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Broadcasting Autonomous Traffic
Monitoring Aerial Network
(BATMAN)
Mei Cheong
Michael Evans
Elise Fahy
Amelia Greig
Natasha Parker
ii
iii
Disclaimer
We the authors declare that the following work is our own unless otherwise stated and must
not be reproduced or copied without the appropriate permission or recognition.
………………………………………………………………
Mei Cheong Date:
1150802
………………………………………………………………
Michael Evans Date:
1164762
………………………………………………………………
Elise Fahy Date:
1161673
………………………………………………………………
Amelia Greig Date:
1149212
………………………………………………………………
Natasha Parker Date:
1153729
iv
Marking Scheme
Group mark
Criteria Mark (total 100)
1. Project definition /10
2. Research activities /15
3. technical calculation /25
4. Drawings /25
5. Format of the report /10
6. Novelty of the solution /15
Project mark
Group member Group mark
(50% x project mark)
Individual mark
(50% x project mark)
Project mark
(Total 100)
1. M Cheong
2. M Evans
3. E Fahy
4. A Greig
5. N Parker
v
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Executive Summary
Traffic congestion in major cities has become a serious problem in recent years. Various
methods to reduce congestion have been implemented such as traffic cameras, road tolls and
aerial traffic monitoring. Aerial traffic monitoring involves flying a camera over the congested
areas, currently by the use of a specially designed helicopter, to divert traffic to less congested
routes. Costs of aerial traffic monitoring could be reduced dramatically by the use of an
unmanned aerial vehicle (UAV) fitted with a high resolution camera, to replace the currently
used helicopters. This project summarises the design of such a UAV.
The UAV is required to be a flying wing configuration, suited to fly at a relatively low altitude
for a period of approximately three hours over densely populated metropolitan areas. The
aircraft will also be designed to be as environmentally friendly as possible, to suit the ever
increasing environmentally based modern society.
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viii
Contents
1 Introduction............................................................................................................................1
2 Literature Review....................................................................................................................3
2.1 Unmanned Aerial Vehicles ...............................................................................................3
2.1.1 Flying Wing UAVs ......................................................................................................4
2.2 Traffic Monitoring ............................................................................................................4
2.2.1 Governmental Traffic Monitoring Cameras ................................................................5
2.2.1 Media Aerial Traffic Monitoring.................................................................................5
2.3 Feasibility Study ...............................................................................................................7
2.4 Statistical Analysis ............................................................................................................8
2.4.1 Empty Weight............................................................................................................9
2.4.2 Wing Span ...............................................................................................................11
2.4.3 Length .....................................................................................................................12
2.4.4 Cruise Speed............................................................................................................12
2.4.5 Endurance ...............................................................................................................13
2.5 Technical Task ................................................................................................................14
2.5.1 Standards ................................................................................................................15
2.5.2 Performance Parameters.........................................................................................16
2.5.2.1 Payload.............................................................................................................16
2.5.2.2 Endurance ........................................................................................................16
2.5.2.3 Range ...............................................................................................................17
2.5.2.4 Speed ...............................................................................................................22
ix
2.5.2.5 Cruise Altitude ................................................................................................. 23
2.5.2.6 Take-off/Landing.............................................................................................. 26
2.5.2.7 Operating Conditions ....................................................................................... 27
2.5.3 Technical level of the product ................................................................................. 28
2.5.4 Economical Analysis................................................................................................ 28
2.5.5 Propulsion Systems................................................................................................. 29
2.5.6 System Requirements ............................................................................................. 29
2.5.7 Reliability and Maintainability................................................................................. 30
3 Conceptual Design ............................................................................................................... 31
3.1 Concept Designs ............................................................................................................ 31
3.1.1 Concept Design 1 – Blended Wing Body.................................................................. 32
3.1.2 Concept Design 2 – Semi-autonomous UAV ............................................................ 33
3.1.3 Concept Design 3 – Simple square edge.................................................................. 34
3.2 Mission Profiles ............................................................................................................. 35
3.3 Weight analysis ............................................................................................................. 36
3.3.1 Avionic Weight Estimation ...................................................................................... 36
3.3.2 Structural Weight Estimation .................................................................................. 37
3.4 Sensitivity...................................................................................................................... 38
3.5 Sizing............................................................................................................................. 40
3.5.1 Sizing to Stall Speed................................................................................................ 40
3.5.2 Drag Polar Estimation ............................................................................................. 41
3.5.3 Sizing to Climb Requirements.................................................................................. 42
3.5.4 FAR 23.65 (AEO) Sizing to Rate of Climb.................................................................. 42
3.5.5 FAR 23.77 (AEO) Sizing for balked landing climb requirements................................ 43
x
3.5.6 FAR 23.65 (AEO) Climb Gradient Sizing ....................................................................44
3.5.7 Climb Sizing Summary..............................................................................................44
3.5.8 Sizing to Cruise Speed..............................................................................................45
3.5.9 Matching Diagram ...................................................................................................45
3.6 Configuration/Planform .................................................................................................47
3.7 Airfoil .............................................................................................................................50
3.7.1 CFD Analysis ................................................................................................................55
3.8 Control Surfaces.............................................................................................................57
3.9 Propulsion......................................................................................................................59
3.9.1 Propellers ................................................................................................................60
3.9.2 Piston Engines .........................................................................................................60
3.9.3 Electric Motors ........................................................................................................61
3.9.4 Solar Powered UAVs................................................................................................61
3.9.5 Jet propulsion..........................................................................................................62
3.9.6 Propulsion System Selection....................................................................................63
3.10 Launch .........................................................................................................................65
3.11 Landing ........................................................................................................................67
3.11.1 Emergency Landing Systems..................................................................................68
3.12 Avionics........................................................................................................................69
3.12.1 Data Link ...............................................................................................................69
3.12.2 Autopilot and Navigation Systems .........................................................................71
3.12.3 Control Actuation Systems.....................................................................................73
3.12.4 Sensors..................................................................................................................75
3.12.5 Electrical Systems ..................................................................................................76
xi
3.13 Materials ..................................................................................................................... 79
3.14 Manufacture ............................................................................................................... 81
3.14.1 Reusable Moulds .................................................................................................. 81
3.14.2 Hand lay-up and pre-impregnated cloth................................................................ 84
3.15 Maintenance ............................................................................................................... 85
3.16 Structural Analysis ....................................................................................................... 86
3.16.1 Wing Structure...................................................................................................... 86
3.16.2 Housing Structure ................................................................................................. 87
4 Weight and Balance Analysis ................................................................................................ 89
4.1 Internal Component Configuration ................................................................................ 90
4.2 Centre of Gravity Determination ................................................................................... 92
4.3 Longitudinal Stability ..................................................................................................... 94
5 Aerodynamic Analysis .......................................................................................................... 97
5.1 Lift Distribution ............................................................................................................. 97
5.2 L/D Determination......................................................................................................... 99
6 Performance Analysis......................................................................................................... 101
6.1 Weight, Wing Loading and Power Loading................................................................... 101
6.2 Sensitivity to new values ............................................................................................. 103
6.3 Cruise Speed and Stall Speed....................................................................................... 104
6.4 Summary..................................................................................................................... 105
7 Technical Drawings ............................................................................................................ 107
7.1 Three View Drawings................................................................................................... 107
7.2 Aircraft Layout Drawing............................................................................................... 107
7.3 Exploded View Assembly Drawing ............................................................................... 107
xii
7.4 Wing Detail Drawing.....................................................................................................107
7.5 Airfoil Drawing .............................................................................................................107
References.............................................................................................................................109
Appendix A – Raw Data for Statistical Analysis .......................................................................119
Appendix B – VTC Charts for various Australian Cities ............................................................121
Appendix C – Hand Calculations.............................................................................................125
C1 – Empirical constants A and B........................................................................................125
C2 – Sizing Calculations ......................................................................................................125
C3 - Centre of Gravity Calculations .....................................................................................127
C4 - Lift Distribution ...........................................................................................................128
C5- L/D Determination .......................................................................................................129
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List of Figures
Figure 2.1: Technology Diagram - Empty Weight ......................................................................10
Figure 2.2: Technology Diagram – Wing Span...........................................................................11
Figure 2.3: Technology Diagram – Length.................................................................................12
Figure 2.4: Technology Diagram – Cruise Speed .......................................................................13
Figure 2.5: Technology Diagram – Endurance...........................................................................14
Figure 2.6: Single range radii marked over city satellite images ................................................19
Figure 2.7: Multiple range radii marked over city satellite images ............................................21
Figure 2.7: A visual terminal chart for the Adelaide airport region............................................23
Figure 2.7: Diagrams of view radius for various altitudes..........................................................26
Figure 3.1: Concept design 1 – Blended Wing Body ..................................................................32
Figure 3.2: Concept design 2 – Semi autonomous UAV.............................................................33
Figure 3.3: Concept design 3 – simple square edge ..................................................................34
Figure 3.4a: Mission Profile for loiter above a city during peak traffic periods ..........................35
Figure 3.4b: Mission profile for cruise out to a remote area, loiter and cruise back ..................35
Figure 3.5: Subsystem mass breakdown of a UAV ....................................................................38
Figure 3.6: Matching diagram ..................................................................................................46
Figure 3.8: Airfoil with reflexed camber line.............................................................................51
Figure 3.9: Airfoil coefficients against angle of attack plot........................................................52
Figure 3.10: Drag polars for prototype airfoils..........................................................................54
Figure 3.11: Eppler 342 airfoil cross-section .............................................................................55
Figure 3.12 : Velocity profileas .................................................................................................56
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Figure 3.13: Pressure profile ................................................................................................... 56
Figure 3.14: The Baldor premium efficiency EM3538 electric AC motor................................... 63
Figure 3.15: ScanEagle UAV catapult launcher......................................................................... 66
Figure 3.16: wing structural components ................................................................................ 87
Figure 4.1: Schematic of Internal Component Configuration Layout ........................................ 92
Figure 4.2: CG Envelope .......................................................................................................... 94
Figure 4.3: Longitudinal Static Margin versus Leading Edge Sweep Angle ................................ 95
Figure 5.1: Span-wise lift distribution along the wing .............................................................. 98
Figure 5.2: Statistical L/D determination graph...................................................................... 100
xvi
List of Tables
Table 2.1: R44 Raven II Specifications.........................................................................................6
Table 2.2: Comparative study of Historical data .........................................................................9
Table 2.3: Take off and empty weight of selected comparison UAVs ........................................10
Table 2.4: Cities grouped by size (adapted from Australian Bureau of Statistics 2009)..............17
Table 2.5: Main on-board avionics systems ..............................................................................30
Table 3.1 – Avionics systems weight distribution......................................................................37
Table 3.2: Configuration drag polars and associated coefficients..............................................41
Table 3.2: Climb sizing summary ..............................................................................................44
Table 3.3: Wing geometry equations and values ......................................................................49
Table 3.3: Comparison of airfoils..............................................................................................52
Table 3.4: Control surface deflection states for various aircraft manoeuvres............................59
Table 3.5: Commercial UAV Launcher specifications ................................................................66
Table 3.5 Data Link Systems.....................................................................................................70
Table 3.6: Autopilot Navigation Systems ..................................................................................72
Table3.7: Servo Motor Specification.........................................................................................74
Table 3.8: UAV Sensor Specifications .......................................................................................75
Table 3.9: UAV Electrical Power Requirements.........................................................................77
Table 3.10: Specifications of Batteries Considered ...................................................................78
Table 3.11 Comparison of Materials.........................................................................................80
Table 3.12: Manufacturing processes.......................................................................................82
xvii
Table 4.1: Aircraft Weight Summary........................................................................................ 89
Table 6.1: Comparison of weight, area and loading values..................................................... 102
Figure 6.2: New matching diagram and design point ............................................................. 103
Table 6.3: Sensitivity parameters .......................................................................................... 103
Table 6.4: New cruise and climb values ................................................................................. 104
Table 6.5: Compliance of performance parameters ............................................................... 105
xviii
Nomenclature
Acronyms
AR - Aspect Ratio
CASA - Civil Aviation Safety Authority
CASR - Civil Aviation Safety Regulations
FOV - Field of View
GC - Centre of Gravity
GCN - Guidance and control interface
CGR - Climb Gradient Ratio
GPS - Global Positioning System
MAC - Mean Aerodynamic Chord/Mean Aerodynamic Centre
MAV - Micro Aerial Vehicle
MTOW - Maximum Takeoff Weight
RPM - Revolutions per Minute
STOL - Short Takeoff and Landing
UAV - Unmanned Aerial Vehicle
VTC - Visual Terminal Chart
VTOL - Vertical Takeoff and Landing
xix
Symbols
A - Aspect Ratio
b - Wing span
c - Chord length
CDo - Drag Polar
CD - Drag coefficient
CL - Lift coefficient
d - Diameter
D - Drag
E - Endurance
e - Oswald efficiency factor
g - Gravitational constant, 9.81m/s^2
I - Moment of inertia
L - Lift
L/D - Lift to drag ratio
n - load factor
ns - safety factor
P - Power
q - Dynamic pressure
R - Range
Re - Reynolds Number
S - Wing Area
xx
SM - Static margin
T - Thrust
V - Velocity
Vcr - Cruise speed
Vto - Takeoff speed
W - Weight
x - Longitudinal axis position
y - Lateral axis/span-wise position
z - Vertical axis position
α - Angle of attack
γ - Climb angle
ε - Efficiency
λ - Taper ratio
μ - Friction co-efficient
ν - Viscosity
ρ - Air density
σ - Density ratio
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Subscripts
airfoil - airfoil
battery - battery
cl - climb
cr - cruise
des - descent
e - empty
f - final
i - initial
l - loiter
los - zero lift produced at wing root
max - maximum
min - minimum
OL - zero lift
p - propeller
p - prop propeller
payload - payload
root - root
stall - stall
static - static
TO - takeoff
v - vertical
xxii
w - main wing
wet - wetted area
wing - main wing section
x - relating to the x axis
y - relating to the y axis
z - relating to the z axis
xxiii
1
1 Introduction
With an increasing number of cars on the road in major metropolitan areas, city traffic
congestion is developing into a serious problem in some areas. Upgrading of
infrastructure is not necessarily a viable option due to space and economical
constraints, therefore other methods to reduce traffic congestion need to be
considered. One such option is that of traffic monitoring through the use of aerial
vehicles or permanent traffic cameras. Traffic monitoring allows motorists to be
warned about the most congested areas in advance, allowing them to pursue alternate
routes, reducing the volume of traffic in problem areas. Another important application
of traffic monitoring is that emergency services vehicles can be deployed along routes
with less traffic as to improve response times.
Aerial traffic monitoring involves an airborne craft surveying the terrain below by the
use of cameras. Currently, aerial traffic monitoring is conducted by specially designed
helicopters capable of carrying a cameraman. A possible cheaper option for aerial
traffic monitoring is the use of specially designed unmanned aerial vehicles (UAVs).
Such a UAV could be not only cheaper than the current helicopters, but also smaller
and quieter. Currently there are no such UAVs in operation.
2
The objective of this project is to design a flying wing UAV for the application of traffic
monitoring. The design process begins with a literature review of current traffic
monitoring methods, followed by a statistic analysis of current UAVs used for similar
applications, before progressing to calculations for each design component, resulting in
the overall design being finalised.
This report does not include information specifically on manufacturing and testing of
the aircraft as this is beyond the scope of the project. Also, some simplifications and
assumptions have been made to simplify the project to a level commensurate to the
current expertise of the design team, able to be complete in the given timeframe.
3
2 Literature Review
A literature review of both unmanned aerial vehicles (UAVs) and current traffic monitoring
methods was conducted with the results summarised below. Several UAVs used for similar
purposes were analysed and statistical graphs of common design features created for use
throughout the design process. Finally, the important requirements of the design were decided
upon and outlined.
2.1 Unmanned Aerial Vehicles
An Unmanned Aerial Vehicle (UAV) is an aerial vehicle that does not require an
onboard crew to operate. It is therefore a useful tool in many operations, especially
military missions such as surveillance over enemy territory or payload drops (MSN
Encarta 2009), as the risk to human life is minimised. UAVs have decreased the
necessity of personnel (particularly in military applications) being put in dangerous
situations as well as decreasing the number of personnel required for such missions.
UAVs can also be extensively used in civil environments such as surveillance and aerial
video transmission due their fast transit ability over any terrain and long range
communications (SPAWAR 2004). Generally UAVs are split into three categories based
on weight, Micro Aerial Vehicles (MAVs) weigh less than 400g, small UAVs ranging from
600g to 2000kg and large UAVs that weight greater than 2000kg. In addition to this
4
UAVs can be split into two main designs, fixed wing and flying wing. A fixed wing UAV
has wings attached to a fuselage that provide lift from the forward momentum of the
aircraft and are the most common form of UAV currently available (MSN Encarta 2009).
2.1.1 Flying Wing UAVs
A flying wing UAV an unmanned aircraft whereby the tail and the fuselage sections are
abandoned and all the necessary components are fitted into a large wing section. Flying
wing UAVs are currently not used extensively as the technology is fairly new onto the
market. Research suggests that flying wing UAVs are simpler to manufacture, more
versatile with the design (Warrick 2008) and have an increased endurance over fixed
wing design (Gizmag 2010). BAE Systems is one of the few companies that have
developed a flying wing UAV, Corax UCAV, which will be used for long endurance and
high reconnaissance missions (Air-attack.com 2010).
2.2 Traffic Monitoring
Traffic flows in both highly populated and remote areas are dynamic and uncertain
environments. Traffic monitoring systems that can accurately and quickly identify
potential hazards, congested areas and blockages can reduce the effect on traffic flow.
Warning motorists about congested areas can improve travel times and stop adverse
situations from developing further by diverting the main traffic flow to an alternate
route. Being able to avoid heavy traffic areas will also improve the response times of
emergency services vehicles. Currently traffic monitoring and reports are performed by
two main mediums, governmental departments and the media.
5
2.2.1 Governmental Traffic Monitoring Cameras
The government and associated local traffic authorities use fixed cameras, located on
major roads to provide real time traffic monitoring. Currently, there are 23 cameras in
Adelaide (Transport SA 2010), 24 cameras in Sydney (Road and Traffic Authority, NSW
n.d.), 24 cameras in Melbourne (VicRoads n.d.) and 21 cameras in Brisbane and
surrounding areas (The State of Queensland Department of Main Roads 2008). Traffic
monitoring cameras are also widely used across the globe. Currently, most large cities
have incorporated cameras for a number of applications most commonly for
monitoring traffic congestion. The Surveillance Camera Players estimate that the
cameras used for such applications can cost up to $90,000USD which results in a rough
cost of $9,000,000USD operational costs annually. It must be noted that these cost
estimates are based on monitoring systems in New York City, USA however this cost
can be used as a comparison to show that the overall cost of using a UAV for traffic
monitoring can be significantly less.
2.2.1 Media Aerial Traffic Monitoring
Media outlets such as news programs and radio stations provide regular traffic updates
during peak periods, such as the morning and afternoon peak hours and during serious
accident events, to warn motorists to avoid certain areas. These reports are provided
either by the general public telephoning the station or by aerial traffic monitoring.
Aerial traffic monitoring over major cities currently involves a crew in a helicopter flying
over the city providing reports through a video link to the ground station. Australia has
6
only one dedicated airborne traffic monitoring company, The Australian Traffic
Network (ATN). The ATN supplies airborne traffic information to 95% of Commercial
Metropolitan Radio Stations, Channel 10 evening news in Sydney, Brisbane and
Melbourne and the breakfast news shows on Channel 7 and Channel 9 (The Australian
Helicopter Directory 2010). Current airborne traffic monitoring is mostly reliant on
specially equipped news helicopters or ‘newscopters’ such as the R44 Raven II
Newscopter, of which the specifications are shown in Table 2.1 below.
Table 2.1: R44 Raven II Specifications
Satellite based monitoring systems have also been implemented in some regions,
however, the major limitation of such a system is that the monitoring provided is
transitory which makes continuous traffic flow a difficult task to track (Runge et al.
2005). This is also quite an expensive option due to the large costs associated with the
manufacture and launch of satellites.
Take-off Weight 2 500lb
Hover Ceiling 8 950 ft
Cruise Speed 113 kts
Range 484 km
Fuel Consumption 0.173 kg/km
Price $744 000 USD
7
2.3 Feasibility Study
There are currently no documented studies of UAVs being used for regular
metropolitan traffic monitoring due to wide-spread regulations prohibiting UAV flight
over populated areas (Ro, Oh & Dong 2007, p1). There have however been a number of
studies into the use of UAVs for surveillance and reconnaissance applications. The
University of Ohio has designed, built and tested surveillance UAV that could achieve a
velocity of 30mph at an altitude of 500ft transmitting data through a 2.4 GHz data link
(Puri 2005). With significantly shorter response time than manned aircraft and better
manoeuvrability than ground based fixed cameras (Puri 2005), the use of UAVs in
regular metropolitan traffic monitoring would seem feasible, pending the update of
regulations concerning unmanned aircraft flying over densely populated areas.
Aside from surveillance and reconnaissance operations, the application of UAV
technology has also been investigated for use in wildlife research and monitoring.
Researchers from the University of Florida proposed the use of UAVs for this
application in 2003 (Jones) and subsequently ran trials the following year (Lee 2004).
These trials were deemed to be successful and such a system viable, should it gain the
support of the wider scientific community (Lee 2004).
The technical constraint that a UAV for this application should have a flying-wing
configuration should be achievable noting that similar designs, such as the Korean
Blended Wing Body (BWB) UAV (Huh & Shim 2009), has previously been shown to be
durable and controllable in similar applications to traffic monitoring. Such a fixed-wing
configuration would also overcome issues of vibration, and therefore image distortion,
encountered by rotary wing surveillance UAVs (Puri 2005).
8
2.4 Statistical Analysis
To allow a basis for the important parameters in the UAV design and to gain a better
understanding of current UAVs on the market a statistical analysis was performed. This
allowed a more feasible design to be created. The statistical analysis included the
critical design parameters, empty weight, wing span, cruise speed, length of the UAV
and endurance.
A comparative study of existing UAVs was performed to create a historical basis for the
project design. Through research, it was noted that the number of flying wing UAVs
currently available for a similar mission were limited. Therefore, in order to improve
the reliability level of the statistics obtained, fixed wing UAVs with a similar mission
were also examined. The data was obtained from an online aircraft database, Flight
Global Aircraft Directory, using data sheets of the respective UAVs which provided the
various aircraft specifications. A list of UAVs considered is shown in Table 2.2 below.
These UAVs were specifically chosen for this statistical analysis as they have all been
designed for similar missions to traffic monitoring. Thus the data obtained will give a
reasonable estimate of the values that should be obtained for a traffic monitoring UAV.
9
Table 2.2: Comparative study of Historical data
Manufacturer Model Range Class
AAI Corporation RQ-7B Shadow 200 Tactical
AAI Corporation Aerosonde 4 LALE
Pioneer UAV RQ-2B Pioneer Tactical
ATE Vulture Tactical
Aerovision Fulmar Tactical
BlueBird Aerosystems Micro B
BAE Systems Phoenix Tactical
BAE Systems Coyote Close Range
An LALE UAV is a Low Altitude and Long Endurance UAV that is generally used for long
endurance missions where continuous sensing is required. Tactical and close range
UAVs are used for low altitude missions and are primarily used with the troops at the
front line with a short endurance (7-20hrs). They are self-sustainable in the field and
are normally very light weight (Defense Update, 2005).
2.4.1 Empty Weight
A basic equation relating the empty and take off weights of an aircraft was obtained
from Roskam (2005) where WO is the takeoff weight and We is the empty weight.
(Equation 2.1)
The variables A and B are empirical constants that are dependent on the type of
aircraft being analysed. The values for these variables are established based on data
collected in the comparative study conducted. For this particular UAV, a technology
diagram which graphs logWo against logWe was created and a linear trend line drawn.
10
Comparing the trend line equation to Roskam’s equation the constants A and B were
determined.
The eight comparison UAVs from Table 2.2 and their respective empty and maximum
take off weights are listed in Table 2.3 below. The logs of these values were taken and
the technology diagram for empty weight of the UAV constructed, as shown in Figure
2.1 below.
Table 2.3: Take off and empty weight of selected comparison UAVs
Model Max Take-Off Weight (kg) log(Wo) Empty Weight (kg) log(We)
RQ-7B Shadow 200 170 2.230449 200.6 2.302330929
Aerosonde 4 12 1.079181 18.1 1.257678575
RQ-2B Pioneer 204.12 2.309886 276 2.440909082
Vulture 150 2.176091 115 2.06069784
Fulmar 19 1.278754 11.5 1.06069784
Micro B 1.1 0.041393 1.1 0.041392685
Phoenix 175 2.243038 220 2.342422681
Coyote 5.5 0.740363 5.5 0.740362689
Figure 2.1: Technology Diagram - Empty Weight
11
The technology diagram for empty weight demonstrates a statistical trend to the
relationship of aircraft empty weight versus maximum take off weight. From the line of
best fit shown in the diagram the empirical constants A and B can be found (For
calculations, see Appendix C):
A= 0.01825
B = 0.97609
Therefore logWo = 0.01825 + 0.97609 logWe (Equation 2.2)
2.4.2 Wing Span
The wing spans of the eight comparison UAVs were graphed against their takeoff
weights in a log-log format with the results shown in Figure 2.2. A line of best fit for the
data points indicates a nearly linear relationship between the wing span of the UAVs
with their takeoff weights. From this trend line, a rough estimation of the wing span of
the UAV design can be obtained. The values of wing span for each of the eight
comparison UAVs can be found in Appendix A.
Figure 2.2: Technology Diagram – Wing Span
12
2.4.3 Length
The logs of the lengths of the UAVs were plotted against the logs of the respective
takeoff weights with the results shown in Figure 2.3. The line of best fit shown on the
graph provides a good estimation for aircraft length when designing the UAV. The
values of UAV length for each of the eight comparison UAVs can be found in Appendix
A.
Figure 2.3: Technology Diagram – Length
2.4.4 Cruise Speed
Due to the nature of the mission profile of the UAV design, the cruise speed of the
aircraft retains a significant importance to the overall concept design. As such, a
technology diagram comparing information regarding existing UAVs’ design cruise
speeds and takeoff weights was constructed and is shown in Figure 2.4 below. The logs
of each axis were used to provide a linear relationship, in turn providing a quick
13
reference to feasible design parameters. The values of cruise speed for each of the
eight comparison UAVs can be found in Appendix A.
Figure 2.4: Technology Diagram – Cruise Speed
2.4.5 Endurance
Aircraft endurance is the final design parameter for this statistical analysis. A
technology diagram for the endurance of the eight existing UAVs was constructed and
is shown in Figure 2.5 below. The relationship using a linear line of best fit is not as
good as the previous technology diagrams; however the results can still be used for an
approximate estimate of design endurance. The values of endurance for each of the
eight comparison UAVs can be found in Appendix A.
14
Figure 2.5: Technology Diagram – Endurance
2.5 Technical Task
The interest in UAVs has grown extensively with the recent rapid developments in
avionics and micro-scale technologies, and their increasing contribution to the aviation
industry. UAVs are generally defined as a reusable, unmanned vehicle capable of
controlled, sustained, level flight and mission capable. Previously limited to military
applications, civil usage of UAVs is gaining public appeal as an alternative method of
enhancing traffic monitoring systems.
UAVs therefore present the potential to be implemented as a flexible and highly
responsive mobile aerial system suitable for traffic monitoring purposes. Aside from
enhancing current traffic monitoring systems, these UAVs would enable prompt
emergency service responses and provide a cost effective method to monitor traffic in
15
rural and remote areas that were previously economically unfeasible. The purpose of
this work is a design study of a small flying wing UAV capable of supporting current
traffic monitoring systems. The small flying wing UAV should be low cost, easy to
manufacture, operate and maintain and provide a quick, non-permanent mobile
surveillance capability.
2.5.1 Standards
The Civil Aviation Safety Authority (CASA) provides the regulations for all aircraft and
aviation in Australia. The most important standards are the Civil Aviation Safety
Regulation (CASR) Part 21, a general section about certification procedures for aircraft
products and parts, and CASR Part 101 subpart 101.F: regulations for UAVs. The main
points to consider for the design of a flying wing UAV for traffic monitoring are:
- UAVs cannot currently be operated over populous areas. That is, an area where
there is sufficient population density that in the event of a fault or failure, an
unreasonable risk is posed to life, safety or property.
- A person must not operate a UAV within 30 metres of another person who is
not directly associated with the operation of the UAV.
- UAVs must be operated within a CASA-approved area, and may only be
operated outside this area if away from populous areas.
- Larger UAVs, currently defined as being over 150kg, must have an airworthiness
or experimental certificate and maintenance must be approved by CASA.
- Presently, there are no airworthiness, manufacturing or maintenance
regulations specifically for small UAVs.
16
CASA are currently working on updating the UAV standards (Carr 2007), especially
design standards and legal certification, but these are not high priority and could take
several years to be produced.
In terms of the UAV being designed in this task, it will be classified as a small UAV under
CASA’s current standards, so there are currently no design standards to adhere to. At
present, the aim of this UAV, which is to monitor traffic in cities, would be prohibited
by CASA since UAVs cannot be operated over populous areas. For this project however,
it will be assumed that the regulations will be changed to allow such operation, so that
this UAV would be able to operate and hence there would be a market for it.
2.5.2 Performance Parameters
2.5.2.1 Payload
As the aircraft is a flying wing UAV, there is no crew or passenger payload. The avionics
systems, including the camera, are included in the gross weight of the aircraft and are
not considered payload. Therefore, for this aircraft and mission there is no payload to
consider.
2.5.2.2 Endurance
Most peak traffic periods, such as the morning and afternoon ‘peak hours,’ are
generally no longer in duration than 3 hours. The morning peak hour is assumed to
begin at 6:30am and continue to 9:30am while the afternoon peak hour is assumed to
begin at 4:00pm and continue until 7:00pm in major cities. Also, any major traffic
incident usually affects traffic for only one or two hours depending on the severity and
magnitude of the incident. Therefore, the flying wing UAV will be designed for a
17
minimum endurance of 3 hours. This will allow the UAV to operate continuously
throughout both peak traffic periods and monitor most major traffic incidents for the
duration.
2.5.2.3 Range
The flying wing UAV for traffic monitoring will be designed for use in major Australian
cities, since the layout of these cities is familiar to the design team and easily
accessible. The capital cities of each Australian state and territory have been grouped
into three size designations according to area and population, since, in Australian cities,
there is a correlation between the two properties. These groupings can be seen in table
2.4.
Table 2.4: Cities grouped by size (adapted from Australian Bureau of Statistics 2009)
Size City
Large
(> 3 million people)
Sydney
Melbourne
Medium
(1-3 million people)
Brisbane
Perth
Adelaide
Small
(< 1 million people)
Canberra
Hobart
Darwin
To estimate flight distances for the UAV, circles with a radius of 10km or 15km were
drawn onto areas of city satellite images (Google Maps 2010). The circumference of
these circles approximates the distance flown in one pass over the city, and the design
will be intended for one pass per hour, equivalent to one revolution through the circle
on the map. This method was selected for estimating distance because mapping
18
specific flight paths was deemed unnecessary for this part of the project. The UAV,
when in use, could be sent along main roads and track where traffic looks to be built up
in real time, allowing an operator or inbuilt flight control system to guide it to where
monitoring is needed. Also, in the case of an emergency, for example a road accident,
the UAV could receive orders from the ground to monitor that region. Figures 2.6 a) to
c) below show the placement of circles of 10km and 15km radii over a city from each
size category. The orange circles represent the 10km radii and the red circles the 15km
radii.
(a)
19
(b)
(c)
Figure 2.6: Single range radii marked over city satellite images for a) a large city (Sydney), b) a medium city
(Adelaide) and c) a small city (Hobart) (Google 2010)
20
By observation of the city layouts with a single circle, three UAVs would be
recommended for large cities, two for medium cities and one for small cities. The maps
were adjusted to show the new coverage of the UAVs, as shown in figure 2.7 a) to c).
For Sydney, one UAV could operate north of the harbour, one UAV to the south and
one UAV over the western suburbs so that all of the main roads in and out of the
central business district are within range of a UAV.
(a)
21
(b)
(c)
Figure 2.7: Multiple range radii marked over city satellite images for a) a large city (Sydney), b) a medium city
(Adelaide) and c) a small city (Hobart) (Google 2010)
22
Assuming two passes of the region every hour to provide regular thirty minute traffic
updates, the total range can be estimated as the circumference of a circle over the
region, travelled twice every hour for three hours.
15km Radius:
kmhrskmR 5663152215 =×××= π
10km Radius:
kmhrskmR 3773102210 =×××= π
2.5.2.4 Speed
In order to provide regular traffic updates, the UAV needs to complete two passes of
the region each hour. As above, assuming the flight path to be a circle of either a 10km
or a 15km radius, the speed is required to be sufficient to allow the UAV to complete
two circles of the circumference every hour.
hrkmhr
kmV /0.189
1
152215 =××= π
hrkmhr
kmV /6.125
1
102210 =××= π
Therefore, the UAV must travel at a velocity of 189km/hr or 102kts for a circular flight
path of 15km radius, or velocity of 126km/hr or 68kts for a circular flight path of 10km
radius. From the previous statistical data collected about similar UAVs, a cruise speed
of 102kts for a small UAV is quite high and could make the design more complicated
and expensive to satisfy. Therefore the UAV will be designed to operate over a 10km
radius, assuming a circular flight path.
23
2.5.2.5 Cruise Altitude
The flying wing UAV should be designed to operate outside of controlled airspace to
enable easier operating conditions. Controlled airspace floor varies in height depending
on location and proximity to nearby airports. In the immediate vicinity of an airport,
controlled airspace extends to the ground (Air Services Australia 2010a). Away from
airports, the floor height of controlled airspace varies depending on the surroundings.
Visual Terminal Charts (VTCs) of Australian airports provide a visual representation of
the controlled airspace coverage around the airport. The VTC for Adelaide is shown in
figure 2.7 below with additional VTC’s for Hobart, Canberra, Sydney and Melbourne
provided in Appendix B. An analysis of the controlled airspace requirements
throughout these regions is outlined below.
Figure 2.7: A visual terminal chart for the Adelaide airport region
24
There is a large circular region surrounding Adelaide airport where the controlled
airspace extends to ground level. This region covers the city centre, as well as the
majority of the southern suburbs. To the north of the city, there is another smaller
region where controlled airspace extends to the ground in the area surrounding
Parafield airfield. In addition, there is a prohibited area to the north of Parafield airfield
due to the military base, Edinburgh. As the majority of traffic flow problems will occur
around the city centre and immediate surrounds, neither the Parafield airfield nor
Edinburgh prohibited areas will affect operation of the UAV. However, as the Adelaide
airport controlled airspace region extends completely over the city and immediate
surrounds, the UAV would need to operate within the controlled airspace in the
Adelaide region.
For the remaining cities, the results are summarised below.
Sydney: Controlled airspace due to Sydney Airport extend to ground level and covers
the region south of the harbour, from the shoreline to inland to the Bankstown region.
The immediate city region north of the harbour is covered by a controlled airspace
floor level of 500ft, with the northern suburbs covered by two regions of 700ft and
1000ft. Over the far western suburbs, there are a large assortment of prohibited areas
intermixed with a controlled airspace region with a floor level of 2500ft.
Melbourne: Controlled airspace extends to ground level over the northern suburbs,
western suburbs and the western side of the city due to Tullamarine airport. There is
another smaller region of ground level airspace in the south east due to Moorabbin
airport. The south and east of city centre are covered by a 1500ft controlled airspace
floor level. The reminding suburbs, including Port Melbourne, are covered by a 2500ft
controlled airspace floor level. There are large prohibited areas over Port Philip Bay and
25
to the South East of the greater city region, however neither of these regions will affect
the UAV’s flight path.
Canberra: Controlled airspace for Canberra Airport extends to ground level covering
the city centre and immediate surrounding suburbs. The remaining suburbs to the
north, west and south are covered by a controlled airspace floor height of 3500ft.
There are two prohibited areas, one small region north of airport and a larger region to
south east however, neither will affect the UAV’s flight path.
Hobart: Any city region to the north of the Derwent river is covered by ground level
controlled airspace. Any region south of the river is outside controlled airspace, or in
region A, in which the controlled airspace is only above 40000ft (Air Services Australia
2010a).
For all five cities analysed, there is a large region of the city covered by controlled
airspace the extends to ground level. In these areas the UAV would be flying within
controlled airspace regardless of the altitude. Over the remaining areas, the controlled
airspace generally has a floor level of 2500ft, therefore the UAV should be designed to
operate at an altitude less that 2500ft to meet with airspace requirements.
In order to visually monitor traffic, the UAV must operate below the cloud base.
According to a 1988 report prepared for the United States Department of Energy, low
level cloud base heights over land are approximately 480, 520 and 590 metres for
cumulonimbus, stratus and stratocumulus (excluding fog) and cumulus clouds
respectively (Warren et al 1988, p. 33b). In foggy conditions the UAV would not be able
to maintain visual contact with the ground, not only rendering it incapable of
monitoring traffic but posing a potential safety risk. From these values, the predicted
operating altitude of the UAV could therefore be between 450 and 600 metres
depending on the weather on a given day.
26
In addition to the airspace and cloud cover concerns, the UAV needs to fly at a height
whereby the cameras view of the traffic is reasonable. If the UAV operates too low,
buildings may obstruct the view of certain areas and the radius of the camera view will
be quite low. If the UAV operates at a high altitude the view will be broader but the
quality due to the large distance to the ground. By basic trigonometry, flying at a height
of 100m, 500m and 1000m, assuming a camera view angle of 90o, allows coverage of
200m, 1000m and 2000m respectively, as shown in figure 2.7.
Figure 2.7: Diagrams of view radius for various altitudes
Therefore, the UAV should be designed to fly at an approximate altitude of 500m to
remain below the controlled airspace level for most main cities and the average cloud
base height, while still maintaining enough altitude to provide appropriate coverage of
the area below.
2.5.2.6 Take-off/Landing
Due to the constraints of operation within a city, it is highly unlikely that there will be
enough room available for a runway take-off. This includes a grass take-off, such as
27
from a lawn or median strip. As the UAV will not need to be very large, it is feasible to
use either a hand launch or catapult launch. These require much less space and can be
launched from anywhere. For the same reasons a runway take-off is not feasible, a
runway landing is not feasible. There is a high possibility there will not be space
available for a long landing in the city centres, where these UAV’s will mostly operate.
A landing option where space is not an issue, such as a net catch is more feasible. This
also negates the need for landing gear, simplifying the design further.
2.5.2.7 Operating Conditions
The UAV should be designed to operate under the temperature, wind and rain
conditions expected in Australia. From analysis of weather statistics gathered by the
Australian Bureau of Meteorology in capital cities, weather stations 009034, 014016,
023000, 040214, 066062, 070014, 086071 and 094029, it can be seen that the
temperatures of major centres have not exceeded 50°C (323K) nor gone below
negative 20°C (253K) at ground level (Australian Government: Bureau of Meteorology
2010b-i). It is foreseeable, however, that during an extended loiter period, a UAV could
significantly exceed these temperatures due to solar heat radiation (up to a maximum
average of 3.30 MJ/m2/hr recorded in Brisbane in January) (Australian Government:
Bureau of Meteorology 2010c).
The Beaufort Wind Scale defines maintained wind speeds of 48-55 knots as a ‘Storm’
resulting in ‘considerable structural damage’ and beyond this are ‘Violent Storm’ and
‘Hurricane’ categorised by wind speeds of 56-63 and 64+ knots respectively (Australian
Government: Bureau of Meteorology 2010a). Unidirectional winds of these speeds in
conjunction with mild rain should theoretically not be problematic for an automatically
28
piloted UAV although operating in gusty, high speed winds (as winds of these speeds
would be in real conditions) would be hazardous and not recommended.
2.5.3 Technical level of the product
The UAV does not need to be technically superior to all other UAV’s in the market, the
task required of this UAV is relatively simple in current aviation terms. Therefore the
UAV should be designed to be simplistic but reliable. There should be some room for
generational evolution, such as the camera being changed to match evolving camera
markets, and the endurance, range and speed to be increased to match urban sprawl.
2.5.4 Economical Analysis
Current traffic monitoring by news programs involve a helicopter costing $744,000USD
(or around $900,000AUD) (from table 2.1) requiring at least two humans to operate,
one pilot, one cameraman and one reporter. The cost of the salaries for the three
personnel further increases the cost. For the government traffic monitoring
departments the current use of cameras costs $900,000USD (or around
$1,200,000AUD) per year to install and maintain (from section 2.2). Therefore if the
UAV can be made for less than $900,000AUD, it will be more economically viable than
the current technologies.
29
2.5.5 Propulsion Systems
Modern society has become very energy and emission conscious and so to have the
UAV more easily accepted it should be designed to be as clean as possible. A battery
powered electric UAV is a suitable design. It is clean, quiet, low maintenance and due
to the small size of the UAV, power is not an issue.
2.5.6 System Requirements
The flying wing UAV will require control surfaces to allow stability and control during
flight, with surfaces to control pitch, roll and yaw. Flying wing UAVs generally do not
have a rudder or flaps, and the elevators and ailerons are combined into a joint system
called an elevon (Waszak and Jenkins 2001).
The main avionic systems on board will include the components shown in table 2.5. The
UAV should be designed to fly a pre-determined flight path on autopilot covering the
usual congestion areas, however, there needs to be an override system so the UAV can
be manually remote controlled in case there is a particular area that requires more
focus. A global positioning system (GPS) is required to assist in navigation.
30
Table 2.5: Main on-board avionics systems
GPS Location, navigation and autopilot reference
Autopilot Allows the UAV to navigate independently
Remote Control Piloting Can override the autopilot if the pre-determined flight path needs
to be changed
Computer Controls all onboard avionics system
Transmitter Allows communication between aircraft and ground control
Actuators Operates the control surfaces
Camera Captures images of traffic situation to transmit to ground control
2.5.7 Reliability and Maintainability
As the UAV will be flying over highly populated areas, it is of utmost importance the
UAV does no fail mid-flight and crash to the ground. If a crash did occur, due to the high
population density it is most likely the UAV will cause severe damage to property or
human, possibly even involving the loss of human life. Therefore reliability must be
considered a vital factor.
As the UAV will be being operated by media and/or government officials it must have
low maintenance requirements. There are currently no CASA regulations regarding
maintenance schedules for UAV’s under 150kg. However, rigid guidelines for
maintenance and servicing must be developed to eliminate the chance of the UAV
failing in any fashion.
31
3 Conceptual Design
The conceptual design depends on the requirements stated in the technical task
(section 2.5) and requires many calculations and comparisons of the various
components. Weight, sensitivity and sizing calculations must be performed before the
design is chosen. Then a general body shape will be chosen, with the airfoil, planform
and control surface designs to follow. Propulsion and avionics systems should be
specified along with take-off and landing methods. The following section details the
development of the design for the traffic monitoring UAV, along with material
selections and a structural analysis.
3.1 Concept Designs
Before the main design can be chosen and the relevant parameters specified, concept
designs need to be devised as a basis for the main design. Three basic concept designs
are outlined below with sketches.
32
3.1.1 Concept Design 1 – Blended Wing Body
Lift to drag in an important parameter desirable to maximise in the design of an
aircraft. Methods of increasing the lift to drag ratio [L/D] is by minimising both the
number of protuberances out of the surface of the vehicle and the cross-sectional area
normal to the flow of air. One configuration of flying-wing aircraft, like that specified
for airborne traffic monitoring, is the blended-wing-body (BWB) which does not have a
separate fuselage section and so increases the area of lifting surfaces proportional to
the overall surface area of the aircraft. This two-view concept sketch displays a BWB
UAV with inset antennae for data transmission and reception, tricycle landing gear for
emergency landings, thereby protecting the belly of the vehicle and two vertical airfoils
in order to assist in vertical stability. The craft is a compromise between stability and
manoeuvrability, with wings swept backwards to increase lateral stability. In the event
of a belly landing, or sufficiently rough landing conditions to cause the UAV to become
uncontrollable, the vehicle’s negative dihedral (anhedral) angle would serve as
protection for the avionics contained within the belly at the expense of the wing
structures.
Figure 3.1: Concept design 1 – Blended Wing Body
33
3.1.2 Concept Design 2 – Semi-autonomous UAV
The second concept design that can be seen in figure 3.2 below incorporates a number
of components that are necessary for a semi-autonomous UAV. Firstly there is an
onboard navigational GPS system and avionics to ensure the UAV can be programmed
to navigate where required. This also requires antennae for data transmission and
reception on the ground. Tricycle landing gear is installed for emergency landing
scenarios and as a compromise with stability and manoeuvrability, the UAV has swept
backward wings. The design includes winglets on the wingtips to increase the lift
generated. Winglets are also an economical way of reducing induced drag power
requirements (Dube 2010). Winglets can optimise drag over a large operation range
rather than just a single point which is advantageous to the application of a traffic
monitoring UAV.
Figure 3.2: Concept design 2 – Semi autonomous UAV
34
3.1.3 Concept Design 3 – Simple square edge
The third concept design involves a simple square wing design with a zero sweep angle.
The propeller is located at the rear of the aircraft along the centerline as to protect the
propeller during landing and emergency landings. There are four control surfaces, two
one each wing trailing edge, to control yaw, pitch and roll. The slight downward
inflection at the wing tips is to increase stability. The design is relatively simple for ease
of design and manufacture. There is no landing gear; the UAV will be launched with a
catapult and landed using a net catch or airbag landing negating the need for landing
gear. This means the base of the UAV will need to be structurally enhanced to avoid
damage to the internal components during an emergency.
Figure 3.3: Concept design 3 – simple square edge
35
3.2 Mission Profiles
Two mission profiles are considered. The first is for the main task of loitering above a
city during peak traffic periods for several hours at a time (figure 3.4a). The second is
for use during highway or rural accidents, where the UAV can be deployed, cruise to
the accident area, then loiter to observe the traffic conditions below before cruising
back in to the ground base (figure 3.4b).
Figure 3.4a: Mission Profile for loiter above a city during peak traffic periods
Figure 3.4b: Mission profile for cruise out to a remote area, loiter and cruise back
36
3.3 Weight analysis
A normal weight analysis uses the known payload and crew weights, along with an
estimate of fuel used for the mission to determine both the empty weight and the
maximum takeoff weight statistically. However, as this design has neither a crew nor
payload and will be operated by electric power instead of combustible fuels, an
alternate method is used.
The project design employs a combination of known values for some components of
the aircraft weight such as the onboard avionics and statistical methods for the
structural weight of the aircraft. Since the avionics are the main components on the
aircraft of which the weight is known, estimation of the entire onboard electronic
system will be conducted prior to the overall aircraft weight being done.
3.3.1 Avionic Weight Estimation
A good approximation of the total weight taken by the onboard electronics was
obtainable using specifications of products provided by the manufacturers. Preliminary
calculations were based on the summation of the weight of all the avionics onboard to
obtain a good first estimate for the aircraft takeoff weight. The overall distribution of
the weight taken up by avionics onboard the UAV design is as shown in table 3.1:
37
Table 3.1 – Avionics systems weight distribution
Component Weight (kg)
Datalink (Communications) 0.3
Navigation (Autopilot) 0.016
Control Surface Actuator (Servo Motors) x4 0.068
Sensor (Camera) 0.465
Electrical Supply (Batteries) x4 1.616
Total Weight 2.465
Known weights on the aircraft was determined to be 2.465kg. Estimation of the
unknown structural weight of the aircraft will now be undertaken using statistical
method.
3.3.2 Structural Weight Estimation
According to the works of Beard et al. (2005), the breakdown of the overall mass of a
UAV can be divided into four major subsystems namely the airframe, the propulsions
systems, the guidance and control interface (GNC) and the payload. According to Figure
3.5 below, the majority of the weight distribution of a UAV (40%) is taken by the
propulsions system. It was also noted that 11% of the overall UAV weight was taken by
the payload while 21% was allocated to the GNC of the aircraft design to allow the UAV
to be fully autonomous. The remaining 28% of the total mass was allocated to the
airframe of the UAV design. Utilizing this statistical approximation, a preliminary weight
estimation for the UAV design was obtained.
38
Figure 3.5: Subsystem mass breakdown of a UAV (Beard et al, 2005)
The calculation of the weight of the avionics onboard the UAV design can be taken to
account for the weight of the payload and GNC subsystems as described in the works of
Beard et al. (2005). The weight of the avionics, found to be 2.465 kg therefore
constitutes 32% of the overall UAV design weight. It was then inferred from these
results that an approximate total weight of the UAV design would be 7.7kg. This
approximation will hence forth be used in further design works and analysis.
3.4 Sensitivity
A sensitivity analysis is done in order to identify the parameters to which the take-off
weight is most dependent on. As the UAV being designed is to be powered electrically
through batteries the sensitivity can be calculated with regards to the payload weight
only. The sensitivity to other parameters, specifically the specific fuel consumption,
39
L/D, propeller efficiency and cruise speed, are not applicable in this case as this is done
for range and endurance cases which are dependent on the engine type.
The following formula is used to calculate the sensitivity to payload weight
(Equation 3.1)
The variable B was determined in the statistical analysis section and was found to be
0.97609. The variables C and D are determined as follows
(Equation 3.2)
(Equation 3.3)
In the case of a battery powered the batteries are considered to be the fuel. As the
batteries do not deplete or reduce over time as fuel does the terms Mreserve, Munusable
and Mff are not applicable and hence C=1. There is also no crew aboard a UAV so
D=WPL= 2.465kg = 5.435lbs. WTO is found from the statistical analysis section to be
7.7kg which is 16.9785lbs.
From equation 3.1:
This result means that for every 1lb that is added to the UAV the take-off weight will
increase by 3.295lbs.
40
3.5 Sizing
The following sub-sections detail the sizing calculations for stall speed, takeoff and
landing distances, climb and cruise. There are no defined standards specific to UAV
climb, cruise, stall and takeoff and landing distances, so the FAR 23 small aircraft
standards were deemed adequate to size the UAV. As the weight of the battery does
not change through the flight as combustible fuel reserves do, the takeoff and landing
weights are equal and no weight corrections will be needed.
Sizing for takeoff and landing distances will not be taken into account, since early
design decisions for alternate takeoff and landing methods that do not require a
runway were made. Time to climb sizing is included in the FAR 23 rate of climb sizing
section. Unless otherwise stated, the units throughout this section are lbs/ft2 for W/S
and lbs/hp for W/P.
3.5.1 Sizing to Stall Speed
The UAV was sized for a maximum stall speed of 60 km/h at WTO. The FAR 23 standard
for aircraft less than 12500 lbs is 61 knots (113 km/h), which is too great for a UAV
being launched by catapult. The chosen value for design lift coefficient CL max in clean
configuration is 1.2, based on data from Roskam (2005). The stall speed (equation 3.4)
gives a limiting value for wing loading W/S of 4.254.
(Equation 3.4)
2/264.4 ftlbsS
W=
41
3.5.2 Drag Polar Estimation
To estimate the drag polar equations for all necessary configurations in takeoff, landing
and clean flight, values for aspect ratio (A), Oswald efficiency factor (e) and variation in
CD0 are required. Values for e and ΔCD0 (Roskam 2005) considering simple flaps and a
small aircraft size have been assumed. CL max values have been estimated from previous
UAV designs, including the Hy-Five tailless UAV (Bayliss et al 2008) and A is assumed to
be 10, since a larger aspect ratio will provide a larger wing span. The initial ratio of
wetted wing area to reference wing area has been estimated at 2, since there is very
little surface area not associated with the wing surfaces. Equivalent skin friction drag
Cfe has been estimated at 0.006, providing the drag coefficient at 0° angle of attack, CDo,
of 0.012 in clean configuration. Table 3.2 summarises the various lift and drag
coefficients, drag polars, and Oswald efficiency factors for the three planned
configurations: clean, takeoff flaps and gear up, and landing flaps and gear up. Only
‘gear up’ configurations are used because the aircraft will not have landing gear, and
therefore finding ‘gear down’ configuration parameters is redundant.
Table 3.2: Configuration drag polars and associated coefficients
Configuration CD0 ΔCD0 CD0,total e CD CL
max
(1) Clean 0.012 0 0.012 0.85 0.012 + CL2/13.35 1.2
(2) Take off
flaps/gear up
0.012 0.010 0.022 0.8 0.022 + CL2/12.57 1.4
(3) Landing
flaps/gear up
0.012 0.055 0.067 0.8 0.067 + CL2/11.78 1.4
42
The CL and CD values for clean configuration are used to calculate the design L/D ratio of
10. This fits in the range of design L/D ratios for homebuilt, single and twin engine
propeller aircraft, which are between 8 and 10 (Roskam 2005). Even though the UAV
has some different characteristics than these larger aircraft, this range is a good
benchmark to meet in early conceptual design work.
3.5.3 Sizing to Climb Requirements
The UAV has been sized according to the FAR23 climb requirements, as there is no
specific climb sizing regulations for UAVs. The UAV has been considered a one-engine
aircraft, so all of the ‘One Engine Inoperative’ (OEI) requirements can be neglected.
For each condition, the propeller efficiency has been assumed as 0.7, even though this
value is quite high for a UAV application. The lift and drag coefficients were taken from
table 3.2 and the density ratio (σ) is taken to be 1 since standard sea level conditions
are assumed. The sections that follow are descriptions of the climb sizing procedures.
3.5.4 FAR 23.65 (AEO) Sizing to Rate of Climb
The FAR23 rate of climb regulations state that the minimum climb rate of an aircraft at
sea level is 300 fpm at a steady climb angle of at least 1:12 (Roskam 2005). The aircraft
configuration for FAR23.65 is landing gear retracted, flaps in the take-off position and
engines up to maximum continuous power (Roskam 2005), so configuration 2 from
table 3.2 will be used. To size for this requirement, the rate of climb is used to
43
determine RCP, the rate of climb parameter, in equation 3.5, and this value is entered
into equation 3.6 to obtain a relationship between W/P and W/S.
(Equation 3.5)
(Equation 3.6)
3.5.5 FAR 23.77 (AEO) Sizing for balked landing climb requirements
The FAR 23.77 sizing for balked landing with all engines operative requires the aircraft
to have a minimum steady climb angle of 1:30 at sea level with engines on takeoff
power, landing flaps and landing gear deployed. The drag polar for this configuration
must be calculated once CL land has been estimated, leading to an estimate for L/D. The
climb gradient parameter can be found from equation 3.7, setting the climb gradient to
1/30 radians, then a second equation for the climb gradient parameter (equation 3.8)
can be used to find another relationship between W/P and W/S.
(Equation 3.7)
(Equation 3.8)
44
3.5.6 FAR 23.65 (AEO) Climb Gradient Sizing
The FAR23 climb regulations also specify the required climb gradient, which, for an
aircraft at sea level, is 1:12 (or 1/12 radians). The aircraft configuration is the same as
for FAR 23.65 rate of climb sizing. The sizing method is the same as for FAR 23.77 climb
gradient, using equations 3.7 and 3.8, but using a climb gradient of 1:12.
3.5.7 Climb Sizing Summary
Table 3.2 summarises the equations for two unknown variables, W/P and W/S, for each
part of FAR 23 climb sizing. These equations are plotted on the matching diagram with
W/S as the independent variable (x-axis) and W/P as the dependent variable (y-axis).
Table 3.2: Climb sizing summary
FAR 23 Climb sizing parameters
FAR 23 section Equation
Rate of Climb
FAR 23.65
Climb Gradient
FAR 23.77 Climb Gradient
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3.5.8 Sizing to Cruise Speed
Sizing the UAV to cruise speed will also be undertaken using a classical FAR23 approach
as suggested in Roskam (2005). The equation relating W/S and W/P for cruise speed is
given in equation 3.9.
(Equation 3.9)
Here, σ is the density ratio between the air density at cruise altitude (1500 ft) and sea
level, and Ip is several combined terms collectively known as the power index. This can
be found from statistics (Roskam 2005) using the calculated cruise speed of 115 ft/s.
Therefore the cruise speed sizing equation that will be placed on the matching diagram
is:
P
W
S
W106.0=
3.5.9 Matching Diagram
The matching diagram in figure 3.6 shows all of the required equations obtained from
preliminary sizing of the aircraft. From analysis of the matching diagram, the design
values of W/S and W/P can be obtained.
46
Figure 3.6: Matching diagram
The design point was chosen as the intersection between the cruise speed sizing line
and the stall speed sizing line, as the met area is the triangle in the lower left region of
the plot. The values at this point are:
47
2/263.4 ftlbsS
W =
hplbsP
W/3.40=
Using a takeoff weight of 7.7 kg, the wing area was calculated to be 3.98 ft2 (0.37 m2),
and using the simple geometric equation A = , where A = 10, the wing span was
found to be 1.92m. This wing span seems appropriate for a UAV with a relatively small
takeoff weight, and compares well to the hypothetical wing span of 2m from the
statistical data in section 2.4.2.
3.6 Configuration/Planform
The geometry of the wing can be planned using the wing area from the sizing section,
selection of some values and calculation of others based on known formulae. Several
important parameters for wing configuration are defined and discussed below (adapted
from Raymer, 2006).
Taper ratio: the ratio of root chord length to tip chord length. Previous design
estimates state that low sweep wings generally have a taper ratio of 0.4-0.5, and wings
with higher sweep have lower ratios of 0.2-0.3 (Raymer 2006). For this design, a taper
ratio of 0.4 seems adequate, as this is a mid-range value.
Aspect ratio: selected as 10 in the sizing section, which is a reasonably high value to
maximise wing span and lift coefficient.
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Thickness ratio: a thickness ratio (t/c) of 14% or over for the airfoil causes increased
separation and increased structural weight, but also causes stall from the trailing edge
and a gradual loss of lift and small change of pitching moment, which is desired. This
will be taken into consideration when selecting an airfoil (see section 3.7).
Wing sweep: the sweep angle is the angle between a line perpendicular to the aircraft
centreline and either the leading edge or the quarter chord line. Sweep helps to delay
stall, and a preliminary quarter chord sweep has been chosen as 20°.
Twist: Applying twist to the wings gives the tip airfoil a different angle of incidence
than the root airfoil. Since the wings will be swept, some wash-out is recommended to
prevent tip stall and lower the wing bending moment. It also decreases weight, but
makes manufacturing more difficult. For this simple design twist would be too
complicated during manufacture for the benefits gains therefore the wings will have
twist and instead, wing tips will be installed to prevent tip stall.
Incidence angle: Generally a consideration for passenger aircraft so will not be
considered in this design.
Dihedral angle and vertical location: no dihedral angle will be added since having
swept wings gives a natural dihedral effect, increasing roll stability.
Simple equations and geometry relate all parameters required for wing planform
design and the calculated values for the flying wing UAV along with their symbolic
representation and associated equation if necessary are provided in table 3.3.
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Table 3.3: Wing geometry equations and values
Wing parameter Equation Value
Wing area, S (from sizing section) 3.98 ft2
Aspect ratio, A (selected in section __) 10
Wing span, b A = b2/S 6.31 ft (1.92 m)
Mean aerodynamic chord, c A = b/c 0.631 ft (19.2 cm)
Root chord, c0
c =
0.85 ft (25.9 cm)
Taper ratio, λ (selected) 0.4
Tip chord, ct λ = ct/c0 0.34 ft (10.4 cm)
Quarter-chord sweep angle 20°
Leading-edge sweep angle 28°
Angle of twist 0
Using these parameters a simple configuration diagram of the wings is constructed, as
shown in figure 3.7.
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3.7 Airfoil
Airfoils define the cross-sectional shape of the wing and the amount of lift the wing can
generate. The airfoil must satisfy lift, drag and moment criteria in all flight
configurations (Raymer 2006). A previously designed airfoil will be selected since airfoil
design is inherently difficult and beyond the scope of this design project. Previous
information and computer simulations were researched as a basis for the airfoil
selection, since more elaborate computational or wind tunnel analysis is beyond the
scope of this project.
The airfoil must satisfy the following conditions:
- Must be able to provide maximum design lift coefficient CL max = 1.2, from
section 3.5.2.
- Must be able to operate in flow with Reynold number (Re) of approximately 5.0
× 105 – choose Reynolds number range of 1.0 × 105 to 2.0 × 106
- Must be able to produce L/D as high as possible, by minimising CD and
maximising CL. At the design lift coefficient, the drag is almost solely from skin
friction drag.
The airfoil properties necessary for a flying wing aircraft are different to an aircraft in
conventional configuration. Since the aircraft is tailless, using a cambered airfoil for
aircraft with a tail would cause instability. In a wing with such an airfoil, the
aerodynamic centre is located in front of the centre of gravity, so when a disturbance is
encountered and the angle of attack increases, the wing will continue to pitch up as it is
unstable (Hepperle 2006). Airfoils for flying wing or other tailless aircraft generally have
a reflexed, or reverse camber line, which gives the airfoil a inflected rear as shown in
figure 3.8. Generally, the reverse camber occurs in the last 25% of the airfoil length.
This causes the aerodynamic centre to be behind the centre of gravity; hence a
51
disturbance that increases angle of attack will cause the nose to pitch down, rendering
the airfoil and wing stable (Hepperle 2006).
Figure 3.8: Airfoil with reflexed camber line (Hepperle, 2006)
Following on from the need for stability in the airfoil is the value of the pitching
moment coefficient Cm. Prior to the airfoil stall angle, this coefficient should have a
negative value, which should change to positive after the stall angle, or else maintain a
negative slope that is very close to zero. In the case of a flying wing, the most desirable
scenario would be to have a moment coefficient very close to zero over the full range
of angles (Raymer, 2006).
The critical Mach number of the airfoil does not need to be considered in this analysis,
since the aircraft will not possess or require the capability to travel in the high subsonic,
transonic or supersonic speed ranges.
Research of previous airfoil designs yielded five prototypes suitable to a flying wing
UAV with a reflexed camber line as discussed above. The airfoils coordinates were
sourced (UIUC Coordinate Database 2010) and plotted and analysed in Javafoil, a free
software determining airfoil properties from uploaded coordinates (Hepperle 2006).
Table 3.3 lists each airfoils’ stall angle, maximum lift coefficient, moment coefficient
and drag coefficient at zero angle of attack. The flow is at a Reynolds number of 1.0 ×
106, which is a higher than expected value for this aircraft, but there is little difference
52
between a Reynolds numbers of 5 × 105 and 1 × 106. The analysis was conducted for
angles of attack between 0° to 20°.
Table 3.3: Comparison of airfoils
Airfoil Stall angle CL CMα CMo CDo
Eppler 334 11° 1.498 -0.059 -0.041 0.01675
Eppler 342 13° 1.468 -0.025 -0.0004 0.00889
Clark YH 14° 1.432 -0.063 -0.016 0.00833
MH 70 12° 1.438 -0.060 -0.048 0.01844
MH 78 19° 1.7 0.011 0.041 0.01695
Figure 3.9 shows a plot of moment and lift coefficients against angle of attack, which
determined the stall angle, maximum lift coefficient and nature of the moment
coefficient before, during and after stall for each airfoil.
Figure 3.9: Airfoil coefficients against angle of attack plot
53
The moment coefficient should take a negative value and have a slope close to zero.
This will provide stability, as positive moment coefficients indicate instability for some
angles of attack and if there are inconsistencies or sharp declines in the slope after the
stall angle, the stall comes from the leading edge or across the entire airfoil, rather
than the preferred trailing edge stall. From figure 3.9, the Eppler 342 airfoil has a
moment coefficient value and slope slightly negative but close to zero, therefore is the
most desirable.
Again from figure 3.9, the lift coefficient and stall angle of each airfoil can be identified
by the turning point of the graph for that airfoil. MH78 has a high stall angle and
corresponding CL, but also a very high CD0 while the Eppler 334 has the lowest stall
angle. The Eppler 342 is the most favourable; it has a moderate stall angle of 13° and a
lift coefficient of 1.47, which is above the CL of 1.2 chosen in the sizing section, allowing
for other parts of the aircraft that could lower the CL value to the design value.
The drag polars for each airfoil are shown in figure 3.10 below with the values of CD0
recorded in table 3.3. The Eppler 342 airfoil and Clark YH airfoil have the lowest CD0
values at around 0.008. These are lower than the design CD0 from the sizing section,
and would therefore be the best airfoils to choose in terms of minimising drag.
54
Figure 3.10: Drag polars for prototype airfoils
From comparison of the five prototype airfoils in Javafoil, the Eppler 342 airfoil was
chosen for the flying wing UAV. The Eppler 342 airfoil has a reflexed camber line, a
moment coefficient slope that is close to zero with values that are slightly negative, a
lift coefficient higher than the design lift coefficient, a moderate stall angle and a low
drag coefficient at zero angle of attack. The thickness ratio is 14.3% indicating the
airfoil will stall from the trailing edge and confirms the small change in moment
coefficient demonstrated. A normalised cross-sectional view of the Eppler 342 airfoil is
shown in figure 3.11.
55
Figure 3.11: Eppler 342 airfoil cross-section (UIUC Coordinates Database, 2010)
3.7.1 CFD Analysis
An in-depth CFD analysis would be conducted as part of the detail design; however, a
simple preliminary analysis was performed to gain initial insight into the flow around
the selected airfoil. The airfoil shape was drawn using CAD software and ANSYS CFX
was used to conduct a two-dimensional analysis at a Reynolds number of 1.0 × 106, and
at angles of attack of zero degrees and 13 degrees (the stall angle). Figures 3.12(a) and
(b) show the velocity profiles at the two chosen angles of attack.
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Figure 3.12 : Velocity profileas at (a) α = 0⁰ and (b) α = 13⁰ (stall)
Figure 3.12(a) shows a boundary layer around the airfoil that remains attached, as
expected at low speeds. The ‘bumps’ as the velocity increases would most likely be due
to the turbulent solver used and mesh quality, which would be refined much more in
further analyses. Figure 3.12(b) shows the separation of the boundary layer at stall,
regions of high velocity on top of the airfoil and lower velocity underneath.
Figure 3.13: Pressure profile at (a) α = 0⁰ and (b) α = 13⁰ (stall)
Figure 3.13(a) shows the highest pressure at the stagnation point, located at the
leading edge of the airfoil, and a low pressure region on top of the airfoil, matching a
57
region of high velocity in figure 3.12(a). At stall, in figure 3.13(b), the stagnation point
has moved to underneath the front edge of the airfoil, and the difference in pressure
generated at stall is evident from the plot. Again, anomalies are most likely the result of
the turbulent solver used or a mesh that could be refined much further, but was good
enough for this simple analysis.
These results are what would be expected from this airfoil with a stall angle of 13
degrees, and form a promising basis for further analysis. Nothing unexpected arose
from CFD analysis of the airfoil at varying angles of attack, which is a good initial test to
see if the airfoil will perform normally once used in practice. Of course, wind tunnel
tests and further CFD analysis would be required to confirm these initial results, but will
not be conducted here.
3.8 Control Surfaces
The UAV will need control surfaces to create stability and allow control during flight.
The BATMAN UAV requires control surfaces to affect changes in pitch, yaw and roll to
allow the aircraft to operate as required. Aircraft usually have four control surfaces to
allow manoeuvring: flaps, elevators, ailerons and a rudder. The flaps provide extra lift
during takeoff and landing as required, elevators control pitch, ailerons control roll and
the rudder controls yaw (Brandt 2004). Flying wing UAVs however often do not have all
four control surfaces as they are not required.
Small flying wing UAVs generally do not have flaps as the planform area is large in
comparison to the aircraft size and weight, therefore it is not required to operate flaps
to gain more lift during takeoff and landing as the wing area by itself provides enough
lift.
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A flying wing UAV by definition has no tail and therefore generally has no rudder. To
control yaw, some flying wing UAVs operate split ailerons, with two control surfaces in
parallel on each wing. One surface deflects downwards and the other simultaneously
deflects upwards creating an airbrake effect. If the split aileron on one wing only is
activated, the aircraft will yaw in that direction.
Flying wing UAV’s usually do not use separate ailerons and elevators. The ailerons and
elevators are combined to become ‘elevons’ that control both pitch and roll. Elevons
are located on the trailing edge of each wing. If both elevons are deflected in the same
direction the aircraft will change pitch. If one elevon is deflected downwards while the
other deflected upwards the aircraft will roll (Waszak and Jenkins 2001).
The BATMAN UAV will not require flaps to operate due to its small size and weight
simplifying the design greatly. Also, as the BATMAN UAV will be launched using a
catapult method, standard takeoff procedures do not apply. The pitch and roll will be
controlled by two sets of elevons located on the trailing edge of each wing. The set of
elevons located closest to the wingtips will be a split elevon, similar to the split ailerons
used to control yaw. The surfaces will be able to act together or independently
depending on the requirements.
A summary of the control surfaces activation requirements for certain manoeuvres is
shown in table 3.4. Column 1 shows the aircraft movement required, with columns two
through to seven showing the deflection state of each control surface.
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Table 3.4: Control surface deflection states for various aircraft manoeuvres
Aircraft
Movement
L1 Elevon 1 L2 Elevon
(upper
surface)
L2 Elevon
(lower
surface)
R1 Elevon R2 Elevon
(upper
surface)
R2 Elevon
(lower
surface)
Pitch
Down
Downward Neutral Downward Downward Neutral Downward
Pitch Up Upward Upward Neutral Upward Upward Neutral
Yaw Left Neutral Upward Downward Neutral Neutral Neutral
Yaw Right Neutral Neutral Neutral Neutral Upward Downward
Roll Left Upwards Upwards Neutral Downwards Neutral Downwards
Roll Right Downwards Neutral Downwards Upwards Upwards Neutral
3.9 Propulsion
A propulsion system should be chosen to satisfy any requirements of cruise speed,
ceiling, fuel efficiency, thrust, weight, cost and environmental considerations. For this
UAV the important considerations are cruise speed, weight, cost and environmental
concerns. The cruise speed needs to be a minimum of 68kts (from section 2.5.2.4) to
allow the UAV to complete the required distance over the cities within a reasonable
time. The weight of the propulsion system should be kept to a minimum as the overall
weight of the UAV needs to be kept to a minimum to allow easy handling by any human
being. The cost of the propulsion system should also be kept to a minimum, as to keep
the overall cost of the UAV lower. Finally, as discussed in section 2.5.5, the UAV should
be designed to be as environmentally friendly as possible, to satisfy the demands of the
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environmentally conscientious modern society. As the propulsion system often releases
the most emission to the environment, the propulsion system for this UAV should be
chosen accordingly.
Many different propulsion systems are currently used to power UAVs ranging from
piston engine powered propeller aircraft to solar powered UAVs and jet propulsion
UAVs. A summary of each propulsion type and it’s suitability for a traffic monitoring
application follows.
3.9.1 Propellers
Rotating propellers may be used to power an aircraft, providing the forward
momentum. The propeller can be located either at the leading edge of the wing in
tractor configuration, or the trailing edge of the wing in pusher configuration (Raymer
2006). Pusher propellers are located behind the engine, effectively pushing the aircraft
through the air. For a UAV with a pusher propeller configuration, there is no propeller
wash over the wings increasing the wing efficiency. Tractor configuration is with the
propeller located on the front of the aircraft, pulling the aircraft forward through the
air. Tractor configuration on a flying wing UAV increases aircraft stability. Propellers can
gain their power from either a fuel powered piston engine or an electric motor.
3.9.2 Piston Engines
Piston engines create a mixture of fuel and air, compress the mixture then ignite it to
provide rotary power to the required device, in this case a propeller. The engine can
either be in a linear or rotary configuration (Raymer 2006). Due to the explosive use of
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fuel, the UAV would require both fuel tanks and refuelling sessions. Having to carry fuel
on board would increase weight and size. For aircraft the required fuel is usually Avgas.
Burning Avgas creates by-products of carbon dioxide and carbon monoxide that pollute
the atmosphere (Confederation of Australian Motor Sport 2003). In an environmentally
conscious culture, like modern day society, these by-products are not a desirable
option and may reduce buyers, and present an undesirable image to the general public.
3.9.3 Electric Motors
A cleaner, but often less powerful version of a piston engine is an electric motor. Like
the piston engine, an electric motor provides rotary power to the required device,
however it does not require combustible fuel to operate. An electric motor is powered
by electricity, provided either through mains power or a battery pack. For the purposes
of an airborne aircraft, the motor would have to be run from a battery as mains power
would not be available during flight. Motor speed, and therefore propeller speed, is
controlled by modifying the voltage provided from the battery to the motor (Batill,
Stelmack and Yu 1999).
3.9.4 Solar Powered UAVs
One of the most important parameters of this UAV’s design is endurance. Endurance
largely depends on whether the UAV is solar powered or fuel powered. Endurance of a
fuel powered UAV is limited by the percentage of fuel burned as a fraction of total
weight and thus has a lower endurance limit than a solar powered UAV. Solar powered
UAVs were first brought on the market during the early 1980s. One such solar UAV is
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the High Altitude Solar (HALSOL) UAV. The HALSOL project was developed to
incorporate silicon solar cells that produce up to 12,500W (The Future of Things 2007).
Whilst having unlimited endurance at altitudes up to 100,000 feet (The Future of Things
2007) solar powered UAV require a large wing span, some up to 247 feet, to
accommodate the required amount of solar panels. The Helios UAV has 62,120 bi-facial
solar cells at a rough cost of $20,000AUD per kilowatt (Urban Ecology Australia 2006).
Despite the benefits that arise from solar UAVs the technology is fairly young and to
date there are few commercial or military applications that utilise solar energy as a
power supply (The Future of Things 2007).
Solar powered UAVs are becoming a viable alternative to conventional fuel systems
with advantages such as low operating costs and flexibility in mission tasks (SPAWAR
2004). From a maintenance perspective, solar UAVs are easily recovered and can be
repaired or modified without hassle. Research has stated that such UAVs can be used
for a number of applications such as civil surveillance, research and telecommunication
relays (SPAWAR 2004). A solar powered UAV would provide a cost effective, reliable
and energy efficient method to traffic monitoring in Australian cities.
3.9.5 Jet propulsion
Currently there are very few jet powered UAV’s available, especially those classified as
small. A jet engine would provide ample thrust for the mission, however the fuel
consumption would be high, as would the environmental emissions and the UAV would
be quite loud. As the UAV is designed to fly over populated areas it is best if noise is
kept to a minimum, therefore a jet propulsion system is not considered.
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3.9.6 Propulsion System Selection
An electric motor powered pusher propeller system was chosen for the BATMAN UAV,
as this provided a clean, quiet and efficient propulsion method. An electric motor
negates the requirement of combustible fuel and fuel tanks, whilst also being quieter
than both piston engines and jet engines. Currently, solar cells are not yet efficient
enough resulting in solar powered UAVs being larger and therefore heavier.
From section 3.5.9, W/P=40.3 lbs/hp. Assuming a weight of 7.7kg (from section 3.3.2)
the power (P) required from the engine is therefore 0.42hp. To ensure ample power is
available for the aircraft, an electric motor with 0.5hp was chosen. The motor chosen
was a Baldor premium efficiency AC motor, product number EM3538, providing 0.5hp
at 1740rpm (Baldor 2010) as this was the lightest motor found with the required
power. The motor weighs 10lbs and requires 1.6Amps at 208V. The batteries used to
power this motor will be the same batteries as those chosen to power the avionics
systems, discussed later in section 3.12.5. The motor requires 1.6Amps, so for a mission
duration of 3hrs the requirement is 4.8Amp-Hrs. Therefore, to minimise weight while
maintaining the required power, one zinc-air battery providing 5Amp-Hrs will be used.
This takes the total weight of the system to 10.9lbs.
Figure 3.14: The Baldor premium efficiency EM3538 electric AC motor.
64
The propeller will be located on the centreline of the UAV in the pusher configuration.
As the UAV will be launched and landed using alternate methods, locating the propeller
on the rear of the aircraft will provide greater protection from damage caused during
landing and takeoff. The pusher configuration will also reduce propeller wash over the
wings and body, increasing the efficiency. The propeller will be manufactured ‘in house’
to lower costs and will consist of two metal blades for simplicity and cost, while
retaining a suitable strength. The material used for the blades will be a high strength,
heat treated aluminium alloy, as this is a light material with good strength properties.
To determine the length of the propeller blades, the following equation from Raymer
(2006) was used.
4 PowerKD P= (Equation 3.10)
Where the power is in horsepower and the propeller diameter in feet with Kp a
constant depending on the number of blades; for a 2 blade propeller Kp is 1.7.
ftD 43.15.07.1 4 =×=
To check the blades tips remain below the maximum suggested helical velocity, the
following relationship should be satisfied (Raymer 2006).
fpsVnD 950)( 22 <+π (Equation 3.11)
fpsftftrpm 9506.173sec)/77.114()43.1)1740(( 22 <=+×π
Therefore a propeller diameter of 1.43 feet is suitable.
The final propeller design will be a two blade propeller, manufactured from a high
strength, heat treated aluminium alloy, each with a length of 0.715ft, giving an overall
propeller diameter of 1.43ft.
65
3.10 Launch
As the UAV will be designed to operate within metropolitan regions where the
population density is quite high and real estate is at a premium, alternative launch
methods will be considered to avoid the necessity of a large area for a runway takeoff.
Small UAVs can be launched either through conventional runway takeoffs or can be
hand launched. A conventional runway launch would require large amount of flat space
for the UAV to reach an appropriate takeoff velocity. Although the UAV would be able
to takeoff from a grass or gravel runway, a concrete or bitumen runway would be
better as the surface is smoother and more reliable. This would require a section of
ground to be surfaced appropriately and within the bounds of a city real estate can be
expensive and hard to acquire so this could provide a problem.
Hand launching is a viable option for a UAV this size. It is light enough for an average
human to hold in preparation for launch. However, the propeller could pose a severe
safety hazard if the person was to hold the UAV directly. As safety is a priority concern,
this is not desirable. Small UAVs can also be launched using a catapult. The UAV is
attached to a hydraulic catapult that uses pressure to accelerate the UAV along a rail to
the required velocity at which point the UAV can fly under its own power. Table 3.5
compares three catapult launchers available commercially for UAVs.
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Table 3.5: Commercial UAV Launcher specifications
MicroPilot MP CAT1 BAE Pusher Prop
Launcher2
RUAG Aerospace3
Maximum UAV Weight 20kg 36kg 320kg
Length 4m 3.5m 14m
Catapult Weight 25kg 123kg 3500kg
Power Batteries Batteries Electrical Power
Generator
Maximum Launch
Velocity
23m/s 20m/s 66kts
Price $15000 USD Not Available Not Available
1: http://www.micropilot.com/products-catapult.htm
2: http://www.acrtucson.com/UAV/launcher/Launcher.pdf
3: http://www.uvs-international.org/pdfs/brochures/ruag_lnchr_archer.pdf
Figure 3.15: ScanEagle UAV catapult launcher (wikipedia.org)
67
3.11 Landing
Conventional aircraft feature one of a variety of different landing gear configurations
tailored for the specific aircraft depending on its size, configuration, task and applicable
standard requirements. Retractable landing gear requires a large area inside the
fuselage for storage of the landing gear when retracted, increase the weight of the
aircraft significantly and are quite complex (van Blyenburgh 1999). However, a fixed
landing gear system, such as that of a tricycle configuration, can increase the overall
drag co-efficient by 5-10% (Boschetti, Cárdenas & Amerio 2005). However as the
BATMAN UAV is classified as a small UAV alternate landing methods are feasible
therefore the BATMAN UAV will not require a dedicated landing gear system.
For the same reasons as to why the launch method should be chosen to minimise the
space required for takeoff, the landing method should also be chosen to minimise the
space required for landing. Parachute landings, airbag recovery methods and net catch
landings are three possible options to eliminate the requirement of a runway.
The use of parachute and airbag recovery methods for UAVs in civilian areas had
become more common in recent years. In 2009 the Korea Advanced Institute of Science
and Technology (KAIST) Department of Aerospace Engineering designed a ‘vision-based
landing system for small-size fixed-wing’ UAVs (Huh & Shim). Seventy percent of
manually controlled fixed wing UAV landing accidents occur due to human error (Huh &
Shim 2009). To avoid a GPS reliant autopilot landing systems or relying on manually
controlled landings, Huh and Shim developed a low cost, high reliability recovery
method whereby the UAV visually tracks, and comes to rest on hemispherical airbag
through a forward mounted camera. The airbag is a pre-designated colour to allow the
UAV to visually track the landing area. This method, however, also requires nets to
68
protect against cross winds affecting the landing in addition to a second camera and a
visual recognition system.
A portable net recovery system is a landing system in which the UAV comes to rest in a
net. The net recovery system is reliable, simple and cost-effective for use with the
solely GPS guided or selectively radio controlled UAVs such as BATMAN. No portable
net recovery systems are currently commercially available; all current net systems are
used in military applications. Therefore, the net system would must also be developed
by the UAV company. The net would need to be specially made from high tensile but
flexible synthetic material. Due to the nature of this report, a quote for cost will not be
included as a manufacturer would need to be contacted directly. However, the UAV will
still be designed to land using a portable net recovery system.
3.11.1 Emergency Landing Systems
Due to the operational area being in the urban environment, the UAV must be able to
conduct emergency landings with little effect on the aircraft and environment. There
are two options for the emergency landing procedure, being a belly-landing or small
emergency fixed landing gear. Damage resulting from a belly landing would be
concentrated on the UAV body resulting in a new casing to be built, however
preferable locations for belly-landings such as low-cut grass or water can reduce the
overall damage (Jones 2003). The inclusion of a small fixed landing gear system, could
cause localised damage to the body of the UAV, potentially causing damage to
expensive internal components (Lee 2004). Therefore, the BATMAN UAV will be
designed to perform emergency belly-landings with no landing gear, and as such, the
bottom region of the UAV body will be reinforced to lessen the damage caused. Also a
slight negative wing dihedral could be included to minimise internal damage in the case
69
of an emergency landing, however as this UAV will feature a reinforced underbelly no
anhedral angle will be included in the final design.
3.12 Avionics
To ensure successful completion of the UAV mission, several avionics systems have to
be integrated into the UAV design. These systems include: a data link for
communication and control purposes, autopilot system and satellite navigation for
remote control capabilities, control actuation subsystems to enable flight parameters
of the UAV to be changed during the mission and an electrical supply system capable of
powering the avionics on board for the duration of the mission. Basic requirements for
each of these systems were developed to enable the design requirements to be
fulfilled. The following sections present a detailed discussion of each of these systems.
3.12.1 Data Link
A data link to allow the transmission of digital information between a ground control
station and the UAV is required, to allow both remote control piloting of the UAV and
transmission of the digital video images back to the ground station. Data link assembles
are divided into three main configurations: simplex communications that allow
communication in one direction only, half-duplex communications that allow
communication in both directions but not simultaneously and duplex communications
that allow communication in both directions simultaneously.
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Requirements of the aerial data link required for this UAV are:
1- The link must allow the UAV to accept flight control inputs from ground control
while it is on a mission
2- Allow the UAV to transmit real time video feed (telemetry) to the ground
station
3- The communication between the ground station and the UAV should have a
range of at least 10km.
Specifications for different data link systems currently available that fulfil the mission
requirements are shown in table 3.5 below. The final product should be chosen to
minimise weight but retain maximum system capabilities.
Table 3.5 Data Link Systems Product Starlink Multiband
Digital Data Link
System4
Commtact CTX
(Transmitters) and
CRX (Receivers)5
Commtact Mini Link
Data Systems5
Range 100 km 150 km 15 km
Dimensions 220mm x 220mm x
150mm
2 x (80mm x 84mm x
48mm)
124 mm x 94 mm x
17 mm
Weight 2.1 kg 0.79 kg 0.3 kg
Input Voltage 18 – 36 V (DC) 11 – 32 V (DC) 9 to 60 V (DC)
4 - http://www.tadspec.com/index.php?id=101
5 - http://www.aeronautics-sys.com/?CategoryID=257&ArticleID=183
The Mini Link System from Commtact is the smallest and lightest data link system that
can fulfil the design specifications and therefore was chosen to be the data link
configuration utilised for this UAV. This system is compact, lightweight, highly reliable
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and capable of delivering high quality video and telemetry signals. The system uses
advanced digital video compression and transmission-reception technology between
remote platforms and control stations. The cost of this system is estimated to be
$10000AUD. As the cost of these systems were not easily accessible, cost estimations
were made based on works by Buretta et al. (2003) which analysed a highly similar UAV
design.
3.12.2 Autopilot and Navigation Systems
To enable the UAV to fly, an onboard navigation system is required, that will enable the
UAV to establish its location in terms of map projections or coordinate systems. Such a
feature onboard the UAV is crucial since data collected would be utilised in the field of
geographic information systems (GIS), whereby data from different sources require a
common referencing system prior to being combined and used (MicroPilot, 2010). For
these purposes, a Global Positioning System (GPS) device onboard the UAV would be a
suitable means of providing the UAV with a comprehensive navigation system. A GPS
system tracks the latitude and the longitude coordinates of the UAV on site thus
effectively keeping a continuous track of the geo-referenced position of the UAV.
UAVs require an autopilot subsystem to guide the UAV without needing any manual
input. This system is essentially computer software integrated with a navigation
system. The software will be integrated with the navigation and data link set-up
onboard the UAV to allow it to read the aircraft’s current position and guide the UAV
on its mission task. Recent technological advances in autopilot systems have enabled
thrust control capabilities allowing the aircraft to be flown with a lower fuel-
consumption than a human pilot. The level of control in autopilots are divided into
three tiers: single-axis autopilot controls the aircraft in the roll axis only, double-axis
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autopilot controls the aircraft in the roll and pitch axes and three-axis autopilot allows
control in the roll, pitch and yaw axes.
Autopilot navigation systems from various specialist manufacturers were compared
with the specifications shown in Table 3.6 below. Again, due to the small size of the
UAV, the final decision was based mainly on weight, size and performance.
Table 3.6: Autopilot Navigation Systems Products UAV Navigation
AP04R6
Micro Pilot MP20287 Procerus Kestrel
Autopilot System
v2.238
Dimensions 46.7 mm x 68.0 mm x
74.0 mm
100 mm x 40 mm x
15 mm
54 mm x 35 mm x 12
mm
Weight 0.2 kg 0.028 kg 0.00167 kg
Altitude Limit 20,000 ft 40,000 ft n/a
Input Voltage 7 – 36 V 4.2 – 26 V 6.0 – 24 V
6 - http://www.uavnavigation.com/uavprod/uavprod_03.htm
7- http://www.micropilot.com/products-mp2028g-specs.htm
8 - http://www.procerusuav.com/index.php
The Autopilot system chosen for the UAV design was the Kestrel Autopilot System
v2.23 from Procerus Technologies as it is the smallest and lightest full-featured
autopilot that was could fulfil the design specifications. This particular system has
already been used for surveillance and reconnaissance applications making it ideal to
be implemented into our UAV design (UAV Navigation, 2009). Coupled with a ground
control Virtual Cockpit, this system is easily operated with powerful mission monitoring
capabilities that allow in-flight adjustments. This system works on a switching power
regulation that achieves high efficiency while consuming less power. The range is
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100km and the maximum speed 130m/s, therefore satisfying the requirements of the
estimated 10km radius circular flight path. The cost of the unit is $5000 (Precerus,
2010).
The Kestral Autopilot System has a stability control set-up that controls the airspeed,
altitude and direction of the UAV. Onboard sensors such as an accelerometer, angular
rate sensors and a magnetometer measure these flight parameters. The mission
control interfaces with the navigation control (which utilizes the GPS and data link set-
up) to ensure course corrections are available when necessary, missions are executed
as required and that the UAV responds to ground control inputs. The 4 serial ports
allow for support payload inter-communication and control which proves to be
essential for the UAV mission profile. This system allows the size and weight of the
overall navigation and autopilot set-up to be minimized while maximizing the flexibility,
functionality and expandability of the core avionic system of the UAV.
3.12.3 Control Actuation Systems
Control actuation systems are required on the UAV to enable the autopilot system to
implement mechanical control over the necessary control surfaces in order to vary
flight parameters. The simplest form of control actuation would be the use of servo
motors as they have a wide range of performance specifications required by the
mission profile. Verifications were made to ensure that the servos selected were able
to produce the necessary hinge moments necessary on the UAV for control actuation.
An approximation of 2.0kg/cm hinge moments required for flight were obtained from
research works conducted by Buretta et al (2003) which analysed an aircraft of similar
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weight, dimensions and flight configurations. This value was found to be within the
design torques of servo motors and is therefore used for the actuation of the elevons in
the aircraft design. Servo motors considered for the application of control surfaces for
the aircraft design were as follows:
Table3.7: Servo Motor Specification
Product MOT-301 Mini Servo
(Ocean Controls)9
S9001 Servo Aircraft
Coreless BB (Futaba)10
S3101 Micro Servo
(Active Robots)11
Dimensions 35mm x 16.9mm x
32mm
40mm x 20mm x
36mm
28mm x 13mm x 30mm
Weight 0.026kg 0.048kg 0.017kg
Torque 3.4kg/cm 5.2kg/cm 2.5kg/cm
Voltage 4.8V 6.0V 6.0V
9 - http://www.oceancontrols.com.au/motors/servo/rc_servo_motors.htm
10 -http://www.gpdealera.com/cgi-bin/wgainf100p.pgm?I=FUTM0075
11 - http://www.active-robots.com/products/motorsandwheels/futaba-servomotors.shtml
The servo motor selected to be implemented in the UAV design is the S3101 Micro
Servo from Active Robots. This model was capable of providing the required torque for
the smallest dimensions and the lightest configuration among the models considered.
The aircraft design will require 4 servo motors to enable the elevons of the design to be
fully functional. Four units of this servo motor model will therefore be considered in
the electrical systems calculations. Priced at $55AUD each, the four units required for
the UAV design would sum up to a total of $220AUD for the control actuation systems.
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3.12.4 Sensors
The UAV provides a mobile platform for a multitude of device used for the purpose of
traffic monitoring. The basic sensors required to serve this purpose would be a
coloured camera with a sufficient field of vision (FOV) with optical zoom capabilities in
flight, mounted within an onboard gimbal. The camera must be equipped with image
stabilization capabilities that are capable of producing clear images at an altitude in the
order of 2,000ft. The camera should be equipped with aiming capabilities through the
coupling with the onboard GPS directed autopilot referencing system. The primary
selection specifications of cameras considered is shown in table 3.8.
Table 3.8: UAV Sensor Specifications
Product BTC-88 Micro Pan/Tilt
Unit (BTC-88 – Micro
Pan / Tilt Unit 2008)12
MicroPilot Dayview
(MP-DAYVIEWPTZ
/MP-
NIGHTVIEWPTZ
Stabilized Payload
Cameras 2006)
Cloud Cap TASE (TASE
Small Gyro Stabilized
Camera Gimbal n.d.)13
Weight 0.465 kg less than 0.9 kg 0.9 kg
Dimensions 136 x 89 x 124 mm Ø110 x 210 mm 127 x 112 x 178 mm
Voltage 6 to 12 V 12 V (nominal) 9 to 20 V
Power
Consumption
2.1 W 3 W 10 W typical, up to 18
W
Resolution 752 x 582 pixels 320 x 240 pixels 752 x 582 to 1080i HD
Optical Zoom 10x 25x Dependent on camera
Retractable Yes (Beetle-Wing
doors)
No Additional Assembly
12 - BTC-88 Gimbal 2009, Procerus Technologies
13 - FCB-H11 Brochure 2008, Sony Corporation
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From the three camera and gimbal pairs considered, the system selected for the
project was the BTC-88 Micro Pan/Tilt Unit from Unmanned Airsystems and
Components. This unit is manufactured by a Procerus Technologies partner company,
Brandebury Tool and is hence designed with inherent compatibility with the Kestrel
Autopilot System manufactured by Procerus Technologies (BTC-88 Gimbal 2009). Despite
having a lower video resolution than those offered with the Cloud Cap TASE, the BTC-88 draws
less than a quarter of the typical power of the TASE and, unlike the heavier MicroPilot Dayview,
the BTC-88 is also retractable into its own housing thereby facilitating emergency belly
landings. Whilst there is very little available pricing information for stabilised gimbal camera
systems, Sony recommends list prices of $500USD and $1,260USD for the 752 x 582 pixel FCB-
IX11AP (FCBIX11AP: 10x Color EXview PAL Block Camera n.d.) and the 1080i FCB-H11 (FCBH11:
High Definition Color Block Camera n.d.) block cameras respectively, further justifying the use
of the BTC-88 with a FCB-IX11AP camera. From these costs, it seems reasonable to assume that
the cost of the BTC-88 gimbal system would be in the order of $1,000USD.
3.12.5 Electrical Systems
Conventional methods of providing electrical power to a UAV include carrying sufficient
batteries onboard or a combination of battery supplies with a generator integrated
with the UAV engine. To maintain simplicity of the design, the first method was chosen
to be implemented. Battery systems were therefore chosen to sustain all onboard
electronics for the given mission time frame while having the weight minimized for
flight efficiency purposes.
Evaluation of the total amount of electrical power required by the UAV design was vital
to determine a suitable battery system for the aircraft. Manufacturer’s stated
maximum current drawn for each electrical component were assumed during the
calculations of the required power onboard the UAV design. With the assumption of a
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mission endurance of 3 hours, the required battery capacity was determined to be 18
Amp-hours as seen in Table 3.9 below.
Table 3.9: UAV Electrical Power Requirements
Electronic Component Maximum Current (Ampere)
Autopilot 0.5
Transmitter 3.2
Receiver 0.7
Servo Motors (x4) 0.4
Sensor (i.e. Camera) 0.2
Miscellaneous 1.0
Total 6.0
Total Amps for Entire Mission 18 Amp Hrs
The battery packs chosen for the UAV must be capable of meeting the power
consumption of the electrical components onboard the UAV. Two types of batteries
were evaluated: lithium-ion batteries and zinc-air batteries. Lithium-ion batteries are
rechargeable batteries that utilise the movement of lithium ions from the negative
electrode to the positive electrode during the discharge phase and in reverse when the
battery is being charged. These types of batteries have been found to provide a good
energy-to-weight ratio with a slow loss in charge (BusinnesWire 2010). Lithium-ion
batteries have been used extensively in the fields of aerospace and defence due to
their high energy density capabilities their ability to provide electrical with minimal
noise (CleanTech 2010).
Zinc-air batteries alternatively are batteries which are powered by oxidizing zinc to
produce an electro-chemical reaction thus generating electricity. The electrochemistry
of these types of batteries are very similar to alkaline manganese with the Mn02 being
replaced with oxygen from the atmosphere. This makes zinc-air batteries considerably
78
safer and more environmentally friendly than lithium batteries (Defence Update 2005).
Zinc-air batteries were also found to be very high-powered and lightweight capable of
extending the flight time of small UAVs (The Free Library 2005).
Both types were considered to be implemented into the project design. Specifications
of the models considered are as shown in table 3.10.
Table 3.10: Specifications of Batteries Considered
Battery Type Lithium – Ion14 Zinc – Air15
Model MR-624 (Electric Fuel) FPEVO25-50003S (Flight
Power)
Weight 0.25kg 0.404kg
Dimensions 62.5mm x 38 mm x 74mm 27mm x 47mm x 74mm
Capacity 2.2Amp-hours 5.0Amp-hours
14 - http://www.electric-fuel.com/downloads/MR-624.pdf
15 - http://www.modelflight.com.au/flightpower_evo25_v-power.htm
Comparing both models, with the 18 Amp-hours requirement obtained in the previous
section, 2.25kg worth of batteries would be used if the lithium-ion option was selected
while 1.616kg worth of batteries would be used if the zinc-air option was taken. Due to
stringent weight constraints of the UAV design, 4 units of the zinc-air batteries will be
used to power all onboard electronics of the UAV design. In addition to the two similar
batteries required to power the electric motor (section 3.9.6) the UAV requires a total
of six zinc-air batteries. Priced at $272AUD each, the total cost of the electrical systems
would total $1632AUD for all four units required to enable full functional capabilities of
the design (Flight Power 2009).
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3.13 Materials
The UAV will be launched using a catapult and landed using a net device. If for any
reason the UAV must perform an emergency landing a belly landing will be done. To
ensure that minimal structural damage occurs in such a situation, the material used
must be strong and able to withstand a vast amount of impact. Metals, plastics and
composites are the materials to be considered. Metal alloys are normally used over
pure metals that are generally too soft to use on their own. Metal alloys present more
advantages in terms of combining different metal properties, to achieve the best
combination (Key to Metals 2010). Composites are materials that are composed of two
or more elements that work together to produce material properties that are
independent of the element properties (SP Systems n.d). Reinforcements are normally
used with composites to add strength and stiffness to the matrix material. Generally
composites are used for advanced and complicated technologies that cannot be met
with conventional alloys (Callister 2007). Plastics are generally not used for on their
own for high strength applications but can be used as a matrix material in a composite
scenario. ABS (Acrylonitrile, Butadiene Styrene) is the most common type of plastic
that can be used for a composite procedure.
Composites have the benefit of being able to produce a material to meet the desired
performance criteria, whilst with metals the component must be varied to obtain the
characteristics that are required (Dorgham 1986). Composites also have the added
advantage of formability; almost any shape can be conceived (Dorgham 1986). They
also “enjoy the best of both worlds, with many of the most useful features of both
metals and plastics” (Maxwell 1994).
In modern UAV design composites are generally used for their superior strength.
Composites can also be quite cost effective as seen by National University of Singapore
students Jayabalan N, Horng L & Leng G who utilised cardboard, balsa wood, paper
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mâché and laminate resins. Composites have resulted in better fuel consumption and
reduced drag coefficients when used in conjunction with good aerodynamic body
designs (Dorgham 1986). They also minimise the permanent damage occurs during
minor accidents (Foale 2006). Composites enable more complicated structures to be
engineered that have reduced weight and increased strength at a lower cost.
In finalising the material to be used, mechanical properties must be assessed to
determine what is most important and which material will satisfy the requirements. In
the case of a traffic monitoring UAV the density, tensile strength and cost will be
analysed as these are the important factors. Aluminium alloy will be the metal
considered in addition to ABS plastic. For a civilian application organic fibre
reinforcement is commonly used, specifically glass due to its low cost but adequate
strength properties for a less rigorous operating condition in comparison to a military
application (Borchardtm 2004). Based on this the composite analysed will be a glass (E)
reinforced composite.
Table 3.11 Comparison of Materials
Material Cost
(USD/kg)
Tensile
Strength (MPa)
Density
(kg/m3)
Source
Metal 2.0909 7-7001 2,700 AAD, 2007 and Metal
Prices, 2010
Composite 1.47-2.95 2000 2,550 Azom, 2010
Plastic 2.21-2.95 45 1,050 Kopeliovich, 2008
Composites present the best strength and weight characteristics for the cost and hence
shall be used for the manufacture of the traffic monitoring UAV. There is a wide
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selection of composites that can be used as well as resins and reinforcements.
Aluminium reinforcement works well with Kevlar structures as the combination
provides a rigid structure that is capable of withstanding a reasonable load in addition
to adverse environmental conditions (Buretta et al 2003). Aluminium is readily
available at a low cost while Kevlar is slightly more expensive but has a very high shear
strength, which is very important in the aeronautical industry.
3.14 Manufacture
The UAV can be manufactured at a very reasonable cost if performed in house. Labour
costs are much lower than if outsourced and hence the primary cost will be from
material selection and any required tooling. Manufacturing can be simplified by
ensuring components are designed to break apart on impact as this will reduce repair
and modification time. The UAV can be manufactured in the methods outlined below.
3.14.1 Reusable Moulds
The flying wing UAV structure will consist of components for the wing and housing,
made from aluminium and composite materials. There are many manufacturing
procedures that could be used depending on cost, time, specific materials and available
equipment. A guide to possible manufacturing procedures for the frame components
and skin is given in table 3.12.
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Table 3.12: Manufacturing processes (adapted from Kalpakjian & Schmid, 2006).
Structural
component
Material Manufacturing process
Spars Aluminium Wrought into the I-beam shape using rolling, extrusion
and drawing and hardened through heat treatment.
Frames
(housing)
Aluminium Wrought into designed curved shape using rolling,
extrusion and drawing and hardened through heat
treatment.
Ribs Aluminium Wrought and heat treated, then cut into specified airfoil
shape using CNC machining.
Skin Fibre-reinforced
composite (e.g.
fibreglass)
See remained of section
Flying wing UAVs can be manufactured using re-usable moulds into which the
framework is set and resin is used to hold the UAV together (Chan et al 2009). The
manufacturing process begins by creating a foam mould of the part to be produced. For
commercial purposes it is more economically viable to have reusable moulds whereby
the UAV may be produced again if required. The foam can be cut in many different
ways of which the three most commonly used and cost effective are: CNC, rig hot-wire
cutting and manual hot-wire cutting.
CNC is a machining process whereby a machine is used to cut out the desired foam
shape in 3D. A 3D model is developed in a CAD package and is used to define the
geometry of the shape (ShopBot 2010). The next step in the process is to develop a 3D
path for the machining tool to follow in order to cut out the desire mould. Once this
has been done the cutting process can begin whereby the CNC has simultaneous
motion in the X,Y and Z axes controlled by a computer for accuracy (ShopBot 2010).
83
While this technique offers the capability to produce complex shapes it is very time
consuming and has a rough cost of US$1/minute (Brown 2007).
Rig hot-wire cutting is a process whereby a wire is heated electrically across a bow and
then used to cut through the foam (Chan et al 2009). The apparatus utilises a pulley
system to ensure a constant cutting speed. However, while this process produces a
high quality final product with an excellent surface finish it is very time consuming and
often this outweighs the benefits of using it as a practical manufacturing technique
(Chan et al 2009). Also in order to use this method many variables, such as temperature
and environmental effects, must be carefully controlled to avoid an uneven finish. This
makes it an undesirable technique to use for large scale and high productivity
applications.
In order to achieve a surface finish to an acceptable standard a manual hot-wire cutting
process can be done. This is essentially the same process as the rig hot-wire cutting
except rather than using a pulley to pull the bow across the foam it is done manually
(Chan et al 2009). This method is quicker and produces a better quality product.
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3.14.2 Hand lay-up and pre-impregnated cloth
Composite material can be laid on in two, cost effective, methods: hand lay-up and pre-
impregnated cloth (Chan et al 2009). Hand lay-up manufacturing is a low cost
technique that many composite components are fabricated by (Cairns 2000). In this
technique, fibre reinforcement is inserted manually into a single sided mould that is
preheated. Once this is done a releasing agent is applied through which the resin is
then forced into the fibre mats using hand rollers (Cairns 2000). The fabric is then left
until the resin has been soaked up and any excess resin eradicated. If necessary this
procedure can be repeated to achieve alternating patterns throughout the composite
layers.
A variation to the above method is known as pre-impregnated cloth, whereby each
individual layer is wet with a gel before being placed in the mould (Cairns 2000). This
technique results little to no extra finishing to be done. Once the gel has been applied
the fabric is then cut to size and the resin is weighed out in the correct proportions to
achieve the necessary resin content, this is done in parts per hundred for curing agents
or by weight for epoxy resin. The resin can then be applied onto the fabric using a
brush or a squeegee. A compaction process is then performed using a roller.
Most UAVs are expected to remain in use for at least 7-10 years as pointed out by
Biesecker from Defence Daily in 2003. For the traffic monitoring application in
Australian major cities it is expected that a maximum of 15 UAVs will be required to be
manufactured. As this is a relatively small number a high labour intensive process that
results in a high quality UAV being produced that requires little surface finishing. Thus a
manual hot-wire cutting process can be utilised to produce the foam core and a hand
lay-up process using pre-impregnated cloth technique can be used for the composite
skin.
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3.15 Maintenance
Currently, CASA has no solid maintenance standards for UAVs. Companies and
researchers understand the risks that are to be taken and are currently abiding by
internal standards. Currently there is discussion regarding whether maintenance
standards should be equivalent to conventional aviation standards but modified slightly
to fit size and mission requirements (Adams 2007). Until legal standards are introduced
for UAV maintenance schedules, a suitable maintenance schedule should be developed
for the customer to abide by.
As the UAV will be mainly operated by media outlets and government officials, the
maintenance schedule should reflect the likely lack of technical knowledge of the
operator. Small maintenance tasks can be performed after every flight, or once a day
for example, such as checking fluid or battery levels and inspecting the aircraft for
damage. These tasks do not require any technical expertise so can be carried out by
anyone. Yearly major services should also be carried out by a professional, licensed
aircraft maintenance engineer (LAME). An appropriate maintenance schedule for the
BATMAN UAV is as follows:
Daily: Inspect aircraft visually for damage or corrosion.
Disconnect the batteries when not in use or charging.
Store the UAV in a location sheltered from the weather, in a clean and dry
location.
Batteries must be recharged after each flight to maximum capacity.
Yearly: A full aircraft service must be performed by a LAME.
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3.16 Structural Analysis
The aircraft can endure loads encountered during flight, including tension,
compression, torsion, bending and shear loads, through good materials selection and
structural design. The layout of the internal wing and housing structures is crucial so
that loads can be dissipated and failure can be avoided.
A full finite element analysis using a program such as ANSYS should be conducted on
the internal structure of the UAV to ensure the loads are not excessive, however due to
the scope of this project and the relatively short timeframe this will not be done here.
3.16.1 Wing Structure
The general structure of an aircraft wing comprises of four parts (Brandt et al, 2004):
- Spars: strong beams extending from root to tip along the length of the wing that
carry forces and moments from the spanwise lift distribution.
- Ribs: airfoil-shaped members that run chordwise through the wing and transfer
chordwise pressure and shear loads from the skin to the spars, as well as
assisting in the reduction of wing twist and torsion. They sometimes have areas
cut out to reduce weight.
- Stringers: additional spanwise stiffeners to transfer loads from the skin to the
spars if necessary, and increase the wing’s buckling strength.
- Skin: the outer layer that wraps around the frame, giving the wing its shape, and
first encounters all loads on the wing.
These components come together in a layout similar to that shown in figure 3.16.
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Figure 3.16: wing structural components (Hieserman, 2005)
The flying wing UAV will have two aluminium I-beam spars, located at 14% and 63% of
the chord. The ribs will also be aluminium and spaced at equal intervals along the
length of the wing, with circular regions cut out to reduce weight if necessary. The
number of ribs and the spacing would be determined by a detailed structural analysis in
later parts of the design process (not in the scope of this project). Stringers would not
be necessary since the wings are quite small and the ribs will provide enough support
to transfer loads between the skin and spars. Skins would be made of glass fibre-
reinforced epoxy (fibreglass) and attached to the ribs by riveting or gluing.
3.16.2 Housing Structure
The flying wing UAV does not have a traditional fuselage, but has a region, referred to
as the ‘housing’, which contains the avionics and propulsion systems between the
wings. The frames in the direction perpendicular to the aircraft centreline will be
specially shaped, almost elliptical aluminium members that attach to the spars on
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either wing. The housing could be a monocoque or a semi-monocoque structure
depending on the weight and how much strength is required. A typical fuselage
structure, featuring longerons and stringers running parallel to the aircraft centreline,
would not be necessary for a flying wing UAV. Monocoque structures have no
longerons or stringers, instead using only the frames and a pre-stressed skin structure
to give torsional and bending stiffness. Semi-monocoque structures have a small
number of longerons and stringers to assist in stiffening the skin (Brandt et al, 2004). As
with the wing structure, there are equations to determine all of the loads and the
positions of load-carrying elements, but they will not be considered at this stage of the
design.
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4 Weight and Balance Analysis
Preliminary weight estimations approximated the total weight of the UAV design to be
7.7kg from section 3.3. Subsequent design analyses were able to present more
accurate weight estimations to certain aircraft design components such as the
propulsions systems. Calculations for the aircraft design weight were finalized by
tabulating the calculated weights of certain components of the aircraft and using
statistical methods to obtain a good approximation for the weights of the remainder of
aircraft components. The summary of the aircraft weight was tabulated in Table 4.1
below:
Table 4.1: Aircraft Weight Summary
Components Weight (kg) Method Used to Obtain Weight of Component
Propulsions System 5.0 Based on battery and propeller selection
Avionics 2.5 Based on battery and avionics selection
Structural Weight 3.0 Based on statistics (Wstructure=0.4WTO)
Total Weight 10.5
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The structural weight was taken as a single value rather than divided into sections as
the aircraft design is a flying wing UAV. Therefore, a breakdown of weight for the wing
and the empennage was not applicable.
4.1 Internal Component Configuration
The positioning of the major components onboard the aircraft was vital in ensuring that
stability of the aircraft design was obtained. A summary of each system and the desired
location follows.
Propulsion System
The position of the batteries and the propeller was dictated by the aircraft flying wing
design and the pusher propeller configuration. The battery unit powering the propeller
was located at the rear of the aircraft with the propeller directly coupled to the output
shaft of the motor. This method of assembly negated the need of heavy and complex
gearbox components that would otherwise be required.
Sensor (Camera)
The sensor, or camera, of the aircraft design was positioned in the front of the aircraft
in order to ensure effective traffic monitoring capabilities of the aircraft as required by
the design specifications. In order maximise the Field of View (FOV) of the aircraft the
camera was position in the front of the aircraft but below the main body to enable the
camera a 360o FOV.
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Avionics
The onboard avionics contributes approximately 20% of the overall aircraft weight
resulting in the careful consideration being given to its placement within the main body
of the aircraft. Due to the flexibility of placement of the avionic systems since it is not
predetermined by other factors, the positioning of these systems was finalized in the
centre of gravity (CG) calculations of the aircraft.
Structural Weight
The structural components onboard the aircraft constituted 40% of the aircraft takeoff
weight. The aircraft design decided that the UAV would be launched by a catapult and
caught by a net system. No landing gears were therefore required on the aircraft design
and the catapult release mechanisms were assumed negligible in CG analysis. Since the
aircraft design utilised a tailless configuration, the longitudinal stability and the static
margin of the aircraft will be manipulated via the positioning of the wings. Moving the
wings aft ensures that the CG and neutral point are further apart hence making the
aircraft more stable.
Figure 4.1 presents a schematic layout of the internal components onboard the aircraft
design. This layout provides the most efficient method of installing the components to
provide the required balance and CG location.
92
Figure 4.1: Schematic of Internal Component Configuration Layout
4.2 Centre of Gravity Determination
In order to determine the stability of the aircraft design, the CG position of the total
aircraft was calculated. The CG of the major components on the aircraft was considered
to be located at the geometrical centre of each component. It was assumed that the
components were evenly distributed about the aircraft such that there was no moment
azimuthally (z-axis) thus negating the necessity of the ZCG analysis. Given that the
aircraft retains a constant weight configuration throughout the entire mission profile,
the CG location of the aircraft would remain constant. The CG of the aircraft was
referenced with respect to a pre-defined coordinate system using equation 4.1.
(Equation 4.1)
The coordinate system was located with the origin at the nose of the aircraft, the x-axis
defined to be in the longitudinal direction, the y-axis in the lateral direction and the z-
93
axis was defined azimuthally. The structural load of the entire aircraft was assumed to
act approximately at the geometrical centre of the aircraft design. The XCG of the entire
aircraft was determined to be 0.192m from the nose of the aircraft (See Appendix C3).
The mean aerodynamic centre (MAC) was considered to be the point through which
the lift of the aircraft acts and where the pitching moment of the aircraft is
independent of angle of attack. For the purpose of this analysis, the wing aerodynamic
centre was assumed to be located at 25% of the MAC. For the flying wing configuration
considered in the project, the MAC coincides with the neutral point of the aircraft
(Equation 4.2).
(Equation 4.2)
The MAC of the aircraft was determined to be 0.246m from the nose of the aircraft.
(See Appendix C3). This verifies that the neutral point of the aircraft is behind of the
CG, coupling with a zero net moment around the CG of the aircraft; the aircraft design
has proven to be a stable trimmed aircraft. The CG lying ahead of the neutral point
provides a negative static margin therefore resulting in longitudinal positive stability of
the aircraft.
Figure 4.2 demonstrates the CG envelope of the aircraft design with the positions of
the CG and the neutral point of the aircraft. Static margin was defined as the difference
between the locations of the CG with the neutral point of the aircraft. The static margin
was determined to be 28% which is sufficiently large to reduce the manoeuverability of
the aircraft. This will not be altered here, however, if problems occurred during flight
testing of the prototype the internal avionics components layout can be altered to
change the neutral point of the aircraft to a location more suitable.
94
Figure 4.2: CG Envelope
4.3 Longitudinal Stability
Having calculated the centre of gravity and neutral point for the initial configuration
design in section 4.2, longitudinal static stability was briefly analysed in order to
investigate pitch controllability as a function of configuration. The BATMAN UAV is a
catapult launched, single-engine blended wing configuration, by definition lacking any
empennage, with a constant-in-flight centre of gravity and no span-wise twist,
therefore the only prominent aspect of configuration design applicable to longitudinal
stability is the variation of static margin with sweep. From geometry, the location of
the neutral point may be determined as 25% of the MAC, which may be located as
distance from the nose from equation 4.3, below from MAC position, , and leading-
edge sweep angle, Λ.
(Equation 4.3)
95
In the case of wing sweep, Λ may realistically be investigated for values less than 30°
(≈0.524 radians) for which values the relationship between Λ and tan(Λ) may
reasonably be assumed to be linear. Using this approximation, and the assumption that
the neutral point for this UAV is always located at 25% MAC, only two points were
required in order to form a plot of static margin versus sweep. The first of these points
was taken to be 28° sweep, 28% static margin, as determined in section 4.2, with the
second calculated from a 0° sweep angle with a static margin of -27% (with the CG of
the structure being located at the intersection of mid-span and mid-chord, resulting in
a shift of less than 5mm forwards in the overall CG of the BATMAN UAV). Plotting a
straight line through these points resulted in Figure 4.3, which shows longitudinal
stability as a function of sweep angle from which it can be seen that a 10% static
margin, which would ease controllability of vehicle in comparison to the current 28%
static margin, would require approximately 17° degrees of leading-edge sweep and that
the BATMAN UAV would be neutrally stable if designed with a leading-edge sweep
angle of approximately 13.5°.
-10
0
10
20
30
40
10 12 14 16 18 20 22 24 26 28 30
Sta
tic
Ma
rgin
(%
MA
C)
Λ(°)
Longitudinal Stability
Figure 4.3: Longitudinal Static Margin versus Leading Edge Sweep Angle
96
97
5 Aerodynamic Analysis
The aerodynamic parameters, lift distribution and L/D of the UAV, will be determined
to allow an understanding of the aerodynamic performance of the aircraft. The
aerodynamic parameters provide important information regarding how the aircraft will
operate and will allow the final performance parameters to be analysed.
5.1 Lift Distribution
The total lift generated by a wing can be separated into two components; a basic lift
component and an additional lift component (Abbott & von Doenhoff 1949). The basic
lift component is a function of twist, wing area, chord, wing span, effective lift-curve
slope and the coefficient Lb as shown in equation 5.1. The coefficient Lb is dependant
on taper ratio and aspect ratio and is determined from table 1 in Abbott and von
Doenhoff (1949).
(Equation 5.1)
98
The additional lift component is a function of wing area, chord, wing span and the
coefficient La as shown in equation 5.2 The The coefficient La is again dependant on
taper ratio and aspect ratio and is determined from table 2 in Abbott and von Doenhoff
(1949).
(Equation 5.2)
The two lift components are combined by equation 5.3 to provide a span-wise lift co-
efficient. A graph depicting the span-wise lift co-efficient for this UAV is shown in figure
5.1 with the calculations in Appendix C4.
Lift Distribution along Wing
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
0 0.5 1 1.5 2 2.5 3 3.5
Distance along wing (ft)
Sec
tio
n L
ift
Co
-eff
icie
nt
Figure 5.1: Span-wise lift distribution along the wing
99
5.2 L/D Determination
During the sizing section, an estimate for L/D of 10 was used as a flying wing UAV can
be considered similar to a homebuilt light aircraft. However, this was an approximation
to allow a basic sizing of the aircraft to be estimated; therefore an actual value for L/D
needs to be calculated to allow the final performance parameters to be specified.
To calculate L/D of an aircraft, the wetted area to reference area needs to be
calculated. Wetted area is described as the surface area of the aircraft that would
become wet if the aircraft was immersed in water (Raymer 2006). The reference area is
described as the area of the top side of the wings, including the area of fuselage
between the two wings. A rough estimate of final body shape was drawn and the
wetted and reference areas calculated from it. The sketch and calculations are included
in Appendix C5.
26.782852 mmSwet =
20.359390 mmS ref =
The ratio of Swet to Sref is therefore
178281.2=ref
wet
S
S
For ease of calculations, an Swet/Sref of 2.2 will be used.
100
The wetted aspect ratio is then 4.55 by equation 5.3.
Wetted Aspect Ratio = 55.42.2
10 ==
ref
wetS
SA
(Equation 5.3)
The maximum L/D of the aircraft is then estimated from a statistical graph such as figure 3.6 in
Raymer (2006). However, as this flying wing UAV has a high aspect ratio and small Swet/Sref
ratio, the graph does not extend to include a wetted aspect ratio of 4.55. Therefore, by
extrapolation of the line for a fixed gear, single propeller aircraft, an L/D maximum is
estimated to be 19 as shown in figure 5.2 below.
Figure 5.2: Statistical L/D determination graph (Raymer, 2006)
101
6 Performance Analysis
The initial sizing analysis led to locating a design point from the matching diagram, as
seen in section 3. The takeoff weight estimate, design wing and power loading, and
specified objectives from earlier sections formed the basis for the configuration design,
where the actual aircraft specifications were obtained. The performance analysis in this
section is akin to ‘sizing in reverse’: using the actual specified values to determine if the
aircraft meets the technical task design requirements and regulations where necessary.
6.1 Weight, Wing Loading and Power Loading
The new takeoff weight was calculated in previous sections from the weights of all
components of the aircraft. The wing area and ratio of wetted area to reference area
were calculated directly from the drawings, and the power of the electric engine was
found in section 3.9. Using these values, the wing loading and power loading of the
aircraft was calculated. The new values are stated in table 6.1 along with the design
values for comparison.
102
Table 6.1: Comparison of weight, area and loading values
Parameter Design value Final value
WTO (lbs) 16.98 23.15
S (ft2) 3.875 3.98
Swet/Sref 2 2.2
W/S (lbs/ft2) 4.263 5.97
W/P (lbs/hp) 40.3 46.3
The increase in takeoff weight combined with a decrease in wing area, has led to a
larger wing loading than the original design value. With increased wing loading the
aircraft will be less susceptible to turbulence and the effects of wind gusts, but also less
manoeuverable. The performance of the aircraft should be improve during adverse
weather conditions with a higher wing loading, but care should still be taken in such
conditions.
The power ratio has also increased due to the increase in takeoff weight, but is still an
acceptable value and the engine is capable of providing enough output power to fly the
aircraft. The modeling of the aircraft using CAD software has led to a slight increase in
the ratio of wetted area to reference area, which shows that the concept sketches
allowed a fairly accurate value to be estimated.
The final matching diagram can be assembled using previous sizing equations from
section 3 and the values calculated in this section. Many parameters previously
estimated are now known from the conceptual design, including L/D, CL, S, Swet/Sref,
Vstall and Vcr. Updating the matching diagram to involve these new values results in a
new matching diagram as shown in figure 6.2 below. The design point marked as a red
cross, corresponding to the new wing and power loading values.
103
Figure 6.2: New matching diagram and design point
6.2 Sensitivity to new values
A new sensitivity analysis was conducted to examine the sensitivity of takeoff weight to
payload weight; the same parameters involved in the original sensitivity analysis in
section 3.4. Sensitivity equations 3.1 to 3.3 were used to calculate these new values
with the results shown in table 6.3. As before, the payload weight is taken to be the
weight of the avionics.
Table 6.3: Sensitivity parameters
Sensitivity parameter Calculated value
B 0.97609
C 1
D (lbs) 5.512
WTO (lbs) 23.15
4.56 lbs
104
Therefore, for every pound added to the UAV, the takeoff weight will increase by 4.56
pounds making the UAV more sensitive to weight increase than in the initial analysis.
6.3 Cruise Speed and Stall Speed
The values for wing and power loading found in section 6.1 and the density ratio from
previous cruise speed sizing (section 3.5.8), enabled a new power index to be
calculated using equation 3.9. The new cruise speed was determined using statistics
from Roskam (2005) and the new value of wing loading with equation 3.4 was used to
calculate the required stall speed. These values and the corresponding equation
between wing and power loading are shown in table 6.4.
Table 6.4: New cruise and climb values
New power index Ip 0.513
New power index
equation
New cruise speed (ft/s) 125
New stall speed (ft/s) 64.7
The stall speed is slightly higher than the estimated design stall speed and is still an
ambitious value for the chosen launch setup, however will be maintained in the design.
The cruise speed for maximum engine output, as calculated in table 6.4, is higher than
the required cruise speed, therefore the aircraft will comfortably achieve the required
speed without excess work by the engine.
Using the maximum cruise speed, an improved range can be calculated to give an
upper bound on the flight path radius and the distance that can be traveled by the
UAV. The flight path radius, with a maximum speed of 137.2 km/h, is now 11 km, which
gives a cruise distance of 415 km over the assumed three hour flight time.
105
6.4 Summary
The compliance of the UAV with performance parameters and guidelines set in the
technical task is summarized in table 6.5. All design values satisfy the required
guidelines outline in the technical task.
Table 6.5: Compliance of performance parameters
Performance Parameter Value Compliance
Weight 23.15 lbs Increased from design value
but still acceptable.
Cruise speed 125 ft/sec Engine capable of achieving
this speed.
Cruise altitude 1500 ft As per technical task.
Flight range 415 km Increased from design value;
batteries capable of supplying
enough power.
Radius of flight path 11 km Increased from design value.
Rate of climb 300 fpm As per sizing section.
Takeoff/landing requirements - Catapult launch and net
landing, as per technical task
and conceptual design.
Flying wing design - Maintained with blended-
wing-body design, seen in
drawings.
Engine Baldor premium efficiency AC
motor, 0.5 hp
Provides required power to
fly UAV.
Fuel type Electric Provided by six zinc-air
batteries.
106
107
7 Technical Drawings
All drawings are attached at the end of the project
7.1 Three View Drawings
Drawing number 1
7.2 Aircraft Layout Drawing
Drawing number 2
7.3 Exploded View Assembly Drawing
Drawing number 3
7.4 Wing Detail Drawing
Drawing number 4
7.5 Airfoil Drawing
Drawing number 5
108
109
References
Abbott, I.H. & von Doenhoff, A.E. 1959 Theory of Wing Sections, Dover Publications, New York.
Adams C. 2007 Technology Focus – Unmanned Aerial Vehicles. Aviation Maintenance
Magazine, Access Intelligence,
<http://www.aviationtoday.com/am/categories/military/Technology-Focus-Unmanned-
Vehicle-Maintenance_16794.html>
Advanced Aluminium Design, 2007, Properties of Aluminium, accessed 15/05/2010,
http://www.aadl.co.uk/properties-of-aluminium.html
Air-Attack.com, 2010, BAE Corax UCAV Factsheet, Air-Attack.com - Military Aviation news and
media, <http://air-attack.com/page/74/BAE-Corax-UCAV.html>
Arjomandi, M, 2010 Aircraft Design Notes. School of Mechanical Engineering, Adelaide
Australian Bureau of Statistics (2009), ‘Australian Demographic Statistics’, Publication 3101.0,
Australia, accessed 20/4/2010
<http://www.abs.gov.au/AUSSTATS/[email protected]/Lookup/3101.0Main+Features1Sep%202009?Op
enDocument>
110
Australian Government: Bureau of Meteorology 2010a, Beaufort Wind Scale, accessed
29/03/2010, http://www.bom.gov.au/lam/glossary/beaufort.shtml
Australian Government: Bureau of Meteorology 2010b, Climate Statistics for Australian
locations: Adelaide West Terrace, accessed 29/03/2010,
http://www.bom.gov.au/climate/averages/tables/cw_023000_All.shtml
Australian Government: Bureau of Meteorology 2010c, Climate Statistics for Australian
locations: Brisbane Regional Office, accessed 29/03/2010,
http://www.bom.gov.au/climate/averages/tables/cw_040214_All.shtml
Australian Government: Bureau of Meteorology 2010d, Climate Statistics for Australian
locations: Canberra Airport Comparison, accessed 29/03/2010,
http://www.bom.gov.au/climate/averages/tables/cw_070014_All.shtml
Australian Government: Bureau of Meteorology 2010e, Climate Statistics for Australian
locations: Darwin Post Office, accessed 29/03/2010, http://www.bom.gov.au/climate
/averages/tables/cw_014016_All.shtml
Australian Government: Bureau of Meteorology 2010f, Climate Statistics for Australian
locations: Hobart (Ellerslie Road), accessed 29/03/2010,
http://www.bom.gov.au/climate/averages/tables/cw_094029_All.shtml
Australian Government: Bureau of Meteorology 2010g, Climate Statistics for Australian
locations: Melbourne Regional Office, accessed 29/03/2010,
http://www.bom.gov.au/climate/averages/tables/cw_086071_All.shtml
Australian Government: Bureau of Meteorology 2010h, Climate Statistics for Australian
locations: Perth Regional Office, accessed 29/03/2010,
http://www.bom.gov.au/climate/averages/tables/cw_009034_All.shtml
111
Australian Government: Bureau of Meteorology 2010i, Climate Statistics for Australian
locations: Sydney (Observatory Hill), accessed 29/03/2010,
http://www.bom.gov.au/climate/averages/tables/cw_066062_All.shtml
Azom, 2010, E-glass Fibre, accessed 15/05/2010,
http://www.azom.com/Details.asp?ArticleID=764
Baldor 2010 Specifications: EM3538. Accessed 03/06/10 online at URL:
http://baldor.com/products/detail.asp?1=1&page=1&catalogonly=1&catalog=EM3538&produc
t=AC+Motors&family=Premium+Efficiency|vw_ACMotors_PremiumEfficiency&winding=&ratin
g=#
Batill, S., Stelmack, M. & Yu, Q. 1999 Multidisciplinary design optimization of an electric-
powered unmanned aerial vehicle. Aircraft Design 2 pp 1-18
Bayliss, R., Collins, C., French, C., Pham, E., Valiyff, A. (2007), ‘Design and Build of a Tailless Fuel
Cell Powered UAV for Surveillance Applications’, viewed on 5 May 2010
<http://www.mecheng.adelaide.edu.au/~marjom01/Honours%20projects/2008%20honours%2
0projects/Design%20and%20build%20of%20a%20fuel-cell%20powered%20UAV.htm>
Biesecker C, 2003, United Industrial Sees $1 Billion In Potential Shadow UAV Sales To Army,
accessed 15/05/2010, http://findarticles.com/p/articles/mi_6712/is_3_218/ai_n28999842/
Borchardt A, 2004, Unmanned Aerial Vehicles spur composite use, accessed 15/05/2010,
http://www.reinforcedplastics.com/view/1423/unmanned-aerial-vehicles-spur-composites-
use/
Brandt, S., Stiles, R., Bertin, J., Whitford, R. (2004), ‘Introduction to Aeronautics: A Design
Perspective 2nd Edition’, American Institute of Aeronautics and Astronautics, Virginia, USA.
112
Brown S, 2007, Forging, Casting and CNC Machining, Harris Cyclery,
<http://www.sheldonbrown.com/dp-forging.html>
BTC-88 – Micro Pan / Tilt Unit 2008, Unmanned Airsystems and Components
BTC-88 Gimbal 2009, Procerus Technologies
Buretta A, Fowler R, Germroth M et al, 2003, Low-cost expendable UAV, Virginia Polytechnic
Institute and State University, pg 44-48
BusinessWire http://satellite.tmcnet.com/news/2010/05/14/4788327.htm
Cairns D & Skramstad J, 2000, Evaluation of Hand lay-up and Resin Transfer Molding in
Composite Wing Turbine Blade Manufacturing, Sandia National Laboratories, USA, pg 1-4
Chan K, Forrester C, Lomas I et al, 2009, 859 Design and Build of a UAV with morphing
configuration, University of Adelaide School of Mechanical Engineering, pg 108-114
CleanTech http://cleantech.com/news/3694/electrovaya-tata-motors-make-electric-indica
Confederation of Australian Motor Sport 2003 Material Safety Data Sheet - Avgas 100 LL
accessed 24 May 2010 online at
http://www.cams.com.au/en/Safety/Safety%20in%20Motor%20Sport/Safety%201st%20at%20
events/~/media/Files/Safety/Material%20data%20Safety%20sheets/MSDS_Avgas100LL.ashx
Defence Material Organisation, 2006, UAV – autonomous decision making Jae Box, Australian
Government Department of Defence,
<http://www.defence.gov.au/teamaustralia/UAV_autonomous_decision-making_(JAE-
Box).htm>
Defence Update http://defense-update.com/products/z/zinc-air-battery-new.htm
113
Dorgham M.A, 1986, Designing with Plastics and Advanced Plastic Composites, Inderscience
Enterprises, Switzerland, pg 107 -128
Dube L, McElroy A & Pepper D, 2010, Sun-Powered Flight, COMSOL, last accessed 06 June 2010,
http://www.comsol.com/stories/unlv_sun_powered_flight/full/
FCB-H11 Brochure 2008, Sony Corporation
FCBH11: High Definition Color Block Camera n.d., Sony Corporation, accessed 26-4-2010
<http://pro.sony.com/bbsc/ssr/cat-industrialcameras/cat-highdefinition/product-FCBH11/>
FCBIX11AP: 10x Color EXview PAL Block Camera n.d., Sony Corporation, accessed 26-4-2010
<http://pro.sony.com/bbsc/ssr/mkt-industrialautomation/mkt-
industrialautomationtraffic/product-FCBIX11AP/>
Foale T, 2006, Motorcycle Handling and Chassis Design the art and science 2nd edition, Tony
Foale Designs, pg 13-13
Gizmag, 2010, VTOL Flying- Wing: a new take on UAV design, Gizmag,
<http://www.gizmag.com/flying-wing-vtol-uav/13962/>
Hepperle, M (2006), Basic Design of Flying Wing Models, viewed on 1 May 2010
<http://www.mh-aerotools.de/airfoils/flywing1.htm>
Jayabalan N, Horng L & Leng G, -, Reverse Engineering and Aerodynamic Analysis of a Flying
Wing UAV, Aeronautical Engineering Group National University of Singapore, pg 5-6
Jayabalan N, Horng L & Leng G, -, Reverse Engineering and Aerodynamic Analysis of a Flying
Wing UAV, Aeronautical Engineering Group National University of Singapore, pg 5-6
114
Jones, GP 2003, ‘The Feasibility of Using Small Unmanned Aerial Vehicles for Wildlife Research’,
MSc Thesis, University of Florida, Florida.
Key to Metals, 2010, Metal Properties, Key to Metals, http://www.keytometals.com
Kalpakjian, S. and Schmid, S. (2006), ‘Manufacturing Engineering and Technology 5th Edition’,
Pearson Education, South Asia.
Kopeliovich D, 2008, Thermoplastic Acrylonitrile-Butadiene-Styrene (ABS), Substances and
Technologies, accessed 15/05/2010,
http://www.substech.com/dokuwiki/doku.php?id=thermoplastic_acrylonitrile-butadiene-
styrene_abs
Lee, K 2004, ‘Development Of Unmanned Aerial Vehicle (UAV) for Wildlife Surveillance’, MSc
Thesis, University of Florida, Florida.
MP-DAYVIEWPTZ/MP-NIGHTVIEWPTZ Stabilized Payload Cameras 2006, MicroPilot.com
MSN Encarta, 2009, UAV (plural UAVs), Encarta World English Dictionary Bloomsbury
Publishing,<http://uk.encarta.msn.com>
Puri, A, 2005, ‘A Survey of Unmanned Aerial Vehicles (UAV) for Traffic Surveillance’, University
of South Florida, Florida.
Raymer, D.P. 2006 Aircraft Design: A Conceptual Approach Fourth Edition. American Institute of
Aeronautics and Astronautics Inc. Virginia
Ro, K, Oh, J & Dong, L 2007, ‘Lessons Learned: Application of Small UAV for Urban Highway
Traffic Monitoring’, paper presented to the 45th AIAA Aerospace Sciences Meeting and Exhibit,
Reno, Nevada, 8 - 11 January 2007.
115
Road and Traffic Authority, NSW n.d., Live Traffic, accessed 21/03/2010,
http://www.rta.nsw.gov.au/trafficreports/
Robinson Helicopter Company 2007, R44 Raven II Newscopter: Specifications, accessed
21/03/2010, http://www.robinsonheli.com/r44newscopter_new_specs.htm
Robinson Helicopter Company 2010, R44 Raven II Newscopter: 2010 Price List, accessed
21/03/2010, http://www.robinsonheli.com/pdf_files/newscopter_pricelist.pdf
Roskam, J. 2005 Aircraft Design Part 1: Preliminary Sizing of Aircraft. DARcorporation, Kansas
ShopBot, 2010, 3D and CNC Routing, ShopBot,< http://www.shopbottools.com/3-
d_work_v2.html >
SP Systems, Composite Engineering Materials, SP Systems
SPAWAR Systems Centre San Diego, 2004, AUMS, SPAWAR,
<http://www.spawar.navy.mil/robots/air/aums/aums.html>
Statzer M, Neblett E, et al, 2003, Low cost expendable UAV Project, Virginia Tech University
Department of Aerospace and Oceanic Engineering, pg 84-86
Strong, A., Snyder, D., Takach, T. (2009), ‘Unmanned Vehicles: Seeing around the Corner and
over the Horizon’, American Composites Manufacturers Association, Virginia, USA.
Surveillance Camera Players, -, on the use of surveillance cameras to enforce traffic laws,
accessed 15/05/2010, http://www.notbored.org/traffic-cameras.html
TASE Small Gyro Stabilized Camera Gimbal n.d., Cloud Cap Technology
116
The Australian Helicopter Directory 2010, Australian Traffic Network, accessed 21/03/2010,
http://www.helidirectory.com/directory/nsw/australian-traffic-network
The Free Library
http://www.thefreelibrary.com/Arotech%27s+Batteries+and+Power+Systems+Division+to+Dev
elop+Zinc-Air...-a0127365440
The Future of Things, 2007, Solar UAV to set a new world record,
<http://thefutureofthings.com/articles/51/solar-uav-to-set-a-new-world-record.html>
The State of Queensland (Department of Main Roads) 2008, Traffic cameras, accessed
21/03/2010, http://www.ourbrisbane.com/transport/traffic-cameras
The UAV, -, UAV Types, The UAV
Transport SA 2010, Traffic Control Centre, accessed 21/03/2010,
http://www.transport.sa.gov.au/transport_network/traffic_ops/traffic_control_cnt.asp
[Unknown author] (2010), UIUC Coordinates Database, viewed on 25 April 2010
<http://www.ae.illinois.edu/m-selig/ads/coord_database.html>
Urban Ecology Australia, 2006, Solar Panels, Urban Ecology Australia,
<http://www.urbanecology.org.au/topics/solarpanels.html>
VicRoads n.d., Live Traffic Cameras, accessed 21/03/2010, http://livetraffic.vicroads.vic.gov.au/
Warick G, 2008, Ares –A Defense Technology, McGraw Hill
<http://www.aviationweek.com/aw/blogs/defense>
117
Williams W, 2003, UAV Handling Qualities….You must be joking, Aerospace Sciences
Corporation, pg 2-3
Wong K.C, 1997, Aerospace Industry Opportunities in Australia- Unmanned Aerial Vehicles
(UAVs) Are they ready this time? Are we?, Royal Aeronautical Society, pg 6
118
119
Appendix A – Raw Data for Statistical Analysis
Weight
Manufacturer Model
Max Take-
Off Weight
(kg) log(Wto)
Empty
Weight
(kg) log(We)
AAI Corporation
RQ-7B
Shadow 200 170 2.230448921 200.6 2.302330929
AAI Corporation Aerosonde 4 12 1.079181246 18.1 1.257678575
Pioneer UAV
RQ-2B
Pioneer 204.12 2.30988556 276 2.440909082
ATE Vulture 150 2.176091259 115 2.06069784
Aerovision Fulmar 19 1.278753601 11.5 1.06069784
BlueBird
Aerosystems Micro B 1.1 0.041392685 1.1 0.041392685
BAE Systems Phoenix 175 2.243038049 220 2.342422681
BAE Systems Coyote 5.5 0.740362689 5.5 0.740362689
Length
Manufacturer Model
Max Take-
Off Weight
(kg) log(Wto) Length (m) log(length)
AAI Corporation
RQ-7B
Shadow 200 170 2.230448921 3.41 0.532754379
AAI Corporation Aerosonde 4 12 1.079181246 2.1 0.322219295
Pioneer UAV
RQ-2B
Pioneer 204.12 2.30988556 4.27 0.630427875
ATE Vulture 150 2.176091259 3.1 0.491361694
Aerovision Fulmar 19 1.278753601 1.23 0.089905111
BlueBird
Aerosystems Micro B 1.1 0.041392685 0.7 -0.15490196
BAE Systems Phoenix 175 2.243038049 3.8 0.579783597
BAE Systems Coyote 5.5 0.740362689 0.9
-
0.045757491
Wing Span
Manufacturer Model
Max Take-
Off Weight
(kg) log(Wto) Span (m) log(span)
AAI Corporation
RQ-7B
Shadow 200 170 2.230448921 4.25 0.62838893
120
AAI Corporation Aerosonde 4 12 1.079181246 2.9 0.462397998
Pioneer UAV
RQ-2B
Pioneer 204.12 2.30988556 5.15 0.711807229
ATE Vulture 150 2.176091259 5.2 0.716003344
Aerovision Fulmar 19 1.278753601 3.1 0.491361694
BlueBird
Aerosystems Micro B 1.1 0.041392685 0.95
-
0.022276395
BAE Systems Phoenix 175 2.243038049 5.5 0.740362689
BAE Systems Coyote 5.5 0.740362689 1.75 0.243038049
Cruise Speed
Manufacturer Model
Max Take-
Off Weight
(kg) log(Wto)
Cruise
Speed (kts)
log(cruise
speed)
AAI Corporation
RQ-7B
Shadow 200 170 2.230448921 90 1.954242509
AAI Corporation Aerosonde 4 12 1.079181246 50 1.698970004
Pioneer UAV
RQ-2B
Pioneer 204.12 2.30988556 65 1.812913357
ATE Vulture 150 2.176091259 65 1.812913357
Aerovision Fulmar 19 1.278753601 54 1.73239376
BlueBird
Aerosystems Micro B 1.1 0.041392685 46 1.662757832
BAE Systems Phoenix 175 2.243038049 85 1.929418926
BAE Systems Coyote 5.5 0.740362689 50 1.698970004
Endurance
Manufacturer Model
Max Take-
Off Weight
(kg) log(Wto)
Endurance
(hrs)
log(enduranc
e)
AAI Corporation
RQ-7B
Shadow 200 170 2.230448921 6 0.77815125
AAI Corporation Aerosonde 4 12 1.079181246 30 1.477121255
Pioneer UAV
RQ-2B
Pioneer 204.12 2.30988556 5 0.698970004
ATE Vulture 150 2.176091259 3.5 0.544068044
Aerovision Fulmar 19 1.278753601 8 0.903089987
BlueBird
Aerosystems Micro B 1.1 0.041392685 1 0
BAE Systems Phoenix 175 2.243038049 4 0.602059991
BAE Systems Coyote 5.5 0.740362689 1.5 0.176091259
121
Appendix B – VTC Charts for various Australian Cities
Melbourne VTC
122
Sydney VTC
123
Canberra VTC
124
Hobart VTC
125
Appendix C – Hand Calculations
C1 – Empirical constants A and B
logWE = y + xlogWTO
logWE = -0.0187 + 1.0245logWTO
logWTO = (logWE +0.0187)/1.0245
logWTO = 0.01825 + 0.97609logWE
Therefore,
A= 0.01825
B = 0.97609
C2 – Sizing Calculations (performed using Excel)
DRAG POLAR ESTIMATION A = 10
C_fe S_wet/S_ref C_D0 A e pi*A*e
0.006 2 0.012 5 0.7 e = 0.7 e = 0.75 e = 0.8 e = 0.85
6 0.75 10.99557 11.78097 12.56637 13.35177
7 0.8 13.19469 14.13717 15.07964 16.02212
8 0.85 15.3938 16.49336 17.59292 18.69248
9 17.59292 18.84956 20.10619 21.36283
10 19.79203 21.20575 22.61947 24.03318
21.99115 23.56194 25.13274 26.70354
FAR 23.65 CGR SIZING A = 10
C_L 1.6
C_D 0.113859
L/D 14.05245
CGR 0.1
CGRP 0.135315
(W/P)(W/S)^(1/2) 98.13368
Pmax/P_TO 89.21243
126
FAR 23.77 CGR SIZING A=10
C_L 1.6
C_D 0.113859
L/D 14.05245
CGR 0.05
CGRP 0.095787
(W/P)(W/S)^(1/2) 138.6306
FAR 23.65 RoC
SIZING TO flaps, no LG
RC (fpm) 300
RCP (33000)^-1 x RC 0.009091
η_p 0.7
C_L^(3/2)/C_D max 1.345(Ae)^(3/4)/C_Do^(1/4) 19.33056
RCP = 0.7 /(W/P) -
(W/S)^(1/2)
/ 367.2807
CRUISE SPEED SIZING
σ (1500 ft) 0.956878418
Speed (ft/s) 115
Speed (mph) 78.40907
I_p (Roskam p.163) 0.48
(I_p)^3 * sigma 0.105823098
1/(I_p)^3*sigma 9.449732796
STALL SPEED SIZING Raymer
density 0.002377
V (km/h) 60
V (fps) 54.678
C_L max (design) 1.4 1.3 1.2
W/S 4.974534682 4.619211 4.26388687
127
C3 - Centre of Gravity Calculations
The weight of each major component onboard the aircraft was considered. The centre of
gravity of each component was assumed to be the geometrical centre of the component itself
(dimensions and masses were obtained from manufacturer data).
The coordinate reference system used had the origin positioned at the nose of the aircraft. The
centre of gravity of the aircraft was therefore determined to be 0.192m from the nose of the
aircraft.
Mean Aerodynamic Calculations
The taper ratio of the aircraft design:
Utilizing the same coordinate system with the origin positioned at the nose of the aircraft, the
mean aerodynamic centre of the aircraft was determined to be 0.246m from the nose of the
aircraft. In the case of this particular aircraft design, given that it is a flying wing without a tail,
the mean aerodynamic centre of the aircraft also coincides with the neutral point of the
aircraft design.
128
C4 - Lift Distribution
c_t/C-o = 0.34/0/.85 = 0.4 and A =10 so from table 1 in Abbott b= 6.31 ft No twist to consider so ε=1 S= 3.98 ft^2 c_o = 0.85 Ft c_t = 0.34 Ft C_L=1.2 1.2 y/(b/2) 0 0.2 0.4 0.6 0.8 0.9 0.95 0.975 y(ft) 0 0.631 1.262 1.893 2.524 2.8395 2.99725 3.07612 L_b -0.321 -0.225 -0.017 0.131 0.195 0.197 0.162 0.115 L_a 1.355 1.265 1.132 0.961 0.748 0.59 0.46 0.343 c 0.85 0.748 0.646 0.544 0.442 0.391 0.3655 0.35275 c_lb 0.23820 0.18973 0.01660 -0.15189 -0.27827 -0.3177 -0.2795 -0.2056 c_la1 1.00548 1.06670 1.10527 1.11424 1.06741 0.95176 0.79382 0.61331 Total 1.44478 1.46977 1.34292 1.18520 1.00263 0.82432 0.67302 0.53034
129
C5- L/D Determination
All measurements are in mm2 calculated from the 3-D ProE model of the aircraft created.
195652 Body (Upper)
14943.1
1170.9
115992
4211.33
1376.4
876.037
8562.62
1066.04
359390 Wing top
52486.744 Rest of wing
9314.35 Prop shaft
17811.0664 prop
Swet = 782852.6 incl prop etc
Swet = 755727.2
Sref = 359390
Swet/Sref = 2.178281 incl prop etc
Swet/Sref = 2.102805