Patent Description:
<CIT> discloses a hot gas generator system which includes a combustor and a condenser, with the combustor connected to the condenser for condensing the product of combustion from the combustor. A hydrogen supply is connected to the condenser and then to the combustor whereby the hydrogen absorbs heat from the combustion product as it condenses and the hydrogen thereby is preheated prior to entering the combustor. An oxygen supply is connected to the combustor for mixing with the hydrogen during combustion. The combustor is part of an integrated heat exchanger/combustor whereby a minor portion of the hydrogen passing through the condenser is used in the combustor for burning purposes and a major portion of the hydrogen is passed through the combustor for superheating the hydrogen prior to delivering the hydrogen to a prime mover, such as a thruster or a turbogenerator of a space platform or the like.

United States patent application <CIT> discloses a method for preparing deep-frozen liquid gas for the purpose of recovering process energy for a downstream process, with which the refrigerating capacity of the deep-frozen liquid gas can also be used in the downstream process. In the disclosure, this is achieved by the fact that the refrigerating capacity of the deep-frozen liquid gas is fed as a heat sink to at least one of the part-steps of the downstream process via at least one heat-exchange medium and, if said heat-exchange medium is not available, the deep-frozen liquid gas is regasified with an additional heat-exchange medium.

United Kingdom patent application <CIT> discloses an engine suitable for high speed (up to say Mach <NUM> cruise) or a trans atmospheric vehicle which comprises a turbojet having an air compressor, a combustor, a first turbine which drives the compressor and a jet pipe which terminates in a variable area nozzle. A source of liquid hydrogen is provided, and the engine has a first heat exchanger cooled by the fuel for cooling the air entering the intake. A bypass valve is provided so as to bypass the heat exchanger to prevent icing in high humidity air. The hydrogen is vaporised in a second heat exchanger in the jet pipe and the gaseous fuel is used to drive the liquid fuel turbo pump and a second turbine which also drives the air compressor. The exhaust from the second turbine is fed to the main combustor to a reheat burner in the jet pipe. Liquid oxygen can be used to cool further the intake air or to cool the turbine components.

European patent application <CIT> discloses a gas turbine engine system which includes a gas turbine engine and a fuel turbine system. The gas turbine engine includes an air inlet, compressor, combustor, turbine, and heat exchange system. The heat exchange system is configured to transfer thermal energy from an inlet air flow or exhaust air flow to a fuel to produce a gaseous fuel that is used to drive a fuel turbine and fuel pump and used for combustion in the gas turbine engine. The fuel turbine is in fluid communication with the heat exchange system and the combustor and configured to extract energy from expansion of the gaseous fuel. The fuel pump is configured to be driven by the fuel turbine and is in fluid communication with the heat exchanger system.

This disclosure relates to fuel delivery for hydrogen-fuelled aero gas turbine engines.

In order to limit emissions of carbon dioxide, use of hydrogen as an alternative to hydrocarbon fuel in gas turbine engines has historically only been practical in land-based installations. Such engines are typically supplied with hydrogen derived from natural gas via concurrent steam methane reformation, which hydrogen is injected into large-volume series staged dry low NOx burners. This type of burner is not suitable for use in an aero engine primarily due to its size and the difficulties in maintaining stable operation during transient manoeuvres.

Experimental programmes have been conducted to develop aero engines operable to be fuelled with hydrogen, however these have typically been high-Mach afterburning turbojets or expander cycles and thus not practical for use on civil airliners operating in the Mach <NUM> to <NUM> regime.

There is therefore a need for technologies for combustion of hydrogen in aero gas turbine installations, in particular around the overall engine cycle to for example minimise fuel consumption, the fuel delivery system to for example meter fuel accurately, and the fuel injection system to for example minimise emissions.

The invention is directed towards fuel delivery systems for delivering hydrogen fuel from a cryogenic storage system to a fuel injection system in a gas turbine engine, gas turbines incorporating such fuel delivery systems, and methods of delivering hydrogen fuel from cryogenic storage systems.

One such fuel delivery system according to claim <NUM> includes a pump, a metering device and a fuel heating system for heating the hydrogen fuel to an injection temperature for the fuel injection system, the fuel heating system comprising a vaporiser provided between the pump and the metering device configured to vaporise liquid hydrogen from the cryogenic storage system; characterised in that the metering device is one of:.

In an embodiment, the vaporiser comprises a fuel offtake for diverting a portion of the hydrogen fuel from a fuel conduit for combustion in a burner located in heat exchange relationship with the fuel conduit.

In an embodiment, the burner is configured to receive pressurised air from a compressor of the gas turbine engine for combustion with the portion of the hydrogen fuel.

In an embodiment, the vaporiser comprises a boil volume or an electric heating element for initial heating of liquid hydrogen if no vaporised hydrogen fuel is available for combustion.

In an embodiment, the heater comprises a fuel offtake for diverting a portion of the hydrogen fuel from a fuel conduit for combustion in a burner located in heat exchange relationship with the fuel conduit.

In an embodiment, the heating system comprises one or more heat exchangers for heating the hydrogen fuel by heat from the gas turbine.

In an embodiment, the one or more heat exchangers are oil-fuel heat exchangers for cooling engine oil or gearbox oil from the gas turbine engine by the hydrogen fuel.

One such gas turbine engine according to claim <NUM> comprises a combustor, a fuel injection system, and a fuel delivery system of the first aspect.

In an embodiment, the burner is configured to receive pressurised air from a compressor of the gas turbine engine for combustion with the portion of the hydrogen fuel.

One such method according to claim <NUM> of delivering hydrogen fuel from a cryogenic storage system to a fuel injection system in a gas turbine engine comprises:.

characterised in that the method comprises:
heating, by a heater, the hydrogen fuel to the injection temperature following metering by the metering device , wherein the metering device is one of:.

In an embodiment, the heating of the hydrogen fuel comprises vaporising the fuel by:.

A hydrogen-fuelled airliner is illustrated in <FIG>. In this example, the airliner <NUM> is of substantially conventional tube-and-wing twinjet configuration with a central fuselage <NUM> and substantially identical underwing-mounted turbofan engines <NUM>.

In the present embodiment, the turbofan engines <NUM> are geared turbofan engines. A hydrogen storage tank <NUM> located in the fuselage <NUM> for hydrogen fuel supply is connected with core gas turbines <NUM> in the turbofan engines <NUM> via a fuel delivery system. In the present embodiment, the hydrogen storage tank <NUM> is a cryogenic hydrogen storage tank and thus stores the hydrogen fuel in a liquid state, in a specific example at <NUM> kelvin. In this example, the hydrogen fuel is pressurised to a pressure from around <NUM> bar to around <NUM> bar, in a specific example <NUM> bar.

A block diagram identifying the flow of hydrogen fuel is shown in <FIG>.

Hydrogen fuel is obtained from the hydrogen storage tank <NUM> by the fuel delivery system <NUM> and supplied to each core gas turbine <NUM>. In the Figure, only one of the core gas turbines is shown for clarity. In this illustrated embodiment, the core gas turbine <NUM> is a simple cycle gas turbine engine. In other embodiments, as will be described with reference to <FIG>, complex cycles may be implemented via fuel-cooling of the gas path.

Referring again to <FIG>, the core gas turbine <NUM> comprises, in fluid flow series, a low-pressure compressor <NUM>, an interstage duct <NUM>, a high-pressure compressor <NUM>, a diffuser <NUM>, a fuel injection system <NUM>, a combustor <NUM>, a high-pressure turbine <NUM>, a low-pressure turbine <NUM>, and a core nozzle <NUM>. The high-pressure compressor <NUM> is driven by the high-pressure turbine <NUM> via a first shaft <NUM>, and the low-pressure compressor <NUM> is driven by the low-pressure turbine <NUM> via a second shaft <NUM>. It will be appreciated that in alternative embodiments, the core gas turbine could be of three-shaft configuration.

As will be described further with reference to <FIG> onward, the fuel injection system <NUM> may be a direct fuel injection system.

As described previously, in the present embodiment, the turbofan engines <NUM> are geared turbofan engines. Thus in operation the low-pressure turbine <NUM> drives a fan <NUM> via a reduction gearbox <NUM>. The reduction gearbox receives input drive from the second shaft <NUM> and provides output drive to the fan <NUM> via a fan shaft <NUM>. In an embodiment, the reduction gearbox <NUM> is an epicyclic reduction gearbox. In a specific embodiment, it is a planetary reduction gearbox. Alternatively, it may be a star reduction gearbox, or a compound epicyclic reduction gearbox. As a further alternative, the reduction gearbox <NUM> could be a layshaft-type reduction gearbox or any other type of reduction gearbox. It will also be appreciated that the principles disclosed herein may be applied to a direct-drive type turbofan engine, i.e. in which there is no reduction gearbox between the low-pressure turbine and the fan.

In operation, the fuel delivery system <NUM> is configured to obtain hydrogen fuel from the hydrogen storage tank <NUM> and provide it to the fuel injection system <NUM> in the core gas turbine <NUM>. <FIG> is a block diagram illustrating the fuel delivery system <NUM> in greater detail.

The fuel delivery system <NUM> comprises a pump <NUM>, a metering device <NUM>, and a fuel heating system for heating the hydrogen fuel to an injection temperature for the fuel injection system <NUM>. In an embodiment, a vent system (not shown) may be included in the fuel delivery system <NUM> close to the fuel injection system <NUM> to vent hydrogen fuel should a rapid shut-off be required, for example in response to a shaft-break event. It is envisaged that the vent system may vent the excess hydrogen fuel into the bypass duct of the turbofan engine <NUM>, or alternatively vent it outside of the nacelle of the engine <NUM>. An igniter may be provided to flare off the excess hydrogen in a controlled manner.

In the present embodiment, the pump <NUM> is high-speed centrifugal pump. In a specific embodiment, it is configured to operate at <NUM> rpm or more. In a specific embodiment, the centrifugal pump comprises an axial inducer to minimise the required inlet pressure and to accommodate multiphase flow in addition to the centrifugal impeller for developing the majority of the required pressure rise. In an alternative embodiment, a piston-type pump could be used.

In an embodiment, the pump <NUM> is located in the hydrogen storage tank <NUM>. In this way leakage of hydrogen fuel past pump seals etc. is accommodated.

In an embodiment, the pump <NUM> is driven by a fuel turbine, as will be described with reference to <FIG>.

Alternatively, the pump <NUM> could be driven by an air turbine supplied with compressor bleed, for example bleed from the high-pressure compressor <NUM>. Alternatively, combustion products from the combustor <NUM> may be used to drive a dedicated turbine for driving the pump <NUM>. In another embodiment, the pump <NUM> is driven via an electrical machine. In an embodiment, the drive means for the pump <NUM> are also located in the hydrogen storage tank <NUM>.

In this embodiment, the metering device <NUM> is configured to meter the required quantity of fuel for the current fuel demand of the core gas turbine <NUM>.

As will be appreciated, it is desirable to increase the temperature of the fuel from the <NUM> kelvin cryogenic storage condition to a temperature much closer to the firing temperature of the core gas turbine; of course this is subject to the constraint of not exceeding the autoignition temperature of the hydrogen fuel prior to admission into the combustor <NUM>. In an example, the injection temperature is from <NUM> to <NUM> kelvin, for example <NUM> kelvin.

The fuel heating system comprises a vaporiser <NUM> for heating of the hydrogen fuel to implement a phase change. This takes place between the pump <NUM> and the metering device <NUM>. In this way the metering device <NUM> meters gaseous hydrogen fuel.

In a non-claimed embodiment, the vaporiser <NUM> is configured to raise the temperature of the hydrogen fuel to the required injection temperature. Thus, in such a configuration, the metering device <NUM> meters the hydrogen fuel at the injection temperature.

According to the invention, the vaporiser <NUM> is configured to raise the temperature of the hydrogen fuel to a metering temperature less than the injection temperature. This could for example be from <NUM> to <NUM> kelvin, for example <NUM> kelvin. This reduces the risk of damage to electronic devices used for sensing temperature, pressure etc..

Further heating is implemented following the metering of hydrogen fuel by the metering device <NUM>. This is achieved with a heater <NUM>. The configuration of the vaporiser <NUM> and heater <NUM> may be substantially similar, and an example will be described further with reference to <FIG>.

Additionally or alternatively, the fuel heating system may comprise one or more heat exchangers for raising the temperature of the hydrogen fuel by use of rejected heat from the core gas turbine <NUM>. As will be described further with reference to <FIG>, this may be achieved by implementing a complex cycle configuration, for example fuel recuperation, intercooling, etc..

However, even in a simple cycle configuration as contemplated herein, this fuel heating may be achieved by, for example, cooling one or more of the various oil systems in the core gas turbine <NUM>. A specific example of such a configuration is shown in <FIG>, in which the fuel heating system comprises a fuel-oil heat exchanger <NUM> for cooling lubricating oil from the reduction gearbox <NUM>. In an example, even with a <NUM> percent efficient gearset, at maximum thrust it may still be required to reject around <NUM> kilowatts of heat from the gearbox oil system, which represents a significant opportunity for raising the temperature of the hydrogen fuel. It will be appreciated that other engine oil, such as main bearing lubrication oil, may also be cooled in a similar manner. It will also be appreciated that cooling air systems may be cooled in a similar manner, with high-pressure compressor <NUM> discharge air being cooled by heat exchange with the hydrogen fuel prior to being delivered to the high-pressure turbine <NUM> for cooling thereof.

In a simple cycle configuration it has been determined that due to the significant heat capacity of the hydrogen fuel, even if it is utilised as a heatsink for engine waste heat, it will still not reach the required injection temperature without implementation of the vaporiser <NUM> and optionally the heater <NUM>, depending on the chosen metering temperature. Further, even in a complex cycle configuration in which the heat of combustion products is recuperated into the hydrogen fuel, it has been determined that at certain points in the operational envelope there will be insufficient heat output from the engine to raise the fuel temperature to the injection temperature. Such occasions may include, for example, ground start, in-flight relight, end of cruise idle, etc..

An example configuration of the vaporiser <NUM> is shown in <FIG>. Such a configuration may also be used for the heater <NUM>.

The vaporiser <NUM> comprises an offtake <NUM> from a main fuel conduit <NUM>. The amount of hydrogen bled from the main fuel conduit <NUM> is controlled by a valve <NUM>. In operation, of the order of around <NUM> percent of the hydrogen fuel flow through the main fuel conduit <NUM> is bled for use in the vaporiser <NUM>.

As described previously, hydrogen has very high specific and latent heat capacities; however as a gas it has a very low molecular weight and density, and thus it can be challenging to exchange heat in a compact way. Thus the vaporiser <NUM> vaporises the hydrogen fuel in the main fuel conduit <NUM> by combustion of the bled fuel in a burner <NUM> located in heat exchange relationship with the main fuel conduit <NUM>. In the present embodiment, the burner <NUM> is concentric around the main fuel conduit <NUM>, although it will be appreciated that other arrangements are possible.

In the present embodiment, air for combustion with the bled hydrogen fuel is bled from the high-pressure compressor <NUM>. Alternatively, it may be bled from the low-pressure compressor <NUM>. It will be appreciated that the air for combustion could be obtained from any other suitable location.

In the present example, the air and the bled hydrogen fuel are mixed in a premixer <NUM>, although in alternative embodiments it may be directly co-injected into the burner with the hydrogen fuel instead. Combustion products from the burner <NUM> are, in an embodiment, exhausted into the bypass duct of the turbofan engine <NUM>. Alternatively, they may be exhausted outside the nacelle.

It should be understood that, in the present example, the products of combustion from the burner <NUM> are not mixed with the fuel in the main fuel conduit <NUM>. In this respect, the vaporiser <NUM> therefore differs from a pre-burner system as used in staged combustion cycle rocket engines.

In steady state, there is enough heat emanating from the burner <NUM> to ensure vaporisation of the small amount of bled hydrogen fuel. At engine start or other cold conditions for example, the vaporiser <NUM> comprises a preheater <NUM> to ensure vaporisation of the bled hydrogen fuel prior to mixing with air in the premixer <NUM>. In a specific embodiment, the preheater <NUM> comprises an electric heating element, for example a coil. Alternatively, the preheater <NUM> could be simply configured as a boil volume, in which the ambient conditions therein contain sufficient enthalpy to boil the initial flow of bled hydrogen fuel prior to delivery to the premixer <NUM> and the burner <NUM>.

Embodiments of the metering device <NUM> are illustrated in Figures 6A and 6B.

Fuel flow on a conventional liquid-fuelled aero engine is typically controlled by means of a pressure regulating valve and a profiled translating spill valve which returns a proportion of the flow supplied by the pump back to the pump inlet. However, because hydrogen has an extremely low density and viscosity, it has a strong tendency to leak through any gap. A control system that relies on close clearances to minimise leakages will be highly problematic with hydrogen as the fuel, since there will be significant leakage with even very tight clearances and the significant thermal variations in a hydrogen system will preclude very tight clearances.

The metering device <NUM> uses a fixed orifice which inherently has no moving parts and may therefore be sealed.

A first embodiment of the metering device <NUM> is shown in <FIG> and comprises a choked sonic orifice <NUM> located in the main fuel conduit <NUM>. Thus, in operation, the flow is through the orifice is choked, i.e. it has a Mach number of <NUM>. The flow is therefore a function only of the area of the orifice and upstream pressure and temperature, measured in this embodiment by a sensor <NUM>. In order to ensure the orifice remains choked, the orifice <NUM> comprises an exit with no expansion, i.e. it is sharp-edged, and the ratio of upstream to downstream pressures is set to be at least the critical pressure ratio which, for hydrogen (a diatomic gas) is around <NUM>.

Flow control is then achieved simply by adjusting the upstream pressure delivered by the pump <NUM>, the upstream temperature being measured and the orifice area being known.

As an alternative, the metering device <NUM> could comprise a fixed but unchoked orifice across which a pressure differential may be measured across upstream and downstream taps using an appropriate sensor. Mass flow may then be derived with knowledge of upstream and downstream pressures and temperatures and the geometry of the fixed orifice.

As described previously, it is envisaged that the fuel delivery system <NUM> and fuel injection system <NUM> may be used in an embodiment of the core gas turbine <NUM> implementing a simple cycle as described with reference to <FIG>, possibly with fuel cooling of engine or gearbox oil or cooling air. Alternatively, the core gas turbine engine <NUM> may implement a complex cycle.

A first embodiment of such a complex cycle is shown in <FIG> with like reference numerals used for matching features. In this example, the turbofan engine <NUM> and core gas turbine <NUM> are unchanged from their arrangement in <FIG>, save for the addition of a recuperator <NUM> located between the low-pressure turbine <NUM> and core nozzle <NUM>. The recuperator <NUM> forms part of the fuel heating system and is operable to heat hydrogen fuel by the exhaust stream of the core gas turbine <NUM>. In this way, less fuel may be required to heat the hydrogen fuel to the injection temperature, increasing cycle efficiency.

In an embodiment, the recuperator <NUM> is a spiral-wound recuperator, which reduces the likelihood of fracture due to thermal expansion and contraction.

Another embodiment of a complex cycle is shown in <FIG>, which builds on the cycle of <FIG> with the inclusion of a fuel turbine <NUM>. It will be appreciated that substantial energy recovery may be achieved from the exhaust stream if it is accepted that less thrust will be developed by the core nozzle <NUM>. Thus, it is possible to heat the hydrogen fuel beyond the required fuel injection temperature and to recover work in the fuel turbine <NUM>, which may be used to drive a load <NUM>. In this example the load <NUM> is an electrical generator. In a specific embodiment, the electrical generator powers the fuel pump <NUM>. Alternatively, the load could be the second shaft <NUM>, with an appropriate drive mechanism being provided. In this way, the fuel turbine <NUM> augments the low-pressure turbine <NUM>. It will be appreciated that other engine loads such as oil pumps etc. could also be driven by the fuel turbine <NUM>.

Additionally or alternatively, as shown in <FIG> it is possible perform further recuperation by using the hydrogen fuel to cool the combustor <NUM>. Gas turbine combustors feature a liner needs to be cooled to maintain its mechanical integrity.

In conventional liquid-fuelled aero engines the combustor liner is cooled by the airflow drawn from atmosphere and which has passed through the compression system. This is typically via a single pass system in which the air passes through holes in the liner and to enter the main heat release region. Hence this air cannot be part of the combustion process and therefore leads to an increase in emissions and a decrease in cycle efficiency.

Thus, in an embodiment, the hydrogen fuel is flowed around the liner of the combustor <NUM>. This scheme may be achieved by provision of for example helical cooling channels around the combustor <NUM> through which the hydrogen fuel may flow prior to injection.

Additionally or alternatively, as shown in <FIG> it is possible to provide intercooling and twin-pass recuperation.

In this embodiment, an intercooler <NUM> is provided in the interstage duct <NUM> between the low-pressure compressor <NUM> and the high-pressure compressor <NUM> for cooling low-pressure compressor discharge air by the hydrogen fuel. In this way, the amount of compression work required to be performed by the high-pressure compressor <NUM> is reduced.

In this specific embodiment, a second recuperator <NUM> is provided between the low-pressure turbine <NUM> and the recuperator <NUM> for further recuperative heating of the hydrogen fuel.

Thus, in this example, hydrogen fuel is first heated by the recuperator <NUM> to a temperature less than the low-pressure compressor <NUM> discharge air, which heats it further in the intercooler <NUM>. Further heating occurs in the second recuperator <NUM>, which has an inlet temperature higher than the recuperator <NUM>. In this way, the temperature difference between the hydrogen fuel and the core gas turbine exhaust temperature is maximised in each recuperator.

Additionally or alternatively, as shown in <FIG> a sequential combustion arrangement may be implemented to facilitate inter-turbine reheat. It will be appreciated that reheat of this type comprises additional stages of combustion to raise temperatures back to a maximum cycle temperature after a first stage of expansion. Along with intercooling, this moves the overall engine cycle closer to an Ericsson cycle, improving thermal efficiency substantially. In this specific example, the high-pressure turbine <NUM> is a multi-stage turbine and a reheat fuel injection system <NUM> and reheat combustor <NUM> are stationed between two of the stages 208A and 208B of the high-pressure turbine <NUM>. Alternatively, the reheat fuel injection system <NUM> and reheat combustor <NUM> may be stationed between the high-pressure turbine <NUM> and the low-pressure turbine <NUM>.

Due to its wide flammability limits and reaction rates, there is significant risk of flashback in hydrogen fuel injection systems. Thus it is preferable to utilise the direct injection principle with low mixing times and high velocities, as opposed to attempting any form of premixing. In order to minimise formation of oxides of nitrogen, residence time at high temperate must also be minimised. These constraints therefore favour a miniaturisation of the individual fuel injectors, sometimes referred to as "micromix" injectors.

<FIG> illustrate non-claimed examples of fuel injection systems.

<FIG> illustrate two possible arrangements of the fuel injection system <NUM>. It will be appreciated that in the present embodiment the core gas turbine <NUM> employs an annular combustion system, and it will be clear how the principles disclosed herein may be adapted e.g. for tubular systems.

In the embodiment of <FIG>, the fuel injection system <NUM> comprises a full annulus <NUM> of fuel injector blocks <NUM>. In the embodiment of <FIG>, the fuel injection system <NUM> comprises a plurality of sectors <NUM> each comprising a subset of the totality of fuel injector blocks <NUM>. In both embodiments, the fuel injector blocks <NUM> are configured with a geometry that substantially tessellates. It will be appreciated that the embodiment of <FIG> will produce a substantially more uniform circumferential heat-release profile, reducing the danger of hot streaks in the combustor <NUM> and uneven loading of the high-pressure turbine <NUM>, improving performance by reducing cooling requirements.

It is contemplated that the fuel injection system <NUM> would comprise many hundreds or even thousands of fuel injector blocks <NUM>. For example, in an embodiment there are from <NUM> to <NUM> fuel injector blocks, for example <NUM> fuel injector blocks <NUM>.

A first configuration for the fuel injector blocks <NUM> will be described with reference to Figures 13A to 15B. A second configuration for the fuel injector blocks <NUM> will be described with reference to <FIG>. A third configuration for the fuel injector blocks <NUM> will be described with reference to <FIG>.

The first configuration for the fuel injector blocks is shown in <FIG>, and will hereinafter be referred to as a rim injector block <NUM>. A first embodiment of the rim injector block <NUM> is shown in <FIG>, and has a quadrilateral, specifically a square, outer profile in the plane of tessellation. Another embodiment is shown in <FIG>, and has a hexagonal, specifically a regular hexagon, outer profile in the plane of tessellation. It will be appreciated that other outer profiles that tesselate could be used. In this example, the rim injector block <NUM> comprises an air admission duct <NUM> and a fuel admission aperture <NUM>.

Referring now to <FIG>, which is a cross sectional view on I-I of <FIG>, the air admission duct <NUM> has an inlet <NUM> for receiving air A from the diffuser <NUM> and an outlet <NUM> for delivering air into a mixing zone in the combustor <NUM>. The air admission duct <NUM> has a central axis C extending from the inlet <NUM> to the outlet <NUM>. The fuel admission aperture <NUM> is located around the periphery of the outlet <NUM> and is configured to inject hydrogen onto a jet shear layer formed at the outlet <NUM> for mixing in the mixing zone. In the embodiment of <FIG>, the fuel admission aperture <NUM> is configured to inject hydrogen fuel parallel to the central axis C, as shown by arrows F. <FIG> shows the equivalence ratios downstream of the rim injector block <NUM>, and was obtained by a periodic isothermal CFD simulation on this configuration. The air admission duct <NUM> was sized with a <NUM> millimetre diameter. In this example, a uniform equivalence ratio U was achieved within <NUM> millimetres of the injection point.

An alternative configuration of the rim injector block <NUM> is shown in Figure 15A, in which the fuel admission aperture <NUM> is configured to inject hydrogen fuel perpendicular to the central axis C, as shown by arrows F. Figure 15B shows the results of a periodic isothermal CFD simulation on this configuration. Again, the air admission duct <NUM> had a <NUM> millimetre diameter and a uniform equivalence ratio U was achieved within <NUM> millimetres of the injection point.

In both configurations, the injection of fuel onto the jet shear layer from the fuel admission aperture <NUM> minimises flammable mixtures at velocities lower than the turbulent flame speed close to the injector. This reduces the risk of flashback.

A second configuration for the fuel injector blocks is shown in <FIG>, and will hereinafter be referred to as a converging jet injector block <NUM>. The converging jet injector block <NUM> comprises a fuel admission duct <NUM> and a plurality of air admission ducts <NUM> distributed around the air admission duct. In this specific embodiment, the air admission ducts <NUM> are equidistant from the fuel admission duct <NUM>. As described previously, the outer profile of the injector block may be configured such that multiple blocks substantially tesselate adjacent to one another.

Referring now to <FIG>, which is a cross-sectional view on II-II of <FIG>, the fuel admission duct <NUM> has a central axis C and each air admission duct <NUM> has its own respective axis R defined between each duct's respective inlet and outlet. In the present embodiment, the fuel admission duct <NUM> and air admission ducts <NUM> are configured to respectively admit fuel and air without swirl. The respective axis R of each air admission duct <NUM>, is inclined towards the central axis C of the fuel admission duct <NUM>, such that emerging fuel F converges on the air A in the mixing zone.

<FIG> shows the equivalence ratios downstream of the converging jet injector block <NUM>, and was obtained by a periodic isothermal CFD simulation on this configuration. In this example, the fuel admission duct <NUM> was sized with a <NUM> millimetre diameter and the air admission ducts <NUM> were sized with a <NUM> millimetre diameter. A uniform equivalence ratio U was achieved within <NUM> millimetres of the injection point.

A third configuration for the fuel injector blocks is shown in <FIG>, and will hereinafter be referred to as a jet matrix injector block <NUM>. The converging jet injector block <NUM> comprises a fuel admission duct <NUM> and a plurality of air admission ducts <NUM> distributed around the air admission duct. In the present embodiment, the fuel admission duct <NUM> and air admission ducts <NUM> are configured to respectively admit fuel and air without swirl.

In <FIG>, the outer profile of the jet matrix injector block <NUM> is a quadrilateral and in <FIG> it is a hexagon.

In an embodiment, the jet matrix injector block <NUM> comprises a number 2N of air admission ducts <NUM> that meet the following criteria: 2N is even and equal to <NUM> or more, i.e. 2N ≥ <NUM>. Thus for example in <FIG> it can be seen that N = <NUM> and thus there are <NUM> air admission ducts <NUM>, whilst in <FIG> it can be seen that N = <NUM> and thus there are <NUM> air admission ducts <NUM>.

The air admission ducts <NUM> are then distributed around the fuel admission duct <NUM> such that they lie on the periphery of an N-gon (i.e. a polygon with N sides) centred on the central axis C of the fuel admission duct <NUM>. In a specific embodiment, N of the air admission ducts <NUM> are arranged at a respective vertex of the N-gon, and the other N of the air admission ducts <NUM> are arranged on a respective edge of the N-gon. For example, the other N of the air admission ducts <NUM> may be arranged at the midpoint of their respective edge.

Taking <FIG> as a worked example, N = <NUM>, therefore there are 2N = <NUM> air admission ducts <NUM>. These are located on the periphery of a <NUM>-gon, i.e. a square. Four of the air admission ducts <NUM> are located at the vertices (i.e. the corners) of the square, and the other four are located on a respective edge of the square. In this example they are located at a midpoint of the edges.

<FIG> is a cross-sectional view on III-III of <FIG>, in which it may be seen that in this embodiment the air admission ducts <NUM> are configured to lie parallel to the fuel admission duct <NUM>. Thus, each air admission duct <NUM> has a respective axis R defined between each duct's inlet and outlet, whilst the fuel admission duct <NUM> has a central axis C defined between its inlet and outlet. The respective axes R of the air admission ducts <NUM> are parallel with the central axis C.

<FIG> shows the equivalence ratios downstream of the jet matrix injector block <NUM>, and was obtained by a periodic isothermal CFD simulation on this configuration. In this example, the fuel admission duct <NUM> was sized with a <NUM> millimetre diameter and the air admission ducts <NUM> were sized with a <NUM> millimetre diameter. A uniform equivalence ratio U was achieved within <NUM> millimetres of the injection point.

By defining fuel injector blocks <NUM> of small scale relative to the overall size of the fuel injection system annulus, the flow field in the combustor <NUM> becomes self-similar and substantially invariant over different practical sizes. An example is shown in <FIG> of how the core gas turbine <NUM> may undergo a power scaling, i.e. the use of a substantially common design for two different power levels. In this example, the fuel injection system <NUM> is sized for an engine with power P in <FIG>, and a power 2P in <FIG>. However, it will be seen that the size of the fuel injector blocks <NUM> has not changed between the two designs, there has simply been an increase of the number making up the overall annulus.

Thus, in an industrial setting, the design process for a new specification engine may simply comprise obtaining a design of a standardised fuel injector block, such as blocks <NUM>, <NUM> and <NUM>. The standard specification for such an injector block would comprise its capability in terms of fuel mass flow performance and its dimensions and geometry. Engine performance data, typically derived prior to detailed component design, would set the required fuel mass flow requirements for the new engine type.

A simple evaluation of the quantity of standardised fuel injector blocks that meets the fuel mass flow requirements for the engine may then be performed. This would not require any dimensional scaling of the standardised fuel injectors, and indeed this would be discouraged as the flow field would change.

Claim 1:
A fuel delivery system (<NUM>) for delivering hydrogen fuel from a cryogenic storage system (<NUM>) to a fuel injection system (<NUM>) in a gas turbine engine (<NUM>), the fuel delivery system including a pump (<NUM>), a metering device (<NUM>), and a fuel heating system (<NUM>, <NUM>) for heating the hydrogen fuel to an injection temperature for the fuel injection system (<NUM>), the fuel heating system comprising a vaporiser (<NUM>) provided between the pump (<NUM>) and the metering device (<NUM>) configured to vaporise liquid hydrogen from the cryogenic storage system (<NUM>);
characterised in that
the metering device is one of:
a fixed orifice and flow rate is controlled by varying the pressure ratio across the orifice; and
a sonic fixed orifice (<NUM>) configured to operate in a choked condition, and flow rate is controlled by varying pressure upstream of the sonic fixed orifice (<NUM>); and wherein
the vaporiser (<NUM>) is configured to raise the temperature of the hydrogen fuel to a metering temperature less than the injection temperature, and the heating system further comprises a heater (<NUM>) for further heating of the hydrogen fuel to the injection temperature following metering by the metering device (<NUM>).