Business Idea generation

MAJOR PROJECT ON

“DESIGN AND ANALYSIS OF VARIFORM AIRFOIL”

A Dissertation Submitted

In partial fulfilment of the requirement for the award of the degree of

Bachelor of Technology

In

Aeronautical Engineering

By

N.ANUSHA	12N31A2161
P.MANASA	12N31A2168
PRIYANKA TIWARI	12N31A2178
R.SRI DIVYA SAHITHI	12N31A2184

Under the Guidance of

Mrs L. SUSHMA

Assistant Professor

Department of ANE

MRCET

DEPARTMENT OF AERONAUTICAL ENGINEERING

MALLA REDDY COLLEGE OF ENGINEERING AND TECHNOLOGY

(Sponsored by C.M.R. Educational society)

(Affiliated to JNTU, Hyderabad)

ACCREDITED by AICTE-NBA

Maisammaguda, Dhulapally post, Secunderabad-500014

DECLARATION

I hereby declare that the project entitled “DESIGN AND ANALYSIS OF VARIFORM AIRFOIL” submitted to MallaReddy college of Engineering and Technology, affiliated to Jawaharlal Nehru Technological University Hyderabad (JNTUH) for the award of the degree of Bachelor of Technology in aeronautical Engineering is a result of original research work done by us.

It is further declared that the project report or any part thereof has not been previously submitted to any University or Institute for the award of degree or diploma.

N. Anusha (12N31A2161)

P. Manasa (12N31A2168)

Priyanka Tiwari (12N31A2178)

R.Sri Divya Sahithi (12N31A2184)

CERTIFICATE

This is to certify that this is Bonafide record of the project titled” DEISGN AND ANALYSIS OF VARIFORM AIRFOIL” is submitted by N. Anusha, P. Manasa, Priyanka Tiwari, R. Sri Divya Sahithi of Roll no: 12N31A2161, 12N31A2168, 12N31A2178, 12N31A2184 of B.Tech in the partial fulfilment of the requirements for the degree of Bachelor of Technology in Aeronautical Engineering during the year 2015-2016. The results embodied in this project report have not been submitted to any other university or institute for the award of any degree or diploma.

Mrs L. Sushma MNVS. Swetha Bala

M. Tech M. Tech

INTERNAL GUIDE HEAD OF DEPARTMENT

EXTERNAL EXAMINER

ACKNOWLEDGEMENT

First and foremost, I offer my sincere gratitude to my internal guide Mrs L. Sushma, Assistant/Associate Professor of Aeronautical Engineering who has supported us throughout this project with her patience and valuable suggestions.

I would like to express my gratitude to Prof MNVS. Swetha Bala Head of the Department Aeronautical Engineering for her support and valuable suggestions during the dissertation work.

I am also grateful to the Principal Dr V. S. K. Reddy for providing me with all the resources in the college to make my project a success. I thank him for his valuable suggestions at the time of seminars which encouraged me to give my best in the project.

I would also like to thank all the supporting staff of the Dept. of ANE and all other departments who have been helpful directly or indirectly in making the project a success.

I am extremely grateful to my parents for their blessings and prayers for my completion of project that gave me strength to do my project.

.

N. Anusha (12N31A2161)

P. Manasa (12N31A2168)

Priyanka Tiwari (12N31A2178)

R.Sri Divya Sahithi (12N31A2184)

CONTENTS

· DECLARATION

· CERTIFICATE

· ACKNOWLEDGEMENT

· ABSTRACT

· LIST OF FIGURES

· LIST OF TABLES

CHAPTER 1- INTRODUCTION

1.1 What is an Airfoil?

1.2 What is Airfoil geometry?

1.3 What is variform concept?

1.4 Airfoils used

A.) NACA 23015 Airfoil

· NACA Airfoil History

· NACA-5 Digit Series

B.) WORTMANN FX 60-126 Airfoil

CHAPTER 2-LITERATURE SURVEY

CHAPTER 3-INTRODUCTION OF AIRFOIL

3.1 Airfoil

3.2 Airfoil History

3.3 Airfoil Terminology

3.4 How it works?

3.5 Various types of airfoil

· Symmetrical Airfoil

· Non-symmetrical Airfoil

· Flat Bottom Airfoil

· Supersonic Airfoil

· Supercritical Airfoil

3.6 Advantages

3.7 Which cause a Turning Moment?

CHAPTER 4- CFX

4.1 Introduction of CFX

4.2 CFX Documentation

4.3 CFX-Pre

4.4 CFX-Solver

4.5 CFX-Post

4.6 History of CFX

4.7 Ansys Fluent

CHAPTER 5- DESCRIPTION

CHAPTER 6-STEPS INVOLVED IN PROCESS

6.1 CFX

· Co-ordinates of airfoils

· CFX

· Meshing

· Fluent

· Theoretical Results

6.2 Manufacturing of Model

6.3 Experimental Analysis

CHAPTER 7- RESULTS AND DISCUSSSIONS

CHAPTER 8- CONCLUSION AND FUTURE SCOPE

CHAPTER 9- REFERENCES

ABSTARCT

In order to improve the fuel efficiency of an UAV, we propose using a wing that changes it shape during flight to minimize total mission drag the wing morphing approach is referred to a Variform wing concept. The fuel would be stored in balloon-like bladders inside the wing structure. These bladders would shrink as the fuel is consumed. Therefore the wing will be of NACA 5 digit airfoil during take-off with full fuel in the tank. During landing since the fuel if consumed to improve the lift wee vary it to develop an airfoil in wind tunnel model.

Machining is an important manufacturing process that is used in a wide range of applications. From aerospace applications to the manufacturing of energy systems and medical robots, we see a major reliance on machining. In this project we focus on gaining an improved understanding of the machine shop skills that provided us an opportunity to lathe and drill a class of components to specified dimensions and tolerances.

In this paper the two airfoils NACA 5 digit series-23015 and WORTMANN FX60-126 AIRFOIL 2D configurations are modelled using Catia and flow analysis is performed using ANSYS CFX. The coefficient of lift is compared for both the airfoils which are used for take-off and landing cases. Then the CNC coding is performed for the given airfoil configurations.

Due to their current successes, unmanned aerial vehicles (UAVs) are becoming a standard means of collecting information. However, as their mission become more complex and require them to fly farther, UAVs can become large and expensive due to fuel needs. Side stepping the paradigm of a fixed static wing, the variform concept developed in this paper allows for greater fuel efficiency. Bulky wings could morph into sleeker profiles, reducing drag, as they burn fuel. The development of such wings will rely heavily on computational design exploiting state of art optimization technique account for uncertainly and insure reliability.

LIST OF FIGURES

S.NO	NAME OF FIGURE	PAGE NO
FIG 1	AIRFOIL
FIG 2	AIRFOIL GEOMETRY
FIG 3	VARIFORM WING CONCEPT
FIG 4	NACA 23015 AIRFOIL
FIG 5	WORTMANN FX 60-126 AIRFOIL
FIG 6	WORTMANN AIRFOIL IN 3D VIEW
FIG 7	THREE AIRFOILS
FIG 8	AIRFOIL PROFILE
FIG 9	AIRFOIL NOMENCLATURE
FIG 10	DIFFERENT AIRFOIL THICKNESS
FIG 11	WORKING PRINCIPLE
FIG 12	TYPES OF AIRFOILS
FIG 13	CFX OF AIRFOIL
FIG 14	MESHING OF AIRFOIL
FIG 15	CFX SOLVER OF AIRFOIL
FIG 16	CFX POST OF AIRFOIL
FIG 17	MESHING OF NACA 23015
FIG 18	MESHING OF WORTMANN FX 60-126
FIG 19	TOP VIEW OF NACA AIRFOIL
FIG 20	SIDE VIEW OF NACA AIRFOIL
FIG 21	TOP VIEW OF WORTMANN
FIG 22	AIRFOIL WITH COPPER WIRES
FIG 23	WIND TUNNEL
FIG 24	TESTING OF WIND TUNNEL
FIG 25	PRESURE TUBES
FIG 26	NOTING OF PRESSUER HEAD VALUES
FIG 27	NOTING OF PRESSURE VALUES
FIG 28	NOTING OF PRESSURE HEAD VALUES

LIST OF TABLES

S.NO	NAME OF TABLE	PAGE NO
TAB 1	NACA 5 DIGIT POSITION AND MEAN LINE
TAB 2	NACA 5 DIGIT SERIES USES
TAB 3	NACA CO-ORDINATES
TAB 4	WORTMANN CO-ORDINATES
TAB 5	CNC CODING
TAB 6	CALCULATIONS OF NACA AIRFOIL
TAB 7	CALCULATIONS OF WORTMANN AIRFOIL

LIST OF GRAPHS

S.NO	NAME OF GRAPH	PAGE NO
1.	Cl OF NACA 23015
2.	Cd OF NACA 23015
3.	Cl OF WORTMANN FX 60-126
4.	Cd OF WORTMANN FX 60-126
5.	Cl/Cd RATIO OF NACA
6.	Cl/Cd RATIO OF WORTMANN
7.	VARIFORM CL/CD RATIO

CHAPTER 1

INTRODUCTION

The problem is that UAVs today have already been optimized for their current mission objectives. Thus, in order to conduct longer missions, more fuel and, hence, the size of the aircraft would need to be increased, in the proposal, a new paradigm is described that could allow these craft grater range without requiring them to carry more fuel. Alternatively, they could carry less fuel than current models and still achieve the same objectives.

In order to improve the fuel efficiency of UAVs, we propose using a wing that changes its shape during its flight to minimize total mission drag. This wing morphing approach is referred to as the variform wing concept. The fuel would be stored in balloon-like bladders inside the wing structure. These bladders would shrink as the fuel as the fuel is consumed. A bladder’s size and shape, as well as the possibility of multiple bladders inn a wing, will be determined by optimization Techniques. The integrated bladder and elastic structures of the wing provide a truly multidisciplinary problem that involves aerodynamics, structure and controls.

There are other classes of aircraft that could also benefit from the variform concept, such as micro aerial vehicles (MAVs). However, in this project we will focus on applications and examples involving UAVs. The variform concept presented here differs from other current research, like that done by NASA’s Aircraft Morphing Program, which uses piezoelectric to induce shape changes. The use of piezoelectric has been mostly for control. They are using them to rapidly change the shape of the trailing edge of the wings in place of ailerons (or in some cases to change the shape of the whole wing). The variform concept as presented here is to use the consumption of fuel to slowly change the shape of the entire wing throughout the flight to lower the drag on the craft over the entire mission.

Research being done at the Naval Research Laboratory (NRL) also deals with structure plus power concepts. Their program has three main design concepts. Their first concept is replacing some of the passive structure of smaller MAVs with battery material. More simply stated, they incorporate the battery as a structural member of the craft, so the weight of the battery is somewhat offset by lowering the weight of the structural components needed. A second NRL idea uses autophagous structure-fuel components. In this case the structure itself would be a fuel source, so the aircraft would self-consume structure over the course of the mission. Their last concept is another type of variform structure-power idea. In this case the structural members of the wing would be inflated with fuel. The first two concepts conserve constant aerodynamic shapes and thus are not very similar to the work presented in this paper. The third concept is similar in that the aerodynamic shape is altered via the fuel supply. However, the variform concept that is presented in this paper does not restrict the fuel to be inflated into structural members or in any certain configuration, but allows for an optimization process to choose where and how the structure should deform.

In this paper the variform concept is introduced in more detail, along with a description of the research that has been completed and an outline of future work. Initial efforts have focused on the development of a computational fluid dynamics model for predicting aerodynamic forces and performance. Future efforts will include the development of a finite element model of the bladder-filled elastic wing structure. The CFD and FEA codes will then be integrated to form a multidisciplinary model of the variform concept. Shape optimization techniques will be used to optimize the uninflected airfoil geometry and the uninflected bladder geometry. The optimization will seek to maximize the variform UAV’s range while accounting for system uncertainties. The model will allow users to explore different structural materials and variform design concepts.

1.1 What is an Airfoil?

A structure with curved surfaces designed to give the most favourable ratio of lift to drag in flight, used as the basic form of the wings, fins, and tail planes of most aircraft.

An airfoil is the term used to describe the cross-sectional shape of an object that, when moved through a fluid such as air, creates an aerodynamic force. Airfoils are employed on aircraft as wings to produce lift or as propeller blades to produce thrust. Both these forces are producing perpendicular to the air flow. Drag is a consequence of the production of lift/thrust and acts parallel to the airflow. Other airfoil surfaces include tail planes, fins, winglets, and helicopter rotor blades. Control surfaces (e.g. ailerons, elevators and rudders) are shaped to contribute to the overall airfoil section of the wing or empennage.

Fig1. Airfoil

1.2 What is Airfoil Geometry?

Airfoil geometry can be characterized by the coordinates of the upper and lower surface. It is often summarized by a few parameters such as: maximum thickness, maximum camber, position of max thickness, position of max camber, and nose radius. One can generate a reasonable airfoil section given these parameters. This was done by Eastman Jacobs in the early 1930's to create a family of airfoils known as the NACA Sections.

Fig2. Airfoil Geometry

1.3 What is variform Concept?

The variform wing is simply a wing that changes shape as fuel is consumed in order to maximize the lift to drag ratio. Maximizing this ratio leads to a great increase in the range of the aircraft. It is a way of making the aircraft go substantially further on the same amount or less fuel. For example, say we start with a NACA 23015 wing cross-section. As the fuel is used up, the wing could morph into the shape of a FX 60-126. This is captured in Figure 2, where the outside line is the larger NACA airfoil and the inner solid section is the sleeker shape.

Fig3. Variform wing concept

This shape changing could be done in a variety of ways. One way this change could occur would be to store the fuel in balloon like bladders that interact with the structure of the wing. When the bladders are filled the shape would look like the outer profile in Figure 2, and when empty the shape would look like the inner solid-filled shape. The simplest bladder configuration, as shown in Figure 3A, would just be an oval or any simple geometric shape. However, to achieve greater control of how the wing changes over time, a non-symmetric shape. Could be used as the bladder (Figure 3B), or even possibly multiple bladders of different size and shape seen in part C of the same figure.

1.4 Airfoils Used

A.) NACA 23015 AIRFOIL:

The NACA airfoils are airfoil shapes for aircraft wings developed by the National Advisory Committee for Aeronautics (NACA). The shape of the NACA airfoils is described using a series of digits following the word "NACA". The parameters in the numerical code can be entered into equations to precisely generate the cross-section of the airfoil and calculate its properties.

Fig4. NACA Airfoil

The early NACA airfoil series, the 4-digit, 5-digit, and modified 4-/5 digit, were generated using analytical equations that describe the camber (curvature) of the mean-line (geometric centerline) of the Airfoil section as well as the section's thickness distribution along the length of the airfoil. Later families, including the 6-Series, are more complicated shapes derived using theoretical rather than geometrical methods. Before the National Advisory Committee for Aeronautics (NACA) developed these series, airfoil design was rather arbitrary with nothing to guide the designer except past experience with known shapes and experimentation with modifications to those shapes.

· NACA Airfoil History:

During the 1930's several families of airfoils and camber lines were developed by the National Advisory Committee for Aeronautics (NACA). Many of these airfoil shapes have been successfully used over the years as wing and tail sections for general aviation and military aircraft, as well as propellers and helicopter rotors. The ordinates for numerous specific airfoils of these series at a coarse set of data points were published in a series of NACA reports. However, when performing parametric studies on effects of such variables as thickness, location of maximum thickness, leading- edge radius, location of maximum camber and others, it is not always easy to obtain the ordinates of the desired shapes rapidly and accurately. To remedy this problem the NASA Langley Research Center sponsored the development of computer programs for generation of ordinates of standard NACA airfoils.

Two separate programs were written by Charles Ladson and Curler Brooks of the NASA Langley Research Center in 1974-1975. The first was documented in NASA TM X-3284 and produces ordinates for NACA 4-digit, 4-digit modified, 5-digit, and 16-series airfoils. These thickness families are defined by algebraic equations. These thickness families are combined with appropriate mean lines to produce the final thick cambered airfoil.

The second program was documented in NASA TM X-3069 and produces ordinates for NACA 6-series and 6A-series airfoils. Unlike the other airfoils, these thickness distributions are not defined by algebraic equations, but use complex variable mapping of a circle into an airfoil shape. These thicknesses are combined with 6-series mean lines to produce the final thick cambered airfoil.

In December 1996, NASA published a new report outlining the theory behind the NACA airfoil sections and a revised computer program incorporating the features of both of the 1974-1975 programs. This report is designated TM-4741 and you can download a copy (PDF, 293KB) from the NASA document server. The program may be available from NASA.

· NACA 5 Digit series:

The NACA Five-Digit Series uses the same thickness forms as the Four-Digit Series but the mean camber line is defined differently and the naming convention is a bit more complex. The first digit, when multiplied by 3/2, yields the design lift coefficient (cl) in tenths. The next two digits, when divided by 2, give the position of the maximum camber (p) in tenths of chord. The final two digits again indicate the maximum thickness (t) in percentage of chord. For example, the NACA 23012 has a maximum thickness of 12%, a design lift coefficient of 0.3, and a Maximum camber located 15% back from the leading edge. The steps needed to calculate the coordinates of such an airfoil are:

1. Pick values of x from 0 to the maximum chord c.

2. Compute the mean camber line coordinates for each x location using the following equations,

And since we know p, determine the values of m and k1 using the table showed below.

Tab1. NACA 5 Digit position and Mean line

3. Calculate the thickness distribution using the same equation as the Four-Digit Series

4. Determine the final coordinates using the same equations as the Four-Digit Series.

Family	Advantages	Dis advantage	Applications
5-digit	1. Higher maximum lift coefficient. 2. Low Pitching moment 3. Roughness has little effect.	1. Poor Stall behavior. 2. Relatively high Drag.	1. General aviation 2. Piston-powered bombers, transports 3. Commuters 4. Business jets

Tab2. NACA 5 digit series uses

Today, airfoil design has in many ways returned to an earlier time before the NACA families were created. The computational resources available now allow the designer to quickly design and optimize an airfoil specifically tailored to a particular application rather than making a selection from an existing family.

B.)WORTMANN FX 60-126 AIRFOIL:

The Wortmann airfoils were designed in the early 1960's by Herr Doktor F. X. Wortmann of the Technischen Hochschule in Stuttgart, West Germany. They are primarily intended for sailplane and other low Reynolds number applications. They are laminar flow sections, medium to high cambered, have relatively small leading edge radii, feature camber relief between 40% and 70% of chord and have relatively thin trailing edges. They have plenty of camber in the nose, which makes them good at low Reynolds numbers; however, the pitching moments are very high on these airfoils.

Fig5. Wortmann Airfoil

This is not a big disadvantage on sailplanes, where the tail lengths are long (typically 4 to 5 chord lengths), but on general aviation applications, where typical tail lengths are 3C (3 chord lengths) and less, the resultant trim drag becomes a considerable disadvantage. The only GA application that makes sense with these airfoils is for airplanes without flaps, since the high camber produces relatively high CL max without flaps. High performance (laminar flow) GA applications without flaps are rare, however.

Fig6. Wortmann in 3D View

Fig7. Comparison of Three Airfoils by Mean Camber-Line

Figure 1 is a sketch of the three airfoils, for comparison. Also shown is the mean (camber) line for each. Notice the typical camber relief on the Wortmann airfoil between .40C and .70C. The idea is that this region of negative camber (negative lift) will delay the trip of the laminar boundary layer to turbulent flow, thus extending the laminar flow region, reducing drag. This "double hump" camber is common to all Wortmann airfoils, and has been copied on some of the recent NASA NLF airfoils. Unfortunately, it is not a good idea, especially since it penalizes the performance with flaps. Conventionally cambered airfoils have positive camber over 100% of their chord length, and are thus able to produce higher CL max with flaps, other things being equal.

Computations of drag polar for low speed wortmann sailplane airfoil are compared with both wind tunnel and flight tests results. Excellent Correlation was shown to exist between computations and flight results except when separate flow regimes were encountered.

CHAPTER 2

LITERATURE SURVEY

T .Gultop, (2005) studied the impact of perspective degree on Airfoil performance. The reason for this study was to focus the ripple conditions not to be kept up throughout wind tunnel tests. These studies indicate that aero elastic insecurities for the changing arrangements acknowledged showed up at Mach number 0.55, which was higher than the wind tunnel Mach number point of confinement velocity of 0.3.

Sanjay Goel, (2008) devised a method of optimization of Turbine Airfoil using Quansi –3D analysis codes. He solved the complexity of 3D modeling by modeling multiple 2D airfoil sections and joining their figure in radial direction using second and first order polynomials that leads to no roughness in the radial direction.

Mr. Arvind, (2010) researched on NACA 4412 airfoil and analyzed its profile for consideration of an airplane wing. The NACA 4412 airfoil was created using CATIA V5 And analysis was carried out using commercial code ANSYS 13.0 FLUENT at an speed of 340.29 m/sec for angles of attack of 0 ̊, 6, 12 and 16 ̊. K-ε turbulence model was assumed for Airflow. Fluctuations of static pressure and dynamic pressure are plotted in form of filled contour.

By J. Fazil and V. Jayakumar, (2011) concluded that despite the fact that it is less demanding to model and make an airfoil profile in CAD environment utilizing camber cloud of focuses, after the making of vane profile it is exceptionally troublesome to change the state of profile for dissection or improvement reason by utilizing billow of focuses. In the paper, they examined and depicted the making of airfoil profile in CAD environment utilizing the control purpose of the camber profile. By method for changing the qualities of control focuses the state of the profile could be effectively changed and additionally the outline of the cambered airfoil is created without influencing the essential airfoil geometry. In the said paper, the Quintic Reverse Engineering of Bezier bend recipe is utilized to discover the camber control focuses the current camber cloud of focuses.

Mr. Mayurkymar kevadiya, (2013) studied the NACA 4412 airfoil profile and recognized its importance for investigation of wind turbine edge. Geometry of the airfoil is made utilizing GAMBIT 2.4.6. Also CFD investigation is done utilizing FLUENT 6.3.26 at different approaches from 0 ̊ to 12 ̊.

CHAPTER-3

AIRFOIL

3.1 Introduction to airfoil:

One of the most spectacular things to view is the structure and the body of an aeroplane. Its concept has always been scintillating and technical. It all started with the answer to how birds can fly. All of us do know that only when an object overcomes the earth’s natural gravitational pull, it tends to fly. The wing of an aircraft helps in gliding it through the wind and also in its landing and take-off. The shape of such an important component of the aircraft makes a lot of impact on its movements. This shape is what is called an airfoil.

Fig8. Airfoil Profile

3.2 Airfoil History:

The earliest serious work on the development of airfoil sections began in the late 1800's. Although it was known that flat plates would produce lift when set at an angle of incidence, some suspected that shapes with curvature that more closely resembled bird wings would produce more lift or do so more efficiently. H.F. Phillips patented a series of airfoil shapes in 1884 after testing them in one of the earliest wind tunnels in which "artificial currents of air (were) produced from induction by a steam jet in a wooden trunk or conduit." Octave Chanute writes in 1893, "...it seems very desirable that further scientific experiments be made on concavo-convex surfaces of varying shapes, for it is not impossible that the difference between success and failure of a proposed flying machine will depend upon the sustaining effect between a plane surface and one properly curved to get a maximum of 'lift'."

3.3 Airfoil Terminology:

Several terms are used to describe airfoils (Dynamic Flight, 2002).

· Leading Edge = Forward edge of the airfoil

· Trailing Edge = Aft edge of the airfoil

· Chord = Line connecting the leading and trailing edge. Denotes the length of the airfoil

· Mean Camber Line = Line drawn half way between the upper and lower surface of the airfoil. Denotes the amount of curvature of the wing

· Point of Maximum Thickness = Thickest part of the wing expressed as a percentage of the chord

By altering each of the above features of an airfoil, the designer is able to adjust the performance of the wing so that it is suitable for its particular task. For example, a crop duster may have a thick, high camber wing that produces a large amount of lift at low speed. Alternatively, a jet would have a thin wing with minimal camber to allow it to cruise at high speeds.

Fig9. Airfoil Nomenclature

The shape of the airfoil is defined using the following geometrical parameters. The mean camber line or mean line is the locus of point’s midway between the upper and lower surfaces. Its shape depends on the thickness distribution along the chord. The thickness of an airfoil varies along the chord. It may be measured in either of two ways:

· Thickness measured perpendicular to the camber line. This is sometimes described as the "American convention"

· Thickness measured perpendicular to the chord line. This is sometimes described as the "British convention".

Fig10. Different Airfoil Thickness

3.2 How it Works?

· The basic principle behind an airfoil is described by Bernoulli’s theorem. Basically this states that total pressure is equal to static pressure (due to the weight of air above) plus dynamic pressure (due to the motion of air).

· Air that travels over the top surface of the airfoil has to travel faster and thus gains dynamic pressure. The subsequent loss of static pressure creates a pressure difference between the upper and lower surfaces that is called lift and opposes the weight of an aircraft (or thrust that opposes drag).

· As the angle of attack (the angle between the chord line and relative air flow) is increased, more lift is created. Once the critical angle of attack is reached (generally around 14 degrees) the airfoil will stall.

Fig11. Working principle

3.4 various types of an Airfoil:

There are mainly two types of airfoil.

· Symmetrical airfoil

· Nonsymmetrical airfoil

· Flat Bottom airfoil

· Supersonic airfoil

· Supercritical airfoil

SYMMETRICAL AIRFOIL:

· The symmetrical airfoil is distinguished by having identical lower and upper surfaces. The mean camber line or chord line are the same on a symmetrical airfoil, and it produce no lift at zero AOA. Most light aircrafts incorporate symmetrical airfoils in main rotor blades.

NONSYMMETRICAL AIRFOILS:

· The nonsymmetrical airfoil has different upper and lower surfaces, with a greater curvature of the airfoil above the chord line than below.

· The mean camber line and the chord line are different. The Nonsymmetrical airfoil can produce useful lift at zero AOA.

· The advantages are more lift production at a given AOA than a symmetrical design, an improved lift-to-drag ratio, and better stall characteristics.

· The disadvantages are Centre of pressure travel of up to 20 present of the chord line and greater production cost.

FLAT BOTTOM AIRFOIL:  

Flat bottoms are usually seen in trainer flights. They look extremely thin. Its bottom is flat and top is curved. Flat bottoms are speeding sensitive. They are similar to symmetrical airfoils. When power and speed is added it produces great lift

SUPERSONIC AIRFOIL: 

A supersonic airfoil is used to generate lift at supersonic speeds. Its need arises when an aircraft is operated consistently in supersonic range.

SUPERCRITICAL AIRFOIL:

A supercritical is designed to delay the drag in the transonic speed range are a few to name.  A supercritical is designed to delay the drag in the transonic speed range. They have a flat upper surface, a highly cambered aft and a greater leading edge radius.

Fig12. Types of airfoils

3.5 Advantages of an airfoil:

1. Cambered airfoils (asymmetric) are the kind which can generate a lift at a zero angle of attack

2. It can increase traction of a vehicle by creating a down force.

3. The angles of attack can be increased by symmetrical airfoil.

3.6 What causes a Turning moment?

  An airfoil has 3 forces. Lift, weight and drag. The lift is usually placed on the same spot as the weight, which is when the airfoil is stable and the plane has no problem, but usually when the lift is placed after weight force it produces instability in the airfoil which in turn produces a turning moment. This turning moment is compensated with the downward pushing force.

CHAPTER-4

CFX

4.1 Introduction to CFX:

ANSYS CFX and CFD-Post are supported in Release 12.0 of ANSYS Workbench, which is built on a new framework while leveraging the strength of ANSYS core applications, solvers, and associated tools with a new workflow and simulation project management capability.

Fig13. CFX of Airfoil

Multiple Configurations and Re-Meshing:

Multiple configurations can be defined to set up a continuous simulation during which the mesh topology or the physics changes during a simulation. For configurations with re-meshing can be defined using either ICEM CFD replay files, or a user defined mode where a script can be referenced to run external meshing processes or load alternative mesh files. This new functionality is defined in CFX-Pre, but also carries through with additional related

Options in the CFX-Solver and CFX-Solver Manager, as well as CFD-Post. In the CFX-Solver, this includes the ability to retain all possible data from input files used when interpolating from one mesh to another, and the possibility of using multiple initial values files for interpolation. For CFD-Post, this means the option to post-process multiple configurations as a continuous transient run, when applicable.

4.2 CFX Documentation:

There have been numerous incremental improvements to the organization and content of the user documentation, to improve clarity and make it more user-friendly. The default and context-sensitive help continues to use CHM (on Windows) and JAR (on UNIX) files, but the documentation in PDF files can be found directly from the Help menus of all ANSYS CFX components.

In addition to adapting the existing tutorials to take advantage of new features in ANSYS CFX and Release 12.0 of ANSYS Workbench, there are new tutorials for ANSYS CFX including:

· Equilibrium and non-equilibrium predictions of steam flowing through an axial turbine

· Modeling of flow in a gear pump using immersed solids.

· Calculation of a drop curve for caveating flow in a pump

· Flow in a spray dryer Modeling of a coal combustor

· Modeling of flow in a steam jet

· Using CFD-Post for analyzing flow in a mixing elbow

· Post processing of flow in a centrifugal compressor using CFD-Post

4.3 CFX Pre:

In addition to supporting the various new physics models, numerous further additions and improvements have been made in CFX-Pre.

Material:

An enhancement to CFX-Pre in this release is the mechanism for defining materials used in a simulation, which has been modified with the aim of making material definition more flexible. Users now define a name for a fluid and then assign a material to it. This permits, for example, the material used in a simulation to be changed more easily, and also helps facilitate the definition of reacting mixtures.

CEL and Expression Editing:

A new extension is available for defining logical expressions, including support for conditional statements. To aid in the definition of expressions, improvements have also been made to the presentation and syntax highlighting in the expression editor. In addition, existing expressions and variables are available on the right-click menu in the expressions widget in all editors.

Execution Control:

CFX-Pre can now optionally be used to edit or define most settings that traditionally have only been available in the Solver Manager. In addition, there are now options to launch the solver and monitor a run directly, rather than visiting the ‘define run' panel in the solver manager. Also now available at the same location is an option for node re-ordering, to change the order of the vertex data written for the solver and potentially improve the solver speed.

Mesh Manipulation:

All mesh transformations can now be dynamically previewed to give immediate visual feedback on the defined action before applying it. Transformations specific to turbo machinery meshes are also available in the general mesh transformation editor, as are scaling and reflecting of both the original mesh and copies of it.

Fig14. Meshing of Airfoil

4.4 CFX-Solver

This section highlights the new features supported in Release 12.0 of the CFX-Solver.

Efficiency and Accuracy:

A number of enhancements have been made in the core numerical algorithms, typically leading to improvements in efficiency and accuracy in the range of 10-20% or more in comparison with the previous release.

A new iteratively bounded high resolution advection/transient scheme has been introduced, primarily to allow more robust and accurate solution of turbulence quantities. Parallel improvements have been made in a number of areas. In the partitioning step, both the partition file sizes and partitioning times are dramatically reduced, and weighting factors have been made available for all partitioning methods.

For transient rotor-stator simulations, the defaults are changed in an effort to significantly reduce memory usage and CPU time during the parallel run. A new partitioning method for transient rotor-stator cases has also been added. This method reduces the number of overlap vertices required for the parallel run by creating banded partitions along the domain interface while using one of the regular partitioning methods in the interior.

Fig15. CFX solver of airfoil

Immersed Solids:

This new option is available to more easily capture the effect of complex geometry motion on flow: meshes of solid regions can be defined as immersed solids, and the solver tracks the overlap of these immersed solids with the background fluid mesh. In regions of overlap, the fluid flow is given the velocity of the solid, thereby having the presence and/or motion of the solid influence the flow. This method of capturing the interaction between fluids and solids does not involve re-meshing, so the fluid mesh must not conform to the solid boundaries. Therefore, the motion that can be simulated is unlimited, with the caveat that the model is not applicable in all situations (for example, it does not resolve near-wall turbulent conditions on the immersed solid).

Particle Tracking Extensions:

In addition to new particle injection options for points and hollow cones, Release 12.0 introduces a further option for primary break-up, the turbulence-induced atomization model of Huh & Gosman to account for turbulence effects in nozzles and improve predictions of the initial spray angle. Particle-wall interaction can be modeled with the Elsässer model (to include wall effects like roughness and temperature), and a quasi-static wall film model can be used to model the changes in heat and mass transfer due to the presence of a wall film. Other enhancements include further controls on particle termination, the ability to have coefficients of restitution be a function of time, and additional options to control the particle data available for post-processing. The stochastic particle-particle collision model, which extends the applicability.

Boundary Conditions:

A number of additional options are available in specifying boundary conditions. At wall boundaries, users now also have the options of explicitly specifying the wall shear or a finite slip condition. At supersonic inlets, the conditions can now be specified using total pressure, static pressure, total temperature, and flow direction. And at opening boundary conditions, a new option called “Entrainment” has replaced the previous options “Static Pressure for Entrainment” and “Opening Pressure for Entrainment”, with the user being able to additionally specify whether the pressure condition applied is total or static when the flow is into the domain.

Material Properties:

Extensions to the material property definitions include support for two-interval NASA Format polynomial coefficients for specific heat capacity in real gas models such as the Aungier Redlich Kwong model. Together with an additional library of materials in CFX-Pre, this allows combustion to be combined with real gas equations of state more seamlessly. To improve the predictions of gas viscosity of pure substances, the Interacting Sphere viscosity model is added and used by default in the real gas combustion library. In addition, a couple of options previously only available as beta features have been fully released: the standard Redlich Kwong and Peng Robinson equations of state, and the built-in non Newtonian dynamic viscosity models.

4.5 CFD-Post

This section highlights the new features supported in Release 12.0 of CFD-Post. A number of new post-processing features have been added, and the name of CFX-Post has changed to become CFD-Post, as it evolves to become the common post-processor for both ANSYS CFX and ANSYS FLUENT.

Comparison Mode:

A new file comparison mode is available, in which comparisons and difference plots can be made between different solutions on the same mesh (at different times in a transient solution or with two different results), and between different solutions on different meshes.

Feature Extraction:

A new visualization object is provided to allow users to easily identify vortex core regions in the flow based on a variety of different derived variables such as swirling strength, Eigen helicity, and others.

Iso Clips:

A new visualization object is the Iso clip, which greatly simplifies the creation of surfaces, planes, or lines bounded by any solution or geometry variable.

Turbo Post

Several new report templates have been added and include support for multiple components/blade rows. Machine types that can be used with these reports include axial compressors, centrifugal compressors, compressible and incompressible flow turbines, and pumps.

Chart Improvements

The user interface to create charts has been re-designed to become more intuitive and consistent with other common charting tools. In addition, histogram charts can now be created, and FFTs are available for spectral analysis.

Color Maps

A color map editor has been added to simplify the creation and modification of color maps, and permit transparency levels to be set.

Fig16. CFX-Post of Airfoil

4.6 History of CFX:

In 2003, ANSYS purchased CFX, one of the world’s leading fluid dynamics tools, from AEA Technology PLC for $21.7 million. ANSYS CFX software 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 solutions quickly and robustly. 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. The solver and its many physical models are wrapped in a modern, intuitive, and flexible GUI and user environment, with extensive capabilities for customization and automation using session files, scripting and a powerful expression language.

But ANSYS CFX is more than just a powerful CFD code. Integration into the ANSYS Workbench platform, provides superior bi-directional connections to all major CAD systems, powerful geometry modification and creation tools with ANSYS Design Modeler, advanced meshing technologies in ANSYS Meshing, and easy drag-and-drop transfer of data and results to share between applications. For example, a fluid flow solution can be used in the definition of a boundary load of a subsequent structural mechanics simulation. A native two-way connection to ANSYS structural mechanics products allows capture of even the most complex fluid–structure interaction (FSI) problems in the same easy-to-use environment, saving the need to purchase, administer or run third-party coupling software.

4.7 ANSYS FLUENT:

ANSYS FLUENT software contains the broad physical modeling capabilities needed to model flow, turbulence, heat transfer, and reactions for industrial applications ranging from air flow over an aircraft wing to combustion in a furnace, from bubble columns to oil platforms, from blood flow to semiconductor manufacturing, and from clean room design to wastewater treatment plants. Special models that give the software the ability to model in-cylinder combustion, aero acoustics, turbo machinery, and multiphase systems have served to broaden its reach.

Today, thousands of companies throughout the world benefit from the use of ANSYS FLUENT software as an integral part of their design and optimization phases of product development. Advanced solver technology provides fast, accurate CFD results, flexible moving and deforming meshes, and superior parallel scalability. User-defined functions allow the implementation of new user models and the extensive customization of existing ones. ANSYS FLUENT’s interactive solver set-up, solution, and post-processing make it easy to pause a calculation, examine results with integrated post-processing, change any setting, and then continue the calculation within a single application. Case and data files can also be read into ANSYS CFD-Post for further analysis with advanced post-processing tools and to compare results from different cases side-by-side.

The integration of ANSYS FLUENT into ANSYS Workbench, provides users with superior bi-directional connections to all major CAD systems, powerful geometry modification and creation with ANSYS Design Modeler and advanced meshing technologies in ANSYS Meshing. It also allows data and results to be shared between applications using an easy drag-and-drop transfer (e.g. to use a fluid flow solution in the definition of a boundary load of a subsequent structural mechanics simulation). Combine these benefits with the extensive range of physical modeling capabilities and fast, accurate CFD results that ANSYS FLUENT software has to offer and you have one of the most comprehensive software packages for CFD modeling available in the world today.

CHAPTER 5

PROBLEM DESCRIPTION

In the present study we explored the science of airfoil and the several ways to analyze and validate the airfoil. In current technology variform airfoil concept is very intriguing for which we have done this project on airfoils which changes its shape initially from NACA 23015 to final shape of wortmann as the mission profile of the UAV continues. We will analyze the aerodynamic characteristics of both the airfoil for several angles of attacks (AOA) in CFX and validate the results using Experimental results obtained

Initially, we selected the 2 best airfoils which have high lift and fuel efficiency. NACA 23015 is a moderate performance airfoil which has more drag as the angle of attack increases.

Wortmann FX60-126 is a fighter aircraft capable of flying at high speeds and high fuel efficiency. Combining these 2 models in the variform concept we can get an airfoil whose drag is reduced and capable of high efficiency. We have done the ICEM CFD of both the airfoils separately and calculated the flow analysis of each of them. After the meshing is done, the files are imported to FLUENT where solution in done and contours and graphs are calculated. Both the coefficient of lift (Cl) is compared and the best is selected.

CNC coding is done for both the airfoils by using coordinates and determined by using design and analysis of variform airfoil. We will import both the airfoils into Ansys fluent and calculating the Coefficient of lifts and drags values through graphs. By these values and coordinates we will do coding on both the airfoils NACA and WORTMANN and prepare a model of airfoils shapes. These airfoils are tested in wind tunnel; by calculate pressure values we can know the difference of both the airfoil by observing coefficient of lift and the best is selected.

In order to improve the fuel efficiency of an UAV, we propose using a wing that changes it shape during flight to minimize total mission drag the wing morphing approach is referred to a variform wing concept. The fuel would be stored in balloon-like bladders inside the wing structure. These bladders would shrink as the fuel is consumed. Therefore the wing will be of NACA 5 digit airfoil during takeoff with full fuel in the tank. During landing since the fuel is consumed to improve the lift we vary it to Wortmann FX 60-126

CHAPTER 6

STEPS INVOLVED IN PROCESS

6.1 COMPUTATIONAL FLUID ANALYSIS:

STEP 1: A.) By taking of the co-ordinates of both NACA 23015 and WORTMANN FX 60-126 from GOOGLE and calculated its CNC coding data.

The co-ordinates are here:-

NACA 23015

1.0000 0.0016 0.9500 0.0112 0.9000 0.0204 0.8000 0.0373 0.7000 0.0525 0.6000 0.0661 0.5000 0.0774 0.4000 0.0859 0.3000 0.0905 0.2500 0.0908 0.2000 0.0892 0.1500 0.0852 0.1000 0.0764 0.0750 0.0690 0.0500 0.0589 0.0250 0.0444 0.0125 0.0334 0.0000 0.0000 0.0125 -0.0154 0.0250 -0.0225 0.0500 -0.0304 0.0750 -0.0361 0.1000 -0.0409 0.1500 -0.0484 0.2000 -0.0541 0.2500 -0.0578 0.3000 -0.0596 0.4000 -0.0592 0.5000 -0.0550 0.6000 -0.0481 0.7000 -0.0391 0.8000 -0.0283 0.9000 -0.0159 0.9500 -0.0090 1.0000 -0.0016

Table 3.NACA coordinates

WORTMANN FX 60-126

0.0000000 0.0000000 0.0010700 0.0067500 0.0042800 0.0134900 0.0096100 0.0209600 0.0170400 0.0280200 0.0265300 0.0349300 0.0380600 0.0417400 0.0515600 0.0480800 0.0669900 0.0545700 0.0842700 0.0602100 0.1033200 0.0658500 0.1240800 0.0707700 0.1464500 0.0755500 0.1703300 0.0795800 0.1956200 0.0832700 0.2222100 0.0861500 0.2500000 0.0885900 0.2788600 0.0901900 0.3086600 0.0913000 0.3392800 0.0916000 0.3705900 0.0913800 0.4024500 0.0904100 0.4347400 0.0889300 0.4673000 0.0867900 0.5000000 0.0842500 0.5327000 0.0811800 0.5652600 0.0778100 0.5975500 0.0740200 0.6294100 0.0699400 0.6607200 0.0654900 0.6913400 0.0608200 0.7211400 0.0558900 0.7500000 0.0508400 0.7777900 0.0456700 0.8043800 0.0405500 0.8296700 0.0355200 0.8535500 0.0307000 0.8759200 0.0261100 0.8966800 0.0218100 0.9157300 0.0177700 0.9330100 0.0141200 0.9484400 0.0108400 0.9619400 0.0079800 0.9734700 0.0055400 0.9829600 0.0035300 0.9903900 0.0019800 0.9957200 0.0008800 0.9989300 0.0002400 1.0000000 0.0000000

Table 4: wortmann coordinates

B.) The sketch for aerofoil geometry and the excel chart shows the co-ordinates for a particular section. The profile of the particular cross section requires only two co-ordinates i.e. combination of X- and Yu- for upper part or X- and YL- for lower part. The upper part and lower part machining can be done separately by holding the job in a stationary fixture and ball nose cutter. The G code programming is a combination of X- and Yu- coordinates for upper part i.e. in each block the program X-co-ordinates   and corresponding Yu- co-ordinates are written with and block end.. For lower half the job is tilted by 180 degrees and the machining is carried out with lower co-ordinates. The YL-co-ordinate becomes YL +ve.

The CNC coding was based on Lathe and Milling machines.

x	Yc	Yt	d(Yc)/dx	θ	Xu	Yu	X L	Y L
0	0	0	0.1	0.099669	0	0	0	0
0.01	0.000988	0.019877	0.0975	0.097193	0.008071	0.02077	0.011929	-0.0188
0.02	0.00195	0.027531	0.095	0.094716	0.017396	0.029357	0.022604	-0.02546
0.03	0.002888	0.033135	0.0925	0.092238	0.026948	0.035882	0.033052	-0.03011
0.04	0.0038	0.037657	0.09	0.089758	0.036625	0.041305	0.043375	-0.03371
0.05	0.004688	0.041471	0.0875	0.087278	0.046385	0.046001	0.053615	-0.03663
0.06	0.00555	0.044772	0.085	0.084796	0.056208	0.050161	0.063792	-0.03906
0.07	0.006388	0.047673	0.0825	0.082314	0.06608	0.053899	0.07392	-0.04112
0.08	0.0072	0.050251	0.08	0.07983	0.075993	0.057291	0.084007	-0.04289
0.09	0.007988	0.052558	0.0775	0.077345	0.085939	0.060388	0.094061	-0.04441
0.1	0.00875	0.054632	0.075	0.07486	0.095914	0.063229	0.104086	-0.04573
0.11	0.009488	0.056504	0.0725	0.072373	0.105914	0.065843	0.114086	-0.04687
0.12	0.0102	0.058195	0.07	0.069886	0.115936	0.068253	0.124064	-0.04785
0.13	0.010888	0.059725	0.0675	0.067398	0.125978	0.070477	0.134022	-0.0487
0.14	0.01155	0.061109	0.065	0.064909	0.136036	0.07253	0.143964	-0.04943
0.15	0.012188	0.062359	0.0625	0.062419	0.14611	0.074426	0.15389	-0.05005
0.16	0.0128	0.063487	0.06	0.059928	0.156198	0.076173	0.163802	-0.05057
0.17	0.013388	0.064501	0.0575	0.057437	0.166297	0.077782	0.173703	-0.05101
0.18	0.01395	0.065409	0.055	0.054945	0.176408	0.07926	0.183592	-0.05136
0.19	0.014488	0.066219	0.0525	0.052452	0.186528	0.080615	0.193472	-0.05164
0.2	0.015	0.066936	0.05	0.049958	0.196657	0.081852	0.203343	-0.05185
0.21	0.015488	0.067566	0.0475	0.047464	0.206794	0.082977	0.213206	-0.052
0.22	0.01595	0.068114	0.045	0.04497	0.216938	0.083995	0.223062	-0.0521
0.23	0.016388	0.068584	0.0425	0.042474	0.227088	0.08491	0.232912	-0.05214
0.24	0.0168	0.068981	0.04	0.039979	0.237243	0.085726	0.242757	-0.05213
0.25	0.017188	0.069309	0.0375	0.037482	0.247403	0.086448	0.252597	-0.05207
0.26	0.01755	0.06957	0.035	0.034986	0.257567	0.087077	0.262433	-0.05198
0.27	0.017888	0.069767	0.0325	0.032489	0.267734	0.087618	0.272266	-0.05184
0.28	0.0182	0.069904	0.03	0.029991	0.277904	0.088073	0.282096	-0.05167
0.29	0.018488	0.069984	0.0275	0.027493	0.288076	0.088445	0.291924	-0.05147
0.3	0.01875	0.070008	0.025	0.024995	0.29825	0.088736	0.30175	-0.05124
0.31	0.018988	0.06998	0.0225	0.022496	0.308426	0.088949	0.311574	-0.05097
0.32	0.0192	0.0699	0.02	0.019997	0.318602	0.089086	0.321398	-0.05069
0.33	0.019388	0.069773	0.0175	0.017498	0.328779	0.089149	0.331221	-0.05037
0.34	0.01955	0.069598	0.015	0.014999	0.338956	0.08914	0.341044	-0.05004
0.35	0.019688	0.069378	0.0125	0.012499	0.349133	0.089061	0.350867	-0.04969
0.36	0.0198	0.069116	0.01	0.01	0.359309	0.088912	0.360691	-0.04931
0.37	0.019888	0.068812	0.0075	0.0075	0.369484	0.088697	0.370516	-0.04892
0.38	0.01995	0.068467	0.005	0.005	0.379658	0.088416	0.380342	-0.04852
0.39	0.019988	0.068084	0.0025	0.0025	0.38983	0.088072	0.39017	-0.0481
0.4	0.02	0.067664	0	0	0.4	0.087664	0.4	-0.04766
0.41	0.019994	0.067208	-0.00111	-0.00111	0.410075	0.087202	0.409925	-0.04721
0.42	0.019978	0.066717	-0.00222	-0.00222	0.420148	0.086695	0.419852	-0.04674
0.43	0.01995	0.066193	-0.00333	-0.00333	0.430221	0.086143	0.429779	-0.04624
0.44	0.019911	0.065636	-0.00444	-0.00444	0.440292	0.085547	0.439708	-0.04572
0.45	0.019861	0.065048	-0.00556	-0.00556	0.450361	0.084908	0.449639	-0.04519
0.46	0.0198	0.06443	-0.00667	-0.00667	0.46043	0.084229	0.45957	-0.04463
0.47	0.019728	0.063782	-0.00778	-0.00778	0.470496	0.083508	0.469504	-0.04405
0.48	0.019644	0.063106	-0.00889	-0.00889	0.480561	0.082748	0.479439	-0.04346
0.49	0.01955	0.062402	-0.01	-0.01	0.490624	0.081949	0.489376	-0.04285
0.5	0.019444	0.061672	-0.01111	-0.01111	0.500685	0.081112	0.499315	-0.04222
0.51	0.019328	0.060915	-0.01222	-0.01222	0.510744	0.080238	0.509256	-0.04158
0.52	0.0192	0.060134	-0.01333	-0.01333	0.520802	0.079328	0.519198	-0.04093
0.53	0.019061	0.059327	-0.01444	-0.01444	0.530857	0.078382	0.529143	-0.04026
0.54	0.018911	0.058497	-0.01556	-0.01555	0.54091	0.077401	0.53909	-0.03958
0.55	0.01875	0.057644	-0.01667	-0.01667	0.550961	0.076386	0.549039	-0.03889
0.56	0.018578	0.056768	-0.01778	-0.01778	0.561009	0.075337	0.558991	-0.03818
0.57	0.018394	0.05587	-0.01889	-0.01889	0.571055	0.074254	0.568945	-0.03747
0.58	0.0182	0.05495	-0.02	-0.02	0.581099	0.073139	0.578901	-0.03674
0.59	0.017994	0.05401	-0.02111	-0.02111	0.59114	0.071992	0.58886	-0.036
0.6	0.017778	0.053049	-0.02222	-0.02222	0.601179	0.070813	0.598821	-0.03526
0.61	0.01755	0.052068	-0.02333	-0.02333	0.611215	0.069603	0.608785	-0.0345
0.62	0.017311	0.051067	-0.02444	-0.02444	0.621248	0.068363	0.618752	-0.03374
0.63	0.017061	0.050047	-0.02556	-0.02555	0.631279	0.067092	0.628721	-0.03297
0.64	0.0168	0.049008	-0.02667	-0.02666	0.641306	0.06579	0.638694	-0.03219
0.65	0.016528	0.04795	-0.02778	-0.02777	0.651331	0.064459	0.648669	-0.0314
0.66	0.016244	0.046874	-0.02889	-0.02888	0.661354	0.063099	0.658646	-0.03061
0.67	0.01595	0.04578	-0.03	-0.02999	0.671373	0.06171	0.668627	-0.02981
0.68	0.015644	0.044668	-0.03111	-0.0311	0.681389	0.060291	0.678611	-0.029
0.69	0.015328	0.043539	-0.03222	-0.03221	0.691402	0.058844	0.688598	-0.02819
0.7	0.015	0.042393	-0.03333	-0.03332	0.701412	0.057369	0.698588	-0.02737
0.71	0.014661	0.041229	-0.03444	-0.03443	0.711419	0.055866	0.708581	-0.02654
0.72	0.014311	0.040048	-0.03556	-0.03554	0.721423	0.054334	0.718577	-0.02571
0.73	0.01395	0.038851	-0.03667	-0.03665	0.731424	0.052774	0.728576	-0.02487
0.74	0.013578	0.037636	-0.03778	-0.03776	0.741421	0.051187	0.738579	-0.02403
0.75	0.013194	0.036405	-0.03889	-0.03887	0.751415	0.049572	0.748585	-0.02318
0.76	0.0128	0.035157	-0.04	-0.03998	0.761405	0.047929	0.758595	-0.02233
0.77	0.012394	0.033893	-0.04111	-0.04109	0.771392	0.046259	0.768608	-0.02147
0.78	0.011978	0.032612	-0.04222	-0.0422	0.781376	0.044561	0.778624	-0.02061
0.79	0.01155	0.031315	-0.04333	-0.04331	0.791356	0.042836	0.788644	-0.01974
0.8	0.011111	0.030001	-0.04444	-0.04442	0.801332	0.041082	0.798668	-0.01886
0.81	0.010661	0.02867	-0.04556	-0.04552	0.811305	0.039302	0.808695	-0.01798
0.82	0.0102	0.027323	-0.04667	-0.04663	0.821274	0.037493	0.818726	-0.01709
0.83	0.009728	0.025959	-0.0477	-0.04774	0.831239	0.035657	0.828761	-0.0162
0.84	0.009244	0.024577	-0.04889	-0.04885	0.8412	0.033793	0.8388	-0.0153
0.85	0.00875	0.023179	-0.05	-0.04996	0.851158	0.0319	0.848842	-0.0144
0.86	0.008244	0.021764	-0.05111	-0.05107	0.861111	0.02998	0.858889	-0.01349
0.87	0.007728	0.020331	-0.05222	-0.05217	0.87106	0.028031	0.86894	-0.01258
0.88	0.0072	0.01888	-0.05333	-0.05328	0.881006	0.026054	0.878994	-0.01165
0.89	0.006661	0.017412	-0.05444	-0.05439	0.890947	0.024047	0.889053	-0.01073
0.9	0.006111	0.015926	-0.05556	-0.0555	0.900883	0.022012	0.899117	-0.00979
0.91	0.00555	0.014421	-0.05667	-0.05661	0.910816	0.019948	0.909184	-0.00885
0.92	0.004978	0.012897	-0.05778	-0.05771	0.920744	0.017854	0.919256	-0.0079
0.93	0.004394	0.011355	-0.05889	-0.05882	0.930668	0.01573	0.929332	-0.00694
0.94	0.0038	0.009794	-0.06	-0.05993	0.940587	0.013576	0.939413	-0.00598
0.95	0.003194	0.008213	-0.06111	-0.06104	0.950501	0.011392	0.949499	-0.005
0.96	0.002578	0.006611	-0.06222	-0.06214	0.960411	0.009177	0.959589	-0.00402
0.97	0.00195	0.00499	-0.06333	-0.06325	0.970315	0.00693	0.969685	-0.00303
0.98	0.001311	0.003348	-0.06444	-0.06436	0.980215	0.004652	0.979785	-0.00203
0.99	0.000661	0.001685	-0.06556	-0.06546	0.99011	0.002342	0.98989	-0.00102
1	0	0	-0.06667	-0.06657	1	0	1	0

Table 5. CNC coding

STEP 2: CFX

In Ansys basing on the coordinates, we have performed CFX for both the airfoils i.e. NACA 23015 and WORTMANN FX 60-126 and given boundary conditions for all the sides i.e. Top, Bottom and Sides and also created surface to it.

STEP 3: MESHING OPERATION

Then calculated meshing operation for both the airfoils by satisfying its boundary condition and their parameters and the file is saved.

Fig17. Meshing of NACA 23015

Fig18. Meshing of WORTMANN FX 60-126

STEP 4: FLUENT

Now the saved files of both the airfoils will be imported to the fluent solver separately, read the file and changes the parameters for calculating the output. Such as,

· Fluid Type: Air

· Flow conditions: Pressure Top

Pressure Bottom

· Boundary conditions: Velocity-1 Mach i.e. 340.29m/s

· Output Plots: Lift, Drag and residual graphs.

· Calculations are based on pressure type.

The outputs are the following from the fluent are:

NACA 23015 AIRFOIL

Graph1: Coefficient of lift of NACA

Graph2. Coefficient of drag of NACA

WORTMANN FX 60-126

Graph3. Coefficient of lift of WORTMANN

Graph4. Coefficient of drag of WORTMANN

STEP 5: THEORETICAL RESULTS

From the above graphs, results that are shown for both the airfoils are theoretical values. And by this we can say that wortmann airfoil is having more coefficient of lift compared to naca airfoil theoretically which is formed during take-off and landing of our variform airfoil.

NACA 23015

· Coefficient of Lift (Cl) =0.72 (Low)

· Coefficient of Drag (Cd) =2.250 (more)

WORTMANN FX 60-126

· Coefficient of Lift (Cl) =2.20e-02 (more)

· Coefficient of Drag (Cd) =0.60e+00 (low)

6.2 MANFACTURING OF MODEL

By using CNC coding and taking certain measurements i.e. Length, Width, Thickness, Angle of attack, Mach number. We have modelled the airfoils by using Lathe and Milling machines.

The Model is prepared with some materials like wood, copper lids, wood powder, Plasters, sanding papers, fevicol and copper tubes and the figures are shown below.

Fig19. Top view of NACA

Fig20. Side view of NACA

Fig21. Top view of WORTMANN

Fig22. Airfoil with copper wires

6.3 EXPERIMENTAL ANALYSIS

The two types of airfoils are tested in Wind Tunnel which is Low speed capable of 1.8 Mach for getting pressure values. We have tested this in our college lab and the figures are shown below.

Fig23. Wind Tunnel

Fig24. Testing in wind tunnel

Fig25. Pressure tubes

Fig26. Noting Pr values

Fig27. Noting Pr values

Fig28. Noting Pr head values

From these we got h1 and h2 pressure head values at zero degrees angle of attack and lift and drag values, by using these values we have calculated Cl and Cd (Coefficient of Lift, Coefficient of drag and also Maximum Lift and Maximum Drag values) from both of the airfoils.

These are the calculations and values which are shown below

Table6. Calculations of NACA Airfoil

Table7. Calculation of WORTMANN Airfoil

By using this we plotted graph, in which we took Cl at x-axis and Cd at y-axis.

Graph5. Cl/Cd Ratio of NACA 23015

Graph6. Cl/Cd ratio of WORTMANN FX 60-12

By this we got Practical results as,

NACA 23015 AIRFOIL

· Coefficient of Lift (Cl) =1.1

· Coefficient of Drag (Cd) =3.50

WORTMANN FX 60-126

· Coefficient of Lift (Cl) =1.22

· Coefficient of Drag (Cd) =0.280

RESULTS AND DISCUSSIONS

As we are doing variform concept by comparing these theoretical and practical values of Cl and Cd values for both the airfoils, we can say that coefficient of lift and coefficient of drag are more for NACA 23015 and WORTMANN FX 60-126 has more coefficient of lift but reduced by drag (Cd). So the range of variform airfoil is increased by maximum L/D ratio.

And we have a variform airfoil graph by comparing both the airfoils and are shown below.

Graph7. Variform Cl/Cd ratio

CONCLUSION AND FUTURE SCOPE

Even with no optimization done on the variform wing in this example case, the range and endurance increase was within the expected range. The increase in both distance and time was quite significant and would greatly benefit many UAV missions.

The example case shows that the variform concept readily extends missions and therefore, a more complete model and analysis should be completed. The work illustrated in this paper lays the foundation for further progress on this topic and shows that the concept is worth more investigation.

Further work should be directed towards developing a structural model and finding suitable materials that will allow for this deformation with the needed stiffness. Building the full coupled MDO analysis and parameterizing the outer airfoil shape and inner bladder shape also need further study.

The majority of the research remaining for this morphing wing design will be in developing the rest of the software model for the complete system analysis and then employing optimization methods that account for uncertainty8 to produce a robust variform wing with maximized range. Once all the systems have been modeled and integrated, this MDO problem could also serve as a test problem for comparing different optimization method