Patent Description:
The propulsion system for commercial aircraft typically includes one or more aircraft engines, such as turbofan jet engines. The aircraft engine(s) may be mounted to a respective one of the wings of the aircraft, such as in a suspended position beneath the wing using a pylon. These engines may be powered by aviation turbine fuel, which is typically a combustible hydrocarbon liquid fuel, such as a kerosene-type fuel, having a desired carbon number and carbon to hydrogen ratio. Such fuel produces carbon dioxide emissions upon combustion and improvements to reduce such carbon dioxide emissions in commercial aircraft are desired.

<CIT> relates to an axial-flow pyrospin combustor comprising inner and outer combustor liners and a plurality of pyrospin effusion holes. The inner liner is coaxially mounted inside the outer liner, about a central combustor axis. The pyrospin effusion holes are formed in at least one of the outer combustor liner and the inner combustor liner. Each of the pyrospin effusion holes has a down angle and a back angle, which control a global swirl flow about the central axis, and promote film cooling without detachment.

Features and advantages of the present disclosure will be apparent from the following, more particular, description of various exemplary embodiments, as illustrated in the accompanying drawings, wherein like reference numbers generally indicate identical, functionally similar, and/or structurally similar elements. A gas turbine according to the invention is defined in claim <NUM>.

Features, advantages, and embodiments of the present disclosure are set forth or apparent from a consideration of the following detailed description, drawings, and claims. Moreover, it is to be understood that the following detailed description is exemplary and intended to provide further explanation without limiting the scope of the disclosure as claimed.

Various embodiments are discussed in detail below. While specific embodiments are discussed, this is done for illustration purposes only. A person skilled in the relevant art will recognize that other components and configurations may be used without departing from the spirit and scope of the present disclosure.

As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

The terms "coupled," "fixed," "attached," "connected," and the like, refer to both direct coupling, fixing, attaching, or connecting as well as indirect coupling, fixing, attaching, or connecting through one or more intermediate components or features, unless otherwise specified herein.

The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

Accordingly, a value modified by a term or terms, such as "about," "approximately," and "substantially" are not to be limited to the precise value specified. For example, the approximating language may refer to being within a one, two, four, ten, fifteen, or twenty percent margin in either individual values, range(s) of values and/or endpoints defining range(s) of values.

The term "bypass ratio," unless stated otherwise, means the bypass ratio at take off conditions. The term bypass ratio as used herein means the ratio between the mass flow rate of air flow accelerated by the engine that bypasses the engine core to the mass flow rate of the air flow entering the engine core. For example, in an exemplary engine such as the turbofan engine <NUM> depicted in <FIG> and discussed further below, the bypass ratio is the ratio of the mass flow rate of the air flow entering the bypass air flow passage <NUM> to the mass flow rate of the air flow entering the core air flow path <NUM>. The bypass ratio can also be estimated as a ratio of the area of an inlet to the bypass duct (e.g., inlet of the bypass air flow passage <NUM>, discussed below) or an area swept by a rotor (e.g., the area swept by fan blades <NUM>, discussed below) to the area of the inlet to the engine core (e.g., inlet of the core air flow path <NUM>).

The term "thrust," unless stated otherwise, means the maximum thrust at take off. This meaning of thrust is adopted when computing a core airflow parameter (relationship (<NUM>), below).

Here and throughout the specification and claims, range limitations are combined, and interchanged. Such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise.

Combustible hydrocarbon liquid fuel, such as Jet-A fuel, has long been used in gas turbine engines and the components of gas turbine engines, particularly, the combustor, have been designed for such fuels. A hydrogen fuel may be utilized to eliminate carbon dioxide emissions from commercial aircraft. Hydrogen fuel, however, poses a number of challenges as compared to combustible hydrocarbon liquid fuel, such as Jet-A fuel. Hydrogen fuel, for example, is a highly reactive fuel that burns at higher temperatures than combustible hydrocarbon liquid fuel. Hydrogen fuel also has much higher flame speeds. For example, the laminar flame speed for a hydrogen fuel of diatomic hydrogen is an order of magnitude greater than the laminar flame speed for Jet-A fuel.

When testing hydrogen fuel in current gas turbine engines with rich burn combustors, we, the inventors, observed that the higher combustion temperature of hydrogen fuel results in increased production of nitrogen oxides ("NOx"), as compared to combustible hydrocarbon liquid fuel. We also observed in our testing that NOx emissions are sensitive to combustor residence time. As noted above, hydrogen fuel is highly reactive (relative to other fuels) with a wide range of flammability limits and very high flame speeds, resulting in a very short hydrogen flame close to the front end of the combustor. With such a short flame, the post-flame residence time increases for combustors designed for Jet-A fuel. These findings resulted in a realization that when designing a hydrogen fuel combustor to meet NOx emission targets, the combustor residence time needs to be reduced by more than about fifty percent. To find a suitable combustor design for gas turbine engines using hydrogen fuel, we conceived of a wide variety of combustors having different shapes and sizes in order to determine which embodiment(s) were most promising for a variety of contemplated engine designs and thrust classes. The various embodiments, as described herein and as shown in the figures, are combustors that are sized to meet NOx emissions targets.

<FIG> is a perspective view of an aircraft <NUM> that may implement various preferred embodiments. The aircraft <NUM> includes a fuselage <NUM>, wings <NUM> attached to the fuselage <NUM>, and an empennage <NUM>. The aircraft <NUM> also includes a propulsion system that produces a propulsive thrust required to propel the aircraft <NUM> in flight, during taxiing operations, and the like. The propulsion system for the aircraft <NUM> shown in <FIG> includes a pair of engines <NUM>. In this embodiment, each engine <NUM> is attached to one of the wings <NUM> by a pylon <NUM> in an under-wing configuration. Although the engines <NUM> are shown attached to the wing <NUM> in an under-wing configuration in <FIG>, in other embodiments, the engine <NUM> may have alternative configurations and be coupled to other portions of the aircraft <NUM>. For example, the engine <NUM> may additionally or alternatively include one or more aspects coupled to other parts of the aircraft <NUM>, such as, for example, the empennage <NUM> and the fuselage <NUM>.

As will be described further below with reference to <FIG>, the engines <NUM> shown in <FIG> are gas turbine engines that are each capable of selectively generating a propulsive thrust for the aircraft <NUM>. The amount of propulsive thrust may be controlled at least in part based on a volume of fuel provided to the gas turbine engines <NUM> via a fuel system <NUM>. The fuel is stored in a fuel tank <NUM> of the fuel system <NUM>. As shown in <FIG>, at least a portion of the fuel tank <NUM> is located in each wing <NUM> and a portion of the fuel tank <NUM> is located in the fuselage <NUM> between the wings <NUM>. The fuel tank <NUM>, however, may be located at other suitable locations in the fuselage <NUM> or the wing <NUM>. The fuel tank <NUM> may also be located entirely within the fuselage <NUM> or the wing <NUM>. The fuel tank <NUM> may also be separate tanks instead of a single, unitary body, such as, for example, two tanks each located within a corresponding wing <NUM>.

Although the aircraft <NUM> shown in <FIG> is an airplane, the embodiments described herein may also be applicable to other aircraft <NUM>, including, for example, helicopters and unmanned aerial vehicles (UAV). The aircraft discussed herein are fixed-wing aircraft or rotor aircraft that generate lift by aerodynamic forces acting on, for example, a fixed wing (e.g., wing <NUM>) or a rotary wing (e.g., rotor of a helicopter), and are heavier-than-air aircraft, as opposed to lighter-than-air aircraft (such as a dirigible). The engine <NUM> may be used in various other applications including stationary power generation systems and other vehicles beyond the aircraft <NUM> explicitly described herein, such as boats, ships, cars, trucks, and the like.

<FIG> is a schematic, cross-sectional view of one of the engines <NUM> used in the propulsion system for the aircraft <NUM> shown in <FIG>. The cross-sectional view of <FIG> is taken along line <NUM>-<NUM> in <FIG>. For the embodiment depicted in <FIG>, the engine <NUM> is a high bypass turbofan engine that is referred to as a turbofan engine <NUM> herein. The turbofan engine <NUM> has an axial direction A (extending parallel to a longitudinal centerline axis <NUM>, shown for reference in <FIG>), a radial direction R, and a circumferential direction. The circumferential direction (not depicted in <FIG>) extends in a direction rotating about the axial direction A. The turbofan engine <NUM> includes a fan section <NUM> and a turbomachine <NUM> disposed downstream from the fan section <NUM>.

The turbomachine <NUM> depicted in <FIG> includes a tubular outer housing or nacelle <NUM> and an inlet <NUM>. Within the housing <NUM> there is an engine core, which includes, in a serial flow relationship, a compressor section including a booster or low-pressure (LP) compressor <NUM> and a high-pressure (HP) compressor <NUM>, a combustion section <NUM> (also referred to herein as a combustor <NUM>), a turbine section including a high-pressure (HP) turbine <NUM> and a low-pressure (LP) turbine <NUM>, and a jet exhaust nozzle section <NUM>. The compressor section, the combustor <NUM>, and the turbine section together define at least in part a core air flow path <NUM> extending from the inlet <NUM> to the jet exhaust nozzle section <NUM>. The turbofan engine <NUM> further includes one or more drive shafts. More specifically, the turbofan engine includes a high-pressure (HP) shaft or spool <NUM> drivingly connecting the HP turbine <NUM> to the HP compressor <NUM>, and a low-pressure (LP) shaft or spool <NUM> drivingly connecting the LP turbine <NUM> to the LP compressor <NUM>.

The fan section <NUM> shown in <FIG> includes a fan <NUM> having a plurality of fan blades <NUM> coupled to a disk <NUM>. The fan blades <NUM> and the disk <NUM> are rotatable, together, about the longitudinal centerline axis <NUM> by the LP shaft <NUM>. The booster <NUM> may also be directly driven by the LP shaft <NUM>, as depicted in <FIG>. The disk <NUM> is covered by a rotatable front hub <NUM> aerodynamically contoured to promote an air flow through the plurality of fan blades <NUM>. Further, an annular fan casing or outer nacelle <NUM> is provided, circumferentially surrounding the fan <NUM> and/or at least a portion of the turbomachine <NUM>. The nacelle <NUM> is supported relative to the turbomachine <NUM> by a plurality of circumferentially spaced outlet guide vanes <NUM>. A downstream section <NUM> of the nacelle <NUM> extends over an outer portion of the turbomachine <NUM> so as to define a bypass air flow passage <NUM> therebetween.

The turbofan engine <NUM> is operable with the fuel system <NUM> and receives a flow of fuel from the fuel system <NUM>. As will be described further below, the fuel system <NUM> includes a fuel delivery assembly <NUM> providing the fuel flow from the fuel tank <NUM> to the turbofan engine <NUM>, and, more specifically, to a plurality of fuel nozzles <NUM> that inject fuel into a combustion chamber <NUM> of the combustor <NUM>.

The turbofan engine <NUM> also includes various accessory systems to aid in the operation of the turbofan engine <NUM> and/or an aircraft including the turbofan engine <NUM>. For example, the turbofan engine <NUM> may include a main lubrication system <NUM>, a compressor cooling air (CCA) system <NUM>, an active thermal clearance control (ATCC) system <NUM>, and a generator lubrication system <NUM>, each of which is depicted schematically in <FIG>. The main lubrication system <NUM> is configured to provide a lubricant to, for example, various bearings and gear meshes in the compressor section, the turbine section, the HP spool <NUM>, and the LP shaft <NUM>. The lubricant provided by the main lubrication system <NUM> may increase the useful life of such components and may remove a certain amount of heat from such components through the use of one or more heat exchangers. The compressor cooling air (CCA) system <NUM> provides air from one or both of the HP compressor <NUM> or LP compressor <NUM> to one or both of the HP turbine <NUM> or LP turbine <NUM>. The active thermal clearance control (ATCC) system <NUM> acts to minimize a clearance between tips of turbine blades and casing walls as casing temperatures vary during a flight mission. The generator lubrication system <NUM> provides lubrication to an electronic generator (not shown), as well as cooling/heat removal for the electronic generator. The electronic generator may provide electrical power to, for example, a startup electrical motor for the turbofan engine <NUM> and/or various other electronic components of the turbofan engine <NUM> and/or an aircraft including the turbofan engine <NUM>.

Heat from these accessory systems <NUM>, <NUM>, <NUM>, and <NUM>, and other accessory systems, may be provided to various heat sinks as waste heat from the turbofan engine <NUM> during operation, such as to various vaporizers <NUM>, as discussed below. Additionally, the turbofan engine <NUM> may include one or more heat exchangers <NUM> within, for example, the turbine section or jet exhaust nozzle section <NUM> for extracting waste heat from an air flow therethrough to also provide heat to various heat sinks, such as the vaporizers <NUM>, discussed below.

The fuel system <NUM> of this embodiment is configured to store the fuel for the turbofan engine <NUM> in the fuel tank <NUM> and to deliver the fuel to the turbofan engine <NUM> via the fuel delivery assembly <NUM>. The fuel delivery assembly <NUM> includes tubes, pipes, and the like, to fluidly connect the various components of the fuel system <NUM> to the turbofan engine <NUM>. As discussed above, the turbofan engine <NUM>, and, in particular, the combustor <NUM> discussed herein may be particularly suited for use with hydrogen fuel (diatomic hydrogen). In the embodiments shown in <FIG>, the fuel is a hydrogen fuel comprising hydrogen, more specifically, diatomic hydrogen. In some embodiments, the hydrogen fuel may consist essentially of hydrogen.

The fuel tank <NUM> may be configured to hold the hydrogen fuel at least partially in the liquid phase and may be configured to provide hydrogen fuel to the fuel delivery assembly <NUM> substantially completely in the liquid phase, such as completely in the liquid phase. For example, the fuel tank <NUM> may have a fixed volume and contain a volume of the hydrogen fuel in the liquid phase (liquid hydrogen fuel). As the fuel tank <NUM> provides hydrogen fuel to the fuel delivery assembly <NUM> substantially completely in the liquid phase, the volume of the liquid hydrogen fuel in the fuel tank <NUM> decreases and the remaining volume in the fuel tank <NUM> is made up by, for example, hydrogen in the gaseous phase (gaseous hydrogen). As used herein, the term "substantially completely" as used to describe a phase of the hydrogen fuel, refers to at least <NUM>% by mass of the described portion of the hydrogen fuel being in the stated phase, such as at least <NUM>%, such as at least <NUM>%, such as at least <NUM>%, such as at least <NUM>%, such as at least <NUM>%, or such as at least <NUM>% by mass of the described portion of the hydrogen fuel being in the stated phase.

To store the hydrogen fuel substantially completely in the liquid phase, the hydrogen fuel is stored in the fuel tank <NUM> at very low (cryogenic) temperatures. For example, the hydrogen fuel may be stored in the fuel tank <NUM> at about -<NUM> degrees Celsius or less at atmospheric pressure, or at other temperatures and pressures to maintain the hydrogen fuel substantially in the liquid phase. The fuel tank <NUM> may be made from known materials such as titanium, Inconel®, aluminum, or composite materials. The fuel tank <NUM> and the fuel system <NUM> may include a variety of supporting structures and components to facilitate storing the hydrogen fuel in such a manner.

The liquid hydrogen fuel is supplied from the fuel tank <NUM> to the fuel delivery assembly <NUM>. The fuel delivery assembly <NUM> may include one or more lines, conduits, etc., configured to carry the hydrogen fuel between the fuel tank <NUM> and the turbofan engine <NUM>. The fuel delivery assembly <NUM> thus provides a flow path of the hydrogen fuel from the fuel tank <NUM> to the turbofan engine <NUM>. The hydrogen fuel is delivered to the engine by the fuel delivery assembly <NUM> in the gaseous phase, the supercritical phase, or both (e.g., the gaseous phase and the supercritical phase). The fuel system <NUM> thus includes a vaporizer <NUM> in fluid communication with the fuel delivery assembly <NUM> to heat the liquid hydrogen fuel flowing through the fuel delivery assembly <NUM>. The vaporizer <NUM> is positioned in the flow path of the hydrogen fuel between the fuel tank <NUM> and the turbofan engine <NUM>. The vaporizer <NUM> may be positioned at least partially within the fuselage <NUM> or the wing <NUM> (both shown in <FIG>), such as at least partially within the wing <NUM>. The vaporizer <NUM> may, however, be positioned at other suitable locations in the flow path of the hydrogen between the fuel tank <NUM> and the turbofan engine <NUM>. For example, the vaporizer <NUM> may be positioned external to the fuselage <NUM> and the wing <NUM> (both shown in <FIG>) and positioned at least partially within the pylon <NUM> (<FIG>) or the turbofan engine <NUM> (<FIG>). When positioned in the turbofan engine <NUM>, the vaporizer may be located in the nacelle <NUM>, for example. Although only one vaporizer <NUM> is shown in <FIG>, the fuel system <NUM> may include multiple vaporizers <NUM>. For example, when a vaporizer <NUM> is positioned in the turbofan engine <NUM> or in the pylon <NUM> and functions as a primary vaporizer configured to operate once the turbofan engine <NUM> is in a thermally stable condition, another vaporizer <NUM> is positioned upstream of the primary vaporizer and proximate to the fuel tank <NUM>, and functions as a primer vaporizer during start-up (or prior to start-up) of the turbofan engine <NUM>.

The vaporizer <NUM> is in thermal communication with at least one heat source <NUM>, <NUM>. In this embodiment, the vaporizer <NUM> is in thermal communication with a primary heat source <NUM> and an auxiliary heat source <NUM>. In this embodiment, primary heat source <NUM> is waste heat from the turbofan engine <NUM>, and the vaporizer <NUM> is, thus, thermally connected to at least one of the main lubrication system <NUM>, the compressor cooling air (CCA) system <NUM>, the active thermal clearance control (ATCC) system <NUM>, the generator lubrication system <NUM>, and the heat exchangers <NUM> to extract waste heat from the turbofan engine <NUM> to heat the hydrogen fuel. In such a manner, the vaporizer <NUM> is configured to operate by drawing heat from the primary heat source <NUM> once the turbofan engine <NUM> is capable of providing enough heat, via the auxiliary heat source <NUM>, to the vaporizer <NUM>, in order to facilitate operation of the vaporizer <NUM>.

The vaporizer <NUM> may be heated by any suitable heat source, and, in this embodiment, for example, the auxiliary heat source <NUM> is a heat source external to the turbofan engine <NUM>. The auxiliary heat source <NUM> may include, for example, an electrical power source, a catalytic heater or burner, and/or a bleed air flow from an auxiliary power unit. The auxiliary heat source <NUM> may be integral to the vaporizer <NUM>, such as when the vaporizer <NUM> includes one or more electrical resistance heaters, or the like, that are powered by the electrical power source. In this configuration the auxiliary heat source <NUM> may provide heat for the vaporizer <NUM> independent of whether or not the turbofan engine <NUM> is running and can be used, for example, during start-up (or prior to start-up) of the turbofan engine <NUM>.

As noted, the vaporizer <NUM> is in communication with the flow of the hydrogen fuel through the fuel delivery assembly <NUM>. The vaporizer <NUM> is configured to draw heat from at least one of the primary heat source <NUM> and the auxiliary heat source <NUM> to heat the flow of hydrogen fuel from a substantially completely liquid phase to a substantially completely gaseous phase or to a substantially completely supercritical phase.

The fuel system <NUM> also includes a high-pressure pump <NUM> in fluid communication with the fuel delivery assembly <NUM> to induce the flow of the hydrogen fuel through the fuel delivery assembly <NUM> to the turbofan engine <NUM>. The high-pressure pump <NUM> may generally be the primary source of pressure rise in the fuel delivery assembly <NUM> between the fuel tank <NUM> and the turbofan engine <NUM>. The high-pressure pump <NUM> may be configured to increase a pressure in the fuel delivery assembly <NUM> to a pressure greater than a pressure within the combustion chamber <NUM> of the combustor <NUM> of the turbofan engine <NUM>, and to overcome any pressure drop of the components placed downstream of the high-pressure pump <NUM>.

The high-pressure pump <NUM> is positioned within the flow of hydrogen fuel in the fuel delivery assembly <NUM> at a location downstream of the vaporizer <NUM>. In this embodiment, the high-pressure pump <NUM> is positioned external to the fuselage <NUM> and the wing <NUM>, and is positioned at least partially within the pylon <NUM>, or at least partially within the turbofan engine <NUM>. More specifically, the high-pressure pump <NUM> is positioned within the turbofan engine <NUM>. With the high-pressure pump <NUM> located in such a position, the high-pressure pump <NUM> may be any suitable pump configured to receive the flow of hydrogen fuel in substantially completely a gaseous phase or a supercritical phase. In other embodiments, however, the high-pressure pump <NUM> may be positioned at other suitable locations, including other positions within the flow path of the hydrogen fuel. For example, the high-pressure pump <NUM> may be located upstream of the vaporizer <NUM> and may be configured to receive the flow of hydrogen fuel through the fuel delivery assembly <NUM> in a substantially completely liquid phase.

The fuel system <NUM> also includes a metering unit in fluid communication with the fuel delivery assembly <NUM>. Any suitable metering unit may be used including, for example, a fuel metering valve <NUM> placed in fluid communication with the fuel delivery assembly <NUM>. The fuel delivery assembly <NUM> is configured to provide the fuel metering valve <NUM>, and the fuel metering valve <NUM> is configured to receive hydrogen fuel. In this embodiment, the fuel metering valve <NUM> is positioned downstream of the high-pressure pump <NUM>. The fuel metering valve <NUM> is further configured to provide the flow of the hydrogen fuel to the turbofan engine <NUM> in a desired manner. The fuel metering valve <NUM> is configured to provide a desired volume of the fuel at, for example, a desired flow rate, to a fuel manifold <NUM> of the turbofan engine <NUM>. The fuel manifold <NUM> then distributes (provides) the hydrogen fuel received to a plurality of fuel nozzles <NUM> (see <FIG>) within the combustion section <NUM> of the turbofan engine <NUM> where the hydrogen fuel is mixed with compressed air, and the mixture of hydrogen fuel and compressed air is combusted to generate combustion gases that drive the turbofan engine <NUM>. Adjusting the fuel metering valve <NUM> changes the volume of fuel provided to the combustion chamber <NUM> (see <FIG>) of the combustor <NUM> and, thus, changes the amount of propulsive thrust produced by the turbofan engine <NUM> to propel the aircraft <NUM>.

Although the turbofan engine <NUM> is shown as a direct drive, fixed-pitch turbofan engine <NUM>, in other embodiments, a gas turbine engine may be a geared gas turbine engine (i.e., including a gearbox between the fan <NUM> and shaft driving the fan, such as the LP shaft <NUM>), may be a variable pitch gas turbine engine (i.e., including a fan <NUM> having a plurality of fan blades <NUM> rotatable about their respective pitch axes), etc. Additionally, in still other exemplary embodiments, the exemplary turbofan engine <NUM> may include or be operably connected to any other suitable accessory systems. Additionally, or alternatively, the exemplary turbofan engine <NUM> may not include or be operably connected to one or more of the accessory systems <NUM>, <NUM>, <NUM>, and <NUM>, discussed above.

The turbofan engine <NUM> discussed herein is an example of the engine <NUM> in which the combustors <NUM> discussed herein may be used. In other embodiments, other suitable engines may be utilized with aspects of the present disclosure. For example, <FIG> show an unducted single fan (USF) engine <NUM> that may be used as the engine <NUM> of the aircraft <NUM> and implement the fuel system described above, and combustor designs discussed further below. <FIG> is a perspective view of the USF engine <NUM> and <FIG> is a cross-sectional view taken along a line <NUM>-<NUM> in <FIG>.

The USF engine <NUM> includes a housing <NUM>. The housing <NUM> may be formed of a nacelle <NUM> and spinner <NUM>. The nacelle <NUM> and/or the spinner <NUM> house internal components of the USF engine <NUM>. For example, the nacelle <NUM> houses a torque producing system <NUM> coupled to a shaft <NUM>. The torque producing system <NUM> in the embodiments discussed herein is a gas turbine engine, such as the turbomachine <NUM> discussed above with reference to <FIG> and, thus, the nacelle <NUM> of this embodiment is similar to the tubular outer housing <NUM> discussed above. As the turbomachine <NUM> used as the torque producing system <NUM> of the USF engine has the same or similar components and features as the turbomachine <NUM> discussed above, a detailed description of the components of the turbomachine <NUM> used in of the USF engine <NUM> is omitted.

The torque producing system <NUM> and the shaft <NUM> are configured to operate (e.g., to rotate) the spinner <NUM>. One or more fan blades <NUM> are coupled to the spinner <NUM>. More specifically, the spinner <NUM> includes a fan hub <NUM>, and the fan blades <NUM> are coupled to the fan hub <NUM>. The spinner <NUM> rotates with respect to the nacelle <NUM>. Coupled to the nacelle <NUM> may be one or more outlet guide vanes <NUM>. In this embodiment, the outlet guide vanes <NUM> are positioned aft of the fan blades <NUM>. During operation, the one or more fan blades <NUM> (by virtue of the connection to the spinner <NUM>) rotate circumferentially around a longitudinal centerline <NUM>, in this embodiment, and the nacelle <NUM> is stationary such that the one or more outlet guide vanes <NUM> do not rotate around the longitudinal centerline <NUM> and are, thus, stationary with respect to rotation about the longitudinal centerline <NUM>. Although the outlet guide vanes <NUM> are stationary with respect to the longitudinal centerline <NUM>, the outlet guide vanes <NUM> are capable of being rotated or moved with respect to the nacelle <NUM>, for example, in the direction A of <FIG>.

During operation of the USF engine <NUM>, air flows from the left side of <FIG> toward the right side of <FIG>. A portion of the air flow may flow past the fan blades <NUM> and the outlet guide vanes <NUM>. A portion of the air flow may enter the nacelle <NUM> through the annular inlet <NUM> to be mixed with the hydrogen fuel for combustion in a combustor <NUM> of the USF engine <NUM> and exit through an outlet <NUM>. The outlet guide vanes <NUM> may be movable with respect to the nacelle <NUM> to guide the air flow in a particular direction. Each outlet guide vane <NUM> may be movable to adjust the lean, pitch, sweep, or any combination thereof, of the outlet guide vane <NUM>.

In the embodiment shown in <FIG>, a forward end or front portion of the housing <NUM> includes the one or more fan blades <NUM> and the one or more outlet guide vanes <NUM>. In other embodiments, the one or more fan blades <NUM> and the one or more outlet guide vanes <NUM> may have a different arrangement with respect to the housing <NUM>. For example, the one or more fan blades <NUM> and the one or more outlet guide vanes <NUM> may be located on an aft end or rear portion of the housing <NUM>, such as coupled to a rear portion of the housing <NUM>.

In other embodiments, an engine according to the disclosure may be configured to have either the stationary vanes positioned forward of the rotating blades <NUM> (thus, the blades <NUM> are inlet guide vanes) or both the blades <NUM> and blades <NUM> configured to operate in a counter-rotating fashion. Either "pusher" or "puller" configurations are contemplated. In each of these alternative embodiments, the fuel delivery system <NUM> and combustor <NUM>, as described in great detail below, may be used. An example of a suitable engine configuration for a counter-rotating engine is shown and described in <FIG> and col. <NUM>, line <NUM> through col. <NUM>, line <NUM> of <CIT>, hereby incorporated by reference for all purposes. Alternative embodiments of the USF engine <NUM> are shown and described in <FIG>, <FIG>, and <FIG> and col. <NUM>, line <NUM> through col. <NUM>, line <NUM> of <CIT>, hereby incorporated by reference for all purposes.

In further embodiments, a turbojet engine <NUM> may be used as the engine <NUM>. <FIG> is a schematic, cross-sectional view of the turbojet engine <NUM>. The cross-sectional view of <FIG> is similar to <FIG>, which is taken along line <NUM>-<NUM> in <FIG>. The turbojet engine <NUM> includes the same or similar components of the turbomachine <NUM> of the turbofan engine <NUM> and a detailed description of these components is omitted. An exemplary turbojet engine <NUM> may not include a fan with bypass duct. An exemplary turbojet engine <NUM> may have high velocity exhaust from the engine, which produces a majority of the thrust for the turbojet engine <NUM>. In still further embodiments, other suitable gas turbine engines, such as a turboshaft engine, a turboprop engine, and the like, may be utilized with aspects of the present disclosure.

As noted above, we conceived of a wide variety of combustors having different shapes and sizes. <FIG> show various combustor shapes that can suitably be used as the combustor <NUM> for the gas turbine engines <NUM> discussed herein. <FIG> are a detail views showing detail <NUM> in <FIG>, and, as <FIG> is a cross-sectional view, <FIG> are also cross-sectional views. <FIG> shows a first combustor <NUM>. <FIG> shows a second combustor <NUM>. <FIG> shows a third combustor <NUM>. <FIG> shows a fourth combustor <NUM>. <FIG> shows a fifth combustor <NUM>. Although the shapes of these combustors <NUM>, <NUM>, <NUM>, <NUM>, <NUM> differ, each of these combustors <NUM>, <NUM>, <NUM>, <NUM>, <NUM> has similar components, and common reference numerals are used in <FIG> to for the same or similar components of these combustors <NUM>, <NUM>, <NUM>, <NUM>, <NUM>. Accordingly, the following detailed description of the first combustor <NUM> also applies to the second combustor <NUM>, the third combustor <NUM>, the fourth combustor <NUM>, and the fifth combustor <NUM>. Some components, such as the combustor casing <NUM>, for example, may not be shown in each figure, but such components may nevertheless be applicable to the combustors <NUM>, <NUM>, <NUM>, <NUM>.

As shown in <FIG>, combustor <NUM> includes a combustor casing <NUM> and a combustor liner <NUM>. The combustor casing <NUM> of this embodiment has an outer casing <NUM> and an inner casing <NUM>, and the combustor liner <NUM> of this embodiment has an outer liner <NUM> and an inner liner <NUM>. A combustion chamber <NUM> is formed within the combustor liner <NUM>. More specifically, the outer liner <NUM> and the inner liner <NUM> are disposed between the outer casing <NUM> and the inner casing <NUM>. The outer liner <NUM> and the inner liner <NUM> are spaced radially from each other such that the combustion chamber <NUM> is defined therebetween. The outer casing <NUM> and the outer liner <NUM> form an outer passage <NUM> therebetween, and the inner casing <NUM> and the inner liner <NUM> form an inner passage <NUM> therebetween. In this embodiment, the combustor <NUM> is a single annular combustor, but, in other embodiments, the combustor <NUM> may be any other combustor, including, but not limited to a double annular combustor.

The combustion chamber <NUM> has a forward end <NUM> (downstream end) and an aft end <NUM> (upstream end). The fuel nozzle <NUM> is positioned at the forward end <NUM> of the combustion chamber <NUM>. The fuel nozzle <NUM> of this embodiment is part of a swirler/fuel nozzle assembly <NUM>. In this embodiment, when the combustor <NUM> is an annular combustor <NUM>, a plurality of fuel nozzles <NUM> is arranged in an annular configuration with the plurality of fuel nozzles <NUM> (the swirler/fuel nozzle assemblies <NUM>) aligned in a circumferential direction of the combustor <NUM>.

As discussed above, the compressor section, the combustor <NUM>, and the turbine section form, at least in part, the core air flow path <NUM> extending from the annular inlet <NUM> to the jet exhaust nozzle section <NUM>. Air entering through the annular inlet <NUM> is compressed by blades of a plurality of fans of the LP compressor <NUM> and HP compressor <NUM>. A cowl assembly <NUM> is coupled to the upstream ends of outer liner <NUM> and the inner liner <NUM>, respectively. An annular opening <NUM> formed in the cowl assembly <NUM> enables compressed air from the compressor section (indicated by arrow B) to enter the combustor <NUM>. The compressed air flows through the annular opening <NUM> to support combustion. Another portion of the compressed air flows around the outside of the combustor liner <NUM> through the outer passage <NUM> and the inner passage <NUM>. This air is introduced into the combustion chamber <NUM> through a plurality of circumferentially spaced dilution holes <NUM> formed in the combustor liner <NUM> at positions downstream of the fuel nozzle <NUM>.

An annular dome plate <NUM> extends between, and is coupled to, outer liner <NUM> and the inner liner <NUM> near their upstream ends. The plurality of circumferentially spaced swirler/fuel nozzle assemblies <NUM> is coupled to dome plate <NUM>. Each swirler/fuel nozzle assembly <NUM> receives compressed air from the annular opening <NUM>. The swirler/fuel nozzle assembly <NUM> includes a swirler <NUM> that is used to generate turbulence in the air. The fuel nozzle <NUM> injects fuel into the turbulent air flow and the turbulence promotes rapid mixing of the fuel with the air. The resulting mixture of fuel and compressed air is discharged into combustion chamber <NUM> and combusted in the combustion chamber <NUM>, generating combustion gases (combustion products), which accelerate as the combustion gases leave the combustion chamber <NUM>.

A turbine nozzle <NUM> is disposed at the outlet of the combustion chamber <NUM>. The turbine nozzle <NUM> may be a stage <NUM> turbine nozzle. The turbine nozzle <NUM> is coupled to outer liner <NUM> and the inner liner <NUM> at the downstream (aft) ends of each of the outer liner <NUM> and the inner liner <NUM>. The turbine nozzle <NUM> of this embodiment includes an outer band <NUM> and an inner band <NUM> coupled to outer liner <NUM> and the inner liner <NUM>, respectively. The turbine nozzle <NUM> also includes a leading edge <NUM>, which in this embodiment is the location where the turbine nozzle <NUM> is coupled to outer liner <NUM> and the inner liner <NUM>, and the outer band <NUM> and the inner band <NUM> each has the leading edge <NUM>. The turbine nozzle <NUM> further includes a plurality of circumferentially spaced vanes <NUM> extending between the outer band <NUM> and the inner band <NUM>. The vanes <NUM> extend in a generally radial direction. The vanes <NUM>, and the turbine nozzle <NUM>, is a static component and the vanes <NUM> may be cured to direct (e.g., spin or swirl) the combustion gases to turn the turbines (e.g., drive the turbine blades) of the first stage of the HP turbine <NUM>. In this embodiment, the turbine section is a multi-stage turbine and these combustion gases will drive subsequent stages of the HP turbine <NUM> and the LP turbine <NUM>. The turbine nozzle <NUM> may, thus, also be referred to as a stage one nozzle (S1N). As discussed above the HP turbine <NUM> and the LP turbine <NUM>, among other things, drive the LP compressor <NUM> and HP compressor <NUM>.

As noted above, we realized that when designing hydrogen fuel combustor to meet NOx emission targets, the combustor residence time needs to be reduced. We sized the combustor <NUM>, and more specifically, the combustor liner <NUM> for various gas turbine engines and flow rates. These different embodiments are shown below in Table <NUM> and were developed for different bypass ratios and thrust classes of engines, characterized by the core airflow. In particular, we considered the height H, also referred to as burner dome height, of the combustion chamber <NUM> and the length L, also referred to as burner length, of the combustion chamber. Diluents could be used to suppress the temperature, and, thus, NOx production, in the combustion chamber <NUM> when hydrogen is used as the fuel. With the combustor sized as described in these embodiments, hydrogen fuel can be used without the need of diluents. In some embodiments, no diluent is added to the combustion chamber <NUM> and the fuel is substantially completely diatomic hydrogen without diluent. As used herein, the term "substantially completely," as used to describe the amount of a particular element or molecule (e.g., diatomic hydrogen), refers to at least <NUM>% by mass of the described portion of the element or molecule, such as at least <NUM>%, such as at least <NUM>%, such as at least <NUM>%, such as at least <NUM>%, such as at least <NUM>%, or such as at least <NUM>% by mass of the described portion of the element or molecule.

<FIG> illustrate how the height H and length L may be determined for the different shapes of combustion liners <NUM> shown in these figures. The height H of the combustion chamber <NUM> is taken at the forward end <NUM> of the combustion chamber <NUM>. The height H is the maximum height between an inner surface of the outer liner <NUM> and an inner surface of the inner liner <NUM> at the forward end <NUM> of the combustion chamber <NUM>. The height H is measured along a line (referred to as a forward line <NUM>, herein) that is generally orthogonal the inner surfaces of the outer liner <NUM> and the inner liner <NUM>. The forward line <NUM> may be orthogonal to a central axis <NUM> of the fuel nozzle assembly <NUM> and/or the fuel nozzle <NUM>. In this manner, the height H may be orthogonal to the central axis <NUM>. In some embodiments, the height H measured using with the forward line is the maximum height of the combustion chamber <NUM> and may also be the maximum dome height of the combustion chamber <NUM>.

The length L of the combustion chamber <NUM> is the distance between forward line <NUM> and the leading edge <NUM> of the turbine nozzle <NUM>. As with the height H, a line (referred to as the aft line <NUM>, herein) can be drawn from the leading edge <NUM> at the outer liner <NUM> and leading edge at the inner liner <NUM>. Each of the forward line <NUM> and the aft line <NUM> has a midpoint (midpoint <NUM> and midpoint <NUM>, respectively) that is halfway between the outer liner <NUM> and the inner liner <NUM>. The length L can be measured from the midpoint <NUM> of the forward line <NUM> to the midpoint <NUM> of the aft line <NUM>. The midpoint <NUM> may be the midspan height of the turbine nozzle <NUM>.

When developing a gas turbine engine, the interplay between components can make it particularly difficult to select or to develop one component during engine design and prototype testing, especially, when some components are at different stages of completion. For example, one or more components may be nearly complete, yet one or more other components may be in an initial or preliminary phase such that only one (or a few) design parameters are known. It is desired to arrive at what is possible at an early stage of design, so that the down selection of candidate optimal designs, given the tradeoffs, become more possible. Heretofore, the process has sometimes been more ad hoc, selecting one design or another without knowing the impact when a concept is first taken into consideration. For example, various aspects of the fan <NUM> design, the HP compressor <NUM> design, and/or the LP compressor <NUM> design may not be known, but such components impact the core air flow through the core air flow path <NUM>, and, thus, may influence the design of the combustion chamber <NUM>.

We desire to narrow the range of configurations or combination of features that can yield favorable results given the constraints of the design, feasibility, manufacturing, certification requirements, etc., early in the design selection process to avoid wasted time and effort. During the course of the evaluation of different embodiments as set forth above, we, the inventors, discovered, unexpectedly, that there exists a relationship between the burner length and the burner dome height, which uniquely identifies a finite and readily ascertainable (in view of this disclosure) number of embodiments suitable for a particular architecture that can meet NOx emissions for hydrogen fuel and provide desired flame residence times. This relationship is referred to by the inventors as the combustor size rating (CSR) (in), and is defined according to the following relationship (<NUM>) between burner length L (in) and burner dome height H (in): <MAT>.

As discussed further below, we have identified a range of the Combustor Size Ratings that enable a combustion chamber <NUM> to be designed for a gas turbine engine <NUM> using hydrogen fuel. This relationship is applicable over a wide range of thrust class and engine designs. Using this unique relationship, a combustor <NUM> design can be developed early in the design process that meets NOx emissions targets and reduces engine weight for gas turbine engines using hydrogen fuel.

Table <NUM> describes exemplary embodiments <NUM> to <NUM> identifying the CSR for various hydrogen fuel burning engines. The embodiments <NUM> to <NUM> may be engines with either rich burn combustors or lean burn combustors. Each of embodiments <NUM> to <NUM> burns hydrogen fuel. Embodiments <NUM> to <NUM> may represent any of the engines described with respect to <FIG> and can be applied to any of the combustion chamber <NUM> shapes shown in <FIG>. In Table <NUM>, the CSR is determined based on the relationship (<NUM>) described above. A core air flow parameter (CAFP) (kN) is defined according to the following relationship (<NUM>) between thrust (kN) and bypass ratio, both at take off. <MAT> The burner length is the length L identified with respect to <FIG>, and in the embodiments <NUM> to <NUM> is between <NUM> (two inches) and <NUM> (six inches). In embodiments <NUM> to <NUM>, the burner length squared may be between <NUM><NUM> (six square inches) and <NUM><NUM> (thirty-five square inches). The burner dome height is the height H identified with respect to <FIG>, and in the embodiments <NUM> to <NUM> is between <NUM> (two and one half inches) and <NUM> (six inches).

The length L may be between <NUM> (<NUM> inches) and <NUM> (<NUM> inches). The length L may be between <NUM> (two inches) and <NUM> (three inches). The length L may be between <NUM> (two and one half inches) and <NUM> (three and one half inches). The height H may be between <NUM> (<NUM> inches) and <NUM> (<NUM> inches). The height H may be between <NUM> (two and one half inches) and <NUM> (six inches). The height H may be between <NUM> (two and one half inches) and <NUM> (five inches). The height H may be between <NUM> (four inches) and <NUM> (five inches). The burner length squared may be between <NUM> (<NUM> inches) and <NUM> (<NUM> inches). The burner length squared may be between <NUM><NUM> (six square inches) and <NUM><NUM> (thirty-five square inches). The burner length squared may be between <NUM><NUM> (six square inches) and <NUM><NUM> (twenty square inches). The burner length squared may be between <NUM><NUM> (six square inches) and <NUM><NUM> (twelve square inches). The burner length squared may be between <NUM><NUM> (eight square inches) and <NUM><NUM> (twelve square inches). The burner length squared and the height may be any values such that the CSR is less than <NUM> (seven inches). The burner length squared and the height may be any values such that the CSR is less than <NUM> (six inches).

<FIG> represents, in graph form, the burner length, squared, as a function of the burner dome height. <FIG> shows that the burner length, squared, may be changed based on the burner dome height. An area <NUM> may present the boundaries of burner length, squared, as a function of burner dome height in which a particular combustor is designed. <FIG> represents, in graph form, the CSR as a function of core air flow parameter. Table <NUM> and <FIG> show that CSR may be changed based on a thrust class, as characterized by the core air flow parameter, of an engine. An area <NUM> may present the boundaries of CSR as a function of The core air flow parameter in which a particular combustor is designed.

As shown in <FIG>, the CSR is less than <NUM> (seven inches) for every core air flow. That is, the CSR is less than <NUM> (seven inches) for every thrust class of engine. The CSR may be between <NUM> (<NUM> inches) and <NUM> (<NUM> inches). The CSR may be between <NUM> (one inch) and <NUM> (seven inches). The CSR may be between <NUM> (one and one half inches) and <NUM> (seven inches). The CSR may be between <NUM> (two inches) and <NUM> (seven inches). The CSR may be between <NUM> (two inches) and <NUM> (six inches). The CSR may be between <NUM> (one inches) and <NUM> (five inches). The CSR may be between <NUM> (two inches) and <NUM> (five inches). The CSR may be between <NUM> (three inches) and <NUM> (five inches). The core air flow parameter may be less than sixty kN. The core air flow parameter may be between five kN and <NUM> kN. The core air flow parameter may be between two and one half kN and sixty kN. The core air flow parameter may be between ten kN and twenty kN. The core air flow parameter may be between thirty kN and forty-five kN.

Claim 1:
A gas turbine engine (<NUM>; <NUM>; <NUM>) comprising:
a hydrogen fuel delivery assembly (<NUM>) configured to deliver a hydrogen fuel flow;
a compressor section (<NUM>, <NUM>) configured to compress air flowing therethrough to provide a compressed air flow; and
a combustor (<NUM>; <NUM>; <NUM>; <NUM>; <NUM>; <NUM>) including a combustion chamber (<NUM>); wherein
the combustor (<NUM>; <NUM>; <NUM>; <NUM>; <NUM>; <NUM>) includes an outer liner (<NUM>) and an inner liner (<NUM>), each of the outer liner (<NUM>) and the inner liner (<NUM>) having an inner surface;
the combustion chamber (<NUM>) has a forward end (<NUM>);
the combustion chamber (<NUM>) is defined between the outer liner (<NUM>) and the inner liner (<NUM>);
the combustion chamber (<NUM>) has a burner length L and a burner dome height H, the burner dome height H being a maximum height between the inner surface of the outer liner (<NUM>) and the inner surface of the inner liner (<NUM>) at the forward end (<NUM>) of the combustion chamber (<NUM>);
the combustion chamber (<NUM>) is configured to combust a mixture of the hydrogen fuel flow and the compressed air flow; and
the combustion chamber (<NUM>) has a combustor size rating between <NUM> (one inch) and <NUM> (seven inches), the combustor size rating being a ratio of the burner length, squared, to the burner dome height.