Patent Description:
A gas turbine engine generally includes a turbomachine and a rotor assembly. Gas turbine engines, such as turbofan engines, may be used for aircraft propulsion. In the case of a turbofan engine, the turbomachine includes a compressor section, a combustion section, and a turbine section in serial flow order, and the rotor assembly is configured as a fan assembly.

During operation, air is compressed in the compressor and mixed with fuel and ignited in the combustion section for generating combustion gases which flow downstream through the turbine section. The turbine section extracts energy therefrom for rotating the compressor section and fan assembly to power the gas turbine engine and propel an aircraft incorporating such a gas turbine engine in flight.

At least certain gas turbine engines include a fuel cell assembly operable therewith.

<CIT> relates to a solid oxide fuel cell for placement in a gas turbine combustor. <CIT> relates to a gas turbine system in which the compressor produces compressed air for a solid oxide fuel cell generator and a topping combustor further heats the hot gas from the solid oxide fuel cell generator before it is expanded in the turbine.

Reference will now be made in detail to present embodiments of the disclosure, one or more examples of which are illustrated in the accompanying drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the disclosure.

As used herein, the term "line" may include a hose, pipe, or other fluid conduit that carries a fluid.

The term "at least one of" in the context of, e.g., "at least one of A, B, and C" or "at least one of A, B, or C" refers to only A, only B, only C, or any combination of A, B, and C.

The present disclosure is generally related to gas turbine engines having fuel cell assemblies operable therewith. Particularly, the present disclosure is related to heating one of more fluids supplied to the fuel cell assembly to increase the overall efficiency of the fuel cell assembly and the gas turbine engine. For example, fuel cell assemblies may rely on recuperation of heat from the turbine exhaust gas to heat the fuel and air prior to use in the fuel cell. This may be done via a large heat exchanger thermally coupled to the exhaust gases. However, such heat exchangers may require low Mach numbers in order to meet other mechanical limitations of the heat exchanger. As such, the present disclosure includes a turbine section having turbine blades arranged in counter rotating stages. The counter rotating stages do not have a turbine vane disposed therebetween, which advantageously results in a large reduction in the Mach number across the counter rotating stages. This reduction in the Mach number allows for a large heat exchanger to thermally couple to the exhaust gases downstream of the counter rotating stages, thereby allowing for a large heat extraction from the exhaust gases, which increases the efficiency of the heat exchanger and the gas turbine engine. Particularly, the reduction in the Mach number across the counter rotating stages allows for a large heat exchanger to be employed in the flowpath without causing blockage issues, which results in higher efficiency heat extraction and higher efficiency gas turbine engine operation.

Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures, <FIG> is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of <FIG>, the gas turbine engine is a high-bypass turbofan jet engine <NUM>, referred to herein as "gas turbine engine <NUM>. " As shown in <FIG>, the gas turbine engine <NUM> defines an axial direction A (extending parallel to a longitudinal centerline <NUM> provided for reference), a radial direction R, and a circumferential direction (i.e., a direction extending about the axial direction A; not depicted). In general, the gas turbine engine <NUM> includes a fan section <NUM> and a turbomachine <NUM> disposed downstream from the fan section <NUM>.

The exemplary turbomachine <NUM> depicted generally includes a substantially tubular outer casing <NUM> that defines an annular inlet <NUM>. The outer casing <NUM> encases, in serial flow relationship, a compressor section <NUM> including a booster or low pressure (LP) compressor <NUM> and a high pressure (HP) compressor <NUM>, a combustion section <NUM>, a turbine section including a high pressure (HP) turbine <NUM> and a low pressure (LP) turbine <NUM>, and an exhaust section <NUM>. The compressor section, the combustion section <NUM>, and the turbine section together define a core flowpath <NUM> extending from the annular inlet <NUM> through the LP compressor <NUM>, the HP compressor <NUM>, the combustion section <NUM>, the HP turbine <NUM>, the LP turbine <NUM> and the exhaust section <NUM>. The core flowpath may generally be bound by the engine frame, engine shaft, engine liner, or other components. A high pressure (HP) shaft or a spool <NUM> drivingly connects the HP turbine <NUM> to the HP compressor <NUM>. A low pressure (LP) shaft or a spool <NUM> drivingly connects the LP turbine <NUM> to the LP compressor <NUM>.

For the embodiment depicted, the fan section <NUM> includes a fan <NUM> having a plurality of fan blades <NUM> coupled to a disk <NUM> in a spaced apart manner. As depicted, the fan blades <NUM> extend outwardly from disk <NUM> generally along the radial direction R. The fan blades <NUM> and disk <NUM> are together rotatable about the longitudinal centerline <NUM> by the LP shaft or the spool <NUM>.

Referring still to the exemplary embodiment of <FIG>, the disk <NUM> is covered by a rotatable spinner cone <NUM> aerodynamically contoured to promote an airflow through the plurality of fan blades <NUM>. Additionally, the exemplary fan section <NUM> includes an annular fan casing or outer nacelle <NUM> that circumferentially surrounds the fan <NUM> and/or at least a portion of the turbomachine <NUM>. For the embodiment depicted, the outer nacelle <NUM> is supported relative to the turbomachine <NUM> by a plurality of circumferentially-spaced outlet guide vanes <NUM>. Moreover, a downstream section <NUM> of the outer nacelle <NUM> extends over an outer portion of turbomachine <NUM> so as to define a bypass airflow passage <NUM> therebetween.

During operation of the gas turbine engine <NUM>, a volume of air <NUM> enters the gas turbine engine <NUM> through an associated inlet <NUM> of the outer nacelle <NUM> and/or the fan section <NUM>. As the volume of air <NUM> passes across the fan blades <NUM>, a first portion of air <NUM> from the volume of air <NUM> is directed or routed into the bypass airflow passage <NUM> and a second portion of air <NUM> from the volume of air <NUM> is directed or routed into the LP compressor <NUM>. The ratio between the first portion of air <NUM> and the second portion of air <NUM> is commonly known as a bypass ratio. The pressure of the second portion of air <NUM> is then increased as it is routed through the HP compressor <NUM> and into the combustion section <NUM>, where it is mixed with fuel and burned to provide combustion gases <NUM>.

The combustion gases <NUM> are routed through the HP turbine <NUM> where a portion of thermal and/or kinetic energy from the combustion gases <NUM> is extracted via sequential stages of HP turbine stator vanes <NUM> that are coupled to an inner casing (not shown) and HP turbine rotor blades <NUM> that are coupled to the HP shaft or spool <NUM>, thus causing the HP shaft or spool <NUM> to rotate, thereby supporting operation of the HP compressor <NUM>.

The combustion gases <NUM> are then routed through the LP turbine <NUM> where a second portion of thermal and kinetic energy is extracted from the combustion gases <NUM> via first stage rotor blades <NUM> and second stage rotor blades <NUM> that rotate together in a first direction around the longitudinal centerline <NUM>, and final stage rotor blades <NUM> that rotate in a second direction around the longitudinal centerline <NUM> opposite the first direction. First turbine guide vanes <NUM> are disposed upstream of the first stage rotor blades <NUM>, and second turbine guide vanes <NUM> are disposed between the first stage rotor blades <NUM> and the second stage rotor blades <NUM>. The first stage rotor blades <NUM> and the second stage rotor blades <NUM> are connected via a first spool <NUM> that rotates in the first direction around the longitudinal centerline <NUM> and is coupled to a gearbox <NUM>. The final stage rotor blades <NUM> are connected to a second spool <NUM> that rotates in the second direction and is also coupled to the gearbox <NUM>. The gearbox <NUM> is further coupled to the LP shaft or spool <NUM>, such that the first stage rotor blades <NUM>, the second stage rotor blades <NUM>, and the final stage rotor blades <NUM> drive the LP shaft or spool <NUM> to rotate through the gearbox <NUM>. Thus, the LP turbine <NUM> supports operation of the LP compressor <NUM> and/or rotation of the fan <NUM>.

The combustion gases <NUM> are subsequently routed through the exhaust section <NUM> of the turbomachine <NUM> to provide propulsive thrust. The exhaust section <NUM> includes a heat exchanger <NUM> disposed immediately downstream of the final stage rotor blades <NUM>.

Simultaneously, the pressure of the first portion of air <NUM> is substantially increased as the first portion of air <NUM> is routed through the bypass airflow passage <NUM> before it is exhausted from a fan nozzle exhaust section <NUM> of the gas turbine engine <NUM>, also providing propulsive thrust. The HP turbine <NUM>, the LP turbine <NUM>, and the exhaust section <NUM> at least partially define a hot gas path for routing the combustion gases <NUM> through the turbomachine <NUM>.

Referring still to <FIG>, the gas turbine engine <NUM> additionally includes a fuel delivery system <NUM>. The fuel delivery system <NUM> generally includes a fuel source <NUM>, such as a fuel tank, and one or more fuel delivery lines <NUM>. The one or more fuel delivery lines <NUM> provide a fuel flow through the fuel delivery system <NUM> to the combustion section <NUM> of the turbomachine of the gas turbine engine <NUM>. The combustion section <NUM> includes an integrated fuel cell and combustor assembly <NUM>. The one or more fuel delivery lines <NUM>, for the embodiment depicted, provide a flow of fuel to the integrated fuel cell and combustor assembly <NUM>.

The exemplary gas turbine engine <NUM> depicted in <FIG> is by way of example only, and in other exemplary embodiments, the gas turbine engine <NUM> may have any other suitable configuration. For example, in other exemplary embodiments, the gas turbine engine <NUM> may instead be configured as any other suitable turbomachine including, e.g., any other suitable number of shafts or spools, and excluding, e.g., the fan <NUM> and/or including, e.g., a gearbox between the fan <NUM> and the LP shaft or spool <NUM>, a variable pitch fan <NUM>, etc. Accordingly, in other exemplary embodiments, the gas turbine engine <NUM> may instead be configured as, e.g., a turbojet engine, a turboshaft engine, a turboprop engine, etc., and further may be configured as an aeroderivative gas turbine engine or an industrial gas turbine engine.

Referring now to <FIG>, a simplified schematic view of a gas turbine engine <NUM> in accordance with an exemplary aspect of the present disclosure is provided. The exemplary gas turbine engine <NUM> depicted in <FIG> may be configured in substantially the same manner as exemplary gas turbine engine <NUM> described above with reference to <FIG>.

For example, as is shown, the gas turbine engine <NUM> generally includes a compressor section <NUM> and a turbine section <NUM> arranged in a serial flow order and coupled to one another via one or more shafts. In some embodiments, the compressor section <NUM> may include a compressor <NUM>, and the turbine section <NUM> may include a turbine <NUM>. For example, in some embodiments, the compressor section <NUM> may include a low pressure compressor and a high pressure compressor (such as the LP compressor <NUM> and the HP compressor <NUM> shown in <FIG>). Similarly, in such embodiments, the turbine section <NUM> may include a LP turbine and an HP turbine (such as the LP turbine <NUM> and the HP turbine <NUM> shown in <FIG>).

In exemplary embodiments, as shown in <FIG>, the turbine <NUM> may include turbine blades arranged in counter rotating stages <NUM>. In some embodiments, the turbine <NUM> may be a low pressure turbine. The counter rotating stages <NUM> may include a first-direction turbine stage <NUM> that rotates in a first direction at a first speed. Additionally, the counter rotating stages <NUM> may include a second-direction turbine stage <NUM> positioned downstream of the first-direction turbine stage <NUM> that rotates in a second direction opposite the first direction at a second speed. In many embodiments, the second speed may be lower than the first speed. For example, the first-direction turbine stage <NUM> may rotate in one of a clockwise direction or a counterclockwise direction, and the second-direction turbine stage <NUM> may rotate in the other of the clockwise direction or the counterclockwise direction. As will be appreciated, the counter rotating stages <NUM> of turbine blades may advantageously expand the turbine working fluid to a larger annulus, thereby lowering the Mach number. Reduction in the Mach number of the turbine working fluid may allow for additional thermal energy to be extracted, thereby increasing the efficiency of the heat exchanger <NUM>. Additionally, the counter rotating stages <NUM> may advantageously reduce the overall size of the aircraft engine by eliminating the need for a large diffuser section. In this way, the gas turbine engine <NUM> described herein may be more axially compact than prior designs.

As shown in <FIG>, the first-direction turbine stage <NUM> may be coupled to a first shaft or a high speed shaft <NUM>, such that the first-direction turbine stage <NUM> and the first shaft <NUM> rotate together in the first direction. Similarly, the second-direction turbine stage <NUM> may be coupled to a second shaft or low speed shaft <NUM>, such that the second-direction turbine stage <NUM> and the second shaft <NUM> rotate together in the second direction. The first shaft <NUM> and the second shaft <NUM> are not coupled to one another. In various embodiments, the second shaft <NUM> may be coupled to a generator <NUM>, such that the generator <NUM> may convert the rotational energy of the second shaft <NUM> into electrical energy. In many embodiments, the compressor <NUM> may be coupled to the high speed shaft <NUM>, such that the first-direction turbine stage <NUM> powers the compressor <NUM>.

In exemplary embodiments, a recuperator or heat exchanger <NUM> may receive exhaust gases from the turbine <NUM> downstream of the counter rotating stages <NUM> of turbine blades. For example, in some embodiments, the heat exchanger <NUM> may be disposed in the core flowpath <NUM> downstream (e.g., immediately downstream) of the counter rotating stages <NUM>. Alternatively, the heat exchanger <NUM> may be in fluid communication with the turbine <NUM> via an exhaust collection duct <NUM> (such that the heat exchanger <NUM> is positioned outside of the core flowpath <NUM>). For example, the exhaust collection duct <NUM> may extend from the turbine <NUM> to the heat exchanger <NUM>. Particularly, the exhaust collection duct <NUM> may extend from the turbine <NUM> downstream of the second-direction turbine stage <NUM> to the heat exchanger <NUM>, such that the exhaust collection duct <NUM> receives exhaust gases from downstream of the counter rotating stages <NUM>. Pulling exhaust gases from downstream of the second-direction turbine stage <NUM> is advantageous because the exhaust gases in this location have a reduced Mach number from having passes through the counter rotating stages <NUM>, and as such, the exhaust gases may be used in the heat exchanger <NUM>. Additionally, positioning the heat exchanger <NUM> outside of the core flowpath <NUM> may advantageously reduce duct losses within the core flowpath <NUM>.

In many embodiments, the gas turbine engine <NUM> may further include a fuel cell assembly <NUM> fluidly coupled to one or more fluid supply lines <NUM> for receiving one or more input fluids. The fuel cell assembly <NUM> may provide one or more output products <NUM> to the turbine section <NUM>. For example, the one or more output products <NUM> may be provided directly to the turbine section <NUM> as a working fluid for expansion in the turbine <NUM> of the turbine section <NUM>. Alternatively, or additionally, the output products <NUM> may be combusted (e.g., via an igniter) prior to entrance into the turbine section <NUM>. In exemplary embodiments, the fuel cell assembly <NUM> may be configured as a solid oxide fuel cell ("SOFC") assembly that includes a cathode, an anode, and an electrolyte. As will be appreciated, a SOFC is generally an electrochemical conversion device that produces electricity directly from oxidizing a fuel. Generally, fuel cell assemblies, and in particular fuel cells, are characterized by the electrolyte material utilized. The SOFC's of the present disclosure may generally include a solid oxide or ceramic electrolyte. This class of fuel cells generally exhibit high combined heat and power efficiency, long-term stability, fuel flexibility, and low emissions.

In many embodiments, the fluid supply lines <NUM> may include a fuel supply line <NUM> and an air supply line <NUM>. The fuel supply line <NUM> may extend from a fuel supply <NUM> to the fuel cell assembly <NUM>. The fuel supply <NUM> may be a container, tank, or other supply of fuel. The fuel supply line <NUM> may provide a fuel to the fuel cell assembly <NUM>, e.g., the fuel may be provided to the anode side of the fuel cell assembly <NUM>. The air supply line <NUM> may extend from the compressor <NUM> of the compressor section <NUM> to the fuel cell assembly <NUM>. The air supply line <NUM> may be a bleed air line from the compressor <NUM>, such that the air supply line <NUM> provides a portion of compressed air from the compressor to the fuel cell assembly. In many embodiments, air may be provided to the cathode side of the fuel cell assembly <NUM>.

In exemplary implementations, the fuel cell assembly <NUM> converts the fuel received by the fuel supply line <NUM> to the anode and the air received by the air supply line <NUM> to the cathode into electrical energy. For example, a fuel cell power output <NUM> may be generated by the fuel cell assembly <NUM> in the form of DC current. This fuel cell power output <NUM> may be directed to a power converter <NUM> in order to change the DC current into AC current that can be effectively utilized by one or more subsystems. In particular, for the embodiment depicted, the electrical power is provided from the power converter <NUM> to an electric bus <NUM>. The electric bus <NUM> may be an electric bus dedicated to the gas turbine engine <NUM>, an electric bus of an aircraft incorporating the gas turbine engine <NUM>, or a combination thereof. The electric bus <NUM> is in electric communication with one or more additional electrical devices <NUM>, which may be a power source, a power sink, or both. For example, the additional electrical devices <NUM> may be a power storage device (such as one or more batteries), an electric machine (an electric generator, an electric motor, or both), an electric propulsion device, etc. In many implementations, the generator <NUM> may produce a generator power output <NUM>, which may be provided to the power converter <NUM> (or alternatively, directly to the electric bus <NUM>) for use with the one or more additional electrical devices <NUM>.

Preheating the air and the fuel supplied to the fuel cell assembly <NUM> may advantageously increase the efficiency of the fuel cell assembly <NUM>. As such, as shown in <FIG>, the heat exchanger <NUM> may be thermally coupled to the one or more fluid supply lines <NUM> of the fuel cell assembly <NUM>. For example, the heat exchanger <NUM> may be separately fluidly coupled to each of the fuel supply line <NUM>, the air supply line <NUM>, and the exhaust collection duct <NUM>, to transfer heat from the exhaust gases in the exhaust collection duct <NUM>, the air in the air supply line <NUM>, and the fuel in the fuel supply line <NUM>. In this way, the heat exchanger <NUM> may be a three-fluid heat exchanger capable of transferring heat between the air, fuel, and exhaust gases. Spent exhaust gases <NUM> (i.e., exhaust gases having traveled through the heat exchanger <NUM>) may be exhausted to the atmosphere.

<FIG> is a schematic view of a portion of a turbine section <NUM> and an exhaust section <NUM> of a gas turbine engine according to one or more embodiments. More specifically, <FIG> shows a turbine <NUM> and a heat exchanger <NUM> disposed downstream of the turbine <NUM> within the exhaust section <NUM> according to one or more embodiments. In particular embodiments, the turbine <NUM> may be an LP turbine <NUM> having the configuration shown in <FIG>. The turbine section <NUM> is generally configured as part of a gas turbine engine defining a radial direction R and an axial direction A.

The turbine <NUM> includes counter rotating stages <NUM>, which include a first-direction turbine stage <NUM> and a second-direction turbine stage <NUM>. The first-direction turbine stage <NUM> may include the first stage rotor blade <NUM> and the second stage rotor blades <NUM>. The second-direction turbine stage <NUM> may include the final stage rotor blades <NUM>. More particularly, the turbine <NUM> may include, in serial flow order and along the axial direction A, first turbine guide vanes <NUM>, first stage rotor blades <NUM>, second turbine guide vanes <NUM>, second stage rotor blades <NUM>, and final stage rotor blades <NUM>. The exhaust section <NUM> according to one or more embodiments is disposed downstream of the turbine <NUM> in the axial direction A and includes the heat exchanger <NUM> also downstream of the final stage rotor blades <NUM> in the axial direction A.

The turbine <NUM> may form part of the core flowpath <NUM> of the gas turbine engine <NUM> (e.g., defined collectively by the compressor section, the combustion section, the turbine section, and the exhaust section). In many embodiments, as shown in <FIG>, the heat exchanger <NUM> may be disposed in the core flowpath <NUM>. In other embodiments, as shown in <FIG>, the heat exchanger <NUM> may be disposed outside of the core flowpath <NUM>. According to one or more embodiments, the heat exchanger <NUM> is a frame-integrated heat exchanger that is integrated into a frame <NUM> of the turbomachine <NUM>. According to one or more embodiments, the heat exchanger <NUM> is a heat exchanger formed separately from the frame <NUM> of the turbomachine <NUM>.

According to one or more embodiments, the first turbine guide vanes <NUM> are directly upstream of the first stage rotor blades <NUM>, the first stage rotor blades <NUM> are directly upstream of the second turbine guide vanes <NUM>, the second turbine guide vanes <NUM> are directly upstream of the second stage rotor blades <NUM>, and the second stage rotor blades <NUM> are directly upstream of the final stage rotor blades <NUM>, and the final stage rotor blades <NUM> are directly upstream of the heat exchanger <NUM>. According to one or more embodiments, the first turbine guide vanes <NUM> and the second turbine guide vanes <NUM> are stationary.

While <FIG> and <FIG> show two stages of turbine rotor blades <NUM>, <NUM> with a single stage of turbine guide vanes <NUM>, therebetween and upstream of the final stage rotor blades <NUM>, one or more embodiments may include additional stages of turbine rotor blades and turbine guide vanes. For example, the turbine <NUM> may further include third turbine guide vanes and third stage rotor blades in serial order downstream of the second stage rotor blades <NUM> and may further include fourth turbine guide vanes and fourth stage rotor blades in serial order downstream of the third stage rotor blades, and so on.

The first-direction turbine stage <NUM>, which includes the first stage rotor blades <NUM> and the second stage rotor blades <NUM>, may rotate in a first direction around the longitudinal centerline <NUM>. The first stage rotor blades <NUM> and the second stage rotor blades <NUM> are connected to each other via a first spool <NUM> that is driven by the first stage rotor blades <NUM> and the second stage rotor blades <NUM> to rotate in the first direction around the longitudinal centerline <NUM>. The first spool <NUM> may be coupled to a gearbox <NUM> as shown in <FIG>. Alternatively, as shown in <FIG>, the first spool <NUM> may be coupled to the first shaft <NUM> to rotate (and power) the compressor <NUM>. If more than two stages of rotor blades are disposed upstream of the final stage rotor blades <NUM>, the additional stage(s) of rotor blades are also connected to the first stage rotor blades <NUM> and the second stage rotor blades <NUM> via the first spool <NUM>.

The final stage rotor blades <NUM> are connected to a second spool <NUM> and rotate in a second direction around the longitudinal centerline <NUM>, which is opposite the first direction. The final stage rotor blades <NUM> are connected to the second spool <NUM> that is driven by the final stage rotor blades <NUM> to rotate in the second direction. The second spool <NUM> may also be coupled to the gearbox <NUM> as shown in <FIG>. Alternatively, as shown in <FIG>, the second spool <NUM> may be coupled to the second shaft <NUM> for generation of electrical power via the generator <NUM>.

As shown in <FIG>, the final stage rotor blades <NUM> have a significantly greater height in the radial direction R than the first stage rotor blades <NUM> and the second stage rotor blades <NUM>. The final stage rotor blades <NUM> also have a significantly greater height in the radial direction R than final stage rotor blades of conventional high speed low pressure turbines that rotate at the same speed in the same direction as the preceding stages of rotor blades. This greater height is possible due to the reduced speed of the final stage rotor blades <NUM> rotating in the second direction, as the reduced speed reduces the stresses experienced by the final stage rotor blades <NUM> compared to the aforementioned final stage rotor blades of conventional high speed low pressure turbines. Furthermore, the reduction in stresses experienced by the final stage rotor blades <NUM> may also enable materials that could not withstand the greater stresses experienced by the final stage rotor blades of conventional high speed low pressure turbines. According to one or more embodiments, the first stage rotor blades <NUM>, the second stage rotor blades <NUM>, and the final stage rotor blades <NUM> are formed of nickel alloys, or are formed of a material that comprises nickel alloys. According to one or more embodiments, the final stage rotor blades <NUM> are formed of a different material from the first stage rotor blades <NUM> and the second stage rotor blades <NUM>. According to one or more embodiments, the final stage rotor blades <NUM> are formed of titanium aluminide or a material comprising titanium aluminide. As titanium aluminide is lighter than nickel alloys, forming the final stage rotor blades <NUM> of titanium aluminide instead of nickel alloys, enabled by the lower speed of the final stage rotor blades <NUM>, results in significant weight savings that in turn enables a more efficient gas turbine engine <NUM>.

According to one or more embodiments, the lower speed of rotation of the final stage rotor blades <NUM> also reduces a Mach number of the combustion gases exiting the turbine <NUM> and entering the exhaust section <NUM>. For example, combustion gases may exit conventional high speed low pressure turbines and enter the exhaust section at a Mach number equal to around <NUM>/<NUM> Ma. According to one or more embodiments, the turbine <NUM> is structured such that combustion gases exit the turbine <NUM> and enter the exhaust section <NUM> at a Mach number of <NUM>/<NUM> Ma or lower. With a Mach number of <NUM>/<NUM> Ma or less, a frame-integrated heat exchanger that is integrated into the frame of the turbomachine <NUM> may be employed as the heat exchanger <NUM> in the exhaust section <NUM>. According to one or more embodiments, the turbine <NUM> is structured such that combustion gases exit the turbine <NUM> and enter the exhaust section <NUM> at a Mach number of <NUM>/<NUM> Ma or less. With a Mach number of a <NUM>/<NUM> Ma or lower, a traditional heat exchanger separate from the frame of the turbomachine <NUM> may be employed as the heat exchanger <NUM> in the exhaust section <NUM>.

Additionally, due to the counter-rotation between the final stage rotor blades <NUM> and the stage of the rotor blades immediately upstream of the final stage rotor blades <NUM>, no guide vane is required therebetween. Thus, compared to a conventional high speed turbines in which the final stage rotor blades rotate in the same direction as the other stages of rotor blades, a guide vane can be removed, such that a more axially compact turbine <NUM> may be formed. Additionally, the more axially compact turbine <NUM> may in turn create more space in the axial direction A for the heat exchanger <NUM>. A larger heat exchanger <NUM> extracts more heat from the exhaust section <NUM>.

<FIG> each illustrate a schematic view of various embodiments of a gas turbine engine <NUM>. <FIG> and <FIG> each illustrate a schematic view of comparative examples, not according to the claimed invention, of a gas turbine engine <NUM>. Each of these gas turbine engines <NUM> may be utilized in an aircraft for generating thrust. The exemplary gas turbine engine <NUM> depicted in <FIG> may each be configured in substantially the same manner as exemplary gas turbine engine <NUM> described above with reference to <FIG>. The gas turbine engine <NUM> may define a longitudinal centerline <NUM>. Additionally, the gas turbine engine <NUM> may define a cylindrical coordinate system having an axial direction A that extends along the longitudinal centerline <NUM>, a radial direction R that extends orthogonally to the longitudinal centerline <NUM>, and a circumferential direction C that extends around the longitudinal centerline <NUM>.

The gas turbine engine <NUM> may generally include a compressor section <NUM>, a combustion section <NUM>, a turbine section <NUM>, and an exhaust section <NUM> arranged in a serial flow order. In some embodiments, the compressor section <NUM> may include a compressor <NUM>, and the turbine section <NUM> may include a turbine <NUM>. For example, in some embodiments, the compressor section <NUM> may include a low pressure compressor and a high pressure compressor (such as the LP compressor <NUM> and the HP compressor <NUM> shown in <FIG>). Similarly, in such embodiments, the turbine section <NUM> may include a LP turbine and an HP turbine (such as the LP turbine <NUM> and the HP turbine <NUM> shown in <FIG>).

In exemplary embodiments, the compressor section <NUM>, the combustion section <NUM>, the turbine section <NUM>, and the exhaust section <NUM> may define (e.g., collectively define) a core flowpath <NUM>. The majority of the air may move through the core flowpath <NUM>, e.g., the air may be received and compressed by the compressor section <NUM>, utilized in the combustion section <NUM>, expanded through the turbine section <NUM>, and exhausted out the exhaust section <NUM>.

In many embodiments, the gas turbine engine <NUM> may further include a fuel cell assembly <NUM> fluidly coupled to one or more fluid supply lines <NUM> for receiving one or more input fluids. As shown in <FIG>, the fuel cell assembly <NUM> may be included in the combustion section <NUM> and may provide output products <NUM> to a combustor <NUM> of the combustion section <NUM>. The combustor <NUM> may ignite the output products <NUM> and provide the combustion gases to the turbine section <NUM> for expansion in the turbine <NUM>. Alternatively, as shown in <FIG>, the fuel cell assembly <NUM> may provide one or more output products <NUM> directly to the turbine <NUM> of the turbine section <NUM>. Additionally, as shown in <FIG>, the fuel cell assembly <NUM> may be disposed in the core flowpath <NUM>, such that the fuel cell assembly <NUM> may be supplied with air from the compressor <NUM> in the core flowpath <NUM> or cooled by the air flowing thereover in the core flowpath <NUM>. In many embodiments, the fluid supply lines <NUM> may include a fuel supply line <NUM> and an air supply line <NUM>. The fuel supply line <NUM> may extend from a fuel supply <NUM> to the fuel cell assembly <NUM>. The fuel supply <NUM> may be a container, tank, or other supply of fuel. The fuel supply line <NUM> may provide a fuel to the fuel cell assembly <NUM>, e.g., the fuel may be provided to the anode side of the fuel cell assembly <NUM>. The air supply line <NUM> may extend from the compressor <NUM> of the compressor section <NUM> to the fuel cell assembly <NUM>. The air supply line <NUM> may be a bleed air line from the compressor <NUM>, such that the air supply line <NUM> provides a portion of compressed air from the compressor <NUM> to the fuel cell assembly <NUM>. In many embodiments, air may be provided to the cathode side of the fuel cell assembly <NUM>.

As shown in <FIG> and <FIG>, the gas turbine engine <NUM> may include a heat exchanger <NUM> that receives exhaust gases from downstream of the counter rotating stages <NUM> of turbine blades. That is, the exhaust gases received by the heat exchanger <NUM> may come from the turbine <NUM> downstream of the counter rotating stages <NUM>. For example, the heat exchanger <NUM> may be in fluid communication with the exhaust section <NUM> via an exhaust collection duct <NUM>. For example, the exhaust collection duct <NUM> may extend from the exhaust section <NUM> to the heat exchanger <NUM>. Particularly, the exhaust collection duct <NUM> may extend from the exhaust section <NUM> downstream of the second-direction turbine stage <NUM> to the heat exchanger <NUM>, such that the exhaust collection duct <NUM> receives exhaust gases from downstream of the counter rotating stages <NUM>. Pulling exhaust gases from downstream of the second-direction turbine stage <NUM> is advantageous because the exhaust gases in this location have a reduced Mach number from having passes through the counter rotating stages <NUM>, and as such, the exhaust gases may be used in the heat exchanger <NUM>.

The heat exchanger <NUM> may thermally couple the air supply line <NUM>, the fuel supply line <NUM>, with the exhaust gases downstream of the counter-rotating stages <NUM>. In many embodiments, as shown in <FIG> and <FIG>, the heat exchanger <NUM> may be disposed outside of the core flowpath <NUM>. Additionally, the heat exchanger <NUM> may be disposed upstream (or axially inward) of the fuel cell assembly <NUM> with respect to the direction of air through the core flowpath <NUM>. Particularly, the heat exchanger <NUM> may be disposed upstream of the fuel cell assembly <NUM> and downstream of the compressor <NUM>.

In exemplary embodiments, as shown in <FIG>, and in a comparative example, not according to the claimed invention, as shown in <FIG>, the turbine <NUM> may include turbine blades arranged in counter rotating stages <NUM>. For example, the counter rotating stages <NUM> may include a first-direction turbine stage <NUM> that rotates in a first direction at a first speed. Additionally, the counter rotating stages <NUM> may include a second-direction turbine stage <NUM> positioned downstream of the first-direction turbine stage <NUM> that rotates in a second direction opposite the first direction at a second speed.

As shown in <FIG>, the first-direction turbine stage <NUM> may be coupled to a first shaft or a high speed shaft <NUM>, such that the first-direction turbine stage <NUM> and the first shaft <NUM> rotate together in the first direction. Similarly, the second-direction turbine stage <NUM> may be coupled to a second shaft or low speed shaft <NUM>, such that the second-direction turbine stage <NUM> and the second shaft <NUM> rotate together in the second direction. The first shaft <NUM> and the second shaft <NUM> are not coupled to one another. In various embodiments, as shown in <FIG>, the second shaft <NUM> may be coupled to a generator <NUM>, such that the generator <NUM> may convert the rotational energy of the second shaft <NUM> into electrical energy. In such embodiments, as shown in <FIG> and <FIG>, the generator <NUM> may be disposed downstream of the turbine section <NUM> and/or the exhaust section <NUM> with respect to the airflow direction through the core flowpath <NUM> (e.g., the generator <NUM> may be disposed on a hot-side of the gas turbine engine <NUM>). In other embodiments, as shown in <FIG>, the generator <NUM> may be disposed upstream of the turbine section <NUM> and/or the combustion section <NUM> with respect to the airflow direction through the core flowpath <NUM> (e.g., the generator <NUM> may be disposed on a cold-side of the gas turbine engine <NUM>). Alternatively, or additionally, as shown in <FIG>, the second shaft <NUM> may be coupled to a booster fan <NUM>, which may have a construction similar to the fan <NUM> described above with reference to <FIG>. In many embodiments, the compressor <NUM> may be coupled to the high speed shaft <NUM>, such that the first-direction turbine stage <NUM> powers the compressor <NUM>.

In particular embodiments, such as that shown in <FIG>, or in comparative examples, not according to the claimed invention, such as those shown in <FIG> and <FIG>, the gas turbine engine <NUM> may further include a heat recovery system <NUM>. The exemplary heat recovery system <NUM> is generally configured to extract heat from a heat source (e.g., a heat source not fully utilizing the heat being extracted therefrom) and transfer such extracted heat to a heat sink, such that the heat sink may more efficiently utilize such extracted heat.

The heat recovery system <NUM> generally includes a heat source exchanger <NUM> (i.e., a heat exchanger configured to extract heat for the heat recovery system <NUM> from a heat source of the gas turbine engine <NUM>), a heat sink exchanger <NUM> (i.e., a heat exchanger configured to transfer heat from the heat recovery system <NUM> to a heat sink of the gas turbine engine <NUM>), a thermal transfer bus <NUM>, and a pump <NUM>. Each of these components is described in greater detail as follows. In some embodiments, as shown in <FIG>, the heat recovery system <NUM> may include a single heat sink exchanger <NUM>. In some comparative examples, not according to the claimed invention, as shown in <FIG> and <FIG>, the heat recovery system <NUM> may include a first heat sink exchanger <NUM> and a second heat sink exchanger <NUM>.

In many embodiments, the heat source exchanger <NUM> is in thermal communication with the exhaust section <NUM> of the gas turbine engine <NUM> such that the heat source exchanger <NUM> extracts heat from the exhaust section <NUM>. The heat source exchanger <NUM> may be integrated into a strut extending through the exhaust section <NUM> or a liner defining at least in part the exhaust section <NUM>, or alternatively may be positioned at any other suitable location in thermal communication with an airflow/ gases through the exhaust section <NUM> of the gas turbine engine <NUM>. In many embodiments, the heat source exchanger <NUM> may be disposed in the core flowpath <NUM>.

The thermal transfer bus <NUM> may circulate a thermal fluid between the heat source exchanger <NUM> and the one or more heat sink exchangers. The thermal transfer bus <NUM> may include a supply line <NUM> and a return line <NUM>. The supply line <NUM> may extend from the heat source exchanger <NUM> to the one or more heat sink exchangers, and the return line <NUM> may extend from the one or more heat sink exchangers to the heat source exchanger <NUM>. While the supply line <NUM> and the return line <NUM> are depicted in phantom (i.e., dashed), it should be appreciated that the supply line <NUM> and the return line <NUM> are solid fluid conduits extending between one or more locations. The pump <NUM> may be disposed on the supply line <NUM>. Additionally, an expansion device <NUM> (e.g., a turbine) may be disposed in fluid communication on the return line <NUM>. For example, the expansion device <NUM> may be in fluid communication with the thermal transfer bus <NUM> downstream of the one or more heat sink exchangers and upstream of the heat source exchanger <NUM>. With such an embodiment, the expansion device <NUM> may extract additional energy from the thermal transfer fluid, increasing the efficiency of the heat recovery system <NUM> and the gas turbine engine <NUM>.

For example, the heat source exchanger <NUM> may be disposed in fluid communication on the thermal transfer bus <NUM> and positioned downstream of the counter rotating stages <NUM>. Particularly, as shown in <FIG> and <FIG>, the heat source exchanger <NUM> may be disposed in the exhaust section <NUM> immediately downstream of the second-direction turbine stage <NUM>, such that the heat source exchanger <NUM> extracts heat from the exhaust gases downstream of the second-direction turbine stage <NUM> (once the Mach number of the exhaust gases has dropped).

Referring specifically to <FIG>, the heat sink exchanger <NUM> may be disposed in fluid communication on the thermal transfer bus <NUM> and positioned upstream of the fuel cell assembly <NUM>. In such embodiments, the heat sink exchanger <NUM> may be thermally coupled to the one or more fluid supply lines <NUM> of the fuel cell assembly <NUM>. For example, as shown in <FIG>, the heat sink exchanger <NUM> may be disposed in the core flowpath <NUM> upstream of the fuel cell assembly <NUM> and downstream of an outlet to the compressor <NUM>, such that all the air from the compressor <NUM> passes through the heat sink exchanger <NUM> prior to entrance into the fuel cell assembly <NUM>. Additionally, the heat sink exchanger <NUM> may be disposed in thermal communication on the fuel supply line <NUM>, in order to preheat fuel being supplied from the fuel supply <NUM> to the fuel cell assembly <NUM>.

Referring now to <FIG> and <FIG>, the heat sink exchanger may be a first heat sink exchanger <NUM> and a second heat sink exchanger <NUM>. In such examples, instead of a single heat sink exchanger being thermally coupled to both the fuel supply line <NUM> and the air supply line <NUM> (or the air in the core flowpath <NUM>) as shown in <FIG> and described above, the first heat sink exchanger <NUM> may thermally couple to the fuel supply line <NUM>, and the second heat sink exchanger <NUM> may thermally couple to the air supply line <NUM>.

For example, the first heat sink exchanger <NUM> may be disposed in fluid communication on the thermal transfer bus <NUM> and positioned in thermal communication on the fuel supply line <NUM>. Similarly, the second heat sink exchanger <NUM> may be disposed in fluid communication on the thermal transfer bus <NUM> and positioned in thermal communication on the air supply line <NUM>. Particularly, the first heat sink exchanger <NUM> and the second heat sink exchanger <NUM> may each be disposed outside of the core flowpath <NUM>.

In exemplary implementations, the first heat sink exchanger <NUM> and the second heat sink exchanger <NUM> may separately thermally couple to the fuel supply line <NUM> and the air supply line <NUM>, which advantageously allows for the first heat sink exchanger <NUM> and the second heat sink exchanger <NUM> to be differently sized. Additionally, the first heat sink exchanger <NUM> and the second heat sink exchanger <NUM> may be configured to transfer different amounts of heat to the respective lines to which the exchangers are coupled.

Additionally, as shown in <FIG> and <FIG>, the combustion section <NUM> may include a combustor <NUM> having a fuel nozzle <NUM> that may inject a mixture of fuel and air into a combustion chamber. The mixture of fuel and air may be combusted and the combustion gases may be provided to the turbine <NUM>. As shown the fuel supply line <NUM> may include a first branch <NUM> and a second branch <NUM>. The first branch <NUM> may extend between the first heat sink exchanger <NUM> and the combustor <NUM>, and the second branch <NUM> may extend between the first heat sink exchanger <NUM> and the fuel cell assembly <NUM>.

In such examples, as shown in <FIG> and <FIG>, the fuel cell assembly <NUM> may not be included in the combustion section <NUM> (i.e., the fuel cell assembly <NUM> may not be disposed in the core flowpath <NUM>). Rather, the gas turbine engine <NUM> may include an auxiliary turbine <NUM> disposed outside of the core flowpath <NUM>, and the fuel cell assembly <NUM> may power the auxiliary turbine <NUM>. The auxiliary turbine <NUM> may include an auxiliary stage <NUM> (e.g., an auxiliary stage of turbine blades) coupled to the generator <NUM> via an auxiliary shaft <NUM>. The auxiliary turbine <NUM> may be disposed radially outward of the turbine <NUM>. Additionally, as shown in <FIG>, the auxiliary turbine <NUM> may be generally axially aligned with the turbine <NUM> and may include a plurality of turbine blades. The plurality of turbine blades in the auxiliary turbine <NUM> may be smaller than the turbine blades in the turbine <NUM>.

The auxiliary turbine <NUM> may be powered by the fuel cell assembly <NUM>. For example, as shown in <FIG> and <FIG>, the fuel cell assembly <NUM> may be disposed forward of the auxiliary turbine <NUM> with respect to the axial direction A. Alternatively stated, the fuel cell assembly <NUM> may be disposed upstream of the auxiliary turbine <NUM>. In this way, the auxiliary turbine <NUM> may receive the output products <NUM> of the fuel cell assembly <NUM>, which may act as the working fluid to power the auxiliary turbine <NUM>.

As shown in <FIG> and <FIG>, the fuel cell power output <NUM> and the generator power output <NUM> may be directed to a power converter <NUM> in order to change the DC current into DC current or AC current that can be effectively utilized by one or more subsystems. In particular, for the example depicted, the electrical power is provided from the power converter <NUM> to an electric bus <NUM>. The electric bus <NUM> may be an electric bus dedicated to the gas turbine engine <NUM>, an electric bus of an aircraft incorporating the gas turbine engine <NUM>, or a combination thereof. The electric bus <NUM> is in electric communication with one or more additional electrical devices <NUM>, which may be a power source, a power sink, or both. For example, the additional electrical devices <NUM> may be a power storage device (such as one or more batteries), an electric machine (an electric generator, an electric motor, or both), an electric propulsion device, etc..

As shown in <FIG>, the first shaft <NUM> may be coupled to the compressor <NUM>, and the second shaft <NUM> may be coupled to the booster fan <NUM>, which may have a construction similar to the fan <NUM> described above with reference to <FIG>. Alternatively, as shown in <FIG>, the turbine <NUM> may not include counter rotating stages <NUM>, such that a singular shaft couples the turbine <NUM> to the compressor <NUM>. In such examples, the heat source exchanger <NUM> may be disposed downstream of the turbine <NUM>.

<FIG> each illustrate a schematic view of a portion of a turbine section <NUM> of a gas turbine engine according to one or more alternative embodiments of the present disclosure. The turbine section <NUM> may be incorporated into any of the gas turbine engines <NUM> described above with reference to <FIG> and <FIG>. More specifically, <FIG> each illustrate various embodiments of a turbine <NUM> and a heat exchanger <NUM>. In some embodiments, the heat exchanger <NUM> may be the heat source exchanger <NUM> described above.

The turbine <NUM> includes counter rotating stages <NUM>, which include a first-direction turbine stage <NUM> and a second-direction turbine stage <NUM>. The first-direction turbine stage <NUM> may include the first stage rotor blade <NUM> and the second stage rotor blades <NUM>. The second-direction turbine stage <NUM> may include the final stage rotor blades <NUM>. More particularly, the turbine <NUM> may include, in serial flow order and along the axial direction A, first turbine guide vanes <NUM>, first stage rotor blades <NUM>, second turbine guide vanes <NUM>, second stage rotor blades <NUM>, and final stage rotor blades <NUM>. In some embodiments, as shown in <FIG>, a third turbine guide vane <NUM> may be disposed axially between the second stage rotor blades <NUM> and the final stage rotor blades <NUM>.

As shown in <FIG>, the first-direction turbine stage <NUM> may be coupled to a gearbox <NUM> via a first spool <NUM>, and the second-direction turbine stage <NUM> may be coupled to a gearbox <NUM> via a second spool <NUM>. According to one or more embodiments, the gearbox <NUM> may be structured as a planetary gear system in which that the first spool <NUM> is fixed or connected to a ring gear, the second spool <NUM> is fixed or coupled to a sun gear, and a generator shaft <NUM> is fixed or coupled to the planet gears. In this way, the first spool <NUM> may rotate the fastest, the second spool <NUM> may rotate the slowest, and the generator shaft <NUM> may rotate at a speed between the first spool and the second spool. The generator shaft <NUM> may be coupled to the generator <NUM> to convert rotational energy into electrical energy. Alternatively, as shown in <FIG>, first spool <NUM> may be coupled to the first shaft <NUM> in order to rotate (and power) the compressor <NUM>, and the second spool <NUM> may be coupled to the second shaft <NUM> for generation of electrical power via the generator <NUM>.

Claim 1:
An aircraft engine comprising:
a compressor section (<NUM>) comprising a compressor (<NUM>);
a turbine section (<NUM>) downstream of the compressor section (<NUM>), the turbine section (<NUM>) including a turbine having turbine blades arranged in counter rotating stages (<NUM>);
one or more fluid supply lines (<NUM>);
a fuel cell assembly (<NUM>) fluidly coupled to the one or more fluid supply lines (<NUM>) for receiving one or more input fluids, the fuel cell assembly (<NUM>) in fluid communication with the turbine section (<NUM>) to provide one or more output products (<NUM>) to the turbine section (<NUM>); and
a heat exchanger (<NUM>) in fluid communication with the turbine downstream of the counter rotating stages (<NUM>) of turbine blades to receive exhaust gases from the turbine, the heat exchanger (<NUM>) thermally coupled to the one or more fluid supply lines (<NUM>) of the fuel cell assembly (<NUM>).