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A MINI PROJECT REPORT

ON

DESIGN OF A RC AIRCRAFT

PRESENTED BY

ROOPAL RAJ

1109140080

SANDEEP KR. MISHRA

1109140085

SHREYA SRIVASTAVA

1109140098

2013-14

MINI PROJECT CO-ORDINATOR

NEERAJ VERMA

JSS MAHAVIDYAPEETHA

JSS ACADEMY OF TECHNICAL EDUCATION, NOIDA

DEPARTMENT OF MECHANICAL ENGINEERING Session 2013-2014

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JSS MAHAVIDYAPEETHA

JSS ACADEMY OF TECHNICAL EDUCATION

DEPARTMENT OF MECHANICAL ENGINEERING

CERTIFICATE

This is to certify that Mini-Project Report entitled DESIGN OF A RC

AIRCRAFT which is submitted by the students Roopal Raj, Sandeep Kr.

Mishra, Shreya Srivastava in Department of Mechanical Engineering JSS

Academy of Technical Education, Noida during their 6th

semester B. Tech. degree

course of U. P. Technical University, is a record of the candidates work carried out

by them under my supervision.

Date:

Place: Noida Mini Project Co-ordinator

(NEERAJ VERMA)

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ABSTRACT

A radio-controlled (model) aircraft (often called RC aircraft or RC plane) is controlled remotely

by a hand-held transmitter and a receiver within the craft. The receiver controls the

corresponding servos that move the control surfaces based on the position of joysticks on the

transmitter, which in turn affect the orientation of the plane.

Flying RC aircraft as a hobby has been growing worldwide with the advent of more efficient

motors (both electric and miniature internal combustion and jet engines), lighter and more

powerful batteries and less expensive radio systems.

Scientific, government and military organizations are also utilizing RC aircraft for experiments,

gathering weather readings, aerodynamic modeling and testing, and even using them as drones or

spy planes.

The RC fixed winged Aircraft project is based on designing a light weight glider electronically

controlled using a remote control with operating frequency of 2.4 GHz. The project concentrated

on designing the mechanical part of the model. It was the stepwise execution of procedure to

synthesize the glider. This project did not concentrated on electronics components as the

components used are readily available in markets and need not be programmed by the users.

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TABLE OF CONTENTS

Introduction . 6

Airfoil Selection... 7

Control Surfaces.. 10

Stability Concepts .. 11

Design Methodology . 14

Electronics components . 19

Conclusion. 20

References . 22

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LIST OF FIGURES

Figure 1 Airfoil terminology (Page No. 7)

Figure 2 Formation of boundary layer (Page No. 9)

Figure 3 Empennage. (Page No. 10)

Figure 4 Motion. (Page No. 11)

Figure 5 - Neutral point (Page No. 13)

Figure 6 - Angle of attack (Page No. 14)

Figure 7 - Design Layout (Page No. 14)

Figure 8 NACA Profile (Page No. 15)

Figure 9 Flow on JAVA Applet (Page No. 16)

Figure 10 CFD on Ansys (Page No. 17)

Figure 11 - Foil Sim Results (Page No. 18)

Figure 12 Auto Desk Velocity Field (Page No. 19)

Figure 13 Auto Desk Pressure Field (Page No. 20)

Figure 14 Rendered View (Page No. 21)

Figure 15 Actual Image(Page No. 21)

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INTRODUCTION:

A fixed-wing aircraft's wings, horizontal, and vertical stabilizers are built with airfoil-shaped

cross-sections, as are helicopter rotor blades. Swimming and flying creatures and even many

plants and sessile organisms employ airfoils/hydrofoils: common examples being bird wings, the

bodies of fish. An aircraft is used for sail purpose, war purpose and also used as cargo carrier.

Aircraft moves with the help of generated aerodynamic force (called lift) perpendicular to the

motion of the aircraft. Aircraft experiences an opposite force to the direction of motion called as

drag force. Designers always try to keep maximum lift and minimum drag in any kind of aircraft.

A RC aircraft has all the dynamic characteristics which are present in actual aircraft. An aircraft

has some control surfaces through which it can be maneuver in air. Designing of RC Aircraft is

the best way to understand the property of any particular airplane configuration or the effect of

change in control surfaces or other parameters on its flying characteristic.

Airfoil design is a major facet of aerodynamics. Various airfoils serve different flight regimes.

Asymmetric airfoils can generate lift at zero angle of attack, while a symmetric airfoil may better

suit frequent inverted flight as in an aerobatic airplane. In the region of the ailerons and near

a wingtip a symmetric airfoil can be used to increase the range of angles of attack to avoid spin-

stall. Thus a large range of parameters on flight if an aircraft depends. Subsonic, transonic,

supersonic planes possess different geometry and different design methodology. RC Aircrafts

can be used to carry some mass of small value or also use by defense authority by applying a spy

camera on it.

In this report the whole design methodology are explained including several analysis of wing and

analysis of flow of air around the aircraft.

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Airfoil selection:

An airfoil is the 2D cross-section shape of the wing, which creates sufficient lift with minimal

drag. The geometry of the airfoil is described with a variety of terms. A key characteristic of an

airfoil is its chord. We thus define the following concepts:

The leading edge is the point at the front of the airfoil that has maximum curvature.

The trailing edge is defined similarly as the point of maximum curvature at the rear of the

airfoil.

The chord line is a straight line connecting the leading and trailing edges of the airfoil.

The chord length, or simply chord, , is the length of the chord line and is the characteristic

dimension of the airfoil section

The mean camber line is the locus of points midway between the upper and lower surfaces.

Its exact shape depends on how the thickness is defined.

Two key parameters to describe an airfoils shape are its maximum thickness (expressed as a

percentage of the chord), and the location of the maximum thickness point (also expressed as a

percentage of the chord).

Fig. 1

Finally, important concepts used to describe the airfoils behavior when moving through a fluid

are

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The aerodynamic center, a point lies on the chord-wise length about which the pitching

moment is independent of the lift coefficient and the angle of attack.

Types of an Airfoil:

Airfoil is classify on basis of following parameters-

On the basis of shape

Symmetrical Airfoil - It is symmetrical about its chord i.e. camber and chord lines are

same in case of symmetrical airfoil. It is use in aircraft having low Mach number. The

stability of such kind of airfoil is good.

Asymmetrical Airfoil - It is asymmetrical about its chord i.e. camber and chord lines are

not same in case of asymmetrical airfoil. The maneuverability of aircraft having this kind

of airfoil is more.

On the basis of speed

Low speed Airfoils which are suitable for low speed aircraft comes under this section.

Ex. Symmetrical airfoil.

High speed airfoils which are suitable for high speed aircraft comes under this section.

In any fluid airfoil mainly experience two types of forces force of lift and force of drag. An

airfoil-shaped body moved through a fluid produces an aerodynamic force. The component of

this force perpendicular to the direction of motion is called lift. The component parallel to the

direction of motion is called drag. Subsonic flight airfoils have a characteristic shape with a

rounded leading edge, followed by a sharp trailing edge, often with asymmetric camber. Foils of

similar function designed with water as the working fluid are called hydrofoils.

The lift on an airfoil is primarily the result of its angle of attack and shape. When oriented at a

suitable angle, the airfoil deflects the oncoming air, resulting in a force on the airfoil in the

direction opposite to the deflection. This force is known as aerodynamic force and can be

resolved into two components: Lift and drag. Most foil shapes require a positive angle of attack

to generate lift, but cambered airfoils can generate lift at zero angle of attack. This "turning" of

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the air in the vicinity of the airfoil creates curved streamlines which results in lower pressure on

one side and higher pressure on the other. This pressure difference is accompanied by a velocity

difference, via Bernoulli's principle, so the resulting flow field about the airfoil has a higher

average velocity on the upper surface than on the lower surface. The lift force can be related

directly to the average top/bottom velocity difference without computing the pressure by using

the concept of circulation and the Kutta-Joukowski theorem.

Fig.2

Drag Coefficient:

The drag coefficient cd is defined as:

where:

is the drag force, which is by definition the force component in the direction of the

flow velocity,

is the mass density of the fluid,

is the speed of the object relative to the fluid and

is the reference area.

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The reference area depends on what type of drag coefficient is being measured. For automobiles

and many other objects, the reference area is the projected frontal area of the vehicle. This may

not necessarily be the cross sectional area of the vehicle, depending on where the cross section is

taken. For example, for a sphere (note this is not the surface area = ).

For airfoils, the reference area is the plan form area. Since this tends to be a rather large area

compared to the projected frontal area, the resulting drag coefficients tend to be low: much lower

than for a car with the same drag and frontal area, and at the same speed.

Lift Coefficient:

The lift coefficient CL is defined by,

Where:

is the lift force, is fluid density, is true airspeed, is plan form area and is the

fluid dynamic pressure. The lift coefficient can be approximated using the lifting-line theory,

numerically calculated or measured in a wind tunnel test of a complete aircraft configuration.

Control Surfaces:

Empennage:

The empennage also known as the tail or tail

assembly, of most aircraft gives stability to the

aircraft, in a similar way to the feathers on an arrow.

Most aircraft feature empennage

Incorporating vertical and horizontal stabilizing

surfaces which stabilize the flight dynamics of

pitch and yaw, as well as housing control surfaces. In

Fig. 3

spite of effective control surfaces, many early aircraft that lacked stabilizing empennage were

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virtually unflyable. Today, only a few (often relatively unstable) heavier than air aircraft are able

to fly without empennage.

Ailerons:

These are the control surfaces situated at the trailing edge of the wing to give it the roll motion.

Ailerons move in opposite direction to each other i.e they have differential action.

Rudder:

It is a control surface unites with vertical stabilizer to control the yaw motion of the aircraft.

Elevator:

It is a control surface unites with horizontal stabilizer to control the pitching motion of the

aircraft.

Stability Concept:

The aircraft's response to momentary disturbance is associated with its inherent degree of

stability built in by the designer, in each of the three axes, and occurring without any reaction

from the pilot. There is another condition affecting flight, which is the aircraft's state of trim or

equilibrium (where the net sum of all forces equals

zero). Some aircraft can be trimmed by the pilot to

fly 'hands off' for straight and level flight, for climb

or for descent. Free flight models generally have to

rely on the state of trim built in by the designer and

adjusted by the rigger, while the remote controlled

models have some form of trim devices which are

adjustable during the flight.

An aircraft's stability is expressed in relation to each axis: Fig. 4

- lateral stability (stability in roll), directional stability (stability in yaw)

- longitudinal stability (stability in pitch).

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Lateral and directional stabilities are inter-dependent. Stability may be defined as follows:

- Positive stability: tends to return to original condition after a disturbance.

- Negative stability: tends to increase the disturbance.

- Neutral stability: remains at the new condition.

- Static stability: refers to the aircraft's initial response to a disturbance.

A statically unstable aircraft will uniformly depart from a condition of equilibrium.

- Dynamic stability: refers to the aircraft's ability to damp out oscillations, which

depends on how fast or how slow it responds to a disturbance.

A dynamically unstable aircraft will (after a disturbance) start oscillating with increasing

amplitude. A dynamically neutrally stable aircraft will continue oscillating after a disturbance

but the amplitude of the oscillations will not change. So, a statically stable aircraft may be

dynamically unstable. Dynamic instability may be prevented by an even distribution of weight

inside the fuselage, avoiding too much weight concentration at the extremities or at the CG.

Also, control surfaces' max throws may affect the flight stability, since a too much control throw

may cause instability, e.g. Pilot Induced Oscillations (PIO). Static stability is proportional to the

stabilizer area and the tail moment. You get double static stability if you double the tail area or

double the tail moment. Dynamic stability is also proportional to the stabilizer area but increases

with the square of the tail moment, which means that you get four times the dynamic stability

if you double the tail arm length. However, making the tail arm longer or increasing the stabilizer

area will move the mass of the aircraft towards the rear, which may also mean the need to make

the nose longer in order to minimize the weight required to balance the aircraft...

A totally stable aircraft will return, more or less immediately, to its trimmed state

without pilot intervention.

Lateral stability is achieved through dihedral, sweepback, keel effect and proper distribution of

weight. The dihedral angle is the angle that each wing makes with the horizontal (see

Wing Geometry). If a disturbance causes one wing to drop, the lower wing will receive more lift

and the aircraft will roll back into the horizontal level. A sweptback wing is one in which the

leading edge slopes backward. When a disturbance causes an aircraft with sweepback to slip or

drop a wing, the low wing presents its leading edge at an angle more perpendicular to the

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relative airflow. As a result, the low wing acquires more lift and rises, restoring the aircraft to its

original flight attitude.

Longitudinal stability depends on the location of the center of gravity, the stabilizer area and how

far the stabilizer is placed from the main wing. Most aircraft would be completely unstable

without the horizontal stabilizer. Non-symmetrical cambered airfoils have a higher lift

coefficient, but they also have a negative pitching moment (Cm) tending to pitch nose-down, and

thus being statically unstable, which requires the counter moment produced by the horizontal

stabilizer to get adequate longitudinal stability. The stabilizer provides the same function in

longitudinal stability as the fin does in directional stability. Symmetrical (zero camber) airfoils

have normally a zero pitching moment, resulting in neutral stability, which means the aircraft

goes wherever you point it. Reflexed airfoils (with trailing edge bent up) have a positive pitching

moment making them naturally stable; they are often used with flying wings (without the

horizontal stabilizer). It is of crucial importance that the aircraft's Centre of Gravity (CG) is

located at the right point, so that a stable and controllable flight can be achieved. In order to

achieve a good longitudinal stability, the CG should be ahead of the Neutral Point (NP), which is

the Aerodynamic Centre of the whole aircraft. NP is the position through which all the net lift

increments act for a change in angle of attack.

The major contributors are the main wing,

stabilizer surfaces and fuselage. The bigger the

stabilizer area in relationship to the wing area and

the longer the tail moment arm relative to the

wing chord, the farther aft the NP will be and the

farther aft the CG may be, provided it's kept

ahead of the NP for stability. The simplest way

of locating the aircraft's NP is by using the areas

of the two horizontal lifting surfaces (main wing

and stab) and locate the NP proportionately along the distance between the main wing's AC point

and the stab's AC point. NP distance to the main wing's AC point would be:

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D = L (stab area) / (main wing area + stab area) as shown on the picture below:

For symmetric airfoils the Absolute angle of attack is equal to the Geometric angle of attack

whereas for asymmetric (cambered) airfoils these two angles are different, since

these airfoils still produce lift at zero Geometric Angle of Attack as shown below.

Methodology:

For designing of airfoil we select the following layout of process.

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Now here is the method to design of an airfoil for a RC plane having mass = 1kg and wing span

of 1 meter. For the given specification of the aero plane we started the designing process from

JAVA FOIL, software from NASA to get the exact profile of a particular airfoil.

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Fig.8

In the above there are 61 points which forms a particular NACA0012 profile.

1. Once we get the NACA0012 profile we analyses its velocity flow field and pressure with the help of JAVA FOIL and ANSYS FLUENT as shown in given

figure.

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Fig.9

The above figure shows the result of pressure field around NACA0012 at 4 angle of attack,

Reynolds no. = 4000 and mach no. = .008 Corresponding values of Cl and Cd is tabulated. The Cp ( center of pressure ) can be observed

close to the leading edge.

JAVA FOIL uses several traditional methods for airfoil analysis. The following two methods

build the backbone of the program:

The potential flow analysis is done with a higher order panel method (linear varying

vortices distribution). Taking a set of airfoil coordinates, it calculates the local, in viscid

flow velocity along the surface of the airfoil for any desired angle of attack.

The boundary layer analysis module steps along the upper and the lower surfaces of the

airfoil, starting at the stagnation point. It solves a set of differential equations to find the

various boundary layer parameters. It is a so called integral method. The equations and

criteria for transition and separation are based on the procedures described by Eppler.

Compared with CalcFoil, this module has been completely rewritten and cleaned up.

A standard compressibility correction according to Karman and Tsien has been implemented to

take mild Mach number effects into account.

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As long as the flow stays subsonic (V below V* in the Velocity diagram), the results should be

fairly accurate. Usually this means Mach numbers between zero and 0.5.

The next step is to find the velocity field lines in Ansys fluent

Fig.10

ANSYS Fluent software is an integral part of the design and optimization of the given

configuration of airfoil. It has advanced solver technology which provides accurate CFD results,

flexible moving and deforming meshes and superior parallel scalability.

From the above analysis results we can conclude about the stagnation point and velocity of air

around NACA0012. Once we get the pressure field flow simulation results and velocity flow

results we can calculate the value of lift and drag force apply on the airplane.

Here it is clearly visible that Cp lies very close to c/4 distance from the leading edge.

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In order to achieve the final conclusion from FOIL-SIM III, this is software to calculate the value

of forces acting on an aerodynamic body. With this software you can investigate how an

aircraft wing produces lift and drag by changing the values of different factors that affect lift and

the factors that affect drag. FoilSim III is the latest (July 2010) version of the FoilSim family of

interactive simulations. The different versions of FoilSim require different levels of knowledge

of aerodynamics, experience with the package, and familiarity with computer technology.

Fig.11

The above image shows the final outcomes that are obtain from FOIL SIM III. The span length is

= 1.06m, Area require to lift the plane = 0.188m and Chord length is 0.178m.

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To get the dimension of empennage we opted some thumb rules that stabilize the relation

between the dimension of wing and horizontal and vertical stabilizer. From mentioned method

we get the area of horizontal and vertical stabilizer as 0.94 m and 0.47 m respectively.

Now in the direction of getting the fuselage length we use the formula given in the section

Stability concepts as mentioned above. The length of fuselage comes out to be 0.70m and to

decide its shape we finally considered two important points one is flow simulation and other is

ease of fabrication.

Here is the result of velocity flow simulation on Autodesk wind tunnel tester :

Fig.12

In ordered to get pressure distribution at the contour points of our aircraft we did a pressure flow

analysis on the same software as mention above. The software which we use for flow simulation

was Flow Design (formerly Project Falcon), which is part of the Digital Prototyping solution.

Simulate airflow and wind tunnel testing around buildings, vehicles, outdoor equipment, aircrafts

or any other virtual structure. Fast feedback and intuitive controls help from which we got deep

design insight.

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Fig.13

Electronic Components:

One, 6 channel 4 Ghz Receiver and transmitter.

Four servos of having maximum 60 rotation.

One brushless motor of 700gms thrust capacity.

One Electronic speed controller of maximum 20Amp capacity current.

One Lithium polymer battery of capacity 11.1 volt, 3 cells and 1500mah

One 8*4 plastic propeller.

Conclusion:

From above methodology and concepts we finally successfully designed a RC Aircraft which has

following characteristics:

Maximum velocity of 36 kmph.

Wight = 1kg.

Wing loading capacity of 650gms.

Range of remote is about 1.5 km.

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Here is the final rendered view on SOLIDWORKS and original picture of the plane.

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References:

John J. Bertin Aerodynamics for Engineers Fourth Edition U.S Air force Academy.

John D. Anderson Introduction to flight Dynamics Third edition.

Wikipedia

Softwares used :

- JAVA Applet

- Ansys

- CATIA

- Airfoil sim III

-AutoDesk Flow Design

-Solidworks.