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 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 “ DESIGN OF A RC AIRCRAFT” AIRCRAFT” which is submitted by the students – Roopal Raj, Sandeep Kr. Mishra, Shreya Srivastava in Department of Mechanical Engineering JSS th
Academy of Technical Education, Noida during their 6 semester B. Tech. degree course of U. P. Technical University, is a record of the candidates work carried out by them under my supervision 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 ad vent 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 ……………………………………………………….
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Airfoil Selection……..……………………………………………. Selection ……..…………………………………………….
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Control Surfaces………………………………………………… Surfaces………………………………………………….. ..
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Stability Concepts ………………………………………………….. 11 Design Methodology ……………………………………………….
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Electronics components ……………………………………………. …………………………………………….
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Conclusion………………………………………………………. Conclusion……………………………………………………….
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References ……………………………………………………….
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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)
No. 13) Figure 5 - Neutral point (Page No. 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)
18) Figure 11 - Foil Sim Results (Page No. 18) Figure 12 – Auto Desk Velocity Field (Page No. 19)
No. 20) Figure 13 – Auto Desk Pressure Field (Page No. No. 21) Figure 14 – Rendered View (Page No. Figure 15 – Actual Image(Page No. 21) 21)
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INTRODUCTION: A fixed-wing aircraft's aircraft's wings, horizontal, wings, horizontal, and vertical and vertical stabilizers are built with airfoil-shaped cross-sections, as are helicopter are helicopter rotor blades. Swimming and flying creatures and even many plants and sessile 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 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 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 spinavoid spinstall. 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 The leading edge is the point at the front of the airfoil that has maximum curvature. The trailing The trailing edge is defined similarly as the point of maximum curvature at the rear of the airfoil. The chord 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 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 airfoil’s shape are its maximum 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 airfoil’s behavior when moving through a fluid are 7
The The aerodynamic center, 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 parametersOn 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 – speed – Airfoils Airfoils which are suitable for low speed aircraft comes under this section. Ex. Symmetrical airfoil.
High speed – speed – airfoils 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 a fluid produces produces an aerodynamic an aerodynamic force. force. The component of this force perpendicular force perpendicular to the direction of motion is called lift. called lift. The component parallel to the direction of motion is called drag. called drag. Subsonic flight airfoils have a characteristic shape with a rounded leading edge, followed by a sharp trailing edge, often with asymmetric with asymmetric camber. camber. Foils Foils of similar function designed with water as the working fluid are called hydrofoils. called hydrofoils. The lift on an airfoil is primarily the result of its angle 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 components: Lift and drag. and drag. Most Most foil shapes require a positive angle of attack to generate lift, but cambered but cambered airfoils can generate lift at zero angle of attack. This "turning" of 8
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 via Bernoulli's principle, 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 circulation and the Kutta-Joukowski the Kutta-Joukowski theorem.
Fig.2
Drag Coefficient:
The drag coefficient cd is defined as:
where: is the drag the drag force, which force, which is by definition the force component in the direction of the flow velocity, is the mass the mass density of the fluid, is the speed the speed of the object relative to the fluid and
is the reference area. 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, airfoils, the reference area is the plan 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 the lift force,
is fluid is fluid density, density,
is true is true airspeed,
is plan is plan form area and
is the
fluid dynamic fluid dynamic pressure. The pressure. The lift coefficient can be approximated app roximated using the lifting-line the lifting-line theory, numerically calculated or measured in a wind 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 ya w 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 deg ree 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 equa ls zero). Some aircraft can be trimmed by the pilot to fly 'hands off' for straight and level flight, for climb climb or for descent. Free flight models generally h ave 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: - lateral stability (stability in roll), directional stability (stability in yaw) - longitudinal stability (stability in pitch).
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Fig. 4
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 h ow 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 prev ented 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 a rm 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 rece ive 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 o ften 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 equa l 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 de sign 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 exa ct 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 4º angle of attack, Reynold’s 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 a irfoil 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 va lue 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 (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.
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The above image shows the final outcomes that are obtain from FOIL SIM III. Th e span length is = 1.06m, Area require to lift the plane = 0.188m² 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 m² and 0.47 m² 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 :
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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. ab ove. The software which we use for flow simulation was Flow Design (formerly Project Falcon), which is part of the Digital 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 d eep 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 capa city 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.
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