THE AERODYNAMIC ANALYSIS OF BWB BASELINE II E5-6 UAV WITH CANARD ASPECT RATIO (AR) OF 6 AT ANGLE OF...

MUHAMMAD HILMI BIN ABD ADZIS 2009405194 JULY2012

i

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


ii

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


iii

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


iv

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.


v

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.


vi

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


vii

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


viii

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


tunnel experiment

Figure 4.5 Graph of CM versus angle of attack obtained 51


tunnel experiment


ix

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)



top of BWB (right)



top of BWB (right)



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)



of BWB (right)


x



of BWB (right)



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)


xi

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


1

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


2

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.


3

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.


4

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.


5

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)


6

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.


7

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.


8

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.


9

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


10

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.


11

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.

http://en.wikipedia.org/wiki/Navigator

http://en.wikipedia.org/wiki/Aviator


12

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.

http://en.wikipedia.org/wiki/Navier%E2%80%93Stokes_equations

http://en.wikipedia.org/wiki/Navier%E2%80%93Stokes_equations

http://en.wikipedia.org/wiki/Euler_equations_(fluid_dynamics)

http://en.wikipedia.org/wiki/Full_potential_equation

http://en.wikipedia.org/w/index.php?title=Linearized_potential_equations&action=edit&redlink=1


13

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


14

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


15

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


16

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


17

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].


18

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


19

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.


20

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.


21

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


22

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


23

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


24

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


25

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


26

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


27

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.


28

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.


29

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.


30

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


31

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.


32

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.


33

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


34

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


35

xvii. The last step for HEXTREM is to run the CFD simulation until finish.

Figure 3.21 CFD simulation process is running


36

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


37

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


38

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.


39

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.


40

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


41

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


42

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.


43

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


44

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


45

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.


46

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


47

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


48

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


49

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


50

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


51

Figure 4.4 Graph of CD versus angle of attack obtained from the simulation


Figure 4.5 Graph of CM versus angle of attack obtained from the simulation


-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


52

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.


53

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.


54

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


55

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



56

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



57

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.


58

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)


59

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)


60

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)


61





Figure 5.8: Static pressure contour of 11o canard angle viewed from cutting plane of



62

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.


63

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)


64

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)


65

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.


66

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)


67

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


68

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.


69

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.


70

REFERENCES

[1] A. C. Kermode, “Mechanics of Flight Tenth Edition”, Pearson Education Ltd.,

1996.

[2] A.M.E. Mamat, et al., "Aerodynamics of Blended Wing Body (BWB) Unmanned

Aerial Vehicle (UAV) using Computational Fluid Dynamics (CFD)," Journal of

Mechanical Engineeringvol. 5,pp. 15-27, 2008.

[3] Antony Jameson, “A Perspective on Computational Algorithms for Aerodynamic

Analysis and Design”, Department of Aeronautics and Astronautics, Stanford

University, Stanford, USA

[4] Barnard R. H. & Philpott D. R., “Aircraft Flight”, 2nd edition, Longman

[5] David F. Anderson and Scott Eberhardt, “Understanding Flight”, McGraw-Hill,

New York, 2000

[6] John D. Anderson Jr, Fundamentals of Aerodynamics Fourth Edition, McGraw-

Hill, New York, 2001

[7] John J. Bertin, “Aerodynamics For Engineers”,4th edition. Prentice Hall


71

[8] Liebeck, R.H., Page, M.A., and Rawdon, B.K. (1998). Blended-Wing-Body

Subsonic Commercial Transport. AIAA Paper 98-0438

[9] Mamat, A. M. I., Mohd Nasir, R. E., and Ngah, Z. (2005). Aerodynamics Of

Blended Wing Body (BWB) Unmanned Aerial Vehicle Using Computational Fluid

Dynamics (CFD). UiTM Malaysia – IRDC Research No. 600-IRDC/ST 5/3/1025

[10] Muhammd Nazreen B Zulkarnain, The Aerodynamic Analysis of BWB Baseline

II E5-8 UAV with Canard Aspect Ratio (AR) of 8 at Angle of Attack of 10 deg at 0.1

Mach Number through CFD Simulation at different canard setting

angles,Undergraduate Thesis, July 2012, Faculty of Mechanical Engineering,

Universiti Teknologi MARA Malaysia.

[11] N Qin, A Vavalle, A Le Moigne, M Laban, KHackett, P Weinerfelt.

Aerodynamics Studies forBlended Wing Body Aircraft. 9th

AIAA/ISSMOSymposium on Multidisciplinary Analysis andoptimization, 4 – 6

September (2002), Atlanta,Georgia.

[12] Nizammuddin Bin Daud, The Aerodynamic Analysis of BWB Baseline II E5-6

UAV with Canard Aspect Ratio (AR) of 6 at Angle of Attack of 10 degree at 0.1 Mach

Number through CFD Simulation at different canard setting angles, Undergraduate

Thesis, July 2012, Faculty of Mechanical Engineering, Universiti Teknologi MARA

Malaysia.


72

[13] Rizal E. M. Nasir, Wahyu Kuntjoro, Wirachman Wisnoe, Zurriati Ali, Nor F.

Reduan, Firdaus Mohamad, Shahrizal Suboh. PreliminaryDesign of Baseline-II

Blended Wing Body(BWB) Unmanned Aerial Vehicle (UAV):Achieving Higher

Aerodynamic EfficiencyThrough Planform Redesign and Low FidelityInverse Twist

Method. Proceedings of EnCon2010, 3rd Engineering Conference on Advancement

in Mechanical and Manufacturing for Sustainable Environment, 14-16 April (2010),

Kuching, Sarawak, Malaysia,

[14] R. E. M. Nasir, et al., "Aerodynamics of Sub-Sonic Blended Wing Body (BWB)

Unmanned Aerial Vehicle (UAV) Using Computational Fluid Dynamics (CFD),"

inInternational Conference on Mechanical & Manufacturing

Engineering(ICME2008), Universiti Tun HusseinOnn Malaysia (UTHM), Malaysia.,

2008.

[15] R. E. M. Nasir, et al., "The effect of canard on Aerodynamics and Static Stability

of Baseline-II Blended Wing-Body Aircraft at Low Subsonic Speed," in Conference

on Engineering and TechnologyEducation,World Engineering Congress 2010,

Kuching, Sarawak,Malaysia, 2010.

[16] Rizal E. M. Nasir, Wahyu Kuntjoro, Wirachman Wisnoe, Zurriati Ali, Nor F.

Reduan, Firdaus Mohamad, The Aerodynamics Performance of Blended Wing Body

Baseline-II E2, In proceedings of 2011 International Conference on Computer and

Communication Devices (ICCCD 2011)


73

[17] Roy Y. Myose, Shigeo Hayashibara, Ping-Chian Yeong, and L. Scott Miller,

Effect of Canards on Delta Wing Vortex Breakdown During Dynamic

Pitching,Journal Of AircraftVol. 34, No. 2, March – April 1997

[18] W. Wisnoe, et al., "Wind Tunnel Experiments of Blended Wing Body (BWB)

Unmanned AerialVehicle (UAV) At Loitering Phase," inProceedings of International

Conference on Mechanical &Manufacturing Engineering (ICME2008), Universiti

Tun HusseinOnn Malaysia (UTHM), Malaysia., 2008.

[19] Wan Zulhazri, The Aerodynamic Analysis of BWB Baseline II E5-8 UAV with

Canard Aspect Ratio (AR) of 8 at Angle of Attack of 0 degree at 0.1 Mach number

through CFD Simulation at different canard setting angles, Undergraduate Thesis,

July 2012, Faculty of Mechanical Engineering, Universiti Teknologi MARA

Malaysia.


74

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


75


5

6

7

8

10

11


76

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


-6

-4

-2

0

2

4


77


5

6

7

8

10

11

Date post:	29-Jul-2015
Category:	Documents
Upload:	hilmi-adzis
View:	383 times
Download:	8 times

THE AERODYNAMIC ANALYSIS OF BWB BASELINE II E5-6 UAV WITH CANARD ASPECT RATIO (AR) OF 6 AT ANGLE OF...

Documents