GB2634034A

GB2634034A - Apparatus

Info

Publication number: GB2634034A
Application number: GB2314763.0A
Authority: GB
Inventors: Nese Francesco; Stonham Joseph; Brooks Ashley; Wood Norman
Original assignee: GKN Aerospace Services Ltd
Current assignee: GKN Aerospace Services Ltd
Priority date: 2023-09-26
Filing date: 2023-09-26
Publication date: 2025-04-02
Also published as: WO2025068700A1; GB202314763D0

Abstract

An electrical aircraft propulsion arrangement 200 comprising a fuel cell 204 arranged to generate electrical power for use in propelling an aircraft, a gas source 202 providing gas to the fuel cell for use in generating electrical power, a centrifugal compressor to compress the gas prior to provision to the fuel cell, and an axial compressor arranged to compress the gas prior to provision to the fuel cell. The axial compressor is upstream of the centrifugal compressor. The gas may be ambient air, oxygen or hydrogen. An energy demand sensor may detect energy demand of the aircraft and a controller may be arranged to control an amount of compression of gas by the centrifugal compressor and the axial compressor in response to a detected energy demand.

Description

Apparatus

Technical Field

The present invention is concerned with green propulsion systems for aircraft. The present invention relates to an energy management system and a turbo-compressor arrangement for use in such an energy management system. The present invention also considers air condition and thermal management within propulsion arrangements for aircraft.

According to most estimates, airline traffic is set to double every fifteen years providing a significant increase in the operation of land-based and, subsequently, airborne propulsion systems and therefore the production of associated emissions. Emissions are known to be harmful whether produced at ground level or at altitude.

In order to meet targets for reduction of emissions set by the International Air Transport Association, the use of alternate fuels has been identified as a possible avenue of exploration. Alternate fuels include biofuels, synthetic kerosene, compressed natural gas. In addition, the ACARE roadmap for 2050 identifies the need and sets objectives for significant reductions for a range of emissions. It is widely recognised that the opportunities to come close to or achieve these targets are limited.

To solve these issues a number of propulsion systems have been employed in different aircraft. Most systems use fossil fuel sources for economic reasons and also due to their very high energy density and specific energy. The prevalence of the gas turbine has also led to fossil fuels being a desirable propulsion mechanism for aircraft. This has led to developments for improving the performance of fossil fuel burning gas turbines.

Some aircraft use electrical systems to provide propulsion. For example, electrical energy can be provided to a propulsor from an energy store or energy source and used to generate thrust.

Energy sources such as fuel cells can be used to provide electrical energy for propulsion. Electrical energy stores such as batteries may be used to provide electrical energy for propulsion. Use of fuel cells in particular requires control when operated at varying altitudes as would be experienced by fuel cell powered aircraft.

Developments exist that consider small aircraft. The developments used for small aircraft are not feasible for large or commercial aircraft.

Summary of the Invention

Aspects of the invention are set out in the accompanying claims.

Viewed from first aspect there is provided an energy management system for an electrically powered aircraft, the system comprising: at least one fuel cell arranged to generate electrical power; a fluid source arranged to provide fluid to at least one fuel cell for use in generating electrical power; at least one centrifugal compressor arranged to compress a fluid prior to provision to at least one fuel cell; at least one axial compressor arranged to compress a fluid prior to provision to at least one fuel cell, wherein at least one axial compressor is upstream of at least one centrifugal compressor.

Thus, according to an energy management system can be provided for aircraft using electrical propulsion. In particular, this system has been found to be suitable for large aircraft. Use of electrical power for aircraft reduces the environmental impact of air flight. The arrangements disclosed herein may be used in electrical-only powered aircraft and partially electrically powered aircraft ("hybrid aircraft").

Radial compressors are suited to low fluid flow rates where high pressure ratio is needed, but can become mass/volume intensive and less efficient as the size of the aircraft increases.

There is also a limit on the radial impeller geometry that can be designed as the diameter increases. If you want to power a large aircraft this leads to a need for multiple small compressors in parallel.

Multi-stage, axial compressors can achieve the same overall pressure ratio with higher efficiency, lower mass and greater compactness than equivalent radial compressors, but designs are only feasible at high flow rates. This is because the blade geometries become too small, particularly in the final stages of the axial compressor.

By combining the two, for example in a larger aircraft, a feasible design window becomes accessible for the axial compressor to provide initial compression, taking advantage of its efficiency and compactness benefits. The latter compression is performed by a centrifugal ("radial") compressor. This therefore allows both to provide high performance.

In particular, the arrangement disclosed herein is highly effective at providing a suitable compression ratio for air when used at a varying altitude suitable for a full size commercial aircraft. The arrangement disclosed herein is also effective at providing flow rates of fluid to the fuel cell to account for power demands from the aircraft users.

In some examples, the system further comprises an energy demand sensor arranged to detect an energy demand associated with an electrically powered aircraft, a controller arranged to control an amount of compression of fluid from the fluid source by the at least one centrifugal compressor and the at least one axial compressor in response to an energy demand detected by the energy demand sensor.

An example of an energy demand sensor may be a current and/or a voltage sensor arranged to detect electrical properties of the fuel cells.

The controller provides intelligent and reactive control over aspects of the energy management system. In response to detections of an energy demand the compression applied by the compressors can be tailored accordingly. In this way, account can be made of varying power requirements from the system and this power can be provided accordingly. In an aircraft, this may be by virtue of different stages of flight or a change in the air condition or the like and the impact on thrust demand or may be due to greater requirement for power in the cabin or the like.

In some examples, the system further comprises a pressure sensor, wherein the pressure sensor is arranged to provide signals to the controller on a pressure of fluid in the energy management system, and wherein the controller is arranged to control an amount of compression of fluid in response to the signals from the pressure sensor.

The pressure of the fluid to be provided to the fuel cell relates to the power that the fuel cell provides. Therefore, by sensing the pressure of the air to the fuel cell, detecting the energy demands on the fuel cell, the controller can reactively ensure the pressure of the air to the fuel cell is increased (as required) to ensure the fuel cell provides the energy demanded. This is therefore a reactive and highly efficient system. The compressor arrangement allows such power to be provided as would be required for thrust of large commercial aircraft.

This synergistically provides a reliable green system for thrust in an (at least partially) electrically powered aircraft.

In some examples, the system further comprises a motor, wherein the controller is arranged to control the speed and/or rotation of the motor in response to an energy demand detected by the energy demand sensor.

The speed or rotation of the motor can also be controlled by the controller in response to the energy demand on the fuel cell. The motor is able to reclaim what would be waste energy from the system.

The motor may be an electric motor.

As will be discussed, this efficient in itself and is highly efficient when rotating portions of the system are located on a mechanical system that the motor is also located on.

In some examples, at least one axial compressor comprises variable guide vanes comprising a controllably variable geometry, and wherein the controller is arranged to control the variable geometry of the at least one axial compressor in response to an energy demand detected by the energy demand sensor.

Use of variable guide vanes provides control for a user over the compressor ratio, as the user can change the arrangement of the guide vanes. This in turn provides the user with a level of control over the compressor ratio. It may be advantageous for the user to have control over the compressor ratio to allow for account for the change in atmospheric pressure as encountered by aircraft (in a way that ground-based electrically propelled vehicles do not).

In example, the present arrangement uses a coupled compressor arrangement, and the arrangement can take advantage of the additional design point efficiency from using both an axial compressor and a centrifugal compressor. The present arrangements also use variable geometry in compressors in order to enhance operability of the system across the whole mission (relevant for aircraft where the power demands vary significantly). Variable geometry may be provided by moving inlet guide vanes.

Moving the inlet guide vanes changes the relationship between the compressor pressure ratio, flow rate, and rotational speed (i.e. moves the map). It allows multiple design points for the compressor, instead of only one. This results in a) improvement of the off-design efficiency of the compressor, b) avoiding compressor surge whilst minimising wasteful compressor bleed flows, and c) achieving a compact and lightweight solution.

In some examples, the system further comprises a temperature management arrangement comprising a heater arranged to provide thermal energy to a fuel for the at least one fuel cell, and a temperature sensor arranged to detect a temperature of a fluid from the fluid source, wherein the controller is arranged to control the provision of thermal energy by the heater in response to signals from the temperature sensor and the energy demand sensor.

Use of a heater allows a fluid to be provided to the fuel cell be heated to a suitable temperature for use in the fuel cell. When the fluid is provided at high flow rates, there may be not sufficient passive heating to warm the fuel. The fuel may be provided in a cryogenic condition and therefore heating is advantageous for efficient power generation. The heater may heat hydrogen into a gaseous state (or further heat gaseous hydrogen) prior to use in the fuel cell. The heater may be referred to as a "superheater as it may superheat hydrogen.

In some examples, the temperature management arrangement further comprises an inter-stage cooler arranged between the at least one axial compressor and the at least one centrifugal compressor, wherein the inter-stage cooler is arranged to cool a fluid from the fluid source prior to compression by the at least one centrifugal compressor, wherein the controller is arranged to control the provision of cooling by the inter-stage cooler in response to signals from the temperature sensor and the energy demand sensor.

In examples, the inter-stage cooler may make use of a cryogenic fuel for cooling that is also used in the system to provide a fuel to the fuel cell. In this way, the inter-stage cooler can cool the air to be compressed while heating the fuel for use in the fuel cell to an advantageously suitable temperature prior to use in the fuel cell. This provides great levels of efficiencies for the present arrangement.

Use of a cooler of any form provides the advantage of cooling the fluid prior to a compression stage. By cooling the fluid, the work required to further compress the fluid is reduced. As such, it can be highly effective for the controller to be informed of the energy demand, the flow rate of the fluid, the compression on the fluid and if further compression is required to know the temperature of the fluid prior to, e.g., the final compression stage. Cooling the fluid before the final (centrifugal) compression stage reduces the work required to compress the fluid and therefore provides efficiencies.

In some examples, the system further comprises a turbine arranged to receive an output fluid from the at least one fuel cell. The turbine is able to advantageously keep pressure in the system. In some examples, the turbine is at least one of: a variable geometry turbine, wherein the controller is arranged to control the variable geometry of the turbine; and, an axial turbine.

Providing a controllably variable geometry allows efficiencies to be obtained as per the controller understanding the power demands on the system and providing a turbine behaviour as desired to reclaim as much energy as is required in the instance. Varying the geometry also allows control over the regulation of pressure by preventing excessive buildup of pressure in the system by diverting excess gas from the fuel cell exhaust away from the turbine.

In some examples, the system further comprises a bypass valve arranged to control the passage of fluid into and out of the at least one fuel cell, wherein the controller is arranged to control the bypass valve in response to an energy demand detected by the energy demand sensor. In some examples, the bypass valve is controllable by the controller to provide fluid to the turbine in response to an energy demand detected by the energy demand sensor.

Bypass valve advantageously maintains high pressure for turbine, in response to controller signals. A restrictor valve may be used in the same manner to maintain pressure. This system benefits highly from controllable pressure levels throughout the system. This allows for a release of pressure where high energy recovery from the fuel cell exhaust is not required.

In some examples, the system further comprises an air filter to filter fluid provided to the at least one fuel cell.

Certain disclosures herein relate to conditioning of air, prior to providing that air to a fuel cell or fuel cell stack. Conditioning the air may include pressure considerations but also purity of air. Fuel cells may be badly damaged by pollution and contaminants. As such, use of a filter further improves the conditioning of the air prior to introduction to the fuel cell.

In some examples, the system further comprises comprising an air quality sensor to detect an air quality, wherein the controller is arranged to control fluid to pass through the air filter in response to signals from the air quality sensor associated with an air quality below a predetermined air quality.

The controller can reactively employ filtration of air when it is required to prevent contamination and damage to the fuel cell.

In some examples, the system further comprises a flight stage sensor arranged to detect a flight stage associated with an electrically powered aircraft, wherein the controller is arranged to control an amount of compression of fluid from the fluid source by the at least one centrifugal compressor and the at least one axial compressor in response to a flight stage detected by the flight stage sensor The controller can reactively provide compression suitable to the flight stage, or the next stage of flight. When the controller is informed that the system is about to come to cruise from climb, the controller can prepare the compressors for a lower thrust demand. Similarly, preparation for high demand flight stages is provided advantageously by this arrangement.

In some examples, the system further comprises an electrical energy store to store electrical energy, wherein the controller is arranged to provide energy from the electrical energy store in response to an energy demand detected by the energy demand sensor. Use of a battery or the like provides an emergency power store for, e.g. use in in flight safety events. This may also be used to reduce the load on the fuel cell where preferential.

In some examples, at least one axial compressor comprises a multi-stage axial compressor, wherein each stage offers a ratio of compression around 1.1 to 1.7.

The compression total compression ratio may be up to around a total of about 10. This has been found to be particularly effective for larger commercial aircraft. Larger aircraft are not currently seen as compatible with greener fuels due to the large drawbacks associated with power production (gas turbines are therefore the prevailing solution employed in large aircraft). This two-stage axial compressor provides a compression ratio that makes the use of green fuels on larger aircraft increasingly feasible.

In some examples, the at least one centrifugal compressor and the at least one axial compressor are mechanically coupled and wherein the mechanical coupling is provided by a shaft. As noted above, benefits are provided by inclusion on the same shaft including reduction in overall weight (by not having multiple shafts).

View from another aspect there is provided an electrical aircraft propulsion arrangement comprising the energy management system as noted above.

View from another aspect there is provided an aircraft comprising the electrical aircraft propulsion arrangement as noted above.

View from another aspect there is provided a method of managing energy provision for an electrically powered aircraft, the method comprising: providing fluid from a fluid source for use in generating electrical power; compressing, by at least one axial compressor, the fluid from the fluid source; receiving, by at least one centrifugal compressor, the fluid from at least one axial compressor; compressing, by the at least one centrifugal compressor, the fluid from the fluid source; and, providing the fluid to at least one fuel cell arranged to generated electrical power for use in propelling an aircraft.

The proposed system provides improvements of efficiency, operability and providing a more 5 compact design compared to modern machines.

Brief Description of the Drawings

One or more embodiments of the invention will now be described, by way of example only, and with reference to the following figures in which: Figure 1 shows a schematic of a turbo-compressor arrangement according to examples disclosed herein; Figure 2 shows a schematic of a system comprising an energy management system according to examples disclosed herein; and, Figure 3 shows a schematic of an electrical aircraft propulsion arrangement comprising an energy management system according to examples disclosed herein.

Any reference to prior art documents in this specification is not to be considered an admission that such prior art is widely known or forms part of the common general knowledge in the field. As used in this specification, the words "comprises", "comprising", and similar words, are not to be interpreted in an exclusive or exhaustive sense. In other words, they are intended to mean "including, but not limited to". The invention is further described with reference to the following examples. It will be appreciated that the invention as claimed is not intended to be limited in any way by these examples. It will also be recognised that the invention covers not only individual embodiments but also combination of the embodiments described herein.

The various embodiments described herein are presented only to assist in understanding and teaching the claimed features. These embodiments are provided as a representative sample of embodiments only, and are not exhaustive and/or exclusive. It is to be understood that advantages, embodiments, examples, functions, features, structures, and/or other aspects described herein are not to be considered limitations on the scope of the invention as defined by the claims or limitations on equivalents to the claims, and that other embodiments may be utilised and modifications may be made without departing from the spirit and scope of the claimed invention. Various embodiments of the invention may suitably comprise, consist of, or consist essentially of, appropriate combinations of the disclosed elements, components, features, parts, steps, means, etc, other than those specifically described herein. In addition, this disclosure may include other inventions not presently claimed, but which may be claimed in future.

Detailed Description

An invention described herein relates to an energy management system for an electrically powered aircraft and components within what may be a propulsion arrangement. The components may form a turbo compressor arrangement or the like. The disclosure below contains discussions of elements that may stand alone or be included in an electrical aircraft propulsion arrangement. The disclosure contains highly efficient arrangements for energy management that significantly improve the modern arrangements used in commercial aircraft.

Operation of electrically powered aircraft may use fuel cells. Aircraft change altitudes during flight. As such, the condition of the fuel cell during the flight requires consideration. Specifically, examples herein control the pressure difference between a preferred fuel cell air pressure and ambient pressure (which changes with altitude). Further considerations herein relate to air condition prior to provision to a fuel cell (or equally to a fuel cell stack). Such conditions include temperature, humidity and pollutants therein.

Examples herein are suitable for use in commercial aircraft, i.e. aircraft that may be around from 19 to around 96 passenger aircraft (and larger).

For smaller aircraft use of centrifugal compressors has been found to be sufficient for providing air at a suitable pressure for the fuel cell. Such an arrangement does not scale sufficiently for large aircraft at high altitudes.

As such, disclosed herein is a large scale (suitable for e.g. 19, 48 and 96 passenger aircraft) electrically operated electrical aircraft propulsion arrangement that is highly effective at a large ranges of altitudes.

The arrangement disclosed herein addresses challenges faced to design a suitable and efficient air delivery system. The challenges relate to how best to handle parasitic power (power for maintaining auxiliary systems, compressing air etc.) requirements of compressing air as this energy is taken from the power output of the power arrangement of the aircraft. Obtaining suitably compression ratios at varying altitudes is particularly troublesome and arrangements below provide an advantageous approach to achieving suitable electrical power for aircraft at altitude.

Referring now to Figure 1, there is shown a schematic view of a turbo-compressor arrangement according to examples disclosed herein. The turbo-compressor arrangement 100 comprises at least one axial compressor 110 and at least one centrifugal compressor 120. The turbo-compressor 100 is therefore a multi-stage compressor that includes both an axial compression and a centrifugal (also referred to as "radial") compression.

In examples, the turbo-compressor 100 may have a plurality of axial compressors and/or a plurality of centrifugal compressors. The arrangement may be such that the axial compressor (or compressors) are upstream of the centrifugal compressor (or compressors). While in examples there may be axial compressors arranged downstream of the centrifugal compressors, we have found that increased advantageous efficiencies in arrangements where the air is turned fewer times. In the arrangement 100 of Figure 1, therefore, the axial compressor 110 is upstream of (closer to the fluid source than) the centrifugal compressor 120.

The fluid source may be an inlet for ambient air or may be a fluid container such as an oxygen or hydrogen source or the like. The fluid source may have a temperature for fluid of ambient or at a predetermined store temperature.

Also shown in the example of Figure 1, there is a motor 130, a stator 140 (downstream of the motor 130) and a turbine 150 (downstream of the stator 140). The stator 140 may be a variable geometry stator 140 and this will be discussed in more detail below.

The motor 130 acts to provide an additional rotation to the shaft on which the arrangement lies. For smaller aircraft or during ground operations a smaller motor may be used and operated at a very high speed. For larger aircraft (the like of which can benefit from the present arrangement particularly) more power-dense arrangements can be used.

In specific examples, motors up to 200kW may be used for e.g. up to 100 seater aircraft.

In terms of operation and the arrangement in the example of Figure 1, the motor is used in combination with the turbine to turn the shaft. Gas turbine engines derive power for their compressors from the turbine section of the engine. This is not the case for an aircraft fuel cell system, the motor is therefore advantageously used to produce shaft power to drive the compressor, while the turbine provides an assistive contribution. The motor also provides control over the speed of the compressor according to the required operating conditions, including during transient operation.

While the motor may be heavy, which is counterintuitive to an aim of the present invention of enabling efficient green flight, the present arrangement accounts for the additional mass.

The present arrangement has an axial turbine which is controllable using variable geometry. The present arrangement has an efficient and controllable compressor architecture with variable geometry and inter-stage cooling which is also integrated with the cryogen system for additional efficiencies. The present arrangement reduces the need for compressor handling bleeds, and re-uses the bleed air in the turbine where advantageous. The present system has a single-spool system (more mass-efficient to have one motor instead of many), with corresponding fewer losses. As such, the present arrangement provides a series of advantageous elements when considered in combination.

In the example shown in Figure 1, the arrangement 100 further comprises an inlet guide vane (IGV) 160. In examples, the IGV 160 is variable. In order to improve the turbo-compressor operability, variable IGVs may be used before the first compressor stage 110. This broadens the compressor maps and protects the compressor from surge particularly at take-off reducing significantly the need to use bleed air. The axial turbine stage 150 also has a variable geometry stator stage 140 to be able to recover more energy during the different points in the mission and operating conditions. This feature avoids using inefficient waste-gates and can recover more energy if the compressor needs to bleed more air at low power In the example shown in Figure 1, the (at least one) axial compressor 110 may be a two stage axial compressor 110. The axial compressor 110 may comprise a stage one compressor 1101 and a stage two compressor 1102. Further stages or compressors may be used in relation to the compression ratios desired from the arrangement.

In the example shown in Figure 1, the arrangement 100 further comprises inter-stage cooling 180.

The arrangement 100 improves efficiency, operability and provides a more compact design compared to alterative two spool or modern centrifugal machines.

In use, the fluid is provided to the axial compressor arrangement 100 via the IGV for compression. The axial compressor arrangement 100 provides a compression ratio per stage of axial compressor. This may be around 1.1 to 2 per axial stage. In a specific example, the compressor arrangement 100 (including axial and centrifugal) provides a total compression ratio of about 10. The ratio of the centrifugal compression may be around up to 2.5 while the axial compression may be around up to 4. This may be formed in a specific example by three axial compressors with a ratio of around 1.3 each or so. The axial compressor arrangement 100 may compress the fluid provided to it to around 1.5 to around 2.0 bar. This has been found to be particularly suitable for larger sized aircraft such as a commercial aircraft that may seat around 48 passengers at an altitude of around 30,000 to around 40,000 ft.

In the arrangement herein, the absolute pressure required by the fuel cells (or, of course, for total clarity, fuel cell stack or stacks) is irrespective of aircraft size but the numbers of cells/stacks and the overall airflow requirement go up with the power demands. The increased airflow rate needed by the larger aircraft platforms advantageously uses axial compressor stages.

The earlier stages in the arrangement 100 (e.g. the first stage of two or the first and second stages of e.g. five) handle lower pressures and therefore can provide greater proportional compression. Latter stages are required to increase the pressure such that the fluid is of a reasonable pressure for use when the inlet air is very low pressure (such as when at altitude). The higher the top altitude expected for the aircraft, the greater number of compressors or compressor stages are to be used in the arrangement 100. In order to avoid reducing the overall efficiency, pressure ratios for each stage are not too high (not often above 2). The arrangement may be a single spool arrangement with 2, 3 or 4 (or more) compression stages.

Due to the relatively low volumetric flow, the final stage may be a centrifugal stage (i.e. centrifugal compressor 120). The centrifugal compressor 120 may be combined with the inter-stage cooling unit 180. In particular, before the last compression stage an inter-stage cooling 180 may be integrated into the compressor casing 120 (or provided separately).

In more detail, radial compressor mass flow may be the same as through the axial stages but it is of a higher inlet pressure so radial compressor size is volumetrically smaller. An axial compressor would struggle to perform as desired as blade geometries become too small and construction is made far more difficult. Instead, we exploit the higher pressure ratio capability of the radial stage with the lower volumetric flow rate.

To reduce the coolant temperature used in the inter-stage cooling, the coolant may be cooled with cryogenic hydrogen before entering the inter-stage cooling. Such hydrogen will be available in a propulsion arrangement using electrical power generation from a fuel cell. The coolant flow may be regulated so as to reduce the air temperature in the compressor as much as possible. This provides a highly efficient integration of the power management system with the cryogenic fuel system to supply cold coolant for interstage cooling.

The effect of the increased operating altitude (during flight for example) is that the amount of compressor work required to provide the desired pressure increases, and in turn results in a larger temperature increase in the compressed air. As such, the amount of cooling required via the inter-stage cooler is much larger. The present inter-stage cooler arrangement may favourably utilise the cryogenic hydrogen to synergistically provide fuel to the fuel cell and maintain suitable temperatures in the compression stages.

During cold days, the compressor will not raise the temperature as much as on warm days and the stack inlet temperature may be below freezing. The inter-stage cooling can then be used to pre-heat the air and also to reduce the differences in the corrected flow between hot and cold days. The corrected flow is an equivalent flow rate expressed at a reference pressure and temperature, normally sea level ISA, which can make compressor performance comparisons more straightforward.

A single stage turbine 150 is installed on the same shaft to recover energy from the Fuel Cell exhaust (not present in the arrangement of Figure 1 but present in the larger electrical aircraft propulsion arrangement). The turbine 150 is an axial single stage to improve the speed compatibility with compressor stages and recover more energy compared to a centrifugal stage alone.

The electrical aircraft propulsion arrangements shown herein are operable with fuel cells or other electrical generation elements. The arrangements may be used in aircraft that are exclusively electrically powered or for aircraft that are hybrid aircraft, i.e. a partially electrically powered aircraft.

Referring now to Figure 2, there is a shown a schematic of an electrical aircraft propulsion arrangement comprising a turbo-compressor arrangement according to examples. The propulsion arrangement 200 shown in Figure 2 has a turbo-compressor arrangement 201. The turbo-compressor arrangement 201 is the arrangement of Figure 1 (albeit connected to the remaining additional elements within arrangement 200).

The arrangement 200 has an air intake 202 for providing air to the arrangement 200. The air intake 202 is connected to the turbo-compressor arrangement 201. The turbo-compressor arrangement 201 is further connected to an exhaust nozzle arrangement 203 and a fuel cell stack portion 204. The inlet guide vanes of the compressor arrangement 201 receive air from the intake 202. The air intake 202 comprises an air inlet and an air filter. The air filter is arranged to filter air ultimately provided to the at least one fuel cell of the electrical aircraft propulsion arrangement 200.

Fuel cells are sensitive to environmental pollution and contaminants, therefore filter prevents contamination and poisoning the catalyst in the stack. In particular, the filter may remove sulphur dioxide and particulate matter. As such, for longevity of life and increased performance, inclusion of an air filter is advantageous. The air intake 202 is in fluid communication, in relation to the air for use in the overall system, with the turbo-compressor 201.

The turbo-compressor 201 is connected to a motor-controller heat sink and coolant pump arrangement 205. Specifically, this arrangement 205 is connected to the motor of the turbo-compressor arrangement 201 (shown as motor 130 of arrangement 100 of Figure 1). The coolant pump arrangement 205 provides coolant for thermal regulation of certain elements within the system 200.

The turbine (shown as turbine 150 of arrangement 100 of Figure 1) of the turbo-compressor arrangement 201 is connected to an exhaust nozzle arrangement 203. The exhaust nozzle arrangement 203 has a water separator and a cathode exhaust. The exhaust nozzle arrangement 203 may include or be additionally connected to a vent.

The turbo-compressor arrangement 201 is connected to a fuel cell stack portion 204 comprising various elements that are used to provide power for the (at least partially) electrically powered aircraft propulsion system disclosed herein.

The fuel cell stack portion 204 comprises various thermal exchangers such as intercoolers, superheaters, radiators, and heat sinks that may use coolants. This arrangement may be in fluid communication with the inter-stage cooling portion of the turbo-compressor arrangement 201. The superheater may act as an intermediary heat exchanger between the hydrogen supply in the arrangement and other higher temperature coolants. Other coolants may be, for example, the coolants used in the fuel cell stack and the inter-stage cooler prior to the centrifugal compressor.

The fuel cell stack portion 204 comprises various pumps, pressure regulators, injectors and valves that may use gas such as hydrogen alongside at least one fuel cell or at least one fuel cell stack.

The fuel cell stack portion 204 also comprises various components that allow transport of coolant which may be water to and from the fuel cell stack for cooling or removal from the system.

The arrangement 200 provides a high efficiency, single spool arrangement that can provide sufficient electrical power for commercial aircraft. The arrangement 200 utilises benefits from air filtration, thermal control of the air alongside pressure manipulation as required considering the stage of flight. Each of these provides a user with great control over the system and the propulsion provided therefrom.

The arrangement 200 is held on a single spool and therefore the activation of elements is not dependent on altitude rather it is the operating conditions of the elements within the arrangement 200 that change, as controlled by a user or a controller. These change based on changes in the ambient air condition and the flight stage and power requirements.

The system uses controllers to enable a highly efficient system. To set-up preferential operating conditions for the air delivery system at a good level of efficiency, the pressure at the stack inlet, the stack air flow and the compressor bleed air flow are controlled. Bleed air may be used at low flow power settings or during cold start (controlled by a controller). Suitable locations for the control sensors are shown in Figure 3. The cathode exhaust has a variable geometry to reduce the drag at low power and needs to be controlled depending on the ambient pressure, compressor flow and air speed.

Figure 3 shows a schematic similar to Figure 2 with controller locations proposed.

Arrangement 300 indicates locations for a mass flow meter 391 near the fuel cell stack, a mass flow meter 392 near the turbo-compressor arrangement and a static pressure sensor 393.

Target air flow in the fuel cell stack may be by a controller based on the stack current demand. A PID controller on the upstream valve with a flow meter feed-back loop can be used to control the air flow going into the stack. These provide additional controls over the performance of the arrangement.

Based on the target pressure (fuel cell air input pressure) for the power demand and altitude, the controller will command a target speed to the compressor and a PID control loop will be used to control the pressure based on a feed-back loop from the pressure sensor. The turbo-compressor motor may set the target speed by using a torque feed-back loop. The variable geometry settings can be linked to the compressor speed so as to increase the surge margin at lower flow. Additional or alternative options to improve efficiency involve linking the variable geometry setting to the speed and corrected flow.

The total compressor flow, including bleed air, may be controlled with a closed loop PID with the turbine variable geometry setting after the compressor outlet. For a given compressor speed, if the demanded stack flow is lower than the surge margin, then the target flow may be increased accordingly until the total flow measured with the flow meter in (or proximal to) the compressor is higher than the minimum needed to keep a surge margin. Based on the target flow in the compressor, air speed and ambient pressure, the variable geometry exhaust may be set-up to a defined area determined from known tested tables.

Air condition is controlled within the arrangement disclosed herein. Air condition includes pressure, temperature and humidity inter alla. Humidity control is a relevant consideration for some low temperature membrane stacks. Means of controlling humidity involve using humidifiers, recirculating hydrogen and decreasing the coolant temperature.

The air temperature at the fuel cell stack inlet is regulated to avoid creating uneven temperature distribution, overheating or risk of frost during cold days. The air temperature is controlled to be sufficiently high to avoid lowering the stack temperature below the minimum at which it can 20 operate.

The stack temperature may be controlled with a liquid cooled system. The stack temperature may be kept at the target values to ensure that the stack operates efficiently, above the minimum operating temperature and below the maximum allowed. The coolant should not be overheated or allowed to freeze as this advantageously reduces the likelihood of damage to the system from the coolant.

The stack 204 may comprise a heater or superheater 2042 arranged to provide heat to cryogenic fuel entering the fuel cells such that the fuel is conditioned suitable for the fuel cell.

Superheater as used herein primarily relates to a heat exchanger arranged to provide heat to superheat a gaseous state fluid (as will be understood in the art). The cryogenic fuel may enter the superheater 2042 in a partially or fully gaseous state, and exit the superheater in a full gaseous state and at a higher temperature. In this context, superheating refers to the addition of heat to the cryogenic fuel, such that the cryogenic fuel is above its boiling point temperature when leaving the superheater. The cryogenic fuel may be a fluid at a temperature higher than its saturation temperature after superheating.

The superheater 2042 may be a liquid-gas heat exchanger utilising a liquid coolant and the cryogenic fuel as the working fluids, wherein the liquid coolant enters the superheater at a higher temperature than the gaseous cryogenic fuel as it enters it. As shown in Figure 2 (and discussed previously), the coolant for the inter-stage cooler may be the liquid coolant of the superheater 2042 such that, in use, the cryogenic fuel (hydrogen in this example) removes heat from the inter-stage cooler, thereby cooling the corresponding compressor stage(s). Simultaneously, the cryogenic hydrogen is favourably conditioned suitable for use by the fuel cells by the waste heat of the compressor(s), providing an extremely efficient aircraft.

Alternatively or additionally, the liquid coolant is arranged to remove heat from the fuel cells and provide the heat to the cryogenic fuel via the superheater in much the same way as the inter-stage cooler. This further improves the overall efficiency of the aircraft.

Temperatures to be controlled include the turbo-compressor motor and controller temperature, the stack coolant temperature, the stack inlet temperature and the air temperature coming out of the turbo-compressor intercoolers. Temperature sensors can be arranged throughout to obtain relevant data on these. The motor-controller may reject heats with a separate radiator compared to the stack coolant due to overly high temperature target temperature differences. The coolant control can be set with a PID with feed-back loop on the coolant temperature outlet. The pump speed can be changed until the coolant temperature reaches the target values.

The inter-stage cooling may be as much as possible in counter-flow so as to be able to reduce the air temperature significantly.

The inter-stage cooling could be used as air pre-heater during cold days without the need for electrical heating. This reduces the differences between cold and hot days for the last compressor stages, which can be advantageous for lifetime as protecting against the wear and tear of experiencing a wider temperature range of air. The temperature control can be achieved with a PID feed-back loop where the valve regulating the flow can be opened and closed until the target is reached. Where the coolant temperature for the stack is too high, the coolant flow may be cooled with cryogenic hydrogen prior to entering the inter-stage cooler.

The post stage intercooler coolant temperature is preferably maintained at a target temperature below the boiling point for the system to function. The coolant temperature can be controlled by changing the coolant flow. A PID with closed loop on the coolant outlet temperature can be used. The valve is opened or closed until the coolant reaches the desired temperature.

The stack coolant inlet and outlet temperatures may be controlled to improve functioning and efficiency of the system. The coolant outlet temperature can be controlled by changing the flow with different speed command to the pump. The radiator can have multiple control features such as by-pass valves, fans and water spraying a complex control logic with priority order system if needed. Water spaying can be used in the take-off phase above 10 degrees air temperature to avoid freezing. This is preferential for safety concerns. The radiator inlet may be set to predefined values based on power demand, ambient air temperature, pressure and true air speed. It may, therefore, not be used to control coolant temperature.

Depending on whether the coolant temperature is too low or high, the priority order may be changed. If the coolant temperature is above target, the bypass fully diverts the coolant through the radiator and the fan is set to idle as default setting. The duct exhaust keeps opening until the temperature is reached, otherwise the fan speed is increased once it reaches the maximum opening. If exhaust is fully opened and fan at full speed and the altitude and temperature conditions are suitable, water spraying may then be used to lower the temperature. If the coolant temperature is too low, water spraying is disabled, fan is set to idle, then exhaust is fully closed, then fan is turned-off and then the radiator is partly bypassed. This can be controlled by a controller receiving information from the various sensors arranged in the system to provide improved efficiencies and safety from the arrangement.

Hot and cold day requirements are relevant in the present arrangement for aircraft propulsion. The balance of plant start-up procedure can, therefore, be different depending on whether it is a hot day or a cold day.

On a hot day, the coolant has already a certain temperature and compressor may not inject cold air into the stack. It is conceivable though that the stack may still be too cold. At the start there is no electrical loading on the stack and hydrogen is not circulated. The heat sink is fully bypassed and the coolant pump is set to low flow. Electrical heating is switched on the stack.

The compressor is started with the batteries and it starts heating-up the coolant together with the electrical heating of the stack. When the coolant reaches a certain temperature, the stack is started with hydrogen supply and electrical loading ramps-up, the stack electrical loading is set to idle and the coolant starts flowing into the radiator.

For a cold start, the compressor may be started by batteries (or other similar power store, possibly via combustion). The stack and coolant expansion tanks may have electrical heating turned-on, the radiator is bypassed. When the coolant is above zero degrees in the expansion tank, it starts flowing into the stack and intercoolers. When the air at the stack inlet is above around 10 degrees, it starts flowing into the stack. When the coolant reaches a certain temperature, electrical loading is applied on the stack with hydrogen supply and electrical heating in the coolant is switched off. When the coolant reaches a certain temperature at the stack outlet, the system is up and running and the coolant can flow into the radiator. Initially the coolant flows at a lower rate which may be gradually increased until it reaches steady state.

The balance of plant shut-down procedure may follow a certain order and ramp-down sequence to reduce the likelihood of damage to the system. First, the electrical loading and balance of plant is set to minimum power, then electrical loading and hydrogen supply may be cut. The compressor keeps flowing air for some short period of time (e.g. a few seconds) to remove any water from the plant, then it is shut-down. The coolant keeps flowing until it reaches the shut down temperature at the stack outlet.

This arrangement provides high levels of variability if the output of the system and the above provides a series of controller behaviours to provide high levels of reliability and performance from the arrangement. These arrangements use controllers to manipulate the performance including of the cooler arrangements.

The above relates to controlling the performance of fuel cells in light of changing conditions of air during flight. Temperatures may be preferably around 80 to 90 °C. Other aspects of control include the variable IGVs. Varying the guide vanes allows for a change in compressor ratio. Variable as used herein primarily relates to the physical arrangement of the guide vanes. The stack pressure is difficult to control in itself and therefore bypass valves may be used to achieve a high operating range.

Valves or a valve in the above arrangement may be used to provide stack bypass alongside providing compressor handling bleed. Not all of the combinations of air flow rate and pressure ratio needed by the fuel cell stacks during the aircraft mission are accessible by the compressor -there is a limitation imposed by the compressor map, and particularly the "surge line". In order to maintain acceptable margin to surge in these conditions, we accept a higher flow rate from the compressor than is required, and "bleed" the excess off further downstream.

The bleed air may therefore be controllably sent to the turbine (rather than straight to the exhaust). This increases the airflow through the turbine (therefore increasing potential for energy recovery) and if the turbine is mechanically coupled to the compressor it may reduce any deficit between the turbine speed and the turbine inlet flow (improving in turn the efficiency of turbine).

This allows the arrangement to have a highly efficient compressor whilst still achieving the required airflows. For a fuel cell system (as per the above arrangement, rather than e.g. a gas turbine arrangement) on an aircraft where there are no systems that require "bleed air", the need to bleed the compressor for operability reasons may be seen as a disadvantage, because it increases fuel consumption and reduces efficiency of the balance of plant. However, this is used advantageously herein (against the prevalent thinking in such areas) to allow greater benefits in total from the system.

The above utilises a single spool in the arrangement for material savings. The integration of the turbo-compressor into the overall arrangement provides a significant mass saving and is therefore found to be advantageous in this instance.

In order to address various issues and advance the art, the entirety of this disclosure shows by way of illustration various embodiments in which the claimed invention(s) may be practiced and provide for a superior propulsion system for at least partially electrically propelled aircraft. The advantages and features of the disclosure are of a representative sample of embodiments only, and are not exhaustive and/or exclusive. They are presented only to assist in understanding and teach the claimed features. It is to be understood that advantages, embodiments, examples, functions, features, structures, and/or other aspects of the disclosure are not to be considered limitations on the disclosure as defined by the claims or limitations on equivalents to the claims, and that other embodiments may be utilised and modifications may be made without departing from the scope and/or spirit of the disclosure. Various embodiments may suitably comprise, consist of, or consist essentially of, various combinations of the disclosed elements, components, features, parts, steps, means, etc. In addition, the disclosure includes other inventions not presently claimed, but which may be claimed in future.

Further examples of feature combinations taught by the present disclosure are set out in the following set of numbered clauses: 1. An electrical aircraft propulsion arrangement comprising: at least one fuel cell arranged to generate electrical power for use in propelling an aircraft; a fluid source arranged to provide fluid to at least one fuel cell for use in generating electrical power; at least one centrifugal compressor arranged to compress a fluid prior to provision to at least one fuel cell; at least one axial compressor arranged to compress a fluid prior to provision to at least one fuel cell, wherein at least one axial compressor is upstream of at least one centrifugal compressor.

2. An electrical aircraft propulsion arrangement according to clause 1, further comprising an air filter to filter air provided to the at least one fuel cell. 10 3. An electrical aircraft propulsion arrangement according to clause 1 or 2, wherein at least one axial compressor comprises a two-stage axial compressor, wherein each stage offers a ratio of compression around 1.1 to 1.7.

4. An electrical aircraft propulsion arrangement according to any of clauses 1 to 3, further comprising an outlet arrangement comprising: an outlet for providing fluid communication between the electrical aircraft propulsion arrangement and an outer environment; and, a vent.

Such an arrangement allows for control over the distribution and release of output fluids from the energy generating system within the aircraft. For example, fuel cells may release water and some gas. These fluids can be controllably released from the system via the outlet and the vent. Vent allows control over the release of the fluids. Release of water onto a runway in very cold conditions, for example, is less advantageous than in a more controlled manner.

5. An electrical aircraft propulsion arrangement according to any preceding clause, wherein at least one axial compressor is arranged to receive fluid from the fluid source before at least one centrifugal compressor.

This avoids turning the fluid unnecessarily. It is preferential to maximise the benefits of both systems to have a centrifugal compressor as the last step. In most examples above, though not all, the system has a series of axial compressors and finally as singular centrifugal compressor.

6. An electrical aircraft propulsion arrangement according to any preceding clause, wherein the at least one centrifugal compressor and the at least one axial compressor are mechanically coupled.

Other mechanical couplings are possible. This improves the stability of the arrangement particularly between two operational compressors. In this way, the arrangement may use for example one shaft on which one centrifugal compressor and one axial compressor are arranged. This can be used to provide a compact and light arrangement avoiding the use of multiple shafts. Weight savings are of significant importance in the propulsion arrangement being present in a commercial aircraft.

7. An electrical aircraft propulsion arrangement according to clause 6, wherein the mechanical coupling is provided by a shaft.

8. An electrical aircraft propulsion arrangement according to any preceding clause, wherein the axial compressor comprises variable guide vanes.

9. An electrical aircraft propulsion arrangement according to any preceding clause, further comprising at least one of a turbine and a restrictor.

Use of a turbine increases the overall efficiency of the propulsion arrangement by utilising what may otherwise be waste energy. This is one of a number of synergistic components in the arrangement that improve efficiency of the electrical propulsion and therefore increases the viability of the arrangement for use in green commercial aircraft.

10. An electrical aircraft propulsion arrangement according to clause 9, wherein at least one of the turbine and restrictor are arranged on an exhaust flow portion of the electrical aircraft propulsion arrangement and are arranged to provide a pressure difference between the arrangement and external environment conditions.

11. An electrical aircraft propulsion arrangement according to clause 9 or 10, further comprising a variable geometry stator as part of the at least one of the turbine.

This feature further allows greater control for the user over the recovery of waste energy produced during the electrical-motive power generation. These variable aspects synergistically allow a user to tailor the recovery between different stages in the mission (flight stages) and operation conditions (e.g. low quality, low pressure air or high quality, high pressure air, etc.).

Having a variable geometry turbine is particularly advantageous where the turbine is mechanically coupled to the compressor (as per the above single-shaft arrangement), because the turbine speed is governed by the compressor speed. Variations in the exhaust flow rate through the turbine (e.g. when we vary the power level demanded from the fuel cells) can then be difficult to accommodate, limiting energy recovery. Having the variable geometry enhances of the off-design performance of the turbine, which is particularly useful in an aircraft fuel cell context.

12. An electrical aircraft propulsion arrangement according to any of clauses 9-11, wherein the at least one turbine is an axial turbine.

The design space is generally larger for an axial turbine -i.e. the designer can more readily optimise the design to the expected exhaust flow and rotational speeds. This means we can access higher efficiency of turbine recovery at the turbine design point (e.g. cruise) than with a radial compressor. If designed for cruise, we can exploit this to minimise fuel consumption.

13. An electrical aircraft propulsion arrangement according to any preceding clause, further comprising a control bypass valve.

14. An electrical aircraft propulsion arrangement according to any preceding clause, wherein the at least one axial compressor and the at least one centrifugal compressor each further comprise an inter-stage cooler arrangement.

15. An aircraft comprising the electrical aircraft propulsion arrangement of any of clauses 1-14.

16. A method of providing electrical power for aircraft propulsion comprising: providing fluid from a fluid source for use in generating electrical power; compressing, by at least one axial compressor, the fluid from the fluid source; receiving, by at least one centrifugal compressor, the fluid from at least one axial compressor; compressing, by the at least one centrifugal compressor, the fluid from the fluid source; and, providing the fluid to at least one fuel cell arranged to generated electrical power for use in propelling an aircraft.

17. A method according to clause 16, further comprising: conditioning the fluid prior to receiving the fluid at the at least one fuel cell, wherein conditioning comprises at least one of: filtering the fluid; altering a humidity of the fluid; and, altering a temperature of the fluid.

18. A method according to clause 16 or 17, wherein compressing, by at least one axial compressor, the fluid from the fluid source comprises: compressing, via a first stage axial compressor; and, compressing, via a second stage axial compressor, wherein each stage offers a ratio of compression around 1.1 to 1.7.

19. A method according to any of clauses 16-18, further comprising mechanically coupling the at least one centrifugal compressor and the at least one axial compressor.

20. A method according to any of clauses 16-19, further comprising providing the generated electrical power to a propulsor for providing thrust for an aircraft.

Claims

CLAIMS1. An energy management system for an electrically powered aircraft, the system comprising: at least one fuel cell arranged to generate electrical power; a fluid source arranged to provide fluid to at least one fuel cell for use in generating electrical power; at least one centrifugal compressor arranged to compress a fluid prior to provision to at least one fuel cell; at least one axial compressor arranged to compress a fluid prior to provision to at least one fuel cell, wherein at least one axial compressor is upstream of at least one centrifugal compressor.
2. An energy management system according to claim 1, further comprising an energy demand sensor arranged to detect an energy demand associated with an electrically powered aircraft, a controller arranged to control an amount of compression of fluid from the fluid source by the at least one centrifugal compressor and the at least one axial compressor in response 20 to an energy demand detected by the energy demand sensor.
3. An energy management system according to claim 2, further comprising a pressure sensor, wherein the pressure sensor is arranged to provide signals to the controller on a pressure of fluid in the energy management system and wherein the controller is arranged to control an amount of compression of fluid in response to the signals from the pressure sensor.
4. An energy management system according to claim 2 or 3, further comprising a motor, wherein the controller is arranged to control the speed and/or rotation of the motor in response to an energy demand detected by the energy demand sensor.
5. An energy management system according to any of claims 2 to 4, wherein at least one axial compressor comprises variable guide vanes comprising a controllably variable geometry, and wherein the controller is arranged to control the variable geometry of the at least 35 one axial compressor in response to an energy demand detected by the energy demand sensor.
6. An energy management system according to any of claims 2 to 5, further comprising a temperature management arrangement comprising a heater arranged to provide thermal energy to a fuel for the at least one fuel cell, and a temperature sensor arranged to detect a temperature of a fluid from the fluid source, wherein the controller is arranged to control the provision of thermal energy by the heater in response to signals from the temperature sensor and the energy demand sensor.
7. An energy management system according to claim 6, wherein the temperature management arrangement further comprises an inter-stage cooler arranged between the at least one axial compressor and the at least one centrifugal compressor, wherein the inter-stage cooler is arranged to cool a fluid from the fluid source prior to compression by the at least one centrifugal compressor, wherein the controller is arranged to control the provision of cooling by the inter-stage cooler in response to signals from the temperature sensor and the energy demand sensor.
8. An energy management system according to any of claims 2 to 7, further comprising a turbine arranged to receive an output fluid from the at least one fuel cell.
9. An energy management system according to claim 8, wherein the turbine is at least one of: a variable geometry turbine, wherein the controller is arranged to control the variable geometry of the turbine; and, an axial turbine.
10. An energy management system according to claim 8 or 9, further comprising a bypass valve arranged to control the passage of fluid into and out of the at least one fuel cell, wherein the controller is arranged to control the bypass valve in response to an energy demand detected by the energy demand sensor.
11. An energy management system according to claim 10, wherein the bypass valve is controllable by the controller to provide fluid to the turbine in response to an energy demand detected by the energy demand sensor.
12. An energy management system according to any of claims 2 to 11, further comprising an air filter to filter fluid provided to the at least one fuel cell.
13. An energy management system according to claim 12, further comprising an air quality sensor to detect an air quality, wherein the controller is arranged to control fluid to pass through the air filter in response to signals from the air quality sensor associated with an air quality below a predetermined air quality.
14. An energy management system according to any of claims 2 to 13, further comprising a flight stage sensor arranged to detect a flight stage associated with an electrically powered aircraft, wherein the controller is arranged to control an amount of compression of fluid from the fluid source by the at least one centrifugal compressor and the at least one axial compressor in response to a flight stage detected by the flight stage sensor.
15. An energy management system according to any of claims 2 to 14, further comprising an electrical energy store to store electrical energy, wherein the controller is arranged to provide energy from the electrical energy store in response to an energy demand detected by the energy demand sensor.
16. An energy management system according to any of claims 1 to 15, wherein at least one axial compressor comprises a multi-stage axial compressor, wherein each stage offers a ratio of compression around 1.1 to 1.7.
17. An energy management system according to any of claims 1 to 16, wherein the at least one centrifugal compressor and the at least one axial compressor are mechanically coupled and wherein the mechanical coupling is provided by a shaft.
18. An electrical aircraft propulsion arrangement comprising the energy management system of any of claims 1 to 17.
19. An aircraft comprising the electrical aircraft propulsion arrangement of claim 18.
20. A method of managing energy provision for an electrically powered aircraft, the method comprising: providing fluid from a fluid source for use in generating electrical power; compressing, by at least one axial compressor, the fluid from the fluid source; receiving, by at least one centrifugal compressor, the fluid from at least one axial compressor; compressing, by the at least one centrifugal compressor, the fluid from the fluid source; and, providing the fluid to at least one fuel cell arranged to generated electrical power for use in propelling an aircraft.