unmanned aerial vehicle structure and propulsion

8/11/2019 Unmanned Aerial Vehicle Structure and Propulsion

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Unmanned Aerial Vehicle Structure and Propulsion

Submitted byChan Kai Rui Alphonsus

Department of

Mechanical Engineering

In partial fulfillment of the requirements for

Degree of Bachelor of Engineering

National University of Singapore

Session 2011/2012


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TABLE OF CONTENTS

Contents

ABSTRACT ............................................................................................................ ii

ACKNOWLEDGEMENTS ................................................................................... iii

TABLE OF CONTENTS ....................................................................................... iv

LIST OF FIGURES ............................................................................................. viii

LIST OF TABLES ................................................................................................. xi

LIST OF SYMBOLS ............................................................................................ xii

1 INTRODUCTION .......................................................................................... 1

1.1 Background .............................................................................................. 1

1.2 Objective .................................................................................................. 1

1.3 Scope ........................................................................................................ 1

2 LITERATURE REVIEW................................................................................ 2

2.1 Materials and Manufacturing ................................................................... 2

2.1.1 Wood ................................................................................................. 3

2.1.2 Foam .................................................................................................. 4

2.1.3 Coverings/Skin .................................................................................. 6

2.2 Structures and Mechanisms ...................................................................... 8

2.2.1 Structure Design ................................................................................ 8

2.2.2 Wing Structure .................................................................................. 9


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REFERENCES ...................................................................................................... 51

APPENDIX ........................................................................................................... 53

Appendix A: Specifications of Balsa Wood and Extruded Polystyrene Foam [16]

[17] [18] ............................................................................................................ 53

Appendix B: Theoretical wing stress (for a 2G manoeuvre with FoS of 1.5) .. 56

Appendix C: Static Thrust Test ......................................................................... 57

Appendix D: Momentum Theory and Blade Element Theory .......................... 59

A) Momentum Theory ............................................................................. 59

B) Blade Element Theory (BET) ............................................................. 61

Appendix E: BEM theory for calculation of theoretical thrust ......................... 62

Appendix F: Drag Calculation (from Aerodynamics part) ............................... 63

Appendix G: Evaluation of Propulsion (from Optimization part) .................... 64

Appendix H: Structure Analysis from Flight Acceleration Data ...................... 66

Appendix I: Loop and Level Banking from Datalogger ................................... 68

Appendix J: L3 Cutlass[15] .............................................................................. 74


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Figure 24. Flight Test Model 2 with modular design............................................ 43

Figure 25. CAD of Flight Test Model 3 ................................................................ 44

Figure 26. Actual Flight Test Model 3 .................................................................. 45

Figure 27. Flight Test Model 3 in operation ......................................................... 45

Figure 28. Close up photo of motor mount ........................................................... 46

Figure 29. Flight Test Model 4 in operation ......................................................... 46

Figure 30. 3 axis of UAV during operation .......................................................... 47

Figure 31. Comparison of Altitude, Ground Speed and Servo Throttle from

Ardupilot Mission Planner .................................................................................... 48

Figure 32. Schematic Diagram of lift contribution during flight .......................... 49

Figure 33. Comparisons of compression strength of various foams ..................... 54

Figure 34. Comparisons of water absorption rate of various materials ................ 55

Figure 35. Comparison of different insulating materials ...................................... 55

Figure 36. Schematic Diagram of Momentum Theory ......................................... 60

Figure 37. Schematic Diagram of BET ................................................................. 62

Figure 38. Power Velocity Curve.......................................................................... 65

Figure 39. Accelerations in x, y, z axis from Ardupilot Mission Planner............. 66

Figure 40. Take-off acceleration graph ................................................................. 67

Figure 41. Crash-landing Acceleration graph ....................................................... 67

Figure 42. Flight Path of Loop .............................................................................. 68

Figure 43. Head-up Display on Ardupilot during Loop (at time 11:30:01) .......... 68





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Figure 52. Flight path of level banking ................................................................. 73


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LIST OF TABLES

Table 1. Parameters for theoretical thrust computation ........................................ 37

Table 2. Results for 10x6" propeller from static thrust test .................................. 37

Table 3. Comparison of Thrust values .................................................................. 40

Table 4. Properties of Balsa Wood ....................................................................... 53

Table 5. Properties of Plywood ............................................................................. 53

Table 6. Type of Specification of Tian Sheng XPS .............................................. 54

Table 7. Static Thrust Test Results ....................................................................... 57

Table 8. BEM computation table .......................................................................... 62

Table 9: Drag components .................................................................................... 63


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R Propeller Radius, m

r Element Radius, m

Density of Air, kg/m3

Normal Stress (MPa) and Blade Solidity

T Thrust, N

V Free Stream Velocity, m/s

Vi Induced Velocity, m/s

p Pressure Jump


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1 INTRODUCTION

1.1

Background

As technology advances, Unmanned Aerial Vehicles become more popular. It

becomes smaller and more sophisticated to meet the needs of the customers,

mostly for military use, such as reconnaissance and armed attacks. The advantage

of an UAV is the elimination of a pilot on-board. In this way, the UAV will not be

constrained by the physical capabilities of the pilot as the aircraft structure is

usually designed to take more G-force than what a human can take. This will also

eliminate the risk of being captured or killed in an attack or surveillance mission.

1.2Objective

The objectives of this project are to design, build and test a long endurance UAV

for monitoring a fixed operational area. The UAV must fit within one standard

SAF field pack and be easily assembled and disassembled in the field by an

infantry man The UAV should be capable of a near vertical take-off, automatic

transition to horizontal flight and automatic landing/recovery.

There are 5 parts to this project; Aerodynamics, Structure and Propulsion,

Instrumentation, Controls and Testing and Optimisation. Structure and propulsion

of the UAV will be covered in this thesis.

1.3Scope

This thesis will focus on meeting the requirement of structure integrity and thrust

for the UAV through the research and design, selection, manufacture and testing

phases.


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2 LITERATURE REVIEW

2.1Materials and Manufacturing

There are many ways and materials for construction of the UAV. The typical RC

planes are usually made of either balsa wood or foam. This is because they are

cheap and can be easily shaped into different profiles with the use of simple tools.

Also, they have very good strength to weight ratio which is suitable for the

construction of UAV prototypes. When the design is optimised and finalised, the

model can be reconstructed using composites or aluminium for better structure

integrity, together with better power systems.

In order to design a robust aircraft, failure modes such as fracture, buckling,

excessive elastic deformation, fatigue, impact damage, creep and corrosion must

be taken into consideration.

With better materials and a proper design, such failure modes can be prevented.

Worst case scenario, the aircraft will fail safely.

With intensive research and development on advanced composite materials,

design and manufacturing processes, UAVs are also developing and evolving

from the conventional design to more complicated designs which are more

efficient and durable, with shorter manufacturing time. One such manufacturing

process, which is getting popular, is Additive Manufacturing. ASTM recently

adopted this new terminology, replacing the old term, Rapid Prototyping because

this technology is not only used for prototyping but is now used to manufacture

the end products too. It is defined as the process of joining different materials,

layer by layer, to make objects directly from the 3D computerised drawings. This


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is achieved by existing technologies such as the 3D printers, Selective Laser

Melting, Stereo lithography. [1]

2.1.1 Wood

Before foam, modellers use wood to build the aircraft models. Wood is cheap and

has a relatively good strength to weight ratio which was required for an aircraft.

2.1.1.1 Balsa Wood

Balsa wood is a soft textured beige/grey wood with very little grain. Under the

microscope, the cells are very large and porous. Balsa wood is only the third or

fourth lightest wood in the world. However, all the woods lighter than balsa are

too weak for any practical applications. In fact, balsa wood has the best strength to

weight ratio in the commercial market, better than pine, hickory or oak. This

explains the extensive use in model aircrafts.

Balsa can be easily shaped and handled as it is light and soft. Unless it is a big

piece of balsa, it can be cut and shaped by simple hand tools and sanding blocks.

Properties of Balsa wood can be found in Appendix A.

2.1.1.2 Plywood

Plywood is made by stacking layers of veneers and ply at right angle to each other,

relative to the previous layer. This method of layering provides the material with

its strength. The strength of the plywood depends on the number of layers.

Properties of Plywood can be found in Appendix A.


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2.1.2 Foam

In the 1960s, modellers realised that tapered wings are more efficient than

constant chord wings due to the smaller wing tips. When they found out that these

profiles can be made easily with foam, foam wings became popular. [2]

Foam selection depends on the size of the model, type of covering, objectives of

model and manufacturing methods. There are many types of foam with different

properties such as density and colour. However, not all are suitable for the

manufacturing of aircraft models.

Density is defined as the mass per unit volume. Although higher density foams are

harder and stiffer, they weigh more. The higher density foams are typically used

in high aspect ratio wings that are relatively thin. [2]

2.1.2.1 Expanded Polypropylene Foam (EPP)

EPP is a poor construction material because of its low compression ratings. It is

rubbery but light and durable. EPP alone is unable to withstand heavy loads. It is

usually used as a packaging for electronic parts and bumpers for some

automobiles. They are usually used with fiberglass or carbon fiber spars to

improve the stiffness and are usually seen only on small and light airplanes and

gliders as they experience lower stress levels. EPP wings are typically covered

with cloth tape, or low temperature heat shrink coverings. EPP is about 3.5 times

of the price of the normal white Styrofoam. [2]

2.1.2.2 Expanded Polystyrene Foam (EPS)

The white Styrofoam that we used to buy from the books and stationeries stores

for arts and crafts when we were young is actually EPS. They are also used as


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forth, which also explains the name of this foam, Fan Fold Foam. This allows

modellers to bend the foam into the profile they desire.[2]

2.1.2.6 Depron Foam

The foam is extremely lightweight, flexible and strong, and will probably be used

for future ultra-lightweight aircraft models. [2]

Depron is a type of extruded polystyrene (XPS) closed-cell foam sheet material

often used below parquet flooring to reduce step noises. The density of Depron is

about the same as balsa wood. They are usually in 2mm, 3mm and 6mm, in white

and grey colours. [2]

Modellers can use epoxy or foam-safe glues such as polyurethane glue or foam

safe CA glues to assemble the foam components. However, one has to note that

the CA glues would cause the foam to become brittle and may break easily. [2]

When foam-safe glue is applied to the Depron sheets, a relatively strong bond

with the glue is formed. Because of the good bonding between the glue and the

foam, and the high flexibility of the foam, the resulting components made of

Depron foam are relatively tough and almost crash-proof. [2]

2.1.3 Coverings/Skin

Prolonged exposure to the sun would cause discolouration and the surface to be

dirty. Therefore, most foam structures are protected by some coverings. The

covering also provides a major portion of the wings strength, and determines the

type of foam to be used. [2]


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Coverings come in sheets made from different materials such as plastic, wood and

composites. [2]

2.1.3.1 Heat shrink plastic coverings

Heat shrink plastic coverings have been around for years. Low heat versions of

these coverings can be applied directly to most foam if the foam is covered with a

light coat of adhesive. These coverings increase the tensile strength tremendously

but not in the compression strength. This however is not a problem because

modellers can use carbon rods to provide the necessary compression strength. [2]

2.1.3.2 Cloth Tape

Cloth tape is a cheaper alternative to heat shrink coverings. It will still provide

high tensile strength when applied properly.

2.1.3.3 Wood coverings

Balsa, plywood, and obechi wood sheetings are common for models of all sizes.

They provide high compression and some tensile strength. These coverings are

usually finished with plastic coverings or light layers of fiberglass or composite

coverings and paint. This results in extremely strong wings with unlimited design

and colour choices. [2]

2.1.3.4

Composite coverings

Fiberglass, carbon fiber and Kevlar cloth and resin are common on high

performance, high end sailplanes, gliders, aerobatic aircrafts and commercial

UAV's. Although these coverings are the most expensive, they provide the best

strength to weight ratio. [2]


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2.2Structures and Mechanisms

Although it is usually a challenge to achieve efficient aerodynamics solution for

smaller UAV, structures and mechanisms usually benefit when the aircraft is

smaller. Larger aircrafts usually experience higher loads and therefore needs to be

constructed from stronger and stiffer materials so that they can carry the loads

over longer distances without failure by direct loading, bending or buckling. [3]

To prevent the skin between frames from local buckling, longerons are added.

There are other methods to reinforce the structure, but at the expense of cost or

weight. One of them is to use a solid piece of foam and shape it into the desired

fuselage, a shell to house the components. It provides the necessary protection for

the expensive yet fragile equipment. However, the foam surface is soft and prone

to dents and holes. Therefore the skin has to be hardened and reinforced by

coatings; simply by applying a layer of cloth tape or epoxy. The other method

would be to use expensive, light composites to construct the UAV. Not only does

it reduce the weight of the aircraft, it also provides more room for payload or

allows the task to be done efficiently, with a lighter aircraft. [3]

Most small-sized UAVs benefit such that they carry lighter loads and are only

required to travel shorter distances which reduces the probability of failure.

However, it is still important to ensure the robustness of the structure for man-

handling. [3]

2.2.1 Structure Design

Although the usual structural design methods apply similarly to unmanned and

manned aircrafts, their application may be slightly different. When designing any


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UAV structure, on top of the objectives given by the customers, e.g. tube-

launched, the basic requirements such as reliability, accessibility, maintenance

and cost have to be considered. [3]

For manned aircrafts, the components which are required to be accessed are

usually reachable from on-board. Otherwise, they are accessed through removable

panels on the external structure of the aircraft. However, this is kept to a minimum

and structural strength and stiffness and aerodynamics must not be compromised.

[3]

For unmanned aircrafts, because of the size, it is not possible to access from on-

board. As for access through removable panels, the panels must be large enough

for handling and may be too large in proportion with the surrounding structure. It

may weaken the structure. Therefore, a better and efficient solution would be

constructing the airframe into detachable modules, using composites if necessary.

[3]

The structural integrity of the UAV depends largely on the type of foam and spars

used. Aircrafts are usually exposed to 4 forces; thrust, drag, lift and weight. The

thrust force is generated by the motor used. Not only do the structures have to be

able to withstand the lift and drag generated by the cross sectional profile, it must

be able to minimise the impact damage if the plane crashes. [3]

2.2.2 Wing Structure

Material used for construction of the wings relates to overall wing strength. Basic

wing strength relates to tension and compression. When a wing is placed under

load, the lift produced by the wing pulled the wing up and the weight of the


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2.2.4 Wear and Fatigue

Regardless of structure or mechanism, not only does it have to efficiently and

effectively perform its required objective, it also has to continue to perform

reliably for a specified time. [3]

Components used in the design will be subjected to rigorous testing during the

development phase to ensure that they perform till the specified or calculated

operational lifespan. The maximum operating hours for both wear and fatigue

becomes necessary so as to know when to replace the component. [3]

Careful selection of materials to be used is one of the important factors which can

help reduce the chances of fatigue failure. Composites are more resistant to

fatigue therefore they are largely used to both manned and unmanned aircraft

nowadays; they prolong the fatigue lives of airframes. However, if the small UAV

is required to attain a fatigue life comparable to a larger aircraft, steels and

stronger alloys should be considered because they are usually able to tolerate the

stress. [3]

Components such as engines and transmission systems would remain to be

manufactured using metals and alloys. This is because these components are also

exposed to high temperature and wear, on top of the high load densities. [3]

2.2.5 Undercarriage design

For both manned and unmanned aircraft, the main purpose of the undercarriage is

to absorb and dissipate the impact energy on landing and to provide a stable base

for the aircraft. For HTOL aircraft, it provides support for the aircraft accelerate to

its take off speed and when it decelerate after landing. [3]


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Again, a small sized aircraft would favour in the design of the undercarriage

structure as compared to the larger aircrafts. However, for the design of the

undercarriage, there are several important factors that would negate the above

advantage if they are not taken care of. The impact velocity for UAVs is usually

greater. Some of the reasons are as follows. Firstly, the wing loading of UAVs is

much lower and therefore they are more prone to turbulent on landing. Secondly,

if the landing is controlled by the pilot, it all depends on the judgement of him and

is prone to human errors. Furthermore, he is flying the UAV from ground,

judgement from the ground is definitely not better than on-board. [3]

Usually, a designer would recommend designing an undercarriage with a vertical

impact velocity of more than 2m/s and deflection yielding a maximum of 4g

deceleration. There are designs ranging from simple to complex. The simplest

form is usually seen in VTOL where the undercarriage is a fixed tubular skid type

with flexible plates to absorb the touchdown impact energy with sufficient

damping. HTOL usually has a wheeled undercarriage, but served the same

purpose, shock absorption and damping. The wheels are installed so that the

aircraft can move on the ground and also to carry the aircraft to the necessary

take-off speed. However, the wheels on HTOL are generally bigger than those on

VTOL. For both type of aircrafts, the wheels are usually retracted so as to reduce

drag during flight. The retraction of the landing gears is preferably a forward or

sideways so that centre of gravity of the aircraft would not shift backwards,

affecting stability. Also, the wheels vary with the hardness of the landing area, the

softer it is, the bigger the wheels. [3]


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2.3Selection of power systems

An engine is the heart of an aircraft; it provides the required energy for flying. [3]

The power system converts the chemical or electrical energy into mechanical

energy which will be transformed into the lift and thrust force required. The speed

it flies, payload it can carry and type of manoeuvre it can do all depends directly

or indirectly to the type of power system used. Apart from these factors, vibration

from the engine must be taken into account as it may affect the other components

such as the Ardupilot or other sensors.

It may be tempting to get the latest technology or custom made engines for the

UAV, but one must consider the availability of it; the new engine might take a

few months to get shipped over and the project might be delayed. [3] Therefore, it

is probably wiser to get a suitable and available engine in the local market. The

power systems which are currently available in the market will be touched on

below.

In the past, the small engines sold in the hobby shop are not manufactured to any

defined design or reliability standards. The performance data are usually not

available as well. However, now that UAV are further researched, developed and

deployed, better engines with the necessary data are now manufactured by the

existing or new companies.

2.3.1 Piston Engines

In this category, there are 4 sub-types mainly the two-stroke engines, four-stroke

engines, stepped piston engines and the rotary engines.


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Two-stroke engines are generally lighter, simpler as it does not incorporate a

valve control system and it provides a higher power output with a more consistent

torque as there is one power stroke per revolution. However, the efficiency is

lower as fresh charge in the cylinder tends to escape and the other reason is that

the fresh charge mixes with the residual gases from the previous combustion. The

two-stroke engines tend to run hotter and require more cooling than the four-

stroke. [11]

Four-stroke engines are generally bigger and heavier, incorporate the valve

control system thus more complicated. For every 2 revolutions, it delivers one

power stroke therefore power output is lower. However, efficiency is higher than

two-stroke engines. [11]

All in all, better performance and efficiency will come with a price of either

weight or cost. It all depends on the requirements of the UAV. It is however

recommended to use a smaller engine if the aircraft is small and only required to

travel short distances while bigger engine when required to travel far and carry a

heavier payload.

2.3.2 Stepped Piston, two stroke engines

This type of engine is relatively new. It has a cylinder made up of 2 bores of

different diameter, with the smaller one on the top. Air will be compressed in two

stages; the compressed air at the bottom of the cylinder will enter the top section

where it will be compressed again for combustion. With higher compression ratio,

efficiency increases. [3]


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2.3.3 Rotary Engines

Rotary engines used to have issues with the seals between the rotor and the casing.

The engines were prone to wear and engine life is short. Torsional vibration used

to be very high but this problem has been eliminated. Hermes 450 and 180 UAV

of the Israeli are using rotary engines and claimed to have a long life and low

specific fuel consumption. [3]

Although the mass-to-power ratio is low, when using a rotary engine, the

operation has to be paired with a reduction gearbox and better cooling equipment.

The additional equipment would probably increase the mass significantly, towards

the mass of a four-stroke engine. [3]

At the moment, there is only a limited power range for this type of engine, not

available for power rating below 28 kW or above 60kW. [3]

2.3.4

Gas-Turbine Engines

Gas-turbine engines consist of turbo-jet, turbo-shaft and turbo-fan units. They are

generally quieter than piston engines. Turbo-jet works in the form of direct

propulsion whereas for turbo-shaft, the generated energy is used to drive a shaft to

do rotary works, usually used for helicopters. Turbo-fan is somewhat a

combination of the two. [3]

Basically, it incorporates a set of compressors and a set of turbine with a single

output shaft. The bad thing about gas-turbine engines is when the output load

increases, it slows down the compressors and turbine shafts. Thus reducing the

available power needed for accelerating the engine back to its operating speed


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until more fuel is injected for combustion. This causes a lag in time in the rotary

motion, dangerous for aircrafts with propellers or spinning blades. [3]

Nowadays, the turbo-shaft units come in free-power-turbine (FPT) configuration.

The output shaft is an independent shaft, so that it will be affected when output

load is increased. Also, when the fuel supplied is increased, the compressor spool

can be accelerated more rapidly, giving a faster response to the higher power

demand. [3]

In turbo-fan engines, most of the combustion energy forms a jetstream and some

are converted into mechanical energy used to drive a fan at the inlet to take in

more air for combustion. [3]

Turbo-jet engines are said to be most efficient when the least jet velocity is left

behind the aircraft. Gas-turbine engines are most fuel-efficient at its maximum

power, which uses more fuels. Therefore, as a recommendation, turbo-jet engines

are usually used for aircrafts required to fly very fast whereas turbo-shaft and

turbo-fan are usually used for slower planes. The smaller planes usually use

smaller engines like the piston engines because using the gas turbine engines for a

small plane would be exaggerating and wasting fuel. [3]

2.3.5

Electric Motors

This type of motor is different from those mentioned above, electric motors

convert electrical energy to mechanical energy, to drive a rotor, fan or propeller.

This electrical energy usually comes from batteries, solar or fuel cells. Electric

powered aircrafts are getting popular mainly because they are quieter, cleaner and

easier to star and operate than the combustion power systems. Also, it runs at the


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are divided into losses in the windings and losses in the magnetic circuit, termed

as copper losses and iron losses respectively The coreless motors are more

efficient than cored motors because they have lesser iron losses. However,

coreless motors are unable to withstand continuous high RPM and/or loads.

Brushed motors have an advantage and disadvantage; they are reasonably cheap

but the brushes wear out easily. [6]

Brushless motors operate differently. The rotor has permanent magnets attached

along the casing and the stator has coils. They are 3 phase AC synchronous

motors. The electrical speed controller send pulses of current to each set of coils

in a 120 degree pattern, frequency at which the speed controller sends the pulses

would determine the speed of the motor. The supplement of voltage to the coils

will generate a rotating magnetic field, followed by the rotor. There is physically

no contact between the magnets and the coils thus have no frictional parts to wear

out. There are two types of brushless motors, in runner and out runner. For an in

runner brushless motor, the rotor rotates inside the stator which is part of the outer

case. The casing acts as a heat sink, radiating heat off from the coils. For an out

runner brushless motor, the stator is located inside the rotor and the rotor is part of

the outer case which is spinning. This out runner is better than the in runner such

that it provides more torque, capable of driving bigger and more efficient

propellers without a gearbox. [6]

In general, brushless motors are a little more expensive than brushed motor but

have a higher efficiency of 80 to 90%. Unlike the brushed motors, they do not

have brush and any parts in contact with each other apart from the bearings.

Therefore, they are able to last for a long time, probably outlast the plane. [6]


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horizontal stabilisers, with the rudder and elevator as the control surfaces. As for

the wings, ailerons are the control surfaces.

After doing a market research, only 2 types of foam are easily available i.e.

Extruded Polystyrene Foam (XPS) and Expanded Polystyrene Foam (EPP). Also,

XPS has a better compression strength than wood. Therefore, extruded

polystyrene will be used for the manufacturing of the main body, wings and

empennage without the elevator, rudder and the ailerons.

Because of its uniform wall thickness of linked honey-comb structure without any

gap, it is able to withstand high compressive stresses. Also, because of its stable

structure, it possesses a long lifespan with excellent thermal, water and vapour

resistance. On top of the above properties, XPS also has extremely good anti-

corrosive performance; neither will it get moldy, nor will it decompose. Although

prolonged exposure to sunlight would cause it to look dull and dusty, the

mechanical and chemical properties would not be affected. [15]

The combination of the excellent properties as mentioned above makes it superior

and suitable for construction of the UAV wings, empennage and fuselage. Also,

with the help of the automated hot wire cutter, foam is an easier and faster way of

shaping complicated and precise profiles for the 3 components. Therefore, it saves

time and cost.

Specification and comparisons of the foams can be found in Appendix A.

3.1.2 Control surfaces

Balsa and plywood are used for the manufacturing of the control surfaces as they

are structurally more rigid than extruded polystyrene, when the control surfaces


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3.1.5 Adhesives

Some of the adhesives are not suitable for XPS. Some of them would corrode the

foam and create a cavity on the foam surface, as shown below. Although they may

stick initially, after the cavity is formed, the surfaces would no longer be in

contact and thus separate from each other. Based on the design and selection of

materials for the UAV, the followings joints have been permutated for the Lap-

Shear Test.

Figure 1. Reaction between XPS and various types of adhesives

Various joint combinations

1)

Foam to Foam

2) Foam to Balsa

3) Foam to Plywood

4)

Balsa to Plywood

5) Balsa to Balsa

6)

Plywood to Plywood

Various adhesives

1) UHU HART glue

2) DAISO EPS foam glue

3) Hot glue

4)

Cyanoacrylate super glue

5)

Epoxy


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Therefore, only hot glue, epoxy and the foam glue are suitable to join foam to

foam and other materials.

3.1.6

Lap-Shear Test

Figure 2. Specimens for Lap-Shear test

For this experiment, the main objective is to find out the strength of the various

joints using a combination of adhesives and materials. For this experiment, only

XPS, balsa and plywood would be used.

So the end of the specimens would be clamped tight as shown on the figures

below. The machine would then apply a force such that a shear rate of 2mm/min

would occur.

When conducting the test for specimens involving the XPS foam, I observed that

under tensile load, the foam would fracture at a load of approximately 8 to 10kg.

Also, the hot glue and epoxy foam to foam specimens have a much stronger

bonding than the foam glue (eps glue) specimen. However, the strength of the

bonding of foam glue, when used for joints between foam and wood, is

approximately the same as epoxy and hot glue. Judging from the experiment, the

best choice of adhesive for foam is epoxy. As for balsa to other wood, either


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cyanoacrylate glue or epoxy should be used. For plywood to plywood, again,

epoxy is the best choice.

Figure 3. Example of Lap-shear test

Lap Shear Test Results

Figure 4. Lap-Shear test for foam to foam joint

Figure 5. Lap-Shear test for plywood to plywood joint

0

5

10

15

0 1 2 3 4

Load

(KG

)

Extension(mm)

Foam to Foam

epoxy

hot glue

eps glue

0

20

40

60

0 5 10 15 20

Load

(KG)

Extension(mm)

Plywood to Plywood

epoxy

hot glue

uhu

ca glue


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Figure 6. Lap-Shear test for balsa to balsa joint

Figure 7. Lap-Shear test for plywood to balsa joint

Figure 8. Lap-Shear test for plywood to foam joint

Figure 9. Lap-Shear test for balsa to foam joint

0

5

10

15

0 2 4 6 8

Load(K

G)

Extension(mm)

Balsa to Balsa

ca glue

epoxy

hot glue

uhu glue

0

10

20

0 2 4 6 8 10

Load

(KG)

Extension(mm)

Plywood to Balsa

epoxy

hot glue

uhu

ca glue

0

5

10

15

20

0 0.5 1 1.5 2 2.5

Load

(KG)

Extension(mm)

Plywood to Foam

epoxy

hot glue

eps glue

0

5

10

0 1 2 3 4 5 6

Load

(KG)

Extension(mm)

Balsa to Foam

epoxy

hot glue

eps glue


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3.1.7

Wing Structure Stress Analysis

The wings and spars are the components which experience the highest stress as

the lift generated has to be equal to or more than the weight of the aircraft so as to

fly straight and level or ascend in altitude from the excess thrust. 4 carbon fiber

rods are used as spars to reinforce the 2 wings on each side. Carbon fiber rods are

chosen because they are strong and rigid unidirectionally; prevent the wings from

warping during flight. To design and ensure the structure integrity of the wings, a

safety factor of 1.5 is used in the calculations. [4]

Theoretical Stresses

- For high speed flight with FoS of 1.5

From aerodynamics simulation, at free stream velocity of 20m/s, the wings would

experience an upward lift of 14.3N.

Assuming uniform loading on the wings and carbon fiber rods of length 0.5m

each,

Each wing will take14.32 = 7.15

Each carbon fiber rod will take7.15

2= 3.575

Therefore,

() =() (1)

, () = 3.5750.5

= 7.151


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= ()2

2 (2)

, = ()2 2

=7.15

2 < 0.5 >2

= 0.89

= (3)

, =

= 0.890.0025

(0.00250.0015)

= 83.32

, = 83.32 1.5 = 124.98Given that outer radius of rod, c2is 2.5mm and inner radius of rod, c1is 1.5mm

= 43

22 + 12 + 1222 + 12 (4)

, = (22 12)

= (0.00252 0.00152)

= 1.26 1052


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= 172.2

= 1.12Therefore, under 2G and a safety factor of 1.5, the carbon rods would start to yield

but would not fail as the maximum stresses are lesser than the UTS of the carbon

rods.

Experimental Stresses (with FoS of 1.5)

As the stresses obtained from a 2G flight manoeuvre is higher than a constant,

high speed flight at 20m/s, an experimental load testing will be tested based on the

theoretical value obtained from the 2G flight manoeuvre.

,,= 1.23 2 = 2.46

, 2.460.5 9.81 1.5

= 0.75As the wing with two carbon fiber spars did not fail, the structure of the wings is

safe.

3.2Designs of Glide Test Model of UAV

Figure 11. Initial Concept

This is the initial conceptual design of the glide test model to get an idea of how

the position of Center of Gravity (CG) position and Aerodynamic Chord (AC)

affect the movement of the UAV during gliding.


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3.3Selection of Power System and Propeller

3.3.1 Selection of Motor

Among the various types of power systems, the electric model is found to be the

most suitable for the UAV. Therefore the AXI 2814/12 Gold Line is selected. It

is able to provide sufficient thrust for models up to 2kg. It is small and light, with

a weight of 106g. Also, it has high efficiency comparable to the piston motors. It

is simple and cleaner to use as it only requires a battery and an electrical speed

controller to run. Last but not least, it is quieter than the rest of the power systems,

which is good for surveillance UAVs.

Figure 15. AXI 2814/12 Gold Line

3.3.2 Selection of Battery

Batteries are rated by their capacity and discharge rate, usually denoted by the

units of maH and C respectively. There are usually battery recommendations

provided by the company manufacturing the motor.

Figure 16. Recommended setups for AXI 2814/12 Gold Line


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The selection of battery would determine the flight duration.

For example, with a 2 blade propeller mounted onto the AXI 2814/12 motor, at

60% throttle, current drawn from the 2200maH battery is 8.3A. At 100% throttle,

current drawn from the 2200maH battery is 28.5A.

Therefore,

For constant 60% throttle:

Flight time: 8.3 = 3.771 = 0.265 = 0.265 60 = 15.9

Discharge rating of battery to accommodate for 100% throttle:

Minimum no. of C:28.52.2 = 13

3.3.3 Static Thrust Tests

There are two methods available to measure static thrust, by using a balancing

beam or a spring scale. By varying the propeller, which comes with different

diameter, pitch and number of blade, we can measure the thrust generated by the

motor and propeller. Also, by connecting a current sensor to the circuit, we can

obtain the current drawn, voltage and power. The propellers used for this

experiment are 3 blades 11x7, 3 blades 10x7, 2 blades 11x6, 2 blades 10x6, 2

blades 10x5 and 2 blades 9X6.

3.3.4 Balancing beam method

As shown below, the motor is mounted on one side of the beam and the other side

is resting on a digital weighing machine capable of measuring up to 1g. The

weighing machine and the motor are placed at equal distance from the center or


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34

The data for the static thrust test can be found in Appendix C. For experimental

thrust, it is shown that the aircraft has to be throttled up to 40% to 100% in order

to overcome the theoretical drag to fly straight and level or to ascend in altitude.

Figure 18. Balancing Beam Static Thrust Test Setup

Figure 19. Thrust vs Current graph (to determine the most efficient setup)

3.3.5 Spring scale method

The objective of conducting the second method, the spring scale method, is

primarily to compare the accuracy of the results. For this experiment, only the

spring scale, the motor and 10x6 propeller, mounted on the plane, are required.

As shown below, one end of the spring scale would be hooked to a fixed point,

while the other would be attached to the plane. After which, the motor would be


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throttled up based on the throttle gauge on the controller, with 20% increment to

100% increment. The thrust values in newton are then collected. It is observed

that the spring scale method is not as accurate as the balancing beam method as

there are frictional losses; maximum thrust at full throttle was less than 8N.

However, the frictional losses can be reduced by installing wheels at the belly of

the aircraft if permitted.

Figure 20. Spring Scale Static Thrust Test Setup

Figure 21. Thrust vs Throttle graph

3.3.6

Theoretical Thrust

There are several propeller theories available to calculate theoretical thrust. They

are the Rankine-Froude Momentum Theory, Blade Element Theory and

Combined Blade Element Momentum (BEM) Theory. Combined BEM was used

as the primary method to calculate theoretical thrust. The Momentum Theory and

Blade Element Theory can be found in the Appendix D.

0

5

10

0 50 100

Thrust(N)

Throttle (%)

Thrust (N) vs Throttle (%)

Thrust (N)

Linear (Thrust (N))


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3.3.7 Combined Blade Element Momentum Theory (BEM)

This theory is based on Prandtl's hypothesis stating that a section of a finite length

wing behaves as though it were a section of an infinite length wing set at an angle

equal to the effective angle of attack. [10] It is a combination of Momentum

Theory and BEM. Induced velocity can be obtained from the MomentumTheory. and of the blade airfoil can be found from the induced velocity andthe blade angles. The thrust generated by individual blade element is then

calculated and summed up to get the total thrust generated by the propeller.

Assumptions Made:

Constant lift curve slope, a0 = 3.6897 rad-1

Based on APC propeller profile; NACA4412 Cl curve [7]

Linear blade angle distribution

Near hub region (r < 0.012m) of the propeller blade does not generate

thrust

Figure 22. Blade Element angle definition


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Given:

Table 1. Parameters for theoretical thrust computation

Parameter Values

Propeller (10 x 6) Radius, R 0.127 m

Air density, 1.225 kg/m3

Free Stream Velocity, V 17 m/s

Figure 22. Specifications of AXI 2814/12 Gold Line

Table 2. Results for 10x6" propeller from static thrust test

Throttle Position

(%) Thrust (N)

Current

(A)

Voltage

(V)

0 0 0 11.75

20 1.11 1

40 4.93 3.8

60 6.89 8.3

80 10.75 15.2

100 17.58 28.5


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Step 1:

= 1390 11.75 (28.5 53 103) = 14232.9 =

14232.9

60= 237.22

Step 2:

= 1 2 =1

17

2(237.22)(0.021) = 0.5Step 3:

= =2(0.0196)

(0.127) = 0.098Step 4:

=

=

0.021

0.127= 0.165

Step 5:

= 1 + 8()0

=0.857 0.5

1 +8

0.165

sin(0.5)

0.098 3.6897

= 0.13

Step 6:

= = 0.857 0.5 0.13 = 0.227


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Step 7:

= 0 = 3.6897 0.227 = 0.838

Step 8:

From NACA 4412 Cdcurve, at = 0.227, Cd 0.05

Figure 23. Coefficient of Drag Curve for NACA 4412

Step 9:

0 = + = 0.5 + 0.13 = 0.63

Step 10:

=(2222)22 (0 0)

= 2 1.225 (22(237.22)2(0.021)2)(0.13)(0.5) (0.838 cos(0.63) 0.05 sin(0.63)) 0.0196

= 19.49


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Step 11:

=

(

)

(

) = 19.49

= 0.021

2

1

=> = ()0.0320.021

With the same method, we can calculate that when r = 0.032m,

()=0.032 = 40.04Therefore, using Trapezium Rule,

() (b a) ()()2

= (0.032 0.021) 19.49 + 40.042

= 0.327

Step 2 11 are repeated for the remaining blade elements to calculate the

corresponding thrust T produced from each element. The total thrust can then be

obtained by summing up all the element thrust value. The calculated values

can be found in the tables in Appendix E.

Theoretical Thrust= 10.92N

3.3.8

Comparison of Theoretical and Experimental Thrust

Table 3. Comparison of Thrust values

Theoretical Drag (from Aerodyanmics Calculations) 1.13N

Simulated Drag (from Solidworks) 3.8N

Optimised Drag (from Optimisation Program) 1.26N

Required Thrust (FoS of 2) 2.52N

Theoretical Thrust 10.92N

Experimental Thrust 17.58N


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Required thrust equates to the drag experienced by the aircraft, shown in

Appendix F. A factor of safety of 2 is used to ensure that there is sufficient thrust

to overcome the drag. Both the theoretical thrust and experimental thrust exceed

the value of the required thrust; about 4 times for theoretical and 7 times for

experimental thrust. Also, the experimental thrust is about 1.8 times of the weight

of the aircraft, thus sufficient thrust for vertical takeoff. Therefore, the selection of

motor and propeller is suitable. The experimental thrust is 61% more than the

theoretical thrust. The accuracy of the theoretical thrust depends greatly on the

accuracy of the air foil geometry data. Many of the propeller manufacturers in the

market do not state the airfoil geometry that they use as it is their trade secrets.

Although APC propeller manufacturer states that they are using NACA4412, it

might be a modified version of the NACA 4412 profile which will improves the

performance. For the theoretical thrust calculation above, the best information

available is the drag curve of NACA 4412. Also, there are several other

assumptions for the Combined Blade Element Momentum Theory which will

affect the accuracy of the BEM analysis. Therefore, the most accurate method of

measuring thrust is through a static thrust test.

An evaluation of thrust is also done in the Optimization part of the project, shown

in Appendix G.

3.4Flight Test Models

3.4.1 Flight Test Model 1

Knowing that the thrust is sufficient, the motor is then mounted onto the glide test

model with an L-shaped aluminium plate at the front. Servos and rods are


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connected to the vertical and horizontal stabilators (a combination of stabiliser

and its respective control surface such that they move as a whole). Ailerons were

not included in the first model; the vertical stabilator was supposed to perform

yawing primarily and rolling secondarily. However, rolling stability became an

issue. It was difficult to push and pull the foam stabiliator. When the stabiliators

were manufactured with extruded polystyrene initially, the control surfaces kept

warping as the servos pull and push the horns on the surfaces, due to the lack of

structure integrity. As a result, the force applied to the stabiliators was not

distributed uniformly causing them to warp. The motor mount deformed under the

torque generated by the motor and propeller. Also, the instruments were not

secured during flight and thus caused the CG to change, resulting in flight

instability. As for the wings, because there wasnt any reinforcement, they tend to

break upon high impact.

Figure 23. Flight Test Model 1


For the second model of our UAV, the design of the fuselage is similar to the first,

made out of 4 layers of XPS, two thin layers as the exterior and the thicker layers

in the middle so as to act as a base for the instruments. The instruments are not

secured onto a piece of plywood for easy installation and replacement. The motor

is mounted on a firewall made of plywood and balsa. Also, the areas where the


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spars are in contact with the foam are reinforced by sandwiching them between

two layers of plywood. For better load distribution, the stabilators were

remodelled into stabilisers and control surfaces separately and used plywood to

manufacture the control surfaces so that they would be light and rigid. The

problems with this model lie with the motor mount and the outer layers of the

fuselage. Because the contact area between the motor mount and foam is too little,

it separates upon impact and thus shattering the firewall. Due to the thinness of

the foam at the outer layers, they were not able to withstand high stress or impact.

Therefore, it tends to crack after a few crashes.

Figure 24. Flight Test Model 2 with modular design


The third model of the UAV is a totally new and improved design to

accommodate for the entire payload required for an automated flight. The main

frame of the UAV has been replaced from 8mm carbon rods to a 22mm hollow

aluminium rod. The motor mount consists of two parts, the motor will be screwed

onto the female part and then as a whole, it is screwed onto the male part fixed on

the hollow aluminium rod. The fuselage is made out of 3 layers; the top and

bottom pieces are 30mm while the middle piece is 40mm. The 22mm hollow rod


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is fitted into the middle portion of the fuselage so that the thrust line would be in

the middle. Also, the Ardupilot module is placed at the top; battery and camera

modules are placed at the bottom so as to keep the CG close to the thrust line to

prevent the steep pitching up motion upon throttling. The servos for the elevator

and rudder are mounted on a platform instead of embedding them into the

fuselage; separate servos are embedded into the wings for the aileron. The box for

the wing spars is designed such that it can cater for both 12cm and 13cm chord

wings while keeping the same angle of incidence. This is primarily for optimising

the size of the wings. The wings and stabilisers are reinforced with transparent

adhesive tapes so as to improve the surface toughness and smoothness. The

purpose of this model is primarily to determine the stability and competency of

the UAV through the various studies of individual parts of the UAV, i.e. the lift

generated by the wings, the effectiveness of the control surfaces, the thrust

generated by the motor and the propeller, the positioning and the responsiveness

of the instruments, the integrity of the structure, the overall optimisation of the

aircraft. Minor tweaks were made to the model to improve the performance and

stability. This model is much superior compared to the previous models, in term

of structure and performance; instruments are well protected during crashes, the

UAV can be repaired easily and it is able to take off, fly and land smoothly.

Figure 25. CAD of Flight Test Model 3

Camera

Battery

Ardupilot


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Figure 26. Actual Flight Test Model 3

Figure 27. Flight Test Model 3 in operation


This is the final design of the UAV. The folding mechanism for the wings is

installed to the model with some changes to the design. The instruments are still

mounted at the same position, with Ardupilot at the top and battery and camera at

the belly of the UAV. The wings, stabilisers and control surfaces are of the same

dimensions as model 3. However, to reduce weight, only the critical areas are

reinforced with cloth tape, i.e. the leading edges, trailing edges and the tips of the

wings and stabilisers. The fuselage consists of 2 halves of a shell constructed out

of XPS, joined together with magnets and a belly piece cut out from a plastic

bottle, cloth taped onto the two halves. The instruments are now mounted onto

plywood platforms glued to plastic pipe holders secured onto the 22mm

aluminium rod. The motor is mounted onto the firewall, which is simplified by

using a aluminium plate, bent into a U-shaped profile. Two blocks of wood act as

spacer between the gaps and two carbon rods are slotted into the holes such that


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the mount is aligned and secured. Holes are drilled onto the aluminium rod to

secure the carbon fiber spars for the stabilisers, thus eliminating the need of using

plastic holders at the empennage.

Figure 29. Close up photo of method of mounting instruments

Figure 28. Close up photo of motor mount

Figure 29. Flight Test Model 4 in operation

Plastic pipe holders

Plywood platforms

Carbon Rods

Wooden spacers

U-Shaped Plate

Motor

XPS

fuselage

PET

bottle

cut-out

Battery

Ardupilot


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4 POST FLIGHT ANALYSIS

4.1Structure Analysis from Flight Acceleration Data

The figure below shows the 3 axis of acceleration and the corresponding

directions. Ardupilot was mounted and recorded the data during one of the flights.

The data could be found in Appendix H. Take off accelerations data are shown in

Graph A. In the X axis, the acceleration was 18m/s2, which was close to 2G. In

the Y axis, the acceleration was close to zero because it was flying straight. When

the acceleration was a negative value, it was banking left, while a positive value,

it was banking right. Acceleration was greatest in the Z axis at a value of -28m/s 2,

which was close to 3G. Z axis acceleration is negative at take off because it is

ascending and acting against gravity.

In Graph B, it shows a crash landing when the aircraft stalled. Acceleration in the

X axis was as high as 38m/s2, approximately 4G. However, the UAV only

suffered minor damages like dents and cracks on the fuselage and wings from the

crash; the UAV is robust and the fuselage protects the instruments well. Also, due

to its modular design, the UAV components can be easily replaced and is ready to

fly again.

X Axis

Y Axis

Z Axis

Figure 30. 3 axis of UAV during operation


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A loop was also attempted but the values collected are wrong, the values of the

acceleration are drastically big and therefore cannot be used. However, from the

loop performed by the pilot, it is observed that the plane is structurally capable of

doing a loop while maintaining its structural integrity.

A level banking was also performed, at 33 degrees of roll, a radius of 63m, and a

velocity of 10m/s.

,

=

2

= 1.6

/

2

The acceleration experienced was equivalent to 0.16G. 33 degrees was the

maximum rolling angle of the UAV, due to its aerodynamics characteristics;

beyond 33 degrees, there would be rolling instability. Therefore, a steep bank of

45 to 60 degrees cannot be attained for testing.

The screenshots of the loop and the level banking can be found in Appendix I.

4.2Verifying Thrust through Flight Data

Figure 31. Comparison of Altitude, Ground Speed and Servo Throttle from Ardupilot Mission Planner

From the altitude, ground speed and servo throttle data collected, it is observed

that the UAV is cruising at an average speed of 15m/s to 17m/s, from takeoff to

Altitude

Servo Throttle

Ground Speed


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straight and level. From the graph, throttle was at 100% at takeoff hence the steep

increase in altitude. As the pilot reduces the throttle percentage, the altitude climb

rate becomes gentler. However, the speed of the UAV remains at its average

cruising speed. As shown in the figure below, the UAV may be pitching up at a

small angle of 3 to 4 degrees due to relative wind, thus excess thrust is contributed

to lift. The speed of the aircraft is constant because forward thrust is equal to drag,

when thrust is more than drag, the aircraft will accelerate. Therefore ideally, if the

UAV is flying at 0 degree angle of attack, the speed of the aircraft is

proportionally to the thrust and the throttle settings.

Figure 32. Schematic Diagram of lift contribution during flight

5 CONCLUSION AND RECOMMENDATIONS

This project has succeeded in design and building a foldable wing UAV with

deployment of wings with servos which can be automated through Ardupilot. We

have tried to launch the UAV with two methods.

First, to attach a water rocket to the UAV and pressurize the water rocket till it

launches. As the rocket loses water mass, it will decelerate. At that moment, the

UAV would still be accelerating and the water rocket will detach from the UAV at

the attachment points and the UAV would continue to fly with the motor powered

up and the wings deployed. For this method, the attachment points which can only


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be epoxied, hot glued or taped to the bottle have a high tendency to shear off

while the water rocket launches. Alternatively, the water rocket could be

integrated as part of the UAV as a booster for launching.

The other method was to place the UAV onto a ramp which launches the UAV by

elastic energy with the use of rubber bands or bungee cord.

Ideally, the tube-launched UAV should look something like the L3 Cutlass, in

Appendix J. The UAV should be able to be packed into a sacrificial shell which

would be used for combustion launching and after launching, should detach itself

from the UAV.

However, the UAV built had been easily hand launched and took off successfully

with or without the help of auto pilot. However, the Ardupilot does help a lot in

stability and ease of flight. Also, there was sufficient thrust for the UAV to take

off and perform a complete flight test while maintaining the integrity of the

structures under 2G to 3G for flight operations. It was also tested that the UAV

could sustain a 4G impact onto the ground with minimum damage to the

structures and instruments, proving that the design is robust. Repairs could be

done in the field by simply replacing the faulty or damaged components; repair

time has been minimized. Also, with a take-off weight of 1kg, it is very portable

and manageable by an infantryman. Lastly, the UAV was able to carry out its

automated mission to provide surveillance at the selected waypoints.


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fromhttp://www.ewp.asn.au/library/downloads/ewpaa_ply_and_lvl_design.pd

f
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19850021052_1985021052.pdfhttp://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19850021052_1985021052.pdfhttp://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19850021052_1985021052.pdfhttp://www2.l-3com.com/uas/pdf_uas/tech/Cutlass.pdfhttp://www2.l-3com.com/uas/pdf_uas/tech/Cutlass.pdfhttp://www2.l-3com.com/uas/pdf_uas/tech/Cutlass.pdfhttp://www.haowei.com.sg/product_tianshengxps.htmlhttp://www.haowei.com.sg/product_tianshengxps.htmlhttp://www.haowei.com.sg/product_tianshengxps.htmlhttp://www.auszac.com/Balsa%20wood%20Properties%20Guide.pdfhttp://www.auszac.com/Balsa%20wood%20Properties%20Guide.pdfhttp://www.auszac.com/Balsa%20wood%20Properties%20Guide.pdfhttp://www.ewp.asn.au/library/downloads/ewpaa_ply_and_lvl_design.pdfhttp://www.ewp.asn.au/library/downloads/ewpaa_ply_and_lvl_design.pdfhttp://www.ewp.asn.au/library/downloads/ewpaa_ply_and_lvl_design.pdfhttp://www.ewp.asn.au/library/downloads/ewpaa_ply_and_lvl_design.pdfhttp://www.ewp.asn.au/library/downloads/ewpaa_ply_and_lvl_design.pdfhttp://www.ewp.asn.au/library/downloads/ewpaa_ply_and_lvl_design.pdfhttp://www.auszac.com/Balsa%20wood%20Properties%20Guide.pdfhttp://www.haowei.com.sg/product_tianshengxps.htmlhttp://www2.l-3com.com/uas/pdf_uas/tech/Cutlass.pdfhttp://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19850021052_1985021052.pdfhttp://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19850021052_1985021052.pdf


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APPENDIX

Appendix A: Specifications of Balsa Wood and Extruded Polystyrene

Foam [16] [17] [18]

Table 4. Properties of Balsa Wood

Table 5. Properties of Plywood


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Figure 34. Comparisons of water absorption rate of various materials

Figure 35. Comparison of different insulating materials


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Appendix B: Theoretical wing stress (for a 2G manoeuvre with FoS of

1.5)

() =() [1]

, () = 4.9050.5

= 9.811

= ()2

2 [2]

, = ()2 2

=9.81

2 < 0.5 >2

= 1.23

= [3]

, =

= 1.230.0025

(0.00250.0015)

= 114.80

, = 114.80 1.5 = 172.2Given that outer radius of rod, c2is 2.5mm and inner radius of rod, c1is 1.5mm

= 43

22 + 12 + 1222 + 12 [4]


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, = (22 12)

= (0.00252 0.00152)

= 1.26 1052,

=

4 4.905 1.53 1.26 105

0.00252 + (0.0015)(0.0025) + 0.001520.00252 + 0.00152 = 1.12

Appendix C: Static Thrust Test

Table 7. Static Thrust Test Results

No. Motor PropellerThrottlePosition

(%)

Weightof Setup

(kg)

Value onWeighing

Machine

Thrust

(kg)

Thrust

(N)

Current

(A)Voltage (V)

1AXI

28143 blades 10x7" 0.00 0.16 -0.16 0.00 0.00 0.00 12.40

2AXI

28143 blades 10x7" 20.00 0.16 0.43 0.59 5.82 7.50 12.05

3AXI

28143 blades 10x7" 40.00 0.16 1.08 1.25 12.21 14.90 11.55

4AXI

28143 blades 10x7" 60.00 0.16 1.50 1.67 16.34 26.80 10.80

5AXI

28143 blades 10x7" 80.00 0.16 1.86 2.02 19.81 40.50 10.15

6AXI

28143 blades 10x7" 100.00 0.16 1.98 2.14 20.97 48.90 9.80

7AXI

28143 blades 11x7" 0.00 0.18 -0.18 0.00 0.00 0.00 11.65

8AXI

28143 blades 11x7" 20.00 0.18 -0.08 0.10 0.97 1.10 11.60

9AXI

28143 blades 11x7" 40.00 0.18 0.27 0.45 4.39 4.60 11.40


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No. Motor Propeller

Throttle

Position

(%)

Weight

of Setup

(kg)

Value on

Weighing

Machine

Thrust

(kg)

Thrust

(N)

Current

(A)Voltage (V)

10AXI

28143 blades 11x7" 60.00 0.18 0.90 1.08 10.57 13.50 11.00

11AXI

28143 blades 11x7" 80.00 0.18 1.29 1.47 14.44 23.20 10.45

12AXI

28143 blades 11x7" 100.00 0.18 1.82 2.00 19.59 40.30 9.75

14AXI

28142 blades 10x5" 0.00 0.16 -0.16 0.00 0.00 0.00 12.40

15AXI

28142 blades 10x5" 20.00 0.16 0.03 0.19 1.85 2.10 12.10

16AXI

28142 blades 10x5" 40.00 0.16 0.28 0.44 4.30 4.80 11.90

17AXI

28142 blades 10x5" 60.00 0.16 0.64 0.80 7.87 10.50 11.55

18AXI

28142 blades 10x5" 80.00 0.16 0.97 1.13 11.06 16.50 11.05

19AXI

28142 blades 10x5" 100.00 0.16 1.33 1.48 14.56 17.60 11.00

21AXI

28142 blades 11x6" 0.00 0.18 -0.18 0.00 0.00 0.00 12.05

22AXI

28142 blades 11x6" 20.00 0.18 -0.07 0.11 1.03 1.00 11.75

23AXI

28142 blades 11x6" 40.00 0.18 0.22 0.40 3.92 4.80 11.50

24AXI

28142 blades 11x6" 60.00 0.18 0.70 0.88 8.59 12.70 11.15

25AXI

28142 blades 11x6" 80.00 0.18 1.28 1.46 14.27 23.10 10.55

26AXI

28142 blades 11x6" 100.00 0.18 1.43 1.61 15.78 28.00 10.30

28AXI

28142 blades 9x6" 0.00 0.18 -0.18 0.00 0.00 0.00 11.90

29AXI

28142 blades 9x6" 20.00 0.18 -0.07 0.11 1.06 1.30 11.55

30AXI

28142 blades 9x6" 40.00 0.18 0.14 0.32 3.12 3.90 11.40

31AXI

28142 blades 9x6" 60.00 0.18 0.41 0.59 5.78 8.10 11.20

32AXI

28142 blades 9x6" 80.00 0.18 0.74 0.92 9.01 14.80 10.80

33AXI

2814

2 blades 9x6" 100.00 0.18 1.29 1.46 14.32 30.50 10.10


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No. Motor Propeller

Throttle

Position

(%)

Weight

of Setup

(kg)

Value on

Weighing

Machine

Thrust

(kg)

Thrust

(N)

Current

(A)Voltage (V)

35AXI

28142 blades 10x6" 0.00 0.16 -0.16 0.00 0.00 0.00 11.75

36AXI

28142 blades 10x6" 20.00 0.16 -0.05 0.11 1.11 1.00 11.40

37AXI

28142 blades 10x6" 40.00 0.16 0.34 0.50 4.93 3.80 11.25

38AXI

28142 blades 10x6" 60.00 0.16 0.54 0.70 6.89 8.30 11.05

39AXI

28142 blades 10x6" 80.00 0.16 0.93 1.10 10.75 15.20 10.65

41AXI

28142 blades 10x6" 100.00 0.16 1.63 1.79 17.58 28.50 10.15

Appendix D: Momentum Theory and Blade Element Theory

A) Momentum Theory

Momentum Theory is initially developed by W.J.M. Rankine, Alfred George

Greenhill and R.E. Froude for the design of marine propellers. For this theory, the

propeller is assumed to be a uniformly loaded actuator disk with the same

diameter, but of immeasurable small thickness. Also, the working fluid is ideal

and absorbs all the power generated from the engine and dissipates it by

increasing the inlet pressure to the outlet pressure of the propeller.


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Figure 36. Schematic Diagram of Momentum Theory

Thrust produced by the propeller can be represented by:

= = + 12

Where T is thrust

A is the area of the actuator disc

is the pressure differenceis the density of the the working fluidV is the free stream velocity

is the induced velocityThe drawback about this theory is that geometry of the propeller is not taken into

account when calculating its performance. Therefore, it is only able to predict the

maximum efficiency and not the actual efficiency of the propeller.


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B) Blade Element Theory (BET)

The second theory, Blade Element Theory, later developed by William Froude,

David W. Taylor and Stefan Drzewiecki differs from the Momentum Theory. It

takes into account of geometry and aerodynamics characteristics of the propeller

blades. The blade is divided into many elementary strips, each strip having

individual width and chord, denoted by dr and c respectively. The lift and drag of

the blade elements are then calculated to obtain the thrust and torque denoted by T

and Q respectively, where

= = 122( )

= () = 122()

Where T is thrust

L is the lift

is the 2D coefficient of liftD is the drag

is the 2D coefficient of dragis the angle of attackis the density of working fluidis the induced velocityc is the chord


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r is the width

Figure 37. Schematic Diagram of BET

However, this theory requires the and of the blade airfoil profile to calculatethe thrust and torque generated.

Appendix E: BEM theory for calculation of theoretical thrust

Table 8. BEM computation table

Free stream velocity = 17m/s, Throttle setting

= 100%

Element

Radius

(m)

0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.10 0.11 0.12 0.13

Chord

(m)0.02 0.02 0.02 0.03 0.02 0.02 0.02 0.02 0.02 0.01 0.01

(rad) 0.86 0.65 0.52 0.43 0.37 0.32 0.28 0.25 0.23 0.21 0.19

(deg) 49.10 37.50 29.88 24.67 20.93 18.14 15.99 14.29 12.91 11.77 10.81

0.50 0.35 0.27 0.21 0.18 0.15 0.13 0.12 0.11 0.10 0.09

0.10 0.11 0.12 0.13 0.12 0.11 0.11 0.09 0.08 0.05 0.03

x 0.17 0.25 0.33 0.42 0.50 0.58 0.67 0.75 0.83 0.92 1.00

0.13 0.11 0.10 0.09 0.07 0.06 0.05 0.04 0.03 0.02 0.01

0.23 0.20 0.16 0.13 0.12 0.10 0.09 0.09 0.08 0.09 0.08

(deg) 13.03 11.17 9.05 7.48 6.67 5.87 5.30 4.99 4.82 4.94 4.87


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Main Wings

(NACA 8414)

0.13 = 0.14

0.0065 X 1.5 =

0.00975

Horizontal Tail

(SD 8020)

0.026 =0.100.0058 X 1.5 =

0.0087

Vertical Tail (SD

8020)

0.00975 =0.10

0.0058 X 1.5 =

0.0087

Fuselage

(

= 0.08

, l =

0.156m)

0.005945

=1.9540.030X 1.5=

0.045

With reference to values obtained in Table 7:

= 12

1.225 172 0.0109 (0.026 + 0.13) + 12

1.225 172

(0.00975 0.13 + 0.0087 0.026 + 0.0087 0.00975

+ 0.045 0.0059) = 0.627

Appendix G: Evaluation of Propulsion (from Optimization part)

The optimization program minimizes drag, hence power provided by motor

should be sufficient to cater for climbing and near-vertical take-off by comparing

optimized drag-dependent power required and the power available from the motor

based on static thrust tests. The motor AXI2814 and 10x6 propeller was selected

after comparison of several combinations of motor and propeller as it provided the

best combination of maximum thrust and lower thrust-to-power ratio required for

endurance flight. Around the same altitude, static thrust produced is relatively

constant.


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65

=@T=D for level flight: =.

Figure 38. Power Velocity Curve

Area enclosed by P,req and P,avail curves show the excess power available for

climbing and maneuvers. Maximum velocities possible for level flight at different

throttle settings are marked at the intersections of the power curves. At higher

throttle settings, there is more than sufficient power for maneuvers such as

banking the aircraft and climbing even after consideration of a safety factor of 2

for near-vertical take-off.


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A) Take-off

Figure 40. Take-off acceleration graph

Red X Axis Green Y Axis Blue Z Axis

B) Crash-landing

Figure 41. Crash-landing Acceleration graph

Red X Axis Green Y Axis Blue Z Axis


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Figure 44. Head-up Display on Ardupilot during Loop (at time 11:30:02)



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Figure 52. Flight path of level banking


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unmanned aerial vehicle structure and propulsion

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