The present invention relates to systems, arrangements, structures and methods for aircraft. In particular, the present invention relates to systems such as propulsion management systems, cooling systems, energy recovery systems, power management systems, electrical power management systems, propulsion systems, monitoring systems, spar arrangements, fluid fuel monitoring systems, power system structures and associated methods.

BACKGROUND

An aircraft flight includes the following phases: taxiing to the runway; takeoff (which includes the takeoff roll); climb; cruise; descent; final approach; and landing (which includes the landing roll). In the taxiing phase, the aircraft moves to the runway under its own power. The aircraft is positioned on the runway ready for the takeoff roll. In takeoff, thrust from the aircraft propellers or jet engines accelerate the aircraft up to speed during a takeoff roll, and once a suitable speed is achieved the nose of the aircraft is raised to increase lift from the wings and effect take off. In the climb phase, the aircraft climbs to cruise altitude and the aircraft speed is gradually increased. In cruise, the aircraft flies at a required altitude and, typically, at a constant speed. In the descent phase, the aircraft altitude is reduced by increasing drag. In the final approach, the aircraft is aligned with the runway before landing. In the landing phase, the aircraft lands on the runway and performs the landing roll, in which the aircraft is decelerated.

Aircraft noise pollution is an important consideration in aeronautics. Considerable noise is produced by the aircraft and its components during the various phases of aircraft flight. Of particular importance in relation to noise pollution are the phases of takeoff, climb, descent and landing. During these phases, the aircraft may pass close to residential areas. It is therefore common for flying restrictions to be applied to airports. These restrictions often require a night period in which aircraft may not be scheduled to takeoff or land at the airport’s runways. Both propeller and jet aircraft produce large amounts of noise during takeoff and climb. In propeller aircraft, this is because the propellers are operating far from optimum efficiency as a result of the low air speed. In jet aircraft, a large source of takeoff noise is a result of the difference between the air speed and the speed of the jet leaving the engine, which creates turbulence. Aircraft vibrations as a result of the large thrust required during takeoff and climb also produce noise. Both types of aircraft also produce considerable noise during descent and landing. Air brakes may be employed to increase drag, resulting in turbulent air and the creation of noise. Angles of descent are typically shallow as the aircraft must maintain speed to ensure lift, whilst increasing drag through use of the air brakes.

Hybrid and electric aircraft are powered at least in part by batteries. In such aircraft, it is beneficial to monitor and maintain battery health during operation, which may involve monitoring the battery power level and temperature. Batteries under large loads typically require cooling. In hybrid aircraft, the power supply must be managed to ensure that it is being used most effectively. Refuelling aircraft, including charging batteries, conventionally takes place on the ground. Upon landing, aircraft wheel brakes may be applied, which creates significant heat which must be brought below a required temperature before subsequent takeoff. These factors increase “turnaround time”. Turnaround time is the time between landing and a subsequent takeoff, wherein the aircraft is not in operation. Decreasing turnaround time is of high importance to airlines.

It is necessary to monitor, manage, and distribute power effectively from the power sources installed in the aircraft. Moreover, it is important to ensure redundancy in the event of component failure, in order to meet the stringent regulations placed on aircraft operation. Monitoring component health can be advantageous in preventing degradation and failure, and also in optimising power usage from power sources to ensure that despatch reliability and short turnaround time is achieved. Commercially available systems for the aforementioned purposes are not adequately suited for hybrid electric aircraft, and do not provide the levels of redundancy or monitoring capability that is necessary for safe and reliable operation of hybrid electric aircraft.

Monitoring fluid fuel use is of particular importance. Where a plurality of fuel types are to be used, the current approach is to corrupt the fuels such that they have identical properties, such as viscosity and density. This is expensive and is inhibiting progress in the aviation sector. Segregating and isolating components is an important consideration in aircraft design to prevent electrical arcing, which can lead to aircraft failure. Current commercially available power system structures for aircraft are not suitable for use in hybrid electric aircraft to sufficiently minimise component heating and weight of electrical cabling. Moreover, current commercially available power system structures do not provide the level of redundancy necessary for safe and reliable operation of electric and hybrid-electric aircraft.

A more holistic approach to hybrid, or fully electric, aircraft design and manufacture is required, instead of simply trying to use or adapt existing systems. Nevertheless, adapting existing systems is advantageous where safety and reliability of aircraft function can be ensured. Perhaps more generally, there is a drive to reduce pollution and other negative environmental impacts of aircraft and the aerospace industry as a whole.

It is an object of the present invention to provide improved systems, arrangements and structures for aircraft and/or to address one or more of the problems discussed above, or discussed elsewhere, or to at least provide alternative systems, arrangements and structures.

SUMMARY OF THE INVENTION

According to the present invention there is provided a propulsion management system, a cooling system, an energy recovery system, and associated methods, as set forth in the appended claims. Other features of the invention will be apparent from the dependent claims, and the description which follows.

According to a first aspect of the present invention there is provided a propulsion management system for an aircraft, the aircraft comprising one or more wheels, and a ground propulsion system arranged to drive one or more wheels of the aircraft, the propulsion management system being arranged to control the ground propulsion system to drive the one or more wheels of the aircraft during a takeoff roll, the takeoff roll being subsequent to a taxiing phase. Such a construction enables the aircraft to be accelerated during a takeoff roll using a ground propulsion system. This facilitates numerous advantages. An aircraft provided with a propulsion management system can take off in a shorter distance than a conventional aircraft. This allows the aircraft to takeoff from shorter runways, and also allows the aircraft to climb to altitude sooner, resulting in reduced noise at the ground. The ground propulsion system operates efficiently at low forward speeds, producing less noise than an air propulsion system, such as propellers, at low forward speeds. It is therefore well suited to the initial period of a takeoff roll and it is highly advantageous to provide a propulsion management system to control the ground propulsion system for such a purpose. Driving the one or more wheels during a takeoff roll may mean driving the one or more wheels for more than 5%, 10%, 15%, or 50% of the runway length, total takeoff time, total takeoff distance. The propulsion management system may control the ground propulsion system to drive the aircraft to a speed of more than 40, 50, 60 or 70 knots. The ground propulsion system may comprise an energy accumulator. The ground propulsion system may be powered by the energy accumulator. The energy accumulator may be a battery, a flywheel and/or a supercapacitor. Energy accumulators are advantageous as they may be provided with energy in a number of ways, facilitating energy recovery systems. Energy accumulators are also advantageous as they can have a small spatial footprint. Powering the ground propulsion system using the energy accumulator is advantageous as high power can be provided. Emissions are typically lower than for combustions engines. The aircraft may comprise an air propulsion system. The propulsion management system may be arranged to control the air propulsion system. The air propulsion system may comprise one or more propellers or jet engines. The propulsion management system may be arranged to control the air propulsion system to drive the one or more propellers or jet engines. Air propulsion systems propel the aircraft in flight. The propulsion management system being arranged to control the air propulsion system and the ground propulsion system allows the propulsion management system to manage power distribution between the propulsion systems. Propellers are often best suited for domestic, or short-haul, or regional, aircraft. Propellers also facilitate the provision of electric motors which may be powered by an energy accumulator. A propeller pitch may be optimised for a climb phase. Optimising for the climb phase is advantageous as this is an energy intensive phase of flight. Climbing to altitude quickly helps to reduce noise pollution at ground level. The one or more propellers may be fixed pitch propellers. Fixed pitch propellers are simpler in construction than variable pitch propellers. Control systems are not needed, resulting in a cheaper, more reliable and lighter aircraft. Despite their advantages, fixed pitch propellers are typically not employed in commercial aircraft, as they are inefficient at low aircraft forward speeds. By providing a ground propulsion system, efficient operation of fixed pitch propellers is thus facilitated. A better overall balance is struck, based on the use of a ground propulsions system. Also, electrical driving of a fixed pitch propeller may be easier implement, or control, than the use of combustion engines, in terms of output torque and so on. During a takeoff roll, the propulsion management system may be configured to cause the aircraft to move for a first time period using the ground propulsion system. The propulsion management system may be configured to cause the aircraft to move for a second time period using the air propulsion system. This facilitates a takeoff roll in which the aircraft initially travels along the runway using the ground propulsion system, and then using the air propulsion system. Overlap between the time periods result in faster acceleration, and therefore a sooner takeoff. No overlap between the time periods results in smoother acceleration , resulting in reduced aircraft vibrations and consequently a quieter takeoff, whilst providing a level of comfort to passengers or users. The first and second time periods may at least partially overlap. There may be an overlap time period during which the propulsion management system is configured to cause the aircraft to move using the ground propulsion system and the air propulsion system. The propulsion management system may be arranged to cause the ground propulsion system to drive the aircraft wheels independently. Driving the wheels independently allows for the propulsion management system to distribute power independently between the wheels. This is advantageous where wheel slip occurs. This is also advantageous where the aircraft deviates off a desired line, for example when taking off in strong crosswind. The propulsion management system may be arranged to receive a situational condition. The propulsion management system may be arranged to control the ground propulsion system to drive the aircraft wheels independently in response to the situational condition. The situational condition may comprise one or more of: wind-speed, wheel speed, travel path or trajectory. Such control, to take into account such conditions, is simply not possible without direct driving of the wheels. Instead, no such control is possible, or aerodynamic changes need to be implemented, which could comprise other aspect of control such as aerodynamic factors or settings for a takeoff. The situational condition may comprise a runway condition. A runway condition may be representative of the condition of the runway. In particular, the runway condition may inform the propulsion management system that the runway is dry, wet, or icy. Providing the propulsion management system with this information is advantageous as it allows the propulsion management system to adapt control of the aircraft in response to the runway condition. For example, where the runway is wet or icy, the propulsion management system may look to smooth acceleration profiles, or command a lower level of instantaneous ground propulsion system motor torque to ensure wheel traction. The propulsion management system may be arranged to monitor the speed of the one or more wheels. The propulsion management system may be arranged to cause drive to be apportioned between the aircraft wheels in response to the wheel speed. The propulsion management system thus ensures that the aircraft wheels do not slip, and maintain traction with the runway surface. Such control, to take into account such conditions, is simply not possible without direct driving of the wheels.

According to a second aspect of the present invention there is provided an aircraft comprising a propulsion management system according to the first aspect. The aircraft may comprise body-mounted landing gear. Alternatively, the aircraft may comprise landing gear provided on struts extending from beneath the aircraft wings or nacelles.

According to a third aspect of the present invention there is provided a method of aircraft takeoff comprising the steps of: a. providing an aircraft comprising one or more wheels and a ground propulsion system arranged to drive one or more wheels of the aircraft; and b. controlling the ground propulsion system to drive the one or more wheels of the aircraft during a takeoff roll, the takeoff roll being subsequent to a taxiing phase.

The ground propulsion system may be controlled to drive the one or more wheels for a first time period from the start of the takeoff roll.

According to a fourth aspect of the present invention there is provided a propulsion management system for an aircraft, the aircraft comprising one or more wheels, a power supply, a ground propulsion system arranged to drive one or more wheels of the aircraft and an air propulsion system, the propulsion management system being arranged to apportion power from the power supply between the ground propulsion system and the air propulsion system. Apportioning energy between the propulsion system is advantageous as it can lead to the power supply being used most efficiently or effectively. The propulsion management system can manage power apportionment for low noise takeoff, short takeoff, minimum energy usage or other requirements. It also simplifies pilot workload as the propulsion management system enables the aircraft to have the same cockpit control layout as in a conventional aircraft, as it can manage the apportionment autonomously, without input from the pilot. Additionally, the propulsion management system ensures safe operation of the aircraft systems. The propulsion management system may be arranged to apportion power between the ground propulsion system and the air propulsion system during a takeoff roll. A significant amount of power is required during takeoff, and so apportioning energy effectively by using the propulsion management system is highly advantageous. The propulsion management system may comprise a single actuator for the ground and air propulsion system. The propulsion management system may apportion power based on the level of actuation. Thus, the pilot need not focus on power apportionment. Additionally, a standard cockpit layout can be retained. The propulsion management system may be arranged to monitor the energy potential of the power supply. Energy potential may include battery charge level or fuel level. It is advantageous for the propulsion management system to be provided with this information such that it can make informed decisions on how to use the power supply most effectively. The power supply may comprise an energy accumulator and an internal combustion engine. The propulsion management system may be arranged to apportion the supply of power from the energy accumulator and the internal combustion engine. This facilitates efficient use of power. For example, where the battery charge is high, it may be preferable to use the battery to power the propulsion systems. This creates battery “space” which may allow it to be used for energy recovery purposes. On the other hand, if the battery charge is low, the propulsion management system may prioritise battery charging, and operate the internal combustion engine to supply power to the aircraft. Where maximum power is required, the propulsion management system may demand a supply of power from both power sources. The propulsion management system may be configured to control the supply of power from the power supply, in dependence on flight information (for example, automatically accessed by, or provided to, the system). The flight information may be one or more of: acceleration profiles stored in memory which result in: the quietest possible takeoff, the shortest possible takeoff, minimum power usage for takeoff, the smoothest possible acceleration, suitable levels of torque being applied to the wheels to ensure traction with the runway surface is maintained; deceleration profiles stored in memory which result in: the quietest possible landing, the shortest possible landing, minimum power usage to slow the aircraft, the smoothest possible deceleration, suitable levels of torque being applied to the wheels to ensure traction with the runway surface is maintained; power source management requirements, including: preserving battery charge, using excess charge to make capacity for regeneration (including decelerating the aircraft using the energy recovery systems), using excess fuel, battery temperature; flight plan information, including: the distance to the destination, time to destination, turnaround time at destination; airport data, including: runway length at takeoff airport or at destination airport, availability of charging ports at destination; or situational conditions, including: wind speed, wheel speed, travel path and trajectory of the aircraft. Providing such flight information is highly advantageous. Acceleration profiles allow the propulsion management system to apportion power, or use the power supply, in such a manner as to achieve the acceleration profile, which may be an optimum acceleration profile for purpose, such as quiet or short takeoff or minimum power usage. Providing deceleration profiles are similarly advantageous. Flight plan information allows the propulsion management system to apportion power based on journey requirements, and may, for example, look to preserve battery charge if turnaround time at destination is intended to be short. The propulsion management system may be configured to apportion power from the power supply between the ground propulsion system and the air propulsion system based on a required take-off, flight, or landing condition.

According to a fifth aspect of the present invention there is provided an aircraft comprising a propulsion management system according to the fourth aspect. The aircraft may comprise body- mounted landing gear. Alternatively, the aircraft may comprise landing gear provided on struts extending from beneath the aircraft wings or nacelles.

According to a sixth aspect of the present invention there is provided a method of managing propulsion in an aircraft, the method comprising the steps of: providing an aircraft comprising one or more wheels, a power supply, a ground propulsion system arranged to drive one or more wheels of the aircraft, and an air propulsion system; and apportioning power from the power supply between the ground propulsion system and the air propulsion system. The step of apportioning power may comprise: powering the air propulsion system to drive the one or more propellers of the aircraft; and not powering the ground propulsion system. Not powering the ground propulsion system may be advantageous to ensure smooth acceleration, or once the aircraft is travelling at a speed which leads to more efficient air propulsion system operation. The step of apportioning power may comprise: powering the ground propulsion system to drive the one or more wheels of the aircraft; and simultaneously powering the air propulsion system to drive the one or more propellers of the aircraft. Simultaneously powering the ground and air propulsion systems facilitates maximum acceleration. Power may be apportioned to the ground propulsion system to drive the one or more wheels of the aircraft until a threshold aircraft speed is achieved. The threshold aircraft speed may be one where the air propulsion system will operate above a certain level of efficiency. The step of apportioning power may be during a takeoff roll or landing roll.

According to a seventh aspect of the present invention there is provided an energy recovery system for an aircraft, the energy recovery system comprising: a transducer assembly arranged to selectively interact with one or more wheels of the aircraft to extract energy from the wheels; and an energy accumulator for storing the extracted energy. A substantial amount of energy is available for extraction from the aircraft wheels when they are in motion. In particular, during a landing roll, friction brakes heat up significantly. By providing an energy recovery system, this energy may be recovered and used, or stored in an energy accumulator. The energy recovered and stored can be used for subsequent operations and can alleviate the need to charge or supply power to the energy accumulator. This can help to reduce turnaround time. The energy accumulator may be a battery, flywheel or supercapacitor. The transducer assembly may be arranged to convert energy from the wheels into electrical energy. Electrical energy can be advantageously used to charge a battery. The transducer assembly may be arranged to convert rotational kinetic energy into electrical energy. The energy recovery system may be operated during a landing roll. As previously mentioned, a substantial amount of energy is available for extraction during a landing roll, resulting from the high kinetic energy of fast rotating aircraft wheels. The energy recovery system may be operated to provide braking effort during a landing roll. Such a construction is advantageous as it reduces heating of the wheels or any brake pads, recovers energy for subsequent use, as well as providing an easily controllable braking effort. The energy accumulator may comprise a battery. The energy accumulator may be comprised in a power supply for a propulsion system of the aircraft. The recovered energy can thus be used to power the aircraft propulsion system. The propulsion system may comprise a ground propulsion system arranged to drive one or more wheels of the aircraft. The transducer assembly may comprise one or more electric motors. The one or more motors may be arranged to drive the one or more wheels of the aircraft. The electric motors may be operated in reverse during operation of the energy recovery system. A ground propulsion system is advantageous for efficient aircraft takeoff and energy recovery on landing. The transducer assembly may be selectively engageable with the one or more wheels. The energy recovery system may cause the transducer assembly to selectively engage with the one or more wheels. The level of braking force applied to each wheel is thus controllable, ensuring safe deceleration with no skidding, and also ensuring efficient energy recovery. The transducer assembly may be engaged during a landing roll.

According to an eighth aspect of the present invention there is provided an aircraft comprising an energy recovery system according to the seventh aspect of the present invention. The aircraft may comprise body-mounted landing gear. Alternatively, the aircraft may comprise landing gear provided on struts extending from beneath the aircraft wings or nacelles.

According to a ninth aspect of the present invention there is provided a method of energy recovery in an aircraft, the method comprising the steps of: interacting a transducer assembly with one or more wheels of the aircraft to extract energy from the wheels; and storing the extracted energy in an energy accumulator. The transducer assembly may convert energy from the wheels into electrical energy.

According to a tenth aspect of the present invention there is provided an energy recovery system for an aircraft, the energy recovery system comprising: a transducer assembly arranged to selectively interact with one or more propellers of a propulsion system of the aircraft to extract energy from the propellers. Extracting energy from the propellers increases aircraft drag, allowing the aircraft to descend. By providing such an energy recovery system, the level of energy extraction allows the angle of descent to be controlled. A steep angle of descent is facilitated, which is advantageous in reducing noise levels at the ground. Extracting energy from the propellers also helps to slow the aircraft. Use of airbrakes can thus be minimised. Airbrakes create noisy turbulent air, so minimising their use is advantageous. The energy recovery system may comprise an energy accumulator for storing the extracted energy. Storing the extracted energy in an energy accumulator allows the energy to be used for other tasks. The stored energy may, for example, be used for subsequent takeoff procedures where energy is extracted during a preceding descent. This reduces turnaround time at the destination airport as it is not necessary to charge the battery on the ground. Alternatively, the energy may not be stored, and may simply be used to power components of the aircraft. The energy accumulator may be comprised in a power supply for the propulsion system of the aircraft. The energy accumulator may be a battery, a flywheel and/or a supercapacitor. The transducer assembly may be arranged to convert kinetic energy into electrical energy. The transducer assembly may be arranged to control the rotational speed of the one or more propellers. Controlling the rotational speed provides a control of aircraft drag. Increasing drag results in an increased descent angle, which is advantageous in reducing overhead noise. The propeller rotation may be slowed significantly so as to reduce energy expenditure, or, to recover energy. The transducer assembly may be arranged to inhibit motion of the one or more propellers. That is, the propellers may “windmill” or freely rotate. The transducer assembly may inhibit the free rotation. The level of inhibition of the propellers can be controlled so as to provide a desired level of battery charging. The energy recovery system may be operated to increase aircraft drag during a descent. Controlling the rotational speed provides a control of aircraft drag. Increasing drag results in an increased descent angle, which is advantageous in reducing overhead noise. One or more of the propellers may be fixed pitch propellers. Fixed pitch propellers are simple to construct and maintain. They are also lighter and do not require complex control systems.

According to an eleventh aspect of the present invention there is provided an aircraft comprising an energy recovery system according to the tenth aspect of the present invention. The aircraft may comprise body-mounted landing gear. Alternatively, the aircraft may comprise landing gear provided on struts extending from beneath the aircraft wings or nacelles.

According to a twelfth aspect of the present invention there is provided a method of energy recovery in an aircraft, the method comprising the steps of: interacting a transducer assembly with one or more propellers of a propulsion system of the aircraft; and extracting energy from the propellers. The extracted energy may be stored in an energy accumulator. The level of energy extraction may be controlled so as to variably inhibit rotation of the one or more propellers, such that aircraft drag is controllable during a descent. The one or more propellers that the transducer assembly is not interacting with may be driven. The one or more propellers that the transducer assembly is interacting with may not be driven.

According to a thirteenth aspect of the present invention there is provided a propulsion management system for an aircraft, the aircraft comprising an air propulsion system having one or more propellers, the propulsion management system being arranged to control the air propulsion system to drive the one or more propellers in reverse in order to provide a reverse thrust. Providing a reverse thrust is highly advantageous in rapidly decelerating the aircraft. This may advantageously be employed during a landing roll, or in the event of an emergency during takeoff. The propulsion management system may be arranged to control the air propulsion system to drive propellers in reverse during a landing roll. This allows the aircraft to land on short runways, such as in cities or on aircraft carriers. Being able to rapidly bring the aircraft to a halt during a landing roll is also a useful safety feature. The pitch of a fixed or variable pitch propeller may be optimised to provide a desired or necessary level of reverse thrust. A necessary level may be one that provides maximum deceleration, or deceleration that brings the aircraft to a halt in a predetermined distance. The air propulsion system may comprise one or more electric motors drivingly connected to the one or more propellers. The air propulsion system motors may thus be used to drive the propellers, which facilitates simple construction in which only a single set of motors are required for both forward and reverse thrust. Electrically driven propellers may, in general, be far easier to operate in reverse than combustion- engine driven propellers, for example requiring little or no mechanical implementation to achieve. Electrical control may be simpler and more direct. The aircraft may further comprise one or more wheels and a load assembly, and the propulsion management system may be operable to cause the load assembly to apply a load to the wheels of the aircraft to provide a braking effort during a landing roll. The load assembly may be a transducer assembly arranged to interact with one or more wheels of the aircraft to convert energy from the wheels into electrical energy. The electrical energy from the transducer may be supplied to the air propulsion system. Braking of the aircraft as a result of the load assembly will thus recover energy, which is advantageously provided to the propellers for increased braking by reverse thrust. Such a construction may partially or completely remove the need to provide a separate power source to achieve reverse thrust. The electrical energy may be supplied to a power supply for the air propulsion system.

According to a fourteenth aspect of the present invention there is provided an aircraft comprising an energy recovery system according to the thirteenth aspect of the present invention. The aircraft may comprise body-mounted landing gear. Alternatively, the aircraft may comprise landing gear provided on struts extending from beneath the aircraft wings or nacelles.

According to a fifteenth aspect of the present invention there is provided a method of decelerating an aircraft, the aircraft comprising an air propulsion system comprising one or more propellers, the method comprising the step of: operating the propellers in reverse to provide reverse thrust. The propellers may be operated in reverse to provide reverse thrust during a landing roll. Operating in reverse means reversing the rotational direction of the propellers from a rotational direction that produces forward thrust. The method may further comprise a step of adjusting the propeller pitch to provide a desired or necessary level of reverse thrust. The aircraft may further comprise one or more wheels and a transducer assembly arranged to interact with one or more wheels of the aircraft to convert energy from the wheels into electrical energy during a landing roll, thereby providing a braking effort to the aircraft. The method may further comprise the step of: supplying the electrical energy to a power supply for the air propulsion system.

According to a sixteenth aspect of the present invention there is provided a cooling system for cooling a battery in an aircraft using a flow of gas, the cooling system comprising: an inlet, an outlet, and a gas flow path extending between the inlet and outlet, and at least proximal to the battery to cool the battery, and a transducer provided in the flow path, the transducer arranged to convert energy of the gas flow into electrical energy. Such a construction enables thermal management of the battery, and simultaneous energy recovery. Substantial energy subsists in the gas flow (e.g. when the aircraft is in flight), and thus providing a cooling system arranged to convert that energy into electrical energy by virtue of the transducer is highly advantageous. Where the cooling system is configured for in-flight use, it will be possible to convert energy of the gas flow into electrical energy for the majority of, if not all of, the flight. The gas flow will typically be derived by the aircraft moving through the gas, such as air, typically in a flight condition. The battery may be comprised in a power supply for a propulsion system of the aircraft. Ensuring that the propulsion system battery is maintained at a suitable temperature is important as energy storage and battery efficiency is dependent on battery temperature. The gas flow path may be defined by a duct extending between the inlet and outlet. The electrical energy obtained from the transducer may be a function of the battery temperature. That is, a warmer battery may result in, or facilitate, increased energy conversion. This is advantageous as a hotter battery temperature will allow the rate of gas flow to be increased, the gas flow thus having greater kinetic energy as it travels along the gas flow path. The cooling system may be arranged to monitor the temperature of the battery. The cooling system may be configured to control a mass flow of the gas along the gas flow path as a function of the component temperature. Monitoring the battery temperature is important for battery health. A greater mass flow of gas along the gas flow path will be needed if the battery temperature is greater than desired or required. Additionally, the cooling system therefore does not require additional valves or flow regulators, which increases simplicity and reduces aircraft weight. Alternatively, additional valves or flow regulators may add additional degrees of control, or redundancy. The transducer may be a turbine. The turbine may be provided in the gas flow path. The turbine may be arranged to be rotated by the flow of gas. A turbine is advantageously suited for capturing the kinetic energy of the gas flow. The mass flow may be a function of the rotational speed of the turbine. This is advantageous as battery cooling is related to turbine speed. The turbine may be downstream of the battery. This advantageously encourages flow along the gas flow path and over the battery, and at the same time is a good position for recovering energy from the gas. The turbine may be comprised in a wheel of the aircraft. The turbine may be comprised in a wheel hub. A wheel hub may comprise vanes. Advantageously, a separate turbine need not be provided. The wheel hub may spin separately to the wheel. Vanes assist in capturing the kinetic energy of the gas flow. A part of the aircraft wheel may be exposed from the belly fairing such that during flight, the wheel is rotated by ambient airflow past the wheel. This encourages airflow along the airflow path, thus cooling the battery. The cooling system may be arranged to direct the electrical energy from the transducer to the battery or to a propulsion system of the aircraft. Advantageously, the recovered energy can therefore be stored for later use, or used to propel the aircraft. In another example, there may be no battery charging, and the system can simply be employed to cool the battery, and in a controlled manner. Alternatively, the cooled battery and the charging battery may not be the same. One battery may be cooled, and another (different) battery may be charged. The propulsion system may comprise a ground propulsion system arranged to drive one or more wheels of the aircraft. The transducer may be comprised in the one or more wheels.

According to a seventeenth aspect of the present invention there is provided an aircraft comprising a cooling system according to the sixteenth aspect of the present invention. The aircraft may comprise body-mounted landing gear. Alternatively, the aircraft may comprise landing gear provided on struts extending from beneath the aircraft wings or nacelles.

According to an eighteenth aspect of the present invention there is provided a method of cooling a battery in an aircraft using a flow of gas, the method comprising the steps of: directing a flow of gas proximal to the battery to cool the battery; and converting energy of the gas flow into electrical energy using a transducer.

According to a nineteenth aspect of the present invention there is provided a spar arrangement for segregating components installed in an aircraft, the spar arrangement being locatable between the components, to provide one or more voids therebetween.

In one example, the components are formed from conductive materials. In one example, the components comprise one or more of: conductive cables, electric generators, fuel tanks and/or energy storage devices. In one example, the energy storage devices are batteries. In one example, the spar arrangement comprises a first component receiving portion and a second component receiving portion, the one or more voids provided therebetween. In one example, the first component receiving portion is arranged to receive a plurality of conductive cables, the first component receiving portion comprising one or more segregating members for segregating the conductive cables therein. In one example, the spare arrangement comprises two or more spar webs separated from one another to define the voids. In one example, the spar arrangement is formed from an insulating material.

According to a twentieth aspect of the present invention there is provided a spar assembly, comprising a first and second spar arrangement, each spar arrangement being according to the nineteenth aspect, wherein the components comprise a fuel tank, a first conductive cable and a second conductive cable, the first and second spar arrangements being located on opposite sides of the fuel tank, the first spar arrangement being arranged to segregate the first conductive cable from the fuel tank, the second spar arrangement being arranged to segregate the second conductive cable from the fuel tank.

According to a twenty first aspect of the present invention there is provided an aircraft wing comprising one or more spar arrangements, and/or spar assemblies, of the nineteenth and/or twentieth aspect. In one example, the spar arrangement is locatable in the aircraft wing between a fuel tank and one or more conductive cables, to provide one or more voids therebetween. In one example, a first spar arrangement is installed in a leading edge of the aircraft wing and/or a second spar arrangement is installed in a trailing edge of the aircraft wing.

According to a twenty-second aspect of the present invention there is provided an aircraft comprising a spar arrangement, spar assembly, or aircraft wing according to the nineteenth, twentieth and/or twenty first aspects.

According to a twenty-third aspect of the present invention there is provided a method of segregating components installed in an aircraft, the method comprising the steps of: providing one or more spar arrangements; and locating the one or more spar arrangements between the components to provide a void therebetween.

In one example of the method, the components comprise a fuel tank, a first conductive cable and a second conductive cable, and the method comprises: providing a first spar arrangement and a second spar arrangement; locating the first and second spar arrangements on opposite sides of the fuel tank; arranging the first spar arrangement to segregate the first conductive cable from the fuel tank; and arranging the second spar arrangement to segregate the second conductive cable from the fuel tank.

According to a twenty-fourth aspect of the present invention there is provided a fluid fuel monitoring system for a vehicle arranged to be fuelled by a fluid fuel supply, the system comprising: a sensing apparatus arranged to sense a fluid flow property of the fluid fuel supply, the fluid flow property being indicative of the composition of the fluid fuel supply; and a controller arranged to control operation of the vehicle based on the fluid flow property.

In one example, the system comprises an arrangement of conduits arranged to receive fluid fuel from a plurality of fuel tanks and combine the fluid fuels in a single conduit. In one example, the sensing apparatus is arranged to sense a fluid flow property of the combined fuels in the single conduit. In one example, the sensing apparatus is arranged to sense one or more of: fluid viscosity and/or density. In one example, the sensing apparatus comprises one or more flow meters for measuring fluid flow. In one example, measuring fluid flow may mean measuring flow rate. In one example, the sensing apparatus comprises two or more flow meters, and a fluid flow property sensed by each flow meter is compared to determine the composition of the fluid fuel supply. In one example, each flow meter comprises an opening of different diameter. In one example, the fluid fuel supply comprises kerosene and/or biofuel. In one example, the controller is arranged to control operation of fuel pumps, motors and/or electric generators, based on the fluid flow property. In one example, the controller is arranged to control operation of the fuel pumps, motors and/or electric generators in response to a change in the sensed fluid flow property, indicating a change in composition of the fluid fuel supply. In one example, the controller is arranged to control operation of the motors and/or electric generators in response to the change in the sensed fluid flow property to maintain power output by the motors and/or electric generators at a substantially constant level. In one example, the vehicle is an aircraft.

According to a twenty-fifth aspect of the present invention there is provided an aircraft comprising a fluid flow monitoring system according to the twenty-fourth aspect.

According to a twenty-sixth aspect of the present invention there is provided a method of fluid fuel monitoring in a vehicle arranged to be fuelled by a fluid fuel supply, the method comprising: sensing a fluid flow property of the fluid fuel supply, the fluid flow property being indicative of the composition of the fluid fuel supply; and controlling an operation of the vehicle based on the fluid flow property.

According to a twenty-seventh aspect of the present invention there is provided a power system structure for an aircraft comprising a nacelle, the power system structure comprising an electric generator for generating electricity to power a propulsion system of the aircraft, the electric generator being housed in the nacelle.

In one example, the propulsion system comprises an electric motor. In one example, the electric motor is also housed in the nacelle. In one example, the electric generator is arranged to provide power to the electric motor. In one example, the electric generator is arranged to provide power primarily to that electric motor. In one example, the nacelle is an inboard nacelle.

According to a twenty-eighth aspect of the present invention there is provided an aircraft comprising one or more power system structures according to the twenty-seventh aspect.

According to a twenty-ninth aspect of the present invention there is provided a method of structuring a power system for an aircraft comprising a nacelle, the method comprising: providing an electric generator for generating electricity to power a propulsion system of the aircraft; and housing the electric generator in the nacelle.

In one example, the method further comprises housing an electric motor in the nacelle. In one example, the method comprises powering the electric motor using the electric generator.

According to a thirtieth aspect of the present invention there is provided a power system structure for an aircraft, the aircraft comprising a flight control actuator system, the power system comprising one or more energy storage devices arranged to provide power to the flight control actuator system via power cables, the energy storage device being located proximal to the flight control actuator system.

In one example, the flight control actuator system comprises electric actuators and/or electric- hydraulic actuators. In one example, the flight control actuator system is arranged to actuate one or more control mechanisms and/or control surfaces. In one example, the control mechanisms and/or surfaces comprise: rudders, elevators, ailerons, spoilers, landing gear systems, steering mechanisms and/or flaps. In one example, the one or more energy storage devices comprises one or more batteries. In one example, the power system structure comprises a plurality of energy storage devices, preferably two energy storage devices, and the energy storage devices are interconnected such that in the event of failure of one of the energy storage devices, another one of the energy storage devices is arranged to provide power to the flight control actuator system. In one example, the energy storage devices are arranged to provide power primarily to the flight control actuator system. In one example, the energy storage devices are arranged to provide power only to the flight control actuator system. In one example, the energy storage devices are arranged to be trickle-charged by an external power source. In one example, the energy storage devices are arranged to be trickle-charged by one or more of: energy recovery systems, renewable energy systems, electric generators and/or another energy storage device.

According to a thirty-first aspect of the present invention there is provided an aircraft comprising a flight control actuator system and a power system structure according to the thirtieth aspect of the present invention. In one example, the power system structure and flight control actuator system are located at the rear of the fuselage. In one example, the power system structure and flight control actuator system are located in the aircraft empennage, wings and/or belly fairing.

According to a thirty-second aspect of the present invention there is provided a method of structuring a power system for an aircraft comprising a flight control actuator system, the method comprising the steps of: providing a power system comprising one or more energy storage devices; arranging the energy storage devices to provide power to the flight control actuator system via power cables; and locating the energy storage devices proximal to the flight control actuator system.

According to a thirty-third aspect of the present invention there is provided a power system structure for an aircraft, the aircraft comprising a flight control actuator system, the power system structure comprising one or more dedicated energy storage devices arranged to provide power primarily to the flight control actuator system.

In one example, the one or more energy storage devices are arranged to provide power only to the flight control actuator systems. In one example, the flight control actuator system comprises electric actuators and/or electric-hydraulic actuators. In one example, the flight control actuator system is arranged to actuate one or more control mechanisms and/or control surfaces. In one example, the control mechanisms and/or surfaces comprise: rudders, elevators, ailerons, spoilers, landing gear systems, steering mechanisms and/or flaps. In one example, the one or more energy storage devices comprise one or more batteries and/or fuel cells. In one example, the power system structure comprises a plurality of energy storage devices, preferably two energy storage devices, and the energy storage devices are interconnected such that in the event of failure of one of the energy storage devices, another one of the energy storage devices is arranged to provide power to the flight control actuator system. In one example, the energy storage devices are arranged to be trickle-charged by an external power source. In one example, the energy storage devices are arranged to be trickle-charged by one or more of: energy recovery systems, renewable energy systems, electric generators and/or another energy storage device. In one example, the energy storage devices are located proximal to the flight control actuator system.

According to a thirty-fourth aspect of the present invention there is provided an aircraft comprising a flight control actuator system and a power system structure according to the thirty third aspect. In one example, the power system structure and flight control actuator system are located at the rear of the fuselage. In one example, the power system structure and flight control actuator system are located in the aircraft empennage, wings and/or belly fairing.

According to a thirty-fifth aspect of the present invention there is provided a method of structuring a power system for an aircraft comprising a flight control actuator system, the method comprising the steps of: providing a power system comprising one or more energy storage devices; and arranging the energy storage devices to provide power primarily to the flight control actuator system.

According to a thirty-sixth aspect of the present invention there is provided an electrical power management system for an aircraft, the aircraft comprising a plurality of electrical power sources for powering components in the aircraft, the power management system comprising a plurality of power management modules arranged to manage electrical power from the electrical power sources to the components, wherein each of the plurality of power management modules is arranged to manage power from at least one of the electrical power sources.

In one example, the plurality of electrical power sources comprises a plurality of power sources of the same type and/or a plurality of power sources of different types. In one example, the electrical power sources comprise: an energy storage device, an electric generator, a photovoltaic panel and/or a turbine. In one example, each of the plurality of power management modules is arranged to manage power from one electrical power source, such that each power source has a dedicated power management module for managing the power from that power source only during normal operation. In one example, one of the dedicated power management modules is arranged to function as a backup power management module for one other power source in the event of failure of one of the dedicated power management modules, thereby to function as a non-dedicated power management module. In one example, each of the plurality of power management modules is arranged to manage power from a plurality of the electrical power sources. In one example, two or more of the power management modules are interconnected such that in the event of failure of one of the power management modules the other power management module is arranged to function as a backup power management module. In one example, the electrical power management system is arranged to manage electrical power from the electrical power sources to the components in dependence upon a target energy level of one or more of the electrical power sources. In one example, the electrical power management system further comprises at least one environment sensor arranged to monitor an environmental condition of, or in the region of, one or more of the electrical power sources, and the electrical power management system is arranged to manage electrical power from that electrical power source in dependence upon the environmental condition. In one example, the components comprise one or more motors and the power management modules are arranged to manage power from the electrical power sources to the one or more motors. In one example, the one or more motors are comprised in an aircraft propulsion system. In one example, the components comprise one or more energy storage devices. In one example, the one or more energy storage devices are comprised in an aircraft propulsion system.

According to a thirty-seventh aspect of the present invention there is provided an aircraft comprising an electrical power management system according to the thirty-sixth aspect.

According to a thirty-eighth aspect of the present invention there is provided a method of managing electrical power in an aircraft, the aircraft comprising a plurality of electrical power sources for powering components in the aircraft, the method comprising the steps of: providing a power management system comprising a plurality of power management modules arranged to manage electrical power from the electrical power sources to the components; and arranging the power management modules each of the power management modules to manage power from at least one of the electrical power sources.

According to a thirty-ninth aspect of the present invention there is provided a propulsion system for an aircraft, the propulsion system comprising a plurality of electrically powered propulsion apparatus and a plurality of electric generators for providing power to the propulsion apparatus.

In one example, each electrically powered propulsion apparatus comprises one or more of the electric generators. In one example, the electric generator is a turbogenerator. In one example, the electric generator is liquid or gas fuelled. In one example, each electrically powered propulsion apparatus has a dedicated electric generator for providing power only to that propulsion apparatus during normal operation. In one example, at least one of the electric generators is arranged to function as a backup generator for one other electrically powered propulsion apparatus in the event of failure of one of the dedicated electric generators, thereby to function as a non-dedicated generator. In one example, the propulsion system comprises electric generators located either side of the fuselage. In one example, one or more of the electric generators are housed in a nacelle of the aircraft. In one example, one or more of the electrically powered propulsion apparatus further comprises one or more energy storage devices. In one example, the energy storage devices are battery packs and/or fuel cells. In one example, one or more of the electric generators are arranged to provide power to the one or more energy storage devices for subsequent use in providing power to the electrically powered propulsion apparatus. In one example, the electrically powered propulsion apparatus are arranged to operate independently, such that in the event of failure of one of the electrically powered propulsion apparatus the others may continue to operate. In one example, at least one of the electrically powered propulsion apparatus comprises an electric motor for providing drive to a propeller, and the electric generator of the respective apparatus is arranged to provide power to the electric motor.

According to a fortieth aspect of the present invention there is provided an aircraft comprising a propulsion system according to the thirty-ninth aspect.

According to a forty-first aspect of the present invention there is provided a method of aircraft propulsion, the method comprising the steps of: providing a propulsion system comprising a plurality of electrically powered propulsion apparatus, and a plurality of electric generators; and generating electrical power using the electric generator and providing the power to the propulsion apparatus.

According to a forty-second aspect of the present invention there is provided a monitoring system for monitoring an electrical power source in an aircraft, the monitoring system comprising: at least one electrical power source sensor arranged to monitor an operating characteristic of the electrical power source; at least one environment sensor arranged to monitor an environmental condition of, or in the region of, the electrical power source; and a processing module arranged to receive the operating characteristic and the environmental condition from the sensors and process the information, to determine and/or predict future operating characteristics of the electrical power source.

In one example, the electrical power source is an energy storage device and/or an electrical generator. In one example, the operating characteristic is one or more of: current rate of charge, current rate of discharge, maximum possible rate of charge, maximum possible rate of discharge, capacity, charge/discharge cycle, cycle life and/or power density. In one example, the environmental condition is one or more of: component temperature, ambient temperature, humidity and/or pressure. In one example, the sensors are wireless sensors and the processing module is arranged to receive information from the wireless sensors. In one example, the processing module is arranged to process the information by comparing the operating characteristic and environmental condition with information stored in memory. In one example, the processing module is further arranged to receive flight information relating to the aircraft flight path, and the processing module is arranged to process the flight information alongside the operating characteristic and environmental condition information. In one example, the flight information comprises information relating to present and/or future: climb profile and rate, descent profile and rate, altitude, time to a target point in the flight, distance to a target point in the flight, landing conditions and/or infrastructure at the destination. In one example, the processing module is arranged to manage power from the electrical power source in dependence on information processed by the processing module. In one example, the processing module is arranged to preferentially use a particular power source in dependence on information processed by the processing module.

According to a forty-third aspect of the present invention there is provided an aircraft comprising a monitoring system according to the forty-second aspect.

According to a forty-fourth aspect of the present invention there is provided a method of monitoring an electrical power source in an aircraft, the method comprising the steps of: providing at least one electrical power source sensor arranged to monitor an operating characteristic of the electrical power source; providing at least one environment sensor arranged to monitor an environmental condition of, or in the region of, the electrical power source; and receiving the operating characteristic and the environmental condition from the sensors at a processing module, and processing the information to determine or predict future operating characteristics of the electrical power source.

According to a forty-fifth aspect of the present invention there is provided an electrical power management system for managing the level of an energy storage device installed in an aircraft, the electrical power management system comprising: a receiver arranged to receive second flight information relating to a second flight, the second flight subsequent to a first flight; and a processor arranged to process the second flight information to determine a target level of the energy storage device suitable for use in the second flight, the electrical power management system being arranged to manage the level of the energy storage device during the first flight to achieve the target level at a target time in the first flight.

In one example, the electrical power management system is arranged to manage a usage profile of the energy storage device to achieve the target level at the target time. In one example, the level is a charge level of the energy storage device. In one example, the second flight information comprises information relating to requirements for use of the energy storage device during the second flight. In one example, the second flight information comprises information relating to a take-off phase and/or initial climb of the second flight. In one example, the target level is such that the energy storage device is suitable for use in a take-off phase and/or initial climb of the second flight. In one example, the target level is such that the take-off roll and/or initial climb can be performed solely using the energy storage device. In one example, the energy storage device is a battery pack. In one example, the target time in the first flight is the end of the first flight. In one example, the first flight is a current flight and the second flight is the next flight. In one example, the level of the energy storage device is managed by managing power provided to the energy storage device by one or more of: energy recovery systems, electric generators, renewable energy systems. In one example, the electrical power management system is arranged to manage the level of the energy storage device during the first flight by increasing the level of the energy storage device from an initial level at the start of a descent phase to a final, higher, level at the end of the descent phase.

According to a forty-sixth aspect of the present invention there is provided an aircraft comprising an electrical power management system according to the forth-fifth aspect.

According to a forty-seventh aspect of the present invention there is provided a method of managing the level of an energy storage device installed in an aircraft, the method comprising the steps of: receiving second flight information relating to a second flight, the second flight subsequent to a first flight; processing the second flight information to determine a target level of the energy storage device suitable for use in the second flight; and managing the level of the energy storage device during the first flight to achieve the target level at a target time in the first flight.

According to a forty-eighth aspect of the present invention there is provided a power management system for managing a plurality of electrical power sources installed in an aircraft, the power management system comprising: a receiver arranged to receive second flight information relating to a second flight, the second flight subsequent to a first flight; and a processor arranged to process the second flight information to determine second flight usage requirements for one or more of the electrical power sources, the power management system arranged to selectively manage the power supplied by each electrical power source in the first flight to facilitate use of the one or more of the electrical power sources in the second flight according to the second flight usage requirements.

In one example, in response to determining second flight usage requirements for a first one of the electrical power sources, the power management system is arranged to preferentially use a second one of the second electrical power sources over the first one of the electrical power sources in the first flight. In one example, the processor is arranged to process the second flight information to determine a second flight usage profile for one or more of the electrical power source, the power management system arranged to selectively manage the power supplied by each electrical power source in the first flight so that the second flight usage profile is actionable in the second flight. In one example, the second flight information comprises one or more of: flight plans, mission restrictions, weather conditions, details of infrastructure at first flight destination and information relating to a takeoff phase of the second flight. In one example, the power management system facilitates use of the of the one or more of the electrical power sources in the takeoff phase of the second flight. In one example, the plurality of electrical power sources comprises an energy storage device. In one example, the second flight usage requirements require a part of the takeoff phase to be performed solely using the energy storage device, and the power supplied by each electrical power source is managed in the first flight so that the electrical power source is suitable for such use in the second flight. In one example, the energy storage device is a battery pack. In one example, the plurality of electrical power sources comprises one or more electric generators, energy recovery systems and/or renewable energy systems. According to a forty-ninth aspect of the present invention there is provided an aircraft comprising a power management system according to a forty-eighth aspect.

According to a fiftieth aspect of the present invention there is provided a method of managing a plurality of electrical power sources installed in an aircraft, the method comprising the steps of: receiving second flight information relating to a second flight, the second flight subsequent to a first flight; processing the second flight information to determine second flight usage requirements for one or more of the electrical power sources; and selectively managing the power supplied by each electrical power source in the first flight to facilitate use of the one or more of the electrical power sources in the second flight according to the second flight usage requirements.

According to a fifty-first aspect of the present invention there is provided a propulsion management system for an aircraft, the aircraft comprising an air propulsion system having one or more propellers, the propulsion management system being arranged to selectively: interact a transducer assembly with the one or more propellers to extract energy from the propellers; and/or control the air propulsion system to drive the one or more propellers in reverse in order to provide a reverse thrust.

In one example, the propulsion management system further comprises an energy accumulator for storing the extracted energy. In one example, the energy accumulator is comprised in a power supply for the air propulsion system of the aircraft. In one example, the energy accumulator is a battery, supercapacitor and/or flywheel. In one example, the transducer assembly is arranged to convert kinetic energy into electrical energy. In one example, the transducer assembly is arranged to control the rotational speed of the one or more propellers. In one example, the transducer assembly is arranged to inhibit motion of the one or more propellers. In one example, the propulsion management system is operated to increase aircraft drag during a descent. In one example, the propulsion management system is arranged to control the air propulsion system to drive propellers in reverse, and interact the transducer assembly with the one or more propellers to extract energy from the propellers, during a descent phase. In one example, the propulsion management system is arranged to both: control the air propulsion system to drive propellers in reverse, and interact the transducer assembly with one or more propellers to extract energy from the propellers, during a landing roll. In one example, the pitch of a fixed or variable pitch propeller is optimised to provide a desired or necessary level of energy extraction and/or reverse thrust. In one example, the air propulsion system comprises one or more electric motors drivingly connected to the one or more propellers. In one example, the propulsion management system is further selectively arranged to control the air propulsion system to drive the one or more propellers to provide a forward thrust, optionally wherein a first propeller is driven to provide a higher level of thrust than a second, driven, propeller.

According to a fifty-second aspect of the present invention there is provided an aircraft comprising an air propulsion system having one or more propellers and a propulsion management system according to the fifty-first aspect.

According to a fifty-third aspect of the present invention there is provided a method of propulsion management in an aircraft comprising an air propulsion system having one or more propellers, the method comprising the steps of selectively: interacting a transducer assembly with one or more propellers of the air propulsion system to extract energy from the propellers; and/or controlling the air propulsion system to drive the one or more propellers in reverse in order to provide a reverse thrust.

Although some exemplary embodiments of the present invention are shown and described, it will be appreciated by those skilled in the art that various changes and modifications might be made without departing from the scope of the invention, as defined in the appended claims.

Additionally, it will be appreciated that the various aspects and embodiments are closely related in terms of concept and technical implementation, and as a result various features of those aspects and embodiments are clearly combinable with one another, and/or may replace one another, unless such combination or replacement would be understood by the skilled person reading this disclosure to be mutually exclusive.

BRIEF DESCRIPTION OF THE DRAWINGS

For a better understanding of the invention, and to show how embodiments of the same may be carried into effect, reference will now be made, by way of example only, to the accompanying diagrammatic drawings in which:

Figure 1 shows an aircraft according to an embodiment of the present invention;

Figure 2 shows a general propulsion management system according to an embodiment of the present invention;

Figure 3 shows general methodology associated with a propulsion management system according to an embodiment of the present invention;

Figure 4 shows a graph of total instantaneous power supplied by the aircraft propulsion system versus time during a takeoff roll according to an embodiment of the present invention;

Figure 5 shows a graph of aggregate power during a takeoff roll for each of an aircraft comprising both a ground propulsion system and an air propulsion system according to an embodiment of the present invention, and an aircraft comprising only an air propulsion system; Figure 6 shows a graph of speed versus time during a takeoff roll for each of an aircraft comprising both a ground propulsion system and an air propulsion system according to an embodiment of the present invention, and an aircraft comprising only an air propulsion system; Figure 7 shows a takeoff trajectory for each of an aircraft comprising both a ground propulsion system and an air propulsion system according to an embodiment of the present invention, and an aircraft comprising only an air propulsion system;

Figure 8 shows a propulsion management system according to an embodiment of the present invention;

Figure 9 shows an alternative block diagram of the propulsion management system architecture. Figure 10 shows a graph of thrust versus speed for an aircraft comprising both ground and air propulsion systems according to an embodiment of the present invention; and Figure 11 shows a general propulsion management system according to an embodiment of the present invention;

Figure 12 shows general methodology associated with a propulsion management system according to an embodiment of the present invention;

Figure 13 shows plots of propeller advance ratio versus propeller efficiency; Figure 14 shows an energy recovery system provided in the aircraft according to an embodiment of the present invention;

Figure 15 shows a general energy recovery system according to an embodiment of the present invention;

Figure 16 shows general methodology associated with an energy recovery system according to an embodiment of the present invention;

Figure 17 shows a general energy recovery system according to an embodiment of the present invention;

Figure 18 shows general methodology associated with an energy recovery system according to an embodiment of the present invention;

Figure 19 shows a general propulsion management system according to an embodiment of the present invention;

Figure 20 shows general methodology associated with a propulsion management system according to an embodiment of the present invention;

Figure 21 shows a cooling system for cooling the battery of the aircraft propulsion system according to an embodiment of the present invention;

Figure 22 shows a general cooling system according to an embodiment of the present invention; Figure 23 shows general methodology associated with a cooling system according to an embodiment of the present invention;

Figure 24 shows a block diagram of an electrical power management system according to an embodiment of the present invention;

Figure 25 shows a block diagram of an alternative electrical power management system according to an embodiment of the present invention;

Figure 26 shows a general electrical power management system according to an embodiment of the present invention;

Figure 27 shows general methodology principles associated with an electrical power management system according to an embodiment of the present invention;

Figure 28 shows a propulsion system according to an embodiment of the present invention; Figure 29 shows general methodology principles associated with a propulsion system according to an embodiment of the present invention;

Figure 30 shows a schematic monitoring system according to an embodiment of the present invention;

Figure 31 shows general methodology principles associated with a monitoring system according to an embodiment of the present invention;

Figure 32 shows a schematic of an electrical power management system according to an embodiment of the present invention;

Figure 33 shows a diagram of an aircraft flight plan, energy storage device level and generator level, managed by an electrical power management system;

Figure 34 shows general methodology principles associated with an electrical power management system according to the present invention; Figure 35 shows a schematic of a power management system according to an embodiment of the present invention;

Figure 36 shows general methodology principles associated with a power management system according to an embodiment of the present invention;

Figure 37 shows a schematic propulsion management system according to an embodiment of the present invention;

Figure 38 shows general methodology principles associated with a propulsion management system according to an embodiment of the present invention;

Figure 39 shows a spar assembly according to an embodiment of the present invention;

Figure 40 shows a spar arrangement according to an embodiment of the present invention; Figure 41 shows general methodology principles associated with a spar arrangement according to an embodiment of the present invention;

Figure 42 shows a schematic fluid fuel monitoring system;

Figure 43 shows general methodology principles associated with a fluid fuel monitoring system;

Figure 44 shows a schematic power system structure;

Figure 45 shows general methodology principles associated with a power system structure; Figure 46 shows a schematic power system structure;

Figure 47 shows general methodology principles associated with a power system structure; and Figure 48 shows general methodology principle associated with a power system structure. DETAILED DESCRIPTION

Referring to Figure 1, an aircraft 1 is provided with landing gear comprising wheels 2. The aircraft comprises a propulsion system comprising a ground propulsion system 4 and an air propulsion system 6. The ground propulsion system 4 comprises ground propulsion system motors 8 arranged to drive the wheels 2. The air propulsion system 6 comprises propellers 10 and air propulsion system motors 12 arranged to drive the propellers 10. The ground propulsion system 4 and air propulsion system 6 share a power supply 3 comprising an energy accumulator, which in this exemplary embodiment, comprises a battery 14. The battery 14 is configured to supply power to the ground propulsion system motors 8 and to the air propulsion system motors 12.

In this exemplary embodiment, the aircraft 1 comprises body-mounted landing gear. Nevertheless, it will be appreciated by the person skilled in the art that alternative landing gear arrangements may also be employed. For example, the landing gear may be located on struts which extend from beneath the aircraft wings or nacelles.

Additionally, in this exemplary embodiment the air propulsion system 6 comprises four propellers 10. Nevertheless, it will be appreciated by the person skilled in the art that the air propulsion system 6 may comprise any suitable number of propellers 10 and associated air propulsion system motors 12. For example, the air propulsion system may comprise two, four or six propellers 10. In an exemplary embodiment, the power supply 3 further comprises an internal combustion engine 16 drivingly connected to the propellers 10. Whilst primarily provided as a backup power source, the internal combustion engine 16 can assist the propulsion system when high accelerations are demanded, or when battery recharging is preferable. In another embodiment, depending on specific implementations of the invention, the ground propulsion system may be driven by a combustion engine, and/or the air propulsion system may be driven by a combustion engine. As discussed below, however, at least partial electrical implementation may be advantageous. Referring back to Figure 1, a propulsion management system 20 is arranged to control the ground and air propulsion systems 4, 6. As will be described in greater detail herein, the propulsion management system 20 is arranged to control the ground propulsion system 4 to drive the wheels 2 during a takeoff roll. In a taxiing phase, the aircraft 1 is positioned on the runway ready for takeoff. The propulsion management system 20 is then employed for a takeoff roll. When thrust is demanded by the aircraft pilot, the propulsion management system 20 controls the ground propulsion system 4 to drive the wheels 2 by supplying power from the battery 14 to the ground propulsion system motors 10. The aircraft 1 is thus caused to accelerate for a first time period using the ground propulsion system 4 to increase the aircraft speed as it travels along the runway. Once a predetermined or desired speed is achieved by operation of the ground propulsion system 4, the propulsion management system 20 then controls the air propulsion system 6 to drive the propellers 10 by causing a supply of power to be directed from the battery 14 to the air propulsion system motors 12. The aircraft 1 is thus caused to accelerate for a second time period using the air propulsion system 6 to continue to increase the aircraft speed up to, and optionally above, an appropriate takeoff speed. The transition from ground propulsion drive to air propulsion drive is automatic and is handled by the system 20, that is, it does not require any particular or dedicated input from the pilot. Overlap between the aforementioned first and second time periods can be controlled to cause the aircraft 1 to perform a desired takeoff roll. For example, in one exemplary use of the propulsion management system 20, the system 20 causes the ground propulsion system 4 and the air propulsion system 6 to operate with no overlap in operational periods. This smooths out acceleration of the aircraft. In another exemplary use of the propulsion management system, the system causes the ground propulsion system 4 and the air propulsion system 6 to operate with an overlap so as to accelerate the aircraft at a faster rate. It is envisaged that the overlap could be the majority of, or all of, a takeoff roll, which may be particularly advantageous where rapid acceleration is required, for example where the aircraft is to perform a takeoff roll on a short runway. In Figure 2, a general propulsion management system according to an embodiment of the present invention is shown. The propulsion management system 20 is for an aircraft 1 comprising one or more wheels 2, and a ground propulsion system 4 arranged to drive one or more wheels 2 of the aircraft 1. The propulsion management system 20 is arranged to control the ground propulsion system 4 to drive one or more wheels 2 of the aircraft 1 during a takeoff roll, the takeoff roll being subsequent to a taxiing phase. In Figure 3, general methodology principles associated with a propulsion management system according to an embodiment of the present invention are shown. Step 100 comprises providing an aircraft 1 comprising one or more wheels 2 and a ground propulsion system 4 arranged to drive one or more wheels 2 of the aircraft 1. Step 102 comprises controlling the ground propulsion system 4 to drive the one or more wheels 2 of the aircraft 1 during a takeoff roll, the takeoff roll being subsequent to a taxiing phase.

Figure 4 shows a graph of total instantaneous power “TIP” supplied by both the ground and air propulsion systems versus time “t” during a takeoff roll. The ground propulsion system 4 is operational for a first time period 104. The air propulsion system 6 is operational for a second time period 106. As shown in the figure, there is an overlap period 108 between the operational periods, which is advantageous in providing rapid acceleration, and/or a seamless transition between periods, which may offer flexibility in technical implementation or provide an improved user (e.g. passenger) experience. A further advantage of the invention is succinctly shown in Figure 4. The ground propulsion system 4 is more efficient than the air propulsion system 6 at low speeds. The ground propulsion system 4 therefore requires less power to accelerate the aircraft from standstill, as is shown by the lower-level power line during the first time period 104. Once the pre-determined speed is achieved, the propulsion management system 20 causes a transition to the air propulsion system 6, and the air propulsion system 6 is operational for the second time period 106at a constant power level. The shaded area 110 indicates the power that would have been required had the air propulsion system 6 been operational for the entire takeoff roll, and thus indicates the power saved by using providing the propulsion management system 20 to control the ground propulsion system 4 to initially accelerate the aircraft.

Figure 5 shows aggregate power “AP” versus time “t” during a takeoff roll for an aircraft 1 having both ground and air propulsion systems 4, 6 (labelled 120), and an aircraft having only an air propulsion system (labelled 122). It has been found that the aggregate power required during a takeoff roll is lower when the ground propulsion system 4 is initially used to accelerate the aircraft 1 before transitioning to the air propulsion system 6, when compared with initiating the takeoff roll using only the air propulsion system 6. Such a system, according to embodiments of the invention, leads to a reduction in aggregate power of 20% for the same takeoff distance in at least one example. The same or better savings may be achieved, depending on how much use is made of the ground propulsion system, balanced against acceleration that might be too great for comfort of passengers or users.

Figure 6 shows speed “v” of an aircraft 1 having both ground and air propulsion systems 4, 6 (labelled 130), and an aircraft having only an air propulsion system (labelled 132) versus time “t”. The shaded region 134 indicates the range of possible aircraft lift-off speeds. The Figure shows that, in addition to the lower aggregate power, it is possible to achieve lift-off speed sooner using a system according to example embodiments This can lead to a 12.5% reduction in the distance travelled from the start of the takeoff roll to take off, in this example alone. This is advantageous where the aircraft is to take off from a short runway, and also in reducing noise pollution as the aircraft climbs to altitude sooner. The same or better savings may be achieved, depending on how much use is made of the ground propulsion system, balanced against acceleration that might be too great for comfort of passengers or users.

Figure 7 shows a graph of altitude “A” versus distance from start of the takeoff roll “d” and illustrates how operation of the ground propulsion system 4 can be advantageous in reducing noise pollution. In the field, noise pollution measurements are obtained by measuring the noise level at the ground, a fixed distance from the start of the takeoff roll. The reduction in takeoff distance facilitated by the propulsion management system 20 arranged to control the ground propulsion system 4 and the air propulsion system 6 allows the aircraft 1 to begin climbing to altitude sooner (as indicated by line 140) compared with an aircraft 1 having only an air propulsion system 6 (as indicated by line 142). The noise levels measured at a point 144 on the ground from the aircraft 1 following the solid line trajectory are therefore lower. This noise level might be reduced further, if any propellers of the air propulsion system are optimised for the climb phase, as opposed to needing to be optimised for the takeoff roll. The invention facilitates this optimisation, since the ground propulsion system means that the air propulsion system is not needed for the entire (and certainly initial period of) the takeoff roll.

Referring now to Figure 8, a ground propulsion system motor 8 is drivingly connected to each aircraft wheel 2. As a result, the ground propulsion system motors 8 can drive the aircraft wheels independently. A sensor assembly 22 is arranged in use to measure one or more of the wind speed, wheel speeds, travel path and trajectory of the aircraft (e.g. one or more situational properties) and the propulsion management system 20 is arranged to monitor the sensor assembly 22. Crosswinds can cause an aircraft to deviate from a desired takeoff travel path, or be urged towards such a deviation. If the sensor assembly 22 senses such deviation or urging during a takeoff roll, the propulsion management system 20 responds by causing the ground propulsion system to drive the wheels 2 at different speeds, by increasing or decreasing the motor torque as required, so as to return the aircraft 1 to the desired travel path.

Aircraft are often required to take off in, or following, adverse weather conditions. Snow, ice or standing water presents a considerable risk during a takeoff roll and will affect wheel traction. If during a takeoff roll the propulsion management system 20 monitoring the sensor assembly 22 detects a deviation in rotational speed of each the aircraft wheels 2, the propulsion management system 20 responds to apportion drive to the wheels 2 accordingly. For example, when the propulsion management system 20 becomes aware of a wheel having a rotational speed greater than the rotational speed of the other aircraft wheels by a threshold amount (e.g. a slip condition, or a condition where the drive is causing the aircraft to turn unintentionally through the wheel driving), the propulsion management system responds by causing the ground propulsion system 4 to reduce the torque provided to the faster rotating wheel, or perhaps increasing the torque to the slower rotating wheel. The former is more likely, as in the control of automobiles, to prevent or reduce wheelspin. Such a system operates in conjunction with the above described sensor assembly 22 for measuring the aircraft travel path. The propulsion management system 20 can thus vary the torque provided to the wheels 2 in dependence on both the travel path and the rotational speed of the wheels, to ensure that the aircraft remains on the desired takeoff travel path.

Referring back to Figure 8, the propulsion management system 20 is arranged to apportion power from the power supply 3 between the ground propulsion system 4 and the air propulsion system 6. The propulsion management system 20 does so based on input from a single thrust actuator (e.g. lever) 50, as in a conventional aircraft, and apportions power based on the level of actuation of the lever 50. Or, if an aircraft has one lever for each air propulsion engine, the invention envisages that same level being used for any one or more associated ground propulsion drivers. As a result, the pilot does not have to control the power apportionment manually, or at all. The apportionment is in the background and automatic. Additionally, the propulsion management system can be retrofitted to existing aircraft, and the cockpit layout need not be altered or adjusted to accommodate the propulsion management system 20. This system when employed during a takeoff roll is highly advantageous as the pilot may focus on other tasks than energy apportionment. The propulsion management system 20 comprises a processor 23, a memory 25, a communications unit 27 and a user interface 29. The user interface 29 allows information to be provided to the propulsion management system 20 and also allows information relating to the aircraft systems to be accessed by the user. Information can also be uploaded to the propulsion management system 20 via the communications unit 27. The propulsion management system 20 is configured to process and store flight plan information, which may be data relating to a past, present or future flight, in memory.

Figure 9 shows an alternative block diagram of the propulsion management system architecture. The propulsion management system 20 communicates with an operational information unit 62 to access flight information, including flight plans. The propulsion management system communicates with a power information unit 64 to access information relating to the status of a power supply, which includes the battery 14 and the internal combustion engine 16. The propulsion management system 20 is arranged to control the operation of the internal combustion engine 16. The propulsion management system 20 is also arranged to communicate with a flight control computer (FCC) 66, which communicate with a power management and distribution unit 68. The propulsion management system 20 is arranged to output information to a user interface 70.

As mentioned above, the propulsion management system 20 receives information from the power supply 3. Such information includes the energy potential of the power supply 3, which informs the pilot of the battery charge level and, where the aircraft comprises an internal combustion engine 16, the fuel level. On receiving this information, the propulsion management system controls the apportionment of power between the ground propulsion system 4 and the air propulsion system 6, and controls the supply of power from each power source, taking into account flight information such as: a. Acceleration profiles stored in memory which result in: the quietest possible takeoff, the shortest possible takeoff, minimum power usage for takeoff, the smoothest possible acceleration, suitable levels of torque being applied to the wheels to ensure traction with the runway surface is maintained; b. Deceleration profiles stored in memory which result in: the quietest possible landing, the shortest possible landing, minimum power usage to slow the aircraft, the smoothest possible deceleration, suitable levels of torque being applied to the wheels to ensure traction with the runway surface is maintained; c. Power source management requirements, including: preserving battery charge, using excess charge to make capacity for regeneration (including decelerating the aircraft using the energy recovery systems described in detail below), using excess fuel, battery temperature; d. Flight plan information, including: the distance to the destination, time to destination, turnaround time at destination; e. Airport data, including: runway length at takeoff airport or at destination airport, availability of charging ports at destination; or f. Situational conditions, including: wind speed, wheel speed, travel path and trajectory of the aircraft.

Examples of the propulsion management system 20 controlling energy apportionment include: a. where the quietest possible takeoff is required, the propulsion management system prioritises the supply of power to the ground propulsion system for as long as possible, and without apportioning power to the air propulsion system until necessary; b. where the shortest possible takeoff is required, the propulsion management system apportions power between the ground and air propulsion systems to produce maximum acceleration, in particular supplying maximum power to both systems; c. where minimum power usage is required, the propulsion management system apportions between the ground and air propulsion system to produce the most efficient takeoff roll; d. where smooth acceleration is required, the propulsion management system selects an appropriate acceleration profile, apportions power and controls the overlap between operation of the ground and air propulsion management system so that acceleration remains substantially constant throughout the takeoff roll; e. where maximum deceleration, or smooth deceleration, is required, the propulsion management system may control operation of the energy recovery systems described in detail below, apportioning power to the energy recovery systems in addition or instead of to the propulsion systems; f. where it is desirable to preserve battery charge, the propulsion management system causes power to be supplied from the internal combustion engine for the remainder of the flight, or until battery charge is restored or increased by operating one of the recovery systems or cooling system as described in detail below; and g. where it is desirable to minimise turnaround time at the destination, the propulsion management system apportions power to the air propulsion system during a landing roll to provide reverse thrust, and does not apportion power to the energy recovery system interacting with the wheels as described in detail below so as to minimise heating of the aircraft wheels. The propulsion management system 20 may do any or all of the above with reference to information stored in memory in relation to previous flights.

Figure 10 shows a graph of net thrust “T”, produced by energy apportionment, versus speed “v” for an aircraft comprising both ground and air propulsion systems. The upper region 300shows the thrust available from the ground propulsion system 4. The lower region 302 shows the thrust available from the air propulsion system 6. The dashed line 304 shows the net thrust possible whilst adhering to airport noise restrictions. The propulsion management system operates to apportion energy between the ground propulsion system 4 and the air propulsion system 6 so as to adhere to the noise restriction.

In Figure 11, a general propulsion management system according to an embodiment of the present invention is shown. The propulsion management system 20 is for an aircraft 1, the aircraft 1 comprising one or more wheels 2, a power supply 3, a ground propulsion system 4 arranged to drive one or more wheels 2 of the aircraft 1 and an air propulsion system 6. The propulsion management system 20 is arranged to apportion power from the power supply 3 between the ground propulsion system 4 and the air propulsion system 6. In Figure 12, general methodology principles associated with a propulsion management system according to an embodiment of the present invention are shown. Step 150 comprises providing an aircraft 1 comprising one or more wheels 2, a power supply 3, a ground propulsion system 4 arranged to drive one or more wheels 2 of the aircraft 1, and an air propulsion system 6. Step 152 comprises apportioning power from the power supply 3 between the ground propulsion system 4 and the air propulsion system 6. Figure 13 shows plots of propeller advance ratio “J” versus propeller efficiency “h”. Advance ratio is the ratio of the aircraft forward speed to the rotational speed of the propellers. The plots illustrate the relationship between advance ratio and propeller efficiency for various propeller pitch angles “b”. Pitch angle is an important consideration in propeller design because of this relationship. At low advance ratio J (that is, in and around the region 220), which is typically at low aircraft forward speeds, the propeller efficiency h is low, and a significant proportion of power supplied to the air propulsion system 6 is wasted as noise. During a takeoff roll, the propulsion management system 20 causes the aircraft 1 to accelerate using the ground propulsion system 4 and the aircraft forward speed increases. When the propellers 10 are subsequently activated, they are in a region of increased efficiency (that is, in and around the region 222). This means that when the propellers 10 are activated, they produce considerably less noise than if they were operating at low aircraft speed in and around region 220, for example, if the propellers 10 were used to accelerate the aircraft 1 from stationary.

Variable pitch propellers are typically employed to maintain high propeller efficiency across a wide range of propeller advance ratios. This, of course, requires a degree of mechanical and control complexity. Nevertheless, in an exemplary embodiment, the aircraft propellers 10 are fixed pitch propellers. The fixed pitch propellers 10 are optimised for the climb phase of flight. By providing the ground propulsion system, the propeller need not be optimised for low advance ratio, enabling simpler fixed pitch propellers to be provided. Optimising the propellers 10 for the climb phase means that the propeller pitch is set so as to produce minimum noise during climb by operating most efficiently during the range of speeds achieved during the climb phase. This might mean that the propellers, and/or any control systems, may be simpler, cheaper, and even lighter.

Figure 14 shows an energy recovery system 30 provided in the aircraft 1. The energy recovery system 30 is connected to the ground propulsion system motors 10. The energy recovery system 30 causes the electric motors 10 to operate as generators, absorbing or generally using rotational kinetic energy from the aircraft wheels 2 and generating electricity which is fed to the battery 14 where it is stored. This could happen in flight, or on the ground. Advantageously, the energy recovery system 30 is operated to provide braking to the aircraft 1 during a landing roll. By adjusting the load on the motors 10, the energy recovery system 30 controls the rate of deceleration of the aircraft wheels and thus the aircraft. Using the energy recovery system 30 to decelerate the aircraft instead of or alongside friction brakes also reduces the brake temperature, which decreases turnaround time of the aircraft at a destination airfield or airport, since the brakes need less time to return to a safe operating temperature. In Figure 15, a general energy recovery system according to an embodiment of the present invention is shown. The energy recovery system 30 is for an aircraft. The energy recovery system comprises a transducer assembly 11 arranged to selectively interact with one or more wheels 2 of the aircraft to extract energy from the wheels 2. The energy recovery system 30 further comprises an energy accumulator 3 for storing the extracted energy. In Figure 16, general methodology principles associated with an energy recovery system according to an embodiment of the present invention are shown. Step 230 comprises interacting a transducer assembly 11 with one or more wheels of the aircraft to extract energy from the wheels. Step 232 comprises storing the extracted energy in an energy accumulator 3.

In this exemplary embodiment, the aircraft is provided with an energy recovery system 30 in conjunction with the propulsion management system 20, and the energy stored in the battery 14 of the energy recovery system is used to power the ground propulsion system 4 during a subsequent takeoff roll. Whilst in this exemplary embodiment Li-Ion batteries are employed, supercapacitors and flywheels are also a feasible alternative energy storage means. Again, this might reduce turnaround time of the aircraft at a destination airfield or airport, since there is no (or a lesser) need to recharge the battery. The energy recovery system 30 is also connected to the air propulsion system motors 12. The energy recovery system 30 causes the two outermost propeller motors to selectively interact with the propellers to operate as generators, absorbing rotational kinetic energy from the propellers and generating electricity which is fed to the battery 14 where it is stored. This interaction is employed during a descent phase of flight. When the energy recovery system 30 is in interaction with the two outermost propellers, the propulsion management system 20 controls the air propulsion system 6 not to provide drive to those propellers. Instead, the propellers are allowed to rotate freely as they push through the air. The motors variably inhibit motion of the propellers to control the rotational speed of the propellers. By adjusting the load on the motors (e.g. by the generation of electricity), the energy recovery system 30 controls the rate of deceleration of the aircraft 1 by increasing the drag produced by the two peripheral propellers. The ability to increase the drag facilitates a steep angle of aircraft descent. This is highly advantageous in reducing aircraft noise pollution, and possibly allowing more locations to be used as airfields or airports. Adjusting the load on the motors also provides greater or lesser rate of battery charging. Whilst the two outermost propellers are not driven, the innermost two propellers may be driven (perhaps more than usual) so as to maintain a constant air speed. In Figure 17, a general energy recovery system according to an embodiment of the present invention is shown. The energy recovery system 30 is for an aircraft. The energy recovery system comprises a transducer assembly 11 arranged to selectively interact with one or more propellers 10 of a propulsion system 6 of the aircraft to extract energy from the propellers. In Figure 18, general methodology principles associated with an energy recovery system according to an embodiment of the present invention are shown. Step 240 comprises interacting a transducer assembly 11 with one or more propellers 10 of a propulsion system 6 of the aircraft. Step 242 comprises extracting energy from the propellers 10.

The propulsion management system 20 is further arranged to control the air propulsion system 6 to drive the propellers 10 in reverse to provide a reverse thrust. This is particularly advantageous in decelerating the aircraft during a landing roll. The propulsion management system comprises an energy recovery system 30 as described above, and the energy recovery system is connected to the wheels 2 to provide braking effort during a landing roll. As a result, energy recovered during the landing roll is directed to the air propulsion system 6 and used to drive the propellers in reverse to provide a reverse thrust. In an exemplary embodiment, the aircraft comprises such a propulsion management system for driving propellers in reverse and variable pitch propellers. The variable pitch propellers can be adjusted on, or prior to, landing to create a maximum reverse thrust when the propellers are operated in reverse. In Figure 19, a general propulsion management system according to an embodiment of the present invention is shown. The propulsion management system 20 is for an aircraft comprises an air propulsion system 6 having one or more propellers 10. The propulsion management system 20 is arranged to control the air propulsion system 6 to drive the one or more propellers 10 in reverse in order to provide a reverse thrust. In Figure 20, general methodology principles associated with a propulsion management system 20 according to an embodiment of the present invention are shown. Step 250 comprises operating the propellers in reverse to provide reverse thrust.

Figure 21 shows a cooling system 40 for cooling the battery 14 of the aircraft propulsion system. The cooling system 40 is provided in the belly fairing of the aircraft and comprises an inlet 42, of the type developed by the National Advisory Committee for Aeronautics (NACA), opening into an airflow duct 44 which extends through the belly fairing to a NACA outlet 46. The battery 14 is provided with a battery temperature sensor 16. The battery 14 is located upstream in the airflow duct 44. The airflow duct 44 directs cool air from outside the aircraft across the battery 14 to provide cooling. The battery is located close to a outer mould line 48 of the aircraft 1, and so some external skin conductive cooling also takes place. The landing gear is mounted in the belly fairing. The landing gear is retractable into the airflow duct 44, downstream of the battery 14. When the landing gear is in a retracted position, a lower portion of each wheel 2 projects through a lower aperture in the duct 44 and out through the outer mould line 48 of the belly fairing. The airflow duct 44 directs the now heated air towards and past the lower portion of each wheel 2 of the landing gear. The airflow duct 44 connects to the outlet 46 downstream of the wheels 2. The outlet 46 is in a location of low pressure relative to the inlet 42 to encourage air flow through the duct 44. As the lower portion of each wheel 2 projects out through the outer mould line 48 of the belly fairing, in flight airflow around the outside of the aircraft 1 causes the wheels 2 to rotate. Wheel hubs 70 comprise vanes located in or projecting from the side of the wheels to provide an increased surface area for contact with the airflow. The vanes are configured such that the airflow through the airflow duct passing each wheel encourages the wheels to rotate in the same direction as the rotation caused by the outside airflow. It will be appreciated that ‘internal’ airflow (that is, not outside airflow) across and past the battery, and to and past the wheels, may alternatively or additionally cause the wheels to rotate in this way. That is, the wheels need not extend outside of the mould line. The rotating wheels 2 create suction through the duct 44, which draws air through the duct 44 at a greater rate. The rate of airflow through the airflow duct 44 is thus a function of the rotational velocity of the wheels 2. By increasing the rate of airflow, the battery 14 can be cooled more rapidly, and by decreasing the rate of airflow, the battery 14 can be cooled less rapidly, the temperature maintained, or the battery allowed to heat. The rotational velocity of the wheels is controlled by connecting the cooling system to an energy recovery system of the kind as described above, where energy can be extracted from wheel rotation. The temperature of the battery is monitored, and an appropriate load is placed on the motors 8 to increase or decrease the rotational velocity of the wheels, to thereby increase or decrease the rate of airflow through the duct 44 as necessary or desired. The configuration of the cooling system 40 is such that a greater motor load produces an increased rate of battery charging and a reduced motor load produces a reduced rate of battery charging. In Figure 22, a general cooling system according to an embodiment of the present invention is shown. The cooling system 40 is for cooling a battery in an aircraft using a flow of gas. The cooling system comprises an inlet 42, an outlet 46, a gas flow path 44 extending between the inlet 42 and outlet 46, and at least proximal to the battery 14 to cool the battery 14. A transducer 8 is provided in the flow path 44, the transducer 8 being arranged to convert energy of the gas flow into electrical energy. In Figure 23, general methodology principles associated with a cooling system according to an embodiment of the present invention are shown. Step 260 comprises directing a flow of gas proximal to a battery to cool the battery 14. Step 262 comprises converting energy of the gas flow into electrical energy using a transducer 8.

In addition to, and alongside, the ground and air propulsion systems 4, 6, the propulsion management system 20 is also configured to control the cooling system 40 and energy recovery system 30.

Referring back to Figure 1, a fuel tank 500 is installed in each aircraft wing 504. Electrical power sources in the form of electric generators, in particular turbogenerators 506, are received in the aircraft nacelles 508. In this exemplary embodiment, one turbogenerator 506 is received in each inboard nacelle 508. The fuel tanks 500 are arranged to provide a supply of fuel to the turbogenerators 506, for use in generating electrical power. Whilst the turbogenerators 506 and fuel tanks 500 will be described as primarily forming part of the air propulsion system 6, it will be understood that the turbogenerators 506 may function to supply electrical power to the ground propulsion system 4, or any other system or component in the aircraft. It will be understood that the term electric generator can also be used to a fuel cell, which may be appropriately implemented in the present invention. For the avoidance of doubt, the electric generator may be comprised in an energy recovery system installed in the aircraft.

Referring to Figure 24, the aircraft 1 further comprises an electrical power management system 508 comprising a plurality of power management modules 510. The plurality of power management modules 510 are arranged to manage electrical power to aircraft components from electrical power sources 512 installed in the aircraft 1. The power management modules 510 are arranged to manage power from electrical power sources 512 which may include: the turbogenerators 506, power supply 3 comprising battery 14 (which may be a pack of batteries), energy storage devices (including batteries, which may be structural batteries,, supercapacitors and flywheels), photovoltaic panels (not shown), or the energy recovery system 30 including turbines. Batteries 14 and energy storage devices can be charged using renewable sources. Turbogenerators are advantageous as they have high power density. The aircraft components to which power is supplied may form part of the aircraft propulsion systems, including the ground propulsion system 4, or air propulsion system 6, and may be primarily motors comprised in these systems.

The architecture of an electrical power management system 508 installed in the aircraft 1 is described with reference to Figure 24. Each of the plurality of power management modules 510 is arranged to manage power from at least one of the electrical power sources 512 to the components 514. Such a construction provides an advantageous level of redundancy, as power can still be managed from the power sources 512 in the event of failure of one, or more if appropriate, of the power management modules 510. In this exemplary embodiment, the electrical power sources 512 comprise turbogenerators 506, a battery pack 14, and the energy recovery system 30, and the power management modules 510 comprise a first power management module 510a and a second power management module 510b. That is, the plurality of electrical power sources 512 comprise a plurality of power sources of different types (battery pack 14, turbogenerators 506 and energy recovery system 30) and also, as the battery pack comprises a plurality of batteries, a plurality of power sources of the same type. In this exemplary embodiment, the components 514 are motors 8, 12 arranged to drive wheels 2 and propellers 10 of the aircraft 1. Power supplied by the turbogenerator 506 and battery pack 14 is managed by the first power management module 510a, and power supplied by the battery pack 14 and energy recovery system 30 is managed by the second power management module 510b. Thus, there is an advantageous level of redundancy provided by the plurality of power management modules, as failure of either the first or second modules 510a, 510b does not result in a complete loss of power supplied to the motors 8, 12.

An alternative architecture is shown in Figure 25. Each of the plurality of power management modules 510 is arranged to manage power from one electrical power source 512. As a result, each power source 512 has a dedicated power management module 510c, 510d, 510e for managing the power from that power source only during normal operation. Normal operation may include where all power management modules 510 are functioning correctly. Additionally, the power management modules 510 are interconnected (shown in dashed lines). In normal operation, despite being interconnected, the non- dedicated power management module does not manage power from any other electrical power source. However, in the event of a failure of one of the dedicated power management modules, or some other functional change in system setup, another one of the power management modules can function as a backup power management module for at least one other electrical power source. In such an event, that power management module is a non-dedicated power management module. Again, advantageously, such a construction provides a level of redundancy to the system as failure of the dedicated power management modules 510c, 510d, 510e does not result in a loss of power from the corresponding power source 506, 14, 30.

In both architectures of the electrical power management system 508, whilst not being essential to the invention, each of the power management modules 510 is arranged to manage power from a plurality of the electrical power sources 512, whether during normal operation or as a backup. Such functionality may be by virtue of the electrical power sources 512 being configured to be managed by multiple power management modules 510a, 510b, or by virtue of interconnection of power management modules 510c, 510d, 510e. An advantageous level of redundancy or interoperability is thus provided, ensuring safe and reliable aircraft operation. Even without interconnection, each power source 512 having a dedicated power management module 510c, 510d, 510e for managing the power from that power source during normal operation may also be advantageous, so that one failure of one module does not cause an entire system failure.

The electrical power management system 508 is arranged to manage electrical power from the power sources to the components 514, which may comprise motors 8, 12, in dependence upon a target energy level of one or more of the electrical power sources 512. That is, in certain situations it may be desirable to reduce usage of the battery 14, such that a target battery energy level is obtained which is suitable for use in a later phase of aircraft flight. It is often advantageous to facilitate aircraft operation using battery power alone during landing and/or subsequent takeoff, as battery operation is quiet when compared with turbogenerator operation. To achieve this, the power management modules 510, or appropriate dedicated module, may reduce power drawn from the battery 14, and instead look to increase electrical power provided by the turbogenerators 506 and energy recovery systems 30. If necessary, the electrical power management system 508 will also manage power from the turbogenerators 506 and/or energy recovery systems 30 to provide a power supply to the battery 14 for charging, in order to achieve the target battery energy level. That is, it is intended that the scope of the invention extends to the components 514 being one or more energy storage devices, which may be the battery pack 14 comprised in the plurality of electrical power sources 512. Supplying power to the battery in anticipation of subsequent use is highly advantageous in ensuring reliability and allows a reduction in noise pollution and emissions. The energy storage devices, such as battery packs, may be comprised in the aircraft propulsion systems 4, 6, and arranged to provide power to motors 8, 12.

The electrical power management system 508 further comprises environmental sensors 516 for monitoring environmental conditions of, and in the region of, the electrical power sources 512. In one exemplary embodiment, each electrical power source 512 may be provided with one sensor 516a for monitoring the electrical power source temperature, such as an external or internal temperature of the battery pack 14, and one sensor 516b for monitoring the ambient air temperature in the region of the electrical power source 512, such as the ambient air temperature in the region of the battery pack 14. The electrical power management system 508 is arranged to manage electrical power from the battery pack 14 in dependence upon the environmental condition or conditions. Monitoring the temperature of the power source 512 and the ambient air temperature in the region of the electrical power source 512 is important and advantageous in predicting general functionality (e.g. behaviour, degradation, and so on) of the power source 512 and allows power to be managed based on the environmental condition, which protects the power source 512 to ensure safe and reliable aircraft operation.

A general electrical power management system 508 is shown in Figure 26. The electrical power management system 508 is for an aircraft, the aircraft comprising a plurality of electrical power sources 512 for powering components in the aircraft, the power management system comprising a plurality of power management modules 510 arranged to manage electrical power from the electrical power sources 512 to the components, wherein each of the plurality of power management modules 510 is arranged to manage power from at least one of the electrical power sources 512.

In Figure 27, general methodology principles associated with an electrical power management system according to the present invention are shown. The method is for managing electrical power in an aircraft, the aircraft comprising a plurality of electrical power sources for powering components in the aircraft. Step 520 comprises providing a power management system comprising a plurality of power management modules arranged to manage electrical power from the electrical power sources to the components. Step 522 comprises arranging the power management modules so that each of the power management modules is arranged to manage power from at least one of the electrical power sources.

Turning now to Figure 28, an aircraft propulsion system 524 is shown, the propulsion system 524 comprising a plurality of electrically powered propulsion apparatus 528 and a plurality of electric generators 526 for providing power to the propulsion apparatus 528. In this exemplary embodiment, the propulsion system 524 is the air propulsion system 6 described above. The electric generators 526 are turbogenerators, which may be liquid or gas fuelled. Such a construction is advantageous as a plurality of electric generators 526 ensures a level of redundancy in the event of failure of one of the generators, allowing the aircraft to continue to operate. The possibility for “single point failure” is thus mitigated, and so the aircraft may continue to function using generator power.

Each electrically powered propulsion apparatus 528 comprises an electric motor 12 for powering the aircraft propellers 10. Each propulsion apparatus 528 is powered by one or more of the electric generators 526. Each propulsion apparatus 528, or subset of a total number of propulsion apparatus (e.g. those located on either side of the aircraft), has a dedicated electric generator 526a, 526b for providing power only to that propulsion apparatus 528a, 528b during normal operation. Advantageously, this provides a level of redundancy and reduces risk of a “single-point failure”, which would lead to the aircraft 1 being inoperative in the event of failure of one of the generators 526a, 526b. In an exemplary embodiment, at least one of the electric generators 526 is arranged to function as a backup generator for one other electrically powered propulsion apparatus in the event of failure of one of the dedicated electric generators, thereby to function as a non-dedicated generator. Advantageously, this further increases safety and reliability of aircraft function.

In an exemplary embodiment, each propulsion apparatus 528, or subset of a total number of propulsion apparatus (e.g. those located on either side of the aircraft), further comprises one or more energy storage devices 530, for example battery packs 14, arranged to provide power to the motor 12 in addition to that provided by the generator 526. In some circumstances, it may be advantageous to provide power from the generators 526 to the energy storage devices 530 for charging. For example, where the energy storage devices 530 are arranged to provide power to the motors 12, it may be advantageous to provide power to the energy storage devices 530 in anticipation of later use of the energy storage devices 530 to power the motors 12. To do so, the electric generators 526 are further arranged to provide power to the energy storage devices 530. The energy storage devices 530 are arranged for subsequent use in providing power to the electrically powered propulsion apparatus 528. The electric generators 526 are arranged to provide power to the energy storage devices 530 in dependence upon a target energy level of the energy storage device 530. In particular, the target energy level may be an energy level which allows the aircraft 1 to be powered for a part of the flight using only the energy storage devices 530. Such an arrangement is advantageous in reducing aircraft turnaround time, as it may prevent the need to refuel the aircraft 1. Such an arrangement may be further advantageous in reducing overhead noise during takeoff and landing, as operation of energy storage devices 530 is quieter than turbogenerator 526 operation.

The electric generators 526 are located either side of the fuselage 535. This is practical where the generators 526 are arranged to be dedicated power supplies to motors 12 when the motors are positioned either side of the fuselage, but also advantageously facilitates a reduction in the length of high voltage electrical cabling which must be used to provide power to the aircraft motors 12. This reduction in length could reduce cost, weight, heating, design complexity, or improve safety.

The electric generators 526 are housed in a nacelle 508 of the aircraft 1. In an exemplary embodiment, each electric generator 526a, 526b is housed in a nacelle 508. Housing the generators in the nacelles is advantageous for cooling purposes, as a flow of cooling air can be directed into the nacelle and over the generators 526a, 526b for cooling. It is also advantageous for cabling routing purposes. It is typically not possible to house generators in a nacelle of a gas turbine powered aircraft, as the gas turbine engine, including associated conduits and cabling, takes up the entire volume of the aircraft nacelle. The present construction enables an electric generator and motor to both be housed in a nacelle, which has advantages in relation to component cooling, maintenance and power cabling routing.

Each electrically powered propulsion apparatus 528a, 528b is arranged to operate independently. In the event of failure of one of the electrically powered propulsion apparatus 528a, 528b, each other apparatus may continue to operate. This further provides an advantageous level of redundancy to the system. The generators 526a, 526b may be interconnected such that in the event of failure of one of the motors, the generators may supply power to another functioning motor, and in the event of failure of one of the generators, the other generators may provide a backup power supply to the motor that lost power supply.

In Figure 29, general methodology principles associated with a propulsion system according to the present invention are shown. Step 532 comprises providing a propulsion system comprising a plurality of electrically powered propulsion apparatus, and a plurality of electric generators. Step 534 comprises generating electrical power using the electric generator and providing the power to the propulsion apparatus.

Referring to Figure 30, the aircraft 1 further comprises a monitoring system 544 for monitoring an electrical power source 536 in the aircraft. The monitoring system comprises at least one electrical power source sensor 538 arranged to monitor an operating characteristic of the electrical power source 536. In this exemplary embodiment, the electrical power source 536 may be an energy storage device, such as a battery pack 14, and/or an electrical generator, such as a turbogenerator or fuel cell. The monitoring system further comprises at least one environment sensor 540 arranged to monitor an environmental condition of, or in the region of, the electrical power source 536. A processing module 542 is arranged to receive the operating characteristic from the electrical power source sensor 538 and receive the environmental condition from the environment sensor 540 and process the information to determine or predict past, present, or in particular future operating characteristics of the electrical power source 536. Such a system is advantageous as operating characteristics can depend strongly on environmental characteristics. It is also important to monitor component function and predict functionality (e.g. general behaviour or degradation) in order to ensure safe and reliable aircraft operation. The monitoring system 544 facilitates this.

The electrical power source sensor 538 (which may be an arrangement of multiple sensors) is configured to provide operating characteristics which are indicative of, represent, or are themselves one or more of the: current rate of charge, current rate of discharge, maximum possible rate of charge or discharge, power source capacity (which may be maximum capacity), charge/discharge cycle (typically used to specify the expected life of the battery), cycle life (number of complete charge/discharge cycles before its capacity falls under 80 percent of its original capacity) and/or power density (power per unit volume of power source). It is advantageous to obtain information regarding these operating characteristics, as the health of the electrical power sources 536 can be determined from such information. The environment sensor 540 (which may be an arrangement of multiple sensors) is configured to provide environmental conditions which are indicative of, represent, or are themselves one or more of the: component temperature, ambient temperature, humidity and/or pressure. It is advantageous to obtain information relating to these environmental conditions, as these can have a significant impact on the operating characteristics listed above, and thus impact on the reliability of the aircraft function. The sensors 538, 540 may be provided on the component 536, which in this exemplary embodiment is the electrical power source, allowing recording of environmental characteristics of the component or a housing of the component, which will have a considerable effect on the component function. However, it is also important and advantageous to monitor conditions in the region of the component, which will influence the function of the component 536.

In an exemplary embodiment, the sensors 538, 540 are wireless sensors and the processing module 542 is arranged to receive information wirelessly from the wireless sensors. The processing module 542 can thus advantageously be provided remotely from the sensors 538, 540, allowing data collection from all sensors at a central point. The processing module 542 is arranged to process the information from the sensors 538, 540 by comparing the operating characteristic and environmental condition with information stored in memory, which may involve referring to look-up tables or pre determined performance data or graphs.

The aircraft flight path may have an indirect effect on component function. For example, during an ascent to altitude, the ambient air temperature is likely to reduce, whereas during a descent, the ambient air temperature is likely to increase. The processing module 542 is arranged to receive flight information relating to the aircraft flight path. The monitoring system 544 can use this information to determine and/or predict future operating characteristics of the electrical power source 536. To do so, the processing module 542 can be arranged to process the flight information alongside the operating characteristic and environmental condition information. This information may also be compared with the information stored in memory. The flight information comprises information relating to present and/or future details of the flight path, including: climb profile and rate, descent profile and rate, altitude, time to a target point in the flight and/or distance to a target point in the flight. Additionally, flight information may comprise landing conditions and/or infrastructure at the destination. The processing module 542 is arranged to manage power from the electrical power source 536 in dependence on information processed by the processing module 542. The processing module can be arranged to preferentially use a particular power source 536 in dependence on information processed by the processing module. For example, if the processing module determines that the electrical power source 536 temperature is above a threshold, the processing module may cause power to be demanded from an alternative power source, so that the electrical power source 536 may cool down. Advantageous aircraft safety benefits are thus obtained.

In Figure 31, general methodology principles associated with a monitoring system according to the present invention are shown. Step 546 comprises providing at least one electrical power source sensor arranged to monitor an operating characteristic of the electrical power source. Step 548 comprises providing at least one environment sensor arranged to monitor an environmental condition of, or in the region of, the electrical power source. Step 550 comprises receiving the operating characteristic and the environmental condition from the sensors at a processing module and processing the information to determine or predict future operating characteristics of the electrical power source.

Turning now to Figure 32, the aircraft further comprises an electrical power management system 552, for managing the level of an electrical power source 554 installed in an aircraft. The system 552 comprises a receiver 556 arranged to receive second flight information relating to a second flight, the second flight being subsequent to a first flight. A processor 558 is arranged to process the second flight information provided to it by the receiver 556. The processor 558 processes the information to determine a target level of the electrical power source 554 that is or would be suitable for use in the second flight. This could be suitable in terms of being non-zero, or being useful for use in the second flight, or being at a level in the second flight that means that no charging or supplemental power is required to perform a certain function. The system is arranged to manage the level of the electrical power source 554 during the first flight to achieve the target level at a target time in the first flight.

The “level” of the electrical power source 554 may be a charge level or a level of energy potential, depending on the type of electrical power source employed. To manage said level of the electrical power source 554, the system 552 can be arranged to manage a usage profile of the electrical power source 554 to achieve the target level at the target time. The second flight information received and processed comprises information relating to usage requirements of the electrical power source 554 during, or intended to be suitable for use during, the second flight. As alluded to above, it may be particularly important to ensure that the level of the electrical power source 554 is suitable for use during at least part of, if not all of, a takeoff phase and/or initial climb of the second flight. This is because the takeoff phase and initial climb create substantial noise pollution in the vicinity of the airport. Using electrical power stored in an electrical power source to power the aircraft motors is quieter than generating electrical power using electrical generators. For this reason, the system is configured to obtain a target level wherein at least the take-off roll (that is, the journey along the runway) can be performed using solely the electrical power source 554, with no assistance from an electrical generator. It is highly advantageous to manage the level of the electrical power source to facilitate use of the electrical power source to thereby minimise noise pollution. Moreover, use of the electrical power source during the takeoff phase can be advantageous in reducing air pollution close to the ground. Second flight information may relate to constraints at future destinations, for example noise restrictions, runway length limits, en-route weather and traffic. Knowledge of these constraints advantageously allows the system 552 to manage power to ensure the desired takeoff or cruise can be performed. Moreover, knowledge of the constraints advantageously allows the system 552 to manage power to ensure the desired descent can be performed, which may preferentially use the electrical power source 536 so that overhead noise is minimised. Additionally and advantageously electrical energy used during the takeoff phase can be recovered during other flight phases and directed to the electrical power source 536, for example the battery pack, for use in charging.

In an exemplary embodiment, the target time in the first flight is the end of the first flight. The first flight is the current flight, and the second flight is the next flight. By ensuring that the target level is obtained at the end of the current flight, the electrical power sources 554 are advantageously ready for the next flight. Turnaround time is thus minimised as time need not be spent charging or replacing electrical power sources 554. The level of the electrical power source 554 can be managed by managing power provided to the electrical power source by one or more of: energy recovery systems, motors, turbogenerators and/or renewable energy systems, such as solar panels or wind turbines, provided on the aircraft.

In Figure 33, altitude 560 versus distance 562 during an exemplary aircraft flight is shown above a graph of power source characteristics for that same flight. The upper lines 564, 566, 568 in the lower graph indicate battery charge level throughout the duration of the flight, and the lower line 570 is the power generated by the electric generator throughout the duration of the flight. Initially, the initial taxi and take-off is performed solely using battery power for period 572. Battery power is used for an initial climb period, advantageously due to low noise output. Following takeoff, and after an initial climb to altitude, the electric generator is used to assist in powering the aircraft propulsion system. The electric generator is used as a power source for the remainder of the climb and cruise phase of the mission, as shown by the increase in power generated by the generator in 570. Changes in altitude during cruise are typically actuated using the generator. As can be seen in the figure and from the central line 566, the battery level is substantially constant during these phases. However, as indicated by the upper increasing line 568, the battery charge level may increase by directing power from the electrical generator to the battery. Alternatively, the battery charge level may decrease as electrical power is drawn from the battery and used to power aircraft motors and systems, as indicated by line 564. Battery charge level may also fluctuate during to temperature increases and/or decrease throughout the flight.

As the aircraft descends during period 574, energy recovery systems installed in the aircraft can recover electrical energy. The electrical power management system 552 is arranged to manage the level of the electrical power source 554 during the flight by increasing the level of the battery from an initial level at the start of the descent phase to a final, higher, level at the end of the descent phase. The charge level is increased in preparation for final approach, and also in preparation for subsequent (i.e. second) flight takeoff conditions, which may require a short but quiet takeoff. Finally, as the aircraft completes final descent and approach during period 576, battery power is used to power the aircraft propulsion system, thus facilitating a quiet final descent and landing at the destination, thereby minimising noise pollution and allowing the aircraft to operate during flight curfew hours.

In this example, the battery charge level at the end of the flight is substantially equal to the level at the beginning of the flight. Such operation can thus be referred to as “battery neutral operation”. Advantageously, the final charge level of the battery is at a level such that the aircraft can perform the takeoff phase of the next flight solely using battery power, or at least significantly contribute to this phase (e.g. so that little or no charging, or related, is required). Importantly, the system ensures that during all phases of flight, the aircraft is able to travel a certain distance, perform in-flight waiting, landing phase and taxi. The system also ensures that the battery charge level does not fall below a minimum level.

In Figure 34, general methodology associated with an electrical power management system is shown. The method is a method of managing the level of an electrical power source installed in an aircraft during a first flight. Step 578 comprises receiving second flight information relating to a second, subsequent, flight. Step 580 comprises processing the second flight information to determine a target level of the electrical power source suitable for use in the second flight. Step 582 comprises managing the level of the electrical power source during the first flight to achieve the target level at a target time in the first flight.

Turning now to Figure 35, a power management system 584 is shown. The power management system 584 is arranged to manage a plurality of electrical power sources 586 installed in an aircraft. The power management system comprises a receiver 590 arranged to receive second flight information relating to a second flight subsequent to a first flight. The power management system 584 further comprises a processor 592 arranged to receive the second flight information from the receiver 590. The processor 592 processes the second flight information to determine second flight usage requirements for one or more of the electrical power sources 586. The power management system 584 is arranged to selectively manage the power supplied by each electrical power source 588, 594 in the first flight to facilitate use of the one or more of the electrical power sources in the second flight according to the second flight usage requirements. Such a construction allows all power sources to be managed by a single power management system 584, ensuring optimal distribution of power to facilitate safe and reliable aircraft function, and also to allow power to be drawn from power sources appropriately, depending on the second flight information. Selectively managing each power source 588, 594 in this manner advantageously ensures that all power sources operate reliably, and that power source degradation and damage is mitigated. Selective management also ensures that the one or more power sources for which usage requirements have been determined are suitable for use when desired or required. Selective management may comprise selectively choosing to use or not use each power source, which may be autonomously controlled by the system 584.

In response to determining second flight usage requirements for a first one of the electrical power sources 586, the power management system is arranged to preferentially use a second one of the electrical power sources 586 over the first one of the electrical power sources 586 in the first flight. That is, the first one of the electrical power sources may be a battery pack 588, or generally battery 14, and the second one of the electrical power sources may be an electric generator 594, or generally electric generator 506. The power management system 584 may determine that it is of interest to preserve battery power, and thus will preference use of the electric generator 506, and thus use the fluid fuel supply from the fuel tank 500, over use of the battery. Advantageously, this ensures that battery reserves are suitable for safe and reliable use during the second flight, which may be during the takeoff phase, thus enabling battery powered takeoff.

The processor 592 is arranged to process the second flight information to determine a second flight usage profile for one or more of the electrical power sources 586. The power management system is arranged to selectively manage the power supplied by each electrical power source in the first flight so that the determined second flight usage profile or profiles are actionable in the second flight. That is, the power management system selectively manages the power supplied so that the second flight usage profile can be followed in the second flight. As an example, the second flight may require the second flight takeoff phase to be performed solely using battery power, and this will be reflected in the second flight usage profile for the battery 14. The electrical generator usage profile will indicate that the electrical generator 506 is not needed for takeoff, and so the power management system may look to power the aircraft during the first flight substantially using the electrical generator 506, as the electrical generator 506 will not be needed for takeoff. The usage profile for the electrical generator 506 may take an exemplary form of the lower dashed line 570 in Figure 33, which although described in relation to electrical power management system 552, is also applicable here. The electrical generator usage profile may also indicate that the electrical generator 506 is needed for a cruise phase of a certain duration. Using this information, the power management system may look to power the aircraft during a descent phase of the first flight using the battery 14, in order to preserve a necessary amount of fuel to facilitate use of the electrical generator 506 throughout the cruise phase duration in the second flight. This may advantageously mitigate a need to refuel the aircraft at the first flight destination, so that turnaround time is minimised.

The second flight information can comprise flight plans (including re-routing information), mission restrictions, weather conditions (which may include changes to weather conditions and wind information), details of infrastructure at the first flight destination and information relating to a takeoff phase of the second flight. Mission restrictions may relate to constraints at future destinations, such as at the first flight destination (which is the second flight takeoff location). Constraints may include, for example, noise restrictions, runway length limits, en-route weather and traffic. It is beneficial for the system to have some or all of this information available to it. Flight plans allow for appropriate power management. Mission restrictions allow the system to manage power based on certain requirements for use of each power source at certain points in the second flight. For example, it may be important to have sufficient battery and electrical generator power to perform a short takeoff in the second flight, due to a short runway length. Details of infrastructure at the first flight destination allows the system to manage power in the first flight to, for example, mitigate a need to refuel at the first flight destination. This may be advantageous when there are a limited number of refuelling stations at the destination, or all are in use.

By selectively managing the power supplied by each electrical power source 586, the power management system facilitates use of the one or more of the electrical power sources in the takeoff phase of the second flight. That is, the power management system processes the second flight information to determine second flight usage requirements for that electrical power source, to facilitate its use in the takeoff phase of the second flight. Where said electrical power source is a battery, this may facilitate quiet takeoff. Where said electrical power source is an electric generator, this may facilitate short takeoff and high aircraft acceleration, by providing high power to both ground propulsion system 4 and air propulsion system 6. Power can be managed by managing power provided to the energy storage device 588 by one or more of: energy recovery systems 30, motors 12, turbogenerators 506 and/or renewable energy systems, such as solar panels or wind turbines, provided on the aircraft 1.

The plurality of electrical power sources 586 installed in the aircraft preferably comprise an energy storage device 588, which may be a battery pack 14, or the electrical power source 586 may be a fuel cell. Fuel cells can be hydrogen fuelled, which can provide a clean and efficient electrical power supply. Battery packs may advantageously be recharged in-flight using energy recovery systems, renewable energy systems, or even from other electrical power sources, including electric generators. The second flight usage requirements may require a part of the takeoff phase to be performed solely using the energy storage device. This may be advantageous where the power management system looks to obtain quiet takeoff. To facilitate this, the power supplied by each electrical power source is managed in the first flight so that the electrical power source is suitable for such use in the second flight.

The plurality of electrical power sources 586 may comprise one or more electric generators 594, or generally 506, energy recovery systems 30 and/or renewable energy systems. Electric generators 506, which may be referred to as “turbo generators”, are arranged to combust fuel from the fuel tank 500 to provide electrical power. Such power can be used instead of, or to supplement, battery power 588, 14, at various points in the aircraft flight.

To compare and contrast the power management system 584 with the electrical power management system 552, whilst system 552 is arranged to manage an electrical power source to achieve a target level for that power source, the power management system 584 is arranged to manage each of the plurality of electrical power sources 586 to facilitate use of one or more of the electrical power sources 586 according to determined second flight usage requirements for the one or more of the electrical power sources. It is advantageous for the plurality of power sources to comprise at least one energy storage device 588, such as a battery pack, because the energy storage device 588 can be recharged using another one or more of the electrical power sources.

As an example of the power management system 584 in operation, the receiver 590 may receive second flight information relating to a second flight. The second flight information informs the power management system of the runway length at the destination airport, and that a noise restriction is in place during the intended takeoff period of the second flight. The processor processes the information to determine second flight usage requirements for one or more of the electrical power sources. In this case, the processor may determine a usage requirement that the battery charge should be sufficient to power the ground propulsion system for a required time period to achieve takeoff within the length of the runway, and also that the battery charge must be high enough to facilitate the short takeoff solely using battery power, such that aircraft noise is kept to a minimum. To facilitate such use of the battery in the second flight, the power management system selectively manages the power supplied by each electrical power source, for example the electric generators, energy recovery systems, and battery packs, during the first flight, so that the takeoff can be performed as planned.

In Figure 36, general methodology associated with a power management system is shown. The method is a method of managing a plurality of electrical power sources installed in an aircraft. Step 596 comprises receiving second flight information relating to a second flight, the second flight subsequent to a first flight. Step 598 comprises processing the second flight information to determine second flight usage requirements for one or more of the electrical power sources. Step 600 comprises selectively managing the power supplied by each electrical power source in the first flight to facilitate use of the one or more of the electrical power sources in the second flight according to the second flight usage requirements.

Turning now to Figure 37, in another embodiment a propulsion management system 602 is arranged to selectively: interact the transducer assembly 11 with the one or more propellers 10 of the air propulsion system 6, to extract energy from the propellers 10; and additionally, or alternatively, control the air propulsion system 6 to drive the one or more propellers 10 in reverse in order to provide a reverse thrust. Reverse thrust may also be provided by physically rotating the propellers such that the rotational direction remains the same, but thrust is provided in a direction opposite to aircraft travel. Driving in reverse may therefore refer to the fact that the rotational direction can remain the same, but the propellers are rotated appropriately to provide reverse thrust. That is, the propulsion management system 602 may be arranged to selectively: control the energy recovery system 30, described above in relation to Figure 14, and/or to interact the transducer assembly 11 with the one or more propellers 10 of the air propulsion system 6, to extract energy therefrom. The propulsion management system is further arranged to drive the propellers 10 in reverse to provide a reverse thrust, as described above in relation to the propulsion management system 20. The selection of each operation may be autonomous, or user controlled but is preferably performed by the propulsion management system 602 without any user input. Selecting either function may depend on a desired amount of energy recovery, for example battery charging, or on a desired level of aircraft deceleration. The propulsion management system 602 may select to interact the transducer assembly 11 with a first propeller to extract energy therefrom, whilst selecting to control the air propulsion system 6 to drive a second propeller in reverse. Features here are compatible with those described above, and may be combined. This means that the propulsion management system 602 can work in different advantageous modes, and this might mean that constituent parts also work in different, related modes. For example, motors may work, or be driveable, for reverse thrust of a related propeller, forward thrust, or energy recovery, e.g. by ‘windmilling’. This might involve the use of one or more software, electrical, or mechanical modes of the motor, or of control systems associated with the motor. Extracting energy from the propellers means absorbing rotational kinetic energy from the propellers using the transducer assembly 11 and generating electricity which is fed to the battery 14 where it is stored. In one example of operation, reducing the thrust provided by the aircraft causes the aircraft to descend. Reducing the thrust coincides with the application of a load to the motors, by the transducer assembly 11, to extract power therefrom by harnessing the kinetic energy. It will be understood that in a holistic sense this is harnessing, or converting, gravitational potential energy to kinetic energy, thereby facilitating energy recovery by interaction with the transducer assembly.

The propulsion management system 602 comprises an energy accumulator 3 (see earlier Figures) for storing the extracted energy when the extraction mode is selected. The energy accumulator 3 may be comprised in a power supply for the air propulsion system 6. The energy accumulator is preferably a battery pack 14, but may also be a supercapacitor or flywheel.

The transducer assembly 11 is arranged to control the rotational speed of the one or more propellers 10. That is, the propulsion management system 602 can be operated to increase aircraft drag during a descent phase. Such a construction advantageously facilitates steep aircraft descent, and deceleration of the aircraft. The transducer assembly 11 is arranged to inhibit motion of the one or more propellers by interacting therewith, placing a load on the driving motors 12 to extract power therefrom. In one example, this advantageously allows the propellers 10 and motors 12 to function as turbine generators, from which power may be extracted.

The propulsion management system is arranged to both: control the air propulsion system 6 to drive propellers in reverse, and interact the transducer assembly 11 with one or more propellers 10 to extract energy from the propellers, e.g. during a landing roll. This allows both energy extraction and rapid deceleration during the landing roll. Extracting energy, which can then be stored in the energy accumulator 14, during a landing roll, is highly advantageous as the energy can be used to power the air propulsion system 6 and ground propulsion system 4 in a subsequent takeoff phase. The propellers may be fixed or variable pitch propellers, and the pitch of the propeller 10 may be optimised to provide a desired or necessary level of energy extraction and/or reverse thrust. That is, the propellers 10 may be adjustable to control the level of energy extraction and/or reverse thrust. Such adjustment may be automated by the propulsion management system, depending on level of energy extraction required or demanded, or deceleration demanded of the aircraft by the aircraft pilot. That is, the pilot or user might not need to manually implement any selection or mode, and the system itself can seamlessly implement any selection or mode dependent on required system functionality, all without explicit, dedicated user or pilot input. Also, different system or component parts are not required to implement multiple functions, saving space, weight and cost.

As above, the air propulsion system 6 comprises electric motors 12 drivingly connected to the propellers 10 for providing drive thereto. The transducer assembly 11 may function to extract power from the motors 12. In an exemplary embodiment, the motor 12 is an outboard motor, housed in an outboard nacelle 508 of the aircraft 1. Normal function of the air propulsion system 6 may involve driving the propellers 10 to provide a forward thrust. Each propeller can be driven at a different rotational velocity such that each propeller may provide a different level of thrust. In one embodiment, a first propeller is driven to provide a higher level of thrust than a second propeller, and the propellers may be optimised for the level of thrust provided. The system may select the energy extraction functionality or reverse thrust functionality in certain phases of flight, when it is preferable, for example, to charge batteries 14, or to decelerate the aircraft.

In Figure 38, general methodology associated with a propulsion management system in an aircraft comprising an air propulsion system having one or more propellers is shown. Step 604 comprises interacting a transducer assembly with one or more propellers of the air propulsion system to extract energy from the propellers. Additional or alternative step 606 comprises controlling the air propulsion system to drive the one or more propellers in reverse in order to provide a reverse thrust

Turning now to Figure 39, a spar assembly 608 for segregating components installed in an aircraft wing 504 is shown, in another example embodiment. The spar assembly 608 comprises two spar arrangements 610, 612 installed in the aircraft wing 504, one toward the wing leading edge 614 of the wing, and the other toward the trailing edge 616 of the wing. The trailing edge 616 of the wing comprises ailerons and flaps (not shown), moveable by actuators (not shown). The spar arrangements 610, 612 are formed from a plurality of substantially parallel, separated spar webs 618. Voids 620 are defined between the spar webs 618. Each spar arrangement 610, 612 is locatable between components in the aircraft wing 504, and the voids 620 are located between the components. As a result, the spar arrangements 610, 612 not only provide structural support to the wing 504, but also segregate the components and provide improved insulation (e.g. electrical or thermal) between the components, by virtue of the void. Advantageously, this reduces the possibility of electrical arcing between the components, and reduces dangers associated with lightning strikes.

Furthermore, each spar arrangement is formed from an insulating material (e.g. electrical or thermal). The insulation provided by the spar arrangement, and void formed therein, ensures electrical arcing is prevented as electric current is not conducted from the skin 622 of the aircraft to the spar arrangements 610, 612 located between the components. In one example, the components are formed from conductive materials, for example metals, and may be conductive cables, such as power cables 624 arranged to carry power to the aileron or flap actuators. The components may also be electric generators 506, including a fuel cell, fuel tanks 500 and energy storage devices, such as a battery 14. All such components typically comprise metallic parts. As shown in Figure 40, the spar arrangements 610, 612 each have a construction wherein two spar webs 618 are connected by an upper plate 626 and a lower plate 628. A first component receiving portion 630 is defined between the upper plate 626, an outward facing side of a first spar web 618a, and the lower plate 628. A second component receiving portion 632 is defined between the upper plate 626, an outward facing side of a second spar web 618b, and the lower plate 628. The void 620 is defined between the spar webs 618, between the upper plate 626, lower plate 628, and two inwardly facing sides of the spar webs 618. The first component receiving portion 630 is arranged to receive conductive power cables 624, and comprises segregating members 634 in the form of protruding fingers or plates, which extend in the first component receiving portion 630 to segregate the cables 624. Such a construction may be referred to as a “race track”. Advantageously, this segregates the conductive cables 624 to assist with cable cooling and routing. Each spar arrangement therefore provides or facilitates multiple functions, in terms of different types of segregation, insulation, structural support, and so on.

Each spar web may have a thickness of between 2 to 10mm, and the void may have a width, that is distance between the spar webs, or between 10 to 50mm. Such dimensions sufficiently separate the cables 624from the fuel tank 500.

In an exemplary and advantageous embodiment, the spar assembly 608 comprises a first and second spar arrangement 610, 612 used to segregate a fuel tank 500 and conductive power cables 624 at the leading edge 614 and trailing edge 616 of the wing 504. The first spar arrangement 610 is installed toward the leading edge of the wing, and the second spar arrangement 612 is installed toward the trailing edge. The fuel tank 500 is located substantially in the middle of the wing 504, so that the spar arrangements 610, 612 are located either side of the fuel tank 500, on opposite sides of the fuel tank. The first spar arrangement 610 is arranged to segregate a first conductive cable 624a, or plurality of cables, along which current flows in a first direction to electronic components at the wing tip. The second spar arrangement 612 is arranged to segregate a second conductive cable 624b, or plurality of cables, along which current flows in a second direction to electronic components in inboard nacelles, or in the fuselage. It will be appreciated that the first direction and second direction are substantially opposite directions. Thus, the cables 624a, 624b in which current flows in opposite directions are separated and do not travel along the same side of the wing. This improves safety, and provides improved cabling routing, for example for ice protection at the leading edge, and for aileron and flap actuators at the trailing edge.

In Figure 41, general methodology associated with the spar arrangement is shown. The method is of segregating components installed in an aircraft. Step 640 comprises providing a spar arrangement. Step 642 comprises locating the spar arrangement between the components to provide a void therebetween.

Turning now to Figure 42, a fluid fuel monitoring system 644 is shown. The fluid fuel monitoring system 644 is described in relation to being installed in the aircraft 1, however it will be appreciated by the person skilled in the art that the fluid fuel monitoring system 644 is suitable for use in any vehicle arranged to be fuelled by a fluid fuel supply, and in particular a vehicle (or, indeed, object) fuelled with different fluid fuel types. The fluid fuel monitoring system comprises a sensing apparatus 646 arranged to sense a fluid flow property of the fluid fuel supply, the fluid flow property being indicative of the composition of the fluid fuel supply (e.g. not being indicative of a fuel level, or tank capacity). The fluid fuel monitoring system further comprises a controller 648 arranged to control operation of one or more components of the aircraft based on the fluid flow property.

By controlling the aircraft operation based on the fluid flow property, different compositions of fluid fuels (that is, different types and mixtures of fluid fuels) may be used in a single system, and operation of the aircraft adjusted accordingly. In existing systems, fluid fuels must all have the same properties for predictable operation, and thus in a system configured to use multiple fuel types, certain fuels, such as biofuels, are “corrupted” so that they have identical fluid properties to other fuels used in the system, so that aircraft operation need not be adjusted when the biofuel is introduced. However, the present system advantageously allows different compositions to be used by sensing the fluid fuel properties and controlling operation accordingly. Safe and reliable operation is thus ensured. Also, manual control or tweaking by the pilot or user is not needed - the system seamlessly implements the detection and control. The pilot or user will not need to know of the multi-fuel use, and related control changes needed to reflect changes in fuel composition at any time.

The system comprises an arrangement of conduits 650 arranged to receive fluid fuel from a plurality of fuel tanks 652, 654 and combine the fluid fuels in a single conduit 650a. The sensing apparatus 646 is arranged to sense a fluid flow property of the combined fuel in the single conduit 650a. The fluid fuel supply may comprise kerosene, biofuel and/or liquid hydrogen. In one example, the fuels, in particular kerosene and biofuel, may be stored in separate fuel tanks 652, 654. The kerosene and biofuel may be combined in the single conduit 650a, in which sensing apparatus 646 is provided. The composition of the mixture of fuels can be determined from the fluid flow property, and the system can thus control operation of the aircraft 1 based on the fluid flow property i.e. based on the composition of the mixture. The composition of the mixture may be determined by reference to look-up tables, data stored in memory, or the like. Such a construction is highly advantageous as the controller can control operation of fuel pumps, engines, motors 8, 12 and/or electric generators 506 based on the fluid flow property, thus ensuring safe and reliable aircraft operation.

The sensing apparatus 646 is arranged to sense one or more of: fluid viscosity and/or density. These properties can be measured in a simple yet effective manner to allow the composition of the mixture to be determined. The sensing apparatus 646 may comprise one or more flow meters for measuring flow rate. Flow rate is indicative of the composition of the fluid fuel supply. The controller can receive this information to control the aircraft operation based thereon, ensuring safe and reliable aircraft operation. Knowledge of the input fuels is likely to be required, so that the measurements can be used in combination with this knowledge, in order to determine, or more readily determine the fluid flow property that is indicative of composition (including related changes in composition), or system changes needed in response to changes in composition.

The controller 648 is arranged to control operation of the fuel pumps, motors 8, 12 and/or electric generators 506 in response to a change in the sensed fluid flow property, indicating a change in composition of the fluid fuel supply. For example, the fuel pump power may be increased when an increase in density is sensed. Such an increase may indicate an increase in the concentration of biofuel in the mixture. That is, fuel system regulation may be adjusted. Additionally, the combustion process can also be controlled and adjusted based on the sensed fluid flow property, or changes in the fluid flow property. The controller 648 is arranged to control operation of the motors 8, 12 and/or electric generators 506 in response to the change in the sensed fluid flow property to maintain power output by the motors 8, 12 and/or electric generators at a substantially constant level. The controller 648 is also arranged to control operation of the motors 8, 12 and/or electric generators 506 in response to the change in the sensed fluid flow property to deliver a level of power output by the motors 8, 12 and/or electric generators that is commanded by the user of the system. The operation of motors 8, 12 and/or electric generators 506 in response to the change is autonomously controlled by the fluid flow monitoring system 644, without user input.

As contextual background, in some examples of conventional aircraft missions, a mission starts with a known volume of fuel in the tanks. Providing the viscosity and density of that fuel is within limits identified by the flow rate sensors on the aircraft then those sensors can determine the fuel flow rate in flight. That flow rate can be combined with the initial fuel volume to inform the pilots of the fuel usage/remaining throughout the mission. It is for this reason that biofuels are modified to conform to a specific range of viscosity and density. The fundamental nature of flow measurement is the correlation of a pressure drop across an orifice of known dimension, a venturi for example. If such a sensor was used with fuel of a different viscosity or density, then the sensor would give an erroneous reading of fuel flow. This could impact the safe operation of the aircraft.

If two sensors of sensing apparatus 646 with, for example, different orifice (or opening) diameters were used to measure the flow of the same fluid then the error due to a change in physical properties of the fluid would be different for each sensor. The quantified differences from these two independent sensors would enable the variation of individual physical properties of the fluid (e.g. viscosity and density) to be ascertained. That information could then be used to modify fuel pumps, motors 8,12, generators 506 and other fuel system components (via software commands) to ensure that the correct amount of fuel is delivered to the combustor to maintain the safety and performance of the aircraft 1.

The fluid fuel monitoring system 644 thus provides an advantageous system which allows a fuel supply comprising a plurality of fluid fuel types to be monitored and aircraft operation autonomously controlled. It is highly advantageous to provide a system which facilitates use of a mixture of fluid fuel types, especially where the system accommodates the use of biofuels. The system 644 is compatible with, and only requires minor modifications to, existing aircraft architecture.

In Figure 43, general methodology associated with a fluid fuel monitoring system 644 is shown. Step 660 comprises sensing a fluid flow property of the fluid fuel supply, the fluid flow property being indicative of the composition of the fluid fuel supply. Step 662 comprising controlling an operation of the vehicle based on the fluid flow property.

Turning now to Figure 44, a power system structure 664 is shown. The power system structure comprises an electric generator 506 for generating electricity to power a propulsion system of the aircraft. In this example, the propulsion system is the air propulsion system 6, and in particular the air propulsion system motors 12. The electric generator 506 is housed in the nacelle 508, described above. Advantageously, this ensures that the electric generators are located close to the electronics which they are required to power, thus reducing the length and weight of electrical power cabling required to electrically connect the electric generator to the components. Moreover, housing the electric generator 506 in the nacelle 508 is advantageous in passive cooling of the electric generator in flight, as the generator is close to the aircraft skin. Additionally, the electric generator being housed in the nacelle 508 is beneficial during maintenance, as the nacelle 508 is easily accessible and often comprises removable panels. Whilst this might seem like a trivial implementation, this demonstrates the more holistic approach to the design of electric/hybrid aircraft that is disclosed herein, as opposed to simply forcing electric/hybrid onto combustion-centric designs.

The air propulsion system motor 12 is also housed in the nacelle 508, along with the electric generator 506. The electric generator is arranged to provide power primarily to that electric motor. Advantageously, this ensures that the motors 12 and electric generators 506 are located in close proximity, thus reducing the length and weight of electrical power cabling required to electrically connect the electric generator to the motor. Thermal management is also simplified by locating the motors and electric generators in close proximity. The nacelle 508 is an inboard nacelle 508. Such a construction advantageously ensures that, where cabling must extend from the fuselage 535 to the inboard nacelle 508 (e.g. for power supply or drawing power, or for fluid fuel supply), cable or conduit length and weight is minimised.

In Figure 45, a general methodology associated with a power system structure is shown. Step 666 comprises providing an electric generator for generating electricity to power a propulsion system of the aircraft. Step 668 comprises housing the electric generator in the nacelle 508.

Turning now to Figure 46, and with reference to Figure 1, a power system structure 670 for an aircraft 1 is shown. The aircraft 1 comprises a flight control actuator system 672 arranged to actuate control mechanisms and/or control surfaces. For example, the control mechanisms and/or control surfaces comprise: elevators 674 and rudders 676 to control the trajectory of the aircraft 1, ailerons, spoilers, landing gear systems, steering mechanisms and/or other flaps. The flight control actuator system 672 comprises electric or electric-hydraulic actuators. Both types of actuator system are reliable and durable. The power system structure 670 comprises two battery packs 678 arranged to provide power to the flight control actuator system 672 via power cables. The battery packs 678 provide power primarily, which may be only, to the flight control actuator system 672. As such, the battery packs 678 are not relied upon (at all, or heavily) for use in any other system, increasing operating safety. Moreover, this construction minimises the length and weight of power cables that extend through the aircraft 1, as the dedicated battery packs 678 can be placed in close proximity to the flight control actuator system 672 that they are arranged to power.

The battery packs 678 are located proximal to the flight control actuator system 672. Advantageously, by locating the battery pack 678 proximal to the flight control actuator system 672, the length and weight of power cable in the fuselage 535 is minimised. Moreover, use of two battery packs 678 ensures a level of redundancy, such that in the event of failure of one of the battery packs the flight control actuator system may continue to operate safely.

The skilled person will appreciate that the battery packs 678 may be interchanged with any other energy storage device, such as supercapacitors or flywheels, to obtain similar advantages of redundancy. In this example, the battery packs 678 are interconnected such that in the event of failure of one of the battery packs, the other battery pack is arranged to provide power to the flight control actuator system 672. That is, each battery pack 678 can function as a backup for the other. The power system structure may detect the failure of one of the battery packs 678, and command the other battery pack to supply power in place of the failed battery pack. This construction ensures that the flight control actuator system 672 can continue to operate normally and safely.

The battery packs 678 are charged at low voltage by a power supply provided by the energy recovery system 30, renewable energy systems, such as photovoltaics or turbines, electric generators 506 and/or other battery packs 14 installed in the aircraft. Preferably, the battery packs 678 are trickle charged, that is for example, charged at a rate equal to a discharge rate at a low voltage, in order to maintain the batter pack 678 at its substantially fully charged level during flight, or to prevent the battery pack level dropping below a threshold charge level. Both the flight control actuator system 678 and the power system structure 670 are located at the rear of the fuselage 535, and, in an exemplary and advantageous embodiment, are both housed in the empennage. This construction ensures that the length and weight of power cable in the fuselage 535 is minimised. This additionally simplifies cable routing. It will be appreciated that both the flight control actuator system 678 and the power system structure 670 can be housed in the wings and/or belly fairing, and similar advantages will be attained therefrom.

In Figure 47, general methodology associated with a power system structure is shown. The method is for an aircraft comprising a flight control actuator system. Step 680 comprises providing a power system comprising one or more energy storage devices. Step 682 comprises arranging the energy storage devices to provide power to the flight control actuator system via power cables. Step 684 comprises locating the energy storage devices proximal to the flight control actuator system.

In Figure 48, general methodology associated with a power system structure is shown. The method is for an aircraft comprising a flight control actuator system. Step 686 comprises providing a power system comprising one or more energy storage devices. Step 688 comprises arranging the energy storage devices to provide power primarily to the flight control actuator system.

In summary, we provide systems that enable reduction in aircraft noise pollution during the various phases of aircraft flight, including enabling quieter operation during a takeoff roll and a steeper and earlier climb to altitude. We also provide systems which enable quieter and steeper aircraft descent, which are also advantageous in reducing noise pollution. The systems incorporate batteries as a source of power, and thus can be charged using renewable sources, thereby enabling inflight regeneration whilst reducing emissions and carbon footprint. Reduced turnaround time is facilitated as a result of inflight battery charging and by providing a system for reverse thrusting, reducing reliance on wheel braking. Power from power sources in the aircraft can be monitored, managed and distributed effectively to aircraft components as necessary or as desired for safe, reliable and efficient aircraft operation. Redundancy is provided in the event of component failure, further increasing aircraft security and reliability. Component health can be monitored to prevent degradation and failure, and this further assists in optimising power usage to ensure that despatch reliability and short turnaround time is achieved. Above described systems facilitate fuel use monitoring, allowing different fuel types to be compatible with a single system. This overcomes disadvantages with current commercial approach of corrupting the fluid fuel supply. Components can be isolated and segregated by employing the spar arrangement described above. This increases aircraft safety and reliability by reducing the dangers associated with lightning strikes and electrical arcing. Furthermore, the spar arrangement provides for improved cable routing. Power system structures described above facilitate improves cable routing, allowing lengths and weight of power cable to be minimised. Ultimately, this reduces cable heating, and improves aircraft safety. Power system structures described above provide a level of redundancy, ensuring safe and reliable operation of aircraft in the event of component or power source failure. Finally, the systems can be provided in new aircraft or can be retrofitted in existing aircraft to provide improved performance during the various stages of aircraft flight.

At least some of the example embodiments described herein may be constructed, partially or wholly, using dedicated special-purpose hardware. Terms such as ‘component’, ‘module’ or ‘unit’ used herein may include, but are not limited to, a hardware device, such as circuitry in the form of discrete or integrated components, a Field Programmable Gate Array (FPGA) or Application Specific Integrated Circuit (ASIC), which performs certain tasks or provides the associated functionality. In some embodiments, the described elements may be configured to reside on a tangible, persistent, addressable storage medium and may be configured to execute on one or more processors. These functional elements may in some embodiments include, by way of example, components, such as software components, object-oriented software components, class components and task components, processes, functions, attributes, procedures, subroutines, segments of program code, drivers, firmware, microcode, circuitry, data, databases, data structures, tables, arrays, and variables. Although the example embodiments have been described with reference to the components, modules and units discussed herein, such functional elements may be combined into fewer elements or separated into additional elements. Various combinations of optional features have been described herein, and it will be appreciated that described features may be combined in any suitable combination. In particular, the features of any one example embodiment may be combined with features of any other embodiment, as appropriate, except where such combinations are mutually exclusive. Throughout this specification, the term “comprising” or “comprises” means including the component(s) specified but not to the exclusion of the presence of others. Attention is directed to all papers and documents which are filed concurrently with or previous to this specification in connection with this application and which are open to public inspection with this specification, and the contents of all such papers and documents are incorporated herein by reference. All of the features disclosed in this specification (including any accompanying claims, abstract and drawings), and/or all of the steps of any method or process so disclosed, may be combined in any combination, except combinations where at least some of such features and/or steps are mutually exclusive. Each feature disclosed in this specification (including any accompanying claims, abstract and drawings) may be replaced by alternative features serving the same, equivalent or similar purpose, unless expressly stated otherwise. Thus, unless expressly stated otherwise, each feature disclosed is one example only of a generic series of equivalent or similar features. The invention is not restricted to the details of the foregoing embodiment(s). The invention extends to any novel one, or any novel combination, of the features disclosed in this specification (including any accompanying claims, abstract and drawings), or to any novel one, or any novel combination, of the steps of any method or process so disclosed. The invention and above-described features, elements and systems may be adopted or deployed in other transportation and vehicular systems, aerial or otherwise, such as in manned or unmanned drones.