Patent Description:
The overall pressure ratio (OPR) is a measure of the total pressure rise in a gas turbine engine (i.e., a pressure ratio equal to the air pressure discharged from the last compressor stage to the ambient air pressure entering the engine). Generally speaking, as OPR increases, the thermodynamic efficiency of the gas turbine engine increases, enabling the engine to consume less fuel per unit of thrust (i.e., thrust specific fuel consumption or TSFC) than a corresponding engine with lower OPR. However, air temperatures within the gas turbine engine increase with increasing OPR and can produce temperatures within the compressor section and/or turbine section that exceed permissible material and structural limits. Furthermore, the maximum temperature within the compressor and the turbine increase as the ambient temperature increases, adding to the temperature increase associated with the OPR of the engine.

Conventionally, turbine temperatures are maintained within acceptable limits by limiting OPR to a ratio that produces acceptable turbine temperatures for worst case ambient conditions, typically, design conditions corresponding to hot day take-off. While this technique produces a gas turbine engine design that provides an acceptable compromise for a variety of operating conditions, limiting OPR for hot day take-off conditions produces a gas turbine engine that operates at less OPR than otherwise possible at cruise power, reducing engine efficiency when high efficiency, low fuel consumption operation is most advantageous to extend aircraft range or payload capacity.

<CIT> relates to gas turbine engines capable of operating in a high overall pressure ratio (OPR) mode and in a low OPR mode to adapt to the ambient conditions and to provide more efficient operation without exceeding thermal limits of the gas turbine engine.

<CIT> relates to a system and method for diagnosing the health state of an auxiliary power unit power compressor.

<CIT> relates to methods and apparatus for operating gas turbine engines used for aircraft propulsion and auxiliary power.

A gas turbine engine in accordance with a first aspect of the invention is as claimed in claim <NUM>.

Some embodiments of the invention are as claimed in the dependent claims.

As described herein, a gas turbine engine has a boost spool that can be selectively operated to increase overall pressure ratio (OPR) during certain engine power levels (e.g., cruise power) while operating the gas turbine engine without the boost spool during other power levels (e.g., takeoff power). In some configurations, the boost spool includes a load compressor having an inlet shared with the boost compressor. With this configuration, a flow division between the boost compressor and the load compressor can be used to increase operating efficiency of the gas turbine engine. Furthermore, the gas turbine engine can operate within thermal limits when ambient temperature limits the OPR and can operate with greater engine efficiency when ambient temperatures are lower and permit higher OPR operation.

<FIG> is a schematic representation of gas turbine engine <NUM> that includes boost spool <NUM> in accordance with an exemplary embodiment of this disclosure. Gas turbine engine <NUM> is a turboprop engine that includes low pressure spool <NUM>, high pressure spool <NUM>, and power spool <NUM>. Low pressure spool <NUM> includes low pressure compressor <NUM> mechanically and rotationally connected to low pressure turbine <NUM> by shaft <NUM>, and high pressure spool <NUM> includes high pressure compressor <NUM> mechanically and rotationally connected to high pressure turbine <NUM> by shaft <NUM>. Power spool <NUM> includes power turbine <NUM> mechanically and rotationally connected to shaft <NUM>. Bearings <NUM> and <NUM> support shaft <NUM> of low pressure spool <NUM>, and bearings <NUM> and <NUM> support shaft <NUM> of high pressure spool <NUM>, each at one of a forward end, an aft end, or an intermediate location of respective shafts. Power shaft <NUM> is supported by bearings <NUM>, <NUM>, <NUM>, <NUM>, and <NUM> at the forward and aft ends as well as intermediate locations along shaft <NUM>. Low pressure spool <NUM>, high pressure spool <NUM>, and power spool <NUM> are coaxial, each extending along and rotating about centerline <NUM> independently of one another (i.e., without a mechanical connection that couples two or more of low pressure spool <NUM>, high pressure spool <NUM>, and power spool <NUM> in a rotational direction).

Compressors and turbines <NUM>, <NUM>, <NUM>, <NUM>, and <NUM> each include at least one compressor stage or turbine stage, each stage formed by a row of stationary vanes and a row of rotating blades. Blade rotors of compressors stages can be axial compressors, radial compressors, or mixed flow compressors. In the exemplary embodiment depicted by <FIG>, low pressure compressor <NUM> has three axial stages and low pressure turbine <NUM> has one axial stage. High pressure compressor <NUM> has a single radial compressor rotor, and high pressure turbine <NUM> has one axial stage. Power turbine <NUM> has two axial stages. However, in other embodiments, the number of stages in each compressor or turbine, as well as the radial, axial, or mixed configuration can be selected based on the desired pressure ratios as is known in the art.

At times, boost spool <NUM>, low pressure spool <NUM>, high pressure spool <NUM>, and power spool <NUM> may be referred to as a first spool, a second spool, a third spool, and/or a fourth spool in which "first", "second", "third", and "fourth" correspond to one of boost spool <NUM>, low pressure spool <NUM>, high pressure spool <NUM>, and power spool <NUM>. Similarly, "first", "second", "third", and/or "fourth" labels may be used in conjunction with corresponding components of the first spool, the second spool, the third spool, and/or the fourth spool in order to distinguish components of each spool from components of the other spools.

Between high pressure compressor <NUM> and high pressure turbine <NUM>, gas turbine engine <NUM> includes diffuser <NUM> and combustor <NUM> (i.e., primary combustor). Diffuser <NUM> is a radial diffuser positioned radially outward from high pressure compressor <NUM>, a radial compressor rotor. Diffuser <NUM> includes multiple ducts <NUM> (see <FIG>). Collectively, inlets of diffuser ducts <NUM> define an inlet to diffuser <NUM> in which individual duct inlets are distributed circumferentially about an outer periphery of high pressure compressor <NUM>. A first set of ducts 60A extend from high pressure compressor <NUM> to communicate with a location adjacent and upstream from combustor <NUM> while a second set of ducts 60B extends from high pressure compressor <NUM> to form part of inlet duct assembly <NUM> of boost spool <NUM>, which fluidly connects diffuser <NUM> to boost spool <NUM>. Combustor <NUM> includes casing <NUM> that forms an annular combustion chamber <NUM>. Within casing <NUM>, combustor <NUM> includes injectors <NUM> for introducing fuel. Combustor <NUM> discharges to high pressure turbine <NUM>.

Gas turbine engine <NUM> includes torque shaft <NUM> and reduction gearbox <NUM> for driving a propeller (not shown). An end of power shaft <NUM> includes internal spline 73A that mates with external spline 73B of torque shaft <NUM>, rotationally coupling shaft <NUM> to torque shaft <NUM>. Reduction gearbox <NUM> includes input shaft <NUM>, gearing <NUM>, and propeller shaft <NUM>. Torque shaft <NUM> mates with input shaft <NUM> via a flanged or splined connection to transfer rotation of power spool <NUM> to gearing <NUM>, which drives propeller shaft <NUM> in rotation about propeller axis <NUM>. Gearing <NUM> can be a series of spline gears, an epicyclic arrangement, or other gear train. The gear ratio of gearing <NUM> is less than one such that propeller shaft <NUM> rotates at a slower speed than shaft <NUM> of power spool <NUM>. Bearings <NUM> and <NUM> support input shaft <NUM>, and bearings <NUM>, <NUM>, and <NUM> support propeller shaft <NUM>, each bearing supporting input shaft <NUM> of propeller shaft <NUM> with respect to a casing of gas turbine engine <NUM>.

In some embodiments, gas turbine engine <NUM> can include electric machine <NUM>. Electric machine <NUM> can be a motor, a generator, or a motor-generator mounted to low pressure shaft <NUM>, high pressure shaft <NUM>, power shaft <NUM>, or propeller shaft <NUM>. As shown, <FIG> depicts electric machine <NUM> mounted between bearings <NUM> and <NUM>. A rotor of electric machine <NUM> can be mounted to shaft <NUM> of high pressure spool <NUM> while a stator of electric machine <NUM> can be mounted to shaft <NUM> of low pressure spool <NUM>.

Boost spool <NUM> includes variable inlet guide vane stage <NUM>, boost compressor <NUM>, secondary combustor <NUM>, and boost turbine <NUM>. Additionally, boost spool <NUM> includes variable inlet guide vane stage <NUM>, load compressor <NUM>, and electric machine <NUM>. Boost compressor <NUM>, boost turbine <NUM>, and load compressor <NUM> include at least one compressor stage or turbine stage, each stage formed by a row of stationary vanes and a row of rotating blades. Boost compressor <NUM>, boost turbine <NUM>, load compressor <NUM>, and electric machine <NUM> mount to boost shaft <NUM>, each component rotatable about boost centerline <NUM>. Bearings <NUM>, <NUM>, <NUM>, and <NUM> support boost shaft <NUM> with respect to a casing surrounding boost spool <NUM>. The casing can be a boost casing discrete from a casing surrounding low pressure spool <NUM> and high pressure spool <NUM> or boost spool <NUM> can be supported from a casing integrated with a casing or casings surrounding low pressure spool <NUM> and/or high pressure spool <NUM>.

Electric machine <NUM> can be a motor, a generator, or a motor-generator. Electric machine <NUM> can be mounted at any location along shaft <NUM>. As shown in <FIG>, electric machine <NUM> is adjacent to load compressor <NUM>.

Secondary combustor <NUM> is disposed between and communicates with boost compressor <NUM> and boost turbine <NUM>. Like primary combustor <NUM>, secondary combustor <NUM> includes casing <NUM> that forms annular combustion chamber <NUM>. Within annular combustion chamber <NUM>, secondary combustor <NUM> includes injectors <NUM> for introducing fuel.

Each of variable inlet vane stage <NUM> and variable inlet vane stage <NUM> forms an array of circumferentially spaced vanes in communication with boost inlet plenum <NUM> upstream of boost compressor <NUM> and load compressor <NUM>. Vanes of variable inlet guide vane stage <NUM> and vanes of variable inlet guide vane stage <NUM> are rotatable about respective spanwise vane axes. An angular position of variable inlet guide vane stage <NUM> and variable inlet guide vane stage <NUM> ranges between a closed position, a neutral or nominal position, and an open position to vary an open area into boost compressor <NUM> and load compressor <NUM>, respectively.

Boost spool <NUM> receives compressed air from diffuser <NUM> through inlet duct assembly <NUM>, which fluidly connects diffuser <NUM> to inlet duct plenum <NUM>. Outlet duct assembly <NUM> fluidly connects boost turbine <NUM> and load compressor <NUM> to diffuser <NUM> and/or combustor <NUM>. Each of inlet duct assembly <NUM> and outlet duct assembly <NUM> includes one or more ducts, pipes, conduits, and/or manifolds to deliver a portion of gas turbine engine <NUM> flow to or from boost spool <NUM>.

The position and orientation of boost spool <NUM> relative to centerline <NUM> is selected based on the mechanical and/or electro-mechanical coupling. Boost centerline <NUM> can be parallel and offset from centerline <NUM> of gas turbine engine <NUM> as schematically shown by <FIG>. Furthermore, <FIG> shows boost spool <NUM> with a reverse flow orientation (i.e., aft-to-forward flow) such that a flow direction through boost spool <NUM> from compressor <NUM> to turbine <NUM> is opposite a flow direction (i.e., forward-to-aft flow) through gas turbine engine <NUM> from inlet <NUM> to an outlet of gas turbine engine <NUM> downstream from power turbine exit <NUM>. Alternatively, boost centerline <NUM> can be oblique, perpendicular, or eccentric to centerline <NUM>. In some embodiments, boost spool <NUM> is located remotely from low pressure spool <NUM>, high pressure spool <NUM>, and power spool <NUM> while in other embodiments, boost spool <NUM> can be located directly adjacent to or attached to casings surrounding spools <NUM>, <NUM>, and <NUM>. In either of these arrangements, boost spool <NUM> can be an auxiliary power unit (APU).

For all mounting positions of boost spool <NUM>, the location and orientation of boost spool <NUM> permits boost spool <NUM> to receive a compressed air flow from gas turbine engine <NUM> and to discharge an expanded air flow to gas turbine engine <NUM>. Boost spool <NUM> can receive a compressed airflow from any compressor stage of gas turbine engine <NUM> to achieve varying degrees of boost compression. In one exemplary embodiment, boost spool <NUM> receives a compressed air flow from a location that is downstream from the last compressor stage of the gas turbine engine. In the case of gas turbine engine <NUM>, boost spool <NUM> receives airflow from diffuser <NUM> and discharges an expanded airflow to diffuser <NUM>. In other instance, boost spool <NUM> receives airflow from diffuser <NUM> and discharges an expanded airflow to both diffuser <NUM> and combustor <NUM>, which is downstream of high pressure compressor <NUM> and upstream from high pressure turbine <NUM>.

<FIG> is a schematic depicting an exemplary implementation of inlet duct assembly <NUM> and outlet duct assembly <NUM>. Section A-A depicts inlet duct assembly <NUM> taken along line A-A in <FIG>. Section B-B depicts outlet duct assembly <NUM> taken along line B-B in <FIG>. As shown in Section A-A, first ducts 60A and second ducts 60B of diffuser <NUM> are circumferentially distributed about high pressure compressor <NUM>. Inlets 132A of first diffuser ducts 60A and inlets 132B of second diffuser ducts 60B are circumferentially arranged about a radially outer periphery of high pressure compressor <NUM>. Each of inlets 132A are interposed with inlets 132B such that each inlet 132A is disposed between two circumferentially adjacent inlets 132B and each inlet 132B is disposed between two circumferentially adjacent inlets 132A. Each of first diffuser ducts 60A (i.e., primary combustor ducts) extend from high pressure compressor <NUM> towards primary combustor <NUM>. Each of second diffuser ducts 60B (i.e., boost ducts) extend from high pressure compressor <NUM> to connect with inlet duct assembly <NUM>.

Inlet duct assembly <NUM> can include collection manifold <NUM>, which may communicate directly with second diffuser ducts 60B. In other embodiments, inlet duct assembly <NUM> can include inlet branch ducts <NUM> that fluidly connect and extend second diffuser ducts 60B to manifold <NUM>. Collection manifold <NUM> can be a plenum with a cross-sectional area that is fixed along its circumferential length or increases with each or as a function of second diffuser duct 60B or inlet branch ducts <NUM>. In other embodiments, collection manifold <NUM> can be a pipe or duct that has a progressively increasing cross-sectional area along its circumferential length as each second diffuser duct 60B or inlet branch duct <NUM> connects to collection manifold <NUM>. For instance, collection manifold <NUM> can be a series of pipe segments in which each branch duct <NUM> or second diffuser duct 60B joins one of the pipe segments with a y-section or a t-section. Extending from collection manifold <NUM>, some embodiments of inlet duct assembly <NUM> include main inlet duct <NUM> that fluidly connects collection manifold <NUM> to inlet duct plenum <NUM> of boost spool <NUM>. Disposed between variable inlet guide vane stage <NUM> associated with boost compressor <NUM> and variable inlet guide vane stage <NUM> associated with load compressor <NUM>, inlet duct plenum <NUM> extends circumferentially about boost shaft <NUM> to distribute air extracted from diffuser <NUM> evenly along annular inlets to variable inlet guide vane stage <NUM> and variable inlet guide vane stage <NUM>.

Outlet duct assembly <NUM> extends from an outlet of boost spool <NUM> downstream from boost turbine <NUM> to join first diffuser ducts 60A upstream of primary combustor <NUM>. As shown in <FIG>, outlet duct assembly <NUM> can include turbine outlet duct <NUM> extending from the outlet of boost turbine <NUM> to outlet manifold <NUM>, which extends circumferentially about diffuser <NUM>. Outlet manifold <NUM> communicates with boost discharge ducts <NUM>. Boost discharge ducts <NUM> are distributed circumferentially about centerline <NUM> radially outward from first diffuser ducts 60A.

Additionally, outlet duct assembly <NUM> can include load compressor duct <NUM> extending from an outlet of load compressor <NUM> to outlet manifold <NUM>, which extends circumferentially about diffuser <NUM>. Outlet manifold <NUM> communicates with load compressor discharge ducts <NUM>. Discharge ducts <NUM> are distributed circumferentially about centerline <NUM> and are interposed between boost discharge ducts <NUM> radially outward of first diffuser ducts 60A.

Alternatively, outlet duct assembly <NUM> can combine flow from discharged from load compressor <NUM> and boost turbine <NUM> prior to introduction into primary combustor <NUM> as shown in <FIG>. Accordingly, in place of outlet manifolds <NUM> and <NUM>, outlet duct assembly <NUM> includes combined discharge manifold <NUM>. Additionally, outlet duct assembly <NUM> includes ejector <NUM> and mixed flow duct <NUM>. Turbine outlet duct <NUM> extends from an outlet of boost turbine <NUM> to inlet <NUM> of ejector <NUM>, and load compressor duct <NUM> extends from an outlet of load compressor <NUM> to working fluid inlet <NUM> of ejector <NUM>. Working fluid inlet <NUM> is centrally disposed within ejector <NUM>. Accordingly, given proper pressure conditions at inlet <NUM> and working fluid inlet <NUM>, air entering ejector <NUM> from load compressor <NUM> entrains air through inlet <NUM>. Mixed flow duct <NUM> extends from outlet <NUM> of ejector <NUM> to combined discharge manifold <NUM>. Combined boost discharge ducts <NUM> extend from discharge manifold <NUM> towards primary combustor <NUM> along with first diffuser ducts 60A. Combined boost discharge ducts <NUM> are circumferentially distributed about centerline <NUM> and diffuser <NUM> radially outward from first diffuser ducts 60A.

<FIG> is a schematic of controller <NUM> that regulates the operation of gas turbine engine <NUM> and, more particularly, a flow division between boost compressor <NUM> and load compressor <NUM> of boost spool <NUM>. Additionally, controller <NUM> regulates fuel flow rates to primary combustor <NUM> and secondary combustor <NUM> based on one or more engine parameters, aircraft parameters, and/or exterior conditions. Controller <NUM> can include a standalone control unit or a control module incorporated into another control unit. Furthermore, controller <NUM> can be an amalgamation of distinct control units and/or distinct control modules that together perform the functions described in this disclosure. In some embodiments, controller <NUM> can be a full authority digital engine control (FADEC), an electric engine controller (EEC), or an engine control unit (ECU).

Controller <NUM> includes processor <NUM>, memory <NUM>, and input/output interface <NUM>. Processor <NUM> executes one or more control algorithms <NUM> stored within memory <NUM> to output engine control signals <NUM> based on one or more input signals <NUM>. Examples of processor <NUM> can include any one or more of a microprocessor, a controller, a digital signal processor (DSP), an application specific integrated circuit (ASIC), a field-programmable gate array (FPGA), or other equivalent discrete or integrated logic circuitry.

Memory <NUM> can be configured to store information within controller <NUM>. Memory <NUM>, in some examples, is described as computer-readable storage media. In some examples, a computer-readable storage medium can include a non-transitory medium. The term "non-transitory" can indicate that the storage medium is not embodied in a carrier wave or a propagated signal. In certain examples, a non-transitory storage medium can store data that can, over time, change (e.g., in RAM or cache). Memory <NUM> can include volatile and non-volatile computer-readable memories. Examples of volatile memories can include random access memories (RAM), dynamic random-access memories (DRAM), static random-access memories (SRAM), and other forms of volatile memories. Examples of non-volatile memories can include, e.g., magnetic hard discs, optical discs, flash memories, or forms of electrically programmable memories (EPROM) or electrically erasable and programmable (EEPROM) memories.

Input/output interface or I/O interface <NUM> can be a series of input and output channels that electrically communicate with an engine control bus. The engine control bus interconnects controller <NUM> with various components of the gas turbine engine <NUM> described above such that engine control signals <NUM> can be transmitted to individual engine components and input signals <NUM> can be received.

Engine control signals <NUM>, input signals <NUM>, or both engine control signals <NUM> and input signals <NUM> can be an analog signal or a digital signal. For example, an analog signal can be a voltage that varies between a low voltage to a high voltage whereas digital signals can be a series of discrete voltage states distributed over a voltage range. Operatively, engine control signals <NUM> cause various components of gas turbine engine <NUM> to change state or position. For example, engine control signals <NUM> can be used to vary the position of one or more fuel valves to vary the fuel rate entering a combustor. Other examples of engine control signals <NUM> include signals associated with an angular position of variable vane stages. Input signals <NUM> are representative of one of engine parameters, aircraft parameters, and environmental parameters. Exemplary engine parameters include rotational speed of a low pressure spool, high pressure spool, boost spool, and/or fan shaft, the state or position of fuel valves, bleed valves, the state or position of clutch assemblies, propeller blade pitch angle, the temperature or pressure within the compressor, combustor, or turbine, and engine power. Aircraft parameters include various parameters associated with an aircraft such as power lever angle, altitude, pitch angle, yaw angle, roll angle, rate of climb, and airspeed, among other possible parameters. Exterior parameters include ambient temperature and pressure at the inlet of gas turbine engine <NUM>.

In operation, gas turbine engine <NUM> receives ambient air flow <NUM> thorough inlet <NUM>, which communicates with low pressure compressor <NUM>. Rotation of low pressure compressor <NUM> and high pressure compressor <NUM> compresses air flow <NUM>. High pressure compressor <NUM> discharges compressed air flow <NUM> into diffuser <NUM>. Within diffuser <NUM>, compressed air flow <NUM> divides into primary flow <NUM> and secondary flow <NUM>. First diffuser ducts 60A deliver primary flow <NUM> from high pressure compressor <NUM> to combustor <NUM> while the second set of diffuser ducts 60B delivers secondary flow <NUM> to inlet plenum <NUM> of boost spool <NUM> via inlet duct assembly <NUM>. Each of diffuser ducts 60A includes divergent walls that reduce primary flow <NUM> velocity and thereby increase static pressure of flow <NUM> before entering combustor <NUM>. Fuel injected into combustor <NUM> mixes with primary flow <NUM>, and one or more ignitors combust the fuel-to-air mixture to produce a compressed and heated primary flow <NUM> that is discharged into high pressure turbine <NUM>.

Based on the angular position of variable inlet guide vane stage <NUM> and the angular position of variable inlet guide vane stage <NUM>, secondary flow <NUM> is further divided into boost flow <NUM> and load compressor flow <NUM>. Boost flow <NUM> passes through variable inlet guide vane stage <NUM> and into boost compressor <NUM>. As boost compressor <NUM> rotates, boost flow <NUM> is compressed and flows into secondary combustor <NUM>. Within secondary combustor <NUM>, injectors <NUM> introduce fuel and the fuel-to-air mixture is ignited, further heating and increasing pressure of boost flow <NUM>. Boost flow <NUM> discharges from secondary combustor <NUM> into boost turbine <NUM> and imparts work to boost turbine <NUM> to thereby rotate shaft <NUM>, boost compressor <NUM>, load compressor <NUM>, and electric machine <NUM>. Boost flow <NUM> discharges from boost turbine <NUM> into outlet duct assembly <NUM>. Load compressor flow <NUM> passes through variable inlet guide vane stage <NUM> and into load compressor <NUM>. Work imparted to load compressor flow <NUM> by load compressor <NUM> compresses load compressor flow <NUM>. Load compressor flow <NUM> discharges from load compressor <NUM> into outlet duct assembly <NUM>. Outlet duct assembly <NUM> returns boost flow <NUM> and load compressor flow <NUM> to a location upstream from primary combustor <NUM> proximate outlets of first diffuser ducts 60A. Boost flow <NUM> and load compressor flow <NUM> can remain separate until mixing with primary flow <NUM> discharged from first diffuser ducts 60A. Alternatively, boost flow <NUM> and load compressor flow <NUM> can mix prior to combining with primary flow <NUM> of first diffuser ducts 60A.

After mixing, primary flow <NUM>, boost flow <NUM>, and load compressor flow <NUM> enter primary combustor <NUM>. Within primary combustor <NUM>, injectors <NUM> introduce fuel to produce a fuel-to-air mixture and the fuel-to-air mixture ignites. Combined flow <NUM> discharges from combustor <NUM> and enters, in sequential order, high pressure turbine <NUM>, low pressure turbine <NUM>, and power turbine <NUM>. Combined flow <NUM> interacts with vanes and blades of high pressure turbine <NUM> causing rotation of shaft <NUM> about centerline <NUM> and driving rotation of high pressure compressor <NUM>. Similarly, combined flow <NUM> interacting with vanes and blades of low pressure turbine <NUM> cause rotation of shaft <NUM> about centerline <NUM> to drive rotation of low pressure compressor <NUM>. Additionally, combined flow <NUM> interacting with vanes and blades of power turbine <NUM> cause rotation of shaft <NUM> about centerline <NUM> to drive rotation of propeller shaft <NUM> via torque shaft <NUM>, input shaft <NUM>, and gearing <NUM>. Downstream of power turbine <NUM>, combined flow <NUM> discharges from engine <NUM> as exhaust at a location downstream from outlet <NUM>.

The flow division between boost flow <NUM> and load compressor flow <NUM> can be managed by controller <NUM> to achieve increased efficiency of gas turbine engine <NUM> during operation of boost spool <NUM>. For example, controller <NUM> can vary the angular position of variable inlet guide vane stage <NUM> based on the engine pressure ratio of boost spool. The engine pressure ratio of boost spool (EPRB) is determined as the pressure of secondary flow <NUM> within inlet duct assembly <NUM> adjacent to diffuser <NUM> (i.e., pressure P3-<NUM>) divided by pressure of boost flow <NUM> discharged from boost turbine and within outlet duct assembly <NUM>. The pressure of boost flow <NUM> is measured at a location upstream from the discharge end of outlet duct assembly <NUM> (i.e., pressure P5-<NUM>), whether outlet duct assembly <NUM> includes ejector <NUM> or not. The target engine pressure ratio of boost spool <NUM> is selected based on a desired overall pressure ratio of gas turbine engine <NUM> for a given operating condition or power level (i.e., cruise power). For instance, the target engine pressure ratio of boost spool can be from <NUM> to <NUM>.

At the same time, controller <NUM> can vary the angular position of variable inlet guide vane stage <NUM> based on a pressure of load compressor flow <NUM> at a discharge end of outlet duct assembly <NUM> (i.e., pressure P5-<NUM>) relative to a pressure of boost flow <NUM> at a discharge end of outlet duct assembly <NUM> (i.e., pressure P5-<NUM>). For instance, a target discharge pressure for load compressor flow <NUM> can be at least <NUM>% greater than the discharge pressure of boost flow <NUM>. At an upper end, the discharge pressure of load compressor flow can be determined based on pressure loss incurred within inlet duct assembly <NUM> and outlet duct assembly <NUM> as well as any associated pressure loss associated with separating or mixing boost flow <NUM>, load compressor flow <NUM>, primary flow <NUM>, and/or secondary flow <NUM>. Accordingly, load compressor flow <NUM> can be used as a working fluid, entraining boost flow <NUM> into combustor <NUM> as facilitated by ejector <NUM> shown in <FIG> or duct arrangements shown in section A-A and section B-B of <FIG>.

<FIG> is an exemplary enthalpy graph showing relative enthalpy gains and losses associated with operation of gas turbine engine <NUM> and boost spool <NUM>. P1 represents the enthalpy increase provided by propeller (not shown) as driven by propeller shaft <NUM>. Flow faction i represents the portion of flow propellered through the propeller that enters gas turbine engine <NUM> through inlet <NUM>. The reminder of propeller flow (i.e., a flow fraction equal to <NUM>-i) propels gas turbine engine <NUM> and the aircraft in flight. The flow division between core flow or ambient flow i and bypass flow <NUM>-i can be expressed as bypass flow ratio BPR, which is the mass flow rate of bypass flow <NUM>-i divided by the mass flow rate of core flow i. Expressed in terms of bypass flow ratio BFR, core flow i is equal to one divided by the sum of one plus the bypass flow ratio, i.e., <NUM> / (<NUM>+BFR). Bypass flow is equal to the bypass flow ratio divided by the sum of one plus the bypass flow ratio, i.e., BFR / (<NUM>+BFR).

LPC1 and HPC1 represent an increase of enthalpy between inlet <NUM> and an outlet of high pressure compressor <NUM>. At an outlet of high pressure compressor <NUM>, compressed air flow <NUM> divides into secondary flow <NUM> represented by flow fraction (k) and primary flow <NUM> represented by flow fraction [<NUM>-k]. Second diffuser ducts 60B direct secondary flow <NUM> (i.e., flow fraction k) through inlet duct assembly <NUM> to inlet duct plenum <NUM>. A first enthalpy loss (dH1) represents an enthalpy decrease associated with egress from diffuser <NUM> and flow through inlet duct assembly <NUM> associated with secondary flow <NUM> (i.e., flow fraction k). First diffuser ducts 60A direct primary flow <NUM> (i.e., flow fraction <NUM>-k) to a location upstream from primary combustor <NUM>.

The relative positions of variable inlet guide vane stage <NUM> and variable inlet guide duct vane stage <NUM> adjacent to inlet duct plenum <NUM>, partitions secondary flow <NUM> (i.e., flow fraction k) as further divided into boost flow <NUM> represented by flow fraction (k x j) and load compressor flow <NUM> represented by flow fraction (k x (<NUM>-j)). Variable inlet guide vane stage <NUM> directs boost flow <NUM> into boost compressor <NUM>. Boost compressor <NUM> compresses boost flow <NUM> as indicated by enthalpy increase HPC2. Compressed boost flow <NUM> enters secondary combustor <NUM> where it is mixed with fuel and the fuel-air mixture ignited as indicated by enthalpy increase C2. Heated and compressed boost flow <NUM> discharges from secondary combustor <NUM> into boost turbine <NUM>. Boost flow <NUM> expands across boost turbine <NUM> as indicated by the enthalpy decrease HPT2. Work extracted by boost turbine <NUM> drives boost compressor <NUM> and load compressor <NUM>. Additionally, where electric machine <NUM> operates as a generator, boost turbine <NUM> drives electric machine <NUM>. Variable inlet vane stage <NUM> directs load compressor flow <NUM> into load compressor <NUM>. Load compressor <NUM> compresses load compressor flow <NUM> as indicated by enthalpy increase HPC3. In some embodiments, boost flow <NUM> and load compressor flow <NUM> mix prior to reinduction at a location upstream from primary combustor <NUM>. Enthalpy decrease dH2 represents the enthalpy change produced by mixing boost flow <NUM> and load compressor flow <NUM>. Enthalpy decrease dH3 represents the enthalpy change associated with the ingress of secondary flow <NUM> (i.e., mixed boost flow <NUM> and load compressor flow <NUM>) at a location upstream from primary combustor <NUM> and proximate a discharge location of first diffuser ducts 60A.

Load compressor <NUM> can be operated such that enthalpy increase HPC3 equates to the enthalpy losses dH1, dH2, and dH3. Effectively, secondary flow <NUM> discharged from boost spool <NUM> can rejoin primary flow <NUM> with an enthalpy equal to or greater than the enthalpy of secondary flow <NUM> discharged from high pressure compressor <NUM> as shown in <FIG>.

Primary flow <NUM> and secondary flow <NUM> mix upstream from primary combustor <NUM>. Because the pressure P5-<NUM> of boost spool <NUM> is greater than the pressure of diffuser <NUM> (i.e., pressure P3-<NUM>), mixing primary flow <NUM> and secondary flow <NUM> increases enthalpy of the mixed flow as indicated by dH4. Within primary combustor <NUM>, injected fuel mixes with the flow and ignites producing the associated enthalpy increase C1.

Combined flow <NUM> discharges from primary combustor <NUM> into high pressure turbine <NUM> and expands across high pressure turbine <NUM>. HPT1 represents the enthalpy decrease associated with high pressure turbine <NUM>. Work extracted from high pressure turbine <NUM> drives high pressure compressor <NUM>. Downstream from high pressure turbine <NUM>, combined flow <NUM> expands across low pressure turbine <NUM> as represented by enthalpy decrease LPT1. Work extracted from low pressure turbine <NUM> drives low pressure compressor <NUM>.

In some embodiments, low pressure turbine <NUM>, high pressure turbine <NUM>, or both low pressure turbine <NUM> and high pressure turbine <NUM> drive one or more engine accessories and/or electric machine <NUM>. Accordingly, the magnitude of LPT1 and/or HPT1 increases in proportion to the work required to drive these additional components. In other embodiments, electrical output from electric machine <NUM> of boost spool <NUM> can be used to drive electric machine <NUM> as a motor as indicated by dH5 in <FIG>.

Combined flow <NUM> discharged from low pressure turbine <NUM> further expands across power turbine <NUM>. The associated enthalpy decrease PT represents work extracted from combined flow <NUM> that drives a propulsor of an aircraft and the associated drive train components. Where the gas turbine engine is a turboprop engine, such as gas turbine engine <NUM> shown in <FIG>, power turbine <NUM> drives torque shaft <NUM>, input shaft <NUM>, gearing <NUM>, propeller shaft <NUM>, and propeller (not shown). However, in other embodiments, the gas turbine engine can be a turboshaft engine or a bypass engine. In these instances, the gas turbine engine does not include power turbine <NUM> and, instead, the magnitude of LPT1 associated with low pressure turbine <NUM> increases to drive the propulsor (e.g., a fan or main rotor) and associated drive train components (e.g., losses associated with shafts, gearboxes, and bearings). Downstream from power turbine <NUM>, combined flow <NUM> discharges from gas turbine engine <NUM> through exhaust nozzle as represented by enthalpy decrease dH6. As shown in <FIG>, gas turbine engine <NUM> produces imparts a net enthalpy change, dHnet, to the ambient air to propel the aircraft in flight.

Operating gas turbine engine <NUM> with boost spool <NUM> in which boost spool <NUM> includes load compressor <NUM> reduces an average temperature of secondary flow <NUM> reintroduced with primary flow <NUM> at a location upstream from primary combustor <NUM>. Lower temperatures of secondary flow <NUM> relative to operation of boost spool <NUM> in which the entirety of secondary flow <NUM> flows through secondary combustor <NUM> increases operational life of primary combustor <NUM> components. Where mixing of boost flow <NUM> and load compressor flow <NUM> occurs upstream of primary combustor <NUM> with primary flow <NUM>, circumferentially distributing boost flow <NUM> discharged from ducts <NUM> about centerline <NUM> and between adjacent ducts <NUM> that discharge load compressor flow <NUM> contributes to temperature uniformity of the combined flow entering primary combustor <NUM>. Similar benefits are provided by embodiments utilizing ejector <NUM> to mix boost flow <NUM> and load compressor flow <NUM> within outlet duct assembly <NUM> upstream from primary combustor <NUM>. Additionally, ejector <NUM> drives mixing of secondary flow <NUM> with primary flow <NUM>.

Operating boost spool <NUM> with load compressor <NUM> enables primary combustor <NUM> to operate at uncharacteristically low fuel-air-ratios (FAR) relative to conventional gas turbine engines (e.g., without boost spool <NUM> and load compressor <NUM>) at intermediate power levels (e.g., cruise power). Furthermore, operating boost spool <NUM> with load compressor <NUM> enables primary combustor <NUM> to operate at uncharacteristically high fuel-air-ratios (FAR) relative to convention gas turbine engines at high power levels (e.g., takeoff and climb power).

Exemplary fuel-air-ratios for gas turbine engine <NUM> operating with boost spool <NUM> are presented in Table <NUM>, Table <NUM>, and Table <NUM> below. Operation of gas turbine engine <NUM> with boost spool <NUM> and load compressor <NUM> can be characterized by engine operating condition (i.e., power level), a flow split between primary flow <NUM> and secondary flow <NUM> (i.e., k, <NUM>-k), and a flow split of secondary flow <NUM> between boost flow <NUM> and load compressor flow <NUM> (i.e., j, <NUM>-j). For instance, gas turbine engine <NUM> can include, in order of increasing power level, flight idle, cruise, climb, and takeoff operating conditions, among other possible intermediate power levels. The flow split between primary flow <NUM> and secondary flow <NUM> is indicated by flow fraction k (i.e., the fraction of secondary flow taken from ambient or core flow <NUM>). The flow split between boost flow <NUM> and load compressor flow <NUM> is indicated by flow fraction j (i.e., the fraction of boost flow k taken from secondary flow <NUM>).

The fuel-air-ratio, "FAR Boost Spool", corresponds to an exit or outlet of secondary combustor <NUM> and varies between a value necessary for minimum safe operation (. e.g., FAR approximately equal to <NUM>) and a maximum fuel-air-ratio permitted by materials and/or cooling of casing <NUM> of secondary combustor <NUM> (e.g., FAR equal to or greater than <NUM>). The fuel-air-ratio, "FAR Ingress", coincides with a location upstream from primary combustor <NUM> where boost flow <NUM> and load compressor flow <NUM> rejoin primary flow <NUM>. For example, "FAR Ingress" may correspond to a location between outlets of boost outlet ducts <NUM> or load compressor outlet ducts <NUM> and primary combustor <NUM> (see <FIG>) or between outlets of discharge ducts <NUM> and primary combustor <NUM> (see <FIG>). The fuel-air-ratio, "FAR Primary Combustor Exit", is determined at the exit of primary combustor <NUM> and is based on total fuel flow added by primary combustor <NUM> and secondary combustor <NUM> divided by total flow through gas turbine engine <NUM> (e.g., ambient flow <NUM> or combined flow <NUM>). The fuel-air-ratio, "FAR Primary Combustor Exit Excluding Secondary Combustor Fuel Flow", is determined at the exit of primary combustor <NUM> and is based on fuel flow added by primary combustor <NUM> only divided by total flow through gas turbine engine <NUM> (e.g., ambient flow <NUM> or combined flow <NUM>).

As shown by Table <NUM>, Table <NUM>, and Table <NUM>, the fuel-air-ratios at the exit of primary combustor <NUM> (including fuel flow added by secondary combustor <NUM>) is greater than <NUM> for all operational conditions of gas turbine engine <NUM>. At high power operation associated with climb and takeoff power (Table <NUM>), the fuel-air-ratio at the exit of primary combustor <NUM> is greater than <NUM> and is equal to approximately <NUM> in the depicted case. In a gas turbine engine without boost spool <NUM> and load compressor <NUM>, such fuel-air-ratios at high power levels (e.g., takeoff and climb power) cause excessive temperatures within high pressure turbine <NUM>. Accordingly, gas turbine engine <NUM> operating with boost spool <NUM> and load compressor <NUM> can achieve higher fuel-air-ratios at high power levels (e.g., takeoff and climb power) relative to conventional gas turbine engines. Furthermore, gas turbine engine <NUM> operating with boost spool <NUM> and load compressor <NUM> can achieve lower fuel-air-ratios relative to conventional gas turbine engines at cruise power levels. As shown by Table <NUM>, the fuel-air-ratio (FAR) at the exit of primary combustor <NUM> can be less than <NUM> at cruise power (see "FAR Primary Combustor Exit"). The fuel-air-ratio (FAR) at the exit of primary combustor <NUM> excluding secondary combustor fuel flow is between <NUM> and <NUM> at low engine thrust (Table <NUM>) and high engine thrust (Table <NUM>). However, the fuel-air-ratio at the exit of primary combustor <NUM> excluding secondary combustor fuel flow is reduced at cruise power, ranging from <NUM> to <NUM> (Table <NUM>) and contributing to improved efficiency of gas turbine engine.

Claim 1:
A gas turbine engine (<NUM>) comprising:
a first spool (<NUM>) comprising a first compressor (<NUM>) and a first turbine (<NUM>) mounted to a first shaft (<NUM>);
a primary combustor (<NUM>) disposed between the first compressor (<NUM>) and the first turbine (<NUM>);
a diffuser (<NUM>) disposed between the primary combustor (<NUM>) and the first compressor (<NUM>), wherein a primary flow path (<NUM>) of the gas turbine engine (<NUM>) includes, in flow series, the first compressor (<NUM>), the diffuser (<NUM>), the primary combustor (<NUM>), and the first turbine (<NUM>);
a second spool (<NUM>) comprising a second compressor (<NUM>), a second turbine (<NUM>), and a load compressor (<NUM>) mounted to a second shaft (<NUM>);
an inlet plenum (<NUM>) fluidly connected to the second compressor (<NUM>) and the load compressor (<NUM>);
a secondary combustor (<NUM>) disposed between the second compressor (<NUM>) and the second turbine (<NUM>), wherein a secondary flow path (<NUM>) of the gas turbine engine (<NUM>) includes, in flow series, the inlet plenum (<NUM>), the second compressor (<NUM>), the secondary combustor (<NUM>), and the second turbine (<NUM>), and a tertiary flow path (<NUM>) of the gas turbine engine (<NUM>) includes, in flow series, the inlet plenum (<NUM>) and the load compressor (<NUM>);
an inlet duct assembly (<NUM>) extending between and fluidly connecting the diffuser (<NUM>) to the inlet plenum (<NUM>); and
an outlet duct assembly (<NUM>) extending between and fluidly connecting the second turbine (<NUM>) and the load compressor (<NUM>) to the diffuser (<NUM>).