Gas turbine engine with combustion section

Abstract

A gas turbine engine includes a steam generation system and a combustion section having a first combustor and a second combustor in fluid communication with a secondary combustion zone of the first combustor. The second combustor is operable to receive compressor bleed air and fuel, and steam from the steam generation system. During the non-idle operating state and the power boost operating state, when the steam system is operable, steam is provided to the second combustor and then to the secondary combustion zone of the first combustor. When the steam generation system is inoperable, fuel and bleed gas are provided to the second combustor and combusted, and combustion products are provided to the first combustor. When the steam generation system portion is operable, steam, bleed gas, and fuel are provided to the second combustor and combustion products are provided to the first combustor.

Description

Technical Field

[0001] The present disclosure relates to a gas turbine engine having a combustion section. Background Art

[0002] Gas turbine engines typically include a combustor with a swirler that provides a swirling air flow mixed with fuel into a combustion chamber where the fuel and air mixture is ignited and burned. The combustion of the fuel and air mixture in the combustion chamber causes the combustor to emit carbon monoxide (CO) and nitrogen oxides (NO x ). An attempt to reduce CO and NO x A technique for exhaust is to inject steam or water directly into the swirler, for example via the fuel nozzle, to mix with the fuel and air mixture. BRIEF DESCRIPTION OF THE DRAWINGS

[0003] Features and advantages of the present disclosure will be apparent from the following description of various exemplary embodiments as illustrated in the drawings, wherein like reference numbers generally indicate identical, functionally similar, and / or structurally similar elements.

[0004] Figure 1 is a schematic partial cross-sectional side view of an exemplary high bypass turbofan jet engine and steam generation system according to one aspect of the present disclosure.

[0005] Figure 2 According to one aspect of the present disclosure, Figure 1 Schematic diagram of a high-bypass turbofan jet engine and steam generation system.

[0006] Figure 3 According to one aspect of the present disclosure, Figure 1 A partial cross-sectional side view of an exemplary combustion section of a gas turbine engine.

[0007] Figure 4 According to one aspect of the present disclosure, Figure 3 An enlarged detail view of the second burner is shown taken at detail view 204 of FIG.

[0008] Figure 5 According to one aspect of the present disclosure, Figure 3 A partial cross-sectional view of a semi-annular second burner having a plurality of second combustion chambers is provided in FIG. Figure 3 The image is taken at plane 5-5.

[0009] Figure 6 According to another aspect of the present disclosure, Figure 4 A partial cross-sectional side view of an optional second burner of the burner is shown.

[0010] Figure 7 According to one aspect of the present disclosure, Figure 6 A partial cross-sectional view of an optional second combustor liner taken at plane 7-7.

[0011] Figure 8 According to one aspect of the present disclosure, Figure 1 Schematic diagram of the operation of a gas turbine engine in a non-power boost operating state.

[0012] Figure 9 According to one aspect of the present disclosure, when a steam generating system operates to generate steam at a first steam generation level, Figure 1 Schematic diagram of the operation of the gas turbine engine during a non-idle operating state and during a power boost operating state.

[0013] Figure 10 According to one aspect of the present disclosure, Figure 1 Schematic diagram of the operation of the gas turbine engine during a non-idle operating state and during a power boost operating state, but with a complete loss of steam generation by the steam generating system.

[0014] Figure 11 According to one aspect of the present disclosure, Figure 1 Schematic diagram of the operation of the gas turbine engine during a non-idle operating state and during a power boost operating state, but with a partial loss of steam generation from the steam generating system.

[0015] Figure 12 According to one aspect of the present disclosure, the operation Figure 1 A flow chart of the process steps of a method for a gas turbine engine. DETAILED DESCRIPTION

[0016] Features, advantages and embodiments of the present disclosure are set forth or apparent by considering the following detailed description, drawings and claims. Furthermore, the following detailed description is exemplary and is intended to provide further explanation without limiting the scope of the disclosure as claimed.

[0017] Various embodiments are discussed in detail below. Although specific embodiments are discussed, this is for illustrative purposes only. Those skilled in the relevant art will recognize that other components and configurations may be used without departing from the scope of this disclosure.

[0018] As used herein, the terms “first” or “second” may be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0019] The terms "upstream" and "downstream" refer to relative directions relative to the flow of fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction towards which the fluid is flowing.

[0020] The terms "fore" and "aft" refer to relative positions within a turbine engine or vehicle and refer to the normal operating attitude of the turbine engine or vehicle. For example, for a turbine engine, the forward position refers to the position closer to the engine inlet, while the aft position refers to the position closer to the engine nozzle or exhaust.

[0021] The terms “coupled,” “fixed,” “attached,” “connected,” and the like refer to both direct coupling, fixing, attachment, or connection, as well as indirect coupling, fixing, attachment, or connection through one or more intermediate components or features, unless otherwise specified herein.

[0022] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0023] As used herein, the terms "axial" and "axially" refer to directions and orientations extending substantially parallel to the centerline of an aircraft gas turbine engine. Additionally, the terms "radial" and "radially" refer to directions and orientations extending substantially perpendicular to the centerline of an aircraft gas turbine engine. Furthermore, as used herein, the terms "circumferential" and "circumferentially" refer to directions and orientations extending arcuately about the centerline of the turbine engine.

[0024] As used herein, "top" refers to the highest or uppermost point, part, or surface of a component in the orientation depicted in the figures.

[0025] As used herein, "bottom" refers to the lowest or lowermost point, part or surface of a component in the orientation depicted in the figures.

[0026] As used herein, the terms "low," "medium" (or "medium"), and "high," or their respective comparatives (e.g., "lower" and "higher," where applicable), when used with reference to a compressor, combustor, turbine, shaft, fan, or turbine engine component, refer to relative pressures, relative speeds, relative temperatures, or relative power outputs within the engine, unless otherwise specified. For example, a "low power" setting defines an engine or combustor configured to operate at a power output lower than the "high power" setting of the engine or combustor, while a "medium power" setting defines an engine or combustor configured to operate at a power output higher than the "low power" setting and lower than the "high power" setting. The terms "low," "medium" (or "medium"), or "high" in the above terms may additionally or alternatively be understood as relative to a minimum allowable speed, pressure, or temperature, or relative to a minimum or maximum allowable speed, pressure, or temperature for normal, desired, steady-state, or other operation of the engine. The duty cycle of a turbine engine includes, for example, low power operation, medium power operation, and high power operation. Low power operation includes, for example, engine starting, idling, taxiing, and approach. Medium power operation includes, for example, cruising. High power operations include, for example, takeoff and climb.

[0027] As used herein, the term "idle operating state" refers to the operating state of a gas turbine engine during engine startup and operation of the turbine engine at a power level less than ten percent of the turbine engine's full power level capability. For example, the idle operating state may be operating the turbine engine after initial engine startup while the aircraft is parked at an airport gate or awaiting the start of taxi. The idle operating state may also be operating the turbine engine upon arrival at the airport gate and prior to shutdown, or operating the turbine engine while the aircraft is stationary at the end of a flight awaiting the start of taxi to the airport gate.

[0028] As used herein, the term "non-idle operating state" refers to an operating state of a turbine engine in which the turbine engine is operating at a power level greater than ten percent or less than approximately thirty percent of its full power capability. For example, the non-idle operating state includes any of the above-mentioned low-power operations (e.g., including approaches when the power level is greater than ten percent or less than approximately thirty percent of its full power capability), any medium-power operation, or a high-power operation, except for engine start-up and idling. Although the foregoing relates to taxiing, approaching, cruising, taking off, and climbing, these operations are generally applicable to turbine engines that are mounted on an aircraft and operated during the flight of the aircraft in which the turbine engine is mounted. However, the non-idle operating state also applies to turbine engines implemented in applications other than aircraft, such as land-based or sea-based implementations. In land-based or sea-based implementations, the turbine engine can be operated to generate electricity, such as to drive a generator or to drive a mechanical drive system. In land-based or sea-based implementations, the non-idle operating state can be a state in which the power level of the turbine engine is sufficient to provide the power required for generating electricity or the power required for driving a mechanical drive system.

[0029] As used herein, the term "non-boost operating state" refers to an operating state of a gas turbine engine in which normal combustion occurs within a first combustor of the gas turbine engine, but a second combustor is in an idle state such that no fuel or steam is supplied to the second combustor. The non-boost operating state may occur during an idle operating state as defined above, or during a non-idle operating state as defined above (when no power boost is being performed).

[0030] As used herein, the term "power boost operating state" refers to an operating state of a gas turbine engine during a non-idle operating state as defined above, but in which a power or thrust boost is being performed. In the power boost operating state, if the steam generation system is normally producing steam, steam generated by the steam generation system is provided to the steam turbine and the second combustor to provide power or thrust boost. Alternatively, if power boost is being performed but the steam generation system is not producing sufficient steam, the power boost operating state is achieved by performing combustion in the second combustor and providing the combustion products of the second combustor to the first combustor to obtain power or thrust boost.

[0031] As used throughout the specification and claims, approximating language is used to modify any quantitative expression that can be permitted to vary without resulting in a change in the basic function to which it is related. Thus, a value modified by one or more terms (e.g., "about," "approximately," "roughly," and "substantially") is not limited to the precise value specified. In at least some cases, approximating language may correspond to the precision of an instrument used to measure a value, or the precision of a method or machine used to construct or manufacture a component and / or system. For example, approximating language may refer to within one percent, two percent, four percent, ten percent, fifteen percent, or twenty percent of a single value, a range of values, and / or an endpoint of a range defining a value.

[0032] Here and throughout the specification and claims, range limitations are combinable and interchangeable. Unless context or language indicates otherwise, such ranges are defined and include all subranges contained therein. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

[0033] In an aircraft gas turbine engine, the combustor may typically include a swirler that provides a swirling flow of air mixed with fuel into a combustion chamber where the fuel and air mixture is ignited and combusted to produce combustion gases that perform work on a turbine within a turbine section to rotate a turbine rotor. The turbine rotor is typically connected to and drives a compressor rotor that compresses inlet air that is provided to the combustor. Under some operating conditions of the gas turbine engine (e.g., during aircraft takeoff or climb), additional power may be required. When the gas turbine engine increases power, the combustor produces additional carbon monoxide (CO) and nitrogen oxides (NO x ) emissions. An attempt to reduce CO and NOx emissions at high power operation x The technology of emission reduction is to include a steam turbine in the gas turbine engine and inject steam into the steam turbine to add additional power to the gas turbine engine, thereby reducing CO and NO x Steam may also be injected into the combustor to increase the density of the combustion gases, thereby providing additional work to be extracted by the turbine section from the denser exhaust gas. However, the use of a steam turbine is generally only feasible when the steam generation system is operating properly to produce sufficient steam to supply the steam turbine or combustor.

[0034] The present disclosure provides a technique for continuing to provide additional power to a gas turbine engine during non-idle operating conditions, even when the steam generation capacity of a steam generation system is reduced or not generating steam at all. According to the present disclosure, a second burner is arranged in parallel with a first (or main) burner, and a second combustion chamber within the second burner is connected to a first combustion chamber within the first burner. When the steam generation system is operating normally, steam can be provided to the second burner, which can then flow from the second combustion chamber into the first combustion chamber of the first burner, thereby enhancing the thrust of the gas turbine engine during non-idle operating conditions. On the other hand, when the steam generation system is not generating steam, fuel and compressor bleed air can be provided to the second burner to generate combustion products within the second combustion chamber. The combustion products from the second combustion chamber can then flow into the first combustion chamber to mix with the combustion gases generated by the first burner, thereby enhancing the thrust of the gas turbine engine. Furthermore, in situations where the steam generation system is generating steam but at a reduced level, insufficient to provide all the required additional power to the steam turbine, a reduced amount of steam can be provided to the second burner, and the second burner can further utilize the compressor bleed air and fuel to ignite the steam-air-fuel mixture within the second combustion chamber. The products of combustion from the second combustor may then be provided to the first combustion chamber to again enhance the thrust of the gas turbine engine.

[0035] Referring now to the accompanying drawings, Figure 1 is a schematic cross-sectional view of an aircraft gas turbine engine 10 taken along a longitudinal centerline axis 12 (provided for reference) of the aircraft gas turbine engine 10, which may be mounted on an aircraft (not shown) and includes a steam generating system 100 (described below), in accordance with one embodiment of the present disclosure. The present disclosure may be implemented in any of various types of aircraft turbine engines, including high bypass turbofan engines, open rotor turbine engines, turbojets, and turboprop engines. The gas turbine engine 10 may be controlled by a controller 13, which may monitor various systems within the gas turbine engine 10 and provide commands for controlling the operation of the gas turbine engine 10. Figure 1 As shown, the aircraft gas turbine engine 10 has a longitudinal direction L (extending parallel to the longitudinal centerline axis 12 ) and a radial direction R orthogonal to the longitudinal direction L. Generally speaking, the aircraft gas turbine engine 10 includes a fan section 14 and a turbo-engine 16 disposed downstream of the fan section 14 .

[0036] The turbocharger 16 includes a housing 18 that surrounds the turbocharger 16 and is generally tubular and defines an annular inlet 20. Figure 1As schematically shown in FIG, turbocharged engine 16 includes, in series flow relationship, a compressor section 21 including a supercharger or low-pressure compressor (LPC) 22, followed downstream by a high-pressure compressor (HPC) 24; a combustion section 26; a turbine section 27 including a high-pressure turbine (HPT) 28, followed downstream by a low-pressure turbine (LPT) 30; and an exhaust section 31 including one or more exhaust nozzles 32. A high-pressure (HP) shaft 34 drivingly connects HPT 28 to HPC 24 for unified rotation with HPC 24. A low-pressure (LP) shaft 36 drivingly connects LPT 30 to LPC 22 for unified rotation with LPC 22. The compressor section 21, combustion section 26, turbine section 27, and exhaust section 31 (including one or more exhaust nozzles 32) collectively define a turbocharged engine air flow path 33 therethrough.

[0037] for Figure 1 In the illustrated embodiment, the fan section 14 includes a fan 38 (eg, a variable pitch fan) having a plurality of fan blades 40 coupled to a disk 42 in a circumferentially spaced relationship. Figure 1 As shown, fan blades 40 extend outwardly from disk 42 generally in a radial direction R. Each fan blade 40 is rotatable relative to disk 42 about a pitch axis P because fan blades 40 are operably coupled to actuators 44 that are configured to collectively and uniformly change the pitch of fan blades 40. Fan blades 40, disk 42, and actuators 44 are rotatable together about longitudinal centerline axis 12 via fan shaft 45, which is powered by LP shaft 36 across a power gearbox (also referred to as a gearbox assembly 46). Gearbox assembly 46 is Figure 1 It is shown schematically in FIG, but includes a plurality of gears (not shown) for adjusting the rotational speed of the fan shaft 45 and adjusting the rotational speed of the fan 38 relative to the LP shaft 36.

[0038] Still see Figure 1 , the disk 42 is covered by a rotatable fan hub 48 having an aerodynamic profile to facilitate airflow through the plurality of fan blades 40. In addition, the fan section 14 includes an annular fan casing or nacelle 50 that circumferentially surrounds the fan 38 and / or at least a portion of the turbocharger engine 16. The nacelle 50 is supported relative to the turbocharger engine 16 by a plurality of circumferentially spaced struts or outlet guide vanes 52. In addition, a downstream section 54 of the nacelle 50 extends over an outer portion of the turbocharger engine 16 to define a bypass airflow passage 56 therebetween. One or more exhaust nozzles 32 may extend through the nacelle 50 and be formed therein. Figure 1In the embodiment of the present invention, the one or more exhaust nozzles 32 include one or more discrete nozzles that are spaced circumferentially around the nacelle 50. Other arrangements of the one or more exhaust nozzles 32 are also possible, including, for example, a single exhaust nozzle that is annular or partially annular around the nacelle 50.

[0039] During a standard operating mode of aircraft gas turbine engine 10, a volume of air 58 enters aircraft gas turbine engine 10 through nacelle 50 and / or inlet 60 of fan section 14. As volume of air 58 passes through fan blades 40, a first portion of air 58, shown as bypass air 62, is directed or routed into bypass airflow passage 56, while a second portion of air 58, shown as turbocharger inlet air 64, is directed or routed into an upstream section of turbocharger air flow path 33, or more specifically, into annular inlet 20 of LPC 22. The ratio between bypass air 62 and turbocharger inlet air 64 is referred to as the bypass ratio. LPC 22 then increases the pressure of turbocharger inlet air 64 to produce compressed air 65, which is directed through HPC 24, where it is further compressed and then directed to combustion section 26. In combustion section 26, compressed air 65 is mixed with fuel 67 and combusted to produce combustion gases 66 (also known as combustion products). One or more stages may be used in each of LPC 22 and HPC 24, with each subsequent stage further compressing compressed air 65.

[0040] Combustion gases 66 are directed from combustion section 26 into HPT 28 and expand through HPT 28, where a portion of the thermal and / or kinetic energy in combustion gases 66 is extracted via successive stages of HPT stator vanes 68 coupled to casing 18 and HPT rotor blades 70 coupled to a rotor connected to HP shaft 34, thereby rotating HP shaft 34 and thereby supporting operation of HPC 24. Combustion gases 66 are then directed into LPT 30 and further expand through LPT 30. Here, a second portion of the thermal and / or kinetic energy is extracted from combustion gases 66 via successive stages of LPT stator vanes 72 coupled to casing 18 and LPT rotor blades 74 coupled to an LPT rotor connected to LP shaft 36, thereby rotating LP shaft 36 and thereby supporting operation of LPC 22 and rotation of fan 38 via gearbox assembly 46. One or more stages may be used in each of HPT 28 and LPT 30.

[0041] The combustion gases 66 are then directed through one or more exhaust nozzles 32 of the turbocharger engine 16 to provide propulsive thrust. Simultaneously with the turbocharger engine inlet air 64 flowing through the turbocharger engine air flow path 33, the bypass air 62 is directed through the bypass flow passage 56 and then discharged from the fan bypass nozzle 76 of the aircraft gas turbine engine 10, also providing propulsive thrust. The HPT 28, the LPT 30, and the one or more exhaust nozzles 32 at least partially define a hot gas path 78 for directing the combustion gases 66 through the turbocharger engine 16.

[0042] As described above, compressed air 65 is mixed with fuel 67 in combustion section 26 to form a fuel and air mixture, which is then combusted to produce combustion gases 66 (combustion products). Fuel 67 may include any type of hydrocarbon fuel used in turbine engines, such as, for example, sustainable aviation fuel (SAF), including biofuel, JetA, JetA-1, or other hydrocarbon fuels. Other fuel types, which may or may not be hydrocarbon fuels but are typically used in aircraft gas turbine engines, may also be used to implement the present disclosure.

[0043] The compressor section 21 also includes a compressor bleed air system 94, which may be generally disposed at the downstream end of the HPC 24 and is arranged to discharge a portion of the compressed air 65 from the HPC 24 as a flow of compressor bleed air 69. The flow of compressor bleed air 69 is provided to a compressor bleed air valve 95 via a compressor bleed air duct 96. The compressor bleed air valve 95 is controllable to divert at least a portion of the flow of compressor bleed air 69 to the combustion section 26 via a compressor bleed air duct 97. The compressor bleed air valve 95 is also controllable to provide at least a portion of the flow of compressor bleed air 69 to other systems within the gas turbine engine 10 via the compressor bleed air duct 99.

[0044] The aircraft gas turbine engine 10 includes a fuel system 80 for providing fuel 67 to the combustion section 26. The fuel system 80 includes a fuel tank 82 for storing the fuel 67 therein and a fuel delivery system 84. The fuel tank 82 may be located on an aircraft (not shown) to which the aircraft gas turbine engine 10 is attached. Figure 1 8, but the fuel system 80 may include any number of fuel tanks 82 as desired. A fuel delivery system 84 delivers the fuel 67 from the fuel tank 82 to the combustion section 26 via one or more fuel supply lines 85. The fuel delivery system 84 also includes a fuel pump 86 to introduce a flow of fuel 67 through the fuel supply line 85 to the combustion section 26. Thus, the fuel pump 86 pumps the fuel 67 from the fuel tank 82 through the fuel supply line 85 into the combustion section 26.

[0045] The aircraft gas turbine engine 10 of the present disclosure includes a steam generation system 100 in fluid communication with one or more exhaust nozzles 32 and the fan bypass nozzle 76. As will be described in greater detail below, the steam generation system 100 generates steam from the combustion gases 66 as they flow through the steam generation system 100 and may deliver at least a portion of the generated steam to the combustion section 26.

[0046] Figure 1 The aircraft gas turbine engine 10 shown in FIG is by way of example only. In other exemplary embodiments, the aircraft gas turbine engine 10 can have any other suitable configuration. For example, in other exemplary embodiments, the fan 38 can be configured in any other suitable manner (e.g., as a fixed pitch fan) and further can be supported using any other suitable fan frame configuration. In addition, in other exemplary embodiments, any other suitable number or configuration of compressors, turbines, shafts, or combinations thereof can be provided. In other exemplary embodiments, aspects of the present disclosure can be incorporated into any other suitable aircraft gas turbine engine, such as, for example, a turbofan engine, an open rotor turbocharged engine, a propfan engine, and / or a turboprop engine.

[0047] Figure 2 According to one aspect of the present disclosure, Figure 1 Schematic diagram of an aircraft gas turbine engine 10 and steam generation system 100. For clarity, Figure 2 An aircraft gas turbine engine 10 is schematically shown in FIG. Figure 1 Some parts depicted and described are not included in Figure 2 As shown in Figure 2 As shown, the steam generation system 100 includes a boiler 102 , a condenser 104 , a water / exhaust gas separator 106 , a water pump 108 , and a steam turbine 110 .

[0048] As described in further detail below, the boiler 102 is a heat exchanger that evaporates liquid water from a water source to produce steam or water vapor. Thus, the boiler 102 is a steam source. Specifically, the boiler 102 is an exhaust gas to water heat exchanger. The boiler 102 is connected to the hot gas path 78 ( Figure 1 ) and is located downstream of the LPT 30. As described in further detail below, the boiler 102 is also in fluid communication with a water pump 108. The boiler 102 may include any type of boiler or heat exchanger for extracting heat from the combustion gases 66 and evaporating the liquid water into steam or water vapor as the liquid water and combustion gases 66 flow through the boiler 102.

[0049] As described in further detail below, the condenser 104 is a heat exchanger that further cools the combustion gases 66 as they flow through the condenser 104. Specifically, the condenser 104 is an air-to-exhaust heat exchanger. The condenser 104 is in fluid communication with the boiler 102 and is located within the bypass airflow passage 56. The condenser 104 can include any type of condenser configured to condense water from the exhaust gas (e.g., the combustion gases 66).

[0050] The water / exhaust separator 106 is in fluid communication with the condenser 104 for receiving the cooled exhaust gas (combustion gases 66) having condensed water entrained therein. The water / exhaust separator 106 is also in fluid communication with one or more exhaust nozzles 32 and a water pump 108. The water / exhaust separator 106 comprises any type of water separator for separating water from the exhaust gas. For example, the water / exhaust separator 106 may comprise a cyclonic separator that uses vortex separation to separate water from the exhaust gas. In such an embodiment, the water / exhaust separator 106 generates a cyclonic flow within the water / exhaust separator 106 to separate the water from the cooled exhaust gas. Figure 2 , the water / exhaust separator 106 is schematically shown as being located in the nacelle 50, but the water / exhaust separator 106 may be located elsewhere within the aircraft gas turbine engine 10, such as, for example, radially inward of the nacelle 50, closer to the turbocharger engine 16. The water / exhaust separator 106 may be driven for rotation by one of the engine shafts (e.g., the HP shaft 34 or the LP shaft 36) via an accessory gearbox (not shown). As described above, the boiler 102 receives liquid water from a water source to generate steam or water vapor. The water source may be a water storage tank 107 disposed between the water / exhaust separator 106 and the water pump 108. Thus, in Figure 2 In the illustrated embodiment, the water storage tank 107 may be the source of water for the boiler 102 .

[0051] A water pump 108 is in fluid communication with the water storage tank 107 and the boiler 102. The water pump 108 can be any suitable pump, such as a centrifugal pump or a positive displacement pump. The water pump 108 directs the separated water 112 stored in the water storage tank 107 through the boiler 102, where it is converted back into steam 114. The steam 114 is conveyed through the steam turbine 110 via the steam supply line 88 to provide work for driving the steam turbine 110.

[0052] In operation, combustion gases 66 (also referred to as exhaust) flow from the LPT 30 into the boiler 102 and condenser 104. The combustion gases 66 transfer heat to water 111 within the boiler 102 to produce steam 114 within the boiler 102, as described in further detail below. The combustion gases 66 then flow into the condenser 104, which condenses the water contained in the combustion gases 66. Bypass air 62 flows through the bypass airflow passage 56 and through or through the condenser 104, extracting heat from the combustion gases 66, cooling the combustion gases 66, and condensing water from the combustion gases 66 to produce an exhaust-water mixture 116. The bypass air 62 then exits the aircraft gas turbine engine 10 through the fan bypass nozzle 76, as described in detail above, to generate thrust. Thus, the condenser 104 may be positioned within the bypass airflow passage 56.

[0053] The exhaust-water mixture 116 flows into the water / exhaust separator 106. The water / exhaust separator 106 separates water and exhaust gas from the exhaust-water mixture 116 to produce separate exhaust gas 118 and water 112. The exhaust gas 118 is discharged from the aircraft gas turbine engine 10 through one or more exhaust nozzles 32 to generate thrust, as described in detail above. Thus, the boiler 102, the condenser 104, and the water / exhaust separator 106 also define the hot gas path 78 ( Figure 1 ) for directing the combustion gases 66, the exhaust-water mixture 116, and the exhaust gas 118 through the steam generating system 100 of the aircraft gas turbine engine 10.

[0054] The water pump 108 helps push water 112 from the water / gas separator 106 into the water storage tank 107 and through one or more water lines (such as Figure 2 The water 112 is pumped (as shown by the arrow of the water 112) so that the water 112 flows into the boiler 102 and mixes with the water 111. The water 111 flows through the boiler 102, and the combustion gas 66 flowing through the boiler 102 transfers heat to the water 111, causing the water 111 to evaporate and generate steam 114.

[0055] The steam turbine 110 is coupled to the LP shaft 36 but may also be coupled to the HP shaft 34. The steam turbine 110 includes one or more stages of steam turbine blades (not shown) and steam turbine stators (not shown). Steam 114 flows from the boiler 102 to the steam turbine 110 via the steam supply line 88, thereby rotating the steam turbine blades of the steam turbine 110 and generating additional work in the LP shaft 36. Additionally, at least a portion of the steam 114 may flow to the combustion section 26 via one or more combustor steam supply lines 98, and a steam control valve 92 may be provided within the combustor steam supply line 98 to control the flow of steam 114 into the combustion section 26. Of the steam 114 provided to the steam turbine 110, the remaining steam (as steam 120) may then flow from the steam turbine 110 back to the boiler 102 via the one or more steam supply lines 90.

[0056] As mentioned above, the compressor section 21 ( Figure 1 ) includes a compressor bleed air system 94, which may be generally disposed at a downstream end of the HPC 24 and is arranged to discharge some of the compressed air 65 from the HPC 24 as a flow of compressor bleed air 69. The flow of compressor bleed air 69 is provided to a compressor bleed air valve 95 via a compressor bleed air duct 96. The compressor bleed air valve 95 may be controlled to divert at least a portion of the flow of compressor bleed air 69 to the combustion section 26 via a compressor bleed air duct 97. The compressor bleed air valve 95 may also be controlled to provide at least a portion of the flow of compressor bleed air 69 to other systems within the gas turbine engine 10 via the compressor bleed air duct 99.

[0057] Figure 3 According to one aspect of the present disclosure, Figure 1 A partial cross-sectional side view of an exemplary combustion section 26 of a turbocharged engine 16 is shown. Figure 3The exemplary combustion section 26 shown in FIG is depicted as an annular combustion section that extends circumferentially about the longitudinal centerline axis 12. With respect to the combustion section 26, the longitudinal centerline axis 12 may also correspond to the combustor centerline axis 12'. The combustion section 26 includes a combustor outer shell 124 and a combustor inner shell 126, each of which extends annularly about the combustor centerline axis 12'. A first combustor 125 is arranged within the combustor outer shell 124 and the combustor inner shell 126. The first combustor 125 includes a first combustor inner liner 130, a first combustor outer liner 132, and a dome structure 134, each of which extends circumferentially about the combustor centerline axis 12'. A first combustion chamber 131 is defined between the first combustor inner liner 130, the first combustor outer liner 132, and the dome structure 134. As will be described in more detail below, the first combustion chamber 131 can be theoretically divided into a primary combustion zone 133 and a secondary combustion zone 137 located downstream of the primary combustion zone 133. Figure 3 In the figure, the dotted line 135 represents the theoretical division between the primary combustion zone 133 and the secondary combustion zone 137.

[0058] The first combustor outer liner 132 may include various airflow openings therethrough, including a plurality of primary cooling openings 136, a plurality of dilution openings 138 ( Figure 3 Similarly, the first combustor inner liner 130 may include various airflow openings therethrough, including a plurality of primary cooling openings 142, a plurality of dilution openings 144 ( Figure 3 28 ), and a plurality of secondary cooling openings 146. Furthermore, the dome structure 134 may include a plurality of cooling airflow openings 148 therethrough. Furthermore, the downstream end 143 of the combustor outer casing 124 may include a plurality of airflow openings 141 to provide a turbine cooling airflow 183 into the HPT 28, and the downstream end 149 of the combustor inner casing 126 may include a plurality of airflow openings 147 to provide a turbine cooling airflow 189 into the HPT 28.

[0059] The first combustor 125 further includes a plurality of swirler assemblies 156 ( Figure 3 One is shown in FIG), and a plurality of first fuel nozzles 158 ( Figure 3 One is shown in FIG. 1 . As described below, each of the plurality of first fuel nozzles 158 injects fuel 67 into a corresponding one of the swirler assemblies 156 .

[0060] The first combustor inner liner 130 and the first combustor outer liner 132 are connected to the dome structure 134, thereby defining a first combustion chamber 131 therebetween. The first combustor inner liner 130 and the first combustor outer liner 132 extend from the dome structure 134 to the HPT 28 ( Figure 1 ) at the inlet of the combustor, thereby at least partially defining a hot gas path between the dome structure 134 and the HPT 28. In addition, a shroud 152 is connected to the first combustor inner liner 130, the first combustor outer liner 132, and the dome structure 134, thereby defining a plenum 154 therein. The shroud 152 extends circumferentially about the combustor centerline axis 12' and can be formed by a single shroud structure or by multiple shroud structures connected together. The shroud 152 includes a plurality of shroud airflow openings 157 ( Figure 3 15), wherein each opening corresponds to a respective one of the plurality of swirler assemblies 156. Each shroud airflow opening 157 provides airflow therethrough into the plenum 154. The shroud 152 is connected to the combustor casing 124 by shroud mounting arms 153.

[0061] like Figure 3 As shown, the combustor outer shell 124 and the combustor inner shell 126 surround the first combustor outer liner 132 and the first combustor inner liner 130. An outer airflow passage 160 is defined between the combustor outer shell 124 and the first combustor outer liner 132, and an inner airflow passage 162 is defined between the combustor inner shell 126 and the first combustor inner liner 130. A diffuser 164 is coupled to the combustion section 26 between an upstream end 166 of the combustor outer shell 124 and an upstream end 168 of the combustor inner shell 126. A plenum 170 is defined between the upstream end 166 of the combustor outer shell 124 and the upstream end 168 of the combustor inner shell 126. The diffuser 164 provides a flow of compressed air 65 from the HPC 24 into the plenum 170.

[0062] Still refer to Figure 3During operation of the aircraft gas turbine engine 10, compressed air 65 flows through the diffuser 164 and into the plenum 170 of the combustion section 26 to pressurize the plenum 170. A first portion of the compressed air 65 in the plenum 170 (as schematically indicated by the arrow representing compressed air 172) flows from the plenum 170 into the plenum 154 of the shroud 152. The compressed air 172 then flows through the swirler assemblies 156 where it mixes with the fuel 67 provided by the first fuel nozzles 158. A fuel-air mixture 191 is generated within each swirler assembly 156. The fuel-air mixture 191 is then injected from the swirler assemblies 156 into the first combustion chamber 131, where the air / fuel mixture is ignited by an igniter (not shown) and combusts to produce combustion gases 66 within the first combustion chamber 131. A portion of the compressed air 172 within the plenum 154 (schematically indicated by arrows representing cooling air 190) may flow through the cooling air flow openings 148 in the dome structure 134 to provide cooling of the downstream side of the dome structure 134. Figure 3 Although not shown, the dome structure 134 may include a deflector or heat shield on the downstream side to protect the dome structure 134 from the heat generated in the first combustion chamber 131, and the cooling air flow openings 148 will extend through the deflector or heat shield.

[0063] A second portion of the compressed air 65 in the plenum 170 (as schematically indicated by arrows representing compressed air 174 and compressed air 176) can be directed into the outer airflow passage 160 and the inner airflow passage 162, respectively. A portion of the compressed air 174 flowing through the outer airflow passage 160 (schematically shown as cooling air 178) can be directed into the primary combustion zone 133 of the first combustion chamber 131 through the plurality of primary zone cooling openings 136. Another portion of the compressed air 174 flowing through the outer airflow passage 160 (schematically shown as cooling air 182) can be directed into the secondary combustion zone 137 of the first combustion chamber 131 through the secondary zone cooling openings 140. Similarly, a portion of the compressed air 176 flowing through the inner airflow passage 162 (schematically shown as cooling air 184) can be directed into the primary combustion zone 133 of the first combustion chamber 131 through the plurality of primary zone cooling openings 142. Another portion of the compressed air 176 flowing through the inner airflow passage 162 (schematically shown as dilution airflow 186) may be directed into the first combustion chamber 131 through the dilution openings 144 of the first combustor inner liner 130 to provide quenching of the combustion gases 66. Another portion of the compressed air 176 flowing through the inner airflow passage 162 (schematically shown as cooling air 188) may be directed into the secondary combustion zone 137 of the first combustion chamber 131 through the secondary zone cooling openings 146.

[0064] exist Figure 3 In the embodiment, the combustion section 26 further includes a second burner 192 disposed within the outer airflow passage 160. Figure 3 Second combustor 192 is depicted as being arranged within outer airflow passage 160, but second combustor 192 may also be arranged within inner airflow passage 162. Second combustor 192 includes a second combustor casing 194 and a second combustor liner 196 arranged within second combustor casing 194. Second combustor liner 196 defines a second combustion chamber 198 therein. Compressor bleed air duct 97 is connected to second combustor 192 to provide a flow of compressor bleed air 69 to second combustor 192. In addition, combustor steam supply line 98 is connected to second combustor 192 to provide a flow of steam 114 to second combustor 192. Further, second combustor 192 includes a second fuel nozzle 200 arranged to provide a second fuel flow 67b to second combustion chamber 198, and an igniter 202 arranged to provide a spark to second combustion chamber 198 to ignite the fuel-air mixture within second combustion chamber 198 (as described below).

[0065] Figure 4 According to one aspect of the present disclosure, Figure 3 204 . As briefly described above, the second combustor 192 is disposed within the outer gas flow passage 160 between the combustor casing 124 and the first combustor outer liner 132 . The second combustor 192 includes a second combustor housing 194 that may be mounted to the first combustor outer liner 132 via one or more mounting brackets 206 . However, the second combustor housing 194 may be mounted to the combustor casing 124 . The second combustor housing 194 includes a second combustor liner 196 therein that defines a second combustion chamber 198 therein. Figure 4 In the embodiment shown, a single second combustor liner 196 is disposed within the second combustor casing 194. However, as described below, more than one second combustor liner 196 may be disposed within the second combustor casing 194. Figure 4The second combustor liner 196 in FIG. 1 is shown configured as a trapped vortex combustor liner 197, although other types of combustor liners may be implemented in place of a trapped vortex combustor liner. Second combustor liner 196 includes an outlet 199 that extends through second combustor casing 194 and also through first combustor outer liner 132 via dilution openings 138. Outlet 199 provides fluid communication from second combustion chamber 198 to secondary combustion zone 137 of first combustion chamber 131. Second combustor liner 196 may be mounted within second combustor casing 194 via one or more mounting brackets 208 that include one or more openings 209 extending therethrough. A second combustor plenum 210 is defined between second combustor casing 194 and second combustor liner 196 and surrounds second combustor liner 196. Openings 209 through mounting brackets 208 allow air or steam to flow freely throughout second combustor plenum 210 around the outside of second combustor liner 196. Alternatively, the second combustor liner 196 may be integrally formed with the first combustor outer liner 132 rather than being connected to the first combustor outer liner 132 via the mounting bracket 208. For example, the second combustor liner 196 may be welded or brazed to the first combustor outer liner 132.

[0066] like Figure 4 As shown, a compressor bleed air duct 97 extends through the combustor casing 124 and is connected to the second combustor housing 194 to provide fluid communication from the compressor bleed air duct 97 to the second combustor plenum 210. In addition, a combustor steam supply line 98 extends through the combustor casing 124 and is connected to the second combustor housing 194 to provide fluid communication from the combustor steam supply line 98 to the second combustor plenum 210. As described below, in some operating conditions of the gas turbine engine 10, compressor bleed air 69 is provided to the second combustor plenum 210 via the compressor bleed air duct 97, or steam 114 is provided to the second combustor plenum 210 via the combustor steam supply line 98. A second fuel nozzle 200 extends through the combustor casing 124 and through the second combustor liner 196 to provide a second fuel flow 67b to the second combustion chamber 198. The igniter 202 also extends through the combustor casing 124 and the second combustor liner 196 and provides a spark to ignite a fuel-air mixture 218 within the second combustion chamber 198 when combustion is to occur within the second combustion chamber 198 .

[0067] The second combustor liner 196 includes a plurality of openings 212 extending through the second combustor liner 196. Each opening 212 provides fluid communication from the second combustor plenum 210 to the second combustion chamber 198. As described below, when the compressor bleed air 69 is provided to the second combustor plenum 210 via the compressor bleed air duct 97, the compressor bleed air 69 flows from the second combustor plenum 210 into the second combustion chamber 198 through the openings 212. As described above, the second combustor liner 196 may be configured as a trapped vortex combustor liner 197. Therefore, the openings 212 are arranged to pass through the trapped vortex combustor liner 197 so as to generate a trapped vortex flow 214 within the trapped vortex combustor liner 197. Figure 4 In the embodiment, the arrangement of the openings 212 is such that a counterclockwise trapped vortex flow 214 is generated. When the compressor bleed air 69 is provided to the second combustion chamber 198 via the openings 212 and combustion is to take place in the second combustion chamber 198, the second combustor fuel valve 216 is opened to allow the second fuel flow 67b (the first fuel flow 67a is provided to the swirler assembly 156 via the first fuel nozzle 158, as described above for Figure 3 The second fuel stream 67b and the compressor bleed air 69 are mixed together in the trapped vortex flow 214 to produce a fuel-air mixture 218, and the fuel-air mixture 218 is ignited by the igniter 202, thereby producing first combustion products 220 in the second combustion chamber 198. The first combustion products 220 then flow through the outlet 199 into the secondary combustion zone 137 of the first combustion chamber 131 of the first combustor 125. Therefore, when combustion is to be carried out in the second combustor 192 (for example, during a non-idle operating state of the gas turbine engine 10), the first combustion products 220 can be combined with the combustion gases 66 ( Figure 3 ) mixture, thereby increasing the thrust of the gas turbine engine 10.

[0068] Continue the discussion Figure 4 , when steam 114 is supplied to the second burner plenum 210 via the burner steam supply line 98, the steam 114 flows from the second burner plenum 210 into the second combustion chamber 198 through the opening 212. Then, the steam 114 can flow from the outlet 199 of the second combustion chamber 198 into the secondary combustion zone 137 of the first combustion chamber 131 in the first burner 125. In this way, the steam 114 flowing into the first combustion chamber 131 can be combined with the combustion gas 66 ( Figure 3 ) to provide some quenching of the combustion gases 66 and to increase the density of the combustion gases 66, thereby Figure 1 ) during the non-idle operating state to drive the turbine section 27 ( Figure 1 ) of additional kinetic energy.

[0069] In yet another embodiment, when steam 114 is provided to the second combustor plenum 210 and compressor bleed air 69 is provided to the second combustor plenum 210, the compressor bleed air 69 and steam 114 mix together within the second combustor plenum 210 to form a steam-air mixture 222 within the second combustor plenum 210. The steam-air mixture 222 then flows through the opening 212 of the second combustor liner 196 into the second combustion chamber 198. The second fuel nozzle 200 is also capable of providing a second fuel stream 67b into the second combustion chamber 198, and the second fuel stream 67b mixes with the steam-air mixture 222 within the second combustion chamber 198 to form a steam-air-fuel mixture 224. The steam-air-fuel mixture 224 is ignited by the igniter 202 and combusts to produce second combustion products 226. The second combustion products 226 then flow through the outlet 199 into the secondary combustion zone 137 of the first combustion chamber 131. Thus, similar to the first combustion products 220, the second combustion products 226 can react with the combustion gases 66 ( Figure 3 ) mixed to increase the thrust of the gas turbine engine 10.

[0070] Figure 5 According to one aspect of the present disclosure, Figure 3 A partial cross-sectional view of a semi-annular second combustor 192' taken at plane 5-5 of FIG. 1 , wherein the semi-annular second combustor has a plurality of second combustion chambers as Figure 3 As mentioned above, the second burner housing 194 ( Figure 4 ) may include more than one second combustor liner 196 ( Figure 4 ). Figure 5 A second combustor casing 194' is depicted that includes a plurality of second combustor liners, including a second combustor liner 196a and a second combustor liner 196b. The second combustor liner 196a and the second combustor liner 196b may be similar to Figure 4 The second combustor liner 196 in the embodiment of the present invention may be a second combustor liner 196a and a second combustor liner 196b. Figure 4 The second fuel nozzle 200a extends through the combustor casing 124, through the second combustor housing 194' and through the second combustor liner 196a, thereby providing a fuel flow, namely the second fuel flow 67b, to the second combustion chamber 198a of the second combustor liner 196a. Figure 4Similarly, compressor bleed air duct 97 extends through combustor casing 124 and through second combustor housing 194' to provide a flow of compressor bleed air 69 to second combustor plenum 210'. Additionally, combustor steam supply line 98 extends through combustor casing 124 and through second combustor housing 194' to provide a flow of steam 114 to second combustor plenum 210'.

[0071] Each of the second combustor liner 196a and the second combustor liner 196b includes Figure 4 212 therethrough, but also includes other openings. For example, the first sidewall 228 of the second combustor liner 196a includes at least one sidewall opening 213a therethrough, and the second sidewall 230 of the second combustor liner 196a includes at least one sidewall opening 215a therethrough. Similarly, for the second combustor liner 196b, the first sidewall 232 of the second combustor liner 196b includes at least one sidewall opening 213b therethrough, and the second sidewall 234 of the second combustor liner 196b includes at least one sidewall opening 215b therethrough. Figure 4 Similar to the openings 212 of the second combustor liner 196a, any of the compressor bleed air 69, steam 114, or steam-air mixture 222 can flow through sidewall openings 213a, 215a, 213b, and 215b, respectively, into the second combustion chamber 198a and the second combustion chamber 198b. The sidewall openings 213a and 215a of the second combustor liner 196a are arranged to generate a swirl flow 236a within the second combustion chamber 198a, and the sidewall openings 213b and 215b of the second combustor liner 196b are arranged to generate a swirl flow 236b within the second combustion chamber 198b. The swirl flow 236a and swirl flow 236b can help better mix the fuel 67 with the compressor bleed air 69 or the steam-air mixture 222.

[0072] Also like Figure 5 As shown, the second combustor liner 196a has an outlet 199a (similar to Figure 4 The second combustor liner 196b has an outlet 199b (also similar to Figure 4 Each of the outlets 199a and 199b is in fluid communication with the secondary combustion zone 137 of the first combustion chamber 131 of the first burner 125. Figure 4 As mentioned, for Figure 5On the other hand, outlet 199a of second combustor liner 196a provides a flow of any of first combustion products 220a, steam 114a, or second combustion products 226a into secondary combustion zone 137. Similarly, outlet 199b of second combustor liner 196b provides a flow of any of first combustion products 220b, steam 114b, or second combustion products 226b into secondary combustion zone 137.

[0073] like Figure 5 As further shown, the second burner 192' extends partially annularly about the burner centerline axis 12'. Thus, the second burner 192' can be considered a partial burner 193 that does not extend annularly about the burner longitudinal centerline axis 12'. In other words, the second burner 192' is a partial annular burner 195 that extends partially annularly about the burner centerline axis 12'. Although Figure 5 A single second combustor 192' is shown in FIG, but the combustion section 26 may include more than one second combustor 192', wherein the second combustors 192' are circumferentially spaced about the combustor centerline axis 12'. Figure 1 ) may include a plurality of second burners 192 ( Figure 4 ), wherein, likewise, each second combustor 192 is circumferentially spaced about the combustor centerline axis 12 ′.

[0074] Figure 6 According to another aspect of the present disclosure, Figure 4 A partial cross-sectional side view of an optional second burner 237 is shown. Figure 6 , the optional second combustor 237 includes a second combustor casing 238 and a second combustor liner 240, wherein a second combustor plenum 242 is defined between the second combustor casing 238 and the second combustor liner 240. The second combustor casing 238 is similar to Figure 4 A second burner housing 194 is provided. Figure 4 Similar in aspect, Figure 6 In some aspects, the second combustor 237 includes a compressor bleed air duct 97 that extends through the combustor casing 124 and through the second combustor housing 238 to provide a flow of compressor bleed air 69 into the second combustor plenum 242. Additionally, the second combustor 237 includes a combustor steam supply line 98 that extends through the combustor casing 124 and through the second combustor housing 238 to provide a flow of steam 114 into the second combustor plenum 242. In some aspects, as described below, the compressor bleed air 69 and the steam 114 can be provided to the second combustor plenum 242 simultaneously, resulting in a steam-air mixture 244 within the second combustor plenum 242.

[0075] and Figure 4 On the contrary, Figure 6 The second combustor liner 240 is not a trapped vortex combustor liner, but rather an axial flow second combustor liner 246 that includes a dome 248, an upstream liner portion 250, and a downstream liner portion 252 that define a second combustion chamber 253 therein. The upstream liner portion 250 and the downstream liner portion 252 further define an outlet 254 therebetween. The upstream liner portion 250 and the downstream liner portion 252 are arranged at an angle 256 relative to the combustor centerline axis 12' such that the outlet 254 is arranged at an angle 256 to direct flow exiting the outlet 254 in a generally downstream direction 258. The outlet 254 may also be arranged at a circumferential angle relative to a circumferential direction C about the centerline axis 12' (not shown). By arranging the outlet 254 at a circumferential angle, tangential flow from the outlet 254 may be introduced into the secondary combustion zone 137 of the combustor 131 for faster and better mixing with the combustion gases 66.

[0076] The dome 248 includes a plurality of dome openings 260 extending through the dome 248. Each of the plurality of dome openings 260 may be arranged at an angle 261 to provide a swirl flow 263 within the second combustion chamber 253. The plurality of dome openings 260 allow any of the compressor bleed air 69, steam 114, or steam-air mixture 244 to flow from the second combustor plenum 242 into the second combustion chamber 253. Figure 4 Similarly, the second combustor 237 includes at least one second fuel nozzle 200 and an igniter 202, both of which extend through the combustor casing 124, through the second combustor housing 238, and through the dome 248 of the second combustor liner 240. The at least one second fuel nozzle 200 can therefore provide a second fuel stream 67b into the second combustor liner 240 to mix with the compressor bleed air 69 to form a fuel-air mixture 262 or with the steam-air mixture 244 to form a steam-air-fuel mixture 264 within the second combustion chamber 253. The igniter 202 can then provide a spark to ignite the fuel-air mixture 262 to produce first combustion products 266 or ignite the steam-air-fuel mixture 264 to produce second combustion products 268. Depending on the operating state of the second burner 237, as described below, any one of the steam 114, the first combustion product 266, or the second combustion product 268 flows from the outlet 254 into the secondary combustion zone 137 of the first combustion chamber 131 to mix with the combustion gas 66 ( Figure 3 )mix.

[0077] Figure 7 According to one aspect of the present disclosure, Figure 6 A partial cross-sectional view of the optional second combustor liner 240 is shown taken at plane 7-7 of FIG. Figure 7, the plurality of dome openings 260 are shown as being arranged at a tangential angle 270 relative to a radial direction 272 extending from the combustor centerline axis 12′. Furthermore, the second fuel nozzle 200 is arranged at the tangential angle 270. Thus, the plurality of dome openings 260 can provide a flow of any of the compressor bleed air 69, steam 114, or the steam-air mixture 244 into the second combustion chamber 253 in a tangential direction (i.e., tangential to the circumferential direction C about the combustor centerline axis 12′) to allow for better mixing within the second combustion chamber 253.

[0078] Figures 1 to 7 The foregoing description describes various configurations of the combustion section 26 of the gas turbine engine 10. The gas turbine engine 10 can operate in any of a low-power operating state (e.g., during aircraft descent or landing), a normal-power operating state (e.g., during aircraft taxi or cruise), or a high-power operating state (e.g., during aircraft takeoff and climb). The foregoing description also discusses providing either compressor bleed air 69, steam 114, or both compressor bleed air 69 and steam 114 to the second combustor 192 or the second combustor 237. The following description will discuss various operating states of the combustion section 26 in conjunction with the operation of the steam generation system 100, with a focus on the operation of the second combustor 192 or the second combustor 237. More specifically, the following description generally describes a first scenario in which the steam generating system 100 operates normally during a non-idle operating state to provide steam to the second burner 192, a second scenario in which the steam generating system 100 completely (or nearly completely) loses steam generating capacity during a non-idle operating state, and a third scenario in which the steam generating system 100 partially loses steam generating capacity during a non-idle operating state.

[0079] Figure 8 is a schematic diagram depicting operation of the gas turbine engine 10 during a non-boost operating state, in which steam 114 is not provided to the second combustor 192 or the steam turbine 110 , according to one aspect of the present disclosure. Figure 8 The non-power boost operating state of the gas turbine engine 10 in FIG. 1 may be an idle operating state, or may be a non-idle operating state in which thrust boost is not performed and, therefore, steam 114 is not provided to the second combustor 192 or the steam turbine 110. Generally, in the non-power boost operating state, although the steam generation system 100 may operate normally to generate steam 114 ( Figure 2 ), but the second combustor 192 and the steam turbine 110 are usually idle and not operating. Figure 8 As shown above, and as Figures 1 to 3As described above, during operation of the gas turbine engine 10, compressed air 65 is provided from the HPC 24 to the combustion section 26. The compressor bleed air system 94 is operable and can discharge a portion of the compressed air 65 through the compressor bleed air valve 95 to provide compressor bleed air 69 to various other systems (except the second combustor 192). However, during non-power enhancement operating conditions, the compressor bleed air 69 is not provided to the second combustor 192. The fuel delivery system 84 provides the first fuel flow 67a to the swirler assembly 156 through the first fuel nozzle 158. The second combustor fuel valve 216 can also be closed so as not to provide the second fuel flow 67b to the second combustor 192. The fuel-air mixture 191 ( Figure 3 ) is injected into the first combustion chamber 131 of the first combustor 125 and is ignited and combusted to generate combustion gas 66 that flows to the HPT 28 and the LPT 30. The combustion gas 66 flows through the steam generation system 100 and then flows through the exhaust nozzle 32 ( Figure 1 and Figure 2 ). In the non-power-enhanced operating state, the steam generation system 100 can generate steam 114, but the steam control valve 92 is closed, thereby not providing steam 114 to the steam turbine 110 or the second combustor 192. Each of the above operations is controlled by the controller 13 ( Figure 1 )control.

[0080] Figure 9 is a schematic diagram depicting operation of a gas turbine engine 10 during a non-idle operating state and during a power boost operating state when the steam generating system 100 is operable to generate steam 114 at a first steam generation level, in accordance with one aspect of the present disclosure. The first steam generation level generally refers to normal operation of the steam generating system 100 to generate a sufficient amount of steam 114 to enable normal operation of the steam turbine 110 at or near full load capacity to provide power or thrust boost, and normal operation of the second combustor 192 to provide power or thrust boost. In other words, the steam generating system 100 operates without any significant loss of steam generating capacity, even though there may be some small loss of steam generating capacity within the steam generating system 100. Figure 9 , the first combustor 125 operates normally as described above, albeit in a non-idle operating state rather than an idle operating state. With respect to the second combustor 192, when operating in a non-idle operating state, and when the combustion section 26 may require a power or thrust boost, the controller 13 controls the steam generation system 100 to generate steam 114 and controls the steam control valve 92 to provide a flow of steam 114 to the second combustor 192 (schematically shown by the arrow connecting the steam control valve 92 and the second combustor 192), as well as to the steam turbine 110. As described above with respect to Figure 4As described above, for example, steam 114 is provided to the second combustor plenum 210 via the combustor steam supply line 98. The steam 114 flows through the opening 212 in the second combustor liner 196 into the second combustion chamber 198, and then enters the secondary combustion zone 137 of the first combustion chamber 131 from the second combustion chamber 198 via the outlet 199. At the same time, the controller 13 controls the compressor bleed air valve 95 to be closed, thereby prohibiting the compressor bleed air system 94 from providing the flow of compressor bleed air 69 to the second combustor plenum 210. The controller 13 also controls the second combustor fuel valve 216 to be in a closed state, thereby prohibiting the second fuel flow 67b from flowing to the second combustion chamber 198. Therefore, in Figure 9 In another aspect, power or thrust enhancement may be achieved by supplying a flow of steam 114 to the steam turbine 110 and the second combustor 192 .

[0081] Figure 10 1 is a schematic diagram illustrating the operation of the gas turbine engine 10 during a non-idle operating state and a power boost operating state, according to one aspect of the present disclosure, but with a complete loss of steam generating capacity of the steam generating system 100. Here, a complete loss of steam generating capacity is intended to mean that the steam generating system 100 is unable to operate to generate steam 114 (e.g., insufficient water 111 in the boiler 102) and has been shut down by the controller 13, or the steam generating system 100 may be capable of generating only a small amount of steam, but the amount of steam generated is significantly insufficient to operate the steam turbine 110 or to provide power boost by providing steam 114 to the second combustor 192.

[0082] exist Figure 10 In, with Figure 8 and Figure 9 In the same way, the first burner 125 ( Figure 3 ) operates normally as described above, albeit in a non-idle operating state rather than an idle operating state. With respect to the second combustor 192, when the gas turbine engine 10 is operating in a non-idle operating state, and when the combustion section 26 may require a power boost, but the steam generation system 100 has completely lost its steam generation capacity, the controller 13 controls the compressor bleed air valve 95 to open to provide a flow of compressor bleed air 69 to the second combustor 192, and the controller 13 also controls the second combustor fuel valve 216 to open to provide a second fuel flow 67b to the second combustor 192. Simultaneously, since the steam generation system 100 is not generating steam 114 (as indicated by the absence of an arrow labeled steam 114 flowing from the steam generation system 100 to the steam control valve 92), the controller 13 controls the steam control valve 92 to be in a closed state.

[0083] As mentioned above Figure 4As described above, when the second fuel stream 67b and the stream of compressor bleed air 69 are provided to the second combustor 192, the stream of compressor bleed air 69 flows into the second combustor plenum 210 and then enters the second combustion chamber 198 through the opening 212. The second fuel stream 67b is injected into the second combustion chamber 198 and mixed with the compressor bleed air 69 to produce a fuel-air mixture 218. The fuel-air mixture 218 is ignited by a spark from the igniter 202 to produce first combustion products 220. The first combustion products 220 then flow through the outlet 199 into the secondary combustion zone 137 of the first combustion chamber 131. Therefore, by adding the first combustion products 220 to mix with the combustion gases 66, a power (or thrust) enhancement can be achieved, even if the steam generation system 100 completely loses steam generation capacity.

[0084] Figure 11 is a schematic diagram depicting the operation of the gas turbine engine 10 during a non-idle operating state and a power boost operating state, but with the steam generating system 100 operating to generate steam at a second steam generation level that is lower than the first steam generation level. The second steam generation level refers to a situation where the steam generating system 100 suffers a partial but significant loss of steam generating capacity such that the steam generating system 100 is operable to generate steam 114, but at a reduced capacity (e.g., fifty percent or less of capacity). That is, the steam generating system 100 is able to operate the steam turbine 110 at approximately half of its normal full power capacity. Figure 11 In, with Figure 8 、 Figure 9 as well as Figure 10 Similarly, the first combustor 125 operates normally as described above, albeit in a non-idle operating state rather than an idle operating state. With respect to the second combustor 192, when operating in a non-idle operating state, and when the combustion section 26 may require a power or thrust boost, but the steam generation capacity of the steam generation system 100 is partially lost, the controller 13 controls the steam control valve 92 to open so as to provide a flow of steam 114 to the second combustor 192 and the steam turbine 110 (although the flow of steam 114 is reduced). The controller 13 further controls the compressor bleed air valve 95 to open so as to provide a flow of compressor bleed air 69 to the second combustor 192, and the controller 13 also controls the second combustor fuel valve 216 to open so as to provide a second fuel flow 67b to the second combustor 192.

[0085] As mentioned above about Figure 4As described, the flow of compressor bleed air 69 flows into the second combustor plenum 210 via the compressor bleed air duct 97, and the flow of (reduced) steam 114 flows into the second combustor plenum 210 via the combustor steam supply line 98. The compressor bleed air 69 and the steam 114 are mixed together in the second combustor plenum 210 to produce a steam-air mixture 222. The steam-air mixture 222 flows into the second combustion chamber 198 through the opening 212. At the same time, the second fuel flow 67b is injected into the second combustion chamber 198 through the second fuel nozzle 200, and the fuel 67b is mixed with the steam-air mixture 222 in the second combustion chamber 198 to produce a steam-air-fuel mixture 224. The steam-air-fuel mixture 224 is ignited by the igniter 202 and combusted in the second combustion chamber 198 to produce second combustion products 226. The second combustion products 226 flow through the outlet 199 into the secondary combustion zone 137 of the first combustion chamber 131 to mix with the combustion gas 66 ( Figure 3 Thus, by adding the second combustion products 226 to mix with the combustion gases 66, a power (or thrust) enhancement may be achieved even if the steam generating system 100 partially loses steam generating capacity.

[0086] Figure 12 is a flow chart of process steps for a method of operating a gas turbine engine 10 according to one aspect of the present disclosure. Figure 12 The method may be implemented in a gas turbine engine 10, involving the above-mentioned Figures 1 to 11 In step S1201, the gas turbine engine 10 is operated in a non-idle operating state and a non-power-boosted operating state. The operation of the gas turbine engine 10 in the non-idle operating state and the non-power-boosted operating state may correspond to the above description. In step S1202, when operating in the non-idle operating state and the non-power-boosted operating state, the controller 13 ( Figure 1 ) determines whether the gas turbine engine 10 is commanded to operate in a power boost operating state to provide a power (or thrust) boost. The controller 13 may be notified of a power increase by an avionics system within the aircraft, for example, if the power increase exceeds ten percent of the full power capability of the gas turbine engine 10, or exceeds thirty percent of the full power capability of the gas turbine engine 10. As part of this determination, the controller 13 may also utilize avionics signals indicating the flight state of the aircraft, such as whether the aircraft is in a takeoff state or a climb state, or whether the aircraft is in a cruise state or a landing approach state. If the determination in step S1202 is negative, the gas turbine engine 10 continues to operate in a non-idle operating state and a non-power boost operating state in step S1201.

[0087] On the other hand, if the determination in step S1202 is yes, the steam generation system 100 ( Figure 1 、 Figure 2 ) is operating normally. That is, the controller 13 monitors the status of the steam generation system 100 during operation of the gas turbine engine 10, and the controller 13 determines whether the steam generation system 100 is operating normally to generate a sufficient amount of steam 114 to operate the steam turbine 110. If the determination in step S1203 is yes, then in step S1204, the controller 13 controls the gas turbine engine 10 in accordance with the above Figure 9 1 and 2. The described operation is to provide a flow of steam 114 to the second combustor 192 and the steam turbine 110.

[0088] On the other hand, if the determination in step S1203 is NO (i.e., the steam generation system 100 is not operating normally and is not generating steam or the capacity of the generated steam is reduced), then in step S1205, the controller 13 determines whether the steam generation system 100 is partially operating to generate steam 114 (i.e., whether the steam generation capacity is partially lost). If the determination in step S1205 is NO, the controller 13 determines that the steam generation capacity has been completely lost, and in step S1206, controls the gas turbine engine 10 to the above-mentioned control by supplying the compressor bleed air 69 and the second fuel flow 67b to the second combustor 192 while prohibiting the steam 114 from flowing from the steam generation system 100 to the second combustor 192 and the steam turbine 110. Figure 10 Operate in this way.

[0089] When the judgment in step S1205 is yes (ie, the steam generation capacity of the steam generation system 100 is partially lost), then in step S1207, the controller 13 controls the gas turbine engine 10 to Figure 11 The system operates in the manner described, thereby providing a flow of compressor bleed air 69 , a flow of the second fuel 67 b , and steam 114 to the second combustor 192 .

[0090] Further aspects of the disclosure are provided by the subject matter of the following clauses.

[0091] 18. The combustion chamber of claim 17, wherein the combustion chamber has an inlet and outlet openings configured to extend outwardly from the combustion chamber and to provide a plurality of outlet ports therein. The combustion chamber comprises a first combustion chamber, wherein the first combustion chamber comprises an outlet port, and a second combustion chamber comprises an outlet port. The second combustion chamber comprises an outlet port, and wherein the first combustion chamber comprises an outlet port. The second combustion chamber comprises an outlet port, and wherein the first combustion chamber comprises an outlet port. and a steam generating system, wherein the steam generating system comprises a steam generating unit and a steam generating unit; wherein the steam generating unit comprises a steam generating unit and a steam generating unit; and wherein the steam generating unit comprises a steam generating unit and a steam generating unit; and wherein the steam generating unit comprises a steam generating unit and a steam generating unit.

[0092] A gas turbine engine according to any preceding clause, wherein the outlet is arranged to direct a flow of steam or a flow of combustion products resulting therefrom in a downstream direction within the secondary combustion zone of the first combustion chamber.

[0093] A gas turbine engine according to any preceding clause, wherein the compressor bleed air system includes a bleed air valve that controls the flow of the compressor bleed air to the second combustor plenum, and the steam generation system includes a steam control valve that controls the flow of steam to the second combustor plenum.

[0094] A gas turbine engine according to any of the preceding items, wherein the second combustor comprises (i) a plurality of second combustor liners, each of the plurality of second combustor liners defining a corresponding second combustion chamber, each second combustion chamber having a corresponding outlet fluidically connected to the secondary combustion zone of the first combustion chamber, and (ii) a plurality of second fuel nozzles, a corresponding second fuel nozzle of the plurality of second fuel nozzles being arranged to provide the second fuel flow to a corresponding one of the plurality of second combustion chambers.

[0095] The gas turbine engine according to any of the preceding clauses, wherein the steam generation system is arranged downstream of an exhaust section of the gas turbine engine and generates steam in a boiler using exhaust gas from the exhaust section.

[0096] The gas turbine engine according to any of the preceding clauses, wherein the steam generation system further comprises a condenser and a separator, wherein the condenser and the separator utilize the exhaust gas to recover water from the exhaust gas and replenish water in the boiler.

[0097] The gas turbine engine according to any preceding clause, wherein the outlet of the second combustor liner is in fluid communication with the secondary combustion zone of the first combustion chamber.

[0098] A gas turbine engine according to any preceding clause, wherein the outlet is arranged to extend through at least one dilution opening in one of the first combustor outer liner or the first combustor inner liner.

[0099] A gas turbine engine according to any of the preceding items, wherein the combustor outer shell and the combustor inner shell extend annularly around the combustor centerline axis, the first combustor is an annular burner extending annularly around the combustor centerline axis, and the second combustor is a partial burner that does not extend annularly around the combustor centerline axis.

[0100] A gas turbine engine according to any preceding clause, wherein the second combustor is a partial annular combustor extending partially annularly about the combustor centreline axis.

[0101] The gas turbine engine according to any preceding clause, wherein the second combustor casing surrounds the second combustor liner.

[0102] The gas turbine engine according to any preceding clause, wherein the second combustor liner is a trapped vortex combustor liner.

[0103] The gas turbine engine of any preceding clause, wherein the second combustor liner includes a plurality of openings therethrough providing fluid communication between the second combustor plenum and the second combustion chamber.

[0104] A gas turbine engine according to any of the preceding items, wherein, in a non-idle operating state and a power boost operating state of the gas turbine engine, and when the steam generating system is capable of operating to generate the steam at a first steam generation level, (i) the steam flow is provided by the steam generating system to the second burner air chamber, to the second combustion chamber via the plurality of openings, and to the secondary combustion zone of the first combustion chamber, (ii) the compressor bleed air system is prohibited from providing compressor bleed air to the second burner air chamber, and (iii) the second fuel nozzle is prohibited from providing the second fuel flow to the second combustion chamber.

[0105] A gas turbine engine according to any of the preceding items, wherein, in the non-idle operating state and the power boost operating state of the gas turbine engine, and when the steam generating system is capable of operating to generate steam at a second steam generation level lower than the first steam generation level, (iv) the steam is provided by the steam generating system to the second burner air chamber and to the second combustion chamber via the plurality of openings, (v) the compressor bleed air system is enabled and provides the compressor bleed air flow to the second burner air chamber and to the second combustion chamber via the plurality of openings, (vi) the second fuel nozzle is enabled and provides the second fuel flow to the second combustion chamber, the fuel, the compressor bleed air and the steam are mixed in the second combustion chamber to form a steam-air-fuel mixture, and (vii) the steam-air-fuel mixture is ignited and combusted in the second combustion chamber to produce combustion products, which are provided to the secondary combustion zone of the first combustion chamber via the outlet.

[0106] A gas turbine engine according to any of the preceding items, wherein, in the non-idle operating state and the power enhancement operating state of the gas turbine engine, and when the steam generating system does not generate the steam, (viii) the steam generating system is prohibited from providing steam to the second combustor air chamber, (ix) the compressor bleed air system is enabled to provide the compressor bleed air flow to the second combustor air chamber, and the compressor bleed air flows into the second combustion chamber through the multiple openings, (x) the second fuel nozzle is enabled and provides the second fuel flow to the second combustion chamber to produce a fuel-air mixture in the second combustion chamber, (xi) the fuel-air mixture is ignited in the second combustion chamber to produce second combustion products, and (xii) the second combustion products are provided from the second combustion chamber to the secondary combustion zone of the first combustion chamber via the outlet.

[0107] A gas turbine engine as claimed in any preceding clause, wherein the plurality of openings are arranged to generate a swirling flow of the compressor bleed air within the second combustion chamber.

[0108] The gas turbine engine of any preceding clause, wherein the second combustor liner defines a trapped vortex combustor liner, and the swirling flow is a trapped vortex flow of the compressor bleed air within the trapped vortex combustor liner.

[0109] The gas turbine engine of any preceding clause, wherein the second combustor liner defines an axial flow second combustion chamber, and wherein the axial flow second combustor liner includes a dome having a plurality of dome openings extending therethrough.

[0110] The gas turbine engine according to any preceding clause, wherein the plurality of dome openings are arranged at a tangential angle and the second fuel nozzles are arranged at the tangential angle.

[0111] A gas turbine engine comprising: a compressor section having a compressor bleed air system connected thereto; a steam turbine; a combustion section comprising: a combustor outer shell and a combustor inner shell, the combustor outer shell and the combustor inner shell defining a first combustor pressure plenum therein; a first combustor, the first combustor being arranged within the combustor outer shell and the combustor inner shell and having (a) a first combustor outer liner and a first combustor inner liner, the first combustor outer liner and the first combustor inner liner defining a first combustion chamber, the first combustion chamber having a primary combustion zone and a secondary combustion zone located downstream of the primary combustion zone a first-stage combustion zone, and (b) at least one swirler assembly, an outer flow passage defined between the combustor outer shell and the first combustor outer liner, an inner flow passage defined between the combustor inner shell and the first combustor inner liner, the first combustor being arranged to receive a compressed air flow from the compressor section and provide the compressed air flow to the first combustion chamber; and a second combustor disposed in one of the outer flow passage or the inner flow passage, the second combustor having a second combustor casing and a second combustor liner, the second combustor liner defining a second combustion chamber therein, the first combustor A second combustor air chamber is defined between a second combustor casing and a second combustor liner, the second combustor liner having an outlet, the second combustor casing being fluidly connected to the compressor bleed air system to receive a compressor bleed air flow from the compressor bleed air system into the second combustor air chamber, and the outlet providing fluid communication between the second combustion chamber and the first combustion chamber of the first combustor; a first fuel nozzle assembly, the first fuel nozzle assembly being arranged to provide a first fuel flow to the at least one swirler assembly of the first combustor; a second fuel nozzle, the second fuel nozzle being arranged to provide a second fuel flow to the second combustion chamber; and a steam generating system, the steam generating system being arranged to provide a steam flow to the second combustor air chamber and the steam turbine, wherein the second combustor includes (i) a plurality of second combustor liners, each of the plurality of second combustor liners defining a corresponding second combustion chamber, each second combustion chamber having a corresponding outlet fluidly connected to the secondary combustion zone of the first combustion chamber, and (ii) a plurality of second fuel nozzles, a corresponding second fuel nozzle of the plurality of second fuel nozzles being arranged to provide the second fuel flow to a corresponding one of the plurality of second combustion chambers.

[0112] A gas turbine engine according to any preceding clause, wherein the compressor bleed air system includes a bleed air valve that controls the flow of the compressor bleed air to the second combustor plenum, and the steam generation system includes a steam control valve that controls the flow of steam to the second combustor plenum.

[0113] The gas turbine engine according to any of the preceding clauses, wherein the steam generation system is arranged downstream of an exhaust section of the gas turbine engine and generates steam in a boiler using exhaust gas from the exhaust section.

[0114] The gas turbine engine according to any of the preceding clauses, wherein the steam generation system further comprises a condenser and a separator, wherein the condenser and the separator utilize the exhaust gas to recover water from the exhaust gas and replenish water in the boiler.

[0115] The gas turbine engine according to any preceding clause, wherein the outlet of the second combustor liner is in fluid communication with the secondary combustion zone of the first combustion chamber.

[0116] A gas turbine engine according to any preceding clause, wherein the outlet is arranged to extend through at least one dilution opening in one of the first combustor outer liner or the first combustor inner liner.

[0117] A gas turbine engine according to any of the preceding items, wherein the combustor outer shell and the combustor inner shell extend annularly around the combustor centerline axis, the first combustor is an annular burner extending annularly around the combustor centerline axis, and the second combustor is a partial burner that does not extend annularly around the combustor centerline axis.

[0118] A gas turbine engine according to any preceding clause, wherein the second combustor is a partial annular combustor extending partially annularly about the combustor centreline axis.

[0119] The gas turbine engine according to any preceding clause, wherein the second combustor casing surrounds the second combustor liner.

[0120] The gas turbine engine according to any preceding clause, wherein the second combustor liner is a trapped vortex combustor liner.

[0121] The gas turbine engine of any preceding clause, wherein the second combustor liner includes a plurality of openings therethrough providing fluid communication between the second combustor plenum and the second combustion chamber.

[0122] A gas turbine engine according to any of the preceding items, wherein, in a non-idle operating state and a power boost operating state of the gas turbine engine, and when the steam generating system is capable of operating to generate steam at a first steam generation level, (i) the steam flow is provided by the steam generating system to the second burner air chamber, to the second combustion chamber via the plurality of openings, and to the secondary combustion zone of the first combustion chamber, (ii) the compressor bleed air system is prohibited from providing compressor bleed air to the second burner air chamber, and (iii) the second fuel nozzle is prohibited from providing the second fuel flow to the second combustion chamber.

[0123] A gas turbine engine according to any of the preceding items, wherein, in the non-idle operating state and the power boost operating state of the gas turbine engine, and when the steam generating system is capable of operating to generate steam at a second steam generation level lower than the first steam generation level, (iv) the steam is provided by the steam generating system to the second burner air chamber and to the second combustion chamber via the plurality of openings, (v) the compressor bleed air system is enabled and provides the compressor bleed air flow to the second burner air chamber and to the second combustion chamber via the plurality of openings, (vi) the second fuel nozzle is enabled and provides the second fuel flow to the second combustion chamber, the fuel, the compressor bleed air and the steam are mixed in the second combustion chamber to form a steam-air-fuel mixture, and (vii) the steam-air-fuel mixture is ignited and combusted in the second combustion chamber to produce combustion products, which are provided to the secondary combustion zone of the first combustion chamber via the outlet.

[0124] A gas turbine engine according to any of the preceding items, wherein, in the non-idle operating state and the power enhancement operating state of the gas turbine engine, and when the steam generating system does not generate steam, (viii) the steam generating system is prohibited from providing steam to the second combustor air chamber, (ix) the compressor bleed air system is enabled to provide the compressor bleed air flow to the second combustor air chamber, and the compressor bleed air flows into the second combustion chamber through the multiple openings, (x) the second fuel nozzle is enabled and provides the second fuel flow to the second combustion chamber to produce a fuel-air mixture in the second combustion chamber, (xi) the fuel-air mixture is ignited in the second combustion chamber to produce second combustion products, and (xii) the second combustion products are provided from the second combustion chamber to the secondary combustion zone of the first combustion chamber via the outlet.

[0125] A gas turbine engine as claimed in any preceding clause, wherein the plurality of openings are arranged to generate a swirling flow of compressor bleed air within the second combustion chamber.

[0126] The gas turbine engine of any preceding clause, wherein the second combustor liner defines a trapped vortex combustor liner, and the swirling flow is a trapped vortex flow of the compressor bleed air within the trapped vortex combustor liner.

[0127] The gas turbine engine of any preceding clause, wherein the second combustor liner defines an axial flow second combustion chamber, and wherein the axial flow second combustor liner includes a dome having a plurality of dome openings extending therethrough.

[0128] The gas turbine engine according to any preceding clause, wherein the plurality of dome openings are arranged at a tangential angle and the second fuel nozzles are arranged at the tangential angle.

[0129] A method of operating a gas turbine engine, the gas turbine engine comprising: a compressor section having a compressor bleed air system connected thereto; a steam turbine; a combustion section comprising: a combustor outer casing and a combustor inner casing, the combustor outer casing and the combustor inner casing defining a first combustor pressure plenum therein; a first combustor, the first combustor being disposed within the combustor outer casing and the combustor inner casing and having (a) a first combustor outer liner and a first combustor inner liner, the first combustor outer liner and the first combustor inner liner defining a first combustion chamber, the first combustion chamber having a primary combustion zone and a secondary combustion zone downstream of the primary combustion zone, and (b) at least one swirler assembly, an outer flow passage being defined between the combustor outer casing and the first combustor outer liner, an inner flow passage being defined between the combustor inner casing and the first combustor inner liner, the first combustor being arranged to receive a compressed air flow from the compressor section and to provide the compressed air flow to the first combustion chamber; and a second combustor arranged in one of the outer flow passage or the inner flow passage, the second combustor having a second combustor casing and a second combustor liner, the second combustor liner defining a second combustion chamber therein, a second combustor air chamber defined between the second combustor casing and the second combustor liner, the second combustor liner having an outlet, the second combustor casing being fluidly connected to the compressor bleed air system to receive a compressor bleed air flow from the compressor bleed air system into the second combustor air chamber, and the outlet providing fluid communication between the second combustion chamber and the first combustion chamber of the first combustor; a first fuel nozzle assembly, the first fuel nozzle assembly being arranged to provide a first fuel flow to the at least one swirler assembly of the first combustor; a second fuel nozzle, the second fuel nozzle being arranged to provide a second fuel flow to the second combustion chamber; and a steam generating system, the steam generating system being arranged to provide a steam flow to the second combustor air chamber and the steam turbine. The method includes operating the gas turbine engine in a non-idle operating state and a power boost operating state of the gas turbine engine, and when the steam generating system is capable of operating to generate steam at a first steam generation level, (i) the steam flow is provided by the steam generating system to the second burner air chamber, to the second combustion chamber via the plurality of openings, and to the secondary combustion zone of the first combustion chamber, (ii) the compressor bleed air system is prohibited from providing compressor bleed air to the second burner air chamber, and (iii) the second fuel nozzle is prohibited from providing the second fuel flow to the second combustion chamber.

[0130] A method according to the preceding clause, wherein, in a case where the steam generating system is operable to generate steam at a second steam generating level lower than the first steam generating level, (iv) steam is provided by the steam generating system to the second combustor air chamber and provided to the second combustion chamber via the plurality of openings, (v) the compressor bleed air system is enabled and the compressor bleed air flow is provided to the second combustor air chamber and provided to the second combustion chamber via the plurality of openings, (vi) the second fuel nozzle is enabled and the second fuel flow is provided to the second combustion chamber, the fuel, the compressor bleed air and the steam are mixed in the second combustion chamber to form a steam-air-fuel mixture, and (vii) the steam-air-fuel mixture is ignited and combusted in the second combustion chamber to produce combustion products, which are provided to the secondary combustion zone of the first combustion chamber via the outlet.

[0131] A method according to any of the preceding items, wherein, when the steam generating system does not generate steam, (viii) the steam generating system is prohibited from providing steam to the second burner air chamber, (ix) the compressor bleed air system is enabled to provide the compressor bleed air flow to the second burner air chamber, and the compressor bleed air flows into the second combustion chamber through the multiple openings, (x) the second fuel nozzle is enabled and the second fuel flow is provided to the second combustion chamber to produce a fuel-air mixture in the second combustion chamber, (xi) the fuel-air mixture is ignited in the second combustion chamber to produce second combustion products, and (xii) the second combustion products are provided from the second combustion chamber to the secondary combustion zone of the first combustion chamber via the outlet.

[0132] A method according to any preceding clause, wherein the outlet is arranged to direct a flow of steam or a flow of combustion products resulting therefrom in a downstream direction within the secondary combustion zone of the first combustion chamber.

[0133] A method according to any preceding clause, wherein the compressor bleed air system includes a bleed air valve controlling the flow of the compressor bleed air to the second combustor plenum, and the steam generation system includes a steam control valve controlling the flow of steam to the second combustor plenum.

[0134] A method according to any of the preceding items, wherein the second combustor comprises (i) a plurality of second combustor liners, each of the plurality of second combustor liners defining a corresponding second combustion chamber, each second combustion chamber having a corresponding outlet fluidically connected to the secondary combustion zone of the first combustion chamber, and (ii) a plurality of second fuel nozzles, a corresponding second fuel nozzle of the plurality of second fuel nozzles being arranged to provide the second fuel flow to a corresponding one of the plurality of second combustion chambers.

[0135] A method according to any preceding clause, wherein the steam generation system is arranged downstream of an exhaust section of the gas turbine engine and generates steam in a boiler using exhaust gas from the exhaust section.

[0136] The method according to any of the preceding items, wherein the steam generation system further comprises a condenser and a separator, wherein the condenser and the separator utilize the exhaust gas to recover water from the exhaust gas and replenish water in the boiler.

[0137] A method according to any preceding clause, wherein the outlet of the second combustor liner is in fluid communication with the secondary combustion zone of the first combustion chamber.

[0138] A method according to any preceding clause, wherein the outlet is arranged to extend through at least one dilution opening in one of the first combustor outer liner or the first combustor inner liner.

[0139] A method according to any of the preceding items, wherein the burner outer shell and the burner inner shell extend annularly around the burner centerline axis, the first burner is an annular burner extending annularly around the burner centerline axis, and the second burner is a local burner that does not extend annularly around the burner centerline axis.

[0140] A method according to any preceding clause, wherein the second burner is a part annular burner extending partially annularly about the burner centreline axis.

[0141] A method as in any preceding clause, wherein the second combustor casing surrounds the second combustor liner.

[0142] A method as in any preceding clause, wherein the second combustor liner is a trapped vortex combustor liner.

[0143] A method as in any preceding clause, wherein the second combustor liner includes a plurality of openings therethrough providing fluid communication between the second combustor plenum and the second combustion chamber.

[0144] A method as in any preceding clause, wherein the plurality of openings are arranged to generate a swirling flow of the compressor bleed air within the second combustion chamber.

[0145] A method as in any preceding clause, wherein the second combustor liner defines a trapped vortex combustor liner, and the swirling flow is a trapped vortex flow of the compressor bleed air within the trapped vortex combustor liner.

[0146] The method of any preceding clause, wherein the second combustor liner defines an axial flow second combustion chamber, and wherein the axial flow second combustor liner includes a dome having a plurality of dome openings extending therethrough.

[0147] A method as in any preceding clause, wherein the plurality of dome openings are arranged at a tangential angle, and the second fuel nozzle is arranged at the tangential angle.

[0148] A method for operating a combustor of a gas turbine engine in a non-idle operating state and a power boost operating state of the gas turbine engine, the gas turbine engine comprising: a combustion section having a first combustor and a second combustor fluidically connected to a combustion chamber of the first combustor; and a steam generating system, the method comprising, when the steam generating system is capable of operating to generate steam at a first steam generation level, (i) providing a steam flow from the steam generating system to a second combustor air chamber of the second combustor and to the second combustion chamber of the second combustor via a plurality of openings in the second combustor liner, and to a secondary combustion zone of the combustion chamber of the first combustor, (ii) prohibiting a compressor bleed air system from providing compressor bleed air to the second combustor air chamber, and (iii) prohibiting a second fuel nozzle of the second combustor from providing a second fuel flow to the second combustion chamber.

[0149] A method according to the preceding clause, wherein the method further includes, when the steam generating system is capable of operating to generate steam at a second steam generation level that is lower than the first steam generation level, (iv) providing the steam by the steam generating system to the second combustor air chamber and to the second combustion chamber via the plurality of openings, (v) enabling the compressor bleed air system to provide the compressor bleed air flow to the second combustor air chamber and to the second combustion chamber via the plurality of openings, (vi) enabling the second fuel nozzle to provide the second fuel flow to the second combustion chamber, the fuel, the compressor bleed air and the steam mixing in the second combustion chamber to form a steam-air-fuel mixture, and (vii) igniting and combusting the steam-air-fuel mixture in the second combustion chamber to produce combustion products, which are provided to the secondary combustion zone of the first combustion chamber via the outlet.

[0150] A method according to any of the preceding items, wherein the method further includes, when the steam generating system does not generate steam, (viii) prohibiting the steam generating system from providing steam to the second burner air chamber, (ix) enabling the compressor bleed air system to provide the compressor bleed air flow to the second burner air chamber, and the compressor bleed air flows into the second combustion chamber through the multiple openings, (x) enabling the second fuel nozzle to provide the second fuel flow to the second combustion chamber to produce a fuel-air mixture in the second combustion chamber, (xi) igniting and burning the fuel-air mixture in the second combustion chamber to produce second combustion products, and (xii) providing the second combustion products from the second combustion chamber to the secondary combustion zone of the first combustion chamber via the outlet.

[0151] Although the above description is directed to some exemplary embodiments of the present disclosure, it will be clear to those skilled in the art that other changes and modifications may be made without departing from the present disclosure. In addition, even if not explicitly stated above, the features described in conjunction with one embodiment of the present disclosure may also be used in conjunction with other embodiments.

Claims

1. A gas turbine engine, characterized in that: include: a compressor section having a compressor bleed air system connected thereto; steam turbines; Combustion section, including: a burner outer shell and a burner inner shell, said burner outer shell and said burner inner shell defining a first burner plenum therein; a first combustor disposed within the combustor outer casing and the combustor inner casing and having (a) a first combustor outer liner and a first combustor inner liner, the first combustor outer liner and the first combustor inner liner defining a first combustion chamber, the first combustion chamber having a primary combustion zone and a secondary combustion zone downstream of the primary combustion zone, and (b) at least one swirler assembly, an outer flow passage defined between the combustor outer casing and the first combustor outer liner, and an inner flow passage defined between the combustor inner casing and the first combustor inner liner, the first combustor being arranged to receive a flow of compressed air from the compressor section and to provide the flow of compressed air into the first combustion chamber; and a second combustor disposed within one of the outer flow passage or the inner flow passage, the second combustor having a second combustor casing and a second combustor liner, the second combustor liner defining a second combustion chamber therein, a second combustor plenum defined between the second combustor casing and the second combustor liner, the second combustor liner having an outlet, the second combustor casing being in fluid communication with the compressor bleed air system to receive a compressor bleed air flow from the compressor bleed air system into the second combustor plenum, and the outlet providing fluid communication between the second combustion chamber and the first combustion chamber of the first combustor; a first fuel nozzle arranged to provide a first fuel flow to the at least one swirler assembly of the first combustor; a second fuel nozzle arranged to provide a second fuel flow to the second combustion chamber; and A steam generating system is arranged to provide a flow of steam to the second combustor plenum and the steam turbine.

2. The gas turbine engine according to claim 1, wherein: Wherein the outlet is arranged to direct a flow of steam or a flow of combustion products resulting therefrom in a downstream direction within the secondary combustion zone of the first combustion chamber.

3. The gas turbine engine according to claim 1, wherein: The compressor bleed air system includes a bleed air valve that controls the flow of the compressor bleed air to the second combustor plenum, and the steam generation system includes a steam control valve that controls the flow of steam to the second combustor plenum.

4. The gas turbine engine according to claim 1, wherein: wherein the second combustor comprises (i) a plurality of second combustor liners, each of the plurality of second combustor liners defining a corresponding second combustion chamber, each second combustion chamber having a corresponding outlet fluidically connected to the secondary combustion zone of the first combustion chamber, and (ii) a plurality of second fuel nozzles, a corresponding second fuel nozzle of the plurality of second fuel nozzles being arranged to provide the second fuel flow to a corresponding one of the plurality of second combustion chambers.

5. The gas turbine engine according to claim 1, wherein The steam generation system is arranged downstream of an exhaust section of the gas turbine engine and generates steam in a boiler using exhaust gas from the exhaust section.

6. The gas turbine engine according to claim 5, characterized in that The steam generation system further includes a condenser and a separator, and the condenser and the separator utilize the exhaust gas to recover water from the exhaust gas and replenish the water in the boiler.

7. The gas turbine engine according to claim 1, wherein: The outlet of the second combustor liner is in fluid communication with the secondary combustion zone of the first combustion chamber.

8. The gas turbine engine according to claim 7, characterized in that Wherein the outlet is arranged to extend through at least one dilution opening in one of the first combustor outer liner or the first combustor inner liner.

9. The gas turbine engine according to claim 1, wherein: The burner outer shell and the burner inner shell extend annularly around the burner centerline axis, the first burner is an annular burner extending annularly around the burner centerline axis, and the second burner is a partial burner not extending annularly around the burner centerline axis.

10. The gas turbine engine according to claim 9, characterized in that The second burner is a partial annular burner, which extends partially annularly around the burner centerline axis.

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