bwb project report

Visvesvaraya Technological University

Belgaum, Karnataka-590 014

A Project Report on

“DESIGN AND CFD ANALYSIS OF BLENDEDWING BODY WITH HIGH LIFT DEVICE”

Project Report submitted in partial fulfillment of the requirement for the award ofthe degree of

Bachelor of EngineeringIn

Aeronautical Engineering

Submitted byK P Sindhu 1SC10AE015Mamatha C D 1SC09AE018

Under the Guidance of

External Guide: Internal Guide:

Mr. Sayee Chandrashekar Mouli Mr. Vikram VJet wings Technology, Bangalore. Lect. Dept. of AE, SCTIT, Bangalore.

S.C.T Institute of Technology, Bangalore-560 075

Department of Aeronautical Engineering

2013-14

S.C.T Institute of Technology, Bangalore-560 075

Department of Aeronautical Engineering

S.C.T.I.T

CertificateThis is to certify that the project work entitled “DESIGN

AND CFD ANALYSIS OF BLENDED WING BODY WITH HIGH

LIFT DEVICE” carried out by Miss. K P Sindhu, USN:1SC10AE015,

and Miss. Mamatha C D, USN:1SC09AE018, are bonafide students of

S.C.T Institute of Technology, in the partial fulfillment for the award

of Bachelor of Engineering in Department of Aeronautical Engineering

of the Visvesvaraya Technological University, Belgaum during the

year 2013-14. It is certified that all corrections/suggestions indicated

for Internal Assessment have been incorporated in the Report deposited

in the departmental library. The project report has been approved as it

satisfies the academic requirements in respect of Project work

prescribed for the Bachelor of Engineering Degree.

Mr. Vikram V Prof. S Narayanaswamy Dr. Sohan Kumar Gupta

Internal Guide Head of Department Principal

External Viva Examiner Signature with Date:

1.

2.

DECLARATION

We, the students of final semester of Aeronautical Engineering

Department, S.C.T Institute of Technology, Bangalore-560 075 declare that

the work entitled “DESIGN AND CFD ANALYSIS OF BLENDED

WING BODY WITH HIGH LIFT DEVICE” has been successfully

completed under the guidance of our internal guide Mr.Vikram V, Lecturer,

Aeronautical Department, S.C.T Institute of Technology, Bangalore and our

external guide Mr. Sayee Chandrashekar Mouli, Jet Wings Technology,

Bangalore. This dissertation work is submitted to Visvesvaraya

Technological University in partial fulfillment of the requirements for the

award of Degree of Bachelor of Engineering in Aeronautical Engineering

during the academic year 2013-2014. Further the matter embodied in the

project report has not been submitted previously by anybody for the award of

any degree or diploma to any university.

Place:

Date:

Team members:

1. K P SINDHU 1SC10AE0152. MAMATHA C D 1SC09AE018

Acknowledgement

The satisfaction and euphoria that accompany the successful completion of any

task would be incomplete without the mention of people, who are responsible for the

completion of the project and who made it possible, because success is outcome of hard

work and preservance, but stead fast of all is encouraging guidance. So with gratitude we

acknowledge all those whose guidance and encouragement served us to motivate towards

the success of the project. We would like to first and foremost thank God almighty for

successfully completing the project prescribed for the academic year 2013-14.

We take this opportunity to thanks beloved Dr. Sohan Kumar Gupta, Principal,

SCTIT, Bangalore for providing excellent academic environment in the college and his

never-ending support for the B.E program.

We would like to convey our sincere gratitude to Prof. S Narayanaswamy ,

HOD of Aeronautical Engineering, SCTIT, and Bangalore for his support and

encouragement given to carry out the project.

We wish to express our heartfelt thanks to our internal guide Mr. Vikram V,

Lecturer, Dept. of A.E and Mr. Sayee Chandrashekar Mouli, External guide, Jet

wings technology for their kind co-operation and encouragement given to pursue this

project.

Last but not the least I thank all those persons and well wishers who directly and

indirectly helped, motivated to complete this project successfully.

Abstract

In recent years there has emerged a significant increase of interest in the design of

Blended Wing Body (BWB) aircraft, specifically applied to a large commercial transport

aircraft. The BWB design has been proven to have significant improvements in

aerodynamic efficiency, as compared to the conventional wing fuselage design.

However, due to inability to counteract significant pitching moments there is difficulty in

the design of high lift devices for BWB, especially trailing edge devices. Due to large

wing area increased lift-to-drag ratio, it was found that, in terms of longitudinal stability,

high lift devices could be successfully applied to the aircraft, which would meet the take-

off and landing requirements for a field length comparable to those of current

conventional large transport aircraft.

In this project, we have designed BWB UAV model with and without high lift

devices. First we have made an analysis of BWB UAV without high lift device using

Ansys CFX Solver and then we have analyzed BWB UAV with high lift device at three

deflection angles. The results obtained by both analyses have been compared and changes

in the aerodynamic forces such as Lift, Drag are noted and stall angles for each case are

found using graphs.

Accordingly in present study an attempt has been made to design a Blended Wing

Body UAV using CATIA V5 and analyze it through CFD approach using ANSYS ICEM

CFX 14.5 to analyze the flow pattern, pressure fluctuations and other aerodynamic

characteristics of BWB at subsonic velocities.

Table of contents Page No.

Certificate

Acknowledgement

Abstract

List of Figures

List of Tables

Chapter 1

Introduction……………………………………………………………………………….1

1.1 Blended Wing Body (BWB)………………………………………………………….1

1.2 Formulation of the BWB concept…………………………………………………….2

1.3 Comparision of aerodynamic, inertial and cabin pressure loads……………………..7

1.4 Key concepts of BWB design…………………………………………………...……8

1.5 Advantages of Blended Wing Body aircraft……………………………………...…10

Chapter 2

Highliftdevices(HLD)……………………………………………..………………….…11

2.1 Introduction……...…..………………………………………………………………11

2.2 Purpose of HLD………………………………………………………………………12

2.3 Types of HLD………………...………………………………………………………12

2.4 Flaps………………………………………..……………………………………...…14

2.5 Physics explanation………………………………………………...…………….......15

2.6 Flaps during take off………………………………….………………………………15

2.7 Flaps during landing………………………………………………………….……...16

2.8 Types of flaps……………………….…………………………………..…..……….16

Chapter 3

Computational Fluid Dynamics (CFD)…………………………………………………19

3.1 Introduction………………………………………………………………………....19

3.2 Uses of CFD……………………………………………………..……………….…19

3.3 CFD methodology…………..……………………………..…………..…………....20

3.4 Discretization methods………..………………………………………………….…22

Chapter 4

Literature Survey………………………………………………….……………………25

4.1Wind Tunnel Experiments and CFD Analysis of Blended Wing Body……………25

4.2 Design And Test Of A UAV Blended Wing Body Configuration………..…….…25

4.3 Blended Wing Body Analysis And Design……..…………………………………26

4.4 Conceptual Design And Aerodynamic Study Of Blended Wing Body Business Jetaircraft………………………………………………………………………………….26

4.5 Aerodynamics Of High-Subsonic Blended-Wing-Body Configuration………...27

4.6 A feasibility study of HLD on BWB large transport aircraft……………..……….27

Chapter 5

Software’s used in the project………………….…………………………………..…28

5.1 CATIA………………………………………………………………………….…28

5.2 ANSYS…………………………………………………………………………....29

Chapter 6

Design Process………………………..………………………………………………33

6.1 Airfoil Selection…………………………………………………………………..33

6.2 Coordinates of MH-45 Airfoil………………………………………….……...…34

6.3 Geometry Parameterization……….……………………….………………..……36

6.4 CATIA V5……………………………………………..……………….………...37

Chapter 7

Meshing Process………………………………………………………..…………….40

7.1 Ansys ICEM CFD……………………………………………………..…………40

7.2 Meshed Models Of BWB UAV………………………………………..………...44

Chapter 8

Solution, Results and discussion……………………………..………………….….49

8.1 ANSYS CFX……………………………………..……………………………...49

8.2 Results…………………………………..……………………….………..….…..50

8.3 Graphs obtained from the above tables………………………………...….……..55

8.4 Contours obtained from the CFX RESULTS………………………………....…59

8.5 Comparitions of results. obtained………………………………………………..63

Chapter 9

Conculsion………….………………………………………………………….……64

Chapter 10

References…………………….…………………………………..………….……...65

List of Figures Page No.

FIG 1.1: A Blended Wing Body aircraft………………………………...…1

FIG 1.2: Aircraft design evolution, the first and second 44 ye…………….2

FIG 1.3: Early Blended Configuration concept………………………...….3

FIG 1.4: Early configuration with cylindrical pressure vessel and engines

burried in the wing root……………….…………………………….….….4

FIG 1.5: Effect of body type on surface area……………………….…..…5

FIG 1.6: Effect of wing/body on surface area……………………….…....6

FIG 1.7: Effect of engine installation on surface area….…………….…....6

FIG 1.8: Effect of controls integration on surface area……………...…….7

FIG 1.9: Comparision of aerodynamic,inertial,and cabin pressure loads….8

FIG 1.10: The Blended Wing Body aircraft…………………..…….……...9

FIG 2.1: Conventional aircraft moments………………..………...……….11

FIG 2.2: Plain flap..............................................................................16

FIG 2.3: Split flap…………………………………...………….…...……..17

FIG 2.4: Slotted flap…………………………………...….….……..…..…17

FIG 2.5: Flower flap…………………………………………….……..…..18

FIG 2.6: Gouge flap…………………………………………….……….…18

FIG 5.1: Structure of ANSYS CFX……………………………………….30

FIG 6.1: MH-45 Airfoil……………………………………..……………..34

FIG 6.2: Geometry of BWB UAV……………………………………..….36

FIG 6.3: Catia Designed Model of BWB half model without flaps..……..38

FIG 6.4: Catia Designed Model of BWB half model with flaps having 5deg

deflection…………………………………………………………………..38


deflection………………………………………………………………..…39


deflection………………………………………………………………….39

FIG 7.1: ICEM CFD half model of BWB UAV………….………….……40

FIG 7.2: Meshed model of half model BWB UAV……………….………42 FIG 7.3: Meshed model of BWB UAV without flaps………………….....44

FIG 7.4: Meshed model of BWB UAV with 5deg deflection of flap……..44

FIG 7.3: Meshed model of BWB UAV with 10deg deflection of flap…....45

FIG 7.6: Meshed model of BWB UAV with 20deg deflection of flap….…45

FIG 8.1: CFX ANALYZED BWB UAV……………..………………...49

FIG 8.2 to 8.17: Graphs obtained………………..……………………....55

FIG 8.18, 8.19: Pressure and Mach number contours for BWB UAV without

Flap at 0 deg AOA…………………………..……..…………………....59

FIG 8.20, 8.21: Pressure and Mach number contours for BWB UAV with 5 deg

Flap at 0 deg AOA………………………………………………………60


Flap at 0deg AOA………………………………………….……………61


Flap at 0deg AOA…………………………………………….…………62

FIG 8.26: Comparative graph of CL vs AOA for different flap angle.….63

List of Tables Page No. Table 1: Airfoil details……………………………………….……………33

Table 2: Airfoil coordinates……………………………………………….34

Table 3: Geometry details…………………………………………………37

Table 4: Domain physics for BWB………………………………………..45

Table5: Boundary Physics or BWB……………………………………….47

Table 6: Obtained and calculated forces for BWB UAV without flap...…51

Table 7: Obtained and calculated forces for BWB UAV with 5 deg deflection of

flap…………………………………………………………………………52

Table 8: Obtained and calculated forces for BWB UAV with 10deg deflection of

flap………………………………………………………………………….53

Table 9: Obtained and calculated forces for BWB UAV with 20 deg deflection of

flap………………………………………………………………………….54

Table 10: Comparison table………………………………………………...63

Chapter 1

Introduction to

BWB

CHAPTER 2

HIGH LIFT DEVICES

CHAPTER 3

Computational

Fluid Dynamics

CHAPTER 4

LITERATURE

SURVEY

CHAPTER 5

SOFTWARE USED

IN THE PROJECT

CHAPTER 6

DESIGN PROCESS

CHAPTER 7

MESHING

PROCESS

CHAPTER 8

SOLUTION,

RESULTS AND

ANALYSIS

CHAPTER 9

CONCLUSION

CHAPTER 10

REFRENCE

Design and CFD analysis of Blended Wing Body with High Lift Devices 2014

Dept. of Aeronautical Engineering, SCTIT, Bangalore Page 1

CHAPTER 1

INTRODUCTION

1.1 Blended Wing Body (BWB)

Blended wing body or Hybrid Wing Body aircraft have a flattened and airfoil

shaped body, where fuselage is merged with wing and tail to form a single entity.BWB is

a hybrid of flying-wing aircraft and the conventional aircraft where the body is designed

to have a shape of an airfoil and carefully streamlined with the wing to have a desired

planform.

If the wing in conventional aircraft is the main contributor to the generation of lift,

the fuselage of BWB generates lift together with the wing thus increasing the effective

lifting surface area. The streamlined shape between fuselage and wing intersections

reduces interference drag, reduces wetted surface area that reduces friction drag while the

slow evolution of fuselage-to-wing thickness by careful design may suggest that more

volume can be stored inside the BWB aircraft, hence, increases payload and fuel capacity.

FIG 1.1: A Blended Wing Body aircraft.



The BWB concept aims at combining the advantages of a flying wing with the

loading capabilities of a conventional airliner by creating a wide body in the center of the

wing to allow space for passengers and cargo. Especially, for very large transport aircraft,

the BWB concept is often claimed to be superior compared to conventional

configurations in terms of higher lift-to-drag ratio and consequently less fuel

consumption.

1.2 Formulation of the BWB concept

FIG 1.2: Aircraft design evolution, the first and second 44 years.

It is appropriate to begin with a reference to the Wright Flyer itself, designed and

first flown in1903. A short 44 year later, the swept-wing Boeing 4-47 took flight. A

comparison of these two airplanes shows a remarkable engineering accomplishment

within a period of slightly more than four decades. Embodiedinthe B-47 are most of the

fundamental design features of a modern subsonic jet transport swept wing and

empennage and podded engines hung on pylons beneath and forward of the wing. The

Airbus A330, designed 44 years after the B-47, appears to be essentially equivalent, as

shown in Fig 1.2.



Thus, in 1988, when NASA Langley Research Center’s Dennis Bushnell asked

the question, “Is there a renaissance for the long- haultransport?” there was cause for

reflection. In response, a brief preliminary design study was conducted at McDonnell

Douglas to create and evaluate alternate configurations.

A preliminary configuration concept, shown in Fig.1.3, was the result. Here, the

pressurized passenger compartment consisted of adjacent parallel tubes, a lateral

extension of the double-bubble concept. Comparison with a conventional configuration

airplane sized for the same design mission indicated that the blended configuration was

significantly lighter, had a higher lift-to-drag ratio, and had a substantially lower fuel

burn.

FIG 1.3: Early Blended Configuration concept



FIG 1.4: Early configuration with cylindrical pressure vessel and engines

burried in the wing root

The performance potential implied by the blended configuration provided the

incentive for NASA Langley Research Center to fund a small study at McDonnell

Douglas to develop and compare advanced technology subsonic transports for the design

mission of 800 passengers and a 7000-n mile range at a Mach number of 0.85. Composite

structure and advanced technology turbofans were utilized.

Defining the pressurized passenger cabin for a very large airplane offers two

challenges. First, the square-cube law shows that the cabin surface area per passenger

available for emergency egress decreases with increasing passenger count. Second, cabin

pressure loads are most efficiently taken in hoop tension. Thus, the early study began with

an attempt to use circular cylinders for the fuselage pressure vessel, as shown in Fig. 1.4.

along with the corresponding first cut at the airplane geometry. The engines are buried

in the wing root, and it was intended that passengers could egress from the sides of both

the upper and lower levels. Clearly, the concept was headed back to a conventional tube



and wing configuration. Therefore, it was decided to abandon the requirement for taking

BWB.

FIG 1.5: Effect of body type on surface area

Three canonical forms shown in Fig.1.5, each sized to hold 800 passengers, were

considered. The sphere has minimum surface area however, it is notstreamlined. Two

canonical streamlined options included the conventional cylinder and a disk, both of

which have nearly equivalent surface area. Next, each of these fuselage is placed on a

wing that has a total surface area of 15,000 ft.

Now the effective masking of the wing by the disk fuselage results in a reduction

of total aerodynamic wetted area of 7000ft compared to the cylindrical fuselage plus wing

geometry, as shown in Fig.1.6.



FIG 1.6: Effect of wing/body on surface area

FIG 1.7: Effect of engine installation on surface area

Next, adding engines (Fig.1.7) provides a difference in total wetted area of 10,200

ft. (Weight and balance require that the engines be located aft on the disk configuration.)

Finally, adding the required control surfaces to each configuration as shown in Fig.1.08

results in a total wetted area difference of 14,300ft 2 or a reduction of 33%. Because the

cruise lift-to-drag ratio is related to the wetted area aspect ratio, b2/Swet, the BWB

configuration implied a substantial improvement in aerodynamic efficiency.



FIG 1.8: Effect of controls integration on surface area

The fuselage is also a wing, an inlet for the engines, and a pitch control surface.

Verticals provide directional stability, control, and act as winglets to increase the effective

aspect ratio. Blending and smoothing the disk fuselage into the wing achieved

transformation of the sketch into a realistic airplane configuration.

1.3 Comparision of aerodynamic, inertial and cabin pressure loads

The unique element of the BWB structure is the center body as the passenger

cabin, it must carry the pressure load bending, and as a wing it must carry the wing

bending load. A comparison of the structural loading of a BWB with that of a

conventional configuration is given in Fig.1.9.



FIG 1.9: Comparision of aerodynamic,inertial,and cabin pressure loads.

1.4 Key concepts of BWB design

Since the initial design of the BWB wing in 1988, it has been refined to its current

state. The principal concept behind the current iteration of the BWB is the blending of

various components of the plane, including the fuselage, wings, and the engines, into a

single lifting surface. As a result, the BWB fuselage is harder to distinguish from the

wing (i.e. it is harder to tell where the wing ends and the fuselage begins).



There are some key concepts to note about the design of the BWB:

i. The BWB is a tailless aircraft: Because of the disc- shaped nature of the

fuselage, the BWB does not have a tail. As a result, the BWB does not

have a rudder.

ii. The engine location of the BWB: Another important characteristic of the

BWB design is position of the engines, are located at the aft sections of

the plane. Because of the weight and balance considerations of the plane,

the engines needed to be place at the rear of the plane. Additionally, with

the engines at the rear of the plane, the fuselage can serve as an inlet for

the intake of air.

iii. Control surfaces: The control surfaces of the wing are located along the

leading and trailing edges of the wing and on the winglets. The number of

control surfaces can vary from 14 to 20 depending on the BWB design.

FIG 1.10: The Blended Wing Body aircraft



1.5 Advantages of Blended Wing Body aircraft

The BWB has several distinct advantages over the conventional tube aircraft.

Some of these advantages are outlined below:

a. Higher fuel efficiency: Initial testing of the BWB aircraft has indicated that it can

have up to a 27% reduction in fuel burn during flight.

b. Higher payload capacity: Due to the blended nature of the fuselage, the fuselage is

no longer distributed along the centerline of the aircraft. As a result, the fuselage is

more spread out, allowing for greater volume and a larger payload capacity.

c. Lower takeoff weight: Early design concepts have determined that the BWB can

have up to a 15% reduction of take-off weight when compared to the conventional

baseline.

d. Lower wetted surface area: The compact design results in a total wetted difference

of 14,300 ft2, a 33% reduction in wetted surface area. This difference implies a

substantial improvement in aerodynamic efficiency.

e. Commonality: One of the greatest advantages of the BWB is commonality of size

and of application. Firstly, the commonality of the components of the airplane will

allow it the payload of the airplane to be varied at little cost. For the 250, 350, and

450 – passenger capacity of the BWB, many components are inter changeable. This

inte changeability serves to drive down the cost of the aircraft. Secondly,

commonality of function allows the BWB to be used in many applications, both

military and civilian. The BWB can be modified to be used as a fighter, troop

transport, tanker, and stand-off bomber in addition to its function as a commercial

airliner.



CHAPTER 2

HIGH LIFT DEVICES

2.1 INTRODUCTION

High-lift devices are moving surfaces or stationary components intended to

increase lift during certain flight conditions. They include common devices such as flaps

and slats, as well as less common features such as leading edge extensions and blown

flaps.

The motivation behind studing high lift device is based on the difficulties involved

in appling them to tail less aircraft as well as their advantage and necessity for large

aircraft in takeoff and landing configurations.

Typically, for a conventional aircraft with a tail, high lift devices can be applied

and moments created by addiational lift are countered by the deflection of the tail as

illustrated in the Fig. 2.1.

FIG 2.1: Conventional aircraft moments

However, with tail less aircraft there is no way to counteracting the pitching

moment created by the high lift devices. Because of this, most blended wing body desing

does not include the high lift devices or only employ simple leading edge slats. Not

having high lift devices results in high angles and velocities for landing and takeoff in

order to achive the required lift. This also creates a higher wing area in order to decrease

the wing loading (W/S) and increase the lift. For large commercial transport aircraft this

effect can be very difficult to handle. Large approach and take-off velocities and angles

not only make the flight uncomfortable but also inculde a significant increase in risk and



safety. Also, because of large size of aircraft to begin with, increasing the wing area

makes airport operation even more difficult.

2.2 Purpose of High lift devices

Aircraft designs include compromises intended to maximize performance for a

particular role. One of the most fundamental of these is the size of the wing, a larger wing

will provide more lift and reduce take-off and landing distance, but will increase drag

during cruising flight and thereby lead to lower than optimum fuel economy. High-lift

devices are used to smooth out the differences between the two goals, allowing the use of

an efficient cruising wing, and adding lift for take-off and landing.

2.3 Types of High Lift Device

2.3.1 Flaps

The most common high-lift device is the flap, a movable portion of the wing that

can be lowered into the airflow to produce extra lift. Their purpose is to re-shape the wing

section into one that has more camber. Flaps are usually located on the trailing edge of a

wing, while leading edge flaps are occasionally used as well. Some flap designs also

increase the wing chord when deployed, increasing the wing area to help produce more

lift such complex flap arrangements are found on many modern aircraft.

The first "travelling flaps" that moved rearward were starting to be used just

before World War II due to the efforts of many different individuals and organizations in

the 1920s and 30s, and have been followed by increasingly complex systems made up of

several parts with gaps in between, known as slotted flaps. Large modern airliners make

use of triple-slotted flaps to produce the massive lift required during take-off.

2.3.2 Slats and slots

Another common high-lift device is the slat, a small airfoil shaped device attached

just in front of the wing leading edge. The slat re-directs the airflow at the front of the

wing, allowing it to flow more smoothly over the upper surface while at a high angle of

attack. This allows the wing to be operated effectively at the higher angles required to

produce more lift. A slot is the gap between the slat and the wing. The slat may be fixed

in position, or it may be retractable. If it is fixed, then it may appear as a normal part of



the leading edge of a wing which has slot. The slat or slot may be either full span, or may

occur on only part of the wing (usually outboard), depending on how the lift

characteristics need to be modified for good low speed control. Often it is desirable for

part of the wing where there are no controls to stall first, allowing aileron control well

into the stall.

The first slats were developed by Gustav Lachmann in 1918 and simultaneously

by Handley-Page who received a patent in 1919, and by the 1930s had developed into a

system that operated by airflow pressure against the slat to close and small springs to

open at slower speeds or automatically when the airflow reached a predetermined angle-

of-attack on the wing, aerodynamic forces would then push the slat out. Modern systems,

like modern flaps, are more complex and are typically deployed hydraulically or with

servos.

2.3.3 Leading edge root extensions

Although not as common, another high-lift device is the leading edge root

extension (LERX) or leading edge extension (LEX). A LERX typically consist of a small

triangular fillet between the wing leading edge root and fuselage. In normal flight the

LERX generates little lift. At higher angles of attack, however, it generates a vortex that

is positioned to lie on the upper surface of the main wing. The swirling action of the

vortex increases the speed of airflow over the wing, so reducing the pressure and

providing greater lift. LERX systems are notable for the potentially large angles in which

they are effective, and are commonly found on modern fighter aircraft.

2.3.4 Boundary layer control and blown flaps

Powered high-lift systems generally use airflow from the engine to shape the flow

of air over the wing, replacing or modifying the action of the flaps. Blown flaps use

"bleed air" from the jet engine's compressor or engine exhaust which is blown over the

rear upper surface of the wing and flap, re-energising the boundary layer and allowing the

airflow to remain attached at higher angles of attack. A more advanced version of the

blown flap is the circulation control wing a mechanism that tangentially ejects air over a

specially designed airfoil to create lift through the Coanda effect.



A more common system uses the airflow from the engines directly, by placing a

flap in the path of the exhaust. The flap requires greater strength due to the power of

modern engines, and most designs deliberately "split" the flap so the portions directly

behind the engines do not move into the airflow.

2.4 Flaps

Flaps are devices used to improve the lift characteristics of a wing and are

mounted on the trailing edges of the wings of a fixed-wing aircraft to reduce the speed at

which the aircraft can be safely flown and to increase the angle of descent for landing.

They shorten take-off and landing distances. Flaps do this by lowering the stall speed and

increasing the drag.

Extending flaps increases the camber or curvature of the wing, raising the

maximum lift coefficient — the lift a wing can generate. This allows the aircraft to

generate as much lift, but at a lower speed, reducing the stalling speed of the aircraft, or

the minimum speed at which the aircraft will maintain flight. Extending flaps increases

drag, which can be beneficial during approach and landing, because it slows the aircraft.

On some aircraft, a useful side effect of flap deployment is a decrease in aircraft pitch

angle, which improves the pilot's view of the runway over the nose of the aircraft during

landing. However the flaps may also cause pitch-up depending on the type of flap and the

location of the wing.

There are many different types of flaps used, with the specific choice depending

on the size, speed and complexity of the aircraft on which they are to be used, as well as

the era in which the aircraft was designed. Plain flaps, slotted flaps, and Fowler flaps are

the most common. Krueger flaps are positioned on the leading edge of the wings and are

used on many jet airliners.



2.5 Physics explanation

The general airplane lift equation demonstrates these relationships:

L=1/2ρV2SCL

where: L is the amount of Lift produced,ρ is the air density,V is the indicated airspeed of the airplane or the Velocity of the airplane, relative

to the airS is the planform area or Surface area of the wing andCL is the lift coefficient, which is determined by the camber of the airfoil used, the

chord of the wing and the angle at which the wing meets the air (or angle of attack)

Here, it can be seen that increasing the area (S) and lift coefficient (CL) allow a

similar amount of lift to be generated at a lower airspeed (V).

Extending the flaps also increases the drag coefficient of the aircraft. Therefore,

for any given weight and airspeed, flaps increase the drag force. Flaps increase the drag

coefficient of an aircraft due of higher induced drag caused by the distorted spanwise lift

distribution on the wing with flaps extended. Some flaps increase the planform area of the

wing and, for any given speed, this also increases the parasitic drag component of total

drag.

2.6 Flaps during take-off

Depending on the aircraft type, flaps may be partially extended for take-off. When

used during take-off, flaps trade runway distance for climb rate using flaps reduces

ground roll and the climb rate. The amount of flap used on takeoff is specific to each type

of aircraft, and the manufacturer will suggest limits and may indicate the reduction in

climb rate to be expected. The Cessna 172S Pilot Operating Handbook generally

recommends 10° of flaps on take-off, especially when the ground is rough or soft.



2.7 Flaps during landing

Flaps may be fully extended for landing to give the aircraft a lower stall speed so

the approach to landing can be flown more slowly, which also allows the aircraft to land

in a shorter distance. The higher lift and drag associated with fully extended flaps allows

a steeper and slower approach to the landing site, but imposes handling difficulties in

aircraft with very low wing loading (the ratio between the wing area and the weight of the

aircraft). Winds across the line of flight, known as crosswinds, cause the windward side

of the aircraft to generate more lift and drag, causing the aircraft to roll, yaw and pitch off

its intended flight path, and as a result many light aircraft have limits on how strong the

crosswind can be, while using flaps. Further more, once the aircraft is on the ground, the

flaps may decrease the effectiveness of the brakes since the wing is still generating lift

and preventing the entire weight of the aircraft from resting on the tires, thus increasing

stopping distance, particularly in wet or icy conditions. Usually, the pilot will raise the

flaps as soon as possible to prevent this from occurring.

2.8 Types of flaps

Plain flap: The rear portion of airfoil rotates downwards on a simple hinge

mounted at the front of the flap. Due to the greater efficiency of other flap types,

the plain flap is normally only used where simplicity is required.

FIG 2.2: Plain flap



Split flap: The rear portion of the lower surface of the airfoil hinges downwards

from the leading edge of the flap, while the upper surface stays immobile. Like the

plain flap, this can cause large changes in longitudinal trim, pitching the nose

either down or up, and tends to produce more drag than lift. At full deflection, a

split flaps acts much like a spoiler, producing lots of drag and little or no lift.

FIG 2.3: Split flap

Slotted flap: A gap between the flap and the wing forces high pressure air from

below the wing over the flap helping the airflow remain attached to the flap,

increasing lift compared to a split flap. Additionally, lift across the entire chord of

the primary airfoil is greatly increased as the velocity of air leaving its trailing

edge is raised, from the typical non-flap 80% of freestream, to that of the higher-

speed, lower-pressure air flowing around the leading edge of the slotted flap. Any

flap that allows air to pass between the wing and the flap is considered a slotted

flap.

FIG 2.4: Slotted flap



Fowler flap: A Split flap that slides backward flat, before hinging downward,

thereby increasing first chord, then camber. The flap may form part of the

uppersurface of the wing, like a plain flap, or it may not, like a split flap, but it

must slide rearward before lowering. It may provide some slot effect.

FIG 2.5: Fowler flap

Gouge flap: A type of split flap that slides backward along curved tracks that

force the trailing edge downward, increasing chord and camber without affecting

trim or requiring any additional mechanisms.

FIG 2.6: Gouge flap



CHAPTER 3

COMPUTATIONAL FLUID DYNAMICS

3.1 INTRODUCTION

Computational fluid dynamics, usually abbreviated as CFD, is a branch of fluid

mechanics that uses numerical methods and algorithms to solve and analyze problems

that involve fluid flows or Computational fluid dynamics (CFD) is a computer- based tool

for simulating the behavior of systems involving fluid flow, heat transfer, and other

related physical processes. It works by solving the equation of fluid flow (in a special

form) over a region of interest, with specified (known) conditions on the boundary of that

region.

Computers are used to perform the calculations required to simulate the

interaction of liquids and gases with surfaces defined by boundary conditions. With high-

speed super computers, better solutions can be achieved. Ongoing research yields

software that improves the accuracy and speed of complex simulation scenarios such as

transonic or turbulent flows. Initial experimental validation of such software is performed

using a wind tunnel with the final validation coming in full-scale testing, e.g. flight tests.

The fundamental basis of almost all CFD problems are the Navier–Stokes

equations, which define any single-phase (gas or liquid, but not both) fluid flow. These

equations can be simplified by removing terms describing viscous actions to yield the

Euler equations. Further simplification, by removing terms describing vorticity yields the

full potential equations. Finally, for small perturbations in subsonic and supersonic flows

(not transonic or hypersonic) these equations can be linearized to yield the linearized

potential equations.

3.2 Uses of CFD

CFD is used by engineers and scientists in a wide range of fields. Typicalapplication includes:

1. Process industry: Mixing vessels, chemical reactors.

2. Building services: Ventilation of buildings, such as atriums.

3. Health and safety: Investigating the effects of fire and smokes.



4. Motor industry: Combustion modeling, car aerodynamics.

5. Electronics: Heat transfer within and around circuit boards.

6. Environmental: Dispersion of pollutants in air or water.

7. Power and energy: Optimization of combustion process.

8. Medical: Blood flow through grafted blood vessels.

3.3 CFD Methodology

CFD can be used to determine the performance of a component at the designstage, or it can used to analyze difficulties with an existing component and lead toimproved design.

For example the pressure drop through a component may be considered excessive.

The first step is to identify the region of interest.

The geometry of the region of interest is then defined. If the geometry already exists inCAD, improves directly. The mesh is then created. After importing the mesh into pre-processor, other elements of the simulation including the boundary conditions (inlet,outlet etc.,) and fluid properties are defined.

The flow solver is run to produce a file of results that contains the variation of velocity,pressure and any other variables throughout the region of interest.

The result can be visualized and can provide the engineer an understanding of thebehavior of the fluid throughout the region of interest.

This can be lead to design modifications which can be tested by changing the geometry ofthe CFD model and seeing the effect.

The process of performing a single CFD simulation is split into four components,

1. Creating the geometry/mesh.

2. Defining the physics of model.

3. Solving the CFD problems.

4. Visualizing the results in the post-processor.

3.3(a) Creating the geometry/mesh

This interactive process is the first pre-processing stage. The objective is toproduce a mesh for input to the physics pre-processor. Before a mesh can be produced, aclosed geometry solid is required. The geometry and mesh can be created in the meshingapplication or any of the other geometry mesh creation tools. The basic steps involve:



1. Defining the geometry of the region of interest.

2. Creating region of fluid flow, solid region and surface boundary names.

3. Setting properties for the mesh.

This pre-processing stage is now highly automated. In CFX geometry can be importedfrom most major CAD packages using native formats, and mesh of the control volumes isgenerated automatically.

3.3(b) Defining the physics of the model

This interactive process is the second pre-processing stage and is used to createinput required by the solver. The mesh files are loaded into physics pre-processor, CFX-pre.

The physical models that are to be included in the simulation are selected. Fluidproperties and boundary conditions are specified.

3.3(c) Solving the CFD problem

The component that solves the CFD problems is called solver. It produces therequired results in a non-interactive/batch process. A CFD problem is solved as follows:

1. The partial differential equations are integrated over all the control volumes in the

region of interest. This is equivalent to applying a basic conservation law (for

example, for mass or momentum) to each control volume.

2. These integral equations are converted to a system of algebraic equation by

generating a set of approximation for the terms in the integral equations.

3. The algebraic equations are solved iteratively.

An iterative approach is required because of the non-linear nature of the equations, and asthe solutions approaches the extra solutions, it is said to converge. Each iteration, anerror, or residual is reported as a measure of the overall conservation of the flowproperties.

How close the final solution is to exact solution on a number of factors, including the sizeand shape of the control volumes and size of the final residuals. Complex physicalprocesses such as combustion and turbulence are often modeled using empiricalrelationships. The approximations inherent in this model also contribute the differencebetween the CFD solutions and the real flow.



The solutions process requires no user interaction and is, therefore usually carried out as abatch process. The solver produces a results file which is then passed to the postprocessor.

3.3(d) Visualizing the result in the post-processor

The post processor is the component used to analyze, visualize and present theresults interactively post-processing includes anything from obtaining point values tocomplex animated sequences.

Examples of some important features of post-processors are:

Visualization of the geometry and control volumes.

Vectors plots showing the direction and magnitude of the flow.

Visualization of the variation of scalar variables (variable which have only

magnitude, not direction, such as temperature, pressure and speed) through the

domain.

Quantitative numerical calculations

Animation

Charts showing graphical plots of variables.

3.4 Discretization Methods

The stability of the selected discretization is generally established numerically

rather than analytically as with simple linear problems. Special care must also be taken to

ensure that the discretization handles discontinuous solutions gracefully. The Euler

equations and Navier-Stokes equations both admit shocks, and contact surfaces.

Some of the discretization methods being used are:

Finite volume method,

Finite element method,

Spectral element method.



3.4(a) Finite Volume Method

The finite volume method (FVM) is a common approach used in CFD codes, as it

has an advantage in memory usage and solution speed, especially for large problems, high

Reynolds number turbulent flows, and source term dominated flows (like combustion).

In the finite volume method, the governing partial differential equations (typically

the Navier-Stokes equations, the mass and energy conservation equations, and the

turbulence equations) are recast in a conservative form, and then solved over discrete

control volumes. This discretization guarantees the conservation of fluxes through a

particular control volume. The finite volume equation yields governing equations in the

form,

Where is the vector of conserved variables, is the vector of fluxes (see Euler

equations or Navier–Stokes equations), is the volume of the control volume element,

and is the surface area of the control volume element.

3.4(b) Finite Element Method

The finite difference method (FDM) has historical importance and is simple to

program. It is currently only used in few specialized codes, which handle complex

geometry with high accuracy and efficiency by using embedded boundaries or

overlapping grids (with the solution interpolated across each grid).

Where, is the vector of conserved variables, and , and are the fluxes in

the , and directions respectively.












form,






















form,













3.4(c) Spectral Element Method

Spectral element method is a finite element type method. It requires the

mathematical problem (the partial differential equation) to be cast in a weak formulation.

This is typically done by multiplying the differential equation by an arbitrary test function

and integrating over the whole domain. Purely mathematically, the test functions are

completely arbitrary - they belong to an infinitely dimensional function space. Clearly an

infinitely dimensional function space cannot be represented on a discrete spectral element

mesh. And this is where the spectral element discretization begins. The most crucial thing

is the choice of interpolating and testing functions. In a standard, low order FEM in 2D,

for quadrilateral elements the most typical choice is the bilinear test or interpolating

function of the form . In a spectral element method

however, the interpolating and test functions are chosen to be polynomials of a very high

order (typically e.g. of the 10th order in CFD applications). This guarantees the rapid

convergence of the method. Furthermore, very efficient integration procedures must be

used, since the number of integrations to be performed in numerical codes is big. Thus,

high order Gauss integration quadratures are employed, since they achieve the highest

accuracy with the smallest number of computations to be carried out. At the time there are

some academic CFD codes based on the spectral element method and some more are

currently under development, since the new time-stepping schemes arise in the scientific

world. You can refer to the C-CFD website to see movies of incompressible flows in

channels simulated with a spectral element solver or to the Numerical Mechanics website

to see a movie of the lid-driven cavity flow obtained with a completely novel

unconditionally stable time-stepping scheme combined with a spectral element solver.





















































CHAPTER 4

LITERATURE SURVEY

4.1 WIND TUNNEL EXPERIMENT AND CFD ANALYSIS OF

BLENDED WING BODY (BWB) UNMANNED AERIAL

VEHICLE(UAV) AT MACH 0.1 and MACH 0.3

By, Wirachman Wisnoe, Rizal Effendy Mohd Nasir, Wahyu Kuntjoro, and Aman

Mohd Ihsan Mamat

This paper reports the aerodynamic performance of UiTM BWB-UAV intended to

be capable for low subsonic operation. The 3-D model generated by CATIA became the

basis of the CFD model for predicting the pressure and flow distributions of the airplane,

which subsequently developed to be the aerodynamic load. Fluent software was employed

in the CFD analysis. Half model of the BWB has been used for wind tunnel tests.

Lift, drag, and pitching moment obtained from wind tunnel experiments have been

studied, analyzed and compared with the CFD results. The experiments have been

conducted around Mach 0.1 and the CFD analysis at Mach 0.1 and 0.3. These Mach

numbers represent the loitering and the cruising phase of the mission profile.

From the CL curves obtained from both CFD and wind tunnel experiments,

coupled with visualization using mini tuft, it can be concluded that this type of BWB can

fly at very high angle of attack. The maximum lift is given for α around 34º-39º. This is

due to the delta wing shape for the proposed BWB model. However, the wing is already

in stall condition at α around 8º, which is considered to be low. This means that the main

contributor of the lift is the aircraft body.

4.2 DESIGN AND TEST OF A UAV BLENDED WING BODY

CONFIGURATION

By, Kai Lehmkuehler, KC Wong and Dries VerstraeteSchool of Aerospace,

Mechanical and Mechatronic Engineering, The University of Sydney, Australia

This paper presented a design and testing of a blended wing body UAV airframe.

The design methodology using fast panel methods has been proven viable for an unusual

configuration. The wind tunnel tests matched the predicted data well and the flight testing



revealed good handling qualities in flight. Some problems during take off and landing due

to the limited aircraft stability and the presence of propulsion effects on the longitudinal

stability remain. The method used to obtain an engineering estimate of these effects has

been proven usuable.

4.3 BLENDED WING BODY ANALYSIS AND DESIGN

By, Mark A potsdam and Robert H Liebeck. McDonnell Douglas Aerospace, Long

Beach,California.

The Blended Wing Body is a novel aircraft configuration offering significant

performance advantages over modren, conventional, transonic transports. Aerodynamic

problems unique to this class of airplane are investigated with the aim of designing an

aerodynamically viable BWB configuration. Using CFD and constrained inverse design

methods.

Inverse design Navier-Stokes codes hanve been successfully applied to the

development of a new BWB configuration. The design is highly integrated and offers

performance improvements of significant proportions. CFD analysis and design methods

have been used to study the priliminary detailed aeerodynamic design of the BWB,

including inboard, kink, and outboard wing design.

4.4 CONCEPTUAL DESIGN AND AERODYNAMIC STUDY OF

BLENDED WING BODY BUSINESS JET

By, Harijono Djojodihardjo and Alvin Kek Leong Wei Universitiy Putra Malaysia.

A Conceptual Design and Aerodynamic Study of Business Jet BWB Aircraft is

carried out focusing on BWB Aerodynamics, including Wing Planform Configuration

and profiles, and their relationship to the Design Requirements and Objectives. Possible

Configuration Variants, Mission profile, Flight Envelope requirements, performance,

stability, as well as the influence of propulsion configuration and noise considerations of

BWB aircrafts are considered and elaborated.

The design of BWB configuration without the fuselage is the major contributor towards

low weight of the overall BWB configuration. This is because fuselage contains about

20% to 30% of overall empty weight of an aircraft which produces high drag yet less lift.



4.5 AERODYNAMICS OF HIGH-SUBSONIC BLENDED-WING-

BODY CONFIGURATIONS

By, Dino Roman, Richard Gilmore, Sean Wakayama, The Boeing Company,

Huntington Beach.

A Mach 0.93 Blended-Wing-Body (BWB) configuration was developed using

CFL3DV6, a Navier-Stokes computational fluid dynamics (CFD) code, in conjunction

with the Wing Multidisciplinary Optimization Design (WingMOD) code, to determine

the feasibility of BWB aircraft at high subsonic speeds. Excluding an assessment of

propulsion airframe interference, the results show that a Mach 0.93 BWB is feasible,

although it pays a performance penalty relative to Mach 0.85 designs. A Mach 0.90

BWB may be the best solution in terms of offering improved speed with minimal

performance penalty.

4.6 A FEASIBILITY STUDY OF HIGH LIFT DEVICES ON

BLENDED WING BODY LARGE TRANSPORT AIRCRAFT

By, Mechanical and Aerospace Engineering, San Jose state university.

The goal of this project was to look at the effect of applying High lift devices to a

blended wing body aircraft, specifically the effects on longitudinal stability.This gives an

idea as to weather or not high lift devices are feasible for this type of aircraft and if the

aircraft meets the requirements for safe take- off and landing

The result of this project shows that the two coonfigurations with only leading

edge devices and only traling edge devices both add a small amount of additional lift

while maintaining stability.



CHAPTER 5

SOFTWARES USED IN PROJECT

5.1 CATIA: Introduction

CATIA (Computer Aided Three Dimensional Interactive Application) is amulti-platform CAD/CAM/CAE commercial software suite developed by the Frenchcompany Dassault Systems. Written in C++ programming language, CATIA is thecornerstone of the Dassault Systems product lifecycle management software suite.CATIA completes in CAD/CAM/CAE market with Siemens NX, Creo Element/Pro, andAutodesk Inventor. CATIA started as an in-house development in1977 by French aircraftmanufacturer Avions Marcel Dassault, at that time customer of the CAD/CAM/CADsoftware to develop Dassault’s Mirage fighter jet, and then was adopted in the aerospace,shipbuilding and other industries.

5.1.1 Scope of this application

Commonly referred to as 3D Product Lifecycle Management software suite,CATIA supports multiple stages of product development (CAX), from conceptualization,design (CAD), manufacturing (CAM) and engineering (CAE). CATIA facilitatescollaborative engineering across disciplines, including surfacing & shape design,mechanical engineering, equipment and systems engineering.

5.1.2 CATIA in aerospace

The Boeing Company used CATIA V3 to develop its 777 airliner, and usedCATIA V5 for the 787 series aircraft. They have employed the full range of DassaultSystems 3D PLM products CATIA, DELMIA, and ENOVIA LCA supplemented byBoeing development applications. The development of the Indian Light Combat Aircrafthas been using CATIA V5.

Chinese Xian JH-7A is the first aircraft developed by CATIA V5, when the

design was completed on September 26, 2000.

European aerospace giant Airbus has been using CATIA 2001 Canadian aircraft

maker Bombardier Aerospace has done all of its aircraft design on CATIA.

The Brazilian aircraft company, EMBRAER, use Catia V4 and V5 to build all

airplanes.

Vought Aircraft Industries use CATIA V4 and V5 to produce its parts.

The Anglo/Italian Helicopter Company, AgustaWestland, use CATIA V4 and

V5 to design their full range of aircraft.



The main supplier of helicopters to the U.S Military forces, Sikorsky Aircraft

corp., uses CATIA as well.

Bell Helicopter, the creator of the bell Boeing V-22 Osprey, has used CATIA

V4, V5 and now V6.

5.1.3 Advantages of CATIA

It is very much necessary in the field of aerospace industries andapplications, because it supports multiple stages of product development.

The CATIA have many advantages when compared to other software:

1. It has multi-platform CAD/CAM/CAE commercial software.

2. It facilitates collaborative engineering across disciplines, including surfacing

and shape design, mechanical engineering, equipment and systems engineering.

3. CATIA offers a unique infrastructure that supports design of large assemblies,

knowledge based design.

4. Coast composites has reduced design time and automated communication,

allowing it to improve response times, take on more projects.

5. Enabling enterprises to reuse product design knowledge and accelerate

development cycles, CATIA helps companies speed their responses to market

needs and helps free users to focus on creativity and innovation.

5.2 ANSYS

ANSYS CFX is a high-performance, general purpose fluid dynamicsprogram that has been applied to solve wide- ranging fluid flow problems for over 20years. At the heart of ANSYS CFX is its advanced solver technology, the key toachieving reliable and accurate solution quickly and robustly. The modern, highlyparallelized solver is foundation for an abundant choice of physical models to capturevirtually any type of phenomena related to fluid flow.



5.2.1 The structure of ANSYS CFX

ANSYS CFX consists of four software modules that take geometry andmesh and pass the information require to perform a CFD analysis.

ANSYS CFX component

FIG 5.1: Structure of ANSYS CFX

5.2.2 ICEM CFD: Introduction

ANSYS ICEM CFD is popular proprietary software packages used forCAD and mesh generation.

Some open source software includes Open FOAM, Feat Flow, and Open FVM etc.Present discussion is applicable to ANSYS ICEM CFD software, it can create a grid likestructured, unstructured, multi-block, and hybrid grids with different cell geometries.

5.2.2(a) Meshes and its types

Mesh is similar to web or net in that it has many attached or wovenstrands. Mesh consists of semi-permeable barrier made of connected strands of metals,fiber, or other flexible/ductile material.

5.2.2(b) Types of mesh

First pass mesh

Triangular surface mesh

Tetrahedral solid mesh

Solid ‘brick’ mesh

Geometry generation software

Mesh generation software

ANSYS CFX-pre (physics pre-processor)

ANSYS CFX-solver (solver)

ANSYS CFD-post (post-processor)

ANSYS CFX-solver manager



Automated mesh generation

Refined mesh

5.2.2(c) Creating a structured grid

The first thing to do when creating a structural grid is to create the geometryor a .tin file in ICEM. You can do this by manually creating it in ICEM or importing datainto ICEM, for example 3-dimensional point data from a .txt file.

The tools available are specified under the geometry tab. There are quite a number oftools and they can be quite useful. However, it is suggested that some planning is donebefore beginning to make geometry. There are tools specifically for curves.

Curves can be split or joined to other curves.

Point can be created at cross-sections of curves’

Surfaces can be created from curves.

All of this gives extra flexibility in the methods of designing a grid.

5.2.2(d) Creating an unstructured grid

Once the curves and surfaces have been created, click the mesh tab -> surfacemesh and define the mesh density on the surfaces.

The surface menu is shown on the right, and to select surfaces, click the button next to itand start selecting surfaces, using middle-click when done. Then select a mesh density(.05 in this case, but will vary with each case) and check remesh selected surface ifneeded and click ok.

Then, click volume mesh, and selecting the method (tetra for tetragonal unstructuredmeshes) to generate the unstructured grid, press ‘ok’ and wait for the grid to be generatedand review the result.

5.2.3 ANALYSIS

We have accomplished CFD analysis in the meshed component with thehelp of ANSYS CFX. It is explained below.

5.2.3(a) CFX Introduction

ANSYS CFX Software is a high-performance, general purpose fluiddynamics program that has been applied to solved wide-ranging fluid flow problems forover 20 years. The modern, highly parallelized solver is the foundation for an abundantchoice of physical models to capture virtually any type of phenomena related to fluidflow.



Integration into the ANSYS workbench planform provides superior bi-directionalconnections to all major CAD systems, powerful geometry modifications and creationtools with ANSYS Design Modeler, advanced meshing, advanced meshing technologiesin ANSYS meshing, and easy drag-and-drop transfer of data and result to share betweenapplications. For example, a fluid flow solution can be used in definition of boundaryload of subsequent structure mechanics stimulation. A negative two way connection toANSYS structure mechanics products allows capture of even the most complex fluidstructural interaction (FSI) problems in same easy-to-use environment, saving the need topurchase, administer or run third-party coupling software.

The ANSYS CFX products allows engineer to test systems in a virtualenvironment. A scalable program has been applied to the stimulation of water flowingpast ship hulls, gas turbine engine (including compressor, combustion chamber, turbinesand after burner), aircraft aerodynamics, pumps, fans, vacuum cleaners and more

Basically, there are three features in CFX as follows:

CFX Pre CFX Solver manager

CFX Post

CFX Pre:

In ICEM CFD we develop the meshes over the model to be analyzed this

in turn after getting the required number of accuracy or quality, the model

is saved in .cfx5 format .this file imported into CFX Pre and required

boundary conditions are given, this is in turn is saved as .cfx format.

CFX Solver Manager:

The resultant of CFX Pre is imported to CFX solver manager which carry outs

the solution iterations. After finishing the required number of iterations or after

meeting the required accuracy the result files are generated.

CFX Post:

The CFX Post is used to visualize the result developed by thesolver. The result can be in a format of the users’ choice like charts, animations,graphs, tables etc…



CHAPTER 6

DESIGN PROCESS

6.1 Airfoil Selection

For 2D airfoil selection in the conceptual design, a basic and simple

approach was adopted by analyzing chosen airfoil using Airfoil Investigation Database

and on-line DesignFOIL software, which are interactive database and programs. Eppler,

Martin Happerle and NACA airfoil series were analyzed for the BWB conceptual design.

The airfoil selection process was focused on the airfoil components to achieve favorable

pressure distribution, maximum lift and minimum drag coefficients.

The Martin Happerle, MH-45 airfoil was best suited for our selected geometry

which is a cambered airfoil and the same airfoil is used for center body, wing root and

wing tip and it has characteristics as follows;

Table 1: Airfoil Details

Parameters Dimensions Parameters Dimensions

Thickness 9.85% C Low moment coefficient, Cm +0.0145

Camber 1.7%C Max CL angle 9.50

Trailing edge angle 4.40 Max L/D 66.664

Lower flatness 66.6% Max L/D angle 6.50

Leading edge radius 0.7% Max L/D CL 0.792

Max CL 0.888 Stall angle 6.50



MH-45 airfoil gives, comparatively high maximum lift coefficient, can be used atReynolds numbers of 100'000 and above and has zero lift angle and has been usedsuccessfully in F3B tailless model airplanes.

FIG 6.1: MH-45 Airfoil

6.2 Coordinates of MH-45 airfoil

Table 2: Airfoil Co-ordinates

X Y X Y X Y

1.00000000 0.0000000 0.21770698 0.06354728 0.23480652 -0.03363521

0.99261598 -0.00017938 0.19881973 0.06240804 0.25458550 -0.03348553

0.97854084 -0.00009869 0.18008494 0.06094538 0.27441348 -0.03321288

0.96156846 0.00052546 0.16154175 0.05913397 0.29428723 -0.03283082

0.94307224 0.00157650 0.14323196 0.05694659 0.31421311 -0.03235163

0.92400403 0.00293829 0.12521660 0.05435169 0.33419194 -0.03178753

0.90464789 0.00454226 0.10759638 0.05132204 0.35421892 -0.03115008

0.88517240 0.00634500 0.09051004 0.04782719 0.37428397 -0.03045087

0.86566466 0.00832085 0.07415938 0.04384423 0.39437412 -0.02969917

0.84614988 0.01046165 0.05895840 0.03940609 0.41448315 -0.02890059

0.82661474 0.01276326 0.04529122 0.03459755 0.43461549 -0.02806037

0.80702606 0.01521504 0.03338913 0.02953397 0.45477836 -0.02718516



0.78734688 0.01780329 0.02403110 0.02472609 0.47496948 -0.02628269

0.76754804 0.02050221 0.01713554 0.02043615 0.49517672 -0.02535960

0.74762467 0.02327930 0.01188947 0.01647650 0.51539045 -0.02441874

0.72760217 0.02610621 0.00787937 0.01289035 0.53561204 -0.02346203

0.70753878 0.02895509 0.00493187 0.00986206 0.55584979 -0.02249252

0.68751592 0.03179249 0.00284325 0.00727975 0.57610687 -0.02151480

0.64772711 0.03731602 0.00144293 0.00489852 0.59637783 -0.02053345

0.62792805 0.03996674 0.00058749 0.00256853 0.61665369 -0.01955073

0.60816817 0.04252671 0.00022317 0.00028353 0.63692823 -0.01856798

0.58844483 0.04498332 0.00038544 -0.00184596 0.65719927 -0.01758624

0.56875417 0.04732603 0.00118455 -0.00374553 0.67746748 -0.01660629

0.54908524 0.04954601 0.00272273 -0.00545642 0.69773586 -0.01562863

0.52942550 0.05163874 0.00510389 -0.00712985 0.71800764 -0.01465263

0.50976714 0.05360070 0.00847008 -0.00898457 0.73828324 -0.01367727

0.49010752 0.05542837 0.01300718 -0.01115140 0.75856026 -0.01270218

0.47044736 0.05711903 0.01880976 -0.01346975 0.77883527 -0.01172777

0.45078952 0.05866975 0.02618997 -0.01591144 0.79910524 -0.01075480

0.43113912 0.06007589 0.03651865 -0.01872425 0.81936849 -0.00978503

0.41150396 0.06133193 0.05064046 -0.02186899 0.83962579 -0.00881954

0.39189352 0.06243131 0.06638508 -0.02468698 0.85988305 -0.00785512

0.35276213 0.06412564 0.08326064 -0.02708845 0.88015055 -0.00688847

0.37231341 0.06336537 0.10106768 -0.02905443 0.90039634 -0.00592428

0.33323972 0.06470335 0.11931857 -0.03057978 0.92052696 -0.00495974

0.31376185 0.06508936 0.13795805 -0.03173142 0.94050753 -0.00396586

0.29435718 0.06527007 0.15696051 -0.03257517 0.96003119 -0.00294225

0.2750461 0.06522556 0.17618669 -0.03315300 0.97789431 -0.00186493

0.25583261 0.06493634 0.19556817 -0.03349854 0.99254086 -0.00071029

0.23671715 0.66758058 0.21511470 -0.03364763 1.00000000 0.00000000



6.3 Geometry Parameterization

Based on the defined scope of the project, we have focused on the geometrical

aspects of BWB wing aircraft. By studying many papers we have selected the BWB UAV

geometry from Design of Blended Wing Body Unmanned Aerial Vehicle by Jeffrey L.

Williams, US, for Catia design.

A Blended Wing Body UAV is disclosed having a novel airfoil profile, wing

configuration, rigging and tractor pull propeller placement that provide improved stability

and safety characteristics over prior art SUAVs and MUAVs of comparable size and

weight. This unique blended wing design includes wing twist on the outboard wing and

an inverted "W" shaped planform to provide lateral and longitudinal stability, and

smooth, even flight characteristics throughout the range of the expected flight envelope.

These flight characteristics are crucial to providing a stable reconnaissance platform with

favorable stall speeds, an increased payload and the ability to hand launch without the

danger of exposing ones hands.

A wing assembly comprising a central main wing having outer edges and external

wings joined to the main wing at the outer edges. In this wing assembly, the airfoil has a

Reynolds number in the range from 20,000 to 100,000.

In this wing assembly the main wing has a pair of flaps on the outboard trailing

edge and the flaps are located at 15% of the chord from trailing edge of the airfoil.

In the figure below given the dimensions are in inches.

FIG 6.2: Geometry of BWB UAV



Table 3: Geometry Details

Parameter Dimension Parameter Dimension

Center chord 0.89m Half span 0.86m

Root chord 0.42m Sweep angle 300

Tip chord 0.26m Dihedral angle 00

Twist angle 00 Aspect Ratio 0.932

The Blended Wing Body geometry is designed using CATIA V5 design software.

6.4 CATIA V5

CATIA (Computer Aided Three-dimensional Interactive Application) is a multi-

platform CAD/CAM/CAE commercial software suite developed by the French company

Dassault Systems. CATIA is the cornerstone of the Dassault Systems product lifecycle

management software suite. CATIA competes in the high-end CAD/CAM/CAE market

with Creo Elements/Pro and NX (Unigraphics).

Commonly referred to as a 3D Product Lifecycle Management software suite,

CATIA supports multiple stages of product development (CAX), including

conceptualization, design (CAD), manufacturing (CAM), and engineering (CAE). CATIA

facilitates collaborative engineering across disciplines, including surfacing & shape

design, mechanical engineering, and equipment and systems engineering.

CATIA provides a suite of surfacing, reverse engineering, and visualization

solutions to create, modify, and validate complex innovative shapes, from subdivision,

styling, and Class A surfaces to mechanical functional surfaces.

It enables the creation of 3D parts, from 3D sketches, sheet metal, composites,

and moulded, forged or tooling parts up to the definition of mechanical assemblies. It

provides tools to complete product definition, including functional tolerances as well as

kinematics

Definition: It facilitates the design of electronic, electrical, and distributed systems

such as fluid and HVAC systems, all the way to the production of documentation for

manufacturing.



CATIA offers a solution to model complex and intelligent products through the

systems engineering approach. It covers the requirements definition the systems

architecture, the behaviour modelling and the virtual product or embedded software

generation. CATIA can be customized via application programming interfaces (API).

CATIA V5 and V6 can be adapted.

The BWB geometries we have designed in Catia V5 are given below;

FIG 6.3: Catia Designed Model of BWB half model without flaps


deflection




deflection


deflection



CHAPTER 7

MESHING PROCESS

7.1 Ansys ICEM CFD

ANSYS ICEM CFD is a popular proprietary software package used for CAD and

mesh generation. Some open source software includes Open FOAM, Feat Flow, and

Open FVM etc. Present discussion is applicable to ANSYS ICEM CFD software. It can

create structured, unstructured, multi-block, and hybrid grids with different cell

geometries.

FIG 7.1: ICEM CFD half model of BWB UAV



7.1.1 Geometric Modelling

ANSYS ICEM CFD is meant to mesh a geometry already created using other

dedicated CAD packages. Therefore, the geometry modelling features are primarily

meant to 'clean-up' an imported CAD model. Never-the-less, there are some very

powerful geometry creation, editing and repair (manual and automated) tools available in

ANSYS ICEM CFD which assist in arriving at the meshing stage quickly. Unlike the

concept of volume in tools like GAMBIT, ICEM CFD rather treats a collection of

surfaces which encompass a closed region as BODY. Therefore, the typical topological

issues encountered in GAMBIT (e.g. face cannot be deleted since it is referenced by

higher topology) don't show up here. The emphasis in ICEM CFD to create a mesh is to

have a 'water-tight' geometry. It means if there is a source of water inside a region, the

water should be contained and not leak out of the BODY.

Apart from the regular points, curves, surface creation and editing tools, ANSYS

ICEM CFD especially has the capability to do BUILD TOPOLOGY which removes

unwanted surfaces and then you can view if there are any 'holes' in the region of interest

for meshing. Existence of holes would mean that the algorithm which generates the mesh

would cause the mesh to 'leak out' of the domain. Holes are typically identified through

the color of the curves. The following is the color coding in ANSYS ICEM CFD, after the

BUILD TOPOLOGY option has been implemented:

Yellow: curve attached to a single surface - possibly a hole exists. In some

cases this might be desirable for e.g., thin internal walls require at least one

curve with single surface attached to it.

Red: curve shared by two surfaces the usual case.

Blue: curve shared by more than two surfaces.

Green: Unattached Curves - not attached to any surface.



7.1.2 Meshing Approach and Mesh

FIG 7.2: Meshed model of half model BWB UAV

There are often some mis-understandings regarding structured/unstructured mesh,

meshing algorithm and solver. A mesh may look like a structured mesh but may/may not

have been created using a structured algorithm based tool. For e.g., GAMBIT is an

unstructured meshing tool. Therefore, even if it creates a mesh that looks like a structured

(single or multi-block) mesh through pain-staking efforts in geometry decomposition, the

algorithm employed was still an unstructured one. On top of it, most of the popular CFD

tools like, ANSYS FLUENT, ANSYS CFX, Star CCM+, Open FOAM, etc. are

unstructured solvers which can only work on an unstructured mesh even if we provide it

with a structured looking mesh created using structured/unstructured algorithm based

meshing tools. ANSYS ICEM CFD can generate both structured and unstructured meshes

using structured or unstructured algorithms which can be given as inputs to structure as

well as unstructured solvers, respectively.



7.1.3 Structured Meshing Strategy

While simple ducts can be modelled using a single block, majority of the

geometries encountered in real life have to be modelled using multi-block strategies if at

all it is possible.

The following are the different multi-block strategies available which can be

implemented using ANSYS ICEM CFD.

O-grid

C-grid

Quarter O-grid

H-grid

7.1.4 Unstructured Meshing Strategy

Unlike the structured approach for meshing, the unstructured meshing algorithm is

more or less an optimization problem, wherein, it is required to fill-in a given space (with

curvilinear boundaries) with standard shapes (e.g., triangle, quadrilaterals-2D; tetrahedral,

hexahedral, polyhedral, prisms, and pyramids - 3D) which have constraints on their size.

The basic algorithms employed for doing unstructured meshing are:

1. Octree (easiest from the user's perspective; robust but least control over the

final cell count which is usually the highest)

2. Delaunay (better control over the final cell count but may have sudden

jumps in the size of the elements)

3. Advancing front (performs very smooth transition of the element sizes and

may result in quite accurate but high cell count)



7.2 MESHED MODELS OF BWB UAV

FIG 7.3: Meshed model of BWB UAV without flaps

FIG 7.4 Meshed model of BWB UAV with 5deg deflection of flap



The above meshed models are then imported into CFX-Pre and the boundary

conditions are specified as in following tables;

Table 4: Domain Physics for BWB

Domain – Air

Type Fluid

Location AIR

Materials

Air Ideal Gas

Fluid Definition Material Library

Morphology Continuous Fluid

Settings

Buoyancy Model Non Buoyant

Domain Motion Stationary

Reference Pressure 1.0000e+00 [Pa]

Heat Transfer Model Total Energy

Turbulence Model k epsilon

Turbulent WallFunctions

Scalable

High Speed Model Off



Table 5: Boundary Physics for BWB

Domain Boundaries

Air

Boundary – Inlet

Type INLET

Location D_INLET

Settings

Flow Regime Subsonic

Heat Transfer Total Temperature

TotalTemperature

2.8720e+02 [K]

Mass AndMomentum

Cartesian Velocity Components

U 0.0000e+00 [m s^-1]

V 6.8059e+01 [m s^-1]

W 0.0000e+00 [m s^-1]

Turbulence Medium Intensity and Eddy Viscosity Ratio

Boundary - Outlet

Type OUTLET

Location D_OUTLET

Settings

Flow Regime Subsonic

Mass AndMomentum

Average Static Pressure

Pressure ProfileBlend

5.0000e-02

RelativePressure

1.0132e+05 [Pa]

PressureAveraging

Average Over Whole Outlet



Boundary – Symm

Type SYMMETRY

Location D_SYMM

Settings

Boundary - Aircraft

Type WALL

LocationW_LEAD, WING_TIP, UPPER_WING,

LOWER_WING, Primitive 2D C, Primitive 2D D

Settings

Heat Transfer Adiabatic

Mass AndMomentum

No Slip Wall

Wall Roughness Smooth Wall

Boundary – Wall

Type WALL

Location D_WALL

Settings

Heat Transfer Adiabatic

Mass AndMomentum

Free Slip Wall



CHAPTER 8

SOLUTION, RESULT AND DISCUSSION

8.1 ANSYS CFX

ANSYS CFX is a commercial Computational Fluid Dynamics (CFD)

program, used to simulate fluid flow in a variety of applications. The ANSYS CFX

product allows engineers to test systems in a virtual environment. The scalable program

has been applied to the simulation of water flowing past ship hulls, gas turbine engines

(including the compressors, combustion chamber, turbines and afterburners), aircraft

aerodynamics, pumps, fans, HVAC systems, mixing vessels, hydro cyclones, vacuum

cleaners, and more.

FIG 8.1: CFX ANALYZED BWB UAV



ANSYS CFX software has its roots in the programs CFX-TASC flow and CFX-4.

CFX-4 was formerly Flow 3D in the United Kingdom and originally developed in-house

for use by the United Kingdom Atomic Energy Authority (UKAEA), and TASC flow

which was developed by Advanced Scientific Computing (ASC), of Waterloo, Ontario,

Canada.

FLOW 3D was commercialized by UKAEA in the late 1980s and early 1990s,

based on other in-house codes. It was renamed as CFX-4 in the mid-1990s, since the

name Flow-3D was already used in North America. The original product offering was

based on a multi-block structured hexahedral code based on a co-located segregated

implementation of the SIMPLE solution method. CFX-4 was very strong in the chemical

process industry and included some of the industry's most advanced multiphase and

chemistry models.

8.2 Results

By using CFX Pre solver we have given the required boundary conditions. Then

solution is done and then analyses is carried out. The results obtained are listed in the

next page………



Table 6: Obtained and Calculated Forces for BWB UAV Without Flap

VelocityV

(m/s)

Angle ofAttack AOA

LiftL

DragD

Total Lift Total Drag

68.06 0 42.0235 14.4337 42.0235 14.4337

68.06 4 188.301 10.0823 187.1391218 23.19199149

68.06 8 334.139 -1.32284 331.0717507 45.18978973

68.06 12 478.954 -19.6886 472.5824351 80.31448884

68.06 16 621.376 -44.8944 609.6821564 128.1066302

68.06 20 759.513 -76.6444 739.927517 187.727544

68.06 24 890.625 -114.276 860.114615 257.8269651

68.06 28 1006.13 -155.475 961.3632132 335.0385423

68.06 32 990.765 -124.458 906.1863192 419.4418122

Surfacearea(m2)

Density(Kg/m3)

Co-efficient of lift Co-efficient ofDrag

CL/CD

0.766154 1.225 0.019332417 0.006640054 2.911485

0.766154 1.225 0.086091154 0.010669203 8.069127

0.766154 1.225 0.152305669 0.020789032 7.326251

0.766154 1.225 0.217405997 0.036947737 5.884149

0.766154 1.225 0.280477113 0.05893395 4.759177

0.766154 1.225 0.340394961 0.086361851 3.941497

0.766154 1.225 0.395685624 0.118610266 3.336015

0.766154 1.225 0.442263852 0.154130545 2.869411

0.766154 1.225 0.41688037 0.192959278 2.160458



Table 7: Obtained and Calculated Forces for BWB UAV With 50 deflection of Flap

VelocityV

(m/s)

Angle ofAttack AOA

LiftL

DragD


68.06 0 125.78 15.9407 125.78 15.9407

68.06 4 270.35 10.2313 268.977891 29.06365662

68.06 8 413.692 -2.12428 409.962191 55.46697565

68.06 12 554.208 -21.0518 546.4756007 94.62612451

68.06 16 689.241 -45.8787 675.1898565 145.8652421

68.06 20 814.583 -75.5378 791.2984211 207.6011363

68.06 24 919.088 -107.244 883.2572426 275.8272478

68.06 28 979.84 -127.338 924.9414496 347.5409424

68.06 32 840.306 -55.5499 742.0732452 398.1548238

Surface area

(m2)

Density(Kg/m3)

Co-efficient oflift

Co-efficient ofDrag

CL/CD

0.766154 1.225 0.057863611 0.007333332 7.890494

0.766154 1.225 0.123740119 0.013370394 9.254785

0.766154 1.225 0.18859829 0.02551693 7.391104

0.766154 1.225 0.25139968 0.043531637 5.775103

0.766154 1.225 0.310613161 0.067103591 4.62886

0.766154 1.225 0.364027542 0.095504464 3.811629

0.766154 1.225 0.406332118 0.126891085 3.202212

0.766154 1.225 0.425508447 0.159882127 2.661388

0.766154 1.225 0.341382078 0.183166449 1.863781




VelocityV

(m/s)

Angle ofAttack AOA

LiftL

DragD


68.06 0 203.91 19.9142 203.91 19.9142

68.06 4 348.869 13.3264 347.0897635 37.62802401

68.06 8 491.543 0.0482399 486.7533276 68.45232808

68.06 12 629.962 -19.7477 620.303319 111.6507301

68.06 16 797.091 -44.5181 778.4875366 176.8984989

68.06 20 879.573 -75.6515 852.408501 229.7206112

68.06 24 971.84 -107.565 931.579808 296.9886802

68.06 28 1068.5 -131.876 1005.355399 385.1545805

68.06 32 840.521 -43.9959 736.1332962 408.0673485

Surface area

(m2)

Density(Kg/m3)

Co-efficient oflift

Co-efficient ofDrag

CL/CD

0.766154 1.225 0.0938064 0.009161294 10.23943

0.766154 1.225 0.159674568 0.01731033 9.224236

0.766154 1.225 0.22392515 0.031490689 7.110837

0.766154 1.225 0.285363254 0.051363607 5.555748

0.766154 1.225 0.358134045 0.081380076 4.400758

0.766154 1.225 0.392140516 0.105680268 3.710631

0.766154 1.225 0.428562346 0.136626153 3.136752

0.766154 1.225 0.46250194 0.17718584 2.610265

0.766154 1.225 0.338649474 0.187726589 1.80395




VelocityV

(m/s)

Angle ofAttack AOA

LiftL

DragD


68.06 0 352.863 34.6662 352.863 34.6662

68.06 4 497.716 26.3652 494.6647592 61.01734514

68.06 8 638.031 11.2139 630.2620803 99.89503807

68.06 12 771.13 -10.4649 756.4570497 150.0790559

68.06 16 892.906 -37.9677 868.7859637 209.6034835

68.06 20 996.764 -70.1161 960.6399648 275.0010857

68.06 24 1062.75 -103.249 1012.876013 337.9054624

68.06 28 1027.44 -98.8858 953.6139649 395.0085582

68.06 32 944.179 -50.1624 827.3100574 457.7644803

Surface area

(m2)

Density(Kg/m3)

Co-efficient oflift

Co-efficient ofDrag

CL/CD

0.766154 1.225 0.162330478 0.015947778 10.17888

0.766154 1.225 0.227564711 0.028070313 8.106953

0.766154 1.225 0.289944665 0.045955539 6.309243

0.766154 1.225 0.347999178 0.069042106 5.040391

0.766154 1.225 0.399674776 0.09642562 4.144902

0.766154 1.225 0.44193113 0.126511019 3.493222

0.766154 1.225 0.46596171 0.155449438 2.997513

0.766154 1.225 0.438698901 0.181719046 2.41416

0.766154 1.225 0.380594272 0.210589171 1.807283



8.3 Graphs obtained from the above tables

For BWB UAV without Flaps

FIG 8.2: CL vs α FIG 8.3: CD vs α

FIG 8.4: CL vs CD FIG 8.5: CL/CD vs α



For BWB UAV with 50 deflection of Flaps





For BWB UAV with 100deflection of Flaps





8.4 Contours obtained from the CFX Results

For BWB UAV without Flaps

FIG 8.18: Pressure Contour for BWB UAV at 00 AOA

FIG 8.19: Mach No Contour for BWB UAV at 00 AOA



For BWB UAV with 50 deflection of Flaps

FIG 8.20: Pressure Contour 00 AOA

FIG 8.21: Mach No Contour at 00 AOA



8.5 Comparision of Results Obtained

Table 10: Comparion table

FIG 8.26: Comparative Graph of CL vs AOA for different flap angles

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

0.5

0 10 20 30 40

CL vs AOA

AOA

CL

Blue Curve:without flap.Red Curve: 5°flap.Green Curve:10° flap.purple curve:20° flap

Conditions CL max CD max CL/CD

Without flap 0.442 0.194 8.06

5° flap 0.425 0.18 9.25

10° deg flap 0.463 0.18 10.23

20° deg flap 0.44 0.2 10.17



CHAPTER 9

Conclusion

This project report gives the aerodynamic performance of BWB-UAV intended to

be capable for low subsonic operation. The 3-D model generated by CATIA became the

basis of the CFD model for predicting the pressure and flow distributions of the airplane,

which subsequently developed to be the aerodynamic load. Ansys CFX software was

employed in the CFD analysis. Half model of the BWB UAV has been used for analyses.

Lift and drag forces were obtained from CFD results. The CFD analysis has been

carried out within Mach 0.1 and 0.3. These Mach numbers represent the loitering and the

cruising phase of the mission profile.

From the CL curves obtained from CFD analyses it can be concluded that this type

of BWB can fly at very high angle of attack. The maximum lift is obtained for α around

280- 300. However, the wing is already in stall condition at α around 8º, which is

considered to be low. This means that the main contributor of the lift is the aircraft body.

From the comparison table.8.1, it can observed that there is an increase in CLmax

and decrease in stall angle after introducing flaps, when compared with BWB UAV

model without flap. And hence there is an improvement in the aerodynamic efficiency.



CHAPTER 10

REFERENCES

1. Design And Test Of A Uav Blended Wing Body Configuration, by, Kai

Lehmkuehler, KC Wong and Dries VerstraeteSchool of Aerospace, The

University of Sydney, Australia.

2. Wind Tunnel Experiments and CFD Analysis of Blended Wing Body (BWB)

Unmanned Aerial Vehicle (UAV) at Mach 0.1 and Mach 0.3, by, Wirachman

Wisnoe, Rizal Effendy Mohd Nasir, Wahyu Kuntjoro, and Aman Mohd Ihsan

Mamat.

3. Blended Wing Body Analysis And Design, by, Mark A potsdam and Robert H

Liebeck. McDonnell Douglas Aerospace, Long Beach,California.

4. A Feasibility Study of High Lift Devices on Blended Wing Body Large Transport

Aircraft, by, Presented by Mechanical and Aerospace Engineering, San Jose state

university.

5. Liebeck, R. H., “Design of the Blended-Wing-Body Subsonic Transport,” 2002

Wright Brothers Lecture, AIAA Paper 2002-0002, Jan. 2002.

6. Liebeck, R. H., Page, M. A., Rawdon, B. K., “Blended-Wing-Body Subsonic

Commercial Transport,” AIAA Paper 98-0438, Jan. 1998.

7. “Blended-Wing-Body Technology Study,” Final Report, NASA Contract NAS1-

20275, Boeing Report CRAD-9405-TR-3780, Oct. 1997.

8. www.wikipedia.com

9. http://cfl3d.larc.nasa.gov/Cfl3dv6/cfl3dv6.html

bwb project report

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