown in Fig.10. As shown in the figure, it is found that the
effective aspect ratio linearly increases with the height of the spacer. In other word, if effective aspect ratio for a
full-span configuration can be predicted with a CFD or something, an optimum height of the spacer is possibly
decided.
As a result of these considerations, it is revealed that an effect of the spacer installation mainly appears as the
increment of the effective aspect ratio and resultant changes in slope of the lift curve and reduction of the induced
drag.
Effective aspect ratio : Ae
10.5
10
9.5
9
8.5
0
30
60
90
120
Height[mm]
150
180
Figure 11. Effective aspect ratio versus height of the spacer.
2. Surface flow pattern
In the previous section, the effect of the spacer installation appears as a result of the change in effective aspect
ratio, and then the change in the profile drag component produced by flow interaction is relatively small because the
induced drag component is dominant in highlift configuration. However, change in profile drag in case the spacer is
used is shown numerically in Murayama et al.17,20, where variation of the drag is attributed to large positive and
negative velocity changes in the plane to assume the symmetric condition by the spacer installation. In the
experiment also, change in local drag is visible in Fig.6 even though it is small compared to induced drag. The
reason for this is considered.
At first, Figure 12 and Figure 13 shows the surface flow patterns around a reading edge of a wing-fuselage
junction region and an aft-fuselage regions for 150mm and 30mm spacer cases at AoA=10deg in order to roughly
observe interaction of the boundary layer flow on the fuselage and the inboard area. In the computational results17
horseshoe vortex formed by the boundary layer on a wind tunnel wall interacts with the fuselage and then changes
the flow field. However in the experiment, it is visible that a stagnation point at the reading edge of a wing-fuselage
junction and a surface flow pattern on the aft-fuselage region seems basically agree between two heights of the
spacer cases. In addition to that, Figure 14 compares the surface flow patterns around the inboard area on the wing
for 150mm and 30mm cases at AoA=10deg. In this pictures also, a difference in a global behavior of the surface
flow pattern on wing and fuselage is not observed. Therefore, it can be said that the local surface flow field around
the fuselage and the inboard area of the wing is not largely affected by the variation of the spacer height.
Secondly, Figure 15 shows static pressure distributions of the fuselage cross sections, which are a fore-fuselage
L-L section and aft-fuselage O-O section shown in Fig.2, when the spacer heights are changed. Results for
AoA=10deg at freestream velocity of 60m/s are shown. Horizontal axis represents the location from fuselage center,
which is normalized by fuselage diameter. It is show that the pressure distribution is largely change at the forefuselage section when the spacer height changes. On the contrary to that, it is not remarkable at the aft-fuselage
section. It can be clearly said that it is due to the interaction of the flow although it is not appear in the surface flow
pattern in Figure 12 and Figure 13. Then, the increment of negative pressure in the fore-fuselage reduces drag in the
inboard area as shown in Fig.6. Furthermore, relation between an increment of the local angle of attack caused by
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American Institute of Aeronautics and Astronautics
the variation of a local lift distribution and a change in a local flow field is possibly considered as a reason for the
change in drag at inboard area. This aspect will be assessed in the work.
This kind of the effect, which is caused by nonlinear change in flow field and is impossible to be explained by
lifting line theory, is difficult to correct. Therefore, it is required to use optimum height of the spacer and estimate
the variation caused by change in flow field. This is more important in the case of lower lift configuration. Figure 16
compares a drag polar for the clean configuration (the slat and the flap are stowed and the nacelle is removed) with
the spacers with 150mm and 60mm heights. In this case, difference of the zero-lift drag, namely the profile drag,
which is due to the spacer, is clearly visible.
Figure 12. Surface flow pattern at wing root region.
(AoA=10deg, upper:150mm spacer, lower:30mm spacer)
Figure 13. Surface flow pattern at aft-fuselage region.
(AoA=10deg, upper:150mm spacer, lower:30mm spacer)
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American Institute of Aeronautics and Astronautics
Figure 14. Surface flow pattern around inboard region on wing.
(AoA=10deg, left:150mm spacer, right:30mm spacer)
-1
-1
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
-0.5
150mm Spacer (WTT)
80mm Spacer (WTT)
30mm Spacer (WTT)
Without Spacer (WTT)
Cp
Cp
-0.5
0
-0.4
-0.2
0
0.2
0.4
0.6
0.5
-0.6
-0.4
-0.2
Z/D
0
0.2
Z/D
Figure 15. Static pressure distributions on the fuselage.
(AoA=10deg, left: fore-fuselage L-L section, right: aft-fuselage O-O section)
2
150mm Spacer (WTT, Cruise config., w/o nacelle)
60mm Spacer (WTT, Cruise config., w/o nacelle)
1.5
1
CL
0.5
-0.6
0
0.5
0
-0.5
-1
0
0.02
0.04
0.06
0.08
0.1
0.12
CD
Figure 16. Effect of the spacer height in clean configuration.
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American Institute of Aeronautics and Astronautics
0.4
0.6
CL [AoA=10.5deg]
C. Estimation of the desirable spacer height using CFD result
This section presents an estimation of a desirable height of the spacer for the half-span model experiment by
comparing the computational result for a full-span model in free-flight condition and a half-span model installed on
non-slip wall condition.
Figure 17 shows lift coefficient versus height of the spacer obtained by two kinds of computations and a wind
tunnel experiment at AoA=10.5deg. At this condition, the lift variation in the experimental and the computation with
the half-span model installed on non-slip wall condition almost agrees, and 75mm spacer seems in good agreement
with the full-span free-flight computation. A lift range which is required as the allowable error for an aerodynamic
measurement in a wind tunnel experiment of an aircraft development27 is superimposed on the graph. Considering
that range, 50mm ~ 110mm can be predicted as the desirable height. On the other hand, from drag point of view,
Figure 18 shows drag coefficient versus the height of the spacer at same lift condition CL=2.43deg. Comparison
between two kinds of computational result shows that the half-span model with 65mm spacer agrees with the fullspan model although computational result for the half-span model has difference from experiment result.
Considering the range of the allowable error described above, 50mm ~ 80mm can be predicted as the desirable
height. As described in the previous section, although absolute values of the aerodynamic coefficient when the
spacer height is changed do not basically agree between experimental results and computational results, the
magnitude of variation with spacer height agree well in Fig.17 and Fig.18.
WTT
CFD_Half-span(on non-slip wall)
CFD_Full-span(FreeFlight)
2.63
2.58
2.53
0
30
60
90
120
Height[mm]
150
180
Figure 17. Lift coefficient versus height of the spacer.
(AoA=10.5deg)
WTT
CFD_Half-span(on non-slip wall)
CFD_Full-span(FreeFlight)
CD [CL=2.43]
0.32
0.3
0.28
0.26
0
30
60
90
120
Height[mm]
150
180
Figure 18. Drag coefficient versus height of the spacer.
(CL=2.43)
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American Institute of Aeronautics and Astronautics
Next, a prediction of a desirable height from an aspect ratio point of view is performed. This prediction allows
overall estimation compared to the way using lift or drag alone, which is discussed above. As shown in Figure 19,
the effective aspect ratio linearly increases with the height of the spacer same as the experimental results. The reason
why the result of 0mm height is not consistent with the line is mainly due to the difference of grid generation near
wind tunnel floor. The result shows that the desirable height is predicted to be 50mm. The curve in the graph may
shifts slightly parallel to the original line because the experiment and computation has some difference. It is a
difference in the nacelle configuration and accuracy of the computational flow field on a flap. As a result, it is
suggested that the spacer height approximately 2 ~ 3 times of the displacement thickness of the floor boundary layer
is a candidate for an optimum configuration for a half-span model experiment. This is close to the result proposed by
William et al.22. However, they didn’t apply the spacer which has the height of 3 ~ 4 times of the displacement
thickness of the floor boundary layer.
WTT
CFD_Half-span(on non-slip wall)
CFD_Full-span(FreeFlight)
Effective aspect ratio : Ae
11
10.5
10
9.5
9
8.5
8
7.5
0
30
60
90
120
Height[mm]
150
180
Figure 19. Effective aspect ratio versus height of the spacer.
VI.
Concluding Remarks
In the research program to make aerodynamic design methodology for an efficient high-lift system in JAXA,
series of lowspeed wind tunnel experiment using a realistic aircraft configuration model with a leading-edge slat,
flaps, a fuselage, a nacelle-pylon, slat tracks and Flap Track Fairings (FTF) has been carried out. Throughout the
wind tunnel experiment champagne in October 2005 - February 2009, an influence of a half model installation was
assessed using four heights of a spacer between bottom of the fuselage and the wind tunnel wall. It is aimed at
reducing the differences between experiment and Reynolds averaged Navier-Stokes computation as for lift and drag
coefficient, and the slope of the lift curve.
Initially, variation in lift, drag, pitching-moment coefficient and static pressure distribution when the spacer
height is changed is observed. Lift and pitching-moment coefficient increases as the spacer becomes higher, but drag
decreases. As for lift curve with angle of attack, the slope also increases with the height of the spacer, which seems
to approximate with a polynomial function. Difference of the aerodynamic coefficients in case the spacer height
from 0mm(without spacer) to 150mm are approximately as follows; ∆CL = 0.08, ∆CD = 0.013, ∆CM = 0.03.
Variation of the aerodynamic characteristics with spacer height in the experiment agrees well with computational
result which is assumed a wind tunnel wall boundary layer. And all of experimental results become closer to the
result of free-flight computation when the spacer height becomes lower.
Pressure on the wing is also changed by the existence of the spacer. The magnitude of the variation is little in
lower angle of attack case, but the pressure on the slat and the leading-edge of the main-wing is deviated depending
on spacer height. Using the pressure result, local lift and drag distribution is obtained. Lift increases continuously
along spanwise direction when the spacer height increases, which shows that the influence of the spacer height
seems to be explained based on the lifting line theory rather than fluid dynamic interaction. But the local drag
increases mainly in the inboard area. This is due to the increment of static pressure on thefuselage cross section in
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American Institute of Aeronautics and Astronautics
front of the wing. It is estimated that the effective angle of attack for 30mm spacer case is 1.5degree and 0.5degree
higher than that of 150mm spacer case at the in- and the outboard sections, respectively.
The variation of the induced drag evaluated by the relation between CD and square of CL shows that an effect of
the spacer installation mainly appears as the increment of the effective aspect ratio and resultant changes in slope of
the lift curve and reduction of the induced drag. And a difference in global behavior of the surface flow pattern on
wing and fuselage is not observed. However, an influence of the flow interaction near wind tunnel wall is not
negligible, because change in the profile drag in the inboard area is observed as descried above, although the profile
drag component is relatively small compared to the induced drag in the case of a highlift configuration.
Finally, an estimation of an desirable height of the spacer for the half-span model experiment by comparing the
computational result for a full-span model in free-flight condition and the half-span model installed on non-slip wall
condition is performed. From lift and drag point of view at certain angle of attack condition, a desirable height of the
spacer is 50mm ~ 110mm, which is 1/6 ~ 1/3 of the fuselage radius. In the case using the effective aspect ratio, the
spacer height approximately 2 ~ 3 times of the displacement thickness of the floor boundary layer is a candidate for
optimum configuration for half-span model experiment.
In the future work, an influence of the flow interaction around the inboard area is assessed in detail using
computational results. Then, the technique to eliminate it or to correct aerodynamic data will be developed.
Acknowledgements
We would like to express our gratitude and appreciation to member of Lowspeed Wind Tunnel Section in Wind
Tunnel Technology Center, Japan Aerospace Exploration Agency, for their generous support and large amount of
efforts throughout the experiment.
Great thanks must also be addressed to Dr. Amemiya, Mr. Hirai and member of Civil Transport Team in
Aviation Program Group and Advance Test Section in Wind Tunnel Technology Center, Japan Aerospace
Exploration Agency, for their specialized and technical support for the whole of this research program.
References
1.
Rudolph PKC. “High-lift Systems on Commercial Subsonic Airliner,”. NASA CR 4746, September 1996.
2.
Meredith, P. T., “Viscous Phenomena Affecting High-Lift Systems and Suggestions for Future CFD Development,” HighLift Systems Aerodynamics, AGARD CP 315, Sep. 1993, pp. 19-1 – 19-8.
3.
F. T. Johnson et al., “Thirty Years of Development and Application of CFD at Boeing Commercial Airplanes, Seattle,”
Computers & Fluids 34, 2005.
4.
C. L. Rumsey and S. X. Ying, “Prediction of high lift: review of present CFD capability,” Progress in Aerospace Science,
Vol.38 145-180, 2002.
5.
http://db-www.larc.nasa.gov/trapwing/archive/register/ [cited 30 December 2005]
6.
F., M., Payne, “High Reynolds Number Studies of Boeing 777-200 High Lift Configuration in the NASA ARC 12’ Pressure
Tunnel and NASA LaRC National Transonic Facility,” AIAA Paper 2000-4220, 2000.
7.
Hansen, H., Thiede, P., Moens, F., Rudnik, R., and Quest, J., “Overview about the European High Lift Research Programme
EUROLIFT,” AIAA Paper 2004-0767, 2004.
8.
Rudnik, R., Geyr, H., Frhr, v., “The European High Lift Project EUROLIFT II – Objectives, Approach, and Structure,”
AIAA Paper 2007-4296, 2007.
9.
Ito, T., Yokokawa, Y., Ura, H., Kato, H., Mitsuo, K., and Yamamoto, K., “High-Lift Device Testing in JAXA 6.5M X 5.5M
Low-Speed Wind Tunnel,” AIAA Paper 2006-3643, 2006.
10.
Yokokawa, Y., Murayama, M., Ito T., and Yamamoto, K., “Experiment and CFD of a High-Lift Configuration Civil
Transport Aircraft Model,” AIAA Paper 2006-3452, 2006.
11.
Murayama, M., Yamamoto, K., and Kobayashi, K., “Validation of Flows on High-Lift Configurations by Structured- and
Unstructured- Mesh Method,” AIAA Paper 2005-1226, 2005.
12.
Murayama, M., Yamamoto, K., and Kobayashi, K., “Validation of Computations Around High-Lift Configurations by
Structured- and Unstructured- Mesh,” Journal of Aircraft, Vol. 43, No. 2, March-April 2006, pp. 395-406.
13.
Murayama, M., Imamura, T., Yamamoto, K., and Kobayashi, K., “Comparison of RANS Simulations of Multi-Element
High-Lift Configurations,” AIAA Paper 2006-1396, 2006.
14.
Murayama, M., Yamamoto, K., Tanaka, K., Yokokawa, Y., Ito, T., “Numerical Simulation around a High-Lift Configuration
of a Civil Aircraft Model,” 5th AWCFD. 2006-3643, 2006.
15.
Murayama, M., Yokokawa, Y., Yamamoto, K., “Validation study of CFD Analysis for High-Lift Systems,” 25th ICAS,
Hamburg, Germany, ICAS-Paper2006-2.3.1, 2006.
16.
Murayama, M., Yokokawa, Y., Yamamoto, K., Ueda, Y., “CFD Validation Study for a High-Lift Configuration of a Civil
Aircraft Model,” AIAA Paper2007-3924, 2007.
16
American Institute of Aeronautics and Astronautics
17.
Murayama, M., Yokokawa, Y., Tanaka, K., Yamamoto, K., Ito, T., “Numerical Simulation of Half-Span Aircraft Model with
High-Lift Devices in Wind Tunnel,” AIAA Paper2008-333, 2008.
18.
Yokokawa, Y., Murayama, M., Kanazaki, M., Murota, K., Ito T., and Yamamoto, K., “Investigation and Improvement of
High-lift Aerodynamic Performances in Lowspeed Wind Tunnel Testing,” 48th Aerospace Sciences Meeting and Exhibit, AIAA
Paper2008-350, 2008.
19.
Yokokawa, Y., Kanazaki, M., Murayama, M., Kato, H., Ito T., and Yamamoto, K., “Investigation of the Flow over
Nacelle/Pylon and Wing Controlled with a Vortex Generation in High-lift Configuration,” ICAS-Paper2008-3.3.3, 2008.
20.
Murayama, M., Yokokawa, Y., Kato, H., Kanazaki, M., Yamamoto, K., and Ito T., “Computational Study for High-lift
Aerodynamics Research in JAXA,” ICAS-Paper2008-2.2.1, 2008.
21.
Kanazaki, M., Jeong, S., and Murayama, M., “High-Lift System Optimization Based on Kriging Model Using High Fidelity
Flow Solver,” Transactions of the Japan Society for Aeronautical and Space Science, Vol. 49, No. 165, 2006.
22.
William, E., Milholen II, Chokani, N., McGhee, Robert, J., “Development of Semi-Span Model Test Techniques,” AIAA
Paper 96-02412, 1996.
23.
Gatlin, G., M., Parker, P., A., and Owens, Jr., L., R., “Development of a SmiI-Span Test Capability at the National Transonic
Facility (Invited),” AIAA Papre2001-0759, 2001.
24.
Milholen, II, William, E., “A Design Methodology for Semi-Span Model Mounting Geometries,” AIAA Paper98-0758, 1998.
25.
Eliasson, P., “Numerical Validation of a Half Model High Lift Configuration in a Wind Tunnel,” AIAA Paper 2007-0262,
2007.
26.
Melber-Wilkending, S. and Wichmann, G., “Application of Advanced CFD Tools for High Reynolds Number Testing,”
AIAA Paper2009-418, 2009.
27.
Payne, F., M., “Low Speed Wind Tunnel Testing Facility Requirements –A Customer’s Perspective,” AIAA Paper99-0306,
1999.
28.
Jewel, B., Barlow, William, H., Rea, Jr., Alan, Pope, “Low-Speed Wind Tunnel Testing,” Third Edition.
29.
Imamura, T., Enomoto, S., Yamamoto, K., “3D Unsteady Flow Computations in a Slat Cove Using Large Eddy Simulation,”
AIAA Paper 2006-2668, 2006.
30.
Ura, H., Yokokawa, Y., Ito, T., “Phased Array Measurement of High Lift Devices in Low Speed Wind Tunnel,” AIAA Paper
2006-2565, 2006.
31.
Kato, H., Watanabe, S., Hashimoto, T., Fujii, K., Ito T., and Yamamoto, K., “PIV Measurement of a High-Lift-Device Model
in JAXA 6.5m × 5.5m Low-Speed Wind Tunnel,” AIAA Paper2007-1064, 2007.
32.
Kato, H., Watanabe, S., Murayama, M., Yokokawa Y., and Ito, T., “PIV Investigation of Nacelle Chine Effects on High- Lift
System Performance,” AIAA Paper2008-240, 2008.
33.
Mitsuo, K., Kurita, M., Kuchi-Ishi, S., Fujii, K., Ito, T., and Watanabe, S., “PSP Measurement of a High-Lift-Device Model
in JAXA 6.5m×5.5m Low-Speed Wind Tunnel,” AIAA Paper2007-1065, 2007.
34.
Nakahashi, K., Togashi, F., Fujita, T., and Ito, Y., “Numerical Simulations on Separation of Scaled Supersonic Experimental
Airplane from Rocket Booster at Supersonic Speed,” AIAA Paper 2002-2843, 2002.
35.
Murayama, M., Togashi, F., Nakahashi, K., Matsushima, K., and Kato, T. “Simulation of Aircraft Response to Control
Surface Deflection Using Unstructured Dynamic Grids,” Journal of Aircraft, Vol. 42, No. 2, March-April 2005, pp. 340-346.
36.
Ito, Y. and Nakahashi, K. “Direct Surface Triangulation Using Stereolithography Data,” AIAA Journal, Vol. 40, No. 3, 2002,
pp. 490-496.
37.
Ito, Y. and Nakahashi, K., “Surface Triangulation for Polygonal Models Based on CAD Data,” International Journal for
Numerical Methods in Fluids, Vol. 39, Issue 1, 2002, pp. 75-96.
38.
Sharov, D. and Nakahashi, K., “A Boundary Recovery Algorithm for Delaunay Tetrahedral Meshing,” Proceedings of 5th
International Conference on Numerical Grid Generation in Computational Field Simulations, Mississippi State, Mississippi, 1996,
pp. 229-238.
39.
Ito, Y. and Nakahashi, K., “Improvements in the Reliability and Quality of Unstructured Hybrid Mesh Generation,”
International Journal for Numerical Methods in Fluids, Vol. 45, Issue 1, May 2004, pp. 79-108.
40.
Obayashi, S. and Guruswamy, G. P., “Convergence Acceleration of an Aeroelastic Navier-Stokes Solver,” AIAA Journal,
Vol. 33, No. 6, 1995, pp. 1134-1141.
41.
Sharov, D. and Nakahashi, K., “Reordering of Hybrid Unstructured Grids for Lower-Upper Symmetric Gauss-Seidel
Computations,” AIAA Journal, Vol. 36, No. 3, 1998, pp. 484-486.
42.
Spalart, P. R. and Allmaras, S. R., “A One-Equation Turbulence Model for Aerodynamic Flows,” AIAA Paper 92-0439, Jan.
1992.
43.
Lei, Z., “Effect of RANS Turbulence Models on Computational of Separated Flows over a Wing-Body Configuration,”
Proceedings of WCCM VI in conjuction with APCOM”04, 2004.
44.
Matsuo, Y., Nakamura, T., Tsuchiya, M., Ishizuka, T., Fujita, N., Ohkawa, H., Hirabayashi, Y., Takaki, R., Yoshida, M.,
Nakamura, K., Yamamoto, K., Suematsu, K., and Iwamiya, T., “Numerical Simulator III – Building a Terascale Distributed
Parallel Computing Environment for Aerospace Science and Engineering,” Proceedings of the Parallel CFD 2002 Conference,
Nara, Japan, Elsevier Science B. V., 2003, pp. 187-194.
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American Institute of Aeronautics and Astronautics