Aircraft propulsion system with an inter-turbine combustor

Abstract

A vehicle propulsion system and a computing system are provided. The propulsion system includes a low-pressure (LP) spool and a core engine having a high-pressure (HP) spool. A frame is positioned between the HP turbine and the LP turbine in a serial flow arrangement. The frame includes an inter-turbine burner that includes struts forming an outlet opening of a core flow path into the propulsion system. A first fuel system is configured to flow liquid fuel to a combustion section to produce a first combustion gas. A second fuel system is configured to flow gaseous fuel to the core flow path via the inter-turbine burner to produce a second combustion gas. The propulsion system forms a rated power output ratio between the core engine and the inter-turbine burner with the LP spool that is between 1.5 and 5.7.

Description

Technical Field

[0001] The present subject matter generally relates to aircraft propulsion systems. The present subject matter particularly relates to structures and methods for engine operation of an aircraft propulsion system. Background Art

[0002] Conventional aircraft propulsion systems are typically configured to generate all levels of thrust from combustion gases from a combustion system positioned between a high-pressure compressor (HPC) and a high-pressure turbine (HPT). Accordingly, the sizes of the HPC, the combustion system, and the HPT are sized to produce the entire range of thrust output or the maximum thrust output.

[0003] Some propulsion systems include a reheat system (e.g., an augmenter or an afterburner) to produce an increased amount of thrust. However, such systems are generally inefficient in terms of fuel consumption, and such systems further produce emissions or noise levels that exceed regulatory levels for emissions and noise, such as for commercial and general aviation aircraft. Reheat systems for non-aircraft gas turbine engines (e.g., industrial gas turbines for power generation) do not need to consider propulsion efficiency and overall aircraft weight, performance, and efficiency.

[0004] Accordingly, there is a need for an improved aircraft propulsion system that can produce a large amount of thrust without adversely affecting the emissions output and fuel consumption. Summary of the Invention

[0005] Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.

[0006] Aspects of the present disclosure relate to an aircraft propulsion system. The propulsion system includes a low-pressure (LP) spool that includes a fan section, an LP compressor, and an LP turbine. A core engine includes a high-pressure (HP) compressor, a combustion section, and an HP turbine. The HP compressor and the HP turbine together form a rotatable HP spool. A frame is positioned between the HP turbine and the LP turbine in a serial flow arrangement. The frame includes an inter-turbine burner that includes struts that form an outlet opening of a core flow path into the propulsion system. A first fuel system includes a first fuel conduit that is in fluid communication with a fuel nozzle at the combustion section and is configured to flow liquid fuel to the combustion section to generate a first combustion gas. A second fuel system includes a second fuel conduit that is in fluid communication with the core flow path via the outlet opening at the inter-turbine burner and is configured to flow gaseous fuel to the core flow path to generate a second combustion gas. The LP compressor, the HP compressor, the combustion section, the HP turbine, the inter-turbine burner, and the LP turbine are in a serial flow arrangement. The propulsion system forms a rated power output ratio between the core engine and the inter-turbine burner with the LP spool that is between 1.5 and 5.7.

[0007] Another aspect of the present disclosure relates to a computing system for an aircraft propulsion system. The computing system includes one or more processors and one or more memories, where the memories are configured to store instructions that, when executed by the processors, cause the propulsion system to operate. The operation includes: flowing liquid fuel to the combustion section of the propulsion system; generating, at the combustion section, a first combustion gas that is 85% or less of the rated power output of the propulsion system; and modulating the rotational speed of the LP spool via a modulated flow of gaseous fuel to the inter-turbine burner to generate a second combustion gas.

[0008] These and other features, aspects, and advantages of the present invention will be better understood with reference to the following description and the appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0009] A complete and enabling disclosure of the present invention, including the best mode thereof, for the ordinary skilled person in the art, is set forth in the specification, which makes reference to the accompanying drawings, in which:

[0010] Figure 1 is an exemplary embodiment of an aircraft including a propulsion system according to aspects of the present disclosure;

[0011] Figure 2 is for an Figure 1 aircraft according to aspects of the present disclosure of a schematic cross-sectional view of a propulsion system;

[0012] Figure 3 is a schematic cross-sectional view of a part of a propulsion system including an embodiment of an inter-turbine combustor according to aspects of the present disclosure; and

[0013] Figure 4 is a flow chart outlining steps of a method for operating a propulsion system according to aspects of the present disclosure.

[0014] Reference numerals reused in this specification and the drawings are intended to represent the same or similar features or elements of the present invention. Detailed Description

[0015] Reference will now be made in detail to embodiments of the present invention, one or more examples of which are illustrated in the drawings. Each example is provided for the purpose of explaining the present invention and not for limiting the present invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made to the present invention without departing from the scope or spirit of the present invention. For example, features shown or described as part of one embodiment can be used with another embodiment to yield yet another embodiment. Accordingly, the present invention is intended to cover such modifications and variations that fall within the scope of the appended claims and their equivalents.

[0016] As used herein, the terms "first", "second", and "third" may be used interchangeably to distinguish one component from another and are not intended to denote the position or importance of the respective components.

[0017] The terms "upstream" and "downstream" refer to the relative direction with respect to the fluid flow in a fluid path. For example, "upstream" refers to the direction from which the fluid flows and "downstream" refers to the direction towards which the fluid flows.

[0018] Embodiments of an aircraft and a propulsion system are provided that include an inter-turbine reheating combustor positioned between a first turbine and a second turbine. A first fuel system supplies liquid fuel to a combustion section to generate combustion gases to the turbines. A second fuel system supplies gaseous fuel to the inter-turbine combustor to selectively generate reheating gases based on specific engine or aircraft operating conditions. Embodiments of the propulsion system are also configured to generate a specific thrust or power output ratio based on the first fuel system and together with the first fuel system and the second fuel system.

[0019] Embodiments of the propulsion system and aircraft provided herein allow for a relatively small core engine size (i.e., the size and power output of the high-pressure spool, combustion section, and high-pressure turbine together), while generating a rated power output similar to that of a larger core engine size by increasing the power extracted from the low-pressure spool. The smaller core engine size allows for reduced fuel consumption, reduced emissions, a greater bypass ratio, and an improved specific fuel consumption rate. The smaller core engine size also allows the propulsion system to be operated as an auxiliary power unit (APU) to power aircraft subsystems, electronics, or provide engine starting power for other propulsion systems without the use of a dedicated APU separate from the propulsion system. Such a system allows for improved overall aircraft efficiency, for example, by eliminating the need or desire for a non-propulsion gas turbine engine.

[0020] Now referring to the drawings, in Figure 1 there is provided an exemplary embodiment of a vehicle 100 that includes a propulsion system 10 having an inter-turbine combustor according to aspects of the present disclosure. In an embodiment, the vehicle 100 is an aircraft that includes an aircraft structure or airframe 105. The airframe 105 includes a fuselage 110 to which wings 120 and a tail 130 are attached. A propulsion system 10 according to aspects of the present disclosure is attached to one or more portions of the airframe. In various embodiments, the aircraft 100 includes a thermal management system 200 that is configured to desirably distribute thermal loads, such as adding or removing heat from one or more fluids or structures (such as, but not limited to, oxidizer at the propulsion system, fuel, lubricant, hydraulic fluid, pneumatic fluid, or cooling fluid for an electric machine, electronics, computing system, environmental control system, gear assembly, or other system or structure).

[0021] In various embodiments, the aircraft 100 includes subsystems that generally define electrical loads that require an input of energy. Such systems include an ice protection system 160, an environmental control system 150, and an avionics system 140. The propulsion system 10 is configured to extract energy from one or more spools to power aircraft subsystems, as described herein. Although some systems may be formed as mechanical systems, the electrification of the systems can reduce the weight and complexity of the aircraft. However, such electrification typically requires a greater energy input, such as from the propulsion system 10 described herein.

[0022] In some cases, the propulsion system 10 is attached to the rear of the fuselage 110. In some other cases, the propulsion system 10 is attached below, above, or through a portion of the wing 120 and / or the empennage 130. In various embodiments, the propulsion system 10 is attached to the airframe 105 via pylons or other mounting structures. In other embodiments, the propulsion system 10 is housed within the airframe, as can be exemplified in certain supersonic commercial aircraft.

[0023] Now referring to Figure 2 , a schematic cross-sectional view of a propulsion system for an aircraft in accordance with an exemplary embodiment of the present disclosure is provided. As Figure 2 shown, the propulsion system 10 defines an axial direction A (extending parallel to the longitudinal centerline 12 provided for reference), a radial direction R, and a circumferential direction (i.e., the direction extending around the axial direction A; not depicted). In various embodiments, the propulsion system 10 is configured as a gas turbine engine, such as a turbofan engine. In a particular embodiment, the propulsion system 10 is a ducted open rotor engine (i.e., without a nacelle surrounding the fan blades). Generally, the propulsion system 10 includes a fan section 14 and a turbine 16 disposed downstream of the fan section 14.

[0024] The exemplary turbine 16 shown generally includes a substantially tubular housing 18 that defines an annular inlet 20. The housing 18 surrounds in a serial flow relationship: a compressor section that includes a first booster or low pressure (LP) compressor 22 and a second high pressure (HP) compressor 24; a combustion section 26; a turbine section that includes a first high pressure (HP) turbine 28 and a second low pressure (LP) turbine 30; and an exhaust nozzle section 32. A high pressure (HP) shaft 34 drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft 36 drivingly connects the LP turbine 30 to the LP compressor 22. The compressor section, the combustion section 26, the turbine section, and the exhaust nozzle section 32 are arranged in a serial flow order and together define a core air flow path 37 through the turbine 16.

[0025] In some embodiments, the propulsion system 10 includes one or more electric motors 370 operatively coupled to a spool of the engine. The electric motor 370 can be operatively coupled to the HP spool, the LP spool, or both, to extract or receive energy from the spool during operation. Additionally, the electric motor 370 can be configured to output or release energy to the spool to initiate or assist in the rotation of the HP spool (e.g., during startup or other desired operations), or to output or release energy to the LP spool during desired operations of the aircraft (e.g., during cruise operation, or transient conditions, or relative bursts of thrust or power output). In various embodiments described herein, the HP spool can be allowed to operate under substantially steady-state conditions, such as to allow for substantially steady-state extraction of energy into the electric motor. The electric motor can release energy to one or more subsystems at the aircraft 100 (e.g., subsystems 140, 150, 160). In particular, embodiments of the propulsion system 10 provided herein allow for increased energy extraction from the HP spool. Further or alternatively, the system 10 can allow for power extraction during ground operating conditions (including ground idle or taxi conditions).

[0026] In certain embodiments, for example Figure 2 as depicted, the fan section 14 can include a variable pitch fan 38. The turbine 16 is operatively coupled to the fan 38 to drive the fan 38. The fan 38 includes a plurality of rotatable fan blades 40 coupled to the disk 42 in a spaced-apart manner. As shown, the fan blades 40 generally extend radially outward from the disk 42 along the radial direction R. By operatively coupling the fan blades 40 to a suitable actuating member 44, each fan blade 40 can rotate relative to the disk 42 about a pitch axis P, and the actuating member 44 is configured to, for example, collectively and uniformly change the pitch of the fan blades 40. The fan blades 40, the disk 42, and the actuating member 44 rotate together about the longitudinal axis 12 via the LP shaft 36 of the cross power gearbox 46. The cross power gearbox 46 includes a plurality of gears for reducing the rotational speed of the LP shaft 36 to a more efficient fan rotational speed. Thus, for the depicted embodiment, the turbine 16 is operatively coupled to the fan 38 via the cross power gearbox 46.

[0027] Still referring to Figure 2, a second portion of the compressed air 64 from the compressor section is mixed with liquid fuel and burned within the combustion section to provide combustion gases 66. The combustion gases 66 are routed from the combustion section 26 through the HP turbine 28, where a portion of the thermal and / or kinetic energy from the combustion gases 66 is extracted via a series of stages of HP turbine stator vanes 68 coupled to the outer casing 18 and HP turbine rotor blades 70 coupled to the HP shaft 34, thus causing the HP shaft 34 to rotate and thereby supporting the operation of the HP compressor 24. The combustion gases 66 are then routed through the LP turbine 30, where a second portion of the thermal and kinetic energy is extracted from the combustion gases 66 via a series of stages of LP turbine stator vanes 72 coupled to the outer casing 18 and LP turbine rotor blades 74 coupled to the LP shaft 36, thus causing the LP shaft 36 to rotate and thereby supporting the operation of the LP compressor 22 and / or the rotation of the fan 38.

[0028] The combustion gases 66 are subsequently routed through the jet exhaust nozzle section 32 of the turbine 16. At the same time, as the first portion of air 62 is routed through the bypass air flow passage 56 before being discharged from the fan nozzle exhaust section 76 of the propulsion system 10, the pressure of the first portion of air 62 is significantly increased, also providing propulsion thrust. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 for routing the combustion gases 66 through the turbine 16.

[0029] It should be understood that Figure 2 the exemplary propulsion system 10 depicted in Figure 2 is a turbofan propulsion system 10 of a relatively high power rating. Thus, when operating at rated speed, the propulsion system 10 can be configured to produce a relatively large amount of thrust. More specifically, when operating at rated speed, the propulsion system 10 can be configured to produce at least about 14,000 pounds of thrust, or at least 18,000 pounds of thrust, or at least 21,000 pounds of thrust, or at least 24,000 pounds of thrust, or at least 30,000 pounds of thrust. Certain embodiments can produce up to 120,000 pounds of thrust at rated speed. Thus,

[0030] It should be understood that other exemplary embodiments of the propulsion system 10 are turboprop propulsion systems 10 of a relatively high power rating. Thus, when operating at rated speed, the propulsion system 10 can be configured to produce a relatively large amount of horsepower. More specifically, when operating at rated speed, the propulsion system 10 can be configured to produce up to 10,000 shaft horsepower (shp). In various embodiments, when operating at rated speed, the propulsion system 10 can be configured to produce at least 2,000 shaft horsepower (shp).

[0031] In addition, it should be understood that Figure 2 the exemplary propulsion system 10 depicted in Figure 2 is only an example, and in other exemplary embodiments, the propulsion system 10 may have any other suitable configuration. For example, in certain exemplary embodiments, the fan may not be a variable pitch fan. Additionally or alternatively, aspects of the present disclosure may be used with any other suitable aero gas turbine engine (such as a turboshaft engine, a turboprop engine, a turbojet engine, etc.). Further embodiments may omit the nacelle surrounding the fan blades, for example, to form an open rotor turbofan engine.

[0032] It should be understood that, as used herein, the rotation and modulation of the speeds of the HP spool and the LP spool correspond to the generation and modulation of output torque, power, or thrust. In the turbofan configuration of the propulsion system, most of the thrust is generated by the rotation of the fan blades via the LP spool. In various embodiments, the remaining thrust is generated by the combustion gases exhausted through the exhaust injection nozzle.

[0033] Now referring to Figure 3 , there is provided Figure 2 a close-up view of a portion of the exemplary propulsion system 10 of Figure 2 . More specifically, Figure 3 a close-up view of the combustion section 26 and the turbine section is provided. In a particular embodiment, the combustion section 26 includes a burner assembly 100. The burner assembly 100 may be configured as a deflagration burner assembly, such as, but not limited to, an annular burner, a dual annular burner, a can-annular burner, a can burner, a vortex burner, or other suitable combustion systems. The burner assembly may be configured as a lean burner, a rich burner, a rich-quench-lean (RQL) burner, or other suitable burner assemblies.

[0034] In one embodiment, the combustion section 26 includes a first fuel conduit formed, for example, by one or more fuel nozzles 124, the one or more fuel nozzles 124 being configured to receive a liquid fuel stream (schematically shown by arrow 352) and supply the liquid fuel to the combustion chamber 114 for combustion or detonation. Although not depicted in further detail, the fuel nozzles 124 may be any suitable type of fuel injector, nozzle, rail, or other liquid fuel dispensing device, atomizing device, or mixing device. In a particular embodiment, the fuel nozzles 124 may be configured for lean or rich mixtures, combustion, or detonation.

[0035] In some embodiments, the burner assembly 100 generally includes a liner 102 extending generally axially between a rear end and a front end, and an outer liner 108 also extending generally axially between a rear end and a front end. The liner 102 and the outer liner 108 together at least partially define a combustion chamber 114 therebetween. The liner 102 and the outer liner 108 are each attached to or integrally formed with an annular dome. More specifically, the annular dome includes an inner dome section 116 integrally formed with the front end 106 of the liner 102 and an outer dome section 118 formed generally with the front end of the outer liner 108. Additionally, the inner dome section 116 and the outer dome section 118 can each be integrally formed (or alternatively can be formed of multiple components attached in any suitable manner), and can each extend circumferentially to define an annular shape.

[0036] However, it should be understood that in other embodiments, the burner assembly 100 may not include the inner dome section 116 and / or the outer dome section 118; may include separately formed inner dome section 116 and / or outer dome section 118 attached to the respective liner 102 and outer liner 108; or may have any other suitable configuration. In other embodiments, the combustion section 26 can be configured as a detonation combustion system, such as a rotating detonation combustion system or a pulse detonation combustion system.

[0037] Still referring to Figure 3 , the burner assembly 100 further includes a plurality of fuel-air mixers spaced circumferentially (not shown) and at least partially positioned within the annular dome. More specifically, the plurality of fuel-air mixers are at least partially disposed radially between the outer dome section 118 and the inner dome section 116. Compressed air from the compressor section of the propulsion system 10 flows into or through the fuel-air mixers, where the compressed air is mixed with fuel and ignited to produce combustion gases 66 within the combustion chamber 114. The inner dome section 116 and the outer dome section 118 are configured to help provide such a flow of compressed air from the compressor section into or through the fuel-air mixers 124. For example, the outer dome section 118 may include a cowl at the front end, while the inner dome section 116 similarly includes an inner cowl at the front end. The cowl and the inner cowl can help direct the flow of compressed air from the compressor section into or through one or more of the fuel-air mixers. However, again in other embodiments, the annular dome can be configured in any other suitable manner.

[0038] Some embodiments of the combustor section 26 or turbine section may include one or more components formed from ceramic matrix composite (CMC) materials. In certain embodiments, both the inner liner 102 and the outer liner 108 are formed from CMC materials. In certain embodiments, the vanes or struts of the frame 300, further described below, are formed from CMC materials. Further embodiments include one or more vane or blade stages of the LP turbine 30 formed from CMC materials. CMC materials are non-metallic materials with the ability to withstand high temperatures. Exemplary CMC materials for such components may include silicon carbide (SiC), silicon nitride, or alumina matrix materials and combinations thereof. Ceramic fibers may be embedded in the matrix, such as oxidation-stable reinforcing fibers, including monofilaments such as sapphire and silicon carbide (e.g., SCS-6 from Textron), and rovings and yarns including silicon carbide (e.g., Ube Industries' and Dow Corning's ), aluminosilicate (e.g., Nextel 440 and 480), and chopped whiskers and fibers (e.g., Nextel 440 and ), and optionally ceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite). For example, in certain embodiments, a fiber bundle including a ceramic refractory coating is formed into a reinforcing tape, such as a unidirectional reinforcing tape. Multiple tapes may be stacked together (e.g., as a laminate) to form a preform component. The fiber bundle may be impregnated with a slurry composition before or after forming the preform. The preform may then be heat treated (e.g., cured or burned out) to produce a high-carbon residue in the preform, and subsequently chemically treated (e.g., melt infiltrated with silicon) to obtain a component formed from a CMC material having a desired chemical composition. In other embodiments, the CMC material may be formed as, for example, carbon fiber cloth rather than tape. Additionally or alternatively, the CMC material may be formed in any other suitable manner or using any other suitable materials.

[0039] Still referring to Figure 3 , and as further discussed above and below, combustion gases 66 flow from the combustor 114 into and through the turbine section of the propulsion system 10, where a portion of the thermal energy and / or kinetic energy from the combustion gases 66 is extracted via a sequential stage of turbine stator vanes and turbine rotor blades within the HP turbine 28 and LP turbine 30. More specifically, as Figure 3 illustrates, combustion gases 66 from the combustor 114 flow into the HP turbine 28 located immediately downstream of the combustor 114, where the thermal energy and / or kinetic energy from the combustion gases 66 is extracted via a sequential stage of HP turbine stator vanes 68 and HP turbine rotor blades 70.

[0040] As also referenced above Figure 2 As discussed above, the HP turbine 28 is coupled via the HP shaft 34 to the HP compressor 24 to form an HP spool or HP rotor that is operable to generally higher maximum speeds than the LP spool formed by the LP compressor 22, LP turbine 30, LP shaft 36, and fan section 14. Accordingly, rotation of the multi-stage HP turbine rotor blades 70 correspondingly rotates the multi-stage HP compressor rotor blades 80.

[0041] Figure 2-3 The exemplary propulsion system 10 is configured to operate to maintain the temperature of the HP turbine 28 below the maximum operating temperature of the various components therein while reducing the cooling flow extracted from the compressor section. In a particular embodiment, the HP turbine includes one or more stages of blades that are formed as substantially solid, impermeable airfoils at the core flow path. In other embodiments, the HP turbine includes one or more stages of blades that are configured to reduce the cooling flow therethrough by reducing or eliminating the cooling flow from the compressor section to the HP turbine, e.g., by reducing or eliminating the amount of air removed from the thermal cycle (i.e., air removed from combustion) to increase engine efficiency.

[0042] Return reference Figure 3 , the turbine section includes an inter-turbine frame 300 positioned between the HP turbine 28 and the LP turbine 30. The frame 300 is configured as a stationary, static support structure that is configured to support one or both of the HP turbine 28 or the LP turbine 30. The frame 300 includes an inter-turbine combustor 310 configured to allow a gaseous fuel flow to enter the core flow path upstream of the LP turbine. The frame forms the combustor and one or more orifices or openings 306 at the vanes or struts 312 of the frame, the one or more orifices or openings 306 being configured to allow the gaseous fuel 362 to flow through the struts 312 into the core flow path.

[0043] In various embodiments, the inter-turbine burner 310 forms a second fuel conduit that is configured to deliver gaseous fuel 362 to the core flow path. In a particular embodiment, the burner 310 is formed as a flame stabilizer at the strut 312 of the frame 300 as compared to, for example, a combustion system formed at a combustion section. In some embodiments, the inter-turbine burner includes struts or vanes formed as airfoils and / or structural members that typically provide, for example, for an inter-turbine frame, an intermediate frame structure, or other support frame. The strut 312 includes a forward edge or leading edge 304 and a rearward edge or trailing edge 302. The strut 312 includes a hollow portion to allow fluid flow therethrough. In some embodiments, the frame 300 includes a lubricant conduit 316 and an air conduit 314 that typically provide, for example, lubricant or air for the bearing assembly 320. The inter-turbine frame 300 may further include a conduit 308 configured to discharge a gaseous fuel 362 flow through an orifice 306 at the strut 312. In a particular embodiment, the orifice 306 is located at the trailing edge 302 of the strut 312 to allow the gaseous fuel to flow therethrough and rearwardly towards the LP turbine 30.

[0044] The combustion section 26 is configured as a deflagration or detonation combustion section. A liquid fuel 352 flow is provided to the combustion section 26 through one or more fuel nozzles 124. The liquid fuel 352 flow is mixed with compressed air from the compressor section and then combusted to produce combustion gases 66. The liquid fuel 352 provided to the combustion section 26 is a liquid injection fuel or an aviation turbine fuel such as a kerosene-based fuel, a naphtha-based fuel, or the like (e.g., Jet-A, Jet-B, JP8, biofuel, synthetic fuel, or other suitable aviation fuel). The gaseous fuel 362 provided to the inter-turbine burner is a gaseous fuel such as hydrogen (H2), natural gas, methane, syngas, or other suitable type of gaseous fuel. The gaseous fuel 362 flow released through the inter-turbine burner 310 between the HP turbine 28 and the LP turbine 30 is mixed with the combustion gases 66 flow.

[0045] It should be understood that the gaseous fuel 362 has a gaseous fuel ignition temperature (i.e., a second ignition temperature) that is less than the liquid fuel ignition temperature (i.e., a first ignition temperature) of the liquid fuel 352. The gaseous fuel 362 also has a gaseous fuel combustion speed (i.e., a second combustion speed) that is greater than the liquid fuel combustion speed (i.e., a first combustion speed) of the liquid fuel 352. The relatively low second ignition temperature limit allows a mixture of the gaseous fuel 362 from the inter-turbine burner 310 and the combustion gases 66 from the combustion section 26 to produce a second combustion gas using the relatively high-speed fluid flow through the turbine section. The flame speed of the second combustion gas produced by the inter-turbine burner 310 may also be greater than the flame speed of the first combustion gas produced by the combustion section 26.

[0046] In some embodiments, the upper flammability limit of the gaseous fuel 362 is greater than that of the liquid fuel. Additionally, in some embodiments, the range of the flammability limits is generally greater than that of the liquid fuel. In some embodiments, the lower flammability limit of the gaseous fuel is lower than the upper flammability limit of the liquid fuel. Further, the gaseous fuel has a higher degree or magnitude of flammability than the liquid fuel. Thus, unlike an afterburner system that utilizes liquid fuel, a mixture of the gaseous fuel and the combustion gas can burn without external ignition (e.g., using an igniter or other energy input).

[0047] The aircraft 100 and the propulsion system 10, either alone or together, include a first fuel system 350 for flowing and distributing the liquid fuel 352 at the combustion section 26, and a second fuel system 360 for flowing and distributing the gaseous fuel 362 at the inter-turbine burner 310. It should be understood that, in contrast to the second fuel system 360, the first fuel system 350 can be further configured to provide the liquid fuel 352 as an actuation fluid and / or a heat exchange fluid (e.g., receiving heat or thermal energy from another fluid or surface). More specifically, the first fuel system 350 can be configured to provide an actuating force or pressure to modulate one or more valves, actuators, doors, openings, nozzles, flow devices, or adjustable areas at the propulsion system, such as variable area nozzles, bleed valves, exhaust nozzles, active clearance control valves or doors, transient or start bleed valves, or other actuatable parts of the propulsion system or the aircraft.

[0048] The embodiments of the aircraft 100 and the propulsion system 10 depicted and described herein can provide improved propulsion system and aircraft efficiency, emissions, or fuel combustion. The inter-turbine burner 310 can increase the LP turbine 30 power extraction at a given high pressure (HP) spool or core engine size (i.e., HP compressor 24, combustion section 26, and HP turbine 38). The second fuel system 360 configured to supply the gaseous fuel 362 to the inter-turbine burner 310 is separate from the first fuel system 350 configured to supply the liquid fuel 352 to the combustion section 26, allowing for an increase in LP turbine power extraction and power output that is greater than the output power from the core engine alone.

[0049] It should be understood that although described as an inter-turbine burner between the HP turbine and the LP turbine, the various embodiments provided herein can include an inter-turbine burner between a first turbine receiving higher pressure combustion gas and a second turbine receiving lower pressure combustion gas. Thus, the various embodiments can include an intermediate pressure (IP) turbine that is generally positioned between the HP turbine and the LP turbine. Particular embodiments can position the inter-turbine burner described herein between the HP turbine and the IP turbine, or between the IP turbine and the LP turbine.

[0050] Further, although described as a conventional turbine rotor, embodiments of the HP turbine or LP turbine provided herein may be configured as interdigitated or bladeless turbine assemblies.

[0051] Embodiments of the aircraft 100 and propulsion system 10 provided herein allow sizing and operating the core engine at steady-state speed and power output, particularly for a hybrid-electric propulsion system, and / or avoiding generating power from a separate auxiliary power unit (APU). In some embodiments, the propulsion system 10 is configured to produce a work split between a core engine including a high-pressure (HP) spool and a combustion section and a low-pressure (LP) spool including an inter-turbine burner. In various embodiments, the propulsion system 10 has a rated power output ratio between the core engine and the inter-turbine burner 310 with an LP spool (i.e., LP turbine 30, LP compressor 22, and fan section 14) between 1.5 and 5.7. In some embodiments, the propulsion system is configured to produce an 85 / 15 work split between the core engine and the LP spool. In other words, the core engine is configured to operate the HP spool at a maximum speed corresponding to 85% of the rated power output of the propulsion system 10. The propulsion system 10 is also configured to produce up to 15% of the rated power output of the propulsion system via the LP spool and the inter-turbine burner using gaseous fuel and combustion gases generated from the core engine. Such a ratio may allow a significant reduction in the thermal load imparted to downstream turbine components from the combustion section 26, which may allow improved durability and reduced cooling flow, which may improve the overall propulsion system efficiency.

[0052] In another embodiment, the propulsion system is configured to produce an 80 / 20 work split between the core engine, the LP spool, and the inter-turbine burner. In yet another embodiment, the propulsion system is configured to produce a 75 / 25 work split between the core engine, the LP spool, and the inter-turbine burner. In yet another embodiment, the propulsion system is configured to produce a 60 / 40 work split between the core engine, the LP spool, and the inter-turbine burner. In still other various embodiments, the propulsion system is configured to generate between 60% and 85% of the maximum power output via combustion gases generated from the core engine and generate the remainder of the maximum power output using gaseous fuel and combustion gases generated from the core engine via the LP spool and the inter-turbine burner.

[0053] In various embodiments, the work is distributed between the remaining difference of the low power output and the maximum power output. In other words, the work distribution is a limit between the low power output operating conditions, above which (via the inter-turbine burner and the gaseous fuel flow), the operating conditions are high power output conditions. In certain embodiments, the maximum power output is in particular the rated power output with reference to the maximum rotational speed of the propulsion system during normal operation. For example, the propulsion system can operate at a rated speed or a rated power output during maximum load operation (e.g., during takeoff operation regarding the Landing Takeoff (LTO) cycle). In certain embodiments, the limit or demarcation of the work distribution of the maximum power output (e.g., typically 60%-85%, such as 85%, or 80%, or 75%, or 60%) corresponds to the cruise or descent operation of the propulsion system and the aircraft with respect to the difference between the LTO cycle and the rated power output of the propulsion system. Thus, certain embodiments of the propulsion system are configured for the maximum rotational speed from the operation of only the first fuel system that provides only liquid fuel (i.e., operation without an inter-turbine burner) to the operation of the core engine corresponding to cruise conditions. In certain embodiments, the propulsion system is configured for the maximum power output or the rated power output from the operation of both the combustion section with the first fuel system and the inter-turbine burner with the second fuel system.

[0054] It should be understood that those skilled in the art understand that the ranges and ratios of the work distribution provided herein correspond to the specific structures and dimensions of the core engine, the inter-turbine burner, and the LP spool. A typical aircraft gas turbine propulsion engine is designed, sized, and configured to produce 100% of the maximum power output via combustion gases that are generated in the combustion section and extracted via the LP spool. Certain aircraft gas turbine propulsion engines utilize an afterburner or a reheating system that is configured to utilize a portion of the liquid fuel, which is typically directed to the main burner in the combustion section and mixed with the combustion gases downstream of the main burner to further produce thrust (i.e., the afterburner). However, such a typical afterburner system is generally not applicable to commercial aircraft or other aircraft subject to emissions output limitations. In addition, such a system that utilizes liquid fuel is generally complex, with an igniter system and complexities associated with the lower flammability of liquid fuel. Such a system typically produces a certain level of emissions, smoke, or noise that may prohibit its use on commercial aircraft.

[0055] Return reference Figure 2, the propulsion system 10 may further include a computing system 210 configured to operate a propulsion system 10 such as described herein. The computing system 210 may correspond to any suitable processor-based device, including one or more computing devices such as those described above. In certain embodiments, the computing system 210 is a full-authority digital engine controller (FADEC) for a gas turbine engine, or other computing module or controller configured to execute instructions for operating a gas turbine engine. For example, Figure 2 FIG. illustrates an embodiment of suitable components that may be included within the computing system 210. The computing system 210 may include a processor 212 and an associated memory 214 configured to perform various computer-implemented functions.

[0056] As shown, the computing system 210 may include control logic 216 stored in the memory 214. The control logic 216 may include instructions that cause the one or more processors 212 to operate when executed by the one or more processors 212. Additionally, the computing system 210 may further include a communication interface module 230. In several embodiments, the communication interface module 230 may include associated electronic circuitry for sending and receiving data. Thus, the communication interface module 230 of the computing system 210 can be used to send data to and / or receive data from the propulsion system 10. Additionally, the communication interface module 230 can also be used to communicate with any other suitable components of the propulsion system 10, such as those described herein.

[0057] It should be understood that the communication interface module 230 can be any combination of suitable wired and / or wireless communication interfaces, and can thus be communicatively coupled to one or more components of the power generation system via wired and / or wireless connections or distributed networks. The communication interface module 230 can include any suitable wired and / or wireless communication links for transmitting communications and / or data, as described herein. For example, the module 230 can include SATCOM networks, ACARS networks, ARINC networks, SITA networks, AVICOM networks, VHF networks, HF networks, Wi-Fi networks, WiMAX networks, gatelink networks, etc.

[0058] A method for operating a propulsion system of an aircraft (hereinafter referred to as "Method 1000") is provided. The method can be carried out using the aircraft and propulsion system as described above or other suitable systems. In a particular embodiment, Method 1000 can be carried out using the computing system 210 of the propulsion system 10 or the aircraft 100, for example, a computer-implemented method. It should be understood that the computing system 210 and Method 1000 provided herein can allow for improved propulsion efficiency, reduced emissions output, and overall improvement in engine and aircraft operation. Certain embodiments can provide benefits specific to the propulsion system and aircraft under limitations in emissions output, noise, or thrust.

[0059] Method 1000 includes flowing liquid fuel to a combustion section of the propulsion system at 1010. Method 1000 includes generating a first combustion gas at 1020 that is 85% or less of the rated power output of the propulsion system in the combustion section. Method 1000 includes modulating the rotational speed of the LP spool at 1030 by modulating the gaseous fuel flow to the inter-turbine burner to generate a second combustion gas, as depicted and described herein.

[0060] In various embodiments, Method 1000 includes operating the core engine and the low-pressure (LP) spool with an inter-turbine burner at a core engine to rated power output ratio of the inter-turbine burner with an LP spool between 1.5 and 5.7 at 1022, as described above. In a particular embodiment, Method 1000 includes operating the high-pressure (HP) spool at a maximum rotational speed between 60% and 85% of the rated power output of the propulsion system at 1024. Method 1000 includes flowing gaseous fuel to the inter-turbine burner at 1026 to generate the rated power output of the propulsion system. Thus, Method 1000 can operate the engine under substantially steady-state operating conditions with a liquid fuel flow of up to 60% to 85% of the rated power output, and Method 1000 can cause the engine to modulate the gaseous fuel flow to generate the remainder or a portion of the rated power output.

[0061] In some embodiments, method 1000 includes operating a high pressure (HP) spool at a steady-state speed at 1040 while modulating the gaseous fuel flow to the inter-turbine burner. In some embodiments, the operation includes receiving a control signal corresponding to a high power operating mode of the propulsion system at 1042. In some embodiments, receiving a control signal corresponding to a high power operating mode includes a rated power operation or a takeoff operating mode of the propulsion system. In other embodiments, the high power operating mode corresponds to climb, descent, or approach or takeoff conditions relative to the LTO cycle. Method 1000 includes, at 1044, flowing gaseous fuel to the inter-turbine burner to generate a second combustion gas corresponding to the difference between the rated power output of the propulsion system and the power output generated by flowing liquid fuel to the combustion section.

[0062] In another embodiment, method 1000 includes receiving a control signal corresponding to a low power operating mode of the propulsion system at 1046. In certain embodiments, the low power operating mode corresponds to cruise conditions relative to the LTO cycle. Method 1000 includes, at 1048, reducing the gaseous fuel flow to the inter-turbine burner to reduce the power output of the propulsion system. In certain embodiments, method 1000 includes operating the high pressure (HP) spool at a steady-state speed at 1050 while reducing the gaseous fuel flow to the inter-turbine burner. In yet another specific embodiment, reducing the gaseous fuel flow to the inter-turbine burner to reduce the power output of the propulsion system corresponds to changing the operating mode of the propulsion system from a high power operating mode to a low power operating mode.

[0063] It should be understood that those skilled in the art will understand the elapsed time, tolerance, range, or deviation of a given speed or power output corresponding to "steady-state" operating conditions. In certain embodiments, those skilled in the art will understand "steady-state" in the context of an aircraft propulsion system. In yet another specific embodiment, those skilled in the art will understand "steady-state", speed, or power output as provided herein in the context of the landing takeoff cycle of an aircraft.

[0064] It should be understood that the embodiments of the propulsion system 10, aircraft 100, and method 1000 provided herein include combinations of elements, subsystems, arrangements, and configurations that provide unexpected benefits over known elements alone or in known arrangements and configurations. For example, it should be understood that having separate fuel systems and methods for control, such as via a first fuel system 350 and a second fuel system 360, and the method 1000 provided herein, introduce elements that may have been considered additional complex or convoluted heretofore, such as discouraging implementation in certain propulsion systems and aircraft (such as commercial or general aviation aircraft). However, as provided herein, the present disclosure describes systems, methods, and specific combinations or arrangements that provide unexpected benefits over the complexity associated with separate fuel systems.

[0065] Such benefits include allowing a substantially steady state speed or operation of the HP spool while increasing and decreasing the power output of the propulsion system. Such benefits can allow one or more propulsion systems of an aircraft to be operated to generate electrical power for aircraft subsystems during idle operating conditions, runway taxi or gate side operations, or other situations where known aircraft propulsion systems may not be operable due to higher fuel consumption compared to using an auxiliary power unit (APU) to generate electrical power for the aircraft or other propulsion systems. Accordingly, embodiments of the propulsion systems and engines provided herein can eliminate the need or desire for an APU in an aircraft, such as to reduce aircraft weight and improve aircraft efficiency.

[0066] Such benefits can also include allowing the core engine to be smaller in size and consume less fuel to produce a rated power output of a known propulsion system having a relatively large core engine. Embodiments provided herein allow the core engine of the propulsion system to perform more typical APU operations and are different from the operations typically performed for aircraft propulsion systems. Additionally, embodiments provided herein allow for improved emissions output over known propulsion systems, such as via a reduced core engine size and improved emissions output from gaseous fuel to produce a rated power output under specific engine operating conditions. Further still, by providing a second fuel system 360 and method 1000 for operating under specific operating conditions, problems associated with gaseous fuel are alleviated compared to using gaseous fuel for substantially all operating conditions.

[0067] This written description uses examples to disclose the invention, including the best mode, and also enables those skilled in the art to practice the invention, including making and using any device or system and performing any incorporated methods. The patentable scope of the invention is defined by the claims and may include other examples that occur to those skilled in the art. If these other examples include structural elements that are identical to the literal language of the claims, or if they include equivalent structural elements that are not substantially different from the literal language of the claims, then these other examples are intended to fall within the scope of the claims.

[0068] Further aspects of the invention are provided by the subject matter of the following clauses:

[0069] 1. An aircraft propulsion system, the propulsion system comprising: a low pressure (LP) spool including a fan section, an LP compressor, and an LP turbine; a core engine including a high pressure (HP) compressor, a combustion section, and an HP turbine, wherein the HP compressor and the HP turbine together form a rotatable HP spool; a frame positioned between the HP turbine and the LP turbine in a serial flow arrangement, wherein the frame includes an inter-turbine burner having struts forming an outlet opening of a core flow path into the propulsion system; a first fuel system including a first fuel conduit in fluid communication with a fuel nozzle at the combustion section, wherein the first fuel system is configured to flow liquid fuel to the combustion section to produce a first combustion gas; a second fuel system including a second fuel conduit in fluid communication with the core flow path via the outlet opening at the inter-turbine burner, wherein the second fuel system is configured to flow gaseous fuel to the core flow path to produce a second combustion gas; wherein the LP compressor, the HP compressor, the combustion section, the HP turbine, the inter-turbine burner, and the LP turbine are in a serial flow arrangement; and wherein the propulsion system includes a rated power output ratio of the core engine to the inter-turbine burner having the LP spool between 1.5 and 5.7.

[0070] 2. The propulsion system according to any one or more of the items herein, wherein the core engine is configured to operate the HP spool at a maximum speed between 60% and 85% of the rated power output of the propulsion system.

[0071] 3. The propulsion system according to any one or more of the items herein, the propulsion system including: a computing system including a processor and a memory, wherein the memory is configured to store instructions that, when executed by the processor, cause the propulsion system to perform operations including: flowing liquid fuel to the combustion section and then producing a first combustion gas at the combustion section corresponding to 85% or less of the rated power output of the propulsion system.

[0072] 4. The propulsion system according to any one or more of the items herein, the operations including: modulating the rotational speed of the LP spool by modulating the gaseous fuel flow to the inter-turbine burner to produce the second combustion gas.

[0073] 5. The propulsion system according to any one or more of the items herein, the operations including: operating the HP spool at a steady-state speed while modulating the gaseous fuel flow to the inter-turbine burner.

[0074] 6. The propulsion system according to any one or more of the articles herein, the operation comprising: modulating a gaseous fuel flow through the inter-turbine burner to vary an output power of the propulsion system.

[0075] 7. The propulsion system according to any one or more of the articles herein, the operation comprising: maintaining a steady-state rotational speed of the HP spool while modulating the gaseous fuel flow.

[0076] 8. The propulsion system according to any one or more of the articles herein, the operation comprising: receiving a control signal corresponding to a high-power operating mode of the propulsion system; and then flowing gaseous fuel to the inter-turbine burner to produce a second combustion gas corresponding to a difference between the rated power output of the propulsion system and a power output produced by flowing liquid fuel to the combustion section.

[0077] 9. The propulsion system according to any one or more of the articles herein, the operation comprising: receiving a control signal corresponding to a low-power operating mode of the propulsion system; and then reducing the gaseous fuel flow to the inter-turbine burner, thereby reducing the power output of the propulsion system.

[0078] 10. The propulsion system according to any one or more of the articles herein, the operation comprising: operating the HP spool at a steady-state rotational speed while reducing the gaseous fuel flow to the inter-turbine burner.

[0079] 11. The propulsion system according to any one or more of the articles herein, wherein the fan section is configured as a non-ducted open rotor.

[0080] 12. The propulsion system according to any one or more of the articles herein, the propulsion system comprising: an electric machine operably coupled to the HP spool.

[0081] 13. A computing system for an aircraft propulsion system, the computing system including one or more processors and one or more memories, wherein the memories are configured to store instructions that, when executed by the processors, cause the propulsion system to perform operations including: flowing liquid fuel to a combustion section of the propulsion system; producing a first combustion gas in the combustion section corresponding to 85% or less of the rated power output of the propulsion system; and modulating a rotational speed of the LP spool by modulating a gaseous fuel flow to an inter-turbine burner to produce a second combustion gas.

[0082] 14. The computing system according to any one or more of the items herein, the operation comprising: operating a high-pressure (HP) spool at a steady-state speed while modulating the gaseous fuel flow to the inter-turbine burner.

[0083] 15. The computing system according to any one or more of the items herein, the operation comprising: receiving a control signal corresponding to a high-power operating mode of the propulsion system; and causing a gaseous fuel flow to the inter-turbine burner to generate the second combustion gas corresponding to the difference between the rated power output of the propulsion system and the power output generated by causing a liquid fuel flow to the combustion section.

[0084] 16. The computing system according to any one or more of the items herein, the operation comprising: receiving a control signal corresponding to a low-power operating mode of the propulsion system; and reducing the gaseous fuel flow to the inter-turbine burner, thereby reducing the power output of the propulsion system.

[0085] 17. The computing system according to any one or more of the items herein, the operation comprising: operating a high-pressure (HP) spool at a steady-state speed while reducing the gaseous fuel flow to the inter-turbine burner.

[0086] 18. The computing system according to any one or more of the items herein, the operation comprising: operating the core engine and the low-pressure (LP) spool having the inter-turbine burner at a rated power output ratio between the core engine and the inter-turbine burner having the LP spool of between 1.5 and 5.7.

[0087] 19. The computing system according to any one or more of the items herein, the operation comprising: operating a high-pressure (HP) spool at a maximum speed between 60% and 85% of the rated power output of the propulsion system.

[0088] 20. The computing system according to any one or more of the items herein, the operation comprising: causing a gaseous fuel flow to the inter-turbine burner to generate the rated power output of the propulsion system.

[0089] 21. The propulsion system according to any one or more of the items herein, comprising the computing system according to any one or more of the items herein.

[0090] 22. The computing system according to any one or more of the items herein, configured to operate the propulsion system according to any one or more of the items herein.

[0091] 23. An aircraft, the aircraft comprising the propulsion system according to any one or more of the items herein.

[0092] 24. An aircraft, the aircraft including the computing system described in any one or more of the items herein.

Claims

1. An aircraft propulsion system, characterized in that, The propulsion system includes: A low-pressure spool, the low-pressure spool including a fan section, a low-pressure compressor, and a low-pressure turbine; A core engine, the core engine including a high-pressure compressor, a combustion section, and a high-pressure turbine, wherein the high-pressure compressor and the high-pressure turbine together form a rotatable high-pressure spool; A frame, the frame being positioned between the high-pressure turbine and the low-pressure turbine in a serial flow arrangement, wherein the frame includes an inter-turbine burner, the inter-turbine burner including struts that form an outlet opening of a core flow path upstream of the low-pressure turbine of the propulsion system; A first fuel system, the first fuel system including a first pipe in fluid communication with a fuel nozzle at the combustion section, wherein the first fuel system is configured to flow liquid fuel to the combustion section to generate a first combustion gas; A second fuel system, the second fuel system including a second pipe in fluid communication with the core flow path via the outlet opening at the inter-turbine burner, wherein the second fuel system is configured to flow gaseous fuel to the core flow path upstream of the low-pressure turbine to generate a second combustion gas; Wherein the low-pressure compressor, the high-pressure compressor, the combustion section, the high-pressure turbine, the inter-turbine burner, and the low-pressure turbine are in a serial flow arrangement; and Wherein the propulsion system includes a rated power output ratio of the core engine to the inter-turbine burner having the low-pressure spool between 1.5 and 5.

7.

2. The propulsion system according to claim 1, wherein Wherein the core engine is configured to operate the high-pressure spool at a maximum rotational speed between 60% and 85% of the rated power output of the propulsion system.

3. The propulsion system according to claim 1, wherein The propulsion system includes: A computing system, the computing system including a processor and a memory, wherein the memory is configured to store instructions that, when executed by the processor, cause the propulsion system to perform operations, the operations including: Flowing liquid fuel to the combustion section and then generating a first combustion gas corresponding to 85% or less of the rated power output of the propulsion system at the combustion section.

4. The propulsion system according to claim 3, characterized in that, The operations include: Modulating the rotational speed of the low-pressure spool via the gaseous fuel flow modulated to the inter-turbine burner to generate the second combustion gas.

5. The propulsion system according to claim 4, wherein The operations include: Operating the high-pressure spool at a steady-state rotational speed while modulating the gaseous fuel flow to the inter-turbine burner.

6. The propulsion system according to claim 3, wherein The operations include: Modulating the gaseous fuel flow through the inter-turbine burner to change the power output of the propulsion system.

7. The propulsion system according to claim 6, characterized in that, The operations include: Maintaining a steady-state rotational speed of the high-pressure spool while modulating the gaseous fuel flow.

8. The propulsion system according to claim 3, wherein The operations include: Receiving a control signal corresponding to a high-power operation mode of the propulsion system; and then Flowing gaseous fuel to the inter-turbine burner to generate the second combustion gas corresponding to the difference between the rated power output of the propulsion system and the power output generated by flowing liquid fuel to the combustion section.

9. The propulsion system according to claim 3, characterized in that, The operations include: Receiving a control signal corresponding to a low-power operation mode of the propulsion system; and then Reduce the gaseous fuel flow to the inter-turbine burner, thereby reducing the power output of the propulsion system.

10. The propulsion system according to claim 9, characterized in that, The operation includes: Operate the high-pressure spool at a steady-state speed while reducing the gaseous fuel flow to the inter-turbine burner.

11. The propulsion system according to claim 1, wherein Wherein the fan section is configured as an unducted open rotor.

12. The propulsion system according to claim 1, characterized in that, The propulsion system includes: An electric motor operably coupled to the high-pressure spool.

13. A computing system for an aircraft propulsion system, characterized in that, The computing system includes one or more processors and one or more memories, wherein the memories are configured to store instructions that, when executed by the processors, cause the propulsion system to perform operations, the operations including: Direct a liquid fuel flow to a combustion section of the propulsion system; Produce a first combustion gas in the combustion section corresponding to 85% or less of the rated power output of the propulsion system; and Modulate the speed of the low-pressure spool by modulating the gaseous fuel flow to the inter-turbine burner to produce a second combustion gas, wherein the gaseous fuel flow exits from an outlet opening of the inter-turbine burner, and wherein the outlet opening is disposed upstream of the low-pressure spool.

14. The computing system according to claim 13, wherein The operation includes: Operate the high-pressure spool at a steady-state speed while modulating the gaseous fuel flow to the inter-turbine burner.

15. The computing system according to claim 13, wherein The operation includes: The operation includes: Receive a control signal corresponding to a high-power operation mode of the propulsion system; and Direct a gaseous fuel flow to the inter-turbine burner to produce the second combustion gas corresponding to the difference between the rated power output of the propulsion system and the power output produced by directing a liquid fuel flow to the combustion section.

16. The computing system according to claim 13, wherein The operation includes: Receive a control signal corresponding to a low-power operation mode of the propulsion system; and Reduce the gaseous fuel flow to the inter-turbine burner, thereby reducing the power output of the propulsion system.

17. The computing system according to claim 16, wherein The operation includes: Operate the high-pressure spool at a steady-state speed while reducing the gaseous fuel flow to the inter-turbine burner.

18. The computing system according to claim 13, wherein The operation includes: Operate the core engine and the low-pressure spool of the inter-turbine burner at a core engine to rated power output ratio of between 1.5 and 5.

7.

19. The computing system according to claim 18, wherein The operation includes: Operate the high-pressure spool at a maximum speed between 60% and 85% of the rated power output of the propulsion system.

20. The computing system according to claim 19, wherein The operation includes: Direct a gaseous fuel flow to the inter-turbine burner to produce the rated power output of the propulsion system.

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