ME 4101 BACHELOR OF ENGINEERING DISSERTATION
Spherical VTOL Unmanned Aerial Vehicle
(Structures and Aerodynamics)
Project AM31
Faculty of Engineering
Department of Mechanical Engineering
Submitted By: CHIA YONG HAN TIMOTHY
Matriculation Number: A0073919U
Session 2012/2013
Page i
SUMMARY
The objective of this project is to design, build and test fly a Spherical-shaped
Vertical Take-Off (VTOL) Unmanned Aerial Vehicle (UAV). This is a 2-men
project which consists of 5 components, namely Propulsions, Electronics, Control
System, Aerodynamics and Structures. This thesis will only cover the
Aerodynamics and Structures components, and a the rest of the components will
be covered in Project AM30.
This newly shaped UAV allows the propellers to be encased within spherical-
shaped struts, which enables the UAV to fly into obstacles without damaging the
propeller. This also makes it safe for the pilot, as well as, anyone in the vicinity.
Running only on a single propeller-motor configuration, the UAV is able to hover,
climb vertically, and transit into translation flight, which is similar to that of a
helicopter.
Apart from that, this UAV boasts a new concept of take-off and landing, which is
not commonly seen in most of the other UAVs. During landing, the UAV rolls
onto the ground until it comes to a standstill. It can then be piloted to an upright
orientation before taking-off again. With a built-in wireless camera, the UAV is
able to navigate within tight spatial constraints, suitable for urban surveillance
operations. The camera is able to provide aerial view of live video footage during
flight.
A total of 3 prototypes were constructed where Prototypes 1 and 2 were
constructed to investigate and study the aerodynamics and structural components
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of the UAV. Only after undergoing thorough experimentation and analysis to
optimise the various component designs, Prototype 3 was finally constructed.
The aerodynamics component of the UAV was first dissected by using equations
to prove the pitch, roll and yaw motion. The control surfaces were optimised
using Computational Fluid Dynamics software during the design stages, and its
results were confirmed by various experiments and flight tests conducted.
The structural component of the UAV was mainly the study of the structural
integrity of the body of the UAV to ensure that it is able to withstand impact
loads. This was done using a simulation software, which provided useful
information of its effects when experiencing an impact load. Similarly, this was
confirmed by experiments which were conducted after the prototypes had been
constructed.
The final product is only made possible by integrating the electronics, control
system and propulsion components.
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ACKNOWLEDGEMENT
This author wishes to express his sincerest appreciation to A/Prof Gerard Leng for
his patience and guidance throughout the course of this project, as well as, staffs
of the Dynamics and Vibrations Laboratory for their administrative and logistical
support.
I would also like to extend my heartfelt thanks to my project partner, Andrew
Ong, who is working on the Electronics and Propulsion portion in Project AM30.
His hard work and sheer dedication to the project allowed for the timely
submission of this project. I would also like to express my gratitude to Andrew,
for all time and sacrifices that he had put in to ensure the success of this project.
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TABLE OF CONTENTS Page No
1) Introduction ................................................................................................................. 1
2) Literature Survey ......................................................................................................... 2
3) Stages and Objectives .................................................................................................. 3
3.1) Flowchart Process ............................................................................................................. 3
3.2) Project Goals ..................................................................................................................... 4
3.3) Objectives .......................................................................................................................... 5
4) Flight Dynamics .......................................................................................................... 6
4.1) Definition of Motion ......................................................................................................... 6
4.2) Propeller Slipstream .......................................................................................................... 9
4.3) Forces and Moments ....................................................................................................... 11
5) Design ........................................................................................................................ 16
5.1) CAD Modelling ............................................................................................................... 16
5.2) Center of Gravity ............................................................................................................. 16
5.3) Structural Simulations ..................................................................................................... 17
5.4) Flow Simulation .............................................................................................................. 19
5.5) Control Surfaces Optimisation ........................................................................................ 20
6) Prototyping ................................................................................................................ 21
6.1) Material Selection ........................................................................................................... 22
6.2) Prototypes Comparison ................................................................................................... 23
6.3) Manufacturing process .................................................................................................... 25
7) Experiments ............................................................................................................... 29
7.1) Thread Test...................................................................................................................... 29
7.2) Bench Test ....................................................................................................................... 30
7.3) Actual Drop Test ............................................................................................................. 31
8) Actual Flight Tests ..................................................................................................... 31
8.1) Indoor Flight Test ............................................................................................................ 32
8.2) Outdoor Flight Test ......................................................................................................... 34
9) Results and Discussion .............................................................................................. 35
9.1) Aerodynamics ................................................................................................................. 35
9.2) Structures......................................................................................................................... 36
9.3) Stability and Control ....................................................................................................... 38
10) UAV Performance ................................................................................................... 43
11) Conclusion ............................................................................................................... 44
12) Recommendations for Further Work ....................................................................... 45
REFERENCES ................................................................................................................. .
APPENDICES .................................................................................................................. .
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LIST OF FIGURES Page No
Fig 1 Japanese Spherical UAV (left) and Fixed-wing VTOL UAV (right) ...................................................... 2
Fig 2 Final UAV Prototype SolidWorks Model ............................................................................................... 6
Fig 3 Yaw Motion ............................................................................................................................................. 7
Fig 4 Pitch Motion ............................................................................................................................................ 7
Fig 5 Roll Motion ............................................................................................................................................ 8
Fig 6 Propeller Slipstream .............................................................................................................................. 9
Fig 7 Hover Front View (Left) and Top View (Right) ................................................................................ 11
Fig 8 Pitch while Hovering, transit to Translational Flight Front View .................................................... 13
Fig 9 Flaps Orientation during Translational Flight Top View ................................................................. 13
Fig 10 Finding Center of Gravity using SolidWorks ..................................................................................... 17
Fig 11 Structural Deformation Front ......................................................................................................... 18
Fig 12 Stress Concentration Lower Struts (Left), Octa Plate (Right) ......................................................... 18
Fig 13 Flap Sizing ........................................................................................................................................... 20
Fig 14 UAV Materials ................................................................................................................................... 23
Fig 15 Prototype 1 ......................................................................................................................................... 23
Fig 16 Prototype 2 ......................................................................................................................................... 24
Fig 17 Prototype 3 ......................................................................................................................................... 25
Fig 18 Constrution Process ........................................................................................................................... 26
Fig 19 Octa Plates ......................................................................................................................................... 27
Fig 20 Central Unit ......................................................................................................................................... 28
Fig 21 Thread Test ......................................................................................................................................... 29
Fig 22 Bench Test .......................................................................................................................................... 30
Fig 23 Actual Drop Test ................................................................................................................................ 31
Fig 24 Wall 'Sticking' .................................................................................................................................... 32
Fig 25 Take off by hand ................................................................................................................................. 33
Fig 26 Stairs Flight ....................................................................................................................................... 33
Fig 27 Outdoor Flight Path ........................................................................................................................... 34
Fig 28 Fracture at lower struts ..................................................................................................................... 38
Fig 29 Longitudinal Stability Zoom Out ..................................................................................................... 38
Fig 30 Longitudinal Stability Zoom In ....................................................................................................... 39
Fig 31 Longitudinal Stability with Disturbances Zoom Out ....................................................................... 40
Fig 32 Longitudinal Stability with Disturbances Zoom In ......................................................................... 40
Fig 33 Directional Stability - Zoom Out ........................................................................................................ 41
Fig 34 - Directional Stability - Zoom In ........................................................................................................ 42
Fig 35 Directional Stability with Disturbance - Zoom Out ........................................................................... 42
Fig 36 Directional Stability with Disturbance - Zoom In .............................................................................. 43
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LIST OF TABLES Page No
Table 1 - Project Phases and Objectives ......................................................................................................... 4
Table 2 Drop Test Conditions ....................................................................................................................... 18
Table 3 Flaps Optimisation ........................................................................................................................... 21
Table 4 Materials Ranking ............................................................................................................................ 22
Table 5 Airflow Velocity (m/s) at different locations ..................................................................................... 35
Table 6 Structural Simulation Results ........................................................................................................... 37
Table 8 UAV Performance ............................................................................................................................ 43
Table 9 Technical Properties (Depron, 2012) .............................................................................................. 2
Table 10 SolidWorks Material Input ................................................................................................................ 2
Table 11 Total Material Costs for Prototype 3 Structures .............................................................................. 3
Table 12 Total Costs Spent on Prototype 3 ..................................................................................................... 3
Table 13 Flow Simulation ................................................................................................................................ 1
LIST OF GRAPHS Page No
Graph 1 Cl vs AOA ........................................................................................................................................ 19
Graph 2 Cd vs AOA ....................................................................................................................................... 19
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LIST OF SYMBOLS
Angle of Attack (AOA) , rad
Flap Deflection, rad
Rudder Deflection, rad
Drag, N
Lift, N
b Wing span, m
Moment about CG
Xb Body Axis, X-Axis
Yb Body Axis, Y-Axis
Zb Body Axis, Z-Axis
Xe Earth Axis, X-Axis
Ye Earth Axis, Y-Axis
Ze Earth Axis, Z-Axis
T Thrust
Tmotor Motor Torque
Vi Induced velocity of air, accelerated across the propeller
Ve Induced exit velocity of air
Cl Lift Coefficient
Cd Drag Coefficient
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SPHERICAL VTOL UAV (STRUCTURES AND AERODYNAMICS)
1) Introduction
The evolution of UAVs started mainly for military purposes during World Wars I
and II. The need to gain aerial superiority at that time led to quick advancements
of unmanned aircrafts, as it meant that no lives were risked during reconnaissance
or sabotage missions. Then in recent years, after the invention of helicopters, the
race to develop fixed wing fighter jets capable of VTOL began to emerge as this
will allow fighter jets to take-off and land using shorter runways, which are also
found on-board aircraft carriers. This capability is also known as short take- off
and landing (STOL). In order to take advantage of both of these capabilities, a
hybrid VTOL UAV was created, with many variants that are currently researched
around the world.
This thesis will discuss the aerodynamics of the UAV which includes the forces
and moments equations governing the pitch, roll and yaw motion of UAV. This is
followed by the design process which includes CAD modelling, structural
analysis using SolidWorks and SolidWorks Motion Simulation. Computation
Fluid Dynamics using SolidWorks Flow Simulation was also done to analyse the
flow behaviour and optimise the control surfaces, before discussing the
manufacture and construction the actual prototype. Experimental data obtained
from the flight tests would be used to confirm the theoretical studies. Finally, a
more intensive date collection of the final prototype would be done to provide
information of the performance and capabilities of the UAV. The portion of
propulsion, electronics and control system is done by Andrew Ong, A0073979H,
project AM30 (AY2012/2013).
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2) Literature Survey
There are two types of VTOL UAVs currently available in the market that are of
interest to this project, which sparked the ideas of the prototype designs. The first
is the spherical UAV which was invented by in Japan. In October 2011, research
engineer, Fumiyuki Sato, working at the Defence Ministrys Technical Research,
presented the worlds first flying spherical UAV in a press conference. It runs on a
single propeller motor and hovers like a helicopter. The spherical UAV is also
claimed to reach speeds of up to 60 kilometres per hour when transiting into
translation flight. Costing about US$1,400, the latest prototype was made merely
from parts bought from a local consumer electronics store in Japan.
Fig 1 Japanese Spherical UAV (left) and Fixed-wing VTOL UAV (right)
The second UAV mimics a fixed wing aircraft capable of VTOL. Similar to the
Japanese spherical UAV, it is propelled by a single rotor which is usually placed
at the tip of the UAV. It is also capable of hovering as well as transiting into
translational flights which looks like a normal fixed wing aircraft.
During the design and prototyping stages, this project aims to investigate the pros
and cons of both designs and finally come up with a final prototype to achieve the
best possible configuration to achieve the objectives which have been set up.
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3) Stages and Objectives
This section describes the initial planning, and a brief summary of the process of
designing before arriving at the final prototype. This is followed by a list of
objectives to define the scope of the project.
3.1) Flowchart Process
Prior to the start of the entire project, the team has developed a flowchart process
to systematically tackle the various objectives, through a series of stages, to be
achieved before arriving at the final prototype.
Figure 1 - Flowchart Process
1) Design/Re-design
2) Prototyping
3) Flight Tests and Troubleshooting
4) Manufacturing of Final Product
5) Final Testing
6) Post Data Processing
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3.2) Project Goals
Below summarises the main goals during the different phases of the entire project:
Table 1 - Project Phases and Objectives
Phase Project Goals
1) Design Structures and Aerodynamics
- Dimensioning/ CAD Modelling
- Structures
- SolidWorks, SolidWorks Motion Simulation and
SolidWorks Flow Simulation
Propulsion and Control
-Sizing of Motor and Propeller Selection
-Calculating required current drawn and selecting
suitable battery
-Integrating Flight Control System with Flight
Controls
2) Prototyping - Integrating flight control electronics system with
UAV structure
3) Flight Tests and
Troubleshooting
- Testing of effectiveness of flight control surfaces
- Systematic Tuning of PID control settings
4) Re-Designing - Final modification of the structural, propulsion and
control designs
5) Manufacturing of
Final Product
- Consolidating modified designs
6) Final Testing - Make final adjustments to the control settings
7) Post Data Processing - Verify that all objectives are met
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3.3) Objectives
This thesis mainly focuses on the structures and aerodynamics of the UAV and
the objectives are as follow:
1) The overall costs for the entire project must be kept below SGD 500.
2) The material used must be lightweight to reduce the all-up weight of the
UAV, so as to increase flight endurance and allow for possible payload.
3) The design of the struts must able to withstand impacts when landing or
due to accidental collision against possible obstacles within an enclosed
Build-Up Area (BUA).
4) The flaps and rudder sizes must be optimised to achieve the highest lift-to-
drag ratio.
5) The final prototype of the UAV must successfully achieve the following
during the test flights:
a. Stability and Control
b. Ability to fly up a flight of stairs.
c. Withstand wall collisions and ability to stick on the surface of the
wall before returning to normal and stable flight.
d. Landing and rolling, followed by a successful take off from the
ground.
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4) Flight Dynamics
This section of the thesis will first introduce the various parts of the UAV and it
will be followed by the response on the motion of the UAV when the relevant
control surfaces are activated. The propeller slipstream theory will also be
explained and finally a theoretical derivation of the motion using equations of
forces and moments.
4.1) Definition of Motion
Typically, for a fixed wing aircraft, the pitch, roll and yaw motions, in the 3 axes,
are controlled by three different sets of control surfaces elevator, aileron and
rudder respectively. However, for the spherical UAV, the motions can be
simplified to 2 axes. This will be further explained in the following sub-section.
Fig 2 Final UAV Prototype SolidWorks Model
Propeller
Flaps (x4)
Rudders (x4)
Central Unit
Struts (x8)
Median Strut
Xe
Ze
Ye
Earth Axes
Xb
Zb
Yb
Body Axes
Octa-Plates (x4)
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4.1.1) Yaw
The yaw motion is controlled by four sets of rudders which have been
programmed to move concurrently when the signal is received. The torque effect
from the counter-clockwise motion of the propeller causes the body of the UAV to
rotate anti-clockwise (about the Z-axis), causing the UAV to yaw. In order to
counter the inherent yaw motion due to the moving propellers, the rudders must
be activated at an angle at its default trim position.
4.1.2) Pitch
Flaps (x2)
Rudders (x4)
Xe
Ze
Earth Axes
Xe
Ye
Earth Axes
Fig 3 Yaw Motion
Fig 4 Pitch Motion
Rudders (x4)
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The pitch motion (about the Y-axis) is controlled by two out of four flaps and four
rudders, which have been programmed to activate in a specific orientation to
cause the motion, as shown. With the Center of Gravity (CG) on the midline of
the UAV, the top two flaps move in opposite directions to the bottom rudders to
create a torqueing effect, thereby allowing the UAV to pitch in the desired
direction.
4.1.2) Roll
Fig 5 Roll Motion
Since the UAV is symmetrically constructed in each quadrant, the roll motion is
similar to the pitch motion, but the coupling moment is about the X-axis. The
pitch and roll motions are essentially describing the same movements. Therefore,
this motion will be defined by pitch in the subsequent sections of the thesis.
Ye
Ze
Earth Axes
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4.2) Propeller Slipstream
Assuming Conservation of Energy, where no loss of kinetic energy is experienced
within the control volume,
Fig 6 Propeller Slipstream
Applying Bernoullis Equation,
Since ,
From Regions (0) to (1): p
V
g p
V
g (4.3.1)
From Regions (2) to (3): p
V
g p
V
g (4.3.2)
Assuming incompressible flow and constant density of air across the propeller,
conservation of mass can be applied,
From Regions (1) to (2):
V A V A (4.3.3)
Since A1 and A2 are the area of the propeller which is represented by the disc, it is
the equal, A1 = A2 = Ad,
V0 (0)
(1)
(2)
V1
V2
V3 (3)
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V V V Vi (4.3.4)
Where Vi, is the induced velocity of the air which is accelerated across the
propeller.
Thrust produced from the propeller disk is due to pressure difference between
regions (1) and (2),
T (p p
) Ad (4.3.5)
Since V3 = V0 + Ve, where Ve is the induced exit velocity of air, and p3 = p0, and
also substituting (4.3.3), (4.3.4) and (4.3.5),
From (4.3.1); p p
V
V
From (4.3.1); p
V
p
V Ve
p
V
p
V
V
V Ve
T
Ad
V
V Ve
T
Ad
V
V
VeV Ve
T
Ad
VeV Ve
T
Ad VeV V
Using Quadratic equations to solve, the exit velocity can be determined by
V V V
T
Ad (4.3.6)
For hovering flight, since V0 = 0,
V T
Ad (4.3.7)
T
AdV
(4.3.8)
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4.3) Forces and Moments
This sub-section of the thesis will explain the response of the control surfaces on
the UAV theoretically, using forces and moments. The following are the
assumptions made:
1) The axial direction of flow of Ve is always parallel to the symmetrical axis
of the UAV body, and the normal velocity, Vn, is negligible because Ve >>
Vn.
2) The direction of airflow is smooth and uniform and is vertically
downwards in the -Zb-direction.
3) The control surfaces are completely submerged in the induced exit airflow,
Ve.
4.3.1) Hover Flight
During hover flight, the velocity of airflow on the control surfaces is the induced
exit velocity of the air produced by the propeller, Ve, because Ve >> V0.
Fig 7 Hover Front View (Left) and Top View (Right)
Ve
T
W
Xe
Z
Earth Axes Xe
Ye
Earth Axes
FD,rudders
FL,rudders
Vn
FL,rudder4
FL,rudder2
FL,rudder1
FL,rudder3
r
Page 12
During hover flight, there is an inherent yaw motion as mentioned in Section
4.1.1). Therefore, all four rudders must be in the orientation shown above at
neutral position to counter the clockwise rotation of the UAV body.
Sum of forces in Ze-Axis;
T W FD,rudders = m . w
Since w = 0;
T = W FD,rudders
Moments about Ze-Axis; (Taking counter-clockwise to be +ve)
Mcg,r = Tmotor - FL,rudders . r
where FL = Ve2 CL S and CL = 2 (for flat plate)
In order for the UAV to yaw in the CCW direction, Tmotor > FL,rudders . r , and for
the UAV to yaw in the CW direction, Tmotor < FL,rudders . r . This can be achieved by
controlling the amount of deflection of the rudders, r. The larger the the
deflection, the greater the FL,rudders.
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4.3.2) Translational Flight
During hover flight, the velocity of airflow on the control surfaces is the induced
exit velocity of the air produced by the propeller, Ve, because Ve >> V0.
Fig 8 Pitch while Hovering, transit to Translational Flight Front View
Fig 9 Flaps Orientation during Translational Flight Top View
During translational flight, the UAV must first pitch as shown in the orientation
above. When the pitching moment reaches equilibrium, the UAV will translate in
the resultant direction.
From left picture in Fig 8, the UAV achieved hover flight first,
Sum of forces in Ze-Axis;
Xe
Ze
Earth Axes
RESULTANT
T
T
V
Xe
Ye
Earth Axes
Ve Ve
FL,rudders
FL,rudders
FL,rudders
FL,flaps
FL,flaps
FL,flaps
FD,frontal
f
r
W
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T - W - FD,rudders - FD,flaps - FDstruts + FL,rudders + FL,flaps + FL,struts = m . w
Since w = 0;
T + FL,rudders + FL,flaps + FL,struts = W + FD,rudders + FD,flaps + FD,struts
However, from Fig 8, after reaching equilibrium when pitching,
Summation of forces in Ze-Axis;
T sin - W - FD,rudders - FD,flaps - FDstruts + FL,rudders + FL,flaps + FL,struts = m . w
T sin - W - D + L = m . w
For the same throttle setting, T, when the UAV achieved hover flight, there will
be a decrease in the overall thrust of T.(1 - sin ), due the change in resultant
thrust vector direction.
Since both (T sin ) and W are constants, the resultant increase or decrease in w is
dependent on D and L which is based on the Lift-to-Drag ratio, dependent
of the angle of attack , for a given Ve and V.
Summation of forces in Xe-Axis;
T cos - FD,struts - FD,rudders - FD,flaps = m . u
T cos - FD,frontal = m . u
where FD,frontal = (Ve + V) 2 Cd S
Moments about Ye-Axis;
Mcw, Flaps & Rudders = Lf,z-axis . (FL,flaps cos f) + Lr,z-axis . (FL,rudders cos r)
Mccw, Flaps & Rudders = Lf,x-axis . (FL,flaps cos f) + Lr,x-axis . (FL,rudders cos r)
Mcw by Propeller = Lp,x-axis . T cos
Mccw by Propeller = Lp,x-axis . T sin
where FL = (Ve + V) 2 CL S
(Taking counter-clockwise to be +ve)
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Mcg = {[ Lf,z-axis. (FL,flaps cos f) + Lr,z-axis . (FL,rudders cos r) ] + [ Lp,x-axis . T cos ]}
- {[ Lf,x-axis. (FL,flaps cos f) + Lr,x-axis . (FL,rudders cos r) ] + [Lp,x-axis . T sin ]}
Theoretically, it is difficult to achieve straight and levelled flight during pitching.
However if L is able to overcome W and D, it will result in an increase in
altitude, which is more ideal. Therefore, it is concluded that the UAV should
flown at a high enough thrust which must be decreased proportionally to
compensate for the increase in Lift at high velocity to achieve a successful
translational flight. This hypothesis will be further elaborated in Section 5.4) Flow
Simulation.
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5) Design
In order to make a fair comparison with the Japanese Sphere, some specifications
of the prototypes were based on the limited knowledge of the Japanese
specifications released. A total of 2 prototypes were made, before arriving at the
final prototype and the diameter of the UAV is fixed at 42 centimeters for all three
prototypes.
5.1) CAD Modelling
The prototypes were modelled using Computer Aided Software SolidWorks, to
speed up the design process by allowing the user to visualise the designs before
manufacturing the actual product. Each parts of the UAV were designed
individually and assembled using the software.
Apart from designing the prototypes to meet the technical specifications, the
design considerations also encompass the manufacturing process. A step-by-step
assembly procedure must be taken into consideration to ensure smooth assembly
as some of the steps require permanent adhesion which is irreversible.
5.2) Center of Gravity
To ensure that the centre of gravity is as close to centre of the sphere as possible,
all the electronics are modelled into individual parts and assigned their individual
masses, which were weighed using an electronic weighing machine. These parts
are then assembled to achieve the desired CG location, at the centroid of the
sphere.
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Fig 10 Finding Center of Gravity using SolidWorks
5.3) Structural Simulations
Before manufacturing the actual prototype, the CAD model is meshed before
undergoing simulations to test the structural designs. The simulation software
used was SolidWorks Motion Simulations.
To ensure that the struts are able to withstand the impact forces due to improper
landing or accidental collisions into obstacles when flying indoors, the UAV is
simulated to undergo a drop test. This provided useful visual aids to identify
locations where high stress concentrations exist, which may lead to possible
fractures. Changes to the designs can be made instantly to improve the structural
integrity of the struts.
5.3.1) Drop Test
The drop test is to simulate the sudden vertical decent of the UAV (in the Ze-
Axis) either due to loss of battery power or during improper landing at 3 different
height of 0.2, 0.5, 1.0 meters. The objective is to find out the location of the UAV
Page 18
which receives the most stress through the simulation. The following were the
boundary conditions applied:
Table 2 Drop Test Conditions
Conditions Inputs
Material: Depron Foam Refer to Appendix A: Depron Foam
Connections Global Contact (Bonded)
CG located at Center of Sphere 600g
Gravity in the Ze-Axis 9.81 m/s2
Fig 11 Structural Deformation Front
Fig 12 Stress Concentration Lower Struts (Left), Octa Plate (Right)
Based on the maximum normal stress criterion, the highest stress concentration
can be observed at the centre of the octa plates as shown in Fig 13. The results
also showed that high stress concentrations, represented by the red/orange regions,
are observed at the lower portion of the struts in Fig 13. The calculated shape of
deformation can also be observed in Fig 11. A more detailed discussion can be
found in 9.2) Structures.
Page 19
5.4) Flow Simulation
SolidWorks Flow Simulation is used to investigate the lift and drag forces at Ve of
13.2m/s while varying the AOA. The highest Lift Coefficient, Cl, can be observed
at approximately 42 degrees, from Graph 1 Cl vs AOA. Also, from Graph 2 Cd vs
AOA, it can be observed that Cd increases with AOA.
Graph 1 Cl vs AOA
Graph 2 Cd vs AOA
To further explain the hypothesis from 4.4.2) Translational Flight, it can be
concluded that at AOA of less than 42o, there will be enough lift force to cause a
resulting increase in altitude since the drag experienced is also lower at smaller
AOA.
0
0.05
0.1
0.15
0.2
0.25
0.3
0 10 20 30 40 50 60 70 80 90
Co
eff
icie
nt
of
Lift
, Cl
AOA, (o)
Cl vs AOA (o)
0
0.1
0.2
0.3
0.4
0 10 20 30 40 50 60 70 80 90
Co
eff
icie
nt
of
Dra
g, C
d
AOA, (o)
Cd vs AOA (o)
Page 20
5.5) Control Surfaces Optimisation
To achieve the highest lift-to-drag ratio, the size of the flaps and rudders must be
optimised. This will ensure that enough lift force is generated by both the flaps
and the rudders to provide sufficient pitching moment for translation flight, and
the size of the rudders is sufficient to counter the inherent yaw motion caused by
the propeller.
5.5.1) Flaps Optimisation
The objective of the flap optimisation is to find the suitable chord length to
provide the highest lift-to-drag ratio. For a fixed spherical UAV diameter of
420mm, the maximum chord length of the flap is 58.5mm. The velocity of
airflow, Ve, just below the propeller, is taken to be 13.2m/s at throttle setting of
50%, which is the approximated cruising speed. Below are the assumptions and
boundary conditions that were applied:
Span = 355
Chord Length, c
Ve = 13.2m/s
Fig 13 Flap Sizing
Page 21
Table 3 Flaps Optimisation
Assumptions/ Boundary Conditions Values
Density of air @ 300K, 1.17714 kg/m3
Dynamic Viscosity of air @ 300K, 1.56E-05 kg/m s
Velocity of air, Ve 13.2 m/s
Span, b 0.355m
AOA, 42o or 0.733 rad
Using SolidWorks Flow Simulation, a parametric study was setup by varying the
chord length between 40mm to 58.5mm.
Results showed that for a smaller chord length, there was only a slight decrease in
the lift-to-drag ratio. However, the resulting Lift force is reduced, compromising
on the resulting pitching moment, which is undesirable. Instead, the maximum
allowable chord length of 58.5mm produced the highest Lift-to-drag ratio.
Therefore, the most suitable chord length of the flap is limited to the diameter of
the sphere at 58.5mm.
In addition, it can be observed that from Appendix C, as AOA increases from 0 to
40 degrees, the simulation results showed that the air flow is still sticking onto
the surface of the flap. However, at 50 degrees onwards, flow separation is
starting to occurs. Therefore, it can concluded that the maximum pitch angle
should not be more than 42 degrees as observed from the Graph 1, as it is the
angle at which the maximum lift force can be generated.
6) Prototyping
There are limitations to how much each prototype can be simulated using software
simulations, and the prototyping stage only begins after intensive simulations
Page 22
have been done. This is where the manufacture and assembly of the prototypes
took place. This section will first describe the key considerations for the material
selection and how each prototype is evolved. This will be followed by a brief
description of the construction process.
6.1) Material Selection
Table 4 Materials Ranking
Materials Cost Ease of fabrication Strength-to-weight
ratio
Depron Foam
(3mm)
Low 1 Very Easy 1 Low 4
Depron Foam
(6mm)
Low 2 Very Easy 2 Low 3
Plywood Medium 3 Easy 3 Medium 2
Carbon Fibre Very High 4 Very difficult 4 High 1
There are various considerations that had to be made when selecting the most
appropriate material to manufacture the struts, which will experience the highest
impact loads.
Since the objective of the project is to keep costs low, costs was the highest
priority and thus Depron was selected over plywood and carbon fibre, which have
the highest strength-to-weight ratio.
Due to time constraint, the time taken for the construction was the next on the
priority list. Depron foam is the easiest choice of material to work with as it can
be cut easily with simple tools pen-knife or foam cutter. Therefore, it was
selected over all the other materials. Below are the materials used for the different
parts of the final prototype:
Page 23
Fig 14 UAV Materials
6.2) Prototypes Comparison
This subsection aims to describe the structures and aerodynamics characteristics
for the first two prototypes and explain the pros and cons of each model, before
arriving at the final prototype.
6.2.1) Prototype 1
Fig 15 Prototype 1
The key difference between Prototype 1 and the two later models is the
construction of the central unit. It is made of a 1.5 litre water bottle which was cut
to a suitable height, and its cap was used to mount the propeller. Since the bottle is
Depron Foam
5.0x0.6mm Carbon Strips
6mm Carbon Rods
Octagonal- Shaped 3mm Plywood
Page 24
shaped similar to that of an aircraft fuselage, its smooth and streamline body
allowed for better airflow. Its advantages and disadvantages are further explained
in 6.3.2. After thoroughly flight testings, the UAV was unable to achieve any
control or stability when the Ardupilot flight control board was installed.
6.2.1) Prototype 2
Fig 16 Prototype 2
Prototype 2 was designed based on the fixed wing VTOL UAV, which was
introduced in 2.1) Literature Survey. During the research, such fixed wing VTOL
UAVs in the market had shown to have high stability during hovering flight. The
construction of Prototype 2 was also much easier and much faster to construct.
It served as a stepping stone to test out the new KK2 Board flight control board
while Prototype 3 was being designed. The larger surface area at the central unit
also contributed to the increase in structural strength of the struts.
The KK2 Board was able to provide stability for the UAV, however, during
subsequent flight testings, it was observed that the UAV was too stable. The
increase in surface are resulted in an increase in front drag during translation
flight, which was undesirable.
Page 25
6.2.1) Prototype 3
Fig 17 Prototype 3
After many design considerations and optimisations from its predecessors,
Prototype 3 was finally constructed. It was installed with the KK2 Board, and
stability was achieved with minor tuning. The strut design was a hybrid of
Prototypes 1 and . It was more hollow at the central unit as compared Prototype
1, to decrease the front drag during translational flight. At the same time, the
assembly of the 8 struts makes up the central unit to install all the electronics. The
circumference of the struts was lined with carbon strips using epoxy as adhesives
to further increase the structural integrity of the UAV. The analysis of the entire
thesis is based on the Prototype 3.
6.3) Manufacturing process
This sub-section aims to briefly describe the process of constructing the
prototypes where similar processes were done for all three prototypes.
Page 26
6.3.1) Struts, Median Strut and Octa Plates
Fig 18 Constrution Process
Using SolidWorks, the part files can be converted into drawing files, which allows
the drawings to be printed in 1:1 scale. The template is then cut out and traced
onto Depron Foam sheets, which are cut by hand using a pen-knife. Slots are
specifically designed and cut to assemble the struts with the octa plates and
median struts, in a jig-saw like manner to properly distribute the stresses. All parts
are then assembled and epoxy is used as the adhesion for those assemblies that
require permanent bonds. The octa plates are laid with fibreglass cloth and cured
using epoxy resin to increase the structure strength as it experiences the largest
stress concentration seen in Fig 19.
1 2
3 4
Page 27
Fig 19 Octa Plates
6.3.2) Central Unit
The central unit is the middle structure of the UAV where all the electronics are
mounted on. Apart from having to withstand torsional loads from the moving
propeller, the central unit must be designed to allow the electronics to be mounted
sturdily.
For the first prototype, an indigenous idea of using a bottle for the central unit
struck the team. The motor fits perfectly on the bottle cap and the torqueing
direction of the propeller tightens the cap further which enhances the design
safety. The cylindrical body of the bottle is ideal for taking torsional loads, and
the electronics can be contained within it. Despite its advantages, a lot of time is
spent on creating brackets for the electronics and fitting them into the bottle. In
addition, holes and slits have to be cut to access the electronics which decreases
the structural integrity of the body. Therefore, the idea was scraped in the
subsequent designs.
Fiberglass cloth with epoxy
Page 28
Fig 20 Central Unit
In the second and third prototypes, the central unit is made from the assembly of
all eight struts. This will allow greater stress distribution of the struts thereby
increasing the structural integrity of the struts. The final improvement to the struts
is seen in the third prototype, where mm carbon fibre strips are stuck onto the
circumference of the struts, using epoxy, to increase the strength of the struts.
Page 29
7) Experiments
This section describes the main experimental test flights which were done after
each prototype was constructed, followed by an actual drop test to validate the
structural integrity of the UAV. The results and discussion will be done in Section
9 of the thesis.
7.1) Thread Test
To investigate if the boundary layer is laminar or turbulent on the control surfaces
and inner surfaces of the UAV, flight threads were attached to the respective
surfaces to observe the behaviour of the threads. If the threads flap in a smooth
and uniform downward direction, it can be concluded that the flow is laminar. If
the threads flap erratically, there is a possible flow separation on the particular
surface has occurred, resulting in turbulent flow in the region.
The throttle setting was increased at intervals from 13.2m/s to 16.8m/s, while
observing the behaviour of the flight threads at 3 different surfaces of the UAV, as
V3 (3)
Flaps
Internal Wall
Rudders
V0 (0)
(1)
(2)
V1 V2
Mid-Section
Fig 21 Thread Test
Flaps
Internal Wall
V2
Page 30
shown in Fig 21. An anemometer was carefully placed just beside the surfaces to
observe the airspeed at the various throttle settings.
7.2) Bench Test
Before the actual test flight, the UAV needs to first attain dynamic stability under
a controlled environment. Dynamic stability refers to the dampening of
oscillations caused by disturbances back to the original equilibrium position, with
time.
In order to prove the forces and moments were valid for the pitching motion, the
Bench test was set up to restrict the UAV to just one Degree of reedom about
the CG. The controller is then activated by the pilot to test if the activation the
correct sets of control surfaces will result in a coupling moment, causing it to
pitch in the desired direction.
Fig 22 Bench Test
Using the same test rig and set-up, but now with the flight control board activated,
a series of P-I-D settings was tested to investigate the effects of the increase or
decrease in the input gain of the P-I-D settings on the response of dampening the
oscillations of the UAV, when disturbed at an angle.
Page 31
The gain values in this experiment were predicted to be overestimated as there are
frictional forces acting at the contacts where the UAV is being hung. However, it
provides useful information of the effects on the change in P-I-D settings which
will reduce the time taken to trim the settings during actual flight tests.
7.3) Actual Drop Test
Fig 23 Actual Drop Test
As mentioned in the design process of the drop test in 5.3.1 Drop Test, and actual
drop test was done for the Final Prototype to test out the structural integrity of the
UAV in the event of a motor failure. The objective of this test is to ensure that the
internal electronics and control surfaces are protected within the spherical struts.
Similar to the simulation done using SolidWorks Simulation, the UAV is released
at 3 different heights of 0.20m, 0.5m and 1m.
8) Actual Flight Tests
Once the UAV was able to achieve decent stability using the bench test, the
UAV will then undergo the actual indoor flight test. The original P-I-D settings
derived from the bench test were no longer valid, as the dampening effect was
accelerated due to the kinetic coefficient of friction at the contact points where the
UAV is pivoted about its CG.
However, since the effects of each setting have been thoroughly investigated
during the bench test, the process of trimming was sped up during the actual test
Page 32
flight. This was done by Andrew Ong, and the results were collated in Project
AM30. This section will describe the indoor and outdoor flight tests to prove its
capabilities, followed by numerical findings to describe the performance of the
UAV.
8.1) Indoor Flight Test
The objective of the indoor flight test was to test fly the UAV under no wind
conditions, and to execute different the various manoeuvres to exhibit the unique
characteristics of the UAV. In addition, the longitudinal and directional stability
was also tested and it will be discussed in section 9.3 of the thesis.
Fig 24 Wall 'Sticking'
1) Unlike most UAVs in the market, this spherical UAV is able to stick
onto walls, and return to normal flight. This is because the propellers are
encased within the spherical struts which protect the propellers and the
electronics. Furthermore, the struts have been designed to withstand
impacts loads as it may be crash into obstacles within constraint areas.
Page 33
Fig 25 Take off by hand
2) There are two ways which the UAV can take off and land. One is by
holding onto it as shown in Fig 25, and the other, on the ground. However,
the unique capability of this UAV is that it is able to land by rolling onto
the ground, and take off again. This is a critical feature, as it gives UAV
the versatility to take off or land in most conditions. Similarly, the struts
have been designed to withstand impacts loads when landing at heights of
up to 1m, as proven during the drop tests.
Fig 26 Stairs Flight
3) The UAV is also able to translate easily up flights of stairs without the
pilot having to increase the throttle. This proved the above hypothesis in
5.4) Flow Simulation. Since the lift force was able overcome the drag
Page 34
force at that particular angle, resulting in a gentle climb during translation
flight.
8.2) Outdoor Flight Test
Fig 27 Outdoor Flight Path
Although the primary role of the UAV is to fly within constraint areas, the
outdoor flight test was done and the flight path was recorded using a GPS at low
wind conditions. With a larger aerial space, there was more room to manoeuvre
the UAV.
The objectives were to test fly the UAV with a planned flight path, as shown in
Fig 27, and test the endurance of the UAV, to find the maximum flight time. The
pilot will manoeuvre the UAV in a V-shaped path, for as long as the battery can
last and the flight time was recorded when the battery ran completely flat. With an
accuracy of about 3 meters, the GPS will record the latitude, longitude and
altitude co-ordinates in real-time and these can be plotted onto Google Maps,
during post processing of the results. The output values of the pitch, roll and yaw
motions are also recorded with time. With this information, the behaviour of UAV
can be studied.
Page 35
9) Results and Discussion
This section aims to compare results from simulations and actual testing and
prove if the initial hypothesis concurred with the experimental results. The
discussion consists of 3 sub-sections Aerodynamics, Structures and Stability.
9.1) Aerodynamics
9.1.1) Thread Test Results
Table 5 Airflow Velocity (m/s) at different locations
Throttle Settings
Surface Locations 50% 60% 70% 80% 90% 100%
Flaps 13.4 13.9 14.6 15.3 15.9 16.6
Mid-Section 13.4 13.8 14.7 15.3 15.9 16.7
Rudders 13.0 13.5 14.3 15.0 15.7 16.4
Internal Wall 6.1 6.2 6.5 6.7 7.0 7.5
From the Thread Test, it was observed that at maximum throttle setting, the flight
threads continue to flap in a uniform and smooth downward direction in the -Zb
direction, depicting laminar flow on the surfaces. This concurs with the
simulation results in 5.4 Flow Simulation, where the UAV was simulated to
experience an airflow velocity from 13.2m/s to 16.6m/s to observe the flow
behaviour.
Another observation made was the change in the velocity of airflow at the four
different locations. When the anemometer was placed just below the propeller, at
the top flaps, the readings observed were similar to the Motor testing readings.
However, there was a slight decrease in velocity when the anemometer was placed
further downstream, at the mid-section and the bottom of the UAV, just below the
Page 36
rudders. This is due to a loss of kinetic energy of the air as it travels further away
from the propeller, as some of the air escapes into the surroundings.
The internal wall, on the other hand, gave much lower readings, and the flight
threads do not flap as much but still remains smooth and laminar. This is because
airflow at this region does not lie within the exit velocity of the propeller as
shown by the control volume defined in the propeller slipstream.
From the thread test, it can be concluded that the assumptions that were made in
4.2) Propeller Slipstream theory were valid.
9.2) Structures
9.2.1) Drop Test Simulation Results
According to Coloumbs criterion, also known as the maximum normal yield
stress criterion, the effective stress can be defined as:
= MAX ( | | , | | , | | )
The simulation calculates the effective stress when subjected to impact loads at
various drop heights, and compares it with the yield stress of the 3mm depron
foam, y, which was defined when initialising the material properties.
Since the likelihood of an engine failure is low, and the pilot has ample time to
react by throttling up when the UAV is flying to low, thereby cushioning the
impact of the descent. If the UAV starts to lose lift too quickly at heights of more
than 1m above the ground, the resultant catastrophic impact would be inevitable.
The team has decided that a safety factor, X, of 1.5 was sufficient. The safety
factor is given by:
Page 37
= y
1.
2.33 MPa
Therefore, the maximum effective stress must not be greater than 2.33 MPa.
Table 6 Structural Simulation Results
Drop Height (m) Max Effective Stress (N/mm2) or
(MPa)
Possible Failure
0.2 0.7 No
0.5 1.2 No
1.0 2.6 Yes
From Table 6, SolidWorks Motion Simulation showed that the maximum
effective stress experienced during a 1.0m drop height was 2.6MPa, but it has not
failed because it is below the yield stress of depron. However, due to a safety
factor of 1.5, there is a possibility of a failure which may occur at the struts.
9.2.2) Actual Drop Test Results
The results of the actual drop test showed that at heights of 0.2m and 0.5m, there
were no signs of cracks or fractures on the struts of the UAV and the electronics
were completely kept intact. However, when the UAV was released at 1m, one of
the struts experienced brittle fracture, resulting in delamination of the carbon
strips which were originally attached to it. There were no damage to the rudders,
and the electronic components were untouched.
Page 38
The location of the fracture is similar to the simulation results, where high stress
concentration was observed at the lower portion of the struts, depicted by the red
regions, as seen in Fig 28. This shows the congruence in both experimental and
simulation results.
Fig 28 Fracture at lower struts
As mentioned during the drop test simulation, the likelihood of the UAV
experiencing a free fall is extremely low as the pilot will instinctively increase
throttle when the UAV starts to decent too quickly, thereby cushioning the free
fall impact. Therefore, it can be concluded that it is safe for the UAV to drop at a
height of 1m.
9.3) Stability and Control
9.3.1) Longitudinal Stability
Fig 29 Longitudinal Stability Zoom Out
Page 39
Longitudinal stability refers to the UAVs ability to return to its equilibrium due
the pitching motion. The first test was to investigate the time taken for the UAV to
return to its trim position at 0o after the UAV has been manually displaced by the
pilot. This was done by deflecting the relevant control surfaces of the UAV to
bring it to approximately 45o pitch angle, and holding it for approximately 2.8
seconds before releasing the control stick to allow the UAV to return to
equilibrium. An interval of 5 seconds was given between the activation of the
control surfaces.
Fig 30 Longitudinal Stability Zoom In
From the above results, it was observed that at approximately 39 seconds, the
control surfaces for pitching were activated and it took about 1.1 seconds before
the UAV maintained at a pitch angle of about 42o. Subsequently, when the control
surfaces were deactivated, the UAV took about 1.4 seconds before returning to its
trim point. From the sinusoidal graph, it can be observed that the system is
underdamped and the oscillations are exponentially decaying, therefore the
damping ratio is between 0 and 1.
Page 40
9.3.2) Longitudinal Stability with Disturbances
Fig 31 Longitudinal Stability with Disturbances Zoom Out
The next test was to test the longitudinal stability with disturbances. This was
done at hover flight, where the UAV was subjected to sudden displacements
which caused the UAV to pitch in a small angle. The pilot will not input any
controls to bring the UAV back to equilibrium, instead, the flight control board
will sense the sudden change in pitching angle and autonomously send signals to
the relevant control surfaces to counter the disturbances. This test was done to
investigate the time taken for the flight control board to react and finally bringing
the UAV to equilibrium.
Fig 32 Longitudinal Stability with Disturbances Zoom In
From the above, it can be observed that at the point of disturbance at 53.1
seconds, it caused the UAV to pitch at an angle of 44o. The UAV was brought
Page 41
back to equilibrium in 1.4 seconds. This concludes that the system is stable in the
longitudinal direction.
9.3.3) Directional Stability
Fig 33 Directional Stability - Zoom Out
Directional stability refers to the UAVs ability to return to equilibrium due to the
yawing motion. However, unlike the pitching motion, the UAV is unable to return
to 0o or its original position. This is because the UAV does not have a Heading
Sensor, thus when the flight control board senses that the yaw motion of UAV
comes to a standstill, it will stop sending signals to the rudders to compensate the
yaw motion. The experiment was carried out by deflecting the rudders to cause
the UAV to move approximately 180o from its starting position, before releasing
the control stick to allow the UAV to return to equilibrium. An interval of 5
seconds was given between the activation of the rudders in opposite directions,
causing the UAV to yaw in counter clockwise and clockwise directions during
each interval.
Page 42
Fig 34 - Directional Stability - Zoom In
It can be observed that at the starting position of about -90o, the rudders were
activated and it took approximately 1.8 seconds before the UAV reaches about
90o. However, the inertia of the yaw motion caused the UAV to travel to 100
o
even after the control stick was released, and it only reached equilibrium after 2.0
seconds later.
9.3.4) Directional Stability with Disturbance
Fig 35 Directional Stability with Disturbance - Zoom Out
Similarly, the UAV was subjected to a disturbance force, at hover flight, with
each disturbance alternating between clockwise and counter-clockwise directions.
This is to investigate the response of the flight control board to bring the UAV to
Page 43
equilibrium when a sudden disturbance force causes the UAV to displace in a yaw
direction, without the pilots control.
Fig 36 Directional Stability with Disturbance - Zoom In
From the graph above, it can be observed that at the original position of 150
degrees, the UAV was displaced due to an external force at 89.1 seconds, causing
the UAV to yaw towards the 0o yaw angle. However, the rudders were able to
react and counter the yawing motion, bringing the UAV to equilibrium at 35
degrees, within 1.9 seconds. This concludes that UAV has directional stability.
10) UAV Performance
The initial objectives of the UAV were set based on some of the capabilities of the
Japanese Sphere and Prototype 3 was used to make comparisons with it.
Table 7 UAV Performance
Description Initial Objectives Prototype 3
Dimension 42cm diameter 42 cm diameter
Maximum All Up
Weight
Under 600 grams 551 grams
Flight Endurance 8 minutes 7:10 minutes
Target Cost Under SGD 500 SGD 445.43
Target Payload 150 grams 150 grams achievable with good
performance
Others Surveillance Capabilities 5.8 GHz wireless camera
surveillance System
Page 44
From Table 7, it can be observed that Prototype 3 met most of the quantitative
objectives. With an initial estimate of just SGD 500 for the total costs of
constructing a spherical UAV in this project, which was way below the US$1,400
used for the Japanese Sphere, only SGD 445.43 was spent on Prototype 3.
The Maximum All Up Weight was 49 grams lesser than the initial objective
which increase the total allowable payload. The 5.8 GHz wireless camera
surveillance system selected was lightweight and provided surveillance
capabilities which allows for possible reconnaissance or during search and rescue
missions.
Although Prototype 3 fared well most objectives, the flight endurance of the UAV
fell short by 50 seconds. The initial estimate of 8 minutes was calculated based on
the specifications of the motor and battery and the all up weight of the UAV. The
estimated all up weight was increased because of the ArduPilot and extra battery
which were mounted to collect data of the flight test, resulting in a decrease in the
total flight endurance.
11) Conclusion
In conclusion, all the objectives of the project have been met. Through both
theoretical and experimental studies of the aerodynamics and structural
components of the UAV, it can be observed that most of the results concur and
this provides useful information of VTOL UAVs.
This uniquely spherical shaped UAV is definitely a class of its own as it has many
characteristics which sets itself apart from most other UAVs currently in the
market.Running only on a single propeller-motor configuration, the UAV is able
to hover, climb vertically, and transit into translation flight, which is similar to
Page 45
that of a helicopter. Apart from that, this UAV boasts a new concept of take-off
and landing, which is not commonly seen in most of the other UAVs. During
landing, the UAV rolls onto the ground until it comes to a standstill. It can then be
piloted to an upright orientation before taking-off again. With a built-in wireless
camera, the UAV is able to navigate within tight spatial constraints, suitable for
urban surveillance operations. The camera is able to provide aerial view of live
video footage during flight.
12) Recommendations for Further Work
Majority of the body of the UAV is made up of Depron Foam. It is light weight
but has low strength-to-weight ratio. Further work can be done to construct the
UAV using CFPR which has much higher strength-to-weight ratio. However, its
only limitation is the costs and manufacturing of the product.
Page A1
REFERENCES
1. Daniel Moll, J. N. (2008). VTOL UAV A Concept Study. Linkping.
2. Depron, i. d. (2012). Depron Foam Technical Information. Retrieved March 15, 2013, from Depron Foam Web site:
http://depronfoam.com/depron-foam/resource/Depron-White-Technical-
Data-Sheet.pdf
3. Nelson, R. C. (1998). Flight Stability and Automatic Control. McGraw-Hill .
4. Zhao, H. W. (2009). Development of a Dynamic Model of a Ducted.
Page A2
APPENDICES
Appendix A: Depron Foam
Table 8 Technical Properties (Depron, 2012)
3mm 6mm
Compressive Stress (@
10% foam deformation):
0.10MPa 0.15MPa
Tensile Stress (@ break,
length direction):
1.30MPa 0.90MPa
Tensile Stress (@ break,
transverse direction):
0.70MPa 0.90MPa
Elongation (@ break,
length direction):
9% 10%
Elongation (@ break,
transverse direction):
12% 12%
Table 9 SolidWorks Material Input
Page A3
Appendix B: Material Costs for Prototype 3
Table 10 Total Material Costs for Prototype 3 Structures
Description of
Items Qty Unit Price (SGD)
Total Amount
(Include GST) (SGD)
Carbon Rod Flat
5x0.6x1200mm 8 3.00 25.68
Carbon Rod Solid
3x1200mm 2 3.80 8.13
MUMEISHA/
Epoxy 5 Min 1 8.50 9.10
Velcro Peel and
Stick Adhesives
(100x200mm)
1 5.50 5.89
Depron Foam
1500X1000X3mm 1.5 5.75 8.63
Total 57.43
Table 11 Total Costs Spent on Prototype 3
S/N Item
Item Weight
(Grams) Quantity Cost (SGD)
1
Prototype 3 Material
Structures (From Table 8) 217 1 57.43
2 Motor Assembly 76 1 24
3 Propellor 8 1 3
4 Electronic Speed Controller 23 1 18
5 Receiver 8 1 30
6 Battery Eliminator Curcuit 13 1 24
7 Servos 9 6 78
8 kK2 Flight Control Board 22 1 33
9 Wiring 5 1 0
11 Counter Balance Weights 20 1 0
12 Battery 82 1 48
13
Camera System (Incl
Cables) 23 1 130
Total: 551 445.43
Page A1
Table 12 Flow Simulation
0 Degrees 10 Degrees 20 Degrees 30 Degrees 40 Degrees
50 Degrees 60 Degrees 70 Degrees 80 Degrees 90 Degrees
Max Velocity (m/s)
0 m/s
Appendix C: Flow Simulation Results
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