UAV for Traffic Monitoring

Broadcasting Autonomous Traffic Monitoring Aerial Network (BATMAN)

Mei Cheong Michael Evans Elise Fahy Amelia Greig Natasha Parker

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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 1150802 Date:

……………………………………………………………… Michael Evans 1164762 Date:

……………………………………………………………… Elise Fahy 1161673 Date:

……………………………………………………………… Amelia Greig 1149212 Date:

……………………………………………………………… Natasha Parker 1153729 Date:

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Marking Scheme

Group mark Criteria 1. Project definition 2. Research activities 3. technical calculation 4. Drawings 5. Format of the report 6. Novelty of the solution Mark (total 100) /10 /15 /25 /25 /10 /15

Project mark Group member Group mark (50% x project mark) 1. M Cheong 2. M Evans 3. E Fahy 4. A Greig 5. N Parker Individual mark (50% x project mark) Project mark (Total 100)

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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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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

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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

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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

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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

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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

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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

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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

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Nomenclature

Acronyms
AR CASA CASR FOV GC GCN CGR GPS MAC MAV Aspect Ratio Civil Aviation Safety Authority Civil Aviation Safety Regulations Field of View Centre of Gravity Guidance and control interface Climb Gradient Ratio Global Positioning System Mean Aerodynamic Chord/Mean Aerodynamic Centre Micro Aerial Vehicle Maximum Takeoff Weight Revolutions per Minute Short Takeoff and Landing Unmanned Aerial Vehicle Visual Terminal Chart Vertical Takeoff and Landing

MTOW RPM STOL UAV VTC VTOL -

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Symbols
A b c CDo CD CL d D E e g I L L/D n ns P q R Re S Aspect Ratio Wing span Chord length Drag Polar Drag coefficient Lift coefficient Diameter Drag Endurance Oswald efficiency factor Gravitational constant, 9.81m/s^2 Moment of inertia Lift Lift to drag ratio load factor safety factor Power Dynamic pressure Range Reynolds Number Wing Area

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SM T V Vcr Vto W x y z

-

Static margin Thrust Velocity Cruise speed Takeoff speed Weight Longitudinal axis position Lateral axis/span-wise position Vertical axis position

α γ ε λ μ ν ρ σ

-

Angle of attack Climb angle Efficiency Taper ratio Friction co-efficient Viscosity Air density Density ratio

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Subscripts
airfoil battery cl cr des e f i l los max min OL p p airfoil battery climb cruise descent empty final initial loiter zero lift produced at wing root maximum minimum zero lift propeller prop propeller payload root stall static takeoff vertical

payload root stall static TO v -

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w wet wing x y z

-

main wing wetted area main wing section relating to the x axis relating to the y axis relating to the z axis

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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.

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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.

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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

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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.

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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

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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

Take-off Weight Hover Ceiling Cruise Speed Range Fuel Consumption Price

2 500lb 8 950 ft 113 kts 484 km 0.173 kg/km $744 000 USD

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.

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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).

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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.

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Table 2.2: Comparative study of Historical data

Manufacturer AAI Corporation AAI Corporation Pioneer UAV ATE Aerovision BlueBird Aerosystems BAE Systems BAE Systems

Model RQ-7B Shadow 200 Aerosonde 4 RQ-2B Pioneer Vulture Fulmar Micro B Phoenix Coyote

Range Class Tactical LALE Tactical Tactical Tactical

Tactical 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.

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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) 2.302330929 1.257678575 2.440909082 2.06069784 1.06069784 0.041392685 2.342422681 0.740362689

RQ-7B Shadow 200 170 Aerosonde 4 RQ-2B Pioneer Vulture Fulmar Micro B Phoenix Coyote 12 204.12 150 19 1.1 175 5.5

2.230449 200.6 1.079181 18.1 2.309886 276 2.176091 115 1.278754 11.5 0.041393 1.1 2.243038 220 0.740363 5.5

Figure 2.1: Technology Diagram - Empty Weight

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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

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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

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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.

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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
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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.

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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

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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 Large (> 3 million people) Medium (1-3 million people)

City Sydney Melbourne 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

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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)

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(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)

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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)

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(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)

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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: R15 = 2 × 2π × 15km × 3hrs = 566km 10km Radius: R10 = 2 × 2π × 10km × 3hrs = 377km

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.
V15 = 2 × 2π × 15km = 189.0km / hr 1hr 2 × 2π × 10km = 125.6km / hr 1hr

V10 =

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.

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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

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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

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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.

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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

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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

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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.

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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.

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Table 2.5: Main on-board avionics systems

GPS Autopilot Remote Control Piloting Computer Transmitter Actuators Camera

Location, navigation and autopilot reference Allows the UAV to navigate independently Can override the autopilot if the pre-determined flight path needs to be changed Controls all onboard avionics system Allows communication between aircraft and ground control Operates the control surfaces 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.

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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.

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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

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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

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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

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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

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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:

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Table 3.1 – Avionics systems weight distribution

Component Datalink (Communications) Navigation (Autopilot) Control Surface Actuator (Servo Motors) x4 Sensor (Camera) Electrical Supply (Batteries) x4 Total Weight

Weight (kg) 0.3 0.016 0.068 0.465 1.616 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.

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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,

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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.

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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)
W = 4.264lbs / ft 2 S

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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 (2) Take off flaps/gear up (3) Landing flaps/gear up

0.012 0.012

0 0.010

0.012 0.022

0.85 0.8

0.012 + CL2/13.35 0.022 + CL2/12.57

1.2 1.4

0.012

0.055

0.067

0.8

0.067 + CL2/11.78

1.4

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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

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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)

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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 FAR 23.65 Rate of Climb Equation

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:
W W = 0.106 S P

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.

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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:

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W = 4.263lbs / ft 2 S W = 40.3lbs / hp P

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 Wing area, S Aspect ratio, A Wing span, b Mean aerodynamic chord, c Root chord, c0 c= Taper ratio, λ Tip chord, ct Quarter-chord sweep angle Leading-edge sweep angle Angle of twist

Equation (from sizing section) (selected in section __) A = b2/S A = b/c

Value 3.98 ft2 10 6.31 ft (1.92 m) 0.631 ft (19.2 cm) 0.85 ft (25.9 cm)

(selected) λ = ct/c0

0.4 0.34 ft (10.4 cm) 20° 28° 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

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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

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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 Eppler 334 Eppler 342 Clark YH MH 70 MH 78

Stall angle 11° 13° 14° 12° 19°

CL 1.498 1.468 1.432 1.438 1.7

CMα -0.059 -0.025 -0.063 -0.060 0.011

CMo -0.041 -0.0004 -0.016 -0.048 0.041

CDo 0.01675 0.00889 0.00833 0.01844 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

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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.

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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.

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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

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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) Downward

R1 Elevon

R2 Elevon (upper surface)

R2 Elevon (lower surface) Downward

Pitch Down Pitch Up Yaw Left Yaw Right Roll Left Roll Right

Downward

Neutral

Downward

Neutral

Upward Neutral Neutral Upwards

Upward Upward Neutral Upwards

Neutral Downward Neutral Neutral

Upward Neutral Neutral

Upward Neutral Upward

Neutral Neutral Downward Downwards Neutral

Downwards Neutral Upwards

Downwards Neutral

Downwards Upwards

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.

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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.
D = K P 4 Power

(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.
D = 1.7 × 4 0.5 = 1.43 ft

To check the blades tips remain below the maximum suggested helical velocity, the following relationship should be satisfied (Raymer 2006).

(πnD) 2 +V 2 < 950 fps

(Equation 3.11)

(π (1740rpm) × 1.43 ft ) 2 + (114.77 ft / sec) 2 = 173.6 < 950 fps

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.

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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 Maximum UAV Weight Length Catapult Weight Power Maximum Launch Velocity Price 20kg 4m 25kg Batteries 23m/s $15000 USD

BAE Pusher Prop Launcher2 36kg 3.5m 123kg Batteries 20m/s Not Available

RUAG Aerospace3 320kg 14m 3500kg Electrical Power Generator 66kts 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)

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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

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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

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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 Digital System4

Multiband Commtact Data Link (Transmitters) CRX (Receivers)5 150 km

CTX Commtact Mini Link and Data Systems5

Range Dimensions

100 km

15 km

220mm x 220mm x 2 x (80mm x 84mm x 124 mm x 94 mm x 150mm 48mm) 0.79 kg 11 – 32 V (DC) 17 mm 0.3 kg 9 to 60 V (DC)

Weight Input Voltage

2.1 kg 18 – 36 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 fuelconsumption 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 0.028 kg 40,000 ft 4.2 – 26 V

54 mm x 35 mm x 12 mm 0.00167 kg n/a 6.0 – 24 V

Weight Altitude Limit Input Voltage

0.2 kg 20,000 ft 7 – 36 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 setup) 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 (Ocean Controls)
9

Servo S9001 Servo Aircraft S3101 Coreless BB (Futaba) x 40mm 36mm 0.048kg 5.2kg/cm 6.0V 0.017kg x 20mm
10

Micro
11

Servo

(Active Robots)

Dimensions

35mm 32mm

x

16.9mm

x 28mm x 13mm x 30mm

Weight Torque Voltage

0.026kg 3.4kg/cm 4.8V

2.5kg/cm 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 MicroPilot

Dayview

Cloud Cap TASE (TASE Small Gyro Stabilized Camera Gimbal n.d.)13

Unit (BTC-88 – Micro (MP-DAYVIEWPTZ/MPPan / Tilt Unit 2008)12 NIGHTVIEWPTZ Stabilized Payload

Cameras 2006) Weight Dimensions Voltage Power Consumption Resolution Optical Zoom Retractable 752 x 582 pixels 10x Yes doors)
12 - BTC-88 Gimbal 2009, Procerus Technologies 13 - FCB-H11 Brochure 2008, Sony Corporation

0.465 kg 136 x 89 x 124 mm 6 to 12 V 2.1 W

less than 0.9 kg Ø110 x 210 mm 12 V (nominal) 3W

0.9 kg 127 x 112 x 178 mm 9 to 20 V 10 W typical, up to 18 W

320 x 240 pixels 25x

752 x 582 to 1080i HD Dependent on camera Additional Assembly

(Beetle-Wing No

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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 FCBIX11AP (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 Autopilot Transmitter Receiver Servo Motors (x4) Sensor (i.e. Camera) Miscellaneous Total Total Amps for Entire Mission

Maximum Current (Ampere) 0.5 3.2 0.7 0.4 0.2 1.0 6.0 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

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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 Model

Lithium – Ion14 MR-624 (Electric Fuel)

Zinc – Air15 FPEVO25-50003S Power) (Flight

Weight Dimensions Capacity

0.25kg 62.5mm x 38 mm x 74mm 2.2Amp-hours

0.404kg 27mm x 47mm x 74mm 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

Density

Source

Strength (MPa) (kg/m3) 7-7001 2,700 AAD, 2007 and Metal Prices, 2010

Metal

2.0909

Composite 1.47-2.95 Plastic 2.21-2.95

2000 45

2,550 1,050

Azom, 2010 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 Spars

Material

Manufacturing process

Aluminium

Wrought into the I-beam shape using rolling, extrusion and drawing and hardened through heat treatment. Wrought into designed curved shape using rolling, extrusion and drawing and hardened through heat treatment. Wrought and heat treated, then cut into specified airfoil shape using CNC machining. See remained of section

Frames (housing)

Aluminium

Ribs

Aluminium

Skin

Fibre-reinforced composite (e.g. fibreglass)

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).

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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 preimpregnated 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 fibrereinforced 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 Propulsions System Avionics Structural Weight Total Weight

Weight (kg) 5.0 2.5 3.0 10.5

Method Used to Obtain Weight of Component Based on battery and propeller selection Based on battery and avionics selection Based on statistics (Wstructure=0.4WTO)

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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.

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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-

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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.

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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 leadingedge sweep angle, Λ. (Equation 4.3)

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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°.

Longitudinal Stability
40
Static Margin (%MAC)

30 20 10 0 -10 10 12 14 16 18 20 Λ(°) 22 24 26 28 30

Figure 4.3: Longitudinal Static Margin versus Leading Edge Sweep Angle

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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)

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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 coefficient. 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
1.6 1.4 Section Lift Co-efficient 1.2 1 0.8 0.6 0.4 0.2 0 0 0.5 1 1.5 2 2.5 3 3.5 Distance along wing (ft)

Figure 5.1: Span-wise lift distribution along the wing

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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.

S wet = 782852.6mm 2

S ref = 359390.0mm 2

The ratio of Swet to Sref is therefore

S wet = 2.178281 S ref

For ease of calculations, an Swet/Sref of 2.2 will be used.

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The wetted aspect ratio is then 4.55 by equation 5.3. Wetted Aspect Ratio =
A S wet S ref = 10 = 4.55 2.2

(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)

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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.

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Table 6.1: Comparison of weight, area and loading values

Parameter WTO (lbs) S (ft2) Swet/Sref W/S (lbs/ft2) W/P (lbs/hp)

Design value 16.98 3.875 2 4.263 40.3

Final value 23.15 3.98 2.2 5.97 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.

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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 B C D (lbs) WTO (lbs)

Calculated value 0.97609 1 5.512 23.15 4.56 lbs

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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 New power index equation New cruise speed (ft/s) New stall speed (ft/s)

0.513

125 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.

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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 Weight Cruise speed Cruise altitude Flight range

Value 23.15 lbs 125 ft/sec 1500 ft 415 km

Radius of flight path Rate of climb Takeoff/landing requirements

11 km 300 fpm -

Flying wing design

-

Engine Fuel type

Baldor premium efficiency AC motor, 0.5 hp Electric

Compliance Increased from design value but still acceptable. Engine capable of achieving this speed. As per technical task. Increased from design value; batteries capable of supplying enough power. Increased from design value. As per sizing section. Catapult launch and net landing, as per technical task and conceptual design. Maintained with blendedwing-body design, seen in drawings. Provides required power to fly UAV. Provided by six zinc-air batteries.

105

106

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

107

108

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118

Appendix A – Raw Data for Statistical Analysis
Weight Max TakeOff Weight (kg) 170 12 204.12 150 19 1.1 175 5.5 Empty Weight (kg) 200.6 18.1 276 115 11.5 1.1 220 5.5

Manufacturer AAI Corporation AAI Corporation Pioneer UAV ATE Aerovision BlueBird Aerosystems BAE Systems BAE Systems Length

Model RQ-7B Shadow 200 Aerosonde 4 RQ-2B Pioneer Vulture Fulmar Micro B Phoenix Coyote

log(Wto) 2.230448921 1.079181246 2.30988556 2.176091259 1.278753601 0.041392685 2.243038049 0.740362689

log(We) 2.302330929 1.257678575 2.440909082 2.06069784 1.06069784 0.041392685 2.342422681 0.740362689

Manufacturer AAI Corporation AAI Corporation Pioneer UAV ATE Aerovision BlueBird Aerosystems BAE Systems BAE Systems Wing Span

Model RQ-7B Shadow 200 Aerosonde 4 RQ-2B Pioneer Vulture Fulmar Micro B Phoenix Coyote

Max TakeOff Weight (kg) 170 12 204.12 150 19 1.1 175 5.5

log(Wto) 2.230448921 1.079181246 2.30988556 2.176091259 1.278753601 0.041392685 2.243038049 0.740362689

Length (m) 3.41 2.1 4.27 3.1 1.23 0.7 3.8 0.9

log(length) 0.532754379 0.322219295 0.630427875 0.491361694 0.089905111 -0.15490196 0.579783597 0.045757491

Manufacturer AAI Corporation

Model RQ-7B Shadow 200

Max TakeOff Weight (kg) 170

log(Wto) 2.230448921

Span (m) 4.25

log(span) 0.62838893

119

AAI Corporation Pioneer UAV ATE Aerovision BlueBird Aerosystems BAE Systems BAE Systems Cruise Speed

Aerosonde 4 RQ-2B Pioneer Vulture Fulmar Micro B Phoenix Coyote

12 204.12 150 19 1.1 175 5.5

1.079181246 2.30988556 2.176091259 1.278753601 0.041392685 2.243038049 0.740362689

2.9 5.15 5.2 3.1 0.95 5.5 1.75

0.462397998 0.711807229 0.716003344 0.491361694 0.022276395 0.740362689 0.243038049

Manufacturer AAI Corporation AAI Corporation Pioneer UAV ATE Aerovision BlueBird Aerosystems BAE Systems BAE Systems Endurance

Model RQ-7B Shadow 200 Aerosonde 4 RQ-2B Pioneer Vulture Fulmar Micro B Phoenix Coyote

Max TakeOff Weight (kg) 170 12 204.12 150 19 1.1 175 5.5

log(Wto) 2.230448921 1.079181246 2.30988556 2.176091259 1.278753601 0.041392685 2.243038049 0.740362689

Cruise Speed (kts) 90 50 65 65 54 46 85 50

log(cruise speed) 1.954242509 1.698970004 1.812913357 1.812913357 1.73239376 1.662757832 1.929418926 1.698970004

Manufacturer AAI Corporation AAI Corporation Pioneer UAV ATE Aerovision BlueBird Aerosystems BAE Systems BAE Systems

Model RQ-7B Shadow 200 Aerosonde 4 RQ-2B Pioneer Vulture Fulmar Micro B Phoenix Coyote

Max TakeOff Weight (kg) 170 12 204.12 150 19 1.1 175 5.5

log(Wto) 2.230448921 1.079181246 2.30988556 2.176091259 1.278753601 0.041392685 2.243038049 0.740362689

Endurance (hrs) 6 30 5 3.5 8 1 4 1.5

log(enduranc e) 0.77815125 1.477121255 0.698970004 0.544068044 0.903089987 0 0.602059991 0.176091259

120

Appendix B – VTC Charts for various Australian Cities
Melbourne VTC

121

Sydney VTC

122

Canberra VTC

123

Hobart VTC

124

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 e 5 0.7 6 0.75 7 0.8 8 0.85 9 10 pi*A*e e = 0.7 10.99557 13.19469 15.3938 17.59292 19.79203 21.99115

C_fe S_wet/S_ref C_D0 A 0.006 2 0.012

e = 0.75 11.78097 14.13717 16.49336 18.84956 21.20575 23.56194

e = 0.8 12.56637 15.07964 17.59292 20.10619 22.61947 25.13274

e = 0.85 13.35177 16.02212 18.69248 21.36283 24.03318 26.70354

FAR 23.65 CGR SIZING A = 10 C_L 1.6 C_D 0.113859 L/D 14.05245 CGR CGRP 0.1 0.135315 (W/P)(W/S)^(1/2) 98.13368 Pmax/P_TO 89.21243

125

FAR 23.77 CGR SIZING A=10 C_L 1.6 C_D 0.113859 L/D 14.05245 CGR CGRP (W/P)(W/S)^(1/2) 0.05 0.095787 138.6306

FAR 23.65 RoC SIZING RC (fpm) RCP η_p C_L^(3/2)/C_D max

TO flaps, no LG 300 (33000)^-1 x RC 0.7 1.345(Ae)^(3/4)/C_Do^(1/4) 19.33056 (W/S)^(1/2) / 0.009091

RCP =

0.7 /(W/P) -

367.2807

CRUISE SPEED SIZING σ (1500 ft) Speed (ft/s) Speed (mph) I_p (Roskam p.163) (I_p)^3 * sigma 1/(I_p)^3*sigma STALL SPEED SIZING Raymer density V (km/h) V (fps) C_L max (design) W/S

0.956878418 115 78.40907 0.48 0.105823098 9.449732796

0.002377 60 54.678 1.4

1.3

1.2 4.26388687

4.974534682 4.619211

126

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.

127

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 y/(b/2) y(ft) L_b L_a c c_lb c_la1 Total 1.2 0 0 -0.321 1.355 0.85 0.23820 1.00548 1.44478

0.2 0.631 -0.225 1.265 0.748 0.18973 1.06670 1.46977

0.4 1.262 -0.017 1.132 0.646 0.01660 1.10527 1.34292

0.6 1.893 0.131 0.961 0.544 -0.15189 1.11424 1.18520

0.8 2.524 0.195 0.748 0.442 -0.27827 1.06741 1.00263

0.9 2.8395 0.197 0.59 0.391 -0.3177 0.95176 0.82432

0.95 2.99725 0.162 0.46 0.3655 -0.2795 0.79382 0.67302

0.975 3.07612 0.115 0.343 0.35275 -0.2056 0.61331 0.53034

128

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

129