Patent Description:
In aerospace applications, hydrogen is often considered as a fuel whenever low carbon emissions are desired. Hydrogen is often stored cryogenically in order to maximize its energy density per unit volume, however it must first be evaporated to a gaseous state before use in combustion. There is always a need for improvements to systems and methods for converting the liquid hydrogen into a gaseous state for use in combustion in the aerospace industry.

<CIT> and <CIT> disclose arrangements of the prior art.

In an aspect of the invention, there is provided a hydrogen fuel system for aircraft according to claim <NUM>.

In optional embodiments, the fuel feed conduit is further defined by, in fluid series downstream of the evaporator, an accumulator, a pressure regulator, a gaseous hydrogen metering unit, a manifold shut-off valve, and a fuel manifold.

In certain optional embodiments, an upstream working fluid conduit in fluid communication with the evaporator and defined at least in part by an electric heat source operative to add heat into the working fluid for heat exchange with the flow of hydrogen passing through the evaporator, and an upstream working fluid pump to drive the working fluid through the evaporator. In certain such optional embodiments, the electric energy source is external to the gas turbine engine, and the electric energy source includes at least one of: an auxiliary power unit, an electric generator, and/or a battery operatively connected to power the electric heat source.

In certain optional embodiments, the engine heat source is turbine exhaust from a turbine exhaust section of the gas turbine engine. In certain optional embodiments, a downstream working fluid conduit defined at least in part by the turbine exhaust section of the gas turbine engine, the gaseous hydrogen heater, and a downstream working fluid pump to drive heated downstream working fluid to the gaseous hydrogen heater to add heat to the flow of gaseous hydrogen passing through the gaseous hydrogen heater. In certain such optional embodiments, the downstream working fluid conduit is at least partially coiled around the turbine exhaust section of the gas turbine engine to add heat from the turbine exhaust section of the gas turbine engine to the flow of downstream working fluid in the downstream working fluid conduit.

In certain optional embodiments, the engine heat source is engine fluid. In certain such optional embodiments, an engine fluid conduit defined at least in part by, in fluid series, an engine fluid source, the gaseous hydrogen heater, and an engine fluid return to add heat from the engine fluid to the flow of gaseous hydrogen passing through the gaseous hydrogen heater.

In certain optional embodiments, the gaseous hydrogen heater is associated with an electric heat source to add heat to the flow of gaseous hydrogen passing through the gaseous hydrogen heater, wherein the electric heat source is external to, but powered by an electric energy module driven the gas turbine engine.

In certain optional embodiments, a gear box operatively connected to be driven by a spool shaft of the gas turbine engine, wherein the electric energy source includes an electric generator operatively connected to be driven by the gear box to power the electric heat source associated with the gaseous hydrogen heater. In certain such optional embodiments, a downstream working fluid conduit defined at least in part by, the gaseous hydrogen heater, the electric heat source to heat the downstream working fluid, a downstream working fluid pump to drive the downstream working fluid to the gaseous hydrogen heater to add heat to the flow of gaseous hydrogen passing through the gaseous hydrogen heater.

These and other features of the optional embodiments of the subject disclosure will become more readily apparent to those skilled in the art from the following detailed description taken in conjunction with the drawings.

Reference will now be made to the drawings wherein like reference numerals identify similar structural features or aspects of the subject disclosure. For purposes of explanation and illustration, and not limitation, an illustrative view of an embodiment of a system in accordance with the disclosure is shown in <FIG> and is designated generally by reference character <NUM>. Other embodiments and/or aspects of this disclosure are shown in <FIG>. Certain embodiments described herein can be used to improve conversion of liquid hydrogen fuel to gaseous hydrogen for combustion.

The present disclosure relates generally to fuel control for gas turbine engines, and more particularly to control of gaseous fuel flow. A gas turbine engine may be fueled with gaseous fuel such as hydrogen gas. It is possible to gasify liquid hydrogen from an aircraft supply through an appropriate fuel pump, heat exchangers, pressure regulator, and metering valves. However, conventional gasification methods may not provide sufficient heat to the liquid hydrogen to offer full or efficient combustion.

In certain embodiments, referring to <FIG>, an aircraft <NUM> can include an engine <NUM>, where the engine <NUM> can be a propulsive energy engine (e.g. creating thrust for the aircraft <NUM>), or a non-propulsive energy engine, and a fuel system <NUM>. As described herein, the engine <NUM> is a turbofan engine, although the present disclosure may likewise be used with other engine types. The engine <NUM> includes a compressor section <NUM> having a compressor <NUM> in a primary gas path <NUM> to supply compressed air to a combustor <NUM> of the aircraft engine <NUM>. The primary gas path <NUM> includes a nozzle manifold <NUM> for issuing fluid to the combustor <NUM>.

The primary gas path <NUM> includes, in fluid communication in a series: the compressor <NUM>, the combustor <NUM> fluidly connected to an outlet <NUM> of the compressor <NUM>, and a turbine section <NUM> fluidly connected to an outlet <NUM> of the combustor <NUM>. The turbine section <NUM> is mechanically connected to the compressor <NUM> to drive the compressor <NUM>.

A fuel feed conduit <NUM> is defined at least in part by a gaseous fuel supply <NUM>, a plurality of fuel nozzles, and the combustor <NUM> of the gas turbine engine <NUM>. In embodiments, the gaseous fuel supply <NUM> can be any suitable gaseous fuel, such as a gaseous pressure and/or temperature regulated fuel supply, which may be or include hydrogen gas.

In embodiments, the fuel feed conduit can further be defined, in fluid series or any suitable order or combination between the combustor <NUM> and the gaseous fuel supply <NUM>, by a fuel shut off valve <NUM>, a fuel pump <NUM>, a liquid/gaseous fuel evaporator/heater <NUM>, a turbine air cooling heat exchanger <NUM>, a gaseous fuel accumulator <NUM>, a pressure regulator <NUM>, a gaseous fuel metering unit <NUM>, and/or a fuel manifold shut off valve <NUM>, and by fuel lines interconnecting therebetween. In certain embodiments, the pre-pressurized gaseous fuel accumulator <NUM> can be used as backup supply pressure source.

As shown in <FIG>, the fuel system <NUM> includes the a evaporator/heater <NUM> where heat is supplued from any suitable heat source <NUM>, for example engine fluid to evaporate/heat the liquid hydrogen therein. The evaporator/heater <NUM> can be an evaporator and/or a heater. For example, if an evaporator, the evaporator <NUM> converts liquid hydrogen to gaseous hydrogen therein. However, in certain embodiments, the evaporator/heater <NUM> is only a heater, in which the heater <NUM> receives heated liquid from an external source to exchange heat with the liquid hydrogen therein, but does not necessarily evaporate the liquid hydrogen or may only evaporate some. Similarly, the evaporator/heater <NUM> can be exclusively a heater if it receives gas solely to add heat to the gas, without any conversion of state. In some instances, it may be that the engine fluid alone is insufficient to entirely evaporate the flow of liquid hydrogen in the evaporator <NUM>, or it may be that the engine fluid alone is unable to heat the liquid and/or gaseous hydrogen to a heat acceptable for desired engine performance. Therefore, a multi-stage heating system (e.g. systems <NUM>-<NUM>) is provided herein.

As shown in <FIG>, in accordance with at least one aspect of this disclosure, there is provided a multi-stage heated fuel system <NUM> for hydrogen fuel for aircraft <NUM>. The pump <NUM> is in fluid communication with the liquid hydrogen tank <NUM> to drive fuel from the liquid hydrogen tank <NUM> to the gas turbine engine <NUM> via the fuel feed conduit <NUM>.

An upstream working fluid conduit <NUM> is defined at least in part by an upstream heater circuit, which branches off of the fuel feed conduit <NUM>. An electric heat source <NUM> (e.g. a resistive heater) is disposed in the upstream heater circuit. The electric heat source <NUM> is in thermal communication with the upstream working fluid flowing through the upstream working fluid conduit <NUM> to add heat to the upstream working fluid.

An upstream working fluid pump <NUM> can be disposed in the upstream heater circuit to drive working fluid through the upstream heater circuit, passing through the electric heat source <NUM> and the evaporator <NUM>. The upstream working fluid conduit <NUM> is in fluidly connected to the evaporator <NUM> to add heat into a flow of hydrogen passing through the evaporator <NUM>. The upstream working fluid conduit <NUM> is fluidly isolated from the flow of liquid hydrogen within the evaporator <NUM> but is in thermal communication with the flow of hydrogen passing through the evaporator <NUM> for exchanging heat from the upstream working fluid to the flow liquid hydrogen passing through the evaporator <NUM>.

An electric energy source <NUM> is associated with the electric heat source <NUM> to power the electric heat source <NUM>. The electric energy source can be external to the gas turbine engine <NUM>, and can be or include at least one of: an auxiliary power unit <NUM>, an electric generator <NUM>, and/or a battery <NUM> operatively connected to power the electric source <NUM>.

Still with reference to <FIG>, according to the claimed invention, the fuel feed conduit <NUM> is further defined by a gaseous hydrogen pump <NUM> downstream of the evaporator <NUM>, fluidly connecting an outlet of the evaporator <NUM> to the gas turbine engine <NUM> to drive gaseous hydrogen from the evaporator <NUM> to the gas turbine engine <NUM>. The fuel feed conduit can be further defined by a gaseous hydrogen heater <NUM> downstream of the gaseous hydrogen pump <NUM> and upstream of the gas turbine engine <NUM>. The gaseous hydrogen heater <NUM> is associated with a downstream heat source <NUM> internal to the engine <NUM> to add heat to the flow of gaseous hydrogen passing through the gaseous hydrogen heater <NUM>.

Referring now to <FIG>, in certain embodiments a system <NUM> can have similar components as in system <NUM>. For brevity, the description of common elements that have been described above are not repeated with respect to <FIG>. In system <NUM>, the downstream heat source <NUM> is compressor bleed air from the compressor section <NUM>. A bleed air conduit <NUM> fluidly connects the compressor section <NUM> with the gaseous hydrogen heater <NUM> to add heat from the compressor section <NUM> to the flow of gaseous hydrogen passing through the gaseous hydrogen heater <NUM>.

According to the claimed invention, an environmental control system return conduit <NUM> conveys cooled compressor bleed air to an environmental control system (ECS) <NUM>. In certain embodiments, a switching module can be disposed in the bleed air conduit <NUM> to allow the option of using either the low pressure compressor bleed air (e.g. lower temperature) as the downstream heat source <NUM>) or high pressure compressor bleed air (e.g. higher temperature) as the downstream heat source <NUM>, depending on the amount of heat needed for the gaseous hydrogen heater <NUM> and the amount of cooling air required for the ECS <NUM>.

Referring to <FIG>, in certain embodiments a system <NUM> can have similar components as in system <NUM>. For brevity, the description of common elements that have been described above are not repeated with respect to <FIG>. In the system <NUM>, the downstream heat source <NUM> can be turbine exhaust from a turbine exhaust section (e.g. turbine outlet <NUM>) of the gas turbine engine <NUM>. A downstream working fluid conduit <NUM> defines a downstream heater circuit which branches off the fuel feed conduit <NUM>. In certain embodiments, the downstream working fluid conduit <NUM> is at least partially coiled (e.g. portion 468a) around the turbine exhaust section <NUM>.

A downstream working fluid pump <NUM> is disposed in the downstream working fluid conduit <NUM> to drive downstream working fluid through the downstream heater circuit, passing through the coiled portion 468a and to the gaseous hydrogen heater <NUM> to add heat to the flow of gaseous hydrogen passing through the gaseous hydrogen heater <NUM>. The downstream working fluid conduit <NUM> is fluidly isolated from the flow of liquid hydrogen within the gaseous hydrogen heater <NUM> but is in thermal communication with the flow of hydrogen passing through the gaseous hydrogen heater <NUM> for exchanging heat from the downstream working fluid to the flow gaseous hydrogen passing through the gaseous hydrogen heater <NUM>.

Referring to <FIG>, in certain embodiments a system <NUM> can have similar components as in system <NUM>. For brevity, the description of common elements that have been described above are not repeated with respect to <FIG>. In system <NUM>, the downstream heat source <NUM> is an engine fluid (e.g. engine oil). An engine fluid conduit is defined at least in part by an engine fluid feed conduit <NUM> and an engine fluid return conduit <NUM> fluidly connecting between the gas turbine engine <NUM> and the gaseous hydrogen heater <NUM> to convey the hot engine fluid from the engine <NUM> to the gaseous hydrogen heater <NUM> to add heat to the flow of gaseous hydrogen passing through the gaseous hydrogen heater <NUM>. An engine pump <NUM> can be disposed in the compressor section <NUM> to drive the engine fluid from the engine <NUM> to the gaseous hydrogen heater <NUM>.

Referring now to <FIG>, in certain embodiments a system <NUM> can have similar components as in system <NUM>. For brevity, the description of common elements that have been described above are not repeated with respect to <FIG>. In system <NUM>, the downstream heat source <NUM> is an electric heat source <NUM> (e.g. resistive heater), where the electric heat source <NUM> is external to, but powered by an electric energy module <NUM> driven by the gas turbine engine <NUM>. For example, a gear box <NUM> is operatively connected to be driven by a spool shaft <NUM> of the gas turbine engine <NUM>. The electric energy module <NUM> includes an electric generator <NUM> operatively connected to be driven by the gear box <NUM> to power the electric heat source <NUM> associated with the gaseous hydrogen heater <NUM>. A downstream working fluid conduit <NUM> defines a downstream heater circuit which branches off the fuel feed conduit <NUM>.

A downstream working fluid pump <NUM> is disposed in the downstream working fluid conduit <NUM> to drive downstream working fluid through the downstream heater circuit, passing through the electric heat source <NUM> to the gaseous hydrogen heater <NUM> to add heat to the flow of gaseous hydrogen passing through the gaseous hydrogen heater <NUM>. The downstream working fluid conduit <NUM> is fluidly isolated from the flow of liquid hydrogen within the gaseous hydrogen heater <NUM> but is in thermal communication with the flow of hydrogen passing through the gaseous hydrogen heater <NUM> for exchanging heat from the downstream working fluid to the flow gaseous hydrogen passing through the gaseous hydrogen heater <NUM>.

In accordance with yet another aspect of this disclosure, which is not part of the claimed invention, there is provided a method of heating fuel in an aircraft (e.g. aircraft <NUM>). The method includes heating a flow of liquid hydrogen with an evaporator (e.g. evaporator <NUM>) in thermal communication with an upstream heat source (e.g. heat source <NUM>) to convert the flow of liquid hydrogen to a flow of gaseous hydrogen. The method includes heating the flow of gaseous hydrogen with a gaseous hydrogen heater (e.g. <NUM>, <NUM>, <NUM>, <NUM>, <NUM>) in thermal communication with a downstream heat source (e.g. <NUM>, <NUM>, <NUM>, <NUM>, <NUM>).

In certain embodiments, the upstream heat source is external to a gas turbine engine (e.g. engine <NUM>) and the downstream heat source is internal to the gas turbine engine. In certain such embodiments, heating the flow of liquid hydrogen includes exchanging heat between an upstream heating working fluid in thermal communication with the upstream heat source and the flow of liquid hydrogen in the evaporator. In certain embodiments, the method includes powering the upstream heat source with at least one of an auxiliary power unit, an electric generator, and/or an electric energy module.

In embodiments, heating the flow of gaseous hydrogen includes exchanging heat between an engine fluid and the flow of gaseous hydrogen in the gaseous hydrogen heater. In embodiments, the engine fluid includes at least compressor bleed air and, optionally, turbine exhaust, and/or engine oil.

In certain embodiments, heating the flow of gaseous hydrogen includes exchanging heat between a downstream working fluid in thermal communication with the downstream heat source and the flow of gaseous hydrogen in the gaseous hydrogen heater. In certain such embodiments, powering the downstream heat source with an electric generator (e.g. generator <NUM>) driven by the gas turbine engine.

In embodiments, liquid hydrogen (LH2) can be pumped from the aircraft cryogenic tanks using an electrically driven LH2 pump. Depending on the heat required to evaporate the LH2, one engine heat source or multiple heat sources can then heat the cold (e.g. <NUM>) liquid hydrogen to convert it from liquid to a gaseous state (e.g. <NUM>) where it can then be combusted in the engine to produce power. In certain configurations, engine bleed or waste heat (e.g. exhaust and oil) can used for evaporation of LH2 removing the need for additional heat energy.

In embodiments, if the engine heat sources (e.g. total thermal energy available) are insufficient to convert the liquid hydrogen to gaseous state and raise its temperature for distribution and combustion in the engine, the heating may occur in multiple stages of the aircraft power plant systems. In certain configurations, the LH2 can first be evaporated to gaseous state using an electrically powered heat exchanger with an intermediate fluid medium to avoid ignition. The electric power supply can come from either an auxiliary power unit (e.g. a thermal engine) and/or electric power storage unit (battery pack). Subsequently, the cold gaseous hydrogen (GH2) can be further heated using heat sources in the engine (e.g. as described herein). In embodiments, the engine heat sources may be used independently, in parallel or in series depending on the final GH2 heat requirements. The multi-stage evaporator and heating approach can be advantageous, for example, by separating the evaporation phase from the engine, the supply of GH2 for the engine is not coupled to the heat output of the engine, meaning for certain low heat output phases (e.g. at engine starting phase, engine ground idle or flight idle running conditions), there is adequate GH2 to start the thermal cycle engine per example.

In embodiments, evaporating the LH2 or heating the GH2 can be accomplished by using hot compressor bleed air, normally used for the aircraft environmental control system (ECS). An additional potential benefit of this configuration can be the ECS bleed air would be cooled by the LH2, thus eliminating the need for a typical ECS pre-cooler. A switching valve can be included in this embodiment to allow for switching between the low pressure compressor bleed air (e.g. lower temperature and lower impact on performance) or high pressure compressor bleed air (e.g. higher temperature more impact on performance), depending on the amount of evaporation needed or cooling needed for the ECS.

In certain embodiments, heat from the exhaust can be used to heat a surrounding fluid (e.g. coiled around the exhaust) which in turn is used to evaporate LH2 in a heat exchanger and change its state from liquid to gaseous form. This source of heat may (if needed) be added in series to the compressor air heat source (e.g. multi-stage heating using three or more heating stages) in order to produce gaseous hydrogen a higher rate or to raise the gaseous hydrogen output temperature higher prior to combustion.

In certain embodiments, hot engine oil can be used in series with any of the other means of heating as described herein. While engine oil heating may be less effective on its own than other methods described above (e.g. due to the lower temperature of the oil compared to bleed air, or exhaust air, for example), engine oil heating can be used in combination with any other method to provide supplemental heat if needed or desired. The heat exchange between the cold LH2 and hot engine oil also additionally provides a means of removing heat from the engine oil which can normally be done using an air-oil heat exchanger. By reducing the size (or eliminating altogether) the air-oil heat exchanger, engine weight and aircraft drag (e.g. smaller cooler inlet) can be reduced.

In certain embodiments, heating the LH2 can be accomplished using an electric resistive heating source powered by the engine (or from another source such as an APU or other electric module) to heat a fluid which would in turn exchange heat with the cold liquid hydrogen. This differs from system <NUM> for example, because it uses an engine mounted generator rather than a separate APU or battery system in the aircraft. This means for heating could also be used in series with the other heating sources and provides the extra advantage of being able to tune the amount of fine tune the amount of heating as required by changing the amount of current delivered to the heat source. By combining these methods (e.g. any or all of the means provided herein with respect to systems <NUM>, <NUM>, <NUM>, <NUM>, <NUM>), the rate of gaseous hydrogen flow required for the engine can be maximized and controlled.

Claim 1:
A hydrogen fuel system for aircraft comprising:
a gas turbine engine (<NUM>);
a fuel feed conduit (<NUM>) defined at least in part by, in fluid series:
a liquid hydrogen tank (<NUM>) fluidly connected to a combustor (<NUM>) of the gas turbine engine;
a liquid hydrogen pump (<NUM>) to drive fuel to the combustor of the gas turbine engine;
an evaporator (<NUM>); and
an electric heat source (<NUM>, <NUM>) in thermal communication with the evaporator to add heat into a flow of hydrogen passing through the evaporator; and
an electric energy source (<NUM>) associated with the electric heat source to power the electric heat source;
wherein the fuel feed conduit is further defined by:
a gaseous hydrogen pump (<NUM>) downstream of the evaporator drive a flow of gaseous hydrogen from the evaporator to the combustor of the gas turbine engine; and
a gaseous hydrogen heater (<NUM>) downstream of the gaseous hydrogen pump and upstream of the gas turbine engine, wherein the gaseous hydrogen heater is associated with an engine heat source (<NUM>) internal to the engine to add heat to the flow of gaseous hydrogen passing through the gaseous hydrogen heater;
and wherein the system is characterized in that:
the engine heat source comprises compressor bleed air from a compressor section (<NUM>) of the gas turbine engine (<NUM>), and in that:
a bleed air conduit (<NUM>) is defined at least in part by the compressor section of the gas turbine engine and the gaseous hydrogen heater to add heat to the flow of gaseous hydrogen passing through the gaseous hydrogen heater and cool the compressor bleed air passing through the gaseous hydrogen heater; and further in that:
an environmental control system return conduit (<NUM>) defined at least in part by the compressor section of the gas turbine engine and an environmental control system (<NUM>) is configured to flow the cooled compressor bleed air to the environmental control system.