Exhaust flow components for gas turbine engines

Abstract

A gas turbine engine includes a turbine defining a core flow therethrough during operation. A first heat exchange assembly is in fluid communication with the turbine for receiving a first exhaust flow from the turbine. A second heat exchange assembly is in fluid communication with the turbine for receiving a second exhaust flow from the turbine. A first flow outlet is provided for receiving the first exhaust flow from the first heat exchange assembly and for supplying the first exhaust flow to a first airflow assembly. A second flow outlet is provided for receiving the second exhaust flow and for supplying the second exhaust flow from the second heat exchange assembly to a second airflow assembly.

Description

Technical Field

[0001] This topic broadly relates to bleed flow components used in gas turbine engines. Background Technology

[0002] A gas turbine engine typically consists of a fan and a turbine. The turbine generally comprises an inlet, one or more compressors, a combustor, and at least one turbine. The compressor compresses air, which is then directed to the combustion chamber where it is mixed with fuel. The mixture is then ignited to generate hot combustion gases. These combustion gases are directed to the turbine, which extracts energy from them to power the compressor and to generate useful work to propel an aircraft in flight and / or power loads such as generators. Exhaust stream assemblies are typically available to discharge the airflow from the turbine for use in various operations. The use of exhaust streams can affect turbine operation, necessitating that exhaust streams be managed in consideration of both exhaust stream and turbine requirements. Attached Figure Description

[0003] The complete and feasible disclosure of this disclosure, including its best mode, is set forth in the specification with reference to the accompanying drawings, for those skilled in the art, wherein:

[0004] Figure 1 This is a schematic cross-sectional view of an exemplary gas turbine engine according to various embodiments of this subject matter.

[0005] Figure 2 This is a schematic diagram of an emission assembly for a gas turbine engine according to an exemplary embodiment of the present disclosure.

[0006] Figure 3 This is a schematic diagram of an emission assembly for a gas turbine engine according to another exemplary embodiment of the present disclosure.

[0007] Figure 4 This is a schematic diagram of an emission assembly for a gas turbine engine according to yet another exemplary embodiment of the present disclosure.

[0008] Figure 5 This is a schematic diagram of an emission assembly for a gas turbine engine according to yet another exemplary embodiment of the present disclosure.

[0009] Figure 6 This is a schematic diagram of an emission assembly for a gas turbine engine according to yet another exemplary embodiment of the present disclosure.

[0010] Figure 7 This is a schematic diagram of an emission assembly for a gas turbine engine according to yet another exemplary embodiment of the present disclosure.

[0011] Figure 8A flowchart of a method for operating a gas turbine engine according to an exemplary aspect of this disclosure is provided.

[0012] Figure 9 A flowchart of a method for operating a gas turbine engine according to another exemplary aspect of this disclosure is provided.

[0013] Figure 10 A flowchart of a method for operating a gas turbine engine according to yet another exemplary aspect of this disclosure is provided.

[0014] Figure 11 A flowchart of a method for operating a gas turbine engine according to yet another exemplary aspect of this disclosure is provided. Detailed Implementation

[0015] Reference will now be made in detail to the present embodiments of this disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerals and letter reference numerals to denote features in the drawings. Similar or analogous reference numerals in the drawings and description have been used to denote similar or analogous portions of this disclosure.

[0016] The term "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as superior or better than other implementations. Furthermore, unless explicitly stated otherwise, all embodiments described herein should be considered exemplary.

[0017] For the purposes described below, the terms “upper,” “lower,” “right,” “left,” “vertical,” “horizontal,” “top,” “bottom,” “lateral,” “longitudinal,” and their derivatives should be associated with the embodiments in which they are oriented in the accompanying drawings. However, it should be understood that various alternative variations may be assumed in the embodiments unless explicitly stated otherwise. It should also be understood that the specific devices shown in the drawings and described in the following description are merely exemplary embodiments of this disclosure. Therefore, the specific dimensions and other physical characteristics associated with the embodiments disclosed herein should not be considered limiting.

[0018] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components.

[0019] The terms "front" and "rear" refer to relative positions within a gas turbine engine or vehicle, and specifically to the normal operating posture of the gas turbine engine or vehicle. For example, in the case of a gas turbine engine, "front" refers to the position closer to the engine inlet, while "rear" refers to the position closer to the engine nozzle or exhaust port.

[0020] The terms "upstream" and "downstream" refer to the relative directions of fluid flow within a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction from which the fluid flows.

[0021] Unless otherwise specified herein, the terms “connection,” “fixed,” “attached to,” etc., refer to both direct connection, fixation, or attachment, and indirect connection, fixation, or attachment via one or more intermediate components or features.

[0022] Unless the context clearly indicates otherwise, the singular forms “a,” “a,” and “the” include plural references.

[0023] In the context of, for example, “at least one of A, B and C” or “at least one of A, B or C”, the term “at least one” means only A, only B, only C, or any combination of A, B and C.

[0024] As used throughout the specification and claims, approximate language is applied to modify any quantitative expression that allows for variation without altering its underlying function. Therefore, values ​​modified by terms such as “about,” “approximate,” and “substantially” are not limited to specified exact values. In at least some cases, approximate language may correspond to the precision of the instrument used to measure the value, or the precision of the method or machine used to construct or manufacture the component and / or system. For example, approximate language may refer to margins of 1%, 2%, 4%, 10%, 15%, or 20%. These approximate margins may apply to a single value, to either end of a range defining a numerical value, or to margins between two ends, and / or between the ends.

[0025] Throughout this specification and claims, scope limitations are combined and interchanged, and unless the context or language otherwise indicates otherwise, such scopes are identified and include all subscopes contained herein. For example, all scopes disclosed herein include endpoints, and endpoints may be combined independently of each other.

[0026] In some exemplary embodiments, the operating temperature of the airflow through the third flow can be below the engine's maximum compressor discharge temperature, and more specifically, below 350 degrees Fahrenheit (e.g., below 300 degrees Fahrenheit, below 250 degrees Fahrenheit, below 200 degrees Fahrenheit, and at least as high as ambient temperature). In some exemplary embodiments, these operating temperatures can facilitate heat transfer to or from the airflow through the third flow and the separated fluid flow. Furthermore, in some exemplary embodiments, under takeoff conditions, or more specifically, under operating conditions of sea-level rated takeoff power, static flight speed, and an ambient temperature of 86 degrees Fahrenheit, the airflow through the third flow can contribute less than 50% (and at least, for example, 2%) of the total engine thrust.

[0027] Furthermore, in some exemplary embodiments, the aforementioned exemplary percentage contribution of the third flow's airflow aspects (e.g., airflow, mixing, or exhaust properties) to the total thrust can be passively adjusted during engine operation or purposefully modified by using engine control features (such as fuel flow, motor power, variable stator, variable inlet guide vanes, valves, variable exhaust geometry, or fluid characteristics) to adjust or optimize overall system performance under a wide range of potential operating conditions.

[0028] The term “turbine” or “turbomachinery” refers to a machine that includes one or more compressors, a heating section (e.g., a combustion section), and one or more turbines that together generate torque output.

[0029] The term "combustion engine" refers to a turbomachinery component used to generate torque output through the force exerted by the combustion reaction. Combustion engines can be described independently of other propulsion or power generation components, such as electric or fuel cell components.

[0030] The term "gas turbine engine" refers to an engine that has a turbine as its power source, in whole or in part. Examples of gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, and hybrid electric versions of one or more of these engines.

[0031] When used with compressors, turbines, shafts, or spool components, unless otherwise specified, the terms “low” and “high,” or their respective comparatives (e.g., “lower” and “higher,” where applicable), refer to relative speeds within the engine. For example, “low-speed turbine” or “low-turbine” defines a component constructed to operate at a rotational speed (such as the maximum permissible rotational speed) lower than that of a “high-speed turbine” or “high-turbine” at the engine.

[0032] Systems and methods are provided for operating an exhaust assembly of a gas turbine engine. The exhaust assembly includes at least one flow tap configured to receive an exhaust flow from the turbine's core flow. This exhaust flow is then redirected to various exhaust flow components of the gas turbine engine or an aircraft equipped with a gas turbine engine, such as anti-icing and / or cabin environmental control systems. This disclosure provides systems and methods to avoid engine operability problems, such as stall and surge, while providing a relatively large volume of exhaust flow. The inventors of this disclosure have discovered that the high percentage of exhaust flow relative to the core flow required to remove some exhaust flow components benefits from the construction and operation of reducing and / or recapturing energy drawn from a given flow tap of the turbine. For example, systems and methods can be provided to achieve high exhaust flow and preferred turbine operation by recapturing energy from the exhaust flow for supplementing turbine power. Additionally or alternatively, the exhaust flow from the compressor section can be supplemented by a post-combustion exhaust flow from downstream of the turbine's combustion section.

[0033] Additionally, the inventors of this disclosure have further discovered that downstream emission flow components can benefit from such systems and methods for recapturing emission flow energy as mentioned above. For example, in energy recapturing operation, pressure and temperature can be reduced to safe levels, thereby providing emission flow resources to components that would otherwise lack sufficient resilience to safely handle these emission flow resources. Therefore, systems and methods for adapting emission flow resources to the requirements of various emission flow components would be useful. Additionally, various emission flow components can be configured to operate under a wider range of conditions and / or with lighter or more cost-effective structures to utilize the treatment of these emission flow resources provided herein.

[0034] For example, in one exemplary aspect of this disclosure, a gas turbine engine is provided having a turbine, an exhaust assembly, and an outlet. The exhaust assembly may be configured to receive an exhaust stream from the turbine and may include an exhaust stream machine (e.g., a bleed air turbine) having a machine outlet and configured to drive a machine load (e.g., an accessory gearbox, motor, etc.) to capture energy in the exhaust stream. Further, the machine outlet is configured to receive an exhaust stream during operation of the gas turbine engine, the mass flow rate of which is equal to at least twelve percent (12%) of the core mass flow rate through the turbine's core flow during, for example, aircraft wing de-icing operations. In this way, the gas turbine engine can be designed to provide a relatively large volume of airflow to the aircraft for, for example, wing de-icing operations, while minimizing the energy loss associated with such a large volume of engine bleed air.

[0035] In another exemplary aspect of this disclosure, a gas turbine engine is provided, comprising a turbine, a first heat exchange assembly, a second heat exchange assembly, a first aircraft flow assembly, and a second aircraft flow assembly. The first heat exchange assembly may be configured to receive a first exhaust flow from the turbine and provide the first exhaust flow to the first aircraft flow assembly. The second heat exchange assembly may be configured to receive a second exhaust flow from the turbine and provide the second exhaust flow to the second aircraft flow assembly. In this way, the parallel flow paths provided to the first and second aircraft flow assemblies can facilitate the efficient use of exhaust flows for different aircraft flow assembly tasks. For example, by separating wing de-icing operations from cockpit environmental control operations, various components along the first and second exhaust flows can be sized, shaped, and constructed to meet different design requirements, resulting in narrower, customized component designs for each task.

[0036] In another exemplary aspect of this disclosure, a gas turbine engine is provided having a turbine, a first flow tip, and a second flow tip. The first flow tip is configured to receive a first exhaust flow from upstream of the combustion section, and the second flow tip is configured to receive a second exhaust flow from downstream of the combustion section. A first flow outlet may be provided in fluid communication with the first flow tip, and a second flow outlet may be provided in fluid communication with the second flow tip. The first and second flow outlets are configured to direct the first and second exhaust flows to at least one aircraft flow assembly. In this way, the exhaust flow from upstream of the combustion section can be supplemented by the exhaust flow from downstream of the combustion section, thereby reducing the upstream exhaust volume required to meet the needs of the same aircraft flow assembly.

[0037] In yet another exemplary aspect of this disclosure, a gas turbine engine is provided, having a turbine, a flow tap in fluid communication with the turbine, and an exhaust assembly. The exhaust assembly may include a machine load and an exhaust flow machine. An exhaust regulator may be provided to regulate the exhaust output supplied to the exhaust flow machine by controlling the capture rate of the exhaust flow by the exhaust flow machine. In this way, the exhaust flow regulator can achieve variable control of the resistance to the flow flowing through the exhaust flow machine, thereby allowing tuning of such a system. For example, by adjusting such a system, the demand for increased machine load and / or the demand for higher pressure flows downstream of the exhaust flow machine can be met.

[0038] Referring now to the accompanying drawings, the same number represents the same element throughout the drawings. Figure 1 This is a schematic cross-sectional view of an aviation gas turbine engine according to an exemplary embodiment of the present disclosure. More specifically, for Figure 1In one embodiment, the aviation gas turbine engine is a high-bypass turbofan jet engine 10, referred to herein as "turbofan engine 10". Figure 1 As shown, the turbofan engine 10 defines an axial direction A (extending parallel to a reference longitudinal centerline 12) and a radial direction R. Generally, the turbofan engine 10 includes a fan section 14 and a turbine 16 disposed downstream of the fan section 14.

[0039] The depicted exemplary turbine 16 generally comprises a basic tubular housing 18 defining an annular inlet 20. The housing 18 surrounds, in a series flow relationship: a compressor section including a boost or low-pressure (LP) compressor 22 and a high-pressure (HP) compressor 24; a combustion section 26; a turbine section including a high-pressure (HP) turbine 28 and a low-pressure (LP) turbine 30; and an exhaust nozzle section 32. The compressor section, combustion section 26, turbine section, and exhaust nozzle section 32 together at least partially define a core airflow path 37 through the turbine 16. A high-pressure (HP) shaft or spool 34 (or more precisely, a high-pressure spool assembly as described below) drives the HP turbine 28 to the HP compressor 24. A low-pressure (LP) shaft or spool 36 drives the LP turbine 30 to the LP compressor 22.

[0040] In the depicted embodiment, fan section 14 includes a variable-pitch fan 38 having a plurality of fan blades 40 spaced apart and coupled to a disk 42. As depicted, the fan blades 40 extend generally radially outward from the disk 42. Since the fan blades 40 are operatively coupled to suitable actuating members 44, each fan blade 40 is rotatable relative to the disk 42 about a pitch axis P, and the actuating members 44 are configured to collectively and uniformly change the pitch of the fan blades 40. The fan blades 40, disk 42, and actuating members 44 are rotatable together about a longitudinal centerline 12 via a power gearbox 46 across a LP shaft 36. The power gearbox 46 includes a plurality of gears for progressively reducing the rotational speed of the LP shaft 36 to a more efficient fan rotation speed.

[0041] Still referencing Figure 1 In an exemplary embodiment, the disc 42 is covered by a rotatable front hub 48 having an aerodynamic profile to facilitate airflow through the plurality of fan blades 40. Additionally, the exemplary fan section 14 includes an annular fan housing or outer nacelle 50 circumferentially surrounding at least a portion of the fan 38 and / or turbine 16. The nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced outlet guide blades 52. Furthermore, the nacelle 50 extends over the outer portion of the turbine 16 to define a bypass airflow passage 56 therebetween.

[0042] During operation of the turbofan engine 10, a certain amount of air 58 enters the turbofan engine 10 through the nacelle 50 and / or the associated inlet 60 of the fan section 14. As the certain amount of air 58 passes through the fan blades 40, a first portion of the air 58, as indicated by arrow 62, is directed or directed into the bypass airflow passage 56, and a second portion of the air 58, as indicated by arrow 64, is directed or directed into the LP compressor 22. The ratio between the first portion of air 62 and the second portion of air 64 is commonly referred to as the bypass ratio. The pressure of the second portion of air 64 then increases as it is directed through the HP compressor 24 and into the combustion section 26, where it mixes with fuel and is burned to provide combustion gases 66. Subsequently, the combustion gases 66 are directed through the HP turbine 28 and the LP turbine 30, where a portion of the thermal and / or kinetic energy from the combustion gases 66 is extracted.

[0043] Then, the combustion gases 66 are directed through the exhaust nozzle section 32 of the turbine 16 to provide propulsive thrust. At the same time, the pressure of the first portion of air 62 increases significantly as it is directed through the bypass airflow passage 56 before being exhausted from the fan nozzle exhaust section 68 of the turbofan engine 10, which also provides propulsive thrust.

[0044] Furthermore, as schematically depicted, the exemplary turbofan engine 10 is part of a gas turbine engine, which further includes various accessory systems to assist the operation of the turbofan engine 10 and / or the aircraft including the exemplary turbofan engine 10. For example, as depicted, the exemplary gas turbine engine further includes a fuel delivery system 70, which is operable in conjunction with the combustion section 26 of the turbine 16 of the turbofan engine 10 for supplying fuel to the combustion section 26. The exemplary fuel delivery system 70 may include one or more fuel delivery lines, fuel pumps (not shown), etc. Further, the exemplary gas turbine engine includes an exhaust assembly 72, which will be described in more detail below. It will be understood that the exhaust assembly 72 generally includes an exhaust flow machine 74 configured to receive exhaust gas flow from the turbine 16. The exhaust flow machine 74 is generally operable to convert energy in the exhaust gas flow into mechanical work, which may also be referred to herein as exhaust output. It should also be understood that exhaust output may represent another form of energy, such as electrical potential energy. Therefore, the exhaust flow machine 74 can generate useful work for the turbofan engine. Additionally, the exhaust flow machine 74 can be operable to reduce the energy of the airflow to a desired level, for example, to meet temperature and / or pressure limitations of downstream components. This reduced-energy airflow can be used for various purposes within, for example, the turbofan engine 10, an aircraft (which houses the turbofan engine 10; not shown), etc., and can be used within the aircraft's environmental control system (ECS), for wing anti-icing (WAI; ​​providing anti-icing operation for the aircraft's wings) and / or nacelle anti-icing (NAI; for the engine nacelle, such as... Figure 1 The turbofan engine 10 is used in the nacelle 50 to provide anti-icing operation, etc.

[0045] Now for reference Figure 2 A schematic diagram of an emission assembly 100 for a gas turbine engine according to an exemplary embodiment of the present disclosure is provided. Figure 2 An exemplary gas turbine engine can be used in conjunction with the above reference. Figure 1 The exemplary turbofan engine 10 described is constructed in essentially the same manner. For example, Figure 2 An exemplary gas turbine engine generally includes a turbine 102. A gas turbine engine can be, for example, a turbofan engine, a turboprop engine, a turboshaft engine, a turbojet engine, etc.

[0046] As mentioned above, with Figure 2 The gas turbine engine that uses the emission components 100 together can be used with the above-mentioned Figure 1 The turbofan engine 10 is constructed in a similar manner. Therefore, as described above, Figure 2The turbine 102 generally includes: a compressor section having a low-pressure (LP) compressor 22 and a high-pressure (HP) compressor 24; a combustion section 26; and a turbine section including a high-pressure (HP) turbine 28 and a low-pressure (LP) turbine 30. The compressor section, combustion section 26, and turbine section are arranged in a serial flow sequence. A high-pressure (HP) shaft or spool 34 drives the HP turbine 28 to the HP compressor 24, and a low-pressure (LP) shaft or spool 36 drives the LP turbine 30 to the LP compressor 22.

[0047] Further as Figure 2 As shown, an electric motor 31 can be provided for connection to the gas turbine engine. As depicted, the motor 31 is mechanically connected to the turbine 102 via an LP shaft 36. It should be understood that the motor 31 can be operable to drive the turbine 102 and / or generate electrical energy from the rotation of the turbine 102. For example, the motor 31 can be operable to provide supplemental drive torque to the turbine 102, thereby increasing engine operability margins, such as reducing the chance of engine stall.

[0048] Furthermore, Figure 2 The arrangement includes a fuel delivery system 104 (which can be connected to...) Figure 1 (Similarly constructed to the fuel delivery system 70), it is capable of operating together with the combustion section 26 of the turbine 102 to supply fuel to the combustion section 26 of the turbine 102. For example... Figure 2 As depicted, a fuel source 106 is provided to supply fuel to the combustion section 26 via a fuel delivery pipeline 108.

[0049] like Figure 2 As depicted, the exemplary discharge assembly 100 includes a discharge flow path 110 in fluid communication with the compressor section of the turbine 102. More specifically, as depicted, the discharge flow path 110 may be in fluid communication with the HP compressor 24. Figure 2 In some embodiments, a flow tap 112 is provided from the HP compressor 24 to the exhaust flow path 110. For example, the flow tap 112 may be provided in a specific stage of the HP compressor 24 (such as the fourth stage). It should be understood that the flow tap 112 may draw pressure from various other pressure sources (e.g., turbine 102). In various embodiments, the source of the flow tap 112 is selected based on the temperature and / or pressure requirements of the exhaust assembly 100. It should also be understood that more than one flow tap 112 may be provided, for example, to meet different temperature and / or pressure requirements of the exhaust assembly 100 at a given engine operating output. Figure 2 As shown, the flow tap 112 can be separated from the exhaust flow path 110 via the tap valve 114, for example, to control the flow rate from the turbine 102 input to the exhaust flow path 110. However, in Figure 2In this embodiment, faucet valve 114 is shown configured as a one-way valve to prevent backflow through faucet 112. It should be understood that various other valve arrangements may be provided, for example, in discharge flow path 110, to control the flow rate through discharge flow path 110.

[0050] It should be understood that other pressure sources can be provided in the discharge assembly 100. For example, a secondary pressure source 116 can provide a secondary tap 118 to the discharge flow path 110. In various embodiments, the secondary pressure source 116 can be a relatively low pressure source, such as a compressor discharge source. Figure 2 As shown, a secondary pressure source 116 is connected to a discharge path 110 via a secondary faucet valve 120. The secondary faucet valve 120 may be operable to facilitate flow from the secondary pressure source 116 into the discharge path 110 and to prevent backflow when the pressure in the discharge path 110 exceeds the pressure of the secondary pressure source 116. It should also be understood that the secondary pressure source 116 can also be used as a discharge destination for the discharge path 110. For example, under high-pressure conditions within the discharge path 110, the secondary faucet valve 120 may be actuated to an open position to release pressure to the relatively lower pressure environment of the secondary pressure source 116.

[0051] Various flow rates within turbine 102 and the exhaust assembly can be quantified by mass flow rate (a function of the mass of airflow per unit time). For example, the core mass flow rate can be based on the total mass flowing into turbine 102 per unit time (i.e., corresponding to...). Figure 1 The second part of air 58, indicated by the middle arrow 64, is used to define this core mass flow rate. This core mass flow rate can be used to compare with the mass flow rate within the emission flow path 110 (e.g., the emission mass flow rate representing the total mass flowing into the emission flow path 110 per unit time).

[0052] One or more flow control devices can be used to regulate the mass flow rate within the discharge flow path 110. For example, such as Figure 2 As shown and as described above, the primary valve 114 and / or the secondary primary valve 120 may be operable to control the airflow rate transferred to the exhaust flow path 110 via the turbine 102. Similarly, as... Figure 2 As shown, a discharge inlet valve 122 may also be provided to regulate the discharge mass flow rate within the discharge flow path 110. As illustrated, the discharge inlet valve 122 is operable to control all pressure sources to the discharge flow path 110. However, it should be understood that the pressure sources (such as flow tap 112 and secondary tap 118) can be controlled independently of each other.

[0053] Still referencing Figure 2The discharge flow path 110 is shown to be in fluid communication with the discharge flow machine 124. As shown, the discharge flow path 110 includes a machine inlet 123 and a machine outlet 125, wherein the discharge flow machine 124 facilitates flow from the machine inlet 123 to the machine outlet 125. (Refer to the above...) Figure 1 The discharge flow machine 124 may be operable to utilize energy in a discharge flow (e.g., a discharge flow supplied from the compressor section of turbine 102 through discharge flow path 110 to the discharge flow machine 124). For example, the discharge flow machine 124 may include a turbine configured to convert energy in the discharge flow path 110 into mechanical work; this may be referred to as a bleed-air turbine or bleed-air expander turbine. It should be understood that this process can also be employed to facilitate the reduction of energy in the discharge flow path 110. For example, the discharge flow machine 124 may be used to reduce pressure or temperature below a safe threshold before the discharge flow reaches one or more components or parts susceptible to damage from overpressure or overtemperature.

[0054] The exhaust flow machine 124 may be operable to transfer energy captured from the exhaust flow to one or more components of an aircraft or gas turbine engine. For example, the exhaust flow machine 124 may be configured to transfer torque generated by the turbine rotation from the exhaust flow. Figure 2 The exhaust flow machine 124 shown is configured to transmit torque to the machine load 126 via a machine transmission 128. In various embodiments, the machine load 126 may be drivably mechanically connected to one or more other components. The machine load 126 is configured to receive energy from the exhaust flow machine 124, such as torque via the machine transmission 128. The machine load 126 may be configured as an electrical device, such as a generator. In this example, the machine load 126 may be configured to generate electricity, which may then be stored or used for various motors, such as motor 31, of the gas turbine engine or associated aircraft. In one embodiment, the generator configuration of the machine load 126 is configured to supply electrical energy for driving the gas turbine engine using motor 31, for example, by driving the HP shaft 34 and / or LP shaft 36. In another embodiment, the machine load 126 is mechanically connected to the HP shaft 34 and / or LP shaft 36 and is configured to drive the turbine 102 using energy supplied from the exhaust flow machine 124 via the machine transmission 128.

[0055] Machine load 126 can be used to drive turbine 102, as described above, directly or indirectly. Therefore, machine load 126 can be operable to increase or maintain the operating speed of turbine 102. Thus, machine load 126 can be used to account for deficiencies in the operation of turbine 102, such as those encountered in the exhaust flow path 110 by diverting a portion of the flow through the turbine's core. In this case, the air / fuel ratio and total flow rate within combustion section 26 may decrease, thereby reducing the operability margin and potentially increasing the likelihood of stall, surge, and / or rich blowout. The reduced operability margin can be accounted for or reversed by increasing the engine speed and thus increasing the flow rate through turbine 102. For example, a higher engine speed of turbine 102 can increase the air / fuel ratio and total flow rate to account for the relatively high flow rate diverted to exhaust flow path 110. Therefore, the arrangement of exhaust flow machine 124 and machine load 126 can achieve a relatively high exhaust flow rate obtained from the core flow through turbine 102. For example, the emission flow may have an emission mass flow rate defined as the percentage of mass entering the emission flow path 110 per unit time relative to the core mass flow rate, which is defined as the mass entering the core flow path 37 of the turbine 102 per unit time (see [reference]). Figure 1 The mass flow rate can be at least 10%, 12%, 14%, 16%, 18%, 20%, or 22% of the core mass flow rate. Various further configurations can be provided to achieve this high-emission-flow arrangement, as will be described below.

[0056] Generally speaking, as described above, a relatively high percentage of the exhaust flow relative to the core flow can facilitate the operation of one or more components of an aircraft or gas turbine engine. For example, a relatively high exhaust flow can enable effective de-icing under harsh conditions and / or with a relatively large surface area to be de-iced. Figure 2 As shown, machine outlet 125 from exhaust flow machine 124 can supply various exhaust flow components. As illustrated, the exhaust flow path 110 downstream of exhaust flow machine 124 extends to aircraft flow assembly 130 via aircraft flow valve 132. Aircraft flow assembly 130 may include various components, such as wing anti-icing (WAI) or cockpit environmental control system (ECS) components. It should be understood that aircraft flow valve 132 can be configured to control upstream and downstream operations based on the flow allowed to pass through aircraft flow assembly 130.

[0057] As described above, the aircraft flow assembly 130 refers to any assembly that utilizes the exhaust flow from the exhaust flow path 110. For example, a WAI arrangement may be provided to maintain a safe temperature for the aircraft wing to avoid icing conditions. An ECS may be provided to maintain safe environmental conditions for the aircraft's internal compartments. It should be understood that the aircraft flow assembly 130 may include various sub-assemblies with associated valves and controllers to maintain desired control without necessarily affecting the operation of upstream components of the exhaust assembly 100. Additionally, the aircraft flow assembly 130 may include at least one exhaust vent (not shown) configured to purge excess exhaust flow to the external environment.

[0058] Still referencing Figure 2 The diagram further illustrates machine outlet 125 from machine load 126 to provide an exhaust flow to starter assembly 134 via starter assembly valve 136. Starter assembly 134 may be operable to start a gas turbine engine, such as another gas turbine engine of the same aircraft equipped with the gas turbine engine and associated exhaust assembly 100. Starter assembly valve 136 may be operable to control starter assembly 134, for example, in an engaged start mode. In an embodiment, starter assembly 134 is a pneumatic starter configured to provide flow and subsequent rotation to the gas turbine engine when starter assembly valve 136 is actuated. It should be understood that starter assembly 134 may include various further downstream valves and controllers to utilize the exhaust flow and control it for preferred start-up operations.

[0059] In various embodiments, one or more further or alternative exhaust flow components may be provided downstream of the exhaust flow machine 124. For example, a separation loop with a corresponding separation valve arrangement may be provided for each of the ECS and WAI arrangements. Additionally, various embodiments may provide exhaust flow components upstream of the exhaust flow machine 124. For example, systems that require or benefit from relatively high pressures or temperatures may operate from flows that have not yet passed through the exhaust flow machine 124. In embodiments, a nacelle anti-icing device (NAI) may be configured to divert the flow upstream of the exhaust flow machine 124 to prevent icing conditions in the nacelle 50 (see [link to documentation]). Figure 1 As will be discussed in more detail below, each of the various exhaust flow components and assemblies can be constructed, sized, and shaped in cooperation with the upstream exhaust flow path 110 to effectively utilize the exhaust flow and maintain the efficient operation of the turbine 102.

[0060] Turn now Figure 3 A schematic diagram of an emission assembly 200 for a gas turbine engine according to another exemplary embodiment of the present disclosure is provided. Figure 3 The exemplary gas turbine engine 10 can be used in conjunction with the above reference. Figure 1The exemplary gas turbine engine 10 described is constructed in essentially the same manner, and Figure 3 The exemplary emission component 200 can be coupled with Figure 2 The exemplary emission component 100 is constructed in a similar manner. For example, Figure 3 An exemplary emission assembly 200 generally includes a turbine 102, which is connected to an emission flow path 210 via a flow tap 212.

[0061] Figure 3 The embodiments described herein are similar to Figure 2 The difference in the embodiment depicted is the provision of a first heat exchange assembly 238 and a second heat exchange assembly 240. It should be understood that the first heat exchange assembly 238 and the second heat exchange assembly 240 are merely exemplary and can be provided in various other numbers and arrangements. For example, at least one of the first heat exchange assembly 238 and the second heat exchange assembly 240 can be configured as an air bleed turbine. Figure 3 As shown, a first heat exchange assembly 238 is provided downstream of the flow tap 212 and the secondary pressure source 216. (Refer to the above...) Figure 2 As described, under certain operating conditions, such as when very high compressor discharge is desired and / or when there is a relatively low discharge flow demand from various discharge flow components, the secondary pressure source 216 can operate as a purging or venting section for pressure from the flow tap 212 or another source. The secondary tap valve 220 can be operable to control the flow into or out of the secondary pressure source 216 through the secondary tap 218. See reference... Figure 2 As described, faucet valve 214 can independently control the flow between faucet 212 and discharge flow path 210. In some embodiments, faucet valve 214 may be a check valve configured to prevent backflow.

[0062] In various embodiments, one or more heat exchange assemblies 238, 240 may be configured to perform additional functions. For example, the first heat exchange assembly 238 may be in temperature communication with the WAI and / or NAI system. In one embodiment, the first heat exchange assembly 238 itself is configured as a WAI system. In this exemplary embodiment, the WAI system can be used to transfer heat for de-icing operations, and then allow the cooled exhaust stream to flow to further components as described herein, such as the cabin ECS downstream of the first heat exchange assembly.

[0063] A primary tap 212 and a secondary tap 218 can be connected to the discharge flow path 210 via a discharge inlet valve 222. The discharge inlet valve 222 can operate as a primary controller for the flow through the discharge flow path 210. Downstream of the discharge inlet valve 222, the discharge flow path 210 continues to a first heat exchange assembly 238. The first heat exchange assembly 238 can be operable to control the nature of the discharge flow before it enters the discharge flow machine 224 via machine inlet 223. For example, the first heat exchange assembly 238 can be used to reduce the heat in the discharge flow before it enters the discharge flow machine 224. In various embodiments, the first heat exchange assembly 238 can be an air-to-air heat exchange assembly, can be liquid-cooled, and / or can be evaporatively cooled using a refrigerant cycle.

[0064] like Figure 3 As shown, one or more exhaust flow components can be configured to receive flow prior to any heat exchange operation. For example, some exhaust flow components may benefit from relatively high temperatures. In the illustrated embodiment, a pre-exchange outlet 242 may be provided upstream of the first heat exchange assembly 238 in the exhaust flow path 210. The pre-exchange outlet 242 is operable to supply uncooled exhaust flow to a drag component 244. The drag component 244 may be any component of the exhaust assembly 200, an associated gas turbine engine, and / or an aircraft configured to receive relatively high-temperature flow. For example, the exhaust assembly 200 may be arranged such that multiple drag components 244 are positioned upstream of the first heat exchange assembly 238 and the exhaust flow machine 224, while the remaining exhaust flow components are positioned downstream of the first heat exchange assembly 238 and the exhaust flow machine 224. In this way, the exhaust assembly 200 can be effectively configured to facilitate useful work from the exhaust flow path prior to potentially less efficient heat and / or pressure conversion operations. However, it should be understood that the first heat exchange component 238 can be configured to perform useful work, for example, by transferring the heat energy removed from the exhaust stream to fuel, aircraft internal air, etc.

[0065] In an embodiment, the resistance component 244 may be as referenced above. Figure 2 A brief description of the cabin anti-icing (NAI) assembly. The NAI assembly of drag component 244 can function similarly to a WAI assembly. However, the NAI assembly may be relatively less susceptible to heat or pressure damage and can therefore be configured to utilize a relatively hot and / or high-pressure exhaust flow. Compared to a WAI assembly, the NAI assembly may also require a relatively low mass flow rate to effectively suppress the risk of icing. Thus, the NAI assembly can be an example of a system that benefits from a separation between heat-resistant components 244, unlike the WAI assembly 230. However, it should also be understood that the NAI assembly can also benefit from operations performed by the exhaust flow mechanism 224.

[0066] Downstream of the first heat exchange assembly 238, the exhaust flow machine 224 is operable to, for example, refer to the above-mentioned... Figure 2 The described machine transmission 228 drives the machine load 226. A machine outlet 225 is provided, and the machine outlet 225 is isolated from further components by one or more valves. It should be understood that a pressure relief valve or vent valve may be provided independently, or a pressure relief valve or vent valve may be incorporated into one or more of the depicted valves to ensure safe operation of the discharge flow machine 224. The depicted valves include an aircraft flow valve 232 that connects the discharge flow path 210 to the aircraft flow assembly, and an initiator assembly valve 236 that connects the discharge flow path 210 to the initiator assembly 234. Figure 3 The structure and Figure 2 The difference in the shown configuration lies in the provision of a second heat exchange assembly 240 between the exhaust flow machine 224 and the aircraft flow assembly 230. For example, the aircraft flow assembly 230 may include one or more components requiring further control over flow properties beyond those described by the reference exhaust flow machine 224 and the first heat exchange assembly 238. However, it should be understood that the first heat exchange assembly 238 and the second heat exchange assembly 240 may be provided alternatively. For example, in a configuration where the exhaust flow machine 224 is constructed for efficient and safe high-temperature operation, only the second heat exchange assembly 240 may be provided, while the first heat exchange assembly 238 may be omitted.

[0067] Turn now Figure 4 A schematic diagram of an emission assembly 300 for a gas turbine engine according to another exemplary embodiment of the present disclosure is provided. Figure 4 An exemplary gas turbine engine can be used in conjunction with the above reference. Figure 1 The exemplary gas turbine engine 10 described is constructed in essentially the same manner. For example, Figure 4 An exemplary emission assembly 300 generally includes a turbine 102, which is connected to a first emission flow path 310 via a first flow tap 312. Figure 4 Implementation examples and Figure 2 and Figure 3 The difference in the embodiment is that a second discharge path 311 is further provided.

[0068] Referring to the first discharge path 310, and as mentioned above... Figure 2Generally described, under certain operating conditions, such as when very high compressor discharge is desired and / or when there is a relatively low discharge flow demand from various discharge flow components, the secondary pressure source 316 upstream of the first discharge flow path 310 can operate as a purging or venting point for pressure from the first flow tap 312 or another source. The secondary tap valve 320 can be operable to control the flow into or out of the secondary pressure source 316 through the secondary tap 318. See reference... Figure 2 As described, the first faucet valve 314 can independently control the flow between the first flow faucet 312 and the first discharge flow path 310. In some embodiments, the first faucet valve 314 may be a check valve configured to prevent backflow. It should be further understood that the second discharge flow path 311 may be separated in a similar manner and may further include such purging or venting facilities.

[0069] First-stage faucet 312 and secondary faucet 318 can be connected to the first discharge flow path 310 via discharge inlet valve 322. Discharge inlet valve 322 can operate as a master controller for the flow through the first discharge flow path 310. Downstream of discharge inlet valve 322, the first discharge flow path 310 continues to a first heat exchange assembly 338. The first heat exchange assembly 338 can be operable to control the nature of the discharge flow downstream of one or more discharge flow components. For example, the first heat exchange assembly 338 can be used to reduce the heat in the discharge flow before it enters the first aircraft flow assembly 346 through the first flow outlet 345. The first aircraft flow assembly 346 can be a cabin environmental control assembly. For example, the first aircraft flow assembly 346 can be a cabin ECS generally as described above. In various embodiments, the first heat exchange assembly 338 can be an air-to-air heat exchange assembly that can be liquid-cooled and / or can be evaporative-cooled using a refrigerant cycle. A first heat exchange assembly valve 339 can be provided upstream of the first heat exchange assembly 338 for independent control of the first heat exchange assembly 338.

[0070] Still referencing Figure 4 Both the first discharge path 310 and the second discharge path 311 are depicted as drawing water from the turbine 102. For example, the turbine 102 may be in direct fluid communication with the first flow tip 312 and the second flow tip 313. As depicted, the first flow tip 312 and the second flow tip 313 draw water from the first flow source 302 and the second flow source 303, respectively. These first flow sources 302 and the second flow source 303 may include various components, such as the compressor section of the turbine 102 (see [link to relevant documentation]). Figure 1-3 In this embodiment, the first flow source 302 and the second flow source 303 are the same source. For example, the first flow tap 312 and the second flow tap 313 may each be drawn from the HP compressor 24 (see [link to relevant documentation]). Figure 1-3The flow can be drawn from, or even drawn from, the same stage, such as the fourth stage of the HP compressor 24. However, it should also be understood that the first flow source 302 can represent a pressure source within the turbine that is different from the second flow source 303. For example, the first flow source 302 can originate from components upstream or downstream of the second flow source 303, relative to the core flow through the turbine 102. Although Figure 4 Although not depicted, it should be understood that at least one of the first flow source 302 and the second flow source 303 may be provided outside the turbine 102, for example, as referenced above. Figure 2 The compressor discharge source described.

[0071] Separating the first flow source 302 and the second flow source 303 can facilitate the separation of the first discharge flow path 310 and the second discharge flow path 311. However, it should also be understood that the first discharge flow path 310 and the second discharge flow path 311 can be separated even in examples where the first flow source 302 and the second flow source 303 are the same. For example, one or more valves, such as check valves, can effectively separate the discharge flow between the first discharge flow path 310 and the second discharge flow path 311. In embodiments where the first flow source 302 and the second flow source 303 are the same, the first tap valve 314 and the second tap valve 315 can cooperate to independently and effectively control the downstream flow in the first discharge flow path 310 and the second discharge flow path 311, respectively.

[0072] By separating the first exhaust flow path 310 from the second exhaust flow path 311, the exhaust flow of the entire exhaust assembly 300 can be controlled to maintain a high level of operability of the turbine 102 while providing a high level of exhaust flow for various exhaust flow components. For example, the aforementioned first exhaust flow path 310 can be optimized for controlling the first aircraft flow assembly 346, including optimizing all flow ducts, valves, and the first heat exchanger 338. When the first aircraft flow assembly 346 does not require exhaust flow, the flow can be cut off independently of the second exhaust flow path 311.

[0073] The second exhaust flow path 311 can be operated to control the second vehicle flow assembly 348. By decoupling the control of the first vehicle flow assembly 346 from the control of the second vehicle flow assembly 348, the first exhaust flow path 310 and the second exhaust flow path 311 can each be configured to efficiently handle their respective tasks. For example, the second vehicle flow assembly 348 can be configured as an anti-icing assembly, such as WAI as described above, and may require a relatively higher mass flow rate of the exhaust flow compared to the first vehicle flow assembly 346. Taking this difference into account, various features of the second exhaust flow path 311 can be configured to differ from corresponding features of the first exhaust flow path 310. For example, the second flow source 303 can be a relatively hot and / or higher pressure source compared to the first flow source 302. In one embodiment, the second flow source 303 originates from a more downstream stage of the same component compared to the first flow source 302. In another embodiment, the second flow source 303 originates from a more downstream component compared to the first flow source 302.

[0074] A second heat exchange assembly 340 is provided in the second discharge path 311. The second heat exchange assembly 340 can be constructed differently from the first heat exchange assembly 338. For example, the second heat exchange assembly 340 can be relatively larger than the first heat exchange assembly 338. In embodiments, the first heat exchange assembly 338 uses a different heat exchange mechanism than the second heat exchange assembly 340. For example, the first heat exchange assembly 338 may only employ air-to-air heat exchange, while the second heat exchange assembly employs liquid and / or evaporative heat exchange.

[0075] The second heat exchange assembly 340 can be in fluid communication with the second discharge path 311 in a manner very similar to that of the first heat exchange assembly 338 in fluid communication with the first discharge path 310. For example, as Figure 4 As shown, the second heat exchange assembly 340 is connected to the second discharge path 311 via a second heat exchange valve 341. Further as... Figure 4 As shown, the second heat exchange assembly 340 is connected to the second aircraft flow assembly 348 through the second flow outlet 347.

[0076] Turn now Figure 5 A schematic diagram of an emission assembly 400 for a gas turbine engine according to another exemplary embodiment of the present disclosure is provided. Figure 5 An exemplary gas turbine engine can be used in conjunction with the above reference. Figure 1 The exemplary gas turbine engine 10 described is constructed in a substantially the same manner. Figure 5 The exemplary emission components can be with Figure 4 It is constructed similarly in the example. Figure 5An exemplary emission assembly 400 generally includes a turbine 102 connected to a first emission flow path 410 via a first flow tap 412, and a second emission flow path 411 connected to the turbine 102 via a second flow tap 413. Figure 5 Implementation examples and Figure 4 The difference in this embodiment is that the second exhaust flow path 411 is connected to the turbine 102 via a second flow tap 413 at a location downstream of the combustion section 26. By providing exhaust flow to the second flow tap 413 from this downstream location, exhaust flow from upstream locations (such as the HP compressor 24) can be reduced while maintaining the desired operation of various exhaust flow components.

[0077] The second-flow nozzle 413 is depicted as drawing from the LP turbine 30; however, it may be further provided in various other locations. For example, the second-flow nozzle 413 may be configured as a scoop located downstream of the combustion section 26. In various embodiments, the second-flow nozzle 413 includes a turbine rear frame scoop and / or a turbine center frame scoop.

[0078] With reference to return flow utilization machines 74, 124, 224 (see Figure 1-3 The energy captured from the exhaust stream can be captured from further exhaust stream paths (such as a second exhaust stream path 411) to reduce operability problems caused by exhaust gas from the first exhaust stream path 410 at the compressor section. Therefore, bleed air from downstream of combustion section 26 can be used to reduce emission requirements from upstream of combustion section 26. (Refer to the above...) Figure 4 The first emission path 410 and the second emission path 411 discussed can be configured to manage their flow sources and associated emission flow components according to requirements. For example, Figure 5 The first exhaust flow path 410 depicted draws from the HP compressor 24 and supplies it to the first aircraft flow assembly 446, which may be a cockpit assembly, such as the cockpit ECS as described above. Figure 5 The second exhaust flow path 411 depicted draws from the LP turbine 30 and supplies it to the second aircraft flow assembly 448, which may be an anti-icing assembly, such as WAI as described above.

[0079] Referring to the first discharge path 410, and as mentioned above... Figure 2Generally described, under certain operating conditions, such as when very high compressor discharge is desired and / or when there is a relatively low discharge flow demand from various discharge flow components, the secondary pressure source 416 upstream of the first discharge flow path 410 can operate as a purging or venting point for pressure from the first flow tap 412 or another source. The secondary tap valve 420 can be operable to control the flow into or out of the secondary pressure source 416 through the secondary tap 418. See reference... Figure 2 As described, the first faucet valve 414 can independently control the flow between the first flow faucet 412 and the first discharge flow path 410. In some embodiments, the first faucet valve 414 may be a check valve configured to prevent backflow. A second faucet valve 415 may be provided to control the flow between the second flow faucet 413 and the second discharge flow path 411. It should be further understood that the second discharge flow path 411 may be separated in a similar manner to the first discharge flow path 410 and may further include such purging or venting facilities.

[0080] First-stage faucet 412 and secondary faucet 418 can be connected to the first discharge flow path 410 via discharge inlet valve 422. Discharge inlet valve 422 can operate as a master controller for the flow through the first discharge flow path 410. Downstream of discharge inlet valve 422, the first discharge flow path 410 continues to a first heat exchange assembly 438. The first heat exchange assembly 438 can be operable to control the nature of the discharge flow downstream of one or more discharge flow components. For example, the first heat exchange assembly 438 can be used to reduce the heat in the discharge flow before it enters the first aircraft flow assembly 446 through the first flow outlet 445. The first aircraft flow assembly 446 can be a cabin environmental control assembly. For example, the first aircraft flow assembly 446 can be a cabin ECS generally as described above. In various embodiments, the first heat exchange assembly 438 can be an air-to-air heat exchange assembly that can be liquid-cooled and / or can be evaporative-cooled using a refrigerant cycle. A first heat exchange assembly valve 439 can be provided downstream of the first heat exchange assembly 438 for independent control of the first heat exchange assembly 438.

[0081] Still referencing Figure 5 Both the first exhaust flow path 410 and the second exhaust flow path 411 are depicted as drawing from the turbine 102. As briefly described above, the first flow tip 412 and the second flow tip 413 draw from a first flow source upstream of the core flow in the combustion section 26 and a second flow source downstream of the core flow in the combustion section 26, respectively. Although in Figure 5 Although not depicted, it should be understood that at least one of the first-stream tap 412 and the second-stream tap 314 can be drawn from outside the turbine 102, for example, as referenced above. Figure 2The compressor discharge source is described above. Figure 5 An embodiment may be configured such that one or both of the first discharge flow path 410 and the second discharge flow path 411 include a pressure source from outside the turbine 102, such as a second pressure source 416.

[0082] First source (in) Figure 5 The image shows HP compressor 24) and the second flow source (in... Figure 5 The separation of the first exhaust flow path 410 and the second exhaust flow path 411 (shown as LP turbine 30) facilitates the separation of the first exhaust flow path 410 and the second exhaust flow path 411. By separating the first exhaust flow path 410 from the second exhaust flow path 411, various exhaust flow properties can be defined relative to each other. For example, a relatively higher pressure and temperature flow can be expected from the second exhaust flow path 411, which receives the flow downstream of the combustion section 26. Additionally, different flow treatments may be required due to the combustion products in the second exhaust flow path 411. For example, the second exhaust flow path 411 may be preferably used for systems less sensitive to combustion products, such as WAI and / or NAI operations.

[0083] A first mass flow rate through the first exhaust flow path 410 can be defined relative to a second mass flow rate through the second exhaust flow path 411. In various embodiments, the second mass flow rate can be greater than the first mass flow rate, for example, due to higher pressure from its supply source. In embodiments, the second mass flow rate is at least twice the first mass flow rate. Temperatures can also be defined for the exhaust flows within the respective exhaust flow paths of the first exhaust flow path 410 and the second exhaust flow path 411. For example, the first exhaust flow path 410 can have a relatively lower temperature compared to the temperature of the second exhaust flow path 411. Similar to the mass flow rates, these relative temperature differences can also originate from the respective supply sources, particularly whether downstream or upstream of the combustion section 26.

[0084] Various features of the different first discharge flow paths 410 and second discharge flow paths 411 can be sized, shaped, and constructed to account for this difference in flow rate and temperature. For example, the first discharge flow path 410 can be constructed for relatively low pressure operation as described above. In embodiments, the cross-sectional area of ​​the various flow conduits along the first discharge flow path 410 can be larger than the corresponding cross-sectional area of ​​the various flow conduits along the second discharge flow path 411. In embodiments, the cross-sectional area of ​​the first outlet 445 of the first discharge flow path 410 is at least twice the cross-sectional area of ​​the second outlet 447 of the second discharge flow path 411.

[0085] Still referencing Figure 5The first outlet 445 of the first discharge path 410 is shown as supplying the first aircraft flow assembly 446 downstream of the first heat exchange assembly 438. (See above reference) Figure 4 As described, the first aircraft flow component 446 may be a cabin environmental control system (ECS) that uses a relatively low emission flow. Figure 5 The embodiment separates the first aircraft flow assembly 446 from combustion products by maintaining separation between the first exhaust flow path 410 and the second exhaust flow path 411. As described above, this parallel flow path separation facilitates the use of flow energy from downstream of the combustion section 26 while preventing combustion products from contaminating components such as the cockpit ECS.

[0086] Figure 5 The second flow outlet 447 supplies the second aircraft flow assembly 448 via the second heat exchange assembly 440, which is controlled by the second heat exchange assembly valve 441. (See above reference...) Figure 4 As described, the second flow assembly 448 may be an anti-icing assembly, such as WAI and / or NAI. The second heat exchange assembly 440 may be sized, shaped, and constructed differently from the first heat exchange assembly 438, for example, to account for temperature differences from the respective flow sources. It should be understood that, at least based on the combustion heat energy flowing to the second heat exchange assembly 440, the second heat exchange assembly 440 may need to be significantly larger than the first heat exchange assembly 438. In one embodiment, the second heat exchange assembly 440 has a heat exchange surface area at least twice that of the first heat exchange assembly 438. In another embodiment, the second heat exchange assembly 440 has a heat exchange volume at least twice that of the first heat exchange assembly 438. The second heat exchange assembly 440 may also include the different heat exchange mechanisms described above to adequately cool the flow passing through it. In one embodiment, the second heat exchange assembly 440 is configured to manage combustion products through one or more of the above-described configurations. Furthermore, greater clean passages and / or corrosion resistance may be provided to facilitate the management of combustion product flow.

[0087] The second flow assembly 448 can also be configured to manage more heat more safely than the first flow assembly 446. For example, the second flow assembly 448 can be configured to have a relatively large flow dispersion volume, similar to the WAI arrangement. More heat-resistant materials can also be used in the second flow assembly 448 compared to the first flow assembly 446.

[0088] Turn now Figure 6 A schematic diagram of an emission assembly 500 for a gas turbine engine according to another exemplary embodiment of the present disclosure is provided. Figure 6 An exemplary gas turbine engine can be used in conjunction with the above reference. Figure 1The exemplary turbofan engine 10 described is constructed in essentially the same manner, and Figure 6 The exemplary emission component 500 can be coupled with Figure 2 Exemplary emission components 100 and / or Figure 3 The exemplary emission component 200 is constructed in a similar manner. For example, Figure 6 An exemplary emission assembly 500 generally includes a turbine 102, which is connected to an emission flow path 510 via a flow tap 512.

[0089] Figure 6 The embodiments described herein are similar to Figure 2 and Figure 3 The difference in the embodiment is that an emission regulator 550 is provided. As shown, the emission regulator 550 is provided downstream of the turbine 102 and upstream of the emission flow machine 524. A flow tap 512 is in fluid communication with the emission regulator 550 via a tap valve 514 and an emission inlet valve 522. It should be understood that various other flow configurations can be provided to the emission regulator 550. For example, a secondary pressure source 516 can further supply the emission regulator 550 via a secondary tap 518 controlled by a secondary tap valve 520, as referenced above. Figure 2 and Figure 3 As described.

[0090] As generally described above, relatively high emission flows can enable operations such as effective de-icing under harsh conditions and / or with a relatively large surface area to be de-iced. Figure 6 As shown, machine outlet 525 from exhaust flow machine 524 can supply various exhaust flow components. As also shown, exhaust flow path 510 downstream of exhaust flow machine 524 extends to aircraft flow assembly 530 via aircraft flow valve 532. Aircraft flow assembly 530 may include various components, such as WAI or cockpit ECS components.

[0091] The emission flow machine 524 can be configured similarly to the above reference. Figure 2 and Figure 3 Such components are supplied in the various embodiments described. For example, the exhaust flow machine 524 may be configured to drive a machine load 526 via a machine transmission 528, wherein the machine load 526 is operable to provide torque to the turbine 102. As described above, the exhaust flow machine 524 can be operable under high bypass flow demand, allowing the turbine to operate within a desired range of operability. However, there may also be situations where the exhaust flow machine 524 does not provide such benefits, such as when the exhaust flow demand is relatively low or when there is a large operability margin in the turbine 102. In such cases, the exhaust regulator 550 may control the output of the exhaust flow machine 524 and subsequently control the machine load 526.

[0092] like Figure 6 As shown, the discharge regulator 550 can facilitate bypassing the discharge flow around the machine 524. For example, a discrete flow path may be provided with the discharge regulator 550, which connects the discharge flow path 510 upstream of the machine inlet 523 to the downstream of the machine outlet 525. In this example, the discharge regulator 550 may include a valve, such as a variable flow valve, which can infinitely regulate the proportion of flow sent to the discharge flow machine 524. Figure 6 As shown, this arrangement of the emission regulator 550 is achieved using a diversion valve 554 upstream of the machine inlet 523, allowing it to regulate the flow supplied to the machine inlet and the diversion path 552. The diversion path 552 exits downstream of the machine outlet 525 to supply various emission flow components. As described below, the emission regulator 550 can also be provided in various alternative configurations.

[0093] In embodiments, the emission regulator 550 may be configured as a component of the emission flow machine 524. For example, the emission regulator 550 may be integrated into the emission flow machine 524. In embodiments, the emission regulator 550 includes variable components of the emission flow machine 524. For example, the emission regulator 550 may control the capture rate of the emission flow machine 524 on the emission flow. As used herein, the capture rate refers to the amount of energy captured by the emission flow machine 524 per mass flow rate as described above. In embodiments, the emission regulator 550 may control this capture rate while maintaining the mass flow rate of the emission flow through the machine outlet 525 of the emission flow machine 524. Therefore, the emission regulator 550 may effectively control the flow resistance within the emission flow machine 524. It should also be understood that the emission flow regulator 550 may generally control the total flow rate downstream of the emission flow machine 524 in its various configurations.

[0094] The integrated configuration of the emission regulator 550 and the emission flow machine 524 may include variable adjustment of the stator of the bleed air expansion turbine configuration of the emission flow machine 524 as generally described above. Additionally, characteristics such as the outlet area of ​​the emission flow machine 524 can be controlled during operation of the emission regulator 550. In this example, the outlet area may be a variable outlet area (achieved by a variable area nozzle), which, assuming the rest of the components remain constant, can be increased to reduce the capture rate. It should be understood that various methods for controlling the capture rate fall within the scope of this disclosure to effectively control how much energy is retained in the emission flow and how much energy is captured, for example, returned to the turbine mechanically and / or electrically.

[0095] Figure 2 and Figure 6This demonstrates that such a configuration can be designed to eliminate the need for separate heat exchange components. For example, the use of exhaust flow machines 124 and 524 can significantly reduce temperature and pressure, making them safe and efficient for downstream operations. (Reference) Figure 6 Machine outlet 525 and / or emission regulator 550 can supply the various emission flow components as described above without further heat exchange operations. As depicted, this downstream flow supplies the aircraft flow assembly 530 via aircraft flow valve 532 and the starter assembly 534 via starter assembly valve 536. However, it should be understood that various other components as described elsewhere herein may be provided.

[0096] Turn now Figure 7 A schematic diagram of an emission assembly 600 for a gas turbine engine according to another exemplary embodiment of the present disclosure is provided. Figure 7 The exemplary gas turbine engine 10 can be used in conjunction with the above reference. Figure 1 The exemplary turbofan engine 10 described is constructed in essentially the same manner, and Figure 7 The exemplary emission component 600 can be coupled with Figure 2 Exemplary emission components 100 and / or Figure 3 The exemplary emission component 200 is constructed in a similar manner. For example, Figure 7 An exemplary emission assembly 600 generally includes a turbine 102, which is connected to an emission flow path 610 via a flow tap 612.

[0097] Figure 7 and Figure 6 The difference lies in the provision of a first heat exchange assembly 638 and a second heat exchange assembly 640. It should be understood that the first heat exchange assembly 638 and the second heat exchange assembly 640 are merely exemplary and can be provided in various other numbers and arrangements. Figure 6 As shown, a first heat exchange assembly 638 is provided downstream of the flow tap 612 and the secondary pressure source 616. (Refer to the above...) Figure 2 As described, under certain operating conditions, such as when very high compressor discharge is desired and / or when there is a relatively low discharge flow demand from various discharge flow components, the secondary pressure source 616 can operate as a purging or venting section for pressure from the flow tap 612 or another source. The secondary tap valve 620 can be operable to control the flow into or out of the secondary pressure source 616 through the secondary tap 618. See reference... Figure 2 As described, faucet valve 614 can independently control the flow between faucet 612 and discharge flow path 610. In some embodiments, faucet valve 614 may be a check valve configured to prevent backflow.

[0098] A primary tap 612 and a secondary tap 618 can be connected to the discharge flow path 610 via a discharge inlet valve 622. The discharge inlet valve 622 can operate as a primary controller for the flow through the discharge flow path 610. Downstream of the discharge inlet valve 622, the discharge flow path 610 continues to a first heat exchange assembly 638. The first heat exchange assembly 638 can be operable to control the nature of the discharge flow before it enters the discharge flow machine 624 via machine inlet 623. For example, the first heat exchange assembly 638 can be used to reduce the heat in the discharge flow before it enters the discharge flow machine 624. In various embodiments, the first heat exchange assembly 638 can be an air-to-air heat exchange assembly, can be liquid-cooled, and / or can be evaporatively cooled using a refrigerant cycle.

[0099] One or more exhaust flow components may be configured to receive the flow before it enters the first heat exchange assembly 638. For example, some exhaust flow components may benefit from relatively high temperatures. In the illustrated embodiment, and generally referred to... Figure 3 As discussed, a pre-exchange outlet 642 may be provided upstream of the first heat exchange assembly 638 in the discharge flow path 610. The pre-exchange outlet 642 is operable to supply uncooled discharge flow to the resistance member 644.

[0100] Still referencing Figure 7 The first heat exchange component 638 may be supplemented or replaced by the second heat exchange component 640. As mentioned above, the second heat exchange component 640 may also be provided independently of the first heat exchange component 638. Any heat exchange component 638, 640 is selected if it can be manufactured at least based on the capacity of the exhaust flow machine 624 and / or the desired output of the machine load 626 connected to the exhaust flow machine 624 via the exhaust flow drive 628.

[0101] As referenced above Figure 6 As described, machine outlet 625 and / or emission regulator 650 can supply various emission flow components. As depicted, this downstream flow supplies aircraft flow assembly 630 via aircraft flow valve 632 and to starter assembly 634 via starter assembly valve 636. However, further... Figure 7 The second heat exchange assembly 640 shown is upstream of the aircraft flow assembly 630. In embodiments, the second heat exchange assembly 640 may be provided upstream of only a portion of the aircraft flow assembly (e.g., the cockpit ECS component of such an assembly), while downstream of another portion (e.g., the WAI component of such an assembly) or on a separate branch relative to the other portion.

[0102] As mentioned above (refer to the reference above) Figure 6 As described, it can be provided downstream of turbine 102. Figure 7 The diversion path 652 is connected to the exhaust flow path 610 via a diversion valve 654 downstream of the first heat exchange assembly 638, as shown. However, it should also be understood that the diversion valve 654 may be provided upstream of the first heat exchange assembly 638. In an embodiment, a second diversion valve (not shown) is provided upstream of the first heat exchange assembly 638. This configuration facilitates further operation to control the energy level of the flow supplied to downstream components (e.g., aircraft flow assembly 630) under different engine operating conditions.

[0103] Now for reference Figure 8 A flowchart of a method for operating a gas turbine engine according to a first exemplary aspect of this disclosure is provided. Figure 8 The method can be used with one or more exemplary gas turbine engines described herein (e.g., Figure 2 or Figure 3 (Exemplary gas turbine engine) used together. However, in other exemplary aspects, Figure 8 This method can be used alternatively with any other suitable gas turbine engine.

[0104] Figure 8 The method shown includes, at 701, discharging a discharge stream from the core stream and through a discharge component. The discharge component used in this method can be any of the discharge components 100, 200, 300, 400, 500, and 600 as described above, or can be constructed in another way, such as combining features of multiple such discharge components. At 702, the method provides a machine load driven by the discharge stream machine. The machine load can be any of the machine loads 126, 226, 526, and 626 as described above, or can be constructed in another way, such as combining features of multiple of those machine loads. Similarly, the discharge stream machine used in this method can be any of the discharge stream machines 124, 224, 524, and 624, or can be constructed in another way, such as combining features of multiple of those discharge stream machines. As described above, it should be understood that the discharge stream machine at 702 is configured to receive the discharge stream from 701.

[0105] Still referencing Figure 8At 703, the method provides guiding the exhaust flow through a machine outlet in fluid communication with the aircraft flow assembly. The machine outlet can be any of the machine outlets 125, 225, 525, 625 as described above, or can be constructed in another way, such as combining features of multiple such machine outlets. Similarly, the aircraft flow assembly can be any of the aircraft flow assemblies 130, 230, 530, 630 as described above, or can be constructed in another way, such as combining features of multiple such aircraft flow assemblies. In various embodiments, the aircraft flow assembly at 703 can include a first aircraft flow assembly and a second aircraft flow assembly, for example, to provide the WAI and cockpit ECS as described in more detail above.

[0106] Still referencing Figure 8 The method at 703, directing the emission flow through the machine outlet, may include, at least under certain operating conditions, directing a given amount of emission flow to the aircraft flow assembly. For example, as described above, when data indicating anti-icing conditions is received, the emission mass flow rate of the emission flow may be at least twelve percent (12%) or at least twenty percent (20%) of the core mass flow rate.

[0107] Now for reference Figure 9 A flowchart of a method for operating a gas turbine engine according to a second exemplary aspect of this disclosure is provided. Figure 9 The method can be used with one or more exemplary gas turbine engines described herein (e.g., Figure 4 or Figure 5 (Exemplary gas turbine engine) used together. However, in other exemplary aspects, Figure 9 This method can be used alternatively with any other suitable gas turbine engine.

[0108] Figure 9 The method shown includes, at 801, operating a turbine having a core flow through it. The turbine used in this method can be the exemplary turbine 102 or can be constructed in various other ways. At 802, a first exhaust flow is received using a first heat exchange assembly. At 804, a second exhaust flow is received using a second heat exchange assembly. It should be understood that the first and second heat exchange assemblies can correspond to those described above, such as any of heat exchange assemblies 238, 240, 338, 340, 438, 440, 638, 640; or can be constructed as any other suitable heat exchange assembly.

[0109] Still referencing Figure 9In this method, at 803, a first exhaust stream is directed to a first vehicle flow assembly. At 805, a second exhaust stream is directed to a second vehicle flow assembly. The vehicle flow assemblies at 803 and 805 can be any of the vehicle flow assemblies 130, 230, 530, and 630 as described above, or can be constructed in another way, such as combining features of multiple of those vehicle flow assemblies. In various embodiments, the first vehicle flow assembly at 803 may include a cockpit ECS, and the second vehicle flow assembly at 805 may include a WAI and / or NAI system, as described in more detail above.

[0110] Now for reference Figure 10 A flowchart of a method for operating a gas turbine engine according to a third exemplary aspect of this disclosure is provided. Figure 10 The method can be used with one or more exemplary gas turbine engines described herein (e.g., Figure 4 or Figure 5 (Exemplary gas turbine engine) used together. However, in other exemplary aspects, Figure 10 This method can be used alternatively with any other suitable gas turbine engine.

[0111] Figure 10 The method shown includes operating a turbine with a core flow passing through it at 901. The turbine used in this method can be the exemplary turbine 102 or can be constructed in various other ways. At 902, a first exhaust flow from upstream of the combustion section is received using a first flow tap. At 904, a second exhaust flow from downstream of the combustion section is received using a second flow tap. It should be understood that the first and second flow taps can correspond to any of the aforementioned flow taps, such as any of flow taps 112, 212, 312, 313, 412, 413, 512, 612; or can be constructed as any other suitable flow tap.

[0112] Still referencing Figure 10 The method involves guiding a first discharge stream from a first flow tip using a first flow outlet at 903. At 905, guiding a second discharge stream from a second flow tip using a second flow outlet. The flow outlets at 903 and 905 can be any of the flow outlets 345, 347, 445, and 447 as described above, or they can be constructed in another way, for example, by combining features of multiple aircraft flow components.

[0113] Still referencing Figure 10The method, at 906, provides receiving a first exhaust stream and a second exhaust stream using at least one aircraft flow assembly. The at least one aircraft flow assembly at 906 can be any of the aircraft flow assemblies 130, 230, 530, 630 as described above, or can be constructed in another manner, such as combining features of multiple such aircraft flow assemblies. In various embodiments, the at least one aircraft flow assembly at 906 can include, for example, referenced above. Figure 9 The method is described in more detail for the first and second aircraft flow components.

[0114] Now for reference Figure 11 A flowchart of a method for operating a gas turbine engine according to a fourth exemplary aspect of this disclosure is provided. Figure 11 The method can be used with one or more exemplary gas turbine engines described herein (e.g., Figure 6 or Figure 7 (Exemplary gas turbine engine) used together. However, in other exemplary aspects, Figure 11 This method can be used alternatively with any other suitable gas turbine engine.

[0115] Figure 11 The method shown includes receiving a portion of the core flow and discharge flow at 1001 using a flow tap, for example from the exemplary turbine 102 described above or other suitable turbine or pressure source. It should be understood that the flow tap at 1001 may correspond to one or more of those described above, such as any of flow taps 112, 212, 312, 313, 412, 413, 512, 612; or may be configured as any other suitable flow tap.

[0116] Still referencing Figure 11 At 1002, the method provides for driving a machine load using the emission output of the emission flow machine. The machine load at 1002 can be any of the machine loads 126, 226, 526, and 626 as described above, or can be constructed in another way, such as combining features of multiple such machine loads. Similarly, the emission flow machine used in this method can be any of the emission flow machines 124, 224, 524, and 624, or can be constructed in another way, such as combining features of multiple such emission flow machines. As described above, it should be understood that the emission flow machine at 1002 is configured to receive an emission flow from 1001.

[0117] Still referencing Figure 11 At 1003, the method provides the use of an emission regulator to regulate the emission output driving the machine load at 1002. It should be understood that the emission regulator at 1003 can be one of emission regulators 550 and 650; or it can be as described in the reference... Figure 6 and Figure 7 The configuration is described in more detail in other ways. Similarly, the capture rate of the exhaust flow machine by the exhaust flow can include any of the above-described configurations, such as at 1003, using variable adjustment of the stator or variable adjustment of the outlet area of ​​the exhaust flow machine's bleed air expansion turbine configuration.

[0118] This written description uses examples to disclose this disclosure, including best practices, and also enables any person skilled in the art to practice this disclosure, including making and using any device or system and methods of making any combination. The patent scope of this disclosure is defined by the claims, but may include other examples that would occur to a person skilled in the art. Such other examples are intended to fall within the scope of the claims if they include structural elements that are not indistinguishable from the literal language of the claims, or if they include equivalent structural elements that are not substantially different from the literal language of the claims.

[0119] Further aspects are provided by the subject matter of the following clauses:

[0120] A gas turbine engine includes: a turbine comprising, in a sequential flow order, a compressor section, a combustion section, and a turbine section, the turbine defining a core flow therethrough during operation, the core flow defining a core mass flow rate; an exhaust assembly including an exhaust flow machine and a machine load, the exhaust flow machine being in fluid communication with the compressor section of the turbine and configured to drive the machine load; and a machine outlet being in fluid communication with the exhaust assembly, the machine outlet defining an exhaust flow therethrough during operation of the gas turbine engine, the exhaust flow defining an exhaust mass flow rate, wherein the compressor section is configured to provide the exhaust flow to an exhaust flow assembly via the exhaust flow machine and the machine outlet, wherein the exhaust mass flow rate is at least 12 percent (12%) of the core mass flow rate.

[0121] The gas turbine engine according to any one of the foregoing clauses, wherein the machine load is configured to drive the turbine.

[0122] The gas turbine engine according to any one of the foregoing clauses, wherein the machine outlet is defined through the exhaust flow therethrough during aircraft wing icing operations, wherein the exhaust mass flow rate is at least 12 percent (12%) of the core mass flow rate during aircraft wing icing operations, and wherein the aircraft flow assembly includes wing anti-icing components.

[0123] The gas turbine engine according to any one of the foregoing clauses, wherein the emission mass flow rate is at least 20 percent (20%) of the core mass flow rate.

[0124] The gas turbine engine according to any one of the foregoing clauses, wherein the aircraft flow assembly includes an aircraft environmental control system, a wing anti-icing assembly, or both.

[0125] The gas turbine engine according to any one of the foregoing clauses, wherein the machine load includes a driveable mechanical connection to the turbine.

[0126] The gas turbine engine according to any one of the foregoing clauses, wherein the machine load includes a generator configured to transmit electricity to an electric motor configured to drive the turbine.

[0127] The gas turbine engine according to any one of the foregoing clauses further includes a first heat exchange assembly disposed between the turbine and the exhaust flow machine in a serial flow sequence.

[0128] The gas turbine engine according to any one of the foregoing clauses further includes a second heat exchange assembly disposed in a serial flow sequence between the exhaust flow engine and the aircraft flow assembly.

[0129] The gas turbine engine according to any one of the foregoing clauses, wherein the machine outlet is configured to provide at least a portion of the exhaust flow to the air starter assembly.

[0130] The gas turbine engine according to any one of the foregoing clauses, wherein the gas turbine engine further includes an emission regulator configured to bypass the emission flow machine from the turbine to the machine outlet using at least a portion of the emission flow.

[0131] A method of operating a gas turbine engine includes: operating a turbine to provide a core flow through the turbine, the core flow defining a core mass flow rate, and the turbine including a compressor section, a combustion section, and a turbine section in a sequential flow order; discharging an exhaust flow from the core flow and through an exhaust assembly, the exhaust flow defining an exhaust mass flow rate, and the exhaust assembly including an exhaust flow machine and a machine load, the exhaust flow machine being in fluid communication with the compressor section of the turbine; driving the machine load using the exhaust flow machine; and directing the exhaust flow through a machine outlet in fluid communication with the exhaust assembly, the machine outlet being configured to provide the exhaust flow to a vehicle flow assembly, wherein the exhaust mass flow rate is at least 12 percent (12%) of the core mass flow rate.

[0132] The method according to any one of the foregoing clauses, wherein the method further includes driving the turbine using the machine load.

[0133] The method according to any one of the foregoing clauses, wherein the method further comprises: receiving data indicating anti-icing conditions of an aircraft; and in response to receiving the data indicating anti-icing conditions of an aircraft, directing the emission flow having an emission mass flow rate of at least 12 percent (12%) of the core mass flow rate to the aircraft flow assembly, wherein the aircraft flow assembly includes a wing anti-icing assembly.

[0134] The method according to any one of the foregoing clauses, wherein the emission mass flow rate is at least 20 percent (20%) of the core mass flow rate.

[0135] The method according to any one of the foregoing clauses, wherein the method further comprises: driving the turbine via a mechanical connection using the machine load.

[0136] The method according to any one of the foregoing clauses, wherein the method further comprises: generating electricity using a generator of the machine load; transmitting electricity from the generator to a motor; and driving the turbine using the motor.

[0137] The method according to any one of the foregoing clauses, wherein the method further comprises: cooling the exhaust stream using a first heat exchange assembly, the first heat exchange assembly being arranged in a serial flow sequence between the turbine and the exhaust stream machine.

[0138] The method according to any one of the foregoing clauses, wherein the machine outlet is configured to provide at least a portion of the exhaust flow to the air starter assembly.

[0139] The method according to any one of the foregoing clauses, wherein the method further comprises: using an emission regulator to bypass the emission flow machine from the turbine to the machine outlet by utilizing at least a portion of the emission flow.

[0140] A gas turbine engine includes: a turbine comprising, in a sequential flow order, a compressor section, a combustion section, and a turbine section, the turbine defining a core flow therethrough during operation; a first heat exchange assembly fluidly connected to the turbine for receiving a first exhaust flow from the turbine; a first outlet fluidly connected to the first heat exchange assembly for receiving the first exhaust flow from the first heat exchange assembly and providing the first exhaust flow to a first vehicle flow assembly; a second heat exchange assembly fluidly connected to the turbine for receiving a second exhaust flow from the turbine; and a second outlet fluidly connected to the second heat exchange assembly for receiving the second exhaust flow and providing the second exhaust flow from the second heat exchange assembly to a second vehicle flow assembly.

[0141] The gas turbine engine according to any one of the foregoing clauses, wherein the first vehicle flow assembly includes a cockpit environment control assembly, and the second vehicle flow assembly includes an anti-icing assembly.

[0142] The gas turbine engine according to any one of the foregoing clauses, wherein the anti-icing component includes a wing anti-icing component.

[0143] The gas turbine engine according to any one of the foregoing clauses, wherein the anti-icing component includes a cabin anti-icing component.

[0144] The gas turbine engine according to any one of the foregoing clauses, wherein the first heat exchange assembly is configured to receive the first exhaust flow from upstream of the combustion section of the turbine, and wherein the second heat exchange assembly is configured to receive the second exhaust flow from downstream of the combustion section of the turbine.

[0145] The gas turbine engine according to any one of the foregoing clauses, wherein the first heat exchange assembly is configured to receive the first exhaust flow from the high-pressure compressor of the compressor section of the turbine.

[0146] The gas turbine engine according to any one of the foregoing clauses, wherein the second heat exchange assembly is configured to receive the second exhaust flow from the low-pressure turbine of the turbine section of the turbine.

[0147] The gas turbine engine according to any one of the foregoing clauses, wherein the first exhaust flow defines a first mass flow rate during operation of the gas turbine engine, and the second exhaust flow defines a second mass flow rate during operation of the gas turbine engine, wherein the second mass flow rate is greater than the first mass flow rate.

[0148] In any of the preceding clauses, the second mass flow rate is at least twice the first mass flow rate.

[0149] The gas turbine engine according to any one of the foregoing clauses, wherein the first outlet includes a first cross-sectional area and the second outlet includes a second cross-sectional area, wherein the first cross-sectional area is larger than the second cross-sectional area.

[0150] The gas turbine engine according to any one of the foregoing clauses, wherein the first cross-sectional area is at least twice the second cross-sectional area.

[0151] The gas turbine engine according to any one of the foregoing clauses, wherein the first heat exchange assembly includes a first heat exchange volume, and the second heat exchange assembly includes a second heat exchange volume, the second heat exchange volume being larger than the first heat exchange volume.

[0152] A method for operating a gas turbine engine, the gas turbine engine including a turbine having a core flow therethrough, the turbine including a compressor section, a combustion section and a turbine section in a serial flow sequence, the method comprising: receiving a first exhaust flow from the turbine using a first heat exchange assembly; directing the first exhaust flow from the first heat exchange assembly to a first aircraft flow assembly using a first flow outlet; receiving a second exhaust flow from the turbine using a second heat exchange assembly; and directing the second exhaust flow from the second heat exchange assembly to a second aircraft flow assembly using a second flow outlet.

[0153] The method according to any one of the foregoing clauses, wherein the first aircraft flow assembly includes a cockpit environment control assembly, and the second aircraft flow assembly includes an anti-icing assembly.

[0154] The method according to any one of the foregoing clauses, wherein the method further comprises: receiving a first exhaust stream from upstream of the combustion section of the turbine using the first heat exchange assembly; and receiving a second exhaust stream from downstream of the combustion section of the turbine using the second heat exchange assembly.

[0155] The method according to any one of the foregoing clauses, wherein the method further comprises: receiving the first discharge stream from the high-pressure compressor of the compressor section of the turbine using the first heat exchange assembly.

[0156] The method according to any one of the foregoing clauses, wherein the method further comprises: receiving the second exhaust flow from the low-pressure turbine of the turbine section of the turbine using the second heat exchange assembly.

[0157] The method according to any one of the foregoing clauses, wherein the first discharge stream includes a first mass flow rate, and the second discharge stream includes a second mass flow rate, wherein the second mass flow rate is greater than the first mass flow rate.

[0158] The method according to any one of the foregoing clauses, wherein the second mass flow rate is at least twice the first mass flow rate.

[0159] According to any one of the preceding clauses, the first heat exchange component includes a first heat exchange volume, and the second heat exchange component includes a second heat exchange volume, the second heat exchange volume being larger than the first heat exchange volume.

[0160] A gas turbine engine includes: a turbine comprising, in a sequential flow order, a compressor section, a combustion section, and a turbine section, the turbine defining a core flow therethrough; a first flow tip configured to receive a first exhaust flow from upstream of the combustion section; a first flow outlet in fluid communication with the first flow tip; a second flow tip configured to receive a second exhaust flow from downstream of the combustion section; and a second flow outlet in fluid communication with the second flow tip; wherein the first flow outlet and the second flow outlet are configured to direct the first exhaust flow and the second exhaust flow to at least one aircraft flow assembly.

[0161] The gas turbine engine according to any one of the foregoing clauses, wherein the first flow outlet is configured to direct the first exhaust flow to a first vehicle flow assembly, and the second flow outlet is configured to direct the second exhaust flow to a second vehicle flow assembly.

[0162] The gas turbine engine according to any one of the foregoing clauses, wherein the first vehicle flow assembly includes a cockpit environment control assembly, and wherein the second vehicle flow assembly includes an anti-icing assembly.

[0163] The gas turbine engine according to any one of the foregoing clauses, wherein the anti-icing component is a wing anti-icing component.

[0164] The gas turbine engine according to any one of the foregoing clauses, wherein the anti-icing component is a cabin anti-icing component.

[0165] The gas turbine engine according to any one of the foregoing clauses, wherein the gas turbine engine further comprises: a first heat exchange assembly configured to receive the first exhaust flow from the first flow tip; and a second heat exchange assembly configured to receive the second exhaust flow from the second flow tip.

[0166] The gas turbine engine according to any one of the foregoing clauses, wherein the first flow tip is configured to receive the first discharge flow from the high-pressure compressor of the compressor section of the turbine.

[0167] The gas turbine engine according to any one of the foregoing clauses, wherein the second flow tip is configured to receive the second exhaust flow from the low-pressure turbine of the turbine section of the turbine.

[0168] The gas turbine engine according to any one of the foregoing clauses, wherein the first exhaust stream from upstream of the combustion section defines a first mass flow rate during operation, and the second exhaust stream from downstream of the combustion section defines a second mass flow rate during operation, wherein the second mass flow rate is greater than the first mass flow rate.

[0169] In any of the preceding clauses, the second mass flow rate is at least twice the first mass flow rate.

[0170] The gas turbine engine according to any one of the foregoing clauses, wherein the first outlet includes a first cross-sectional area and the second outlet includes a second cross-sectional area, wherein the first cross-sectional area is larger than the second cross-sectional area.

[0171] The gas turbine engine according to any one of the foregoing clauses, wherein the first cross-sectional area is at least twice the second cross-sectional area.

[0172] The gas turbine engine according to any one of the foregoing clauses, wherein the first heat exchange assembly includes a first heat exchange volume, and the second heat exchange assembly includes a second heat exchange volume, the second heat exchange volume being larger than the first heat exchange volume.

[0173] A method for operating a gas turbine engine, the gas turbine engine including a turbine having a core flow therethrough, the turbine including a compressor section, a combustion section and a turbine section in a serial flow sequence, the method comprising: receiving a first exhaust flow from upstream of the combustion section using a first flow tip; directing the first exhaust flow from the first flow tip to at least one aircraft flow assembly using a first flow outlet; receiving a second exhaust flow from downstream of the combustion section using a second flow tip; and directing the second exhaust flow from the second flow tip to the at least one aircraft flow assembly using a second flow outlet.

[0174] The method according to any one of the foregoing clauses, wherein the method further comprises: using the first flow outlet to guide the first emission flow from the first flow head to a first aircraft flow assembly; and using the second flow outlet to guide the second emission flow from the second flow head to a second aircraft flow assembly.

[0175] The method according to any one of the foregoing clauses, wherein the method further comprises: receiving the first discharge stream from the high-pressure compressor of the compressor section of the turbine using the first flow tap.

[0176] The method according to any one of the foregoing clauses, wherein the method further comprises: receiving the second discharge stream from the low-pressure turbine of the turbine section of the turbine using the second flow tap.

[0177] The method according to any one of the foregoing clauses, wherein the first aircraft flow assembly includes a cockpit environment control assembly, and the second aircraft flow assembly includes an anti-icing assembly.

[0178] The method according to any one of the foregoing clauses, wherein the anti-icing component is a wing anti-icing component.

[0179] The method according to any one of the foregoing clauses, wherein the anti-icing component is a cabin anti-icing component.

[0180] A gas turbine engine includes: a turbine comprising, in a sequential flow order, a compressor section, a combustion section, and a turbine section, the turbine defining a core flow therethrough during operation; a flow tap in fluid communication with the turbine, the flow tap configured to receive a portion of the core flow therethrough as an exhaust flow; and an exhaust assembly including: a machine load; an exhaust flow machine configured to drive the machine load via the flow tap; and an exhaust regulator configured to regulate an exhaust output supplied to the exhaust flow machine by controlling a capture rate of the exhaust flow by the exhaust flow machine.

[0181] The gas turbine engine according to any one of the foregoing clauses, wherein the gas turbine engine further includes a flow outlet downstream of the exhaust flow machine, wherein the exhaust regulator is further configured to control the capture rate of the exhaust flow machine on the exhaust flow while maintaining the mass flow rate of the exhaust flow through the flow outlet.

[0182] The gas turbine engine according to any one of the foregoing clauses, wherein the emission regulator comprises: a diversion path fluidly connecting the flow head and the flow outlet and bypassing the emission flow machine; and a diversion valve disposed upstream of the emission flow machine, the diversion valve being configured to control the diversion of flow through the diversion path.

[0183] The gas turbine engine according to any one of the foregoing clauses, wherein the emission regulator includes at least one variable emission feature configured to control the capture rate of the emission stream by the emission stream machine.

[0184] The gas turbine engine according to any one of the foregoing clauses, wherein the at least one variable emission feature is configured as a component of the emission flow machine, the component being configured to regulate the flow exiting the emission flow machine.

[0185] The gas turbine engine according to any one of the foregoing clauses, wherein the exhaust flow includes an exhaust flow mass flow rate, and the core flow includes a core mass flow rate, wherein the exhaust mass flow rate is at least 12 percent (12%) of the core mass flow rate.

[0186] The gas turbine engine according to any one of the foregoing clauses further includes an aircraft flow assembly in fluid communication with the emission assembly, wherein the flow assembly includes at least one of: a wing anti-icing assembly; a cabin anti-icing assembly; or a cabin environmental control assembly.

[0187] The gas turbine engine according to any one of the foregoing clauses, wherein the machine load includes a driveable mechanical connection to the turbine.

[0188] The gas turbine engine according to any one of the foregoing clauses, wherein the machine load includes a generator configured to transmit electricity to an electric motor configured to drive the turbine.

[0189] The gas turbine engine according to any one of the foregoing clauses further includes a first heat exchange assembly disposed between the turbine and the exhaust flow machine in a serial flow sequence.

[0190] The gas turbine engine according to any one of the foregoing clauses further includes a second heat exchange assembly disposed in a serial flow sequence between the exhaust flow machine and the flow outlet.

[0191] The gas turbine engine according to any one of the foregoing clauses, wherein the flow outlet is configured to provide at least a portion of the exhaust flow to the air starter assembly.

[0192] A method of operating a gas turbine engine, the gas turbine engine including a turbine having a core flow therethrough, the turbine including a compressor section, a combustion section and a turbine section in a serial flow sequence, the method comprising: receiving a portion of the core flow defining an exhaust flow via a flow tap in fluid communication with the turbine; driving a machine load via an exhaust output through an exhaust flow machine configured to be in fluid communication with the turbine via the flow tap; and regulating the exhaust output by controlling the capture rate of the exhaust flow by the exhaust flow machine using an exhaust regulator.

[0193] The method according to any one of the foregoing clauses, wherein the method further comprises using the emission regulator to control the capture rate of the emission flow machine on the emission flow while maintaining the emission mass flow rate of the emission flow through an outlet located downstream of the emission flow machine.

[0194] The method according to any one of the foregoing clauses, wherein the method further comprises controlling the diversion of flow through a diversion path by means of a diversion valve disposed upstream of the discharge flow machine, the diversion path being fluidly connected to the flow tap and the flow outlet and bypassing the discharge flow machine.

[0195] The method according to any one of the foregoing clauses, wherein the method further comprises controlling the capture rate of the emission stream machine on the emission stream using at least one variable emission characteristic of the emission regulator.

[0196] According to any one of the foregoing clauses, the at least one variable emission feature is configured as a component of the emission flow machine.

[0197] The method according to any one of the foregoing clauses, wherein the method further comprises: adjusting the fluid flow out of the discharge flow machine using the at least one variable discharge characteristic.

[0198] The method according to any one of the foregoing clauses, wherein the emission stream includes an emission stream mass flow rate, and the core stream includes a core mass flow rate, wherein the emission mass flow rate is at least 12 percent (12%) of the core mass flow rate.

[0199] The method according to any one of the foregoing clauses, wherein the method further comprises: generating electricity using a generator of the machine load; transmitting the electricity to a motor; and driving the turbine using the motor.

Claims

1. A gas turbine engine, characterized in that, include: A turbine comprising a compressor section, a combustion section and a turbine section in a sequential flow order, the turbine being defined by a core flow passing through it during operation; A first heat exchange assembly, which is in fluid communication with the turbine, is used to receive a first exhaust flow from the turbine through a first exhaust flow path; A first heat exchange component valve, located upstream of the first heat exchange component, is used to control the first discharge flow to the first heat exchange component; A discharge inlet valve, which is located upstream of the first heat exchange component valve and is configured to control the first discharge flow to the first heat exchange component valve; A first faucet valve, upstream of the discharge inlet valve and configured to control the first discharge flow from the turbine; A secondary faucet valve, upstream of the discharge inlet valve and configured to control the flow between the secondary pressure source and the discharge inlet valve; A first outlet, which is in fluid communication with the first heat exchange assembly, is used to receive the first exhaust flow from the first heat exchange assembly and to provide the first exhaust flow to the first aircraft flow assembly; A second heat exchange assembly, which is in fluid communication with the turbine, is used to receive a second exhaust flow from the turbine via a second exhaust flow path; and A second outlet, which is in fluid communication with the second heat exchange assembly, is used to receive the second exhaust flow and to supply the second exhaust flow from the second heat exchange assembly to the second aircraft flow assembly; Wherein the first flow outlet is not connected to the second aircraft flow component, and wherein the second flow outlet is not connected to the first aircraft flow component; The first emission path and the second emission path operate independently of each other.

2. The gas turbine engine according to claim 1, characterized in that, The first aircraft flow component includes a cockpit environment control component, and the second aircraft flow component includes an anti-icing component.

3. The gas turbine engine according to claim 2, characterized in that, The anti-icing components mentioned above include wing anti-icing components.

4. The gas turbine engine according to claim 2, characterized in that, The anti-icing components mentioned above include cabin anti-icing components.

5. The gas turbine engine according to claim 1, characterized in that, The first heat exchange assembly is configured to receive the first exhaust flow from upstream of the combustion section of the turbine, and the second heat exchange assembly is configured to receive the second exhaust flow from downstream of the combustion section of the turbine.

6. The gas turbine engine according to claim 5, characterized in that, The first heat exchange component is configured to receive the first discharge stream from the high-pressure compressor of the compressor section of the turbine.

7. The gas turbine engine according to claim 6, characterized in that, The second heat exchange component is configured to receive the second exhaust flow from the low-pressure turbine of the turbine section of the turbine.

8. The gas turbine engine according to claim 1, characterized in that, The first exhaust flow defines a first mass flow rate during operation of the gas turbine engine, and the second exhaust flow defines a second mass flow rate during operation of the gas turbine engine, wherein the second mass flow rate is greater than the first mass flow rate.

9. The gas turbine engine according to claim 8, characterized in that, The second mass flow rate is at least twice the first mass flow rate.

10. The gas turbine engine according to claim 1, characterized in that, The first outlet includes a first cross-sectional area, and the second outlet includes a second cross-sectional area, wherein the first cross-sectional area is larger than the second cross-sectional area.

11. The gas turbine engine according to claim 10, characterized in that, The area of ​​the first cross-section is at least twice the area of ​​the second cross-section.

12. The gas turbine engine according to claim 1, characterized in that, The first heat exchange component includes a first heat exchange volume, and the second heat exchange component includes a second heat exchange volume, wherein the second heat exchange volume is larger than the first heat exchange volume.

13. A method for operating a gas turbine engine, the gas turbine engine comprising a turbine having a core flow therethrough, the turbine comprising a compressor section, a combustion section and a turbine section in a sequential flow order, characterized in that, The method includes: The first exhaust flow from the turbine is received via a first exhaust flow path using a first heat exchange component. Using a first flow outlet, the first discharge flow from the first heat exchange assembly is directed to the first aircraft flow assembly, wherein directing the first discharge flow from the first heat exchange assembly to the first aircraft flow assembly includes controlling the first discharge flow using a first heat exchange assembly valve located upstream of the first heat exchange assembly, an discharge inlet valve upstream of the first heat exchange assembly valve, a first tap valve upstream of the discharge inlet valve, and a secondary tap valve upstream of the discharge inlet valve, the secondary tap valve being configured to control the flow between a secondary pressure source and the discharge inlet valve; The second exhaust flow from the turbine is received via a second exhaust flow path using a second heat exchange assembly; and The second exhaust stream from the second heat exchange assembly is directed to the second aircraft stream assembly using the second stream outlet; Wherein the first flow outlet is not connected to the second aircraft flow assembly, and wherein the second flow outlet is not connected to the first aircraft flow assembly; and The first emission path and the second emission path operate independently of each other.

14. The method according to claim 13, characterized in that, The first aircraft flow component includes a cockpit environment control component, and the second aircraft flow component includes an anti-icing component.

15. The method according to claim 13, characterized in that, Further includes: The first exhaust stream from upstream of the combustion section of the turbine is received using the first heat exchange component; and The second heat exchange assembly is used to receive the second exhaust stream from downstream of the combustion section of the turbine.

16. The method according to claim 15, characterized in that, Further includes: The first discharge stream from the high-pressure compressor of the compressor section of the turbine is received using the first heat exchange assembly.

17. The method according to claim 16, characterized in that, Further includes: The second heat exchange assembly receives the second exhaust flow from the low-pressure turbine of the turbine section of the turbine.

18. The method according to claim 15, characterized in that, The first discharge stream includes a first mass flow rate, and the second discharge stream includes a second mass flow rate, wherein the second mass flow rate is greater than the first mass flow rate.

19. The method according to claim 18, characterized in that, The second mass flow rate is at least twice the first mass flow rate.

20. The method according to claim 13, characterized in that, The first heat exchange component includes a first heat exchange volume, and the second heat exchange component includes a second heat exchange volume, wherein the second heat exchange volume is larger than the first heat exchange volume.

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