Patent Description:
There is interest in using hydrogen (H2) as a fuel in aircraft engines. One of the issues with hydrogen is its explosiveness, and for this reason, there remain challenges to using hydrogen as a fuel source in aircraft engines.

<CIT> discloses a prior art gas turbine installation operated with hydrogen-rich fuel gas. <CIT> discloses a hydrogen fuelled gas turbine. <CIT> discloses a hydrogen fueled power plant.

An aspect of the present invention provides an aircraft engine in accordance with claim <NUM>.

Reference is now made to the accompanying figures in which:
<FIG> is a schematic cross sectional view of an aircraft engine with a hydrogen fuel system and a water recovery system.

<FIG> illustrates an aircraft engine <NUM> of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication an inlet <NUM> through which air is drawn into the aircraft engine <NUM>, a compressor section <NUM> for pressurizing the air, a combustor <NUM> in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section <NUM> for extracting energy from the combustion gases. The turbine section <NUM> is configured drive a rotatable load, which is a gearbox <NUM> and propeller <NUM> in the configuration of the aircraft engine <NUM> shown in <FIG>, but which may also be other rotatable loads (e.g. a fan). The compressor and turbine sections <NUM>,<NUM> rotate about a longitudinal center axis <NUM> of the aircraft engine <NUM>.

The combustion gases generated by the combustor <NUM> flow through ducts. Referring to <FIG>, the combustion gases generated flow from the combustor <NUM> through a duct of the turbine section <NUM>, and exit the aircraft engine <NUM> via an exhaust duct <NUM> downstream of the turbine section <NUM>, relative to the flow direction of the combustion gases. The exhaust duct <NUM> is defined by radially inner and outer casing or shroud segments which are concentric about the center axis <NUM>. The volume delimited by the exhaust duct <NUM> has an annular shape about the center axis <NUM>. The combustion gases flow through the exhaust duct <NUM> before leaving the aircraft engine <NUM>. In another possible configuration of the aircraft engine <NUM>, the rotatable load driven by the turbine section <NUM> is a fan, and the aircraft engine <NUM> has a bypass duct <NUM>. The bypass duct <NUM> is shown schematically in <FIG>, and receives part of the bypass air or airflow compressed by the fan in order to divert the airflow around the core of the aircraft engine <NUM>. The volume delimited by the bypass duct <NUM> has an annular shape about the center axis <NUM>.

The fuel which is combusted in the combustor <NUM> of the aircraft engine <NUM> is hydrogen H2. The hydrogen H2 combusted in the combustor <NUM> is in gaseous form (sometimes referred to herein as "gaseous hydrogen GH2"). The hot combustion gases generated by the combustor <NUM> when the gaseous hydrogen GH2 is mixed with air and combusted are therefore almost entirely water vapor WV. The water vapor WV generated by the combustor <NUM> is completely in gaseous form. The expression "almost entirely water vapor" refers to the combustion gases being close to <NUM>% water vapor WV by composition or mass of the combustion gases other than the nitrogen in air. It is understood that there may be trace amounts of other gases in the combustion gases. For example, nitrogen at high temperatures reacts with oxygen to form small amount of nitrogen oxides (NOx). Stated differently, the water vapor WV is the main product of combustion of the hydrogen H2 and oxygen O2 in air, and the combustion gases may contain by-products of combustion in very small amounts. The water vapor WV exits the combustor <NUM> as a flow of water vapor. The combustion disclosed herein of the hydrogen H2 is in contrast to the combustion of conventional hydrocarbon fuels, such as jet fuel, whose main combustion products are in majority carbon dioxide (CO2), small quantity of water vapor and nitrogen in the consumed amount of air.

The aircraft engine <NUM> includes a hydrogen fuel system <NUM> that stores the gaseous hydrogen GH2, and which supplies the gaseous hydrogen GH2 to the combustor <NUM> to be combusted. The hydrogen fuel system <NUM> may include any number of suitable components to achieve this functionality. For example, and referring to <FIG>, the hydrogen fuel system <NUM> includes a gaseous hydrogen accumulator <NUM>. The accumulator <NUM> is any suitable body or reservoir which defines an internal volume that stores and contains the gaseous hydrogen GH2. In an embodiment, the accumulator <NUM> is a temperature-controlled pressure vessel which stores the H2 in its gaseous form GH2 under high pressure and very low temperature. The hydrogen fuel system <NUM> has one or more fuel lines <NUM>. The fuel lines <NUM> operate to carry or convey the hydrogen H2 to a destination. For example, the fuel lines <NUM> include one or more fuel supply lines 24I which are in fluid communication with the accumulator <NUM>, in order to convey the hydrogen H2 from an upstream component to the accumulator <NUM>, such that the accumulator <NUM> can be supplied with the gaseous hydrogen GH2 prior to, or during, operation of the aircraft engine <NUM>. The fuel supply lines 24I may also extend from a hydrogen fuel source <NUM> that may be internal or external to the aircraft engine <NUM>, to a component of the hydrogen fuel source <NUM> to convey liquid hydrogen LH2 to the downstream component. The fuel lines <NUM> include one or more fuel outlet lines 24O which are in fluid communication with both the accumulator <NUM> and a fuel control unit <NUM>, which schedules and meters the amount of gaseous hydrogen GH2 to the combustor <NUM>. The fuel outlet lines 24O extend from the accumulator <NUM> through the fuel control unit <NUM> to the combustor <NUM>, so that the gaseous hydrogen GH2 can be conveyed or carried by the fuel outlet lines 24O to the combustor <NUM> to be combusted. Referring to <FIG>, the hydrogen H2 supplied to the combustor <NUM> via the fuel outlet lines 24O is in gaseous form, such that gaseous hydrogen GH2 is combusted in the combustor <NUM>. The hydrogen fuel system <NUM> may also include other components and features including, but not limited to, pumps, tanks or reservoirs, air lines, fuel purge lines, actuators, nozzles, valves, manifolds, fuel schedule unit, fuel meter unit or any other component common to a fluid system.

Referring to <FIG>, the hydrogen H2 provided by the hydrogen fuel system <NUM> is the only fuel that is combusted in the combustor <NUM> in order to generate the energy needed to drive all of the components of the aircraft engine <NUM>. In an alternate embodiment, the hydrogen fuel system <NUM> is one of the fuel systems of the aircraft engine <NUM>. In this alternate embodiment, the aircraft engine <NUM> has one or more additional fuel systems which supply a hydrocarbon such as jet fuel to be combusted in the combustor <NUM> either with, or separately from, the hydrogen H2. This configuration of two fuel supply systems may be arranged as a dual fuel system that can burn one type of fuel during operation of the aircraft engine <NUM>, or can burn both types of fuels staged or concurrently during operation of the aircraft engine <NUM>. In another alternate embodiment, the hydrogen H2 provided by the hydrogen fuel system <NUM> is the fuel that is combusted in the combustor <NUM> in order to generate the energy needed to drive some of the components of the aircraft engine <NUM>, and the aircraft engine <NUM> has one or more additional fuel systems which supply a hydrocarbon such as jet fuel to be combusted in the combustor <NUM> (or elsewhere) to generate the energy needed to drive other components of the aircraft engine <NUM>. It will thus be appreciated that the hydrogen fuel system <NUM> disclosed herein may be used as the only source of fuel for the aircraft engine <NUM>, or may be used as one of multiples sources of fuel for the aircraft engine <NUM> (which is sometimes referred to as a "hybrid" fuel system or as a "dual-fuel" fuel system).

The aircraft engine <NUM> disclosed herein recuperates and reuses at least some of the water vapor WV generated from the combustion of the hydrogen H2. Referring to <FIG>, the aircraft engine <NUM> includes a water recovery system <NUM>. The water recovery system <NUM> recovers water from the combustion process to be used for the purposes set out in the present disclosure in additional detail below. In an embodiment, the water recovery system <NUM> is the only component of the aircraft engine <NUM>, or of the aircraft to which the aircraft engine <NUM> is mounted, which supplies water for the purposes described in detail below. It will be appreciated that the water recovered by the water recovery system <NUM> may also be used for purposes in addition to those set out below, by the aircraft engine <NUM> or by the aircraft to which the aircraft engine <NUM> is mounted. These other purposes include, but are not limited to, the following: humidity control, grey water uses, as cooling medium in heat exchangers, and as drinking water after being suitably treated.

Referring to <FIG>, the water recovery system <NUM> includes a water vapor collector <NUM>. The water vapor collector <NUM> functions to receive or to collect some or all of the exhaust gas in the exhaust duct <NUM>, which consists mainly of water vapor WV and small amounts of high-temperature non-reacted nitrogen, so that it can be processed and used by other components of the water recovery system <NUM> and as exhaust gas recirculation (EGR) <NUM> as described in greater detail below. The water vapor collector <NUM> may thus take any suitable form to achieve this functionality. For example, the water vapor collector <NUM> may be an annular plenum about the center axis <NUM> into which the water vapor WV is directed. The water vapor collector <NUM> may be an assembly of components, and in such an embodiment may include one or more collector lines <NUM> to divert the collected water vapor WV to other components of the water recovery system <NUM>. Referring to <FIG>, the water vapor collector <NUM> is positioned in the exhaust duct <NUM> downstream of the turbine section <NUM>, relative to the flow direction of combustion gases emanating from the combustor <NUM>. The water vapor collector <NUM> is positioned in the annular volume delimited by the exhaust duct <NUM>. In this position, the water vapor collector <NUM> may have access to the highest concentration of water vapor WV. In an alternate embodiment, the water vapor collector <NUM> is positioned in, or in fluid communication with, the duct of the turbine section <NUM>, for example adjacent to one of the turbine stages of the turbine section <NUM>. In an alternate embodiment, the water vapor collector <NUM> is positioned downstream of, or in fluid communication with, the combustor <NUM> and upstream of the turbine section <NUM>, relative to the flow direction of combustion gases emanating from the combustor <NUM>. The water vapor collector <NUM> may also include other components and features including, but not limited to, compressors, tanks or reservoirs, air lines, actuators, nozzles, valves, manifolds, or any other component common to a fluid system.

Referring to <FIG>, the water recovery system <NUM> includes a condenser <NUM> in fluid communication with an outlet of the water vapor collector <NUM>, such as via the collector lines <NUM>, to receive the water vapour WV from the water vapor collector <NUM>. The condenser <NUM> is a heat exchanger which functions to cool some or all of the water vapor WV to produce liquid water LW and/or cooled or cooler water vapor WVC, thereby condensing a significant amount of the water vapour WV. The condenser <NUM> thus extracts heat from the engine exhaust gas which consists of water vapor WV so that the water vapor WV can undergo a phase change. The water vapor WV is thus cooled by passing through the condenser <NUM>. The change of phase undergone by the water vapor WV may result in the water vapor WV becoming liquid water LW in the condenser <NUM>, in becoming a cooled or cooler water vapor WVC that is at a temperature much lower than the water vapor WV harvested by the water vapor collector <NUM>, or in becoming a mixture of liquid water LW and cooler water vapor WVC. The cooler water vapor WVC may take the form of a mist of water. The condensed liquid water LW and/or cooler water vapor WVC may be stored in a water collector <NUM> of the water recovery system <NUM>, so that it can be used by the water recovery system <NUM> as described in greater detail below. After passing through the water collector <NUM>, the lower temperature nitrogen gas N2 may be dumped to the engine overboard via the overboard line 32O. Referring to <FIG>, the water collector <NUM> is separate from the condenser <NUM> and in fluid communication with the condenser <NUM> to receive and store the liquid water LW and/or cooler water vapor WVC. In an alternate embodiment, the water collector <NUM> is a component or part of the condenser <NUM>. In an embodiment, the condenser <NUM> and the water vapor collector <NUM> are part of a sub-assembly of the water recovery system <NUM>. In an embodiment, the functionalities of the condenser <NUM> and of the water vapor collector <NUM> are performed by a single component or by a single sub-assembly of the water recovery system <NUM>.

Different cooling media may be used to cool the water vapor WV passing through the condenser <NUM>. For example, and referring to <FIG>, the condenser <NUM> is positioned at least partially in, or in fluid communication with, the airflow through the bypass duct <NUM>. The relatively cold ambient air of the airflow of the bypass duct <NUM> is a cooling medium which cools the water vapor WV passing through the condenser <NUM>, thereby transforming the water vapor WV into the liquid water LW and/or cooler water vapor WVC.

In another possible configuration of cooling media, and referring to <FIG>, one or more of the fuel inlet lines 24I extends from the hydrogen fuel source <NUM> to the condenser <NUM>, in order to supply the condenser <NUM> with cryogenic liquid hydrogen LH2. The cryogenic liquid hydrogen LH2 is thus a cooling medium which cools the water vapor WV passing through the condenser <NUM>, thereby transforming the water vapor WV into the liquid water LW and/or cooler water vapor WVC. In this configuration where the cryogenic liquid hydrogen LH2 is the cooling medium, the condenser <NUM> functions as a hydrogen evaporator because it converts the liquid hydrogen LH2 used to cool the water vapor WV into gaseous hydrogen GH2, using the heat released from the hot engine exhaust gas transformation and cooling of the water vapor WV into the liquid water LW and/or cooler water vapor WVC. Another one of the fuel supply lines 24I may extend from the condenser <NUM> to the accumulator <NUM> to transport the heated gaseous hydrogen GH2 back to the accumulator <NUM> where it may be stored, or to transport the gaseous hydrogen GH2 for storage in another body. Thus, in this configuration where the cryogenic liquid hydrogen LH2 is the cooling medium, the heat released by exhaust gas and the water vapor WV can become a heat source in a hydrogen-fueled combustion system that is used to heat up and vaporize the cryogenic liquid hydrogen LH2 into gaseous hydrogen GH2. Since significant amounts of heat may be required to vaporize the cryogenic liquid hydrogen LH2 into the gaseous hydrogen GH2 form required for combustion in the combustor <NUM>, it may be beneficial for at least some of this heat to be recuperated from the condenser <NUM> instead of being otherwise discarded out of the aircraft engine <NUM>. The condenser <NUM> is thus able to use the heat from the hot combustion gases to vaporize the liquid hydrogen LH2 into the gaseous hydrogen GH2 needed for combustion. Thus, the condenser <NUM> allows for the use of liquid hydrogen LH2 to cool the water vapor WV, which in turn heats up the liquid hydrogen LH2 into gaseous hydrogen GH2 before it is combusted. The condenser/evaporator <NUM> is thus capable of providing an alternative, or additional, heat source to evaporate the liquid hydrogen LH2 into gaseous hydrogen GH2. In an embodiment, only the heat recovered in the condenser/evaporator <NUM> is used to vaporize the liquid hydrogen LH2 into gaseous hydrogen GH2. In the condenser <NUM>, the hot exhaust gas losses its heat energy to the cold liquid hydrogen LH2.

The liquid water LW and/or cooler water vapor WVC recovered from the water vapor WV may be used for multiple purposes, some of which are now described in greater detail below.

One of these purposes includes reducing, suppressing or eliminating the undesired flammability and detonability of the hydrogen H2 in areas and regions of the aircraft engine <NUM> and/or aircraft body where auto-ignition, unintended flames, combustion and detonation are unwanted. Referring to <FIG>, the water recovery system <NUM> includes a flame and detonation mitigation system <NUM> (sometimes referred to herein simply as "detonation mitigation system <NUM>"). The detonation mitigation system <NUM> is an assembly of cooperating components which functions to reduce or suppress the risk of the hydrogen H2 auto-ignition, flame, combustion, detonating or exploding in parts of the aircraft engine <NUM> where it is not intended to. The detonation mitigation system <NUM> may include any component and feature, including but not limited to, pumps, lines, tanks or reservoirs, air lines, fuel lines, electrical wirings, actuators, valves, manifolds, or any other component common to a fluid system. Referring to <FIG>, one of the components of the detonation mitigation system <NUM> is one or more spray nozzles 38N having inlets which are in fluid communication with the condenser <NUM> and/or the water collector <NUM>, such as via an outlet line 34A. The spray nozzles 38N (sometimes referred to herein simply as the "nozzles 38N") can be hydraulically, mechanically or electrically controlled and activated, and each nozzle 38N can be sized to spray a different volume flow rate of the liquid water LW and/or cooler water vapor WVC. These nozzle 38N receive, or are provided with, the liquid water LW and/or cooler water vapor WVC, through a water supply system, which may include water and water vapor supply components, including but not limited to, water pump, scheduling controller, metering valves, or any other component common to a water and water vapor supply system, so that the nozzles 38N can spray the liquid water LW and/or cooler water vapor WVC on components of the aircraft engine <NUM> which may present a risk of hydrogen H2 flammability and detonability. For example, and referring to <FIG>, the nozzles 38N are configured to spray one or more components of the aircraft engine <NUM> that are outside of the combustor <NUM>. For example, and referring to <FIG>, the nozzles 38N are configured to spray one or more components of the hydrogen fuel system <NUM>. Referring to <FIG>, the nozzles 38N are configured to spray one or more of the fuel lines <NUM> with the liquid water LW and/or cooler water vapor WVC in order to suppress the flammability and detonability of the hydrogen H2 carried in the sprayed fuel lines <NUM>. The liquid water LW and/or cooler water vapor WVC may be sprayed by the nozzles 38N on the outside of the fuel lines <NUM> to suppress any leak of the hydrogen H2 or to suppress or reduce the ability of the hydrogen H2 to detonate within or outside the fuel lines <NUM>. The water recovery system <NUM> may have a pump, or may use a pump of the aircraft engine <NUM>, in order to pressurize the liquid water LW and/or cooler water vapor WVC so that it can be sprayed by the hydraulically, mechanically or electrically controlled and activated nozzles 38N.

Other components of the aircraft engine <NUM> may also be sprayed with the liquid water LW and/or cooler water vapor WVC by the nozzles 38N of the detonation mitigation system <NUM>. For example, and referring to <FIG>, the nozzles 38N can spray the accumulator <NUM> and/or one or more compartments <NUM> of the hydrogen fuel system <NUM> with the liquid water LW and/or cooler water vapor WVC. The compartments <NUM> may be any partially or fully enclosed volumes of the hydrogen fuel system <NUM>, for example defined by a casing of the aircraft engine <NUM>, into which the hydrogen H2 may leak. Thus, the detonation mitigation system <NUM> allows for liquid water LW and/or cooler water vapor WVC to be sprayed in confined spaces and compartments <NUM> which have, or are exposed to, hydrogen fuel components. These may include, for example, the fuel lines <NUM>, carrying parts, and the outside and inner space between the double-layer or double-walled hydrogen fuel lines <NUM>. The nozzles 38N may also be used to spray components outside of the core of the aircraft engine <NUM>. For example, and referring to <FIG>, a bypass duct fuel line 24BD is one of the fuel lines <NUM> which conveys the gaseous hydrogen GH2 to and/or from the accumulator <NUM>, or to/from another component described herein such as the condenser <NUM>. The bypass duct fuel line 24BD is exposed to the airflow in the bypass duct <NUM>, either because the bypass duct fuel line 24BD is positioned in the bypass duct <NUM> or is in fluid communication with the airflow therein. The nozzles 38N can be used to spray the bypass duct fuel line 24BD with the liquid water LW and/or cooler water vapor WVC to suppress or reduce the flammability and detonability of the hydrogen H2 therein. The nozzles 38N can be used to spray the liquid hydrogen LH2 fuel line 24I which conveys the liquid hydrogen LH2 from the hydrogen fuel source <NUM> to the condenser/evaporator <NUM>. The flame and detonation mitigation system <NUM> may thus be used to reduce or suppress flame and detonation risk outside of the core of the aircraft engine <NUM>, in areas or spaces which are exposed to ambient or bypass airflow. These areas, which may not be airtight, may be continuously flushed with airflow while simultaneously being fed with the liquid water LW and/or cooler water vapor WVC by the detonation mitigation system <NUM>. It will be appreciated that any undesirable or excess accumulation of water provided by the nozzles 38N may be drained.

The detonation mitigation system <NUM> thus allows for the liquid water LW and/or cooler water vapor WVC to be sprayed, via the nozzles 38N, in confined spaces and compartments which have, or are exposed to, hydrogen H2 fuel components in order to impede or suppress the flammability and detonability of the hydrogen H2 in these places. The sprayed liquid water LW and/or cooler water vapor WVC may thus reduce or suppress the flammability and detonation limits of hydrogen H2, and thus serve as an effective hydrogen safety measure which allows for the use of hydrogen H2 as the sole or primary fuel source for the aircraft engine <NUM>. Thus, the aircraft engine <NUM> disclosed herein allows for the use of the water vapor in the exhaust gases from the combustion of hydrogen H2 to reduce or suppress the flammability limits and the detonability limits of the hydrogen H2 in areas of the aircraft engine <NUM> where hydrogen H2 flames and detonations are not desirable or intended.

Since the water vapor WV is abundant in the combustion products of the hydrogen-fueled aircraft engine <NUM>, the water recovery system <NUM> allows for its recuperation so that it can be used for, among other purposes, hydrogen safety measures by reducing or suppressing the flammability and detonability limits of the hydrogen H2 in areas outside of the combustor <NUM>. The water recovery system <NUM> allows for providing these hydrogen safety measures without having to use or carry another diluent, which may contribute to reducing the weight of the aircraft engine <NUM> and its part count. In contrast to ground-based gas turbine engines which may combust hydrogen and use a ground-supplied water source for hydrogen flame and detonation mitigation, the water recovery system <NUM> disclosed herein is the water/cooled vapor source for hydrogen flame and detonation mitigation on the aircraft engine <NUM>, and it is airborne with the aircraft engine <NUM>. The water recovery system <NUM> may thus provided a water source for hydrogen flame and detonation mitigation on the aircraft engine <NUM>, thereby allowing the aircraft engine <NUM> to avoid the volume and weight penalty inherent to a dedicated diluent supply that might otherwise be required. The aircraft engine <NUM> disclosed herein may thus help to address some of the challenges associated with combusting hydrogen H2 to generate propulsive thrust in aero engines.

Another of the purposes for which the liquid water LW and/or cooler water vapor WVC recovered from the water vapor WV may be used is suppressing or reducing the production of nitrogen oxide (NOx) in the combustor <NUM>. Referring to <FIG>, the water recovery system <NUM> includes water injectors <NUM>. The water injectors <NUM> are nozzles or other fluid-injection devices which are in fluid communication with the condenser <NUM> and/or the water collector <NUM>, such as via the outlet line 34A. The water injectors <NUM> receive the liquid water LW and/or cooler water vapor WVC from the condenser <NUM> and/or the water collector <NUM>, then through a water supply system, which may include water and water vapor supply components, including but not limited to, water pump, scheduling controller, metering valves, or any other component common to a water and water vapor supply system, and spray or inject the liquid water LW and/or cooler water vapor WVC into the chamber of the combustor <NUM>. The collected liquid water LW and/or cooler water vapor WVC may thus be injected into the combustor <NUM>, in any desired amount, to manage the heat rise and control the peak combustion temperature in the combustor <NUM> in order to prevent or minimize the formation of NOx, and its emission by the aircraft engine <NUM>. The introduction of the liquid water LW and/or cooler water vapor WVC into the combustor <NUM> to reduce NOx generation may lower the combustor flame temperature, and thus have an impact on the performance or efficiency of the aircraft engine <NUM>. However, this may be an acceptable compromise and tradeoff which may be desirable or necessary to meet regulatory requirements, or to satisfy commitments made to customers of the aircraft engine <NUM>.

Another of the purposes for which the liquid water LW and/or cooler water vapor WVC recovered from the water vapor WV may be used is to direct all or part of the remaining exhaust gas back to the combustor <NUM>, which is commonly referred to as "Exhaust Gas Recirculation" or "EGR", through the lower-temperature EGR line <NUM>. This remaining exhaust gas is mainly the nitrogen in the air that passed through the aircraft engine <NUM>, and the remaining water vapor that has not been condensed by the condenser <NUM> and/or extracted by the detonation mitigation system <NUM>. A customer predetermined or an optimized amount of the recirculating exhaust gas may be injected into the dome region of the combustor <NUM> by one or more EGR injectors <NUM>. The exhaust gas recirculation of mainly lower temperature nitrogen and remaining amount of water vapor may dilute oxygen content inside the combustor <NUM> and lower the H2-air mixture temperature, which may collectively reduce the H2-air reaction rate and the combustion rate. This may lead to lower combustor flowfield peak temperature and subsequently lower the NOx emissions.

Referring to <FIG>, there is disclosed a method (not claimed) of recovering and using the water vapor WV generated by combustion of the hydrogen H2 in the combustor <NUM> of the aircraft engine <NUM>. The method includes collecting the water vapor WV, such as with the water vapor collector <NUM>. The method includes cooling the water vapor WP to generate the liquid water LW and/or cooler water vapor WVC, such as with the condenser <NUM>. The method includes using the liquid water LW and/or cooler water vapor WVC for one or more of the following purposes: <NUM>) to suppress accidental flame and detonation due to a leak of the hydrogen H2 (such as with the detonation mitigation system <NUM>), <NUM>) to provide an alternative heat source to evaporate the liquid hydrogen LH2 into gaseous hydrogen GH2 (such as with the condenser/evaporator <NUM>), and <NUM>) to provide a water source for NOx suppression in the combustor <NUM> (such as with the water injectors <NUM>).

Claim 1:
An aircraft engine (<NUM>) comprising:
a hydrogen fuel system (<NUM>) including an accumulator (<NUM>) of hydrogen, and one or more fuel lines (<NUM>) in fluid communication with the accumulator (<NUM>) and configured to carry the hydrogen;
a combustor (<NUM>) in fluid communication with the one or more fuel lines (<NUM>) to receive the hydrogen from the accumulator (<NUM>), the combustor (<NUM>) operable to combust the hydrogen and produce a flow of water vapor;
a turbine section (<NUM>) downstream of the combustor (<NUM>) for extracting energy from the flow of water vapor; and
a water recovery system (<NUM>) comprising:
a water vapor collector (<NUM>) positioned downstream of the turbine section (<NUM>) for receiving at least part of the water vapor;
a condenser (<NUM>) in fluid communication with the water vapor collector (<NUM>), the condenser (<NUM>) operable to receive and cool in the condenser (<NUM>) the at least part of the water vapor and thereby condense at least part of the at least part of the flow of water vapor; and
a spray nozzle (38N) in fluid communication with the condenser (<NUM>) and operable to spray the condensed part of the at least part of the flow of water vapor onto a component of the aircraft engine (<NUM>), wherein the component comprises one or more of the fuel lines (<NUM>) of the hydrogen fuel system (<NUM>).