FIELD OF INVENTION

The present invention relates to an unmanned aerial vehicle having an on-board hydrogen generator and a method of manufacturing the unmanned aerial vehicle by producing a main body of the unmanned aerial vehicle. In particular, the present invention provides a fixed wing vertical take-off and landing, VTOL unmanned aerial vehicle having on demand supply of hydrogen by an on board hydrogen generator.

BACKGROUND ART

Unmanned Aerial Vehicles, UAV, also commonly known as drones is an aircraft without a human pilot on board. There are numerous applications for UAVs including military aerospace applications, commercial use, monitoring and recovery. Examples of UAV application include hazardous material recovery, traffic monitoring and disaster relief support. Further UAVs allow remote imaging which can be useful in infrastructure condition evaluation.

Due to its mobility and compact size, UAVs can be used in military applications for recognition, environmental observation particularly in farming, maritime surveillance and mine removal activities and non military applications. UAVs are able to assist in environmental surveillance in a highly effective manner as compared to a helicopter which although manoeuvrable, are not always convenient or safe to fly in various terrains.

Throughout the years, UAVs have evolved. While drones and UAVs can be seen as a modern vehicle due to accessibility in owning a drone today, UAVs have been around since the 1800s where the first military use of UAVs, first aerial photograph and first camera on a UAV were first introduced. Since then, many prototypes of UAVs are being constantly invented to become more effective in this much needed and competitive industry.

Conventionally, there are four main types of drones or UAVs that common in the drone industry. The types of drones are multi-rotor drones, fixed wing drones, single rotor helipcopter drones and fixed wing hybrid VTOL drones. A multi-rotor drone comprises fans above their compact body and comprises multiple propellers to provide lift. The multi-rotor drones are usually smaller in size and provides for ease of control. A fixed wing drone requires a runway or catapult for launching as they do not have vertical take off ability. However, the presence of wings in these types of drones keeps the drones in the air and improves aerodynamics. A single rotor helicopter drone is generally a combination of multi-rotor device and large single wing units that are able to carry larger payloads. A fixed wing hybrid VTOL drone is a combination of fixed wing drones with vertical take-off ability of a rotor-based device.

Some major drawbacks in the current technology of unmanned aerial vehicles is in flight time which can be short and its payload capacity or maximum take-off weight (MTOW) which can be small.

Numerous drones or unmanned aerial vehicle have been invented to improve on the basic features of the presently available types of drones or unmanned aerial vehicles. The use of UAV having a configuration of VTOL and fixed wing is promising having the advantage of both VTOL type drones and fixed wing type drones. There have been many drones that improve on the foundation of a VTOL and fixed wing configuration type drone.

One example of a VTOL and fixed wing configuration type drone is disclosed in China Patent Application No. CN 206615386 U, hereinafter referred to as CN 386 U entitled “VTOL fixed wing UAVs screw pitch-changing mechanism and unmanned aerial vehicle’ having a filing date of 8 March 2017, Applicant: Shenzhen Smart Drone UAV Co Ltd. CN 386 U disclose a UAV having a configuration of VTOL and fixed wing. The configuration as disclosed in CN 386 U provides for long endurance and fast speed in flight mode. CN 386 U discloses having two motors at the bottom of the propellers to drive the rotation of the propellers. Besides that, CN 386 U discloses propellers, oar folder, oar folder connecting rods, guide holders, centre mount, pitch change link and back moving spring. The propellers form the main part of the UAV in driving the UAV having VTOL and fixed wing configuration.

Another example of a VTOL and fixed wing configuration type drone is disclosed in United States of America Patent No. US 10472064 B2, hereinafter referred to as US 064 B2 entitled “VTOL fixed wing aerial drone with interchangeable cabins’ having a filing date of 10 April 2018, Applicant: Yu Tian. US 064 B2 an aerial drone with a canard configuration having a left and right canard wing in addition to its main wings fixed on the main body. US 064 B2 further discloses using VTOL having detachable cabins where the cabins have the capacity to hold a passenger or cargo. US 064 B2 discloses using a hybrid engine to produce electricity. Further, US 064 B2 discloses the use of batteries and energy storage units disposed within the cabin. The energy stored is electricity and the storage unit stores batteries which supplies and replenishes energy eliminating the need for the drone to stop to charge its batteries. The drone as disclosed in US 064 B2 can be performed autonomously or actively controlled.

A further example of a VTOL and fixed wing configuration type drone is disclosed in United States of America Patent No. US 3794273 A, hereinafter referred to as US 273 A, entitled “VTOL Rotor Wing Drone Aircraft” having a filing date of 25 September 1972, Applicant: Teledyne Ryan Aeronautical. US 273 A disclose a drone aircraft that has VTOL capability having rotary and fixed wing configuration. US 273 A focus on the ability to provide an aircraft having fixed wings in cruising flight while also having the capability of rotating for vertical take-off and landing where portions of the wings are movable for lift and directional control. The drone aircraft comprises a fuselage for containing equipment and payload and wings. US 273 A disclose a turbojet engine that provides cruising thrust and wing rotating thrust. US 273 A further disclose a wing mounted above the fuselage having three arms and a beam that fits to the underside of the wing that minimizes airflow disturbance.

As outlined above, various drones have been developed having a VTOL and fixed wing type configuration. However, none of the prior arts disclose an unmanned aerial vehicle having an on board hydrogen generator for on demand supply of hydrogen by hydrolysis A hydrogen fuel cell uses hydrogen and air to generate electricity through an electrochemical process. The application of using hydrogen fuel cell enhances flight endurance in powering of the unmanned aerial vehicle. SUMMARY OF INVENTION

The present invention relates to an unmanned aerial vehicle having an on-board hydrogen generator. In particular, the present invention provides a fixed wing vertical take-off and landing, VTOL unmanned aerial vehicle having on demand supply of hydrogen by an on board hydrogen generator.

One aspect of the present invention provides an unmanned aerial vehicle (100) having a fixed wing vertical take-off and landing powered by on demand supply of hydrogen comprising a main body (102) for housing a plurality of components for driving the unmanned aerial vehicle (100) and to support a maximum take-off weight, a plurality of motors (106a, 106b, 106c) for transition of the unmanned vehicle (100) to a forward tilting motion, at least two wings (104) configured to a wing box (104a) of the main body (102) for lifting of the unmanned aerial vehicle (100) where optimum dimensions relating to lift are determined characterized in that the main body comprises at least one autopilot controller for directing and guiding the unmanned aerial vehicle for take-off and landing, an on board hydrogen generator (107e) for supplying hydrogen by hydrolysis for generation of electricity to drive the unmanned aerial vehicle, a cartridge (107a) having Proton-Exchange Membrane Fuel cells, PEMFC, for receiving hydrogen produced by the on board hydrogen generator for energy generation to drive the unmanned aerial vehicle, a plurality of temperature and pressure sensors (107b) for detecting temperature and pressure of the unmanned aerial vehicle, a plurality of batteries (107c) for generation of power when PEM fuel cells reach a power level of at least 200W and below, a fuel cell controller (107d) for detecting levels of the PEM fuel cells wherein a power level of at least 100W will allow for continued usage of the on board hydrogen generator (107e) and a power level of at least 200W will direct energy production using the plurality of batteries (107c);

Another aspect of the present invention provides an unmanned aerial vehicle (100) wherein the plurality of motors (106a, 106b, 106c) are preferably at least three motors.

Yet another aspect of the present invention provides an unmanned aerial vehicle (100) wherein the plurality of motors (106a, 106b, 106c) are positioned preferably at each wing (104) and at a back position of the main body (102). Another aspect of the present invention provides an unmanned aerial vehicle (100) wherein the main body (102) for housing a plurality of components for driving the unmanned aerial vehicle and to support a maximum take-off weight comprises a weight of at least 6.5 kg.

Yet another aspect of the present invention provides an unmanned aerial vehicle (100) of claim 1 , wherein the at least two wings (104) configured to the wing box (104a) of the main body (102) for lifting of the unmanned aerial vehicle where optimum dimensions relating to lift are determined further comprises an optimum dimension of a lift and drag, CL/CD value of 8.94 at an angle of attack of 0°.

Another aspect of the present invention provides an unmanned aerial vehicle (100) wherein the main body (102) for housing a plurality of components for driving the unmanned aerial vehicle and to support a maximum take-off weight comprises a length of at least 2.86 m.

Yet another aspect of the present invention provides an unmanned aerial vehicle (100) wherein the at least two wings (104) configured to the wing box (104a) of the main body (102) for lifting of the unmanned aerial vehicle where optimum dimensions relating to lift are determined further comprises an optimum dimension of a mean aerodynamic chord to determine aerodynamic lift of the at least two is at least 40.5 cm.

Another aspect of the present invention provides an unmanned aerial vehicle (100) wherein the main body (102) for housing a plurality of components for driving the unmanned aerial vehicle and to support a maximum take-off weight is made from carbon fibre composite.

Yet another aspect of the present invention provides an unmanned aerial vehicle (100) wherein the at least two wings (104) configured to the wing box (104a) of the main body (102) for lifting of the unmanned aerial vehicle where optimum dimensions relating to lift are determined is coated with carbon fiber.

Another aspect of the present invention provides an unmanned aerial vehicle (100) wherein the at least two wings (104) configured to the wing box (104a) of the main body (102) for lifting of the unmanned aerial vehicle where optimum dimensions relating to lift are determined, wherein the wing box comprises a front spar (104b), a port keel beam (104d), a starboard keel beam (104c) and a rear spar (104e).

Yet another aspect of the present invention provides a method for manufacturing an unmanned aerial vehicle having a fixed wing vertical take-off and landing powered by on demand supply of hydrogen involves producing a main body of the unmanned aerial vehicle, the method for producing a main body of the , VTOL unmanned aerial vehicle comprising steps of designing a 3-D mould by using a computer aided design, CAD (3102); printing a 3-D mould comprising a plurality of pieces that make up a bottom and a top part of the main body (3104), joining the plurality of pieces by a plurality of steel rods (3106), compressing the plurality of pieces by using pressure sensitive tape (3108); applying glue for covering cracks between the plurality of pieces (3110), applying epoxy steel to seal cracks between the plurality of pieces (3112), applying poly putty to smoothen imperfections on a surface of the main body (3114), sanding the 3-D mould with 200 Cw sandpaper (3116), sanding the 3-D mould with 600 Cw and 800 Cw sandpaper (3118), proceeding with carbon fibre composite process (3120), shaving of an outer perimeter of the main body (3122), spraying the main body with paint (3124). Further, proceeding with carbon fibre composite process comprises steps of waxing the 3-D mould which acts as a release agent to peel off cured ply layers (31202), cutting the cured ply layers to a predetermined size (31204), positioning the cured ply layers in the 3-D mould (31206), adding a peel ply layer to the 3-D mould (31208), adding an infusion mesh layer (31210), positioning a resin feed spiral directly over the infused mesh (31212), applying a vacuum bag taping (31214), positioning and taping the vacuum bag (31216), connecting and sealing a resin feed hose (31218), clamping the resin feed hose (31220) and switching on a vacuum pump to test the vacuum (31222).

Another aspect of the present invention provides a method wherein designing a 3-D mould by using a computer aided design, CAD (3102) comprises steps of sketching a main body design (31022), editing of a generative design excel file to design a wing part of the unmanned aerial vehicle (31024) and using computer aided engineering of predicting aerodynamics and structural dynamics unmanned aerial vehicle (31026).

The present invention consists of features and a combination of parts hereinafter fully described and illustrated in the accompanying drawings, it being understood that various changes in the details may be made without departing from the scope of the invention or sacrificing any of the advantages of the present invention.

BRIEF DESCRIPTION OF ACCOMPANYING DRAWINGS

To further clarify various aspects of some embodiments of the present invention, a more particular description of the invention will be rendered by references to specific embodiments thereof, which are illustrated in the appended drawings. It is appreciated that these drawings depict only typical embodiments of the invention and are therefore not to be considered limiting of its scope. The invention will be described and explained with additional specificity and detail through the accompanying drawings in which:

Figure 1 .Oa illustrates an unmanned aerial vehicle of the present invention.

Figure 1.1 a illustrates a side view of the unmanned aerial vehicle of the present invention.

Figure 1 .1 b illustrates a top view of the unmanned aerial vehicle of the present invention.

Figure 1.1c illustrates a front view of the unmanned aerial vehicle of the present invention.

Figure 2.0 illustrates a general overview of the components of the unmanned aerial vehicle of the present invention.

Figure 3.0a illustrates a top mould fuselage of the unmanned aerial vehicle of the present invention.

Figure 3.0b illustrates a bottom mould of the unmanned aerial vehicle of the present invention.

Figure 3.0c is a table showing the final fuselage UAV configuration dimensions of the unmanned aerial vehicle of the present invention.

Figure 3.1 a shows the application of glue to the mould cracks for the fuselage of the unmanned aerial vehicle of the present invention. Figure 3.1b shows the application of epoxy steel to the mould cracks for the fuselage of the unmanned aerial vehicle of the present invention.

Figure 3.1c shows dried polly putty on the mould for the fuselage of the unmanned aerial vehicle of the present invention.

Figure 3.1 d shows the shaving of the fuselage from the mould of the unmanned aerial vehicle of the present invention.

Figure 3.1 e shows the fuselage treated with poly putty of the unmanned aerial vehicle of the present invention.

Figure 3.1 f shows the white painted top fuselage of the unmanned aerial vehicle of the present invention.

Figure 4.0a illustrates an internal skeleton of the fuselage of the unmanned aerial vehicle of the present invention.

Figure 4.0b illustrates a wing box of the unmanned aerial vehicle of the present invention.

Figure 4.0c illustrates an isometric view of the wing box of the unmanned aerial vehicle of the present invention.

Figure 4.0d illustrates a top view of the wing box of the unmanned aerial vehicle of the present invention.

Figure 4.0e is a table showing wing configuration of the unmanned aerial vehicle of the present invention.

Figure 4.0f is a table showing a comparison of criterias for wings of an unmanned aerial vehicle of different carbon composite.

Figure 5.0 is a table showing components for autopilot and its respective software used in the unmanned aerial vehicle of the present invention. Figure 6.0a illustrates a tilt mechanism of the motors of the unmanned aerial vehicle when in VTOL mode.

Figure 6.0b illustrates a tilt mechanism of the motors of the unmanned aerial vehicle when in forward flight mode.

Figure 7.0 shows an internal configuration of the components within the fuselage of the unmanned aerial vehicle.

Figure 8.0a is a graph showing a standard flight envelope criteria outlined by the Federal Aviation Administration Airworthiness Regulation.

Figure 8.0b is a table showing the parameters for flight envelope calculation of the present invention.

Figure 8.0c is a table showing the parameters for creating the gust lines in the flight envelope.

Figure 8.0d is a graph illustrating the flight envelope detailing the aerostructural limits for the present invention.

Figure 9.0a is a first model design of the unmanned aerial vehicle of the present invention.

Figure 9.0b is a final model design of the unmanned aerial vehicle of the present invention.

Figure 9.0c illustrates a top, side and front view of a Power Ray.

Figure 9.0d is a flowchart showing a method employed for the present invention in designing the fuselage.

Figure 10.0a illustrates force balance on a typical airplane with an empennage.

Figure 10.0b illustrates a reflexed airfoil of the wings. Figure 10.0c is a graph showing a HS522 reflexed airfoil loaded into the analysis tool XFLR5.

Figure 10.0d illustrates the parameter input for obtaining mean aerodynamic length of the present invention.

Figure 10.1a illustrates a tri-copter configuration.

Figure 10.2a illustrates an actuator disc theory force diagram for climb.

Figure 10.2b is a table showing the input and output parameters to determine induced power for climb.

Figure 10.2c illustrates an actuator disc theory force diagram for descent.

Figure 10.2d is a table showing the input and output parameters to determine induced power for descent.

Figure 10.2e is a graph of induced power against diameter of a propeller.

Figure 10.2f illustrates a force-balance diagram for a typical aircraft in flight.

Figure 11 .0a illustrates a computer aided design to produce a 3-D model similar to the Power Ray showing the step of selecting a sketch tracer from the menu.

Figure 11.0b illustrates a computer aided design to produce a 3-D model similar to the Power Ray showing the step of selecting ‘shading with material’ icon.

Figure 11 .0c illustrates a computer aided design to produce a 3-D model similar to the Power Ray showing the step of selecting ‘create an immersive sketch’ icon.

Figure 11. Od illustrates a computer aided design to produce a 3-D model similar to the Power Ray showing the step of adjusting the size of the image. Figure 11 .Oe illustrates a computer aided design to produce a 3-D model similar to the Power Ray showing the step of repeating previous steps to obtain all views.

Figure 11. Of illustrates a computer aided design to produce a 3-D model similar to the Power Ray showing the step of selecting ‘new part’ for designing the present invention.

Figure 11.0g illustrates a computer aided design to produce a 3-D model similar to the Power Ray showing the step of drawing a 3D curve by selecting ‘3D curve’.

Figure 11 .Oh illustrates a computer aided design to produce a 3-D model similar to the Power Ray showing the step of drawing a 3D curve by selecting ‘lock privileged plane orientation parallel to screen.

Figure 11. Oi illustrates a computer aided design to produce a 3-D model similar to the Power Ray showing the step of drawing a 3D curve by using sketch tracing.

Figure 11. Oj illustrates a computer aided design to produce a 3-D model similar to the Power Ray showing the step of selecting ‘Multi-section Surface’ and ‘Fill’ tools.

Figure 11 .Ok illustrates a computer aided design to produce a 3-D model similar to the Power Ray showing the step of closing a surface by selecting ‘Join’ or ‘Healing’ tools.

Figure 11.01 illustrates a computer aided design to produce a 3-D model similar to the Power Ray showing the result of closing a surface.

Figure 11 .0m illustrates a computer aided design to produce a 3-D model similar to the Power Ray showing the result of closing a surface by selecting ‘Split.2’.

Figure 11 .On illustrates a computer aided design to produce a 3-D model similar to the Power Ray showing the result of closing a surface by selecting ‘Healing. T.

Figure 11. Oo illustrates a comparison of the computer aided design of a 3-D model similar to the Power Ray and final Iteration 0 design.

Figure 11 .1a illustrates a browser to the indicated file path in excel for designing the wing. Figure 11.1 b illustrates editing an excel sheet for designing the wing by deleting the cells.

Figure 1 1 .1 c illustrates editing an excel sheet for designing the wing by copying the cells.

Figure 1 1.1d illustrates editing an excel sheet for designing the wing continuously.

Figure 1 1.1 e illustrates an excel sheet of importing the airfoil parameters from the .dat file for designing the wing.

Figure 11.1f illustrates an excel sheet showing replacing the first cell with the intended root cord for designing the wing.

Figure 11.1 g illustrates an excel sheet showing editing the last row of the data set for designing the wing.

Figure 11.1h illustrates selecting ‘options’ under ‘file’ tab for preparing the file to be imported.

Figure 11.1 i illustrates selecting the ‘Data’ box for file to be imported.

Figure 11.1j illustrates access to the ‘Developer’ tab in the Excel sheet.

Figure 1 1.1 k illustrates HS522 airfoil uploaded into CATIA v5 via Excel spreadsheet.

Figure 1 1.11 illustrates two identical airfoil profiles with different chord lengths.

Figure 1 1.1 m illustrates connecting the two identical airfoil profiles using 3D curves.

Figure 1 1 .1 n Illustrates the 3D curve joining the trailing edge (TE) is horizontal.

Figure 1 1 .1 o illustrates creation of a multi section surface in designing the wing.

Figure 1 1 .2a illustrates a data sheet for motor U7 V2.0 KV280 T -Motor. Figure 11 .2b illustrates a data sheet for motor U7 V2.0 KV420 T-Motor.

Figure 11 .2c illustrates a data sheet for motor U7 V2.0 KV490 T -Motor.

Figure 11 .3a is a graph showing lift coefficient against drag coefficient.

Figure 11.3b is a graph showing lift coefficient against angle of attack and a moment coefficient against the angle of attack.

Figure 11 .3c is a graph showing life coefficient against top transition point and lift drag ratio against angle of attack.

Figure 11 .3d illustrates batch analysis with XFLR5.

Figure 11 .4a illustrates fluid flow around a cylinder for validating experimental results.

Figure 11 .4b is a graph of the drag coefficient against Reynolds number.

Figure 11 .4c is a table showing the difference for a small range of Angle of Attack values.

Figure 11.5a illustrates the first step in running Computational fluid dynamics simulation by dragging ‘Fluid flow (Fluent)’ into working environment.

Figure 11.5b illustrates importing the external CAD geometry in running Computational fluid dynamics.

Figure 11 .5c illustrates CAD model imported in ZX plane in preparing the Geometry of the unmanned aerial vehicle.

Figure 11.5d illustrates CAD model imported in preparing the Geometry of the unmanned aerial vehicle showing the splitting of fuselage into two parts. Figure 11.5e illustrates CAD model imported in preparing the Geometry of the unmanned aerial vehicle showing creating the fluid zone.

Figure 1 1.5f illustrates CAD model imported in preparing the Geometry of the unmanned aerial vehicle showing extension of the fluid zone.

Figure 11.5g illustrates CAD model imported in preparing the Geometry of the unmanned aerial vehicle showing subtracting the solid from the fluid zone.

Figure 1 1 .5h illustrates creating named selection for the fluid zone surfaces.

Figure 1 1 .5i illustrates Y+ values and related wall functions.

Figure 1 1.5j illustrates wall spacing value generated as output using a desired y+ value and initial conditions as input.

Figure 1 1 .5k illustrates inflation layers used to capture the turbulent boundary layers.

Figure 11 .51 is a graph showing convergence to a single value with increasing number of elements.

Figure 1 1 .5m illustrates a model set up by selecting a viscous model for the solver.

Figure 1 1 .5n illustrates computational selection to change outlet type.

Figure 1 1 .5o illustrates setting up of initial conditions at the winglet for 0° AOA.

Figure 1 1 .5p illustrates setting reference values.

Figure 1 1 .5q illustrates setting up drag monitor.

Figure 1 1 .5r illustrates initializing the solution.

Figure 1 1 .5s illustrates calculating the iterative solution. Figure 11.6a illustrates side views of the evolution of design of the fuselage of an unmanned aerial vehicle.

Figure 11 .6b is a graph showing lift to drag ratio against iteration progression.

Figure 11 .6c is a table showing winglet analysis.

Figure 11 .6d illustrates a final winglet design.

Figure 11 .6e illustrates lift drag values for different angles of attack.

Figure 11 .6f is a table showing the load cases of the unmanned aerial vehicle.

Figure 11.6g illustrates load cases on the centre wingbox for fixed wing flight cruise and fixed wing acceleration.

Figure 11 .6h illustrates load cases on the centre wingbox for take-off.

Figure 11 .6i illustrates load cases on the centre wingbox for VTOL transition at different angles.

Figure 11 .6j illustrates load cases on the centre wingbox for fixed wing banking.

Figure 11 .6k illustrates load cases on the centre wingbox for gust loads.

Figure 11.61 illustrates load cases on the tail section for steady cruise and fixed wing acceleration.

Figure 11 .6m illustrates load cases on the tail section for transition flight.

Figure 11 .6n illustrates load cases on the tail section for take-off.

Figure 11 .60 illustrates load cases on landing gears for static ground.

Figure 11 .6p illustrates load cases on the landing gears for landing. Figure 11 .7a illustrates an assembly design showing equidistant placement of the motors.

Figure 11 .7b illustrates shows the offset from the fuselage surface.

DETAILED DESCRIPTION OF THE DRAWINGS

The present invention relates to an unmanned aerial vehicle having an on-board hydrogen generator. In particular, the present invention provides a fixed wing vertical take-off and landing, VTOL unmanned aerial vehicle having on demand supply of hydrogen by an on board hydrogen generator. Hereinafter, this specification will describe the present invention according to the preferred embodiments. It is to be understood that limiting the description to the preferred embodiments of the invention is merely to facilitate discussion of the present invention and it is envisioned without departing from the scope of the appended claims.

The present invention provides a hydrogen fuel cell powered tricopter-fixed-wing unmanned aerial vehicle. The unmanned aerial vehicle of the present invention comprises vertical take-off and landing (VTOL) capabilities to a fixed wing configuration. The present invention provides that the powering of the unmanned aerial vehicle is by hydrogen fuel cell where hydrogen and air are used to generate electricity through an electrochemical process by an on board hydrogen generator. The unmanned aerial vehicle of the present invention uses carbon composite mould in creating the fuselage also known as the main body and the wings outer structure. The unmanned aerial vehicle of the present invention is equipped with three motors for VTOL.

The advantages of the unmanned aerial vehicle of the present invention where there is VTOL capabilities coupled with fixed wing configuration is that it eliminates the need for a catapult launch. Further, using an on board hydrogen generator allows an on demand supply of hydrogen which in turn enhances flight endurance. Carbon composite was chosen due to its high structural density. The positioning of the motors, two on either wing at the front and one at the backend of the fuselage is important to ensure a stable flight following the take-off as the drone must be as level to the horizon as possible.

Reference is first made to Figure 1.0a. Figure 1.0a illustrates an unmanned aerial vehicle of the present invention. The different views of the unmanned aerial vehicle are further illustrated in Figure 1.1a, Figure 1.1b and Figure 1.1c. Figure 1.1a illustrates a side view of the unmanned aerial vehicle, Figure 1.1 b illustrates a top view of the unmanned aerial vehicle and Figure 1.1c illustrates a front view of the unmanned aerial vehicle of the present invention. The present invention provides a fixed wing vertical take-off and landing, VTOL unmanned aerial vehicle (100) having on demand supply of hydrogen. The unmanned aerial vehicle comprises a main body (102) for housing a plurality of components for driving the unmanned aerial vehicle (100) and to support a maximum take-off weight. Further, the present invention comprises a plurality of motors (106a, 106b, 106c) for transition of the unmanned vehicle (100) to forward motion by tilting. The plurality of motors (106a, 106b, 106c) of the present invention is at least three motors and is positioned preferably at each wing (104) and at a back position of the main body (102). The present invention also includes at least two wings (104) configured to a wing box (104a) of the main body (102) further detailed in Figure 4.0a, Figure 4.0b, Figure 4.0c and Figure 4.0d, for lifting of the unmanned aerial vehicle (100) where optimum dimensions relating to lift are determined. The two wings (104) configured to the wing box (104a) of the main body (102) for lifting of the unmanned aerial vehicle where optimum dimensions relating to lift are determined further comprises the optimum dimension of a lift and drag, C _L /C _D value of 8.94 at an angle of attack of 0°. The at least two wings (104) configured to the wing box (104a) of the main body (102) comprises the optimum dimension of a mean aerodynamic chord to determine aerodynamic lift of the at least two wings (140) is at least 40.5 cm. The at least two wings (104) is coated with carbon fiber.

In describing further on the main body or fuselage, reference is made to Figure 7.0. Figure 7.0 shows an internal configuration of the components within the fuselage of the unmanned aerial vehicle. The main body or fuselage comprises an autopilot controller for directing and guiding the unmanned aerial vehicle for take-off and landing, an on board hydrogen generator (107e) for supplying hydrogen by hydrolysis for generation of electricity to drive the unmanned aerial vehicle, a cartridge (107a) having Proton- Exchange Membrane Fuel cells, PEMFC, for receiving hydrogen produced by the on board hydrogen generator for energy generation to drive the unmanned aerial vehicle, a plurality of temperature and pressure sensors (107b) for detecting temperature and pressure of the unmanned aerial vehicle, a plurality of batteries (107c) for generation of power when PEM fuel cells reach a power level of at least 200W and below, a fuel cell controller (107d) for detecting levels of the PEM fuel cells wherein a power level of at least 100W will allow for continued usage of the on board hydrogen generator (107e) and a power level of at least 200W will direct energy production using the plurality of batteries (107c). The main body (102) comprises a maximum take-off weight of at least 6.5 kg and is ideally at a length of at least 2.86 m. The main body (102) is preferably made from carbon fibre composite.

Reference is made to Figure 2.0 illustrating a general overview of the components of the unmanned aerial vehicle of the present invention. The main components of the present invention are the aerostructure, the avionics and the propulsion. The aerostructure is made of the fuselage and the wings. The avionics which is the electronic system applied to the unmanned aerial vehicle is the autopilot system. The propulsion or what drives the aircraft of the unmanned aerial vehicle includes a tilt mechanism.

Fuselage Mould

The main material of the fuselage is preferably carbon fibre composite due to its high structural density. Further, the manufacturing process maintains high fidelity to the original surface design of the fuselage since a mould is used. In describing further on the fuselage, reference is made to Figure 3.0a, Figure 3.0b and Figure 3.0c. Figure 3.0a illustrates a top mould fuselage of the unmanned aerial vehicle of the present invention. Figure 3.0b illustrates a bottom mould of the unmanned aerial vehicle of the present invention.

The method for manufacturing an unmanned aerial vehicle of a fixed wing vertical takeoff and landing powered by on demand supply of hydrogen involves producing a main body of the , VTOL unmanned aerial vehicle, The method for manufacturing the fuselage or main body is further described in Figure 3.1g. Prior to to producing a first prototype, a 3-D mould is designed using a computer aided design (3102), To produce the first prototype, a 3-D printed negative mould of the top and bottom part of the fuselage was prepared (3104). At least 24 different pieces for both the top and the bottom part of the fuselage mould was printed and joined (3106) together by 8mm having an 8.2mm tolerance steel rods that run along the length and width of the mould. A threaded steel rod allows clamping action at the ends to compress the spaces between the 3-D printed part and improve mould quality. The wings of the unmanned aerial vehicle are preferably prepared in the same way. Figure 3.0c is a table showing the final fuselage UAV configuration dimensions of the unmanned aerial vehicle of the present invention. The present invention also named MK- 3 is an improvement of a first model design MK-2. The width of the fuselage of Mk-3 is preferably 50.6cm and the length is 91 .41 cm. The maximum height of the fuselage is 16 cm. The wetted are is 0.843 m ² and the cross-sectional area is 0.0588 m ² . The mass of the fuselages is 7.5 kg. Mk-2, an earlier prototype of the present invention was built with a length of 76.6 cm and height of 16 cm.

Upon 3-D printing the first prototype, the fuselage is then joined together to form a single piece by mould treatment process. Pressure sensitive tapes such as masking tape is first used to compress (3108) the 3-D printed parts together. After that, glue is applied as shown in Figure 3.1 a to the cracks between parts (3110). The glue used is preferably BOSSIL 8502 Power glue. However, the type of glue is not limited and may include any adhesive that has the ability to hold the fuselage parts together. After the glue is applied, an epoxy steel of 4 minutes is used to completely seal off the cracks (31 12) as shown in Figure 3.1 b. The next step in mould treatment is to apply Hitary 98 Poly Putty to smoothen any imperfections caused by the 3-D printing process (31 14) as shown in Figure 3.1 c. Once the putty is dried, the moulds are sanded with 200Cw sandpaper (31 16) and subsequently 600Cw & 800Cw type sandpaper with water (31 18). The moulds are cleaned and the carbon fibre composite process can proceed (3120). Figure 3.1 h elaborates on the carbon fibre composite process. The first step is to wax the 3-D mould (31202) which acts as a release agent to peel off cured ply layer. The ply layers are cut to a predetermined size (31204) and positioned in the 3-D mould (31206). The peel ply layers (31208) and the infusion mesh layer (31210) is added to the 3-D mould. A resin feed spiral is positioned directly over the infusion mesh (31212). Then the vacuum bag taping is applied (31214) where the vacuum bag is positioned and taped down (31216). The resin feed hose is connected and sealed (31218) and is further clamped (31220) before switching of the vacuum pump for testing (31222).

Reference is further made to Figure 3.1 i which will be described in further detail in Figure 9.0a to Figure 11 .7b in designing a 3-D mould using a computer aided design. The main steps in designing the 3-D mould using a computer aided design is sketching a main body design, editing a generative deisgn excel file to design a wing part and in the application of computer aided engineering to predict the aerodynamics and structural dynamics of the unmanned aerial vehicle. The results of the process are shown in Figure 3.1 d and Figure 3.1 e. Figure 3.1 d shows the shaving of the fuselage from the mould and Figure 3.1 e shows the fuselage treated with poly putty to smoothen the surface of the fuselage. The fuselage is then sprayed with paint as shown in Figure 3.1 f.

Wing Box of the Fuselage

In describing further on the wing box (104a), reference is made to Figure 4.0a, Figure 4.0b, Figure 4.0c and Figure 4.0d. Figure 4.0a illustrates an internal skeleton of the fuselage of the unmanned aerial vehicle of the present invention. Figure 4.0b illustrates a wing box of the unmanned aerial vehicle of the present invention. Figure 4.0c illustrates an isometric view of the wing box of the unmanned aerial vehicle of the present invention. Figure 4.0d illustrates a top view of the wing box of the unmanned aerial vehicle of the present invention. The wing box comprises a front spar (104b), a port keel beam (104d), a starboard keel beam (104c) and a rear spar (104e).

Reference is made to Figure 4.0e. Figure 4.0e is a table showing wing configuration of the unmanned aerial vehicle of the present invention. As can be seen in the table, the parameters of the wing configuration include the airfoil, full span, root chord, tip chord, taper ratio, sweep angle, mass, centre of gravity an aerodynamic centre. The airfoil selected in the present invention is HS522 for both Mk-3 and its first previous prototype MK-2. For the Mk-3 the full span value is 2.87m, the root chord is 0.536 m, the top chord is 0.234 m, the taper ratio is 0.43, the sweep angle is 23.4°, the mass if 7.5 kg, centre of gravity is 0.455 m and aerodynamic centre is 0.504 m.

The wings of the present invention are reinforced by carbon composite. Certain criterias are shown in Figure 4. Of where a table on a comparison of criterias for wings of an unmanned aerial vehicle of different carbon composite is shown. The criterias that are vital include rigidity, weight, access, ease of manufacture, cost and appearance.

Avionics

The electronic system used in the unmanned aerial vehicle is an autopilot system. The autopilot components and software are shown in Figure 5.0. An autopilot test is performed by establishing waypoints which will determine flight path in the ground control software, GCS. The information will be relayed to the autopilot controller, which feeds the GCS its global positioning system, GPS, and other flight information in realtime. This open-loop control system ensures that the drone will be able to fly autonomously through the designated flight path and also recover, given the external factors that act on the drone. The flight simulator which is the X-Plane 10 v 10.51 is useful in determining the centre of gravity of the drone without any payload, after achieving steady flight.

Propulsion

The driving of the unmanned aerial vehicle is by tilt mechanism. The tilt mechanism allows the transition from vertical take-off and landing to forward flight and vice versa. Reference is made to Figure 6.0a and Figure 6.0b on the positioning of the motor in different modes. Figure 6.0a illustrates a tilt mechanism of the motors of the unmanned aerial vehicle when in VTOL mode. The motor will be in upright position when VTOL takes place and rotates across the x-axis when the forward flight takes place where the motor is perpendicular to the shaft. Figure 6.0b illustrates a tilt mechanism of the motors of the unmanned aerial vehicle when in forward flight mode. The motor will be parallel to the shaft and at the z-axis.

On Board Hydrogen Generation System

The process of hydrolysis within the on board hydrogen generator for driving and powering the unmanned aerial vehicle is described. The process begins in the Hydrolysis Reactor where sodium borohydride, NaBH ₄ , catalyst, de-ionized water, and sodium hydroxide, NaOH, react to form hydrogen. The hydrolysis reaction is an exothermic process which releases heat with temperature of 45 degrees Celsius whic is detected by Thermocouple. The byproduct of the reactions is sodium metaborate, NaBO2. NaBO2 will be stored in the by-product container. The hydrogen gas will then travel to the Gas Cooler via a spiral 6mm length plastic tube submerged in water for the cooling of the generated hydrogen gas. The hydrogen gas cools off from 45 degrees Celsius down to 25 degrees Celsius. The gas then flows into a Gas Dryer where the purity of hydrogen gas needs to be kept at 99.995% which is monitored by Hygrometer. The Gas Dryer contains Calcium Chloride, CaCI ₂ , as the drying agent. The hydrogen gas will then be released based on the fuel cell requirement. The flow rate of the hydrogen gas that feeds into a 200W PEMFC is at 2.7 L/min. If the flow rate falls below or rises above the threshold, a signal will be sent to Pressure Regulator to maintain the pressure of hydrogen gas to correspond with the assigned flow rate. Finally, the gas flow is controlled by the solenoid valve as per instructions from the fuel cell controller.

Vertical Take-off and Landing

The take-off mechanism and landing mechanism used in the present invention involves the servo actuator retractable landing gear system. In Take off mode the landing gear is in deployed mode and retraction will commence upon aircraft transition to forward flight. As for landing, upon transition from fixed wing mode to VTOL, the landing gear is deployed from stowing position.

Satellite Communication

There are three preferred satellites used to interface with broadband network. The satellite communications are IMARSAT, THURAYA and STARLINK. The link and VPN will be connected to the ground station to enhance range capability.

Flight Envelope

Aircrafts must meet performance and design specifications to be certified air-worthy. The specifications are in the form of a flight envelope outlined by the Federal Aviation Administration Airworthiness Regulation as shown in Figure 8.0a. Based on the standard flight envelope the parameters of the unmanned aerial vehicle’s flight envelope were calculated. Figure 8.0b is a table showing the parameters for flight envelope calculation of the present invention. One important parameter is gust line calculation for descent velocity and cruise velocity.

The gust lines are calculated using the FAR Part 25 Airworthiness Standards. where p = density of air at sea level (kgm' ³ )

V= aircraft speed (ms' ¹ ) a = 3-D lift curve slope.

K _g = alleviation factor

U _de = derived gust velocities (15.24 ms' ¹ at V _c , 7.62 ms' ¹ at V _D )

The gust alleviation factor, K _g can be calculated from

Where pg is in turn calculated from

The parameters for creating the gust lines in the flight envelope are calculated where the values are presented in a table as shown in Figure 8.0c. The parameters for creating gust lines in the flight envelope is the density, aircraft speed, lift-curve slope, Inertial Ratio, Gust alleviation factor, maximum and minimum lift coefficient, mass and planform area. The flight envelope that details the aerostructural limits of the present invention is illustrated in the graph of Figure 8.0d.

COMPUTER AIDED DESIGN, CAD REFERENCE

General 3-D design as reference

In accurately forming the dimensions in order to build the unmanned aerial vehicle of the present invention, computer aided designs are used. A 3-D scanned model of a first model, which is an earlier model of the present invention, is shown in Figure 9.0a and is used as a reference that represents the physical model of the unmanned aerial vehicle. Figure 9.0b is a final model design of the unmanned aerial vehicle of the present invention which is also known as the Mk-3 model.

Aerostructure Design as reference

In improving on the current design from the previous first model of the present invention different programme was used in the aerostructure disgn. The fuselage used in the first model consisted of the Skywalker X8 fuselage augmented with a canopy roof with air intakes. In order to better accommodate a hydrogen fuel cell power plant, a blended wing fuselage designed in house was employed. A study on the Power Ray developed by Power Vision in forming the base profile design for the fuselage in the improved model of the present invention, Mk-3 was used. Figure 9.0c illustrates a top, side and front view of a Power Ray. The Power Ray is a hydrodynamic underwater remote operated vehicle, ROV, and was selected due to its fluid dynamic properties. The base design is simulated to obtain aerodynamic properties and is compared to the first model. As illustrated in Figure 9.0d of a flowchart showing a method employed for the present invention in designing the fuselage, improvement in terms of aerodynamic properties that give better performance compared to the first model of the present invention results in proceeding with the rest of the design for building the improved final model of the present invention.

1

THEORY

Wing theory

The configuration used for the present invention is a blended wing body that is defined by the absence of an empennage and smaller wetted area when compared to a conventional configuration as shown in Figure 10.0a which illustrates force balance on a typical airplane with an empennage. Conventional aircrafts have tail sections with vertical stabilizers in order to counteract the moment generated at the aerodynamic centre. For a blended wing body, since there is no vertical stabilizer, a special airfoil known as the reflexed airfoil is used to provide the stabilization. As shown in Figure 10.0b, the reflexed airfoil has a negative camber at the trailing edge,TE. Figure 10.0a and Figure 10.0b show that the moment caused by the weight about the aerodynamic centre is ‘nose-heavy’. However, the negative camber of Figure 10.0b causes a clockwise moment at the reflex about the aerodynamic centre and balances out the nose heavy moment and eliminates the need for a dedicated empennage with vertical stabilizers. The reflexed airfoil selected is HS522.as shown in the graph of Figure 10.0c is a graph that was loaded into the analysis tool XFLR5. For the reflex mechanism illustrated in Figure 10.0b to be effective, the aerodunamic centre, AC of the aircraft must be placed behind the centre of gravity. A browser tool ‘rcwingcog’ is used in determining the aerodynamic centre as shown in Figure 10.0d. The parameter input for obtaining mean aerodynamic length is shown in Figure 10. Od.

Propulsion Theory

The unmanned aerial vehicle of the present invention is propelled in vertical flight by three motor-powered electric propellers as shown in Figure 10.1a. Figure 10.1a is a tricopter configuration showing the forces and moments involved in the analysis. For axial flight, each of the propellers contributes to the minimum lift threshold for 7.5 kg overall mass. Using the actuator disc theory, the take-off and landing of the tri-copter configuration can be modelled to obtain an estimate for the motor power value and propeller diameter.

Reference is made to Figure 10.2a illustrates an actuator disc theory force diagram for climb. The thrust acting on the rotor disc is calculated as follows ^T = ^A (Pi ~ Pu)

On the upper half of the rotor, the total pressure must be the same near the disc and further upstream. This is represented by the following equation for the upper region and the following equation for the lower region

Combining the equations of the upper region and the lower region will produce the following equation

R - Pu = p(y _c v ₂ + ^v 2 ² )

Combining the above equation with the equation on thrust acting on the rotor disc produces the following equation

The mass flow rate equation through the actuator disc is calculated as follows The rate of momentum is the thrust and is represented by the following equation

Combining the equation of and

5 produces the following equation l? ₂ = 2v _{ -

The above equation is then substituted into to produce the following equation v ² + V _c vt ~ v ₀ ² = 0 where .

The induced power for the rotor can then be obtained from p _t = T(y _c + _Vi )

Figure 10.2b is a table showing the input and output parameters to determine induced power for climb. By substituting values from Figure 10.2b, induced power for each rotor can be obtained.

A similar derivation of the climb theory can be deciphered to obtain the following equation for induced power for descent motion Pi = T(-v _D + _Vj )

Figure 10.2c illustrates an actuator disc theory force diagram for descent and Figure 10.2d is a table showing the input and output parameters to determine induced power for descent.

Applying a safety factor of 1.5 which is the minimum power using Autonomous Distress Tracking, ADT is 426 W. The values will change as the propeller diameter and climb/descent speed changes.

A study of induced climb power required in relation to the propeller diameter at a constant climb speed of 8.33 ms' ¹ is further shown in the graph of Figure 10.2e. The graph shows induced power against diameter of a propeller. The graph shows that for the same climb velocity, the greater the diameter of the propeller, the less induced power consumed by the motor.

For forward flight which is usually employed in typical aircrafts, the power required for cruise can be obtained from the following equation

P _T =

When in forward flight, a typical aircraft must have a careful balance of four physical forces which is the lift, drag, weight and thrust as shown in the force-balance diagram of Figure 10.2f. An aircraft’s lift must balance its weight and its thrust must balance its drag.

COMPUTER AIDED DESIGN, CAD, APPLICATION

CAD is a software that assists in the creation, modification, analysis and optimization processes. CAD was used in designing the present invention for aerodynamic surface designs, part designs and preparing geometry for numerical analysis. The main CAD software used is CATIA v5 R21 .

Fuselage Surface Design using CAD

Using section views of the Power Ray as reference and the ‘3D Curve’ tool in CATIA v5, a 3-D model of the Power Ray was produced.

The first part of creating a fuselage design is in sketching the design using a sketch tracer. The next part is in designing the curves of the fuselage before filing the surfaces between the curves. Finally the solid lines are closed for to ensure a full design without gaps.

Sketching

The steps involved in the first step of producing a 3-D model similar to the Power Ray is illustrated in Figure 11.0a, Figure 11.0b, Figure 11.0c, Figure 11. Od, Figure 11. Oe and Figure 11. Of which involves the sketch tracer. Figure 11.0a shows the first step in sketching which involves opening the CATIA v5 R21 and selecting File>Shape>Sketch Tracer. Next, the ‘shade with material’ icon at the bottom taskbar is selected as shown in Figure 11 ,0b. The next step is to select ‘create an immersive sketch’ on the tool bar on the right as shown in Figure 11 .0c for the top view of the model. After the image has been loaded a ‘top’ view can be selected located beside the Shading icon. The step is repeated for ‘front’ view and the ‘side’ views as shown in Figure 11. Oe. The sizing of each view are adjusted as shown in Figure 11. Od to ensure each of the three images with the ‘top’, ‘front’ and ‘side’ views line up. Once the main fuselage is designed, new parts can be added to the design by selecting components>new part as shown in Figure 11 .Of. Curves of the fuselage

For drawing and designing the curves, the icon ‘tools’ is selected before selecting ‘customize’. The ‘commands’ tab is selected where the ‘categories’ will be displayed. The icon ‘All commands’ is selected. Under ‘All commands’ the ‘3D Curve tool can be selected as shown in Figure 11 .0g where the 3-D curve tool is added. Before drawing a 3D curve, changing the views of the design to an isometric view is required and then selecting ‘lock the compass privilege plane as shown in Figure 11. Oh. The 3-D curves can now be drawn and the points in the curves can be matched to the images previously designed in the Sketch tracer that was imported. As shown in Figure 11 .0i, the 3-D curve can be adjusted to orient with the imported sketch.

Filling Surface of the curves

Once the 3-D curves have been connected to form closed geometry the ‘fill’ command is used as shown in Figure 11. Oj. The ‘Multi-section Surface’ & ‘Fill’ tools in the generative shape design are used to create surfaces with the 3-D curves that were earlier created.

Closing a surface of the Design model

Before a computer model of the unmanned aerial vehicle is used in computational fluid dynamic, the surface of the model must be converted to a solid body using the ‘close surface’ or ‘thick surface’ tool to ensure complete join lines and no gaps between any of the surfaces created. As shown in Figure 11.0k, the ‘Join’ or ‘Healing’ command is used to seal gaps between the created surface which allows for the use of ‘Close Surface’ tool to convert the surface into a solid. All surface must be connected before using closing the surface. The entire surface must only have one ‘hole’ which edge must be coplanar as shown in Figure 11.01. The result of closing a surface is shown and the healed surface has only one opening with coplanar points. The highlighted plane is used to cut the healed surface. Using the plane to intersect with the surfaces, a planar surface can be ‘split’ from the main body and the rest of the surface can be closed. The split surface is then closed to obtain a solid part as shown in Figure 1 1.0m and Figure 11. On. Once the surfaces have been closed, the surface is hidden to reveal the solid body and then the file is saved as an STP file. A final product of the unmanned aerial vehicle design is then produced as shown in Figure 11. Oo. Figure 11. Oo shows a comparison of the computer aided design of a 3-D model of the Power Ray and the final Iteration design.

Wing Design Using CAD

To design the wing using CAD software, airfoil data is required to be imported into the system. A file path is selected in the browser C:\Program File (x86)\Dassault Systemes\B21 \intel_a\code\command which is shown in Figure 11.1 a. The browser shows the indicated file path in excel for designing the wing. The excel file ‘GSD_PointSplineLoftFromExce is open for editing of the Generative Shape Design (GSD). Editing of the GSD excel sheet includes deleting the cells as shown in Figure 11.1 b, copying the cells for temporary storage as shown in Figure 11.1c and continuing editing of the excel sheet as shown in Figure 11 .1 d. After that the airfoil parameters are imported from the .dat file as shown in Figure 11.1 e. The excel sheet is then further edited where the first cell is replaced with the intended root chord as shown in Figure 11.1f. The last row of the data set is then further edited as shown in Figure 11.1g. The file is then imported. First, the icon ‘options’ under ‘file’ tab is selected as shown in Figure 11.1 h. Next, the category ‘Data’ is selected for file to be imported as shown in Figure 11.1i. Access to the ‘Developer’ tab in the excel sheet will then be available as shown in Figure 11.1j where the Macro name ‘FeuiU .Main’ is selected. A HS522 aifoil can then be loaded into CATIA v5 via the Excel sheet as shown in Figure 11.1 k. At least wo identical airfoil profiles can be uploaded having different chord lengths as shown in Figure 11.11. The profiles can then be connected using 3-D curves as shown in Figure 11.1 m. To ensure that the wing design does not have a varying Angle of Attack (AoA), the 3-D curve joining the trailing edge (TE) must be horizontal as can be seen in Figure 11 .1 n. A multi-section surface or fill is then created by using the two profiles with the 3-D curves as guidelines as shown in Figure 11 .1 o.

Analytical Solutions

Using the theory as previously mentioned, analytical solutions can be obtained. The analytical solutions allow us to quantify performance. Analytical solution aids in the selection of motors of the present invention. Using the induced power value calculated by applying a safety factor of 1.5, the minimum power required using ADT is 426 W where the climb speed is approximately 8.33 ms ¹ . Figure 11.2a, Figure 11.2b and Figure 11.2c show the data sheet for 3 different motor candidates which are U7 V2.0 KV280 T-Motor, U7 V2.0 KV420 T-Motor and U7 V2.0 KV490 T-Motor respectively.

The factors considered in motor selection are the propeller diameter, operating voltage and operating temperature to produce at least 426 W of power. A smaller diameter will ease geometrical constraints on the fuselage as there will be greater clearance for collision between the propeller and the fuselage. The preferred motor by using the analytical solution is the U7 V2.0 KV420 T-Motor having a 16” x 5.4” Carbon Fibre Propeller.

COMPUTER AIDED ENGINEERING, CAE, APPLICATIONS

In designing the parts of the unmanned aerial vehicle, predicting aerodynamics and structural dynamics is required.

One of the software used in studying the aerodynamics of the unmanned aerial vehicle is the XFLR5. The XFLR5 is used to obtain orders of magnitudes and trends and to understand sensitivity to design parameters before performing airfoil analysis. The results of the airfoil analysis can be seen in Figure 11 .3a, Figure 11 .3b and Figure 11 .3c. Figure 11.3a is a graph showing lift coefficient against drag coefficient. The maximum lift occurs at a drag of about 0.03. Figure 11 .3b is a graph showing lift coefficient against angle of attack on the left and a moment coefficient against the angle of attack on the right. The lift coefficient against angle of attack shows a value of 5.73. Figure 11.3c is a graph showing life coefficient against top transition point on the left and lift drag ratio against angle of attack on the right. The analysis is performed by selecting Polars > Batch Analysis and selecting the settings as shown in Figure 11 .3d.

Computational Fluid Dynamics

To validate the results of the experiment, a fluid flow analysis of a cylinder at Reynolds’ Number, Re = 500000 was performed using ANSYS Fluent. Figure 11.4a illustrates fluid flow around a cylinder for validating experimental results. The results converged to a coefficient of drag, C _D , of 1 .42 which matches the analysis shown in Figure 11 .4b which is a graph of the drag coefficient against Reynolds number. By using the same Re and Mach number, the right set up for computational fluid dynamics is used. In addition, running the set-up for the X8 Skywalker CAD model yields very similar coefficient values when compared to known wind tunnel data. Slight difference may be due to an inherent angel of incidence occurring either during manufacturing of the actual experiment set up or due to inconsistency of the CAD model and the model used in the wind tunnel experiments which originate from different sources. Figure 11.4c is a table showing a small range of Angle of Attack as backed by ANSYS Fluent. The angle of attack is accepted as the initial boundary conditions of the analysis.

The initial conditions calculation is determined by the equation below to calculate the Reynolds’ number.

Running Computational Fluid Dynamics Simulation

To run the CFD simulation, the ‘Fluid Flow, Fluent’ icon is dragged into the working environment as shown in Figure 11.5a. The next step is to import the external CAD geometry as shown in Figure 11 .5b.

Preparing Geometry in Computational Fluid Dynamics

In preparing the geometry of the unmanned aerial vehicle, the ‘Geometry’ section in the new node created is selected. This enables the ANYSYS Design Modeler to open. Then go to File > Import External Geometry File and then select the IGS file previously saved with CATIA. The object in the tree on the left is right clicked and the icon ‘Generate’ is selected. The ZX plane is selected as shown in Figure 11.5c.Next, click create> slice where the XY plane is selected to cut the geometry symmetrically in half. This process reduces meshing time as well as computation time. This splits the fuselage into two parts as shown in Figure 11 .5d. The two split fuselage parts may be suppressed when working on one of the fuselage part. The fluid zone may not be created.

A new sketch is created on the XY plane as the first step in creating the fluid zone as shown in Figure 11 ,5e.The ‘Arc by Center’ is selected to create and arc around the origin. The dimensions of the arc are edited by clicking on ‘Dimensions > Radius’ in the Sketching toolbox. A dimension of about 15 times the height of the fuselage is selected. This increases the volume of the fluid between the fuselage surface and the walls of the enclosure, which makes the wall effects in the fluid negligible. Before exiting the sketch, a line joining both ends of the arc is created. The arc is resolved 90° around the YZ plane which provides the front-section of the intended C-mesh grid fluid zone. The fluid zone can be extended as shown in Figure 11.5f which is important for turbulence modelling. The flat surface of the body is extruded to extend in the x direction which ensures the extended length is about 10 times the length of the fuselage. In determining the fluid surrounding the unmanned aerial vehicle, the solid is subtracted from the fluid zone as shown in Figure 11.5g. To subtract, select ‘Create >Boolean’. Then change the operation to ‘Subtract’ and the fuselage will be subtracted from the fluid zone. This prepares the geometry for meshing.

Meshing

To create mesh, the ANSYS Mesher in ANSYS Workbench is opened by double-clicking the ‘Mesh’ section in the project space. Once the mesh is loaded, the name selections are created as shown in Figure 11 .5h for the fluid zone surfaces.

The ‘Mesh’ section is clicked in the tree and the icon’ generate is selected to generate a basic mesh. To capture turbulent boundary layers, the requirements is to ensure the ‘First Cell Thickness’ is within range of the Y+ values. The accepted Y+ values used for the viscous flow in ANSYS Fluent is shown in the table in Figure 11 .5i.

The mesh is prepared according to the k-Q SST model where y+ <1. To do this an ‘Inflation Layer’ must be added to the mesh. The ‘Mesh’ icon is selected and then the icon ‘Inflation’ For the geometry, the entire fluid zone is selected. For the boundary, the ‘fuselage surface’ is selected. Next, the ‘Inflation Option’ is changed to ‘First Layer Thickness’ and the Y+ values are inserted. Figure 11.5j illustrates wall spacing value generated as output using a desired y+ value and the initial conditions as input. For the present invention the value used is 0.00196 mm and the maximum layers are changed to 35 before selecting the ‘generate’ icon. This produces the inflation layers in the mesh needed to capture the boundary layer as shown in Figure 11 .5k.

Next, the fluid zone region needs to be sized. The icon ‘Mesh’ is clicked and from there, navigated from ‘Mesh’ > Insert > Sizing to apply sizing to the fluid zone geometry. In the selection ‘Element Size’, a size of 35 mm is input. This element size is altered when doing a mesh dependence study for each of the meshes used in designing the unmanned aerial vehicle. Figure 11 .51 is a graph showing convergence to a single value with increasing number of elements which is an example of sizing the fluid zone region. Once the mesh has been generated, the ‘Setup’ in ANSYS Workbench is opened. The icon ‘Double Precision’ option is selected to open the Fluent interface. The icon ‘Models’ is selected to open the viscous model where the most suitable model is selected for the present invention as shown in Figure 11.5m. The ‘k-omega, 2 eqn’ model is selected and the radio option ‘SST’ is selected. This option allows solving of up to Y+ <=1 of the boundary layer providing an accurate prediction at the cost of computation time and is better for flows where a separation is predicted particularly in locating the stall point of the wings. Then, select ‘ok’. Next, the ‘Boundary Conditions’ as shown in Figure 11.5n is selected, and the ‘outlet’ which has been previously defined in the mesher and change the ‘Type’ to ‘outflow’. The icon ‘Ok’ is selected. The boundary conditions at the inlet are specified as shown in Figure 11 .5o for a 0° AOA. The speed depends on the scale of the model and the components of velocity can be changed to obtain an angle of attack. The reference values are then set as shown in Figure 11.5p.ln ‘Reference Values’, the ‘compute from’ parameter is set to ‘inlet’ which will change the ‘velocity’ parameter to the magnitude of velocity at the inlet. The ‘Area’ parameter is changed to the spanwise area of the model which can be obtained from CATIA v5 since it is a lifting body. The ‘Length’ parameter is changed to the root chord of the wing. The drag produced for a reference area using either the frontal area or the planform area is the same with the only difference being the matter of data packaging. Next, the monitors are set up and created for drag, lift, pressure and turbulence as shown in Figure 11.5q. Selecting the right direction is important that is consistent with the imported model and angle of attack needed. The ‘turb-y-plus’ monitor checks for mistakes where an incidence of >=1 a mistake can be detected. Next, the ‘Standard Initialization’ is selected. The ‘Compute From’ parameter is then changed to ‘inlet’ before selecting ‘initialize’ as shown in Figure 11.5r. This provides the initial conditions to calculate the solution. Finally, calculation of the solution is performed by selecting ‘Run Calculation’ having an iteration number of 1000 as shown in Figure 11 .5s.

Computational Fluid Dynamics, CFD, Analysis

Three types of CFD analysis was performed which is the Fuselage Analysis, Winglet Analysis and the UAV Aerodynamic Performance Analysis.

In Fuselage analysis, different fuselage designs were created and tested against the X8 Fuselage where the wings are not combined in the model. The evolution of the design (side views) of the fuselage can be seen in Figure 11 .6a. The fuselage analysis provides qualitative analysis for fuselage aerodynamic performance in a short computation time. The results of the fuselage analysis can be shown in Figure 11 .6b on the lift to drag ratio against iteration progression.

In the Winglet analysis, six different winglets were designed as shown in Figure 11 .6c where the built-in incidence used was 1 .29°, with a Re=500000 which can be increased. The final winglet design as shown in Figure 11.6d was selected having and AOA, Incidence of 0°, 3.79° which is Iteration 4-5 which is the design that produced the best Cl/Cd value.

As the present invention comprises VTOL capability, there must be enough lift at 0° AoA to ensure there is no vertical drop during the transition from VTOL to fixed wing configuration. Hence, a built-in incidence is required where the numerical result is obtained from the UAV Aerodynamic Analysis. To determine the required incidence, different AOA from 0° to 5° was analysed to determine the lift drag values for different angles of attack as shown in Figure 11 .6e. For a lift of 49 N with a 5 kg mass, a C _L value of 0.287 is required. The built-in incidence was increased by 2.5° to provide enough lift at 0° AoA. The final new incidence for the CAD model is 3.79° which completes the UAV Aerodynamic Performance analysis.

Finite Element Analysis, FEA

Finite Element Analysis, FEA, is a type of CAE that simulates forces acting on the UAV during operating conditions. It predicts the limit loads of the UAV before testing and is key in defining flight envelope. A finite element analysis is performed on parts of the design to test if the designs meet the requirements. A factor of safety, known as the Reserve Factor RF of 1 .5 is required to be applied to the limit load and is calculated as follows: fese Factor,

The resulting limit load multiplied by the factor of safety is then taken as the ultimate load. Before performing FEA, the load cases of the operating conditions of the UAV are identified and the description of the load case numbers is shown in Figure 11 .6f. The load cases can be further illustrated in force balance diagrams which show the load path of the forces and reveal how the structures should be connected. The forces transfer through Bending Stress, BS, or Plane Stress, PS. The point at which the stresses change in form is the connection point and physically requires a bracket component. The different force balance diagrams for different load cases are shown in Figure 11.6g illustrating load cases on the centre wingbox for fixed wing flight cruise and fixed wing acceleration, Figure 11.6h illustrating load cases on the centre wingbox for take-off, Figure 11 .6i illustrating load cases on the centre wingbox for VTOL transition at different angles, Figure 11 .6j illustrating load cases on the centre wingbox for fixed wing banking, Figure 11.6k illustrates load cases on the centre wingbox for gust loads, Figure 11.61 illustrating load cases on the tail section for steady cruise and fixed wing acceleration, Figure 11 .6m illustrating load cases on the tail section for transition flight, Figure 11 .6n illustrates load cases on the tail section for take-off, Figure 11 .60 illustrating load cases on landing gears for static ground and Figure 11 .6p illustrating load cases on the landing gears for landing.

Assembly design

In designing a fixed wing VTOL UAV, the placement of components within the structure is important. Figure 11 .7a illustrates an assembly design showing the equidistant placement of the motors of the present invention for VTOL capability. To ensure steady aerodynamics, the load paths of the lifts produced by each propeller must be the same distance, approximately 1100mm separating each motor. Further as shown in Figure 1 .7b, the back propeller is set to have a minimum offset of 1 .4cm from the fuselage surface.

Throughout this specification, unless the context requires otherwise, the word “comprise”, or variations such as “comprises” or “comprising”, will be understood to imply the inclusion of a stated step or element or integer or group of steps or elements or integers, but not the exclusion of any other step or element or integer or group of steps, elements or integers. Thus, in the context of this specification, the term “comprising” is used in an inclusive sense and thus should be understood as meaning “including principally, but not necessarily solely”.