the aerodynamic analysis of bwb baseline ii e5-6 uav with canard aspect ratio (ar) of 6 at angle of...
DESCRIPTION
FINAL YEAR PROJECT THESIS FOR BEng MECHANICALTRANSCRIPT
MUHAMMAD HILMI BIN ABD ADZIS 2009405194 JULY2012
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TABLE OF CONTENTS
CONTENTS PAGE
TABLE OF CONTENTS i
ACKNOWLEDGEMENT iv
ABSTRACT v
LIST OF TABLES vi
LIST OF FIGURES vii
LIST OF ABBREVIATIONS xi
CHAPTER 1 INTRODUCTION
1.0 Project Background 1
1.1 Problem statement 6
1.2 Objective 6
1.3 Scope Of Project 7
1.4 Significance Of Project 7
1.5 Project Methodology 8
CHAPTER 2 LITERATURE REVIEW
2.1 Definition 11
2.1.1 Blended Wing Body (BWB) 11
2.1.2 Unmanned Aerial Vehicle (UAV) 11
2.1.3 Canard 12
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2.1.4 Computational Fluid Dynamics 12
(CFD)
2.2 Research On BWB UAV 13
2.3 Aerodynamic Overview 15
2.3.1 Aerodynamic force on aircraft 15
2.4 Turbulence Model Overview 19
2.4.1 Spalart-Allmaras turbulence model 19
CHAPTER 3 PARAMETER VALIDATION
3.1. Parameter Validation 21
3.1.1 Geometry Selection 21
3.1.2 CFD Software Setup 22
3.1.3 Parameter Validation Procedure 24
3.1.4 The Results 36
3.1.5 Conclusion 38
CHAPTER 4 GRID INDEPENDENCE STUDY
4.1 Introduction to Grid Independence Study 40
4.1.1 What is grid independence study? 40
4.1.2 How to achieve grid independence 40
4.2 Grid Independence Study Process 41
4.2.1 Test with Various Face Settings 41
4.2.2 Test the boundary box size with
various sizes of box 42
4.2.3 Test with various refinement
numbers 47
4.3 Summary of the Parameters Obtained
From GridIndependence Study 49
4.4 Test the parameters obtained in Grid
IndependenceStudy 50
4.5 Conclusion 52
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CHAPTER 5 RESULTS AND DISCUSSIONS
5.1 Aerodynamics Analysis for BWB
Baseline II E5-6 54
5.1.1 Lift Coefficient, CLAnalysis 55
5.1.2 Drag Coefficient, CDAnalysis 56
5.1.3 Pitching Moment Coefficient,
CMAnalysis 57
5.1.4 Lift-to-drag ratio analysis 59
5.1.5 Static Pressure Contour Analysis 60
5.1.6 Mach Number Contour Analysis 63
5.1.7 Velocity Vector Analysis 65
CHAPTER 6 CONCLUSION AND RECOMMENDATIONS
6.1 Conclusion 68
6.2 Recommendations 69
REFERENCES
References 70
APPENDICES
APPENDIX A Static pressure contour for every canard setting 74
angle, δ viewed from the top and bottom surface of
the BWB Baseline II E5-6
APPENDIX B Mach number contour for every canard setting angle, 76
δ viewed from the top and bottom surface of the BWB
Baseline II E5-6
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ACKNOWLEDGEMENT
In the name of Allah, the Most Gracious, Most Merciful, the author would like to
express his greatest acknowledgement for giving all of strengths during the
completion this thesis. First and foremost, i would like to thank to my beloved
parents Abd Adzis and Mek Eshah and not forgotten to Siti Salwa for their support
through their greatest motivation and financial. Thanks to Professor Dr. Wirachman
Wisnoe as the project supervisor and Pn. Zurriati Ali as co-supervisor, on behalf of
their guidance towards the finer aspects of finishing this project paper. They also
showed great patience in helping to understand the aerodynamics behavior of the
aircraft and provide great explanation needed. Despite of that, the author would like
to thanks to the entire team member for their help and support throughout discussion
session among them. Finally, the author would like to express gratitude to all parties
such as my best friends who had contributed directly and indirectly in completion
this project especially all friends for their help and support. Thank you very much,
may God repays for their kindness.
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ABSTRACT
This thesis presents a study of the effects of canard on the aerodynamic of a Blended
Wing Body (BWB) Baseline II E5 Unmanned Aerial Vehicle (UAV).The objective
of this project is to obtain the aerodynamic characteristics such as lift coefficient
(CL), drag coefficient (CD) and pitching moment coefficient (CM) of BWB Baseline II
E5-6 UAV with canard and also to obtain the aerodynamic visualizations such as
pressure contour and Mach number contour. In this project, a pair of canard with
aspect ratio of 6 is fitted on the BWB Baseline II E5. This thesis starts with literature
overview, then followed with parameter validation, grid independence study and
final simulation. Forces and moments are measured and value of CL, CD and CM are
obtained and compared at 0.1 Mach number with respect to variation of canard
setting angle, δ at 0o angle of attack, α. Pressure contours and Mach number
contours, both extracted from CFD analysis, are plotted. An extensive discussion of
the results, conclusion and recommendation is presented at the end of the chapter.
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LIST OF TABLES
TABLE TITLE PAGE
Table 3.1 Atmosphere parameter for the experimental condition 22
Table 3.2 selected force vector values of X and Y component 23
Table 3.3 Comparison between CL and CD 36
Table 4.1 Face setting parameters 41
Table 4.2 seven different box sizes with initial mesh values 43
Table 4.3 parameters for CFD simulation for box test 44
Table 4.4 Results and cell numbers from box test 45
Table 4.4 Results obtained from the refinement number test 47
Table 4.5 summary of the parameters obtained from grid 49
independence study
Table 5.1 results obtained from the simulation 54
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LIST OF FIGURES
FIGURE TITLE PAGE
Figure 1.1 three axes of rotation for aircraft motion 2
Figure 1.2 tail assembly (left) and canard (right) 3
Figure 1.3 BWB Baseline I (left) and BWB Baseline II (right) 4
Figure 1.4 The design of BWB Baseline II E5 5
Figure 1.5 Methodology flowchart 9
Figure 2.1 Force vector on aerofoil 15
Figure 3.1 Half model of BWB Baseline-II E4 21
Figure 3.2 The BWB half model imported into the HEXPRESS 24
Figure 3.3 Boundary box setup 25
Figure 3.4 BWB domain setup (left) and the domain result (right) 26
Figure 3.5 Boundary condition for the BWB 26
Figure 3.6 The initial mesh setup for the BWB 27
Figure 3.7 Mesh adaptation process. Refinement number 28
(upper right) and trimming parameters (bottom)
Figure 3.8 Mesh adaptation result 28
Figure 3.9 Snap to geometry process 29
Figure 3.10 Snap to geometry result 29
Figure 3.11 Mesh optimization process command 30
Figure 3.12 Applying a viscous layer on the mesh process 30
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Figure 3.13 Successful mesh with viscous layer surrounded 30
the BWB
Figure 3.14 Fluid properties selection 31
Figure 3.15 Fluid model properties 32
Figure 3.16 Boundary condition setup 32
Figure 3.17 Initial solution properties setup 33
Figure 3.18 Numerical model properties setup 33
Figure 3.19 Output parameter setup for vector component 34
Figure 3.20 Control variable to set the iteration number 34
and convergence criteria
Figure 3.21 CFD simulation process running 35
Figure 3.22 Graph of CL versus angle of attack 37
Figure 3.23 graph of CD versus angle of attack 37
Figure 4.1 Graph of CL versus number of cell (left) and 46
CD versus number of cell (right)
Figure 4.2 Graph of CL versus refinement numbers (left) 48
and CD versus refinement number (right)
Figure 4.3 graph of CL versus angle of attack obtained 50
from the simulation and from the wind
tunnel experiment
Figure 4.4 Graph of CD versus angle of attack obtained 51
from the simulation and from the wind
tunnel experiment
Figure 4.5 Graph of CM versus angle of attack obtained 51
from the simulation and from the wind
tunnel experiment
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Figure 5.1 graph of CL versus δ (left) and linear region at 55
low canard angle (right)
Figure 5.2 graph of CD versus δ 56
Figure 5.3 graph of CM versus δ (left) and linear region 58
at low canard angle (right)
Figure 5.4 graph of L/D versus δ 59
Figure 5.5 Static pressure contour of 0o canard angle 60
viewed from cutting plane of BWB (left) and
top of BWB (right)
Figure 5.6 Static pressure contour of 8o canard angle 61
viewed from cutting plane of BWB (left) and
top of BWB (right)
Figure 5.7 Static pressure contour of 10o canard angle 61
viewed from cutting plane of BWB (left) and
top of BWB (right)
Figure 5.8 Static pressure contour of 11o canard angle 61
viewed from cutting plane of BWB (left) and
top of BWB (right)
Figure 5.9 Mach number contour of 0o canard angle viewed 63
from bottom surface of BWB (left) and top surface
of BWB (right)
Figure 5.10 Mach number contour of 8o canard angle viewed 64
from bottom surface of BWB (left) and top surface
of BWB (right)
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Figure 5.11 Mach number contour of 10o canard angle viewed 64
from bottom surface of BWB (left) and top surface
of BWB (right)
Figure 5.12 Mach number contour of 11o canard angle viewed 64
from bottom surface of BWB (left) and top surface
of BWB (right)
Figure 5.13 Velocity vector when canard positioned at 11o 66
Figure 5.14 Detail on the velocity vector when canard positioned 66
at 11o (left) and the beginning point of reverse flow
zoomed from the left picture (right)
Figure 5.15 velocity profiles in the boundary layer. 67
The flow start to separate at the 3rd
picture (right)
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LIST OF ABBREVIATIONS
α Angles of attack (AOA)
AR Aspect Ratio
BWB Blended Wing Body
b airfoil span length
c airfoil chord length
CFD Computational Fluid Dynamics
CL Lift coefficient
CLmax maximum lift coefficient
CD Drag coefficient
CM Pitching moment coefficient
D Drag force
δ Canard setting angle
L lift force
L/D Lift to Drag Ratio
M Pitching moment force
Ma Mach number
ρ Density
Re Reynolds number
Sref Reference area
UAV Unmanned Aerial Vehicle
V velocity
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CHAPTER 1
INTRODUCTION
1.0 Project Background
An aircraft motion divided into three primary ways, which are pitch, roll
and yaw. All of these motions rotate about its centre of gravity, in its specific axes.
The longitudinal axis is the axis that extends lengthwise through the fuselage from
the nose to the tail. The lateral axis is the axis that extends crosswise from wingtip to
wingtip. While the vertical axis passes vertically through the centre of gravity. Pitch
is the movement of the aircraft nose up or down in lateral axis, roll is the rotation
around the longitudinal axis, and yaw is movement of the nose to left or right in the
vertical axis
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Figure 1.1 Three axes of rotation for aircraft motion
In conventional aircraft, it has a tail assembly known as empennage, which
included an elevator, horizontal and vertical stabilizer. The horizontal stabilizer and
elevator assembled together in one airfoil. The fixed section at the front is the
horizontal stabilizer while the rear movable section is the elevator. Changing the
angle of deflection at the elevator changes the amount of lift generated by the main
airfoil and cause the pitching motion occurred. When this horizontal stabilizer placed
at the front, it is known as a canard.
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Figure 1.2 Tail assembly (left) and canard (right)
Canard is a horizontal surface mounted ahead of the main wing to give
longitudinal stability and control. It may be a fixed, movable, or variable geometry
surface, with or without control surfaces. There are three major types of canard,
which are control canard, lifting canard and close coupled canard. For longitudinal
control during manoeuvring, control canard primarily is used and it carrying no
aircraft weight in normal flight.
For tailless aircraft, canard is mounted at the in front of the main wing to
achieve the same function as the horizontal stabilizer and elevator at the empennage.
This included the tailless aircraft such as Blended Wing Body (BWB). The BWB
concept was introduced by Robert Liebeck [8] at the McDonnel Douglas Corporation
(now known as Boeing Company) in 1988. The BWB concept is a blends of
fuselage, wing, and the engines into a single lifting surface, which allowing the
maximization of the aerodynamics efficiency.
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In year 2005, researchers from the Faculty Of Mechanical Engineering,
Universiti Teknologi MARA (UiTM) has started conducting research on BWB-UAV
concept and comes with the first Computational Fluid Dynamics (CFD) tested BWB
called BWB Baseline I [2] [9]. Following to the first research, development of BWB-
UAV continued to the second model with different design under code name BWB
Baseline II [13][16].
Figure 1.3 BWB Baseline I (left) and BWB Baseline II (right)
BWB Baseline II was a completely revised, redesigned and has a simpler
planform with slenderer body than its predecessor but still maintaining the wing
span. Until now, BWB Baseline II has evolved from the E1 version until E4 with
several modifications done. In this study, the evolution of the BWB Baseline II will
continue to the latest E5 version.
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Figure 1.4 The design of BWB Baseline II E5
The E5 version has a rectangular canard at the front and twisted wing. The
type of canard is control canard and the design is according to NACA 2415 airfoil
design. This model will be named as BWB Baseline II E5-6 (where after this will
referred as BWB). The E5 represent the fifth generation or evolution of BWB
Baseline II while the number 6 is the canard aspect ratio (AR). The canard setting
angles is set to12various different angles.
There are two types of aerodynamics analysis that can be done, which are
by using wind tunnel and Computational Fluid Dynamics (CFD) software. For this
study, CFD software will be used to obtain the aerodynamics results of the canard.
Because of the simulation must deal with the turbulent flow, it is known that the
turbulent flow in the air is hard to be calculated. Therefore, it is easier to simulate
with the existing turbulence models. In CFD software, it provides a various
turbulence models such as Spalart-Allmaras, k-Epsilon (k-ɛ) and k-Omega (k-ω)
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turbulence model. For this study, only one types of turbulence models will be used
which are Spalart-Allmaras models.
1.1 Problem Statement
The BWB has a several weakness and one of its weaknesses was a stability
issues when pitching up and down. For BWB, its weight during flight mostly is
carried by the main wing and the additional control-canard airfoil will be used to
control the longitudinal movement and pitching motion. Thus, a control-canard
mostly operates only as a surface controller and is usually at zero angle of attack.
Therefore, a study needs to be carried out to determine the effect of the canard when
the canard is set to specific setting angle, shape, and aspect ratio. In addition, the
aerodynamics of the BWB with canard at the in front of its main wing needs to be
determined to get the overall results.
1.2 Objective
The objective of this study is to obtain the lift (CL), drag (CD) and pitching
moment (CM) coefficient of the BWB with canard through the Computational Fluid
Dynamics (CFD) simulation at different canard setting angle. Together with the
simulation, CFD visualizations such as pressure contour, Mach contour, and velocity
vectors will be obtained.
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1.3 Scope of Project
In this study, a rectangular canard with aspect ratio (AR) of 6 will be fitted
on the BWB with 12 various canardsetting angles, δ. Then the test will be done at 0.1
Mach number through the CFD simulation software at 0 degree angle of attack. The
simulation will run in Spalart Allmaras turbulence model.
Aquired data for aerodynamics characteristics such as coefficient of drag
(CD), coefficient of lift (CL) and moment coefficient (CM) will be analyzed on the
graph of CD, CL and CM versus canard setting angle, δ.
1.4 Significance Of Project
The significance of this project is to know the behaviour of the canard when
it is fitted on the BWB Baseline II E5.In this study, the lift coefficient (CL), pitching
moment coefficient (CM), drag coefficient (CD) and lift to drag ratio (L/D) will be
obtained. Furthermore, results obtained from this project can be used as a reference
and comparison to experimental study of the BWB in the future.
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1.5 Project Methodology
This project divided into three main stages, which are the parameter
validation, grid independent study and CFD simulation of BWB Baseline II with
canard. In parameter validation, an objects selection from the existing aerodynamic
analysis that is obtained from the journals or research papers will be done. The object
then is redraw back by using CAD software and CFD simulation will be done by
following the given parameters. The obtained result then is compared with the
original results to determine whether the result is same or better than the referred
analysis. The mesh and CFD setup used in this section will be used in the second
process, which is grid independence study.
In grid independence study, a test of simulation will be done by using
several mesh setting and CFD pre-process setup that obtained from the first
procedure. In this process, a BWB drawing will be converted into a solid model by
using CAD software and converted into a variation of grid or mesh numbers before
simulate in CFD. The best result of the CFD pre-process and mesh setup in this stage
will be used for the final procedure that is the true simulation.
The last process is to run the CFD simulation of the BWB by applying the
mesh and CFD pre-process setup that obtained from the grid independence study. In
this stage, a 12 different mesh filesfor12 different canard setting angles will be
created. Then all of these file will be simulated by using CFD software at 0o angle of
attack by using Spalart Allmaras turbulence model.
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N
Y
N
Y
Figure1.5 Methodology flowchart
Object selection for
parameter validation
CAD drawing
CFD simulation
Obtain BWB drawing
CFD analysis with variation of grid
number
analyze
start
Compare
Continue?
CFD analysis with Spalart-Allmaras with various canard
setting anglesat 0oangle of
attack
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CHAPTER 2
LITERATURE REVIEW
Chapter 2 give details the meaning of the main terminologies such as BWB,
UAV and CFD. Other than that, this chapter will also explain briefly about
aerodynamic fundamentals such as lift, drag, and pitching moment. Some research
that has been done either by the UiTM researchers or outsider so far for BWB II
series will be discussed in this chapter. Final part of this chapter will highlight the
mathematical models such as drag, lift, and moment, and also a tubulence models
which is Spallart-Almaras.
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2.1 DEFINITION
2.1.1 Blended Wing Body (BWB)
BWB is an alternative design features from a traditional separated fuselage
and wing aircraft. The BWB concept is a fusions of fuselage, wing, and the engines
into a single lifting surface, which allowing the boosting of the aerodynamics
efficiency. In reality, the BWB didn’t have normal control device such as ailerons
and tail stabilizer, also lacking of ability when dealing with the important pitching
moments. However, this design is able to provide high lifts that can potentially
decrease the fuel utilization. Besides that, BWB also have reduced surfaced area
where almost conventional aircraft has, so at the same time reduced the skin friction
drag.
2.1.2 Unmanned Aerial Vehicle (UAV)
Unmanned Aerial Vehicle, also known as Unmanned Aircraft System or
Remotely Piloted Aircraft refer to the aircraft which functions either by the remote
control of a navigator or pilotthat is as a self-directing entity. It usually can carry
cameras, sensors, communication equipment and other payloads. An UAV has
capable of controlled, sustained level flight and reciprocating engine. UAVs come in
two variations which is some are controlled from a remote location and others fly
autonomously based on pre-programmed flight plans using more complex dynamic
automation systems.
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2.1.3 Canard
Canard is a horizontal surface mounted ahead of the main wing to give
longitudinal stability and control. It may be a fixed, movable, or variable geometry
surface, with or without control surfaces. There are three major types of canard,
which are control canard, lifting canard and close coupled canard. For tailless aircraft
such as BWB, canard is attached at the in front of the main wing to achieve the same
function as the horizontal stabilizer and elevator at the empennage.
2.1.4 Computational Fluid Dynamics (CFD)
CFD is one of the branches of fluids mechanics that uses numerical methods
and algorithms to solve and analyze problems that involved fluid flows. Computers
are used to perform the millions of calculations required to simulate the interaction of
fluids and gases with the complex surfaces used in engineering. The technique is
very powerful and spans a wide range of industrial and non-industrial application
areas. The fundamental basis of almost all CFD problems is the Navier–Stokes
equations or Lattice Boltzmann methods, which define any single-phase fluid flow.
These equations can be simplified by removing terms describing viscosity to yield
the Euler equations. Further simplification, by removing terms describing vorticity
yields the full potential equations. Finally, these equations can be linearized to yield
the linearized potential equations.
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2.2 RESEARCH ON BWB UAV
The idea about the BWB began in 1988 when Robert Liebeck [8] introduced
the first BWB concept which the aircraft concept blends with the fuselage, wing,
engines into a single lifting surface, allowing the aerodynamic efficiency to be
maximized. It was a totally different idea compare to the mainstream conventional
aircraft that has been designed since the first aircraft was invented. According to
Liebeck[8], it is possible to achieve up to a 33% reduction in surface area. This
reduction comes mainly from the elimination of tail surfaces and engine/fuselage
Ning Qin [11] found that, “To achieve the best aerodynamic performance, the
optimal spanwise lift distribution should be a fine balance of the vortex induced drag
due to lift and the wave drag due to the shock wave formation at transonic speeds.
For the integrated BWB shape, the elliptic distribution should no longer be the target
for minimum drag design”. Since the aerofoil profile design can have a significant
effect on shock alleviation, it is therefore essential that the spanwise loading design
is considered along with the aerofoil profile design. The study also reveals that the
pressure drag is playing a much more important role in the total drag for the BWB
design as compared with the conventional designs due the intrinsic nature of the
lower surface to volume ratio for BWB shape. It is therefore more rewarding to
minimise the pressure drag before skin-friction drag reduction techniques, such as
laminar flow control, are considered.
In UiTM, the development of BWB has been started since 2005. The BWB
designed with 4 meter wingspan and 2 meter length and it was classified as mini
UAV. This BWB research has been done either by CFD, finite element method and
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also by wind tunnel testing. Computational studies of BWB Baseline 1 using CFD
shown that it can fly at a very high angle of attack before it start to stall which is the
highest angle of attack it can achieve was 35° with maximum lift coefficient of 1.03
[2][14]. But the studies also revealed that a small deflection of lift curve occured at
the angle of attack at 8o. Wisnoe et al. [18] found out it is occured due to the flow
separation, which starts to occur on the wing part. It is found from the wind tunnel
experiment where a visualization using mini tuft is done. From the wind tunnel
experiments, the maximum lift-to drag ratio obtained was at 6° angle of attack while
from CFD analysis, it was 3°, which is lower than results from wind tunnel
experiment.
Since 2009, Uitm started to design a new variation of BWB named
Baseline-II. It is totally differents than its predecessor where it was equipped with a
pair of canards in front of its main wings. With a simpler planform, broader-chord
wing and slimmer body but still maintain the wing span, it was a completely revised,
redesigned version of Baseline-I BWB. The main objective and target of this new
design is to boost flight performance at low cruising speed by increasing lift-to-drag
ratio through planform and shape redesign and inverse twist method on airfoils
throughout its span [13].
For a study of a canard, Myose [17] study the effect of canards on delta
wing vortex breakdown during dynamic pitching. He found that the furthest static
breakdown location of canard from the main wing increase more aerodynamic
performance in term of pitching moment. Rizal et al. [15][16] studied the effect of
canard on aerodynamics and static stability of Baseline-II at low sub sonic regime.
They found out that canards can add lift more than it adds drag if suitable canard’s
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setting angle is found. A properly-sized canard with suitable setting angle may
improve (L/D) max further. In 2010, there was modification on the wingspan of the
BWB-Baseline-II whereby part of the wing was twisted down at certain angle and
called Baseline-II E2.
2.3 AERODYNAMIC OVERVIEW
2.3.1 Aerodynamic force on aircraft
The concept of the generated aerodynamic force is when a stream air flows
upside or downside, or an aerofoil moves through the air. Point of impact will occur
when the air separates to flow about the aerofoil. At this point, a higher pressure of
area or stagnation point will be formed. Usually the high pressure area is located
around the lowest part of the leading edge, depend on how is the angle of attack. The
overall force produced by the aerofoil are contributed by this high pressure area[6].
Figure 2.1 Force vector on aerofoil
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The deflection force or impact pressure may exert a zero positive force at
very low or zero angles of attack, or even a downward or negative force. Air passing
over the top of the aerofoil produces aerodynamic force in another way. The shape of
the aerofoil causes a low pressure area above the aerofoil according to Bernoulli's
Principle, and the decrease in pressure on top of the aerofoil exerts an upward
aerodynamic force. Pressure differential between the upper and lower surface of the
aerofoil is quite small. Even a small pressure differential produces substantial force
when applied to the large area of an aerofoil.
The resultant force on airfoil or usually called as aerodynamic force is
divided into two components which are lift and drag. The lift is defined as the
component of force in the plane of symmetry in direction perpendicular to the line of
flight [4][3][1]. For steady level flight, the upward lift force has to be balanced by the
aircraft weight. The formula of lift is
Lift, L = CL
Where L = Lift force (N)
ρ = Density (kg/m3)
V = Velocity (m/s)
S = Reference area (m2)
CL= Lift coefficient
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By producing a greater pressure at the lower surface than the upper surface
of the body, a lift force will be generated. The difference of pressure is achieved
when the air speed at the upper surface is higher compared to the lower surface. Lift
coefficient measures how efficient the wing is changing velocity into lift. The higher
lift coefficient indicates higher efficiency in the aerofoil design compares to aerofoils
with low lift coefficient[4][3][1].
The drag is defined as the resistance or force that opposes the motion of the
airfoil through the air [1][5][7]. It acts on the aerofoil in parallel to the relative wind.
The formula of drag is
Drag, D =
Where
D = Drag force (N)
ρ = Density (kg/m3
)
V = Velocity (m/s)
S = Reference area (m2
)
CD= Drag coefficient
Total Drag produced by an aircraft is the sum of the Profile drag, Induced
drag, and Parasite drag. Total drag is primarily a function of airspeed. The airspeed
that produces the lowest total drag normally determines the aircraft best rate of climb
speed, minimum rate of descent speed for autorotation, and maximum endurance
speed [5][7].
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The pitching moment is defined as the force that tends to push the nose
upwards or downwards. The pitching moment is positive when it tends to push the
nose upwards and negative when the nose tends to go downwards at zero lift
[1][4][5]. The formula of pitching moment is
Pitching moment, M =
where
M = Pitching moment force (N)
ρ = Density (kg/m3
)
V = Velocity (m/s)
S = Reference area (m2
)
CM
= Pitching moment coefficient
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2.4 TURBULENCE MODEL OVERVIEW
2.4.1 Spalart-Allmaras turbulence model
The Spalart-Allmaras turbulence model is the simple one-equation model that
is widely used in computational fluid dynamic. Spalart-Allmaras turbulence model is
one of the suitable and mostly chose in computational fluid dynamic to solve a
modelled transport equation for the kinematic turbulent viscosity due to its stability,
good results produced by it and the time consumption used to solve the problem. The
Spalart-Allmaras model was designed specifically to solve many applications
involving wall-bounded flows mostly for the aerodynamic problems such as
aerospace problems.
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CHAPTER 3
PARAMETER VALIDATION
The methodology of this project consists of three main stages, which are the
parameter validation, grid independence study and CFD simulation of BWB Baseline
II with canard. In parameter validation, an objects selection from the existing
aerodynamic analysis that is obtained from the journals or research papers will be
done. The object then is redraw back by using CAD software and the CFD
simulation will be done with the given parameters. The result from this simulation
then is compared with the original result to determine whether the result is same or
better than the referred analysis. The successful mesh and CFD setup used in this
section will be used in the second process, which is grid independent study.
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3.1 Parameter Validation
3.1.1 Geometry Selection
In parameter validation, the half model of BWB Baseline-II E4 is used
because its aerodynamics performance already known and validated via wing tunnel
testing. This model came from the previous research by Rizal et al. [16]. Also, this
object was chosen due to its similarity with the BWB Baseline-II E5 except it come
without a canard. The dimension of the half model as below:
Reference length, Lref = 0.6548 m
Reference Area, Sref = 1.3205 m2
Figure 3.1 Half model of BWB Baseline-II E4
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3.1.2 CFD Software Setup
For this project, CFD software from NUMECA International is used which
is the NUMECA FINE/HEXA. This CFD software package consists of HEXPRESS
which is the mesher, HEXTREAM for CFD solver and CFVIEW for post processor.
This software only limited to one type of mesh which is the hexa mesh. To enable the
selected geometry to be manipulated in this software, the geometry is converted to
the Parasolid format which is “.x_t” format.
For the experimental condition, a fix parameter is use according to the data
from the journal. The parameters used as below:
Table 3.1 Atmosphere parameter for the experimental condition
Condition value
Atmospheric pressure, Patm 101325 Pa
Air temperature 24oC or 297.5 K (average)
Air density
1.1642 kg/m3
Air kinematic viscosity 1.5482x105
m2/s
Air velocity, V 35 m/s
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To get the value of CL and CD for each angle of attack, the value of X and Y
component of flow direction is set to the angle of attack direction as below:
Table 3.2 Selected force vector values of X and Y component
Angle of
attack
Drag lift
x-component y-component x-component y-component
22 0.927 0.374 -0.374 0.927
26 0.898 0.438 -0.438 0.898
30 0.866 0.499 -0.499 0.866
34 0.829 0.559 -0.559 0.829
36 0.809 0.588 -0.588 0.809
40 0.766 0.643 -0.643 0.766
42 0.743 0.669 -0.669 0.743
44 0.719 0.695 -0.719 0.695
46 0.694 0.719 -0.694 0.719
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3.1.3 Parameter Validation Procedure
Below are the steps involved in the parameter validation proceses:
i. The selected geometry is imported into the HEXPRESS and a boundary
box with size 20 times of the BWB’s nose-to-wing tip length which is
2.25m is created. Then a Boolean operation is applied to the geometry
where the selected geometry which is the BWB Baseline-II E4 is
subtracted from the boundary box.
Figure 3 The BWB half model imported into the HEXPRESS
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Figure 3.3 Boundary box setup
ii. The geometry then is export as a domain. The maximum length of the
domain is divided into 10000 from the original length and the minimum
length is divided into 10, while for the curve chordal and surface plane
tolerance is divided by 10000. The parameters for the domain as below:
Minimum length: 0.00022225
Maximum length: 22.225
Curve chordal tolerance: 0.000111125
Surface plane tolerance: 0.000111125
Curve resolution: 3
Surface resolution: 3
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Figure 3.4 The BWB domain setup (left) and the domain result (right)
iii. The domain file then is imported back into the HEXPRESS for the
meshing process. Before proceed to the meshing process, all boundary
conditions was grouped and named as external for the external, mirror,
and BWB for the solid.
Figure 3.5 Boundary condition for the BWB
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iv. The meshing process started with the initial mesh size selection. For x-
axis and y-axis, the mesh number is set to 12 respectively while for z-
axis; it was set to 7 which resulting the total of initial mesh cell number is
1008. The mesh then is generated.
Figure 3.6 The initial mesh setup for the BWB
v. The process continued to the "adapt to geometry" process. In this process,
the entire BWB surface was selected and activated to change the cell size.
For this step, the cell size number is set to 0.0 for axis X, Y, and Z, and
the refinement number is set to seven. While for the trimming surfaces,
parameters remain as default. The mesh then is generated.
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Figure 3.7 Mesh adaptation process. Refinement number (upper right) and trimming
parameters (bottom)
Figure 3.8 Mesh adaptation result
vi. The next step is process to snap the mesh according to the geometry
shape. The setting is leave as default and then the command is accept.
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Figure 3.9 Snap to geometry process
Figure 3.10 Snap to geometry result
vii. Next, the process continues to the mesh optimization process. All setting
is leave as default. The function of the mesh optimization process is to
remove negative cell, twisted cell and improve the mesh quality.
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Figure 3.11 Mesh optimization process command
viii. The last meshing procedure is to create a viscous layer around the
geometry shape. The number of viscous layer computed automatically
according to the object reference length, kinematic viscosity and stream
velocity, which are 0.6548m,1.5482x105
m2/s and 35m/s.
Figure 3.12 Applying a viscous layer on the mesh process
Figure 3.13 Successful mesh with viscous layer surrounded the BWB
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ix. The successful mesh then exported to the HEXTREAM for computation
process. The mesh parameters obtained is;
Number of cells: 63618
Number of leaf cells: 63618
Number of vertices: 70627
x. In HEXTREM, parameters for the fluid properties such as temperature,
kinematic viscosity, and pressure was set according the data given in the
research journal of the object as stated in table 1.
Figure 3.14 Fluid properties selection
xi. The next step is to set the turbulence model, which is Spalart-Allmaras,
the reference length, and reference velocity. Because the Mach number
used is only 0.1, so the “low speed flow” option was selected.
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Figure 3.15 Fluid model properties
xii. The third step for the HEXTREM process is to set the boundary
conditions. For the solid, which is the BWB, it is remaining as default
with the “compute force and torque” option selected.
Figure 3.16 Boundary conditions setup
xiii. Next, the initial solution parameters are set to same as the fluid properties
given in table 1.
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Figure 3.17 Initial solution properties setup
xiv. For the numerical model parameters, the values remain as default except
for the characteristic velocity which is set to 35 m/s and the “user defined
gauge parameters” option was selected and set with the given pressure
and temperature in table 1.
Figure 3.18 Numerical model properties setup
xv. Next step is to set the output parameters, which is the value of CL, CD and
CM (if any). The value of lift direction and drag direction was set as given
in table 2. The “output of residuals to CFView” and “forces and
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moments” options selected. For the “force and moments”, the value is set
as given in table 1.
Figure 3.19 Output parameter setup for vector component
xvi. For the control variables, the number of iterations is set to 1000 iterations,
while for the convergence criteria is set to -5
Figure 3.20 Control variable to set the iteration number and convergence criteria
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xvii. The last step for HEXTREM is to run the CFD simulation until finish.
Figure 3.21 CFD simulation process is running
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3.1.4 The Results
For the parameter validation process, nine selective angles of attack was
chosen. The results from the CFD process are compared to the wind tunnel results as
below;
Table 3.3 Comparison between CL and CD
Angle of
attack
(degree)
CL
(reference)
CD
(reference)
CL
(obtained)
%
difference
CD
(obtained)
%
difference
22 0.750 0.327 0.785 4.67 0.358 9.48
26 0.835 0.423 0.860 2.99 0.469 10.87
30 0.872 0.511 0.942 8.03 0.580 13.50
34 0.906 0.610 0.948 4.64 0.664 8.85
36 0.916 0.659 0.951 3.82 0.709 7.59
40 0.938 0.767 0.949 1.17 0.801 4.43
42 0.945 0.823 0.945 0 0.847 2.92
44 0.939 0.870 0.937 0.21 0.892 2.53
46 0.936 0.924 0.919 1.82 0.932 0.87
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Figures below shows the graph of CL and CD versus angle of attack.
Figure 3.23: graph of CL versus angle of attack
Figure 3.22: graph of CD versus angle of attack
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
20 25 30 35 40 45 50
CL
angle of attack, α (o)
experimental
value
NUMECA
value
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
20 25 30 35 40 45 50
CD
angle of attack, α (o)
Experimental
value
NUMECA value
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3.1.5 Conclusion
From the results obtained in this parameter validation procedure, the value
of CL and CD from the simulation on the CFD has slightly low percentages of
difference but still not able to be used for the final simulation process. In addition,
grid independence studies need to carry on in order to get result that will not change
with the grid settings. This grid independence study process will be discussed in the
next chapter.
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CHAPTER 4
GRID INDEPENDENCE STUDY
This chapter will discuss about the grid independence study which is the
second step after the parameter validation process. The purpose of grid independence
study is to achieve a high quality mesh setting to be used for the final simulation and
to ensure that the mesh was refined enough to produce adequate results. This chapter
will explain how the grid independence is achieved through the several experiments
using the CFD software NUMECA FINE/HEXA based on the previous BWB
Baseline II E4 used in the parameter validation process.
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4.1 Introduction to Grid Independence Study
4.1.1 What is grid independence study?
To produce a good result on CFD simulation, the mesh or grid must have a
very goodand satisfactory qualityto ensure the result produced from the simulation is
accurate enough. The result also should not change on different grid parameter.
When the solution is not affected by the size or parameter of the grid, it can be said
that the grid is independence. Grid independence study is the process or method to
achieve the condition where the grid is no more affected on the solution even some
change on the grid size, the boundary box or the initial mesh is done. Once the grid
independence is achieved, the grid parameter can be used for the final simulation
process.
4.1.2 How to achieve grid independence
Because obtaining the correct grid setting is the most important things in
this study to obtain a very good result, study on the grid independence takes lot of
time before achieve it. There are several steps that can be used to achieve grid
independence. The steps will be used in this study is;
i) Test with various face settings
ii) Test the boundary box size with various box and initial mesh sizes
iii) Test with various refinement numbers
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4.2 Grid Independence Study Process
4.2.1 Test with Various Face Settings
Nazreen [10] in his study found the suitable face setting that can fit to any
grid settings after done several test on various values of the face settings. Table
below shows the parameters of the face setting according to his study;
Table 4.1 Face setting parameters
min length 0.007
max length /100
curve tolerance /1000
surface tolerance /1000
curve resolution 6
surface resolution 7
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4.2.2 Test the boundary box size with various sizes of box
The first step in the grid independence study process is to determine the
suitable boundary box size to be applied. This step was according to the study done
by Nizamuddin [12]. It is important to get the most suitable boundary box size
because if the boundary box is too big, the numbers of cells will probably also
increase and the simulation running time will be longer. But if the boundary box is
too small, the results of the boundary conditions effects on the subject body such as
angle of attack and the pressure contour will probably can’t be obtained as wanted.
To proceed with the boundary box size testing, the boundary box is divided
into seven differences box sizes with initial size of 40x40x20 (x1 = -20, y1 = -20, z1
= 0 and x2 = 20, y2 = 20, z2 = -20). To get the initial mesh value for each box, each
side of the box was divided by 10. For example, a 40x40x40 box size will gives an
initial mesh value of 4x4x2. For this test, face setting used according to the study
done by Nazreen [10] as stated on the 4.2.1.
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Table below shows the boxes sizes and its initial mesh values:
Table 4.2 Seven different box sizes with initial mesh values
Box no. Box scale Box size Initial mesh
1 1x x1 -20 y1 -20 z1 0
4 4 2 x2 20 y2 20 z2 -20
2 1.5x x1 -30 y1 -30 z1 0
6 6 3 x2 30 y2 30 z2 -30
3 2x x1 -40 y1 -40 z1 0
8 8 4 x2 40 y2 40 z2 -40
4 2.5x x1 -50 y1 -50 z1 0
10 10 5 x2 50 y2 50 z2 -50
5 3x x1 -60 y1 -60 z1 0
12 12 6 x2 60 y2 60 z2 -60
6 3.5x x1 -70 y1 -70 z1 0
14 14 7 x2 70 y2 70 z2 -70
7 4x
x1 -80 y1 -80 z1 0
16 16 8 x2 80 y2 80 z2 -80
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The parameters for CFD simulation used in this studyare same as used in the
parameter validation. The parameters as below:
Table 4.3 Parameters for CFD simulation for box test
Conditions Value
velocity 35 m/s
Angle of attack 0o
Refinement number 8
Pressure 101325 Pa
Density of air 1.1642 kg/m3
Air kinematic viscosity 1.5482x105
m2/s
Reference area 1.3205 m2
Length 0.658 m
Moment centre 1.16 m
Target cell size x : 0.01, y : 0.01, z : 0.01
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The results obtained from the simulation of parameters above as shown in
table below:
Table 4.4 Results and cell numbers from box test
Box no. CL CD CM No. of cell Running time
1 0.310 0.031 -0.140
399,515
1H 26Min
2 0.314 0.031 -0.142
384, 038
2H 34Min
3 0.311 0.031 -0.141
400,332
1H 35Min
4 0.311 0.031 -0.141
386,789
1H 15Min
5 0.315 0.031 -0.143
384,795
1H 21Min
6 0.314 0.031 -0.143
400,937
1H 23Min
7 0.316 0.031 -0.144
386,262
2H 13Min
The results obtained from this study then are plotted on the graph to
determine which the best parameter to be used. Figures below shows the results
obtained from the boundary box test plotted in two graphs which is graph of CL
versus box number and CD versus box number.
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Figure 4.1: Graph of CL versus number of cell (left) and CD versus number of cell
(right)
From the graphs above, it can be seen that the value of CL does not change
much for the boundary box 1 until 4. It can be seen also that the value of CDalmost
nearly constant for any boundary box sizes. From the observation of the graphs, it
can be said that the boundary box number 4 satisfy the conditions where it has an
average number of cells plus with the lowest running time compare to the others
which only takes1 hour and 15 minutes. This result then will be used for the second
test which is the refinement number test.
0
0.05
0.1
0.15
0.2
0.25
0.3
0.35
CL
number of cell
Graph of CL vs number of cell
0
0.005
0.01
0.015
0.02
0.025
0.03
0.035
CD
number of cell
Graph of CD vs number of cell
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4.2.3 Test with various refinement numbers
The second step of grid independence study is to test the refinement number
to find the most suitable refinement number. A suitable refinement number will give
a good result without causing too much cell numbers produced which will lead to the
increasing of simulation running time.
In this test, parameters used in the previous test in section 4.2.1 will be used
with only 1 parameter value changed which is the refinement number. For this
test,Wan Zulhazri [19]suggested that the refinement number is tested from 7 until 13.
Table 4.5 below shows the results obtained from the simulation with 7 difference
refinement numbers.
Table 4.5 Results obtained from the refinement number test
Refinement no. CL CD CM No. of cell
7 0.317 0.039 -0.144 222772
8 0.317 0.029 -0.146 619581
9 0.314 0.025 -0.144 1841955
10 0.314 0.025 -0.144 1841955
11 0.314 0.025 -0.144 1841955
12 0.314 0.025 -0.144 1841955
13 0.314 0.025 -0.144 1841955
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Figure 4.2: Graph of CL versus refinement numbers (left) and CD versus refinement
number (right)
Figure 4.2 above shows the graph of CL and CD versus refinement numbers.
From the observation of the graph, it can be seen that the value of CL and CD are
constant starting from the refinement numbers of 9 until 13. From this observation, it
can be conclude that the most suitable refinement number to be used for the next
final simulation for BWB Baseline II E5-6 is 9. This is together with the parameters
used in the section 4.2.1. These parameters also will be used to run the simulation of
BWB Baseline II E4 and then will be compared with the wind tunnel results in the
next section.
0
0.05
0.1
0.15
0.2
0.25
0.3
0.35
7 8 9 10 11 12 13
CL
Graph of CL vs refinement
numbers
0
0.01
0.02
0.03
0.04
0.05
7 8 9 10 11 12 13
CD
Graph of CL vs refinement
numbers
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4.3Summary of the Parameters Obtained From Grid Independence Study
Below is the summary of the grid parameters obtained from the grid
independence study process. These parameters now are ready to be tested on the
BWB Baseline II E4 where the results obtained from the simulation will be discussed
in the next section.
Table 4.6 Summary of the parameters obtained from grid independence study
Parameter Value
Box size x1 -50 y1 -50 z1 0
x2 50 y2 50 z2 -50
min length 0.007
max length /100
curve tolerance /1000
surface tolerance /1000
curve resolution 6
Surface resolution 7
Initial mesh x 10 y 10 z 5
Refinement number 8
Target cell size x 0.01 y 0.01 z 0.01
Y+ 5
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4.4 Test the parameters obtained in grid independence study
For this section, the parameters obtained from the previous grid
independence study in section 4.1 and mentioned in section 4.3 will be tested on the
BWB Baseline II E4 and then the results will be compared with the results obtained
from the wind tunnel test. Figures below shows the results obtained after the
simulation.
Figure 4.3 Graph of CL versus angle of attack obtained from the simulation
and from the wind tunnel experiment
-0.8
-0.6
-0.4
-0.2
0
0.2
0.4
0.6
0.8
1
1.2
-20 -15 -10 -5 0 5 10 15 20 25 30 35 40 45 50 55
CL
Angle of attack, α (o)
exp
cfd
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Figure 4.4 Graph of CD versus angle of attack obtained from the simulation
and from the wind tunnel experiment
Figure 4.5 Graph of CM versus angle of attack obtained from the simulation
and from the wind tunnel experiment
-0.1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
1.1
-20 -10 0 10 20 30 40 50 60
CL
Angle of attack, α (o)
exp
cfd
-0.45
-0.4
-0.35
-0.3
-0.25
-0.2
-0.15
-0.1
-0.05
0
0.05
0.1
0.15
0.2
-20 -10 0 10 20 30 40 50 60
CM
Angle of attack, α (o) exp
cfd
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From the observation of graphs above, it can be seen that the parameters
obtained from the grid independence study gives a good results for lift coefficient
test and a slightly inaccurate for the drag coefficient and moment coefficient test.
4.5 Conclusion
From the overall results, it can be said that the parameters obtained from the
grid independence study is refined enough and satisfy the required qualities for grid
independence where the grid must independence and the result must not change on
the different grid settings. As the conclusion, these grid independence study achieve
its objective and the grid parameters is said to be ready to use for the final simulation
of the BWB Baseline II E5-6 which will be discussed in the next chapter.
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CHAPTER 5
RESULTS AND DISCUSSIONS
This chapter will discusses the aerodynamics results obtained from the
simulation of BWB Baseline II E5-6 at 0o angle of attack at different canard setting
angles through computational fluid dynamic. Aerodynamics coefficients such as CL,
CD and CM obtained from the simulation will be discussed in detail. This chapter also
will discuss the visualization obtained from the simulation such as Mach number
contour, pressure contour, velocity vector and streamline.
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5.1 Aerodynamics Analysis for BWB Baseline II E5-6
Table below show the results from the CFD simulation of BWB Baseline II
E5-6 with canard aspect ratio of 6 at 0o angle of attack. These results obtained from
12 differences canard setting angles starting from -6o to 11
o.
Table 5.1 Results obtained from the simulation
Canard setting
angle, δ (o)
CL CD CM L/D
-6 0.276 0.044 -0.187 6.337
-4 0.291 0.039 -0.161 7.406
-2 0.303 0.039 -0.131 7.694
0 0.316 0.039 -0.103 8.027
2 0.329 0.039 -0.075 8.447
4 0.341 0.040 -0.047 8.523
5 0.347 0.041 -0.033 8.489
6 0.352 0.042 -0.023 8.410
7 0.361 0.043 -0.007 8.376
8 0.365 0.045 0.00045 8.059
10 0.369 0.055 0.0051 6.743
11 0.368 0.056 0.00192 6.576
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5.1.1 Lift Coefficient, CL Analysis
Lift can be described as the effect of the average pressure differences
between the upper and the lower surface of the airfoil. Figure below shows the
variation of lift coefficient, CL as the canard setting angle, δ increase.
Figure 5.1 Graph of CL versus δ (left) and linear region at low canard angle (right)
From the graph above, it shown that the trend of graph is linear starting
from -4o to 6
obefore the CL values start to increase gradually until the maximum point
at 10o. This linear region is the condition where the flow over the airfoil moves
smoothly and still attached over most of the airfoil surface. The linear region given
0.27
0.28
0.29
0.30
0.31
0.32
0.33
0.34
0.35
0.36
0.37
0.38
-8 -6 -4 -2 0 2 4 6 8 10 12
CL
canard angle, δ (o)
CL = 0.0063δ + 0.3159
0.3
0.31
0.32
0.33
0.34
0.35
-4 -2 0 2 4 6
CL
canard angle, δ (o)
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by the equation of CL = 0.0063δ + 0.3159. The maximum lift, CLmax is at 10o
which is
the δmax before it starts to decreaseat 11o.At this 11
o point, flow over the canard’s top
surface commonly at the trailing edge,start to separate and creating a large wake or
relatively “dead air” behind the airfoil [6]. From the graph also it shows that even the
canard was positioned at the 0o, it still able to generate lift.
5.1.2 Drag Coefficient, CD Analysis
Figure 2 below shows the variation of drag coefficient,CD as the canard
setting angle, δ increases.
Figure 5.2 Graph of CD versus δ
0.038
0.040
0.042
0.044
0.046
0.048
0.050
0.052
0.054
0.056
0.058
-8 -6 -4 -2 0 2 4 6 8 10 12
CD
canard angle, δ (o)
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From the graph, it can be seen that the increament of CD is a little bit slow
for canard setting angle below 8o. The graph also shows that the increament value of
CD between -4o to 2
o are almost constant with very tiny percentage of differences.
The increament of drag value is linear starting from 2o until 7
o. At this condition, the
flow still attached on the BWB wing surface but slowly seperated from the canard
surface. Higher grows rate of CD occured at 8o
and it continue to hike due to the
increasing of air resistance.
Something interesting here is the value of CD at canard angle of 2o
is a little
bit lower than the value at 0o. Even the percentage of difference is just about 1% but
it is still unexpected due to the assumption that the resistance of air at 2o should be
higher than at 0o.
5.1.3 Pitching Moment Coefficient, CM Analysis
Figure 3 below presented the graph of pitching moment coefficient, CM
versus canard setting angle, δ. The pitching moment is measured 1.16 m from the
nose of the BWB.
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Figure 5.3 Graph of CM versus δ (left) and linear region at low canard angle (right)
From the curve above it can be seen that the value of CM increased as the
value of canard setting angle, δ increased until the maximum point it can reach at
10o. The graph also shows that the pitching moment is linear starting from the lowest
δ at -5o
before it starts to deflect at 5o
which corresponds to the flow separation at the
upper surface of the airfoil. Based on the graph it can be said that the canard possible
to change the BWB nose position which is the angle of attack of the BWB while
cruising from 0o angle of attack to certain unknown angle of attack.
-0.20
-0.18
-0.16
-0.14
-0.12
-0.10
-0.08
-0.06
-0.04
-0.02
0.00
0.02
-8 -6 -4 -2 0 2 4 6 8 10 12
CM
canard angle δ (o)
CM = 0.0141δ - 0.1031
-0.2
-0.16
-0.12
-0.08
-0.04
0
-10 -5 0 5 10
CM
canard angle δ (o)
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5.1.4 Lift-to-drag ratio analysis
Lift-to-drag ratio is the amount of lift divided by drag which is CL/CD. It is
one of the important aerodynamics parameter to study to find the optimum flight
configuration of the airplane. Figure 4 below presented the curve of lift-to-drag ratio
(L/D) as a function of canard setting angle, δ.
Figure 5.4 Graph of L/D versus δ
It is noticed that the L/D value increased from the lowest canard setting angle, δ until
it reach the maximum angle at 4o. This point which gives the value of L/D = 8.52
indicated the optimum flight configuration of the BW. The value of L/D continues to
dropped gradually after the 4o angle until 7
o.
6.0
6.2
6.4
6.6
6.8
7.0
7.2
7.4
7.6
7.8
8.0
8.2
8.4
8.6
8.8
-8 -6 -4 -2 0 2 4 6 8 10 12
L/D
canard setting angle, δ (o)
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5.1.5 Static Pressure Contour Analysis
Static Pressure contour is the visualization method that indicates the
pressure distribution around the subject or body simulated on the CFD software. It is
known that the lift only can be generated when there is a pressure differences
between the upper and lower surface of the airfoil. By analyze the static pressure
contour, pressure distribution behavior and characteristic on the BWB can be
obtained. Figures 5.5 until 5.8 below shows the static pressure contour taken at
canard setting angle δ of 0o, 8
o, 10
o and 11
o.
Figure 5.5 Static pressure contour of 0o canard angle viewed from cutting plane of
BWB (left) and top of BWB (right)
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Figure 5.6 Static pressure contour of 8o canard angle viewed from cutting plane of
BWB (left) and top of BWB (right)
Figure 5.7 Static pressure contour of 10o canard angle viewed from cutting plane of
BWB (left) and top of BWB (right)
Figure 5.8: Static pressure contour of 11o canard angle viewed from cutting plane of
BWB (left) and top of BWB (right)
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At 0o, the lift are mostly generated by the main wing of BWB as it can be
seen that the pressure around the upper surface at the leading edge of the wing is
lower than at the upper surface of the canard. Above this angle, the canard starts to
produce lift to the BWB. As can be seen on 8o, the canard upper surface is starts to
be surrounded by a blue color along its leading edge while at the main wing the
pressure is slightly higher compare to the condition at 0o. This indicated that the
pressure around this region is getting lower due to the air resistance caused the flow
over the canard moving slower than the surrounding.
The lift reached its maximum point at 10o as can be seen that the canard
upper surface now is fully covered with a low pressure region. From the cutting
plane, it can be seen that the low pressure region only focused at the leading edge of
the canard and main wing. This lowest pressure region around the top leading edge
of canard and main wing then starts to be fragmented at 11o caused the pressure
difference between the upper and lower canard surface become lower than at 10o.
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5.1.6 Mach Number Contour Analysis
To investigate more detail about what actually happened on the BWB when
the canard change its angle, Mach number contour can be used to see the air speed
profile on the BWB surface especially on the canard surface. Mach number is a ratio
of air speed over the speed of sound. It can be obtained by using the simple equation
of Mach number, M = v/ . Mach number study is one of way to determine the
flow separation phenomenon occurs on the airfoil.
Figures 5.9 until 5.12 below shows the Mach number contours of top and
bottom of BWB surface when the canard positioned at 0o, 8
o, 10
o, and 11
o.
Figure 5.9 Mach number contour of 0o canard angle viewed from bottom surface of
BWB (left) and top surface of BWB (right)
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Figure 5.10 Mach number contour of 8o canard angle viewed from bottom surface of
BWB (left) and top surface of BWB (right)
Figure 5.11 Mach number contour of 10
o canard angle viewed from bottom surface
of BWB (left) and top surface of BWB (right)
Figure 5.12 Mach number contour of 11o canard angle viewed from bottom surface
of BWB (left) and top surface of BWB (right)
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At 0o canard deflection, the low velocity region indicated with the dark blue
color only occurred on the lower surface of the canard and also with a small partition
along the trailing edge at the upper surface of canard. At this point, the flow
separation only developed along the trailing edge of the canard and main wing. At 8o
canard deflection, the low velocity regions begin to spread on the upper surface of
canard while maintaining at the trailing edge of the main wing. At the lower surface
of canard and BWB, the velocity is getting higher. It can be seen that at this point,
the flow separation is getting higher until the maximum lift point the BWB can reach
which is at 10o of canard deflection.
The flow was fully separated from the upper surface of the canard at 11o
canard deflection angle. It can be seen that at 11o, the velocity at the lower surface of
the BWB is getting lower and this low velocity region start to spread from the body
to the main wing. This phenomenon causes the loss of lift to BWB.
5.1.7 Velocity Vector Analysis
Figures below shows the velocity vector viewed on the upper surface of
canard at the cutting plane when the canard was positioned at 11o. It can be seen that
the reverse flow on the boundary layer on the upper surface of the canard starts to
formed due to the flow separation.
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Figure 5.13 Velocity vector when canard positioned at 11o
Figure 5.14 Detail on the velocity vector when canard positioned at 11o (left) and
the beginning point of reverse flow zoomed from the left picture (right)
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According to Anderson [6], flow separation happened to the airfoil upper
surface when the boundary layer on the surface travels far enough against an adverse
pressure gradient. When this happened, the speed of the boundary layer relative to
the object dropped and becomes almost zero. At this situation, the fluid flow
becomes unable to attach to the airfoil surface. Figure below show the process of the
flow separation;
Figure 5.15 Velocity profile in the boundary layer. The flow start to separate at the
3rd
picture (right)
By referring to the Figure 5.14 above, it can be understand that he increasing
of distance downstream will lead to the flow separation on the boundary layer of the
canard upper surface, creating a reverse flow on the bottom of the boundary layer.
This reverse flow keep growing due to the increasing of the distance downstream,
causing the velocity flows on the surface getting lower and slowly transform the
upper surface of the canard into a condition which called as a “dead air”.
Increasing distance downstream
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CHAPTER 6
CONCLUSION AND RECOMMENDATIONS
6.1 Conclusion
The study on the aerodynamic characteristics of BWB Baseline II E5-6 with
canard aspect ratio of 6 at 0o angle of attack by CFD simulation has been done at 12
various canard setting angles. Aerodynamic characteristics obtained from the CFD
simulation such as CL, CD, CM, and L/D was plotted in the related graph to shows the
aerodynamic performance of the BWB.
The investigation on aerodynamics characteristic obtained from the
simulation of CFD shows that the BWB can achieve a maximum coefficient of lift
value, CLmax which is 0.369 at 10o canard setting angle before the flow separation
starts to occure. The separation of flow occured on the upper surface of the canard
only. From the study, it is found that the optimum flight condition of the BWB is at
4o canard setting angle.
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6.2 Recommendations
Based on the results obtained from the simulation, the flow separation is
founded to be occured at 11o canard setting angle. This angle seems like quite low
and should can be extended. So in order to achieve higher lift and delay the flow
separation, further study on modification of the canard type or size can be done such
as by replacing the canard with a thicker airfoil such as NACA 2316, NACA 2416,
or NACA 4415 for example. A thicker airfoil can helps to reduce the drag at the
higher angle of attack or at the higher canard setting angle, and also the blunt leading
edge of the thick airfoil will make the airflow over the upper surface remain attached
to the surface which will resulting of delaying on the flow separation.
For the simulation process using the CFD software, a deeper understanding
on the CFD software used which is NUMECA FINE/HEXA should be done in order
to fully utilized the software to get a better results in the next time. Also, for the
NUMECA software, the running time are very slow especially when simulate or
generate a large number of cell. The problem came from the firewall restriction
problem due to the share license used. This problem need to be solved in order to do
grid independence study with a large number of cell.
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REFERENCES
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[17] Roy Y. Myose, Shigeo Hayashibara, Ping-Chian Yeong, and L. Scott Miller,
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APPENDIX A - STATIC PRESSURE CONTOUR FOR EVERY CANARD
SETTING ANGLE, δ VIEWED FROM THE TOP AND BOTTOM SURFACE
OF THE BWB BASELINE II E5-6
δ Top surface Bottom surface
-6
-4
-2
0
2
4
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δ Top surface Bottom surface
5
6
7
8
10
11
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APPENDIX 2 - MACH NUMBER CONTOUR FOR EVERY CANARD
SETTING ANGLE, δ VIEWED FROM THE TOP AND BOTTOM SURFACE
OF THE BWB BASELINE II E5-6
δ Top surface Bottom surface
-6
-4
-2
0
2
4
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δ Top surface Bottom surface
5
6
7
8
10
11