Patent Description:
Gas turbine engines typically use kerosene fuel which is pumped into a combustion section of an engine for combustion in the presence of compressed air provided into the combustion section from an upstream fan via various compressor stages. The fuel is typically pumped using a low- pressure centrifugal pump and a high-pressure gear pump.

Liquid hydrogen has recently become of interest as a fuel for gas turbine engines but liquid hydrogen presents problems for the existing pumping approaches because of the very low temperature (<NUM>), low lubricity of liquid hydrogen and the differing mechanical properties compared to kerosene.

Hydrogen-burning rockets use high speed turbo pumps driven by a driving turbine powered by the burning of a small proportion of the hydrogen fuel. These pumps are capable of the high rotational speeds (typically <NUM>,<NUM> RPM) necessary to pump liquid hydrogen. However, these high rotational speeds are likely to preclude the use of traditional driving means such as electric drives and gearboxes.

Cabin blower systems are used to pressurise the cabins of aircraft and to provide de-icing of aircraft wings. Cabin blowers are typically driven by one or more associated gas turbine engines of the aircraft. The gas turbine engine may be used to drive a cabin blower compressor in a number of ways (e.g. using electrical power generated by the engine or mechanically). Where mechanical driving of the compressor is employed, drive is typically taken from a shaft of the gas turbine engine via an accessory gearbox. A means of varying the speed of the drive delivered to the compressor is also required; it is not desirable for the cabin air flow and pressure to be determined by the particular operating condition of the gas turbine. Therefore a gearing mechanism such as a continuously variable transmission is also provided in the drive path between the accessory gearbox and compressor.

<CIT> discloses a cabin blower system comprising a compressor and a transmission system including a summing epicyclic gearbox. The summing epicyclic gearbox receives two inputs. The first input is provided mechanically from a first accessory gearbox powered by a shaft from the aircraft engine. The accessory gear box also drives a first electrical machine which converts the mechanical power to electrical power. A power management system interconnects the first electrical machine with a second electrical machine which, in turn is connected to the summing epicyclic gearbox. The second electrical machine converts the electrical power back to mechanical power to provide the second input for the epicyclic gearbox.

The epicyclic gearbox has an output that is a function of the difference between the speeds of the first and second inputs. The second input is a continuously variable positive or negative input which can be used to increase or decrease the mechanical first input as desired and as required by operating conditions.

<CIT> describes a similar system except that the second input is provided from a second shaft of the aircraft engine which powers a second accessory gearbox which, in turn provides power to the first electrical machine.

As well as operating a blower mode, these known compressors can also be operated in a starter mode where they can be used to start the aircraft engine. In this starter mode, the compressor can be operated in reverse (as an expander) and the second electrical machine can be isolated so that the compressed air drives the epicyclic gearbox which in turn drives the accessory gear box and associated shaft. The shaft can, in turn, drive a compressor within the aircraft engine which facilitates starting of the engine.

United States patent application <CIT> discloses a cryogenic fuel system for an aircraft having a turbine engine with a compressor section and a combustion chamber, including a tank for storing cryogenic fuel, a supply line operably coupling the tank to the combustion chamber and a pump coupling the tank to the supply line to pump the cryogenic fuel at high pressure through the supply line where the pump is operably coupled to the compressor such that operation of the turbine engine drives the pump and a method for delivering fuel in a fuel system to a turbine engine.

United States patent application <CIT> discloses a gas turbine engine which includes a main compressor section, a combustor, and a main turbine section. A fuel pump delivers fuel to the combustor. A tap taps air compressed by the main compressor section, and is connected for delivering the tapped air through a first heat exchanger and to a boost compressor. Air downstream of the boost compressor is connected to cool a component. Driving compressed air is connected to be delivered to a power turbine. The power turbine is connected to drive both the boost compressor and the fuel pump.

United States patent application <CIT> discloses a fuel power transfer system for an engine which includes a cryogenic fuel supply, a fuel pump in fluid communication with the cryogenic fuel supply, a multi-position valve in fluid communication with the fuel pump and a combustion chamber of the engine, a fuel turbine operatively coupled to the fuel pump and having a primary discharge port in fluid communication with the combustion chamber, a primary heat exchanger in fluid communication between the multi-position valve and the fuel turbine, and a gearbox operatively coupled to the fuel turbine and the fuel pump and configured to transfer power from the fuel turbine to the engine. There is a need to provide a fuel system for a gas turbine engine that can accommodate the challenges posed by the use of liquid hydrogen fuel.

The present disclosure has been devised with the above considerations in mind.

According to a first aspect there is provided a fuel system for a gas turbine,as set out in claim <NUM>.

By using compressed air flow to drive a driving turbine which, in turn drives the fuel pump, the present inventors have found that the fuel pump can be operated at rotational speeds high enough to pump liquid hydrogen within a gas turbine engine without the need for traditional drives/gears and without the need for burning fuel to drive the driving turbine.

Optional features will now be described.

In some embodiments, the fuel pump may comprise a turbo pump.

In some embodiments, the system may further comprise a fuel reservoir configured to contain liquid hydrogen. For example, the fuel reservoir may be a cryogenic fuel reservoir e.g. cooled to around - <NUM>. In some embodiments, the system further comprises cryogenic pipework connecting the fuel reservoir to the fuel pump.

In some embodiments the fuel reservoir is located proximate e.g. adjacent the fuel pump in order to minimise pipework, e.g. cryogenic pipework, between the fuel reservoir and fuel pump.

The source of compressed air may be provided by a compressor.

In some embodiments the fuel system comprises a compressor i.e. there is a dedicated fuel system compressor for generating the compressed air to drive the driving turbine. The fuel system compressor may have a blower mode for the feed of compressed air to the driving turbine.

The fuel system transmission system may comprise a first accessory gearbox operatively coupled between one of the intermediate- or high-pressure compressor of the gas turbine engine and the epicyclic gearbox. The fuel system transmission system may comprise a first electrical machine connected to the first accessory gear box. In other embodiments, the first electrical machine may be connected to a second accessory gearbox which is operatively connected between the other one of the intermediate- or high-pressure compressor of the gas turbine engine and the epicyclic gearbox. The first electrical machine is configured to convert mechanical power to electrical power and to provide the electrical power to a second electrical machine. The second electrical machine is configured to convert the electrical power (from the first electrical machine) to mechanical power for input into the epicyclic gear box. The fuel system may comprise a fuel system power management system to control the transfer of electrical power between the first and second electrical machines. The fuel system power management system may further comprise a power source e.g. a battery. The power source/battery may be used to drive the driving turbine in a start-up mode to achieve sufficient fuel flow for start-up of the gas turbine engine.

The fuel system compressor may additionally have a starter mode where the fuel system compressor acts in reverse as an expander. The compressed air generated in the starter mode may be provided to drive the driving turbine to provide fuel to the fuel pump during start-up of the gas turbine engine. In the starter mode, the fuel system transmission system may transmit mechanical power to a shaft operatively connected to the intermediate- or high-pressure compressor of the gas turbine engine.

The source of compressed air is provided from a cabin blower system of an aircraft i.e. the fuel system comprises a cabin blower bleed line from a cabin blower compressor of the cabin blower system.

In this way, a portion of the compressed air generated by the cabin blower compressor may be diverted to drive the driving turbine. Cabin blower systems are typically sized to accommodate cabin air needs under operating conditions experienced at the top of descent. Under these operating conditions, fuel demand is at a minimum. At operating conditions where fuel demand is higher, the cabin blower system typically has spare capacity which can be effectively utilised to drive the driving turbine through the cabin blower bleed line.

The cabin blower system comprises a cabin blower compressor and a cabin blower transmission system. The cabin blower compressor may be provided with air from the core compressors or from the engine fan bypass duct.

The cabin blower transmission system is operatively coupled to a compressor stage of the core of the gas turbine engine, e.g. operatively coupled to the intermediate- and/or high-pressure compressor of the gas turbine engine. The cabin blower transmission system comprises a summing epicyclic gearbox operatively coupled to the intermediate- and/or high-pressure compressor of the gas turbine engine. The cabin blower transmission system may comprise a first accessory gearbox operatively coupled between one of the intermediate- or high-pressure compressor of the gas turbine engine and the epicyclic gearbox. The cabin blower transmission system may comprise first electrical machine connected to the first accessory gear box. In other embodiments, the first electrical machines may be connected to a second accessory gearbox which is operatively connected between the other one of the intermediate- or high-pressure compressor of the gas turbine engine and the epicyclic gearbox. The first electrical machine is configured to convert mechanical power to electrical power and to provide the electrical power to a second electrical machine. The second electrical machine is configured to convert the electrical power (from the first electrical machine) to mechanical power for input into the epicyclic gear box. The cabin blower system may be as described in <CIT> and/or <CIT>. For example, the cabin blower compressor may have a blower mode for the feed of compressed air to the cabin blower bleed. It may also have a starter mode where the transmission system transmits mechanical power to a shaft operatively connected to the intermediate- or high-pressure compressor.

In yet further embodiments, the source of compressed air may comprise a core bleed line from the gas turbine engine core. For example, the system may comprise a core bleed line from a compressor stage of the gas turbine engine e.g. a core bleed line from an intermediate pressure compressor and/or a core bleed line from the high-pressure compressor within the gas turbine engine core.

In yet further examples not forming part of the claimed invention, the source of compressed air comprises a mixture of two or more of a dedicated fuel system compressor, a core bleed line and/or a cabin blower bleed line, all of which are described above.

In some embodiments, the fuel system further comprises a fuel evaporator downstream of the fuel pump to evaporate the fuel e.g. liquid hydrogen into a gaseous phase fuel for introduction into the gas turbine engine.

In some embodiments, the fuel system further comprises an exhaust downstream of the driving turbine and upstream of the fuel pump for venting any upstream fuel leakage.

In some embodiments, the fuel system comprises a fuel metering unit for metering flow of fuel to the fuel pump. In some embodiments, the fuel system additionally or alternatively comprises a throttle which may be upstream of the driving turbine to control the flow of compressed air and thus the driving force of the driving turbine which, in turn controls the fuel output of the fuel pump. Alternatively, the throttle may be downstream of the driving turbine to control the output of the driving turbine to the fuel pump.

In a second aspect, there is provided an aircraft comprising a fuel system according to the first aspect.

In some embodiments, the aircraft further comprises a cabin blower system, the cabin blower system comprising a cabin blower compressor, the fuel system comprising a cabin blower feed line for channelling compressed air from the cabin blower compressor to the driving turbine. Further details of the cabin bower system may be as described for the first aspect.

In some embodiments, the aircraft comprises an engine core comprising a compressor stage, the fuel system comprising a core bleed line for channelling compressed air from the compressor stage to the driving turbine. The compressor stage may comprise an intermediate- or a high-pressure compressor.

According to a third aspect, there is provided a method of operating a fuel system in a gas turbine engine, as set out in claim <NUM>.

In some embodiments, the fuel is liquid hydrogen.

The method comprises using a compressor to generate the compressed air for powering the driving turbine.

The method comprises generating the compressed air to drive the driving turbine using a cabin blower system of an aircraft i.e. the method comprises channelling compressed air from a cabin blower compressor through a cabin blower bleed line.

The method comprises driving the blower system compressor through a cabin blower system transmission system operatively coupled to a compressor stage within the engine core e.g. with the intermediate- and/or high-pressure compressor of the gas turbine engine. The cabin blower transmission system is as described above for the first aspect.

In examples not forming part of the claimed invention, the method may comprise channelling compressed air to the driving turbine through a core bleed line from the gas turbine engine core. For example, the method may comprise channelling compressed air from a compressor stage of the gas turbine engine e.g. through a core bleed line from the intermediate pressure compressor and/or a core bleed line from the high-pressure compressor within the gas turbine engine core.

In some embodiments, the method further comprises evaporating the fuel downstream of the fuel pump into a gaseous phase fuel for introduction into the gas turbine engine.

In some embodiments, the method comprises an exhausting any fuel leakage downstream of the driving turbine and upstream of the fuel pump.

In some embodiments, the method comprises metering flow of fuel to the fuel pump. In some embodiments, the method additionally or alternatively comprises a throttling the compressed air upstream of the driving turbine to control the driving force of the driving turbine which, in turn controls the fuel output of the fuel pump.

With reference to <FIG> , a gas turbine engine is generally indicated at <NUM>, having a principal and rotational axis <NUM>. The engine <NUM> comprises, in axial flow series, an air intake <NUM>, a propulsive fan <NUM>, an intermediate-pressure compressor <NUM>, a high-pressure compressor <NUM>, combustion equipment <NUM>, a high-pressure turbine <NUM>, an intermediate-pressure turbine <NUM>, a low-pressure turbine <NUM> and an exhaust nozzle <NUM>. A nacelle <NUM> generally surrounds the engine <NUM> and defines both the intake <NUM> and the exhaust nozzle <NUM>.

The gas turbine engine <NUM> works in the conventional manner so that air entering the intake <NUM> is accelerated by the fan <NUM> to produce two air flows: a first air flow into the intermediate-pressure compressor <NUM> and a second air flow which passes through a bypass duct <NUM> to provide propulsive thrust. The intermediate-pressure compressor <NUM> compresses the air flow directed into it before delivering that air to the high-pressure compressor <NUM> where further compression takes place.

The compressed air exhausted from the high-pressure compressor <NUM> is directed into the combustion equipment <NUM> where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high-, intermediate- and low-pressure turbines <NUM>, <NUM>, <NUM> before being exhausted through the nozzle <NUM> to provide additional propulsive thrust. The high-<NUM>, intermediate- <NUM> and low- <NUM> pressure turbines drive respectively the high-pressure compressor <NUM>, intermediate-pressure compressor <NUM> and fan <NUM>, each by suitable interconnecting shaft.

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. By way of example such engines may have an alternative number of interconnecting shafts (e.g. two) and/or an alternative number of compressors and/or turbines. Further the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan.

<FIG> shows a fuel system <NUM> for a gas turbine engine <NUM>. The system comprises a fuel pump <NUM> for fluid communication with a cryogenic fuel reservoir (not shown) via a cryogenic pipework <NUM>. A driving turbine <NUM> for driving the fuel pump <NUM> is provided upstream of the fuel line <NUM>. The fuel pump <NUM> is a turbo-pump.

The driving turbine <NUM> is powered by compressed air flow within an airflow channel <NUM> generated by a fuel system compressor <NUM> and a fuel system transmission system <NUM>. The fuel system transmission system <NUM> is operatively coupled to a compressor stage within the engine core <NUM>. The fuel system transmission system <NUM> comprises a summing epicyclic gearbox <NUM> operatively coupled to the intermediate- and/or high-pressure compressor <NUM>, <NUM> of the gas turbine engine core <NUM>. The fuel system transmission system <NUM> comprises a first accessory gearbox <NUM> operatively coupled between one of the intermediate- or high-pressure compressor <NUM>, <NUM> of the gas turbine engine core <NUM> and the summing epicyclic gearbox <NUM> to provide a first mechanical input into the summing epicyclic gearbox <NUM>.

The fuel system transmission system <NUM> comprises a first electrical machine <NUM> connected to the first accessory gear box <NUM>. The first electrical machine <NUM> is configured to convert mechanical power from the first accessory gear box <NUM> into electrical power. The first electrical machine <NUM> is configured to provide this electrical power to a second electrical machine <NUM> which is configured to convert the electrical power (from the first electrical machine <NUM>) to mechanical power to provide a second mechanical input into the epicyclic gear box <NUM>.

The fuel system also comprises a fuel system power management system <NUM> to control the transfer of electrical power between the first and second electrical engines <NUM>, <NUM>.

The summing epicyclic gearbox <NUM> has an output that is a function of the difference between the speeds of the first and second inputs. The second input is a continuously variable positive or negative input (as a result of the control by the power management system <NUM>) which can be used to increase or decrease the compressed air output of the fuel system compressor <NUM> as desired and as required by operating conditions.

In addition, a first core bleed line (not shown) channels compressed air from the intermediate pressure compressor <NUM> in the gas turbine engine into the compressed air flow channel <NUM>. A second core bleed line (not shown) channels compressed air from the high-pressure compressor <NUM> in the gas turbine engine into the compressed air flow channel <NUM>.

In other embodiments (not shown), the source of compressed air may additionally or alternatively be provided from a cabin blower system of an aircraft i.e. the fuel system may comprise a cabin blower bleed line from a cabin blower compressor of the cabin blower system.

The cabin blower compressor may be provided with air from the core compressors (as shown in <FIG>) or from the engine fan bypass duct.

By using compressed air flow from the fuel system compressor <NUM> and from the first and second core bleed lines to drive a driving turbine <NUM> which, in turn drives the fuel pump <NUM>, the fuel pump <NUM> can be operated at rotational speeds high enough to pump liquid hydrogen within the gas turbine engine without the need for traditional drives/gears and without the need for burning fuel to drive the driving turbine. Any upstream leakage of hydrogen to the driving turbine <NUM> can be exhausted from the driving turbine <NUM> along with the compressed air via an exhaust <NUM>.

A throttle (not shown) is provided in the compressed air flow channel <NUM> to vary the volume/flow rate of the compressed air flow and thus control the driving power of the driving turbine <NUM> (and thus the output of the fuel pump <NUM>). Such a throttle can negate the need for a fuel metering system.

The summing epicyclic gearbox <NUM> can also be used to reverse the fuel system compressor <NUM> from a blower mode to a start-up mode in which air from the reversed compressor <NUM> can still be used to drive the driving turbine <NUM> and allow operation of the fuel pump during start-up. In addition, isolation of the second electrical machine <NUM> allows the reversed compressor <NUM> to drive first accessory gearbox and, in turn, the operatively coupled engine core compressor <NUM>, <NUM> to assist in engine start up.

While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. Various changes to the described embodiments may be made, provided they fall within the scope of the appended claims defining the invention.

Claim 1:
A fuel system (<NUM>) for a gas turbine engine (<NUM>), the system comprising:
a fuel pump (<NUM>) for fluid communication with a fuel reservoir;
a driving turbine (<NUM>) for driving the fuel pump (<NUM>); and
a source of compressed air flow (<NUM>) to drive the driving turbine (<NUM>); characterised in that:
the source of compressed air flow comprises a cabin blower compressor and a cabin blower bleed line from a cabin blower system; and
a cabin blower transmission operatively coupled to a compressor stage of the gas turbine engine and comprising a summing epicyclic gearbox operatively coupled to a compressor of the gas turbine engine.