BWB Project Report

Visvesvaraya Technological University
Belgaum, Karnataka-590 014
A Project Report on
“DESIGN AND CFD ANALYSIS OF BLENDED
WING BODY WITH HIGH LIFT DEVICE”
Project Report submitted in partial fulfillment of the requirement for the award of
the degree of
Bachelor of Engineering
In
Aeronautical Engineering
Submitted by
K P Sindhu 1SC10AE015
Mamatha C D 1SC09AE018
Under the Guidance of
External Guide: Internal Guide:
Mr. Sayee Chandrashekar Mouli Mr. Vikram V
Jet 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
Certificate
This 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 1SC10AE015
2. 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 Jet
aircraft………………………………………………………………………………….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 3
Computational
Fluid Dynamics

CHAPTER 5
SOFTWARE USED
IN THE PROJECT

CHAPTER 8
SOLUTION,
RESULTS AND
ANALYSIS

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.

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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.

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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

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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

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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.

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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.

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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.

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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).

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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

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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.

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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

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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

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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.

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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.

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2.5 Physics explanation
The general airplane lift equation demonstrates these relationships:
L=1/2ρV2
SCL
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 air
S is the planform area or Surface area of the wing and
CL 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.

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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

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 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

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 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

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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. Typical
application 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.

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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 design
stage, or it can used to analyze difficulties with an existing component and lead to
improved 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 in
CAD, 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 the
behavior of the fluid throughout the region of interest.
This can be lead to design modifications which can be tested by changing the geometry of
the 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 to
produce a mesh for input to the physics pre-processor. Before a mesh can be produced, a
closed geometry solid is required. The geometry and mesh can be created in the meshing
application or any of the other geometry mesh creation tools. The basic steps involve:

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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 imported
from most major CAD packages using native formats, and mesh of the control volumes is
generated automatically.
3.3(b) Defining the physics of the model
This interactive process is the second pre-processing stage and is used to create
input 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. Fluid
properties and boundary conditions are specified.
3.3(c) Solving the CFD problem
The component that solves the CFD problems is called solver. It produces the
required 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 as
the solutions approaches the extra solutions, it is said to converge. Each iteration, an
error, or residual is reported as a measure of the overall conservation of the flow
properties.
How close the final solution is to exact solution on a number of factors, including the size
and shape of the control volumes and size of the final residuals. Complex physical
processes such as combustion and turbulence are often modeled using empirical
relationships. The approximations inherent in this model also contribute the difference
between the CFD solutions and the real flow.

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The solutions process requires no user interaction and is, therefore usually carried out as a
batch process. The solver produces a results file which is then passed to the post
processor.
3.3(d) Visualizing the result in the post-processor
The post processor is the component used to analyze, visualize and present the
results interactively post-processing includes anything from obtaining point values to
complex 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.

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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.
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form,
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form,

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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.
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2014

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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

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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.

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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.

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CHAPTER 5
SOFTWARES USED IN PROJECT
5.1 CATIA: Introduction
CATIA (Computer Aided Three Dimensional Interactive Application) is a
multi-platform CAD/CAM/CAE commercial software suite developed by the French
company Dassault Systems. Written in C++ programming language, CATIA is the
cornerstone of the Dassault Systems product lifecycle management software suite.
CATIA completes in CAD/CAM/CAE market with Siemens NX, Creo Element/Pro, and
Autodesk Inventor. CATIA started as an in-house development in1977 by French aircraft
manufacturer Avions Marcel Dassault, at that time customer of the CAD/CAM/CAD
software 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 facilitates
collaborative 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 used
CATIA V5 for the 787 series aircraft. They have employed the full range of Dassault
Systems 3D PLM products CATIA, DELMIA, and ENOVIA LCA supplemented by
Boeing development applications. The development of the Indian Light Combat Aircraft
has 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.

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 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 and
applications, 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 dynamics
program that has been applied to solve wide- ranging fluid flow problems for over 20
years. At the heart of ANSYS CFX is its advanced solver technology, the key to
achieving reliable and accurate solution quickly and robustly. The modern, highly
parallelized solver is foundation for an abundant choice of physical models to capture
virtually any type of phenomena related to fluid flow.

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5.2.1 The structure of ANSYS CFX
ANSYS CFX consists of four software modules that take geometry and
mesh 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 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 a grid like
structured, 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 woven
strands. 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

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 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 geometry
or a .tin file in ICEM. You can do this by manually creating it in ICEM or importing data
into 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 of
tools and they can be quite useful. However, it is suggested that some planning is done
before 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 -> surface
mesh 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 it
and 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 if
needed and click ok.
Then, click volume mesh, and selecting the method (tetra for tetragonal unstructured
meshes) to generate the unstructured grid, press ‘ok’ and wait for the grid to be generated
and review the result.
5.2.3 ANALYSIS
We have accomplished CFD analysis in the meshed component with the
help of ANSYS CFX. It is explained below.
5.2.3(a) CFX Introduction
ANSYS CFX Software is a high-performance, general purpose fluid
dynamics program that has been applied to solved wide-ranging fluid flow problems for
over 20 years. The modern, highly parallelized solver is the foundation for an abundant
choice of physical models to capture virtually any type of phenomena related to fluid
flow.

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Integration into the ANSYS workbench planform provides superior bi-directional
connections to all major CAD systems, powerful geometry modifications and creation
tools with ANSYS Design Modeler, advanced meshing, advanced meshing technologies
in ANSYS meshing, and easy drag-and-drop transfer of data and result to share between
applications. For example, a fluid flow solution can be used in definition of boundary
load of subsequent structure mechanics stimulation. A negative two way connection to
ANSYS structure mechanics products allows capture of even the most complex fluid
structural interaction (FSI) problems in same easy-to-use environment, saving the need to
purchase, administer or run third-party coupling software.
The ANSYS CFX products allows engineer to test systems in a virtual
environment. A scalable program has been applied to the stimulation of water flowing
past ship hulls, gas turbine engine (including compressor, combustion chamber, turbines
and 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 the
solver. The result can be in a format of the users’ choice like charts, animations,
graphs, tables etc…

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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

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MH-45 airfoil gives, comparatively high maximum lift coefficient, can be used at
Reynolds numbers of 100'000 and above and has zero lift angle and has been used
successfully 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

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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

2014
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

2014
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.

2014
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
FIG 6.4: Catia Designed Model of BWB half model with flaps having 5deg
deflection

2014
deflection
deflection

2014
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

2014
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.

2014
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.

2014
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)

2014
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

2014

2014
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 Wall
Functions
Scalable
High Speed Model Off

2014
Table 5: Boundary Physics for BWB
Domain Boundaries
Air
Boundary – Inlet
Type INLET
Location D_INLET
Settings
Flow Regime Subsonic
Heat Transfer Total Temperature
Total
Temperature
2.8720e+02 [K]
Mass And
Momentum
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 And
Momentum
Average Static Pressure
Pressure Profile
Blend
5.0000e-02
Relative
Pressure
1.0132e+05 [Pa]
Pressure
Averaging
Average Over Whole Outlet

2014
Boundary – Symm
Type SYMMETRY
Location D_SYMM
Settings
Boundary - Aircraft
Type WALL
Location
W_LEAD, WING_TIP, UPPER_WING,
LOWER_WING, Primitive 2D C, Primitive 2D D
Settings
Heat Transfer Adiabatic
Mass And
Momentum
No Slip Wall
Wall Roughness Smooth Wall
Boundary – Wall
Type WALL
Location D_WALL
Settings
Heat Transfer Adiabatic
Mass And
Momentum
Free Slip Wall

2014
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

2014
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………

2014
Table 6: Obtained and Calculated Forces for BWB UAV Without Flap
Velocity
V
(m/s)
Angle of
Attack AOA
Lift
L
Drag
D
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
Surface
area
(m2
)
Density
(Kg/m3
)
Co-efficient of lift Co-efficient of
Drag
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

2014
Table 7: Obtained and Calculated Forces for BWB UAV With 50
deflection of Flap
Velocity
V
(m/s)
Angle of
Attack AOA
Lift
L
Drag
D
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 of
lift
Co-efficient of
Drag
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

2014
deflection of Flap
Velocity
V
(m/s)
Angle of
Attack AOA
Lift
L
Drag
D
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 of
lift
Co-efficient of
Drag
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

2014
deflection of Flap
Velocity
V
(m/s)
Angle of
Attack AOA
Lift
L
Drag
D
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 of
lift
Co-efficient of
Drag
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

2014
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

2014
 For BWB UAV with 50
deflection of Flaps

2014
deflection of Flaps

2014
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

2014
deflection of Flaps
FIG 8.20: Pressure Contour 00
AOA
FIG 8.21: Mach No Contour at 00
AOA

2014
deflection of Flaps
AOA
AOA

2014
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

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