Emission flow assembly for a gas turbine engine

Abstract

A gas turbine engine includes a turbine defining a core flow therethrough during operation. A flow tap is provided in fluid communication with the turbine, wherein the flow tap is configured to receive a portion of the core flow therethrough as a bleed flow. A bleed assembly includes a machine load, a bleed flow machine, and a bleed regulator. The bleed flow machine is disposed in fluid communication with the turbine through the flow tap and is configured to drive the machine load. The bleed regulator is configured to regulate a bleed output provided to the bleed flow machine by controlling a capture rate of the bleed flow machine on the bleed flow.

Description

TECHNICAL FIELD

[0001] The present subject matter relates generally to bleed flow assemblies for gas turbine engines. BACKGROUND

[0002] Gas turbine engines typically include a fan and a turbine machine. The turbine machine generally includes an inlet, one or more compressors, a combustor, and at least one turbine. The compressor compresses air, which is directed to the combustor where it is mixed with fuel. The mixture is then ignited to generate hot combustion gases. The combustion gases are directed to the turbine, which extracts energy from the combustion gases to power the compressor and to produce useful work to propel an aircraft in flight and / or to power a load, such as an electrical generator. A bleed flow assembly can generally be provided to bleed air flow from the turbine machine for use in various operations. The use of bleed flow can impact turbine machine operation such that both the bleed flow and turbine machine demands must be considered to manage the bleed flow. BRIEF DESCRIPTION OF DRAWINGS

[0003] A complete and enabling disclosure of the present disclosure, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification, which is to be taken in conjunction with the appended drawings, wherein:

[0004] Figure 1 is a schematic cross-sectional view of an exemplary gas turbine engine in accordance with various embodiments of the present subject matter.

[0005] Figure 2 is a schematic view of a bleed assembly for a gas turbine engine in accordance with an exemplary embodiment of the present disclosure.

[0006] Figure 3 is a schematic view of a bleed assembly for a gas turbine engine in accordance with another exemplary embodiment of the present disclosure.

[0007] Figure 4 is a schematic view of a bleed assembly for a gas turbine engine in accordance with yet another exemplary embodiment of the present disclosure.

[0008] Figure 5 is a schematic view of a bleed assembly for a gas turbine engine in accordance with still another exemplary embodiment of the present disclosure.

[0009] Figure 6 is a schematic view of a bleed assembly for a gas turbine engine in accordance with still another exemplary embodiment of the present disclosure.

[0010] Figure 7 is a schematic view of a bleed assembly for a gas turbine engine in accordance with still another exemplary embodiment of the present disclosure.

[0011] Figure 8A flow diagram of a method for operating a gas turbine engine according to an example aspect of the present disclosure is provided.

[0012] Figure 9 A flow diagram of a method for operating a gas turbine engine according to another example aspect of the present disclosure is provided.

[0013] Figure 10 A flow diagram of a method for operating a gas turbine engine according to yet another example aspect of the present disclosure is provided.

[0014] Figure 11 A flow diagram of a method for operating a gas turbine engine according to yet another example aspect of the present disclosure is provided. DETAILED DESCRIPTION

[0015] Reference will now be made in detail to the current embodiments of the present disclosure, one or more examples of which are illustrated in the drawings. The detailed description uses numerical and letter designations

[0016] The word "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 preferred or advantageous over other implementations. Furthermore, unless otherwise explicitly provided, none of the examples described herein are intended to be exhaustive or limiting of the possible implementations.

[0017] For purposes of the description hereinafter, the terms "upper", "lower", "right", "left", "vertical", "horizontal", "top", "bottom", "lateral", "longitudinal", and derivatives thereof shall relate to the embodiments as they are oriented in the drawing

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

[0019] The terms "forward" and "aft" refer to relative positions within a gas turbine engine or vehicle and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, for a gas turbine engine, forward refers to a position closer to the engine inlet and aft refers to a position closer to the engine nozzle or exhaust.

[0020] The terms“upstream” and“downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example,“upstream” refers to the direction from which the fluid flows, and“downstream” refers to the direction to which the fluid flows.

[0021] Unless otherwise specified herein, the terms“coupled,”“fixed,”“attached to,” and the like, mean either directly coupled, fixed, or attached by one or more intermediate components or features.

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

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

[0024] Approximating language as used throughout the specification and claims is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is directed. Accordingly, a value modified by a term or terms, such as“about,”“approximately,” and“substantially,” are not limited to the precise value specified. In at least some instances, the approximating language can correspond to the precision of an instrument for measuring the value or the precision of the methods or machines for constructing or manufacturing the components and / or systems. For example, the approximating language can refer to a margin of error within 1%, 2%, 4%, 10%, 15%, or 20%. These approximating margins can apply to individual values, to any one of the endpoints of a range, or to the range itself. For values that are not approximating, the stated value is the exact value.

[0025] Herein and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

[0026] In certain example embodiments, the operating temperature of the airflow through the third stream can be lower than a maximum compressor discharge temperature of the engine, and more particularly, can be lower than 350 degrees Fahrenheit (such as lower than 300 degrees Fahrenheit, such as lower than 250 degrees Fahrenheit, such as lower than 200 degrees Fahrenheit, and at least as high as ambient temperature). In certain example embodiments, these operating temperatures can facilitate heat transfer to or from the airflow through the third stream and the separate fluid stream. Further, in certain example embodiments, the airflow through the third stream can contribute less than 50% of the total engine thrust (and at least, for example, 2% of the total engine thrust) when operating in takeoff conditions, or more particularly, in sea level rated takeoff power, static flight speed, 86 degrees Fahrenheit ambient temperature operating conditions.

[0027] Further, in certain example embodiments, the aforementioned example percentage contribution to total thrust by the third stream of airflow (e.g., airflow, mixing, or exhaust properties), and thereby to overall system performance, can be adjusted passively during engine operation or purposefully modified through the use of engine control features such as fuel flow, electric machine power, variable stators, variable inlet guide vanes, valves, variable exhaust geometry, or fluidic features, to adjust or optimize overall system performance under a wide range of potential operating conditions.

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

[0029] The term “combustion engine” refers to a turbomachinery component used to generate a torque output by forces imparted from a combustion reaction. A combustion engine 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 having a turbomachine as all or a portion of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, and the like, as well as hybrid electric versions of one or more of these engines.

[0031] The terms “low” and “high,” or their respective comparative forms (e.g., more “low” and more “high,” as applicable), when used in conjunction with a compressor, turbine, shaft or spool component, or the like, each refer to a relative speed within the engine, unless otherwise noted. For example, a “low turbine” or “low speed turbine” defines a component configured to operate at a lower rotational speed (such as a maximum allowable rotational speed) than a “high turbine” or “high speed turbine” at the location of 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 a core flow of a turbine. The exhaust flow is then redirected to various exhaust flow components of the gas turbine engine or an aircraft in which the gas turbine engine is installed, such as an anti-icing and / or cabin environmental control system. The present disclosure manages to provide systems and methods to avoid engine operability issues, such as stall and surge issues, while providing a relatively large amount of exhaust flow. The inventors of the present disclosure have discovered that the high percentage of exhaust flow required to remove some exhaust flow components relative to the core flow benefits from a configuration and operation that reduces and / or recaptures energy tapped 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 with post-combustion exhaust flow downstream of the combustion section of the turbine.

[0033] Additionally, the inventors of the present disclosure have further discovered that downstream exhaust flow components can benefit from such systems and methods of recapturing exhaust flow energy as mentioned above. For example, in energy recapturing operations, pressure and temperature can be reduced to safe levels, thereby providing exhaust flow resources to components that would otherwise not have sufficient resilience to safely handle such exhaust flow resources. Accordingly, systems and methods that conform exhaust flow resources to the requirements of various exhaust flow components would be useful. Additionally, various exhaust flow components can be configured to operate under a wider range of conditions and / or operate in a lighter or more cost-effective structure to take advantage of the handling of such exhaust flow resources provided herein.

[0034] For example, in one exemplary aspect of the present disclosure, a gas turbine engine is provided having a turbine, an exhaust assembly, and a flow outlet. The exhaust assembly can be configured to receive an exhaust flow from the turbine and can include an exhaust flow machine (e.g., a bleed air turbine) having a machine outlet and configured to drive a machine load (e.g., an accessory gearbox, an electric machine, etc.) to capture energy in the exhaust flow. Further, the machine outlet is configured to receive the exhaust flow during operation of the gas turbine engine at a mass flow rate equal to at least twelve percent (12%) of a core mass flow rate of a core flow through the turbine, for example, during an aircraft wing ice accretion operation. In this way, the gas turbine engine can be designed to provide a relatively large amount of airflow to an aircraft for wing de-icing operations, for example, while minimizing energy losses associated with such a large amount 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 the embodiment depicted, the aircraft gas turbine engine is a high-bypass turbofan jet engine 10, referred to herein as "turbofan engine 10." As shown in Figure 1 The turbofan engine 10 defines an axial direction A (extending parallel to a reference longitudinal centerline 12) and a radial direction R. In general, the turbofan engine 10 includes a fan section 14 and a turbine section 16 disposed downstream from the fan section 14.

[0039] The exemplary turbine section 16 depicted generally includes a substantially tubular outer casing 18 defining an annular inlet 20. The casing 18 surrounds, in serial flow relationship: a compressor section including a booster 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 injection exhaust nozzle section 32. The compressor section, the combustion section 26, the turbine section, and the injection exhaust nozzle section 32 together at least partially define a core air flowpath 37 through the turbine section 16. A high pressure (HP) shaft or spool 34 (or more precisely, a high pressure spool assembly as described below) drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft or spool 36 drivingly connects the LP turbine 30 to the LP compressor 22.

[0040] For the embodiment depicted, the fan section 14 includes a variable pitch fan 38 having a plurality of fan blades 40 coupled in spaced relation to a disc 42. As depicted, the fan blades 40 extend generally in the radial direction R outwardly from the disc 42. Since the fan blades 40 are operably coupled to suitable actuating members 44 configured to collectively and uniformly vary the pitch of the fan blades 40, each fan blade 40 is rotatable about a pitch axis P relative to the disc 42. The fan blades 40, the disc 42, and the actuating members 44 are rotatable together about the longitudinal centerline 12 by the LP shaft 36 across a power gear box 46. The power gear box 46 includes a plurality of gears for stepwise reducing the rotational speed of the LP shaft 36 to a more efficient rotational fan speed.

[0041] Still referring to the exemplary embodiment of Figure 1 The disc 42 is shrouded 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 casing or outer nacelle 50 circumferentially surrounding at least a portion of the fan 38 and / or the turbine section 16. The nacelle 50 is supported relative to the turbine section 16 by a plurality of circumferentially spaced outlet guide vanes 52. Moreover, the nacelle 50 extends over an outer portion of the turbine section 16 so as to define a bypass airflow passage 56 therebetween.

[0042] During operation of the turbofan engine 10, an amount of air 58 enters the turbofan engine 10 through an associated inlet 60 of the nacelle 50 and / or fan section 14. As the amount of air 58 passes through the fan blades 40, a first portion of the air 58 is directed or channeled into the bypass airflow passage 56 as indicated by arrow 62, and a second portion of the air 58 is directed or channeled into the LP compressor 22 as indicated by arrow 64. The ratio between the first and second portions of air 62 and 64 is commonly referred to as a bypass ratio. The pressure of the second portion of air 64 is then increased as it is channeled through the HP compressor 24 and into the combustion section 26 where it is mixed with and burned a fuel to provide combustion gases 66. Subsequently, the combustion gases 66 are channeled 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] The combustion gases 66 are then channeled through the ejection exhaust nozzle section 32 of the turbine engine 16 to provide propulsive thrust. At the same time, the pressure of the first portion of air 62 is significantly increased as it is channeled 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] Further, as illustratively depicted, the example turbofan engine 10 is part of a gas turbine engine that further includes various accessory systems to assist in the operation of the turbofan engine 10 and / or an aircraft that includes the example turbofan engine 10. For example, as depicted, the example gas turbine engine further includes a fuel delivery system 70 that is operable with a combustion section 26 of a turbine 16 of the turbofan engine 10 for providing fuel to the combustion section 26. The example fuel delivery system 70 can include one or more fuel delivery lines, a fuel pump (not shown), etc. Further, the example gas turbine engine includes an exhaust assembly 72 as will be explained in greater detail below. It will be understood that the exhaust assembly 72 generally includes an exhaust flow machine 74 that is configured to receive an exhaust gas stream from the turbine 16. The exhaust flow machine 74 can generally operate to convert energy in the exhaust gas stream into mechanical work, which can also be referred to herein as an exhaust output. It will also be understood that the exhaust output can represent another form of energy, such as electrical potential energy. Thus, 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 gas stream to a desired level, for example, to satisfy temperature and / or pressure limits of downstream components. Such reduced energy gas streams can be used for various purposes within, for example, the turbofan engine 10, an aircraft that has the turbofan engine 10 installed therein; not shown, etc., can be used within an environmental control system (ECS) of the aircraft, can be used in wing anti-icing (WAI; to provide anti-icing operations for a wing of the aircraft) and / or nacelle anti-icing (NAI; to provide anti-icing operations for a nacelle of the turbofan engine 10, such as Figure 1 nacelle 50 of the turbofan engine 10, etc.

[0045] Referring now to Figure 2 , a schematic view of an exhaust assembly 100 for a gas turbine engine in accordance with example embodiments of the present disclosure is provided. Figure 2 The example gas turbine engine of Figure 1 may be constructed in substantially the same manner as the example turbofan engine 10 described above with reference to Figure 2 For example, The example gas turbine engine of

[0046] may be constructed in substantially the same manner as the example turbofan engine 10 described above with reference to Figure 2 Thus, as described above, Figure 1 Figure 2Turbomachine 102 can generally include a compressor section having a 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. The compressor section, combustion section 26, and turbine section are arranged in serial flow order. A high pressure (HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HP compressor 24, and a low pressure (LP) shaft or spool 36 drivingly connects the LP turbine 30 to the LP compressor 22.

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

[0048] Further, Figure 2 the arrangement includes a fuel delivery system 104 (which can be constructed in a similar manner as fuel delivery system 70 of Figure 1 ) operable with combustion section 26 of turbomachine 102 for providing fuel to combustion section 26 of turbomachine 102. As Figure 2 depicted, a fuel source 106 is provided to supply fuel to combustion section 26 through a fuel delivery line 108.

[0049] As Figure 2 depicted, exemplary exhaust assembly 100 includes an exhaust flow path 110 in fluid communication with the compressor section of turbomachine 102. More particularly, as depicted, exhaust flow path 110 can be in fluid communication with HP compressor 24. In Figure 2 embodiments, a flow tap 112 is provided from HP compressor 24 to exhaust flow path 110. For example, flow tap 112 can be provided at a particular stage of HP compressor 24, such as the fourth stage. It should be appreciated that flow tap 112 can draw from various other pressure sources, such as turbomachine 102. In various embodiments, the source of flow tap 112 is selected based on temperature and / or pressure requirements of exhaust assembly 100. It should also be appreciated that more than one flow tap 112 can be provided, for example, to satisfy different temperature and / or pressure requirements of exhaust assembly 100 at a given engine operating output. As Figure 2 shown, flow tap 112 can be separated from exhaust flow path 110 by a tap valve 114, for example, to control the flow input from turbomachine 102 to exhaust flow path 110. However, in Figure 2In the embodiment shown, the faucet valve 114 is shown configured as a one-way valve configured to prevent backflow through the flow faucet 112. It should be appreciated that various other valve arrangements can be provided, for example in the exhaust flow path 110, to control the flow through the exhaust flow path 110.

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

[0051] The various flows within the turbine 102 and exhaust assembly can be quantified by mass flow rate (a function of the mass of air flow per unit time). For example, a core mass flow rate can be defined based on the total mass flowing into the turbine 102 per unit time (i.e., corresponding to the second portion of air 58 as indicated by arrow 64 in Figure 1 FIG. 1). This core mass flow rate can be used for comparison with the mass flow rate within the exhaust flow path 110 (e.g., an exhaust mass flow rate representing the total mass flowing into the exhaust flow path 110 per unit time).

[0052] One or more flow control devices can be used to regulate the mass flow rate within the exhaust flow path 110. For example, as Figure 2 shown and as described above, the faucet valve 114 and / or the secondary faucet valve 120 can be operable to control the amount of air flow diverted through the turbine 102 to the exhaust flow path 110. Also as Figure 2 shown, an exhaust admission valve 122 can also be provided to regulate the exhaust mass flow rate within the exhaust flow path 110. As shown, the exhaust admission valve 122 is operable to control all pressure sources to the exhaust flow path 110. However, it should be appreciated that the pressure sources (such as the flow faucet 112 and the secondary faucet 118) can be independently controlled from one another.

[0053] Still referring to 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] The machine load 126 can be operable to directly or indirectly drive the turbine machine 102 as described above. Thus, the machine load 126 can be operable to increase or maintain an operating speed of the turbine machine 102. Accordingly, the machine load 126 can be operable to account for deficiencies in the operation of the turbine machine 102, such as those encountered by diverting portions of the core flow through the turbine machine into the bleed flow path 110. In this case, the air / fuel ratio and total flow within the combustion section 26 can be reduced, thereby reducing operability margins and potentially increasing the likelihood of stall, surge, and / or rich blow out. The reduced operability margins can be accounted for or reversed by increasing the rotational speed and thus the flow through the turbine machine 102. For example, a greater rotational speed of the turbine machine 102 can increase the air / fuel ratio and total flow to account for the relatively high flow diverted to the bleed flow path 110. Accordingly, the arrangement of the bleed flow machine 124 and the machine load 126 can enable a relatively high bleed flow to be taken from the core flow through the turbine machine 102. For example, the bleed flow can have a bleed mass flow rate defined as the mass per unit time entering the bleed flow path 110 that is a relatively high percentage of a core mass flow rate defined as the mass per unit time entering the core flow path 37 (see Figure 1 ) of the turbine machine 102. In various embodiments, the bleed mass flow rate can be at least ten percent, twelve percent, fourteen percent, sixteen percent, eighteen percent, twenty percent, or twenty-two percent of the core mass flow rate. As will be described below, various further configurations can be provided to enable such a high bleed flow arrangement.

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

[0057] As noted above, aircraft flow assembly 130 refers to any assembly that utilizes the bleed flow from bleed flow path 110. For example, a WAI arrangement can be provided to maintain a safe temperature of the aircraft wing to avoid icing conditions. An ECS can be provided to maintain a safe environmental condition of the interior cabin of the aircraft. It should be appreciated that aircraft flow assembly 130 can include various sub-assemblies with associated valves and controllers to maintain the desired control without necessarily affecting the operation of the upstream components of bleed assembly 100. Additionally, aircraft flow assembly 130 can include at least one bleed flow bleed-off (not shown) configured to purge excess bleed flow to the outside environment.

[0058] Still referring to Figure 2 , a machine outlet 125 from the machine load 126 is further shown to provide bleed flow to a starter assembly 134 through a starter assembly valve 136. Starter assembly 134 can be operable to start a gas turbine engine, such as another gas turbine engine of the same aircraft on which bleed assembly 100 is installed and associated. Starter assembly valve 136 can be operable to control starter assembly 134, such as to engage a start mode. In embodiments, starter assembly 134 is a pneumatic starter configured to provide flow and subsequent rotation to a gas turbine engine when starter assembly valve 136 is operated. It should be appreciated that starter assembly 134 can include various further downstream valves and controllers to utilize and control bleed flow for preferred start operations.

[0059] In various embodiments, one or more further or alternative bleed flow components can be provided downstream of bleed flow machine 124. For example, a separate circuit with a corresponding separate valve arrangement can be provided for each of the ECS and WAI arrangements. Additionally, various embodiments can provide bleed flow components upstream of bleed flow machine 124. For example, systems that require or benefit from relatively high pressure or temperature can operate from flow that has not yet passed through bleed flow machine 124. In embodiments, a nacelle anti-icing device (NAI) can be configured to tap off upstream of bleed flow machine 124 for preventing icing conditions in nacelle 50 (see Figure 1 As will be discussed in greater detail below, each of the various bleed flow components and assemblies can be configured, sized, and shaped in cooperation with upstream bleed flow path 110 to effectively utilize bleed flow and maintain efficient operation of turbomachine 102.

[0060] Turning now to Figure 3 , a schematic view of a bleed assembly 200 for a gas turbine engine according to another example embodiment of the present disclosure is provided. Figure 3 The example gas turbine engine 10 of 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 discharge flow machine 224 is operable to drive a machine load 226 through, for example, the machine drive 228 described above with reference to Figure 2 A machine outlet 225 is provided and is separated from further components by one or more valves. It will be appreciated that a pressure relief valve or bleed valve can be provided independently or can be incorporated into the one or more valves depicted in order to ensure safe operation of the discharge flow machine 224. The valves depicted include an aircraft flow valve 232 connecting the discharge flow path 210 to the aircraft flow assembly and a starter assembly valve 236 connecting the discharge flow path 210 to the starter assembly 234. Figure 3 The configuration of Figure 2 differed from that shown in that a second heat exchange assembly 240 was provided between the discharge flow machine 224 and the aircraft flow assembly 230. For example, the aircraft flow assembly 230 can include one or more components that require further control of the flow properties beyond that described with reference to the discharge flow machine 224 and the first heat exchange assembly 238. However, it will be appreciated that the first heat exchange assembly 238 and the second heat exchange assembly 240 can instead be provided. For example, in a configuration in which the discharge flow machine 224 is configured for efficient and safe high temperature operation, the second heat exchange assembly 240 can be provided alone with the omission of the first heat exchange assembly 238.

[0067] Turning now to Figure 4 a schematic view of a discharge assembly 300 for a gas turbine engine according to another example embodiment of the disclosure is provided. Figure 4 The example gas turbine engine of Figure 1 may be configured in substantially the same manner as the example gas turbine engine 10 described above with reference to Figure 4 For example, Figure 4 The example discharge assembly 300 of Figure 2 and Figure 3 differed from the example of the embodiment of

[0068] Reference is made to the first discharge flow path 310 and, as above with reference to 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 assembly, the secondary pressure source 316 upstream of the first discharge flow path 310 can operate as a purge or bleed of pressure from the first flow header 312 or another source. The secondary header valve 320 can be operable to control flow into or out of the secondary pressure source 316 through the secondary flow header 318. As described with reference to Figure 2 The first header valve 314 can independently control flow between the first flow header 312 and the first discharge flow path 310. In certain embodiments, the first header valve 314 can be a check valve configured to prevent backflow. It should be further understood that the second discharge flow path 311 can be split in a similar manner and can further include such purge or bleed facilities.

[0069] The first flow header 312 and the secondary flow header 318 can be connected to the first discharge flow path 310 through a discharge entry valve 322. The discharge entry valve 322 can operate as a master controller for flow through the first discharge flow path 310. Downstream of the discharge entry 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 properties 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 heat in the discharge flow prior to entry into a first aircraft flow assembly 346 through a 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, can be liquid cooled, and / or can be evaporative cooled with 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 referring to Figure 4 , the first discharge flow path 310 and the second discharge flow path 311 are each depicted as tapping from the turbine 102. For example, the turbine 102 can be in direct fluid communication with the first flow header 312 and the second flow header 313. As depicted, the first flow header 312 and the second flow header 313 tap from first and second flow sources 302, 303, respectively. These first and second flow sources 302, 303 can include various components, such as compressor sections of the turbine 102 (see Figures 1-3 ) In embodiments, the first and second flow sources 302, 303 are the same source. For example, the first flow header 312 and the second flow header 313 can each tap from the HP compressor 24 (see Figures 1-3suction, and can each even be from the same stage, such as the fourth stage of the HP compressor 24. However, it should also be appreciated that the first flow source 302 can represent a different pressure source within the turbine than the second flow source 303. For example, the first flow source 302 can be from a component upstream or downstream of the second flow source 303 relative to the core flow through the turbine 102. Although Figure 4 not depicted in FIG. 3, it should also be appreciated that at least one of the first flow source 302 and the second flow source 303 can be provided external to the turbine 102, for example as a compressor discharge source as described above with reference to FIG. 2. Figure 2

[0071] The separation of 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 appreciated 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 control downstream flow in the first discharge flow path 310 and the second discharge flow path 311, respectively.

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

[0073] ​The second bleed flow path 311 can be operated to control the second aircraft flow assembly 348. By decoupling control of the first aircraft flow assembly 346 from control of the second aircraft flow assembly 348, the first bleed flow path 310 and the second bleed flow path 311 can each be configured to efficiently handle its respective task. For example, the second aircraft flow assembly 348 can be configured as an ice protection assembly, such as a WAI as described above, and can require a relatively higher mass flow rate of bleed flow as compared to the first aircraft flow assembly 346. In view of this difference, various features of the second bleed flow path 311 can be configured differently than corresponding features of the first bleed flow path 310. For example, the second flow source 303 can be a relatively hotter and / or higher pressure source as compared to the first flow source 302. In an embodiment, the second flow source 303 is from a more downstream stage of the same component as compared to the first flow source 302. In another embodiment, the second flow source 303 is from a more downstream component as compared to the first flow source 302.

[0074] In the second bleed flow path 311, a second heat exchange assembly 340 is provided. The second heat exchange assembly 340 can be configured differently as compared to the first heat exchange assembly 338. For example, the second heat exchange assembly 340 can be relatively larger as compared to the first heat exchange assembly 338. In an embodiment, 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 can employ only 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 bleed flow path 311 in much the same way as the first heat exchange assembly 338 is in fluid communication with the first bleed flow path 310. For example, as shown in Figure 4 the second heat exchange assembly 340 is connected to the second bleed flow path 311 by a second heat exchange valve 341. Further as shown in Figure 4 the second heat exchange assembly 340 is connected to the second aircraft flow assembly 348 by a second flow outlet 347.

[0076] Turning now to Figure 5 a schematic view of a bleed assembly 400 for a gas turbine engine according to another example embodiment of the present disclosure is provided. Figure 5 The example gas turbine engine of Figure 1 may be configured in substantially the same manner as the example gas turbine engine 10 described above with reference to Figure 5 The example bleed assembly of Figure 4 may be configured similarly as in Figure 5The exemplary bleed assembly 400 generally includes the turbine 102 connected to a first bleed flow path 410 through a first flow tap 412, and a second bleed flow path 411 connected to the turbine 102 through a second flow tap 413. Figure 5 Embodiments of the present disclosure differ from Figure 4 Embodiments of the present disclosure differ from

[0077] The second flow tap 413 is depicted as tapping from the LP turbine 30, however, it can be provided further at various other locations. For example, the second flow tap 413 can be configured as a scoop disposed downstream of the combustion section 26. In various embodiments, the second flow tap 413 includes a turbine aft frame scoop and / or a turbine center frame scoop.

[0078] In a manner related to the energy captured from the bleed flow by the bleed flow machine 74, 124, 224 (see Figures 1-3 ), energy can be captured from further bleed flow paths, such as the second bleed flow path 411, to reduce operability issues caused by the bleed air at the compressor section with the first bleed flow path 410. Thus, bleed air from downstream of the combustion section 26 can be used to reduce bleed requirements from upstream of the combustion section 26. As discussed above with reference to Figure 4 , the first bleed flow path 410 and the second bleed flow path 411 can be configured to manage their flow sources and associated bleed flow components requirements. For example, Figure 5 The first bleed flow path 410 depicted in FIG. 4A taps from the HP compressor 24 and feeds to a first aircraft flow assembly 446, which can be a cabin assembly, such as a cabin ECS as described above. Figure 5 The second bleed flow path 411 depicted in FIG. 4A taps from the LP turbine 30 and feeds to a second aircraft flow assembly 448, which can be an anti-icing assembly, such as a WAI as described above.

[0079] Referring to the first bleed flow path 410, and as discussed above with reference to 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 purge or bleed of pressure from the first flow header 412 or another source. The secondary header valve 420 can be operable to control flow into or out of the secondary pressure source 416 through the secondary flow header 418. As described with reference to Figure 2 The first header valve 414 can independently control flow between the first flow header 412 and the first discharge flow path 410. In certain embodiments, the first header valve 414 can be a check valve configured to prevent backflow. A second header valve 415 can be provided to control flow between the second flow header 413 and the second discharge flow path 411. It should be further understood that the second discharge flow path 411 can be split in a similar manner to the first discharge flow path 410 and can further include such purge or bleed facilities.

[0080] The first flow header 412 and the secondary flow header 418 can be connected to the first discharge flow path 410 through a discharge entry valve 422. The discharge entry valve 422 can operate as a master controller for flow through the first discharge flow path 410. Downstream of the discharge entry 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 properties 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 heat in the discharge flow prior to entry into a first aircraft flow assembly 446 through a 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, can be liquid cooled, and / or can be evaporative cooled with 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 referring to Figure 5 , both the first discharge flow path 410 and the second discharge flow path 411 are depicted as tapping from the turbine 102. As briefly depicted and described above, the first flow header 412 and the second flow header 413 tap from a first flow source upstream of the combustion section 26 relative to the core flow and a second flow source downstream of the combustion section 26 relative to the core flow, respectively. Although not depicted in Figure 5 , it should also be understood that at least one of the first flow header 412 and the second flow header 314 can tap from outside of the turbine 102, such as described above with reference to Figure 2A compressor discharge source is described. As noted above, Figure 5 Embodiments can be configured such that one or both of the first and second discharge flow paths 410, 411 include a pressure source from outside the turbomachine 102, such as the second pressure source 416.

[0082] The separation of the first flow source (shown in Figure 5 as the HP compressor 24) and the second flow source (shown in Figure 5 as the LP turbine 30) facilitates the separation of the first and second discharge flow paths 410, 411. By separating the first and second discharge flow paths 410, 411, various discharge flow properties can be defined with respect to one another. For example, a relatively higher pressure and temperature flow from the second discharge flow path 411 can be expected, the second discharge flow path 411 taking flow from downstream of the combustion section 26. Additionally, different flow processing can be required due to the combustion products in the second discharge flow path 411. For example, the second discharge flow path 411 can be preferably used for systems that are less sensitive to combustion products, such as WAI and / or NAI operation.

[0083] A first mass flow rate through the first discharge flow path 410 can be defined with respect to a second mass flow rate through the second discharge flow path 411. In various embodiments, the second mass flow rate can be greater than the first mass flow rate, for example due to the 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 discharge flows within the respective discharge flow paths 410, 411. For example, the first discharge flow path 410 can have a relatively lower temperature compared to the temperature of the second discharge flow path 411. As with the mass flow rates, these relative temperature differences can also come from the respective supply sources, particularly whether downstream or upstream of the combustion section 26.

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

[0085] Still referring to Figure 5, the first flow outlet 445 of the first exhaust flow path 410 is shown to feed a first aircraft flow assembly 446 downstream of the first heat exchange assembly 438. As described above with reference to Figure 4 Figure 5 Embodiments of the present disclosure separate the first aircraft flow assembly 446 from the combustion products by maintaining separation between the first exhaust flow path 410 and the second exhaust flow path 411. As described above with reference to

[0086] Figure 5 The second flow outlet 447 of the second exhaust flow path 411 feeds a second aircraft flow assembly 448 through a second heat exchange assembly 440 controlled by a second heat exchange assembly valve 441. As described above with reference to Figure 4

[0087] The second aircraft flow assembly 448 can also be configured to safely manage more heat than the first aircraft flow assembly 446. For example, the second aircraft flow assembly 448 can be configured to have a relatively large flow dispersion volume, as with WAI arrangements. More heat resistant materials can also be employed in the second aircraft flow assembly 448 relative to the first aircraft flow assembly 446.

[0088] Turning now to Figure 6 , a schematic view of an exhaust assembly 500 for a gas turbine engine according to another example embodiment of the present disclosure is provided. Figure 6 The example gas turbine engine of the present disclosure can be configured in a manner similar to the above-referenced Figure 1 ​​The exemplary turbofan engine 10 is constructed in substantially the same manner as Figure 6 The exemplary bleed assembly 500 can be constructed in a similar manner as Figure 2 The exemplary bleed assembly 100 and / or Figure 3 The exemplary bleed assembly 200. For example, Figure 6 The exemplary bleed assembly 500 generally includes a turbine 102 connected to a bleed flow path 510 by a flow tap 512.

[0089] Figure 6 The embodiment depicted in Figure 2 and Figure 3 differs from the embodiments in Figure 2 and Figure 3 in that a bleed regulator 550 is provided. As shown, the bleed regulator 550 is provided downstream of the turbine 102 and upstream of a bleed flow machine 524. The flow tap 512 is in fluid communication with the bleed regulator 550 through a tap valve 514 and a bleed inlet valve 522. It should be appreciated that various other flow configurations to the bleed regulator 550 can be provided. For example, the secondary pressure source 516 can further feed the bleed regulator 550 through a secondary tap 518 controlled by a secondary tap valve 520, as described above with reference to and

[0090] As generally described above, a relatively high bleed flow can enable operations such as effective de-icing in adverse conditions and / or with a relatively large surface area to be de-iced. As Figure 6 shown, a machine outlet 525 from the bleed flow machine 524 can feed various bleed flow components. As shown, the bleed flow path 510 downstream of the bleed flow machine 524 extends to an aircraft flow assembly 530 through an aircraft flow valve 532. The aircraft flow assembly 530 can include various components such as WAI or cabin ECS components.

[0091] The bleed flow machine 524 can be configured to feed such components in a similar manner as described above with reference to various embodiments in Figure 2 and Figure 3 For example, the bleed flow machine 524 can be configured to drive a machine load 526 through a machine drive 528, where the machine load 526 is operable to provide torque to the turbine 102. As described above, the bleed flow machine 524 can be operable to enable the turbine to operate within a desired operability range at high bypass flow demands. However, there can also be instances where the bleed flow machine 524 does not provide such benefits, such as when bleed flow demands are relatively low or when there is a large operability margin in the turbine 102. In such instances, the bleed regulator 550 can control the output of the bleed flow machine 524 and, in turn, 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] The flow tap 612 and the secondary tap 618 can be connected to the discharge flow path 610 through a discharge entry valve 622. The discharge entry valve 622 can operate as a primary controller for flow through the discharge flow path 610. Downstream of the discharge entry 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 properties of the discharge flow prior to entry into a discharge flow machine 624 through a machine inlet 623. For example, the first heat exchange assembly 638 can be used to reduce heat in the discharge flow prior to entry into 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 evaporative cooled with a refrigerant cycle.

[0099] Prior to entry into the first heat exchange assembly 638, one or more discharge flow components can be configured to receive flow prior to any heat exchange operations being performed. For example, certain discharge flow components can benefit from a relatively high temperature. In the illustrated embodiment, and generally with reference to Figure 3 As discussed, a pre-exchange outlet 642 can 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 a resistance component 644.

[0100] Still referring to Figure 7 The first heat exchange assembly 638 can be supplemented or replaced by a second heat exchange assembly 640. As described above, the second heat exchange assembly 640 can also be provided independently of the first heat exchange assembly 638. The heat exchange assembly is selected if any, 638, 640, can be made based at least on capabilities of the discharge flow machine 624 and / or a desired output of the machine load 626 through the discharge flow transmission 628 to the discharge flow machine 624.

[0101] As described above with reference to Figure 6 The machine outlet 625 and / or the discharge conditioner 650 can supply various discharge flow components. As depicted, the downstream flow is supplied to an aircraft flow assembly 630 through an aircraft flow valve 632 and to a starter assembly 634 through a starter assembly valve 636. However, the second heat exchange assembly 640 shown in Figure 7 The second heat exchange assembly 640 can be provided upstream of the aircraft flow assembly 630, in embodiments. The second heat exchange assembly 640 can be provided upstream of only a portion of the aircraft flow assembly (e.g., a cabin ECS component of such assembly), while downstream of another portion (e.g., a WAI component of such assembly) or on a separate branch relative to another portion, in embodiments.

[0102] As also described above with reference to Figure 6 The turbine 102 can be provided downstream of the starter assembly 634, in embodiments.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 8The method of 703 can include directing the exhaust stream through a machine outlet in fluid communication with an aircraft flow assembly. The machine outlet can be any of the machine outlets 125, 225, 525, 625 described above, or can be configured in another manner, such as by combining features of multiple ones of those machine outlets. Likewise, the aircraft flow assembly can be any of the aircraft flow assemblies 130, 230, 530, 630 described above, or can be configured in another manner, such as by combining features of multiple ones of those aircraft flow assemblies. In various embodiments, the aircraft flow assembly of 703 can include a first aircraft flow assembly and a second aircraft flow assembly, such as to provide a WAI and a cabin ECS as described in greater detail above.

[0106] Still referring to Figure 8 The method of 703 can include directing the exhaust stream through a machine outlet can include, at least under certain operating conditions, directing a given amount of the exhaust stream to the aircraft flow assembly. For example, as described above, the exhaust mass flow rate of the exhaust stream can be at least twelve percent (12%) or at least twenty percent (20%) of the core mass flow rate, such as when data indicative of ice protection conditions is received.

[0107] Referring now to Figure 9 a flowchart of a method for operating a gas turbine engine according to a second example aspect of the present disclosure is provided. Figure 9 The method of 801 can be used with one or more example gas turbine engines described herein (e.g., the example gas turbine engine of Figure 4 or Figure 5 However, in other example aspects, Figure 9 The method of 801 can instead be used with any other suitable gas turbine engine.

[0108] Figure 9 The method illustrated includes, at 801, operating a turbomachine having a core flow therethrough. The turbomachine used in this method can be the example turbomachine 102 or can be configured in various other manners. At 802, a first exhaust stream is received with a first heat exchange assembly. At 804, a second exhaust stream is received with a second heat exchange assembly. It will be appreciated that the first and second heat exchange assemblies can correspond to those described above, such as any of the heat exchange assemblies 238, 240, 338, 340, 438, 440, 638, 640; or can be configured as any other suitable heat exchange assembly.

[0109] Still referring to Figure 9The method of 803, the first bleed stream is directed to a first aircraft flow assembly. At 805, the second bleed stream is directed to a second aircraft flow assembly. The aircraft flow assemblies at 803 and 805 can be any of the aircraft flow assemblies 130, 230, 530, 630 as described above, or can be constructed in another manner, e.g., combining features of multiple those aircraft flow assemblies. In various embodiments, the first aircraft flow assembly at 803 can include a cabin ECS, and the second aircraft flow assembly at 805 can include a WAI and / or NAI system, as described in more detail above.

[0110] Referring now to Figure 10 , a flow diagram of a method for operating a gas turbine engine according to a third example aspect of the disclosure is provided. Figure 10 The method of 803 can be used with one or more example gas turbine engines described herein (e.g., the example gas turbine engine of 801, 802, or 804). However, in other example aspects, Figure 4 or Figure 5 the example gas turbine engine of 801, 802, or 804). However, in other example aspects, Figure 10 the method of 803 can instead be used with any other suitable gas turbine engine.

[0111] Figure 10 The method shown includes, at 901, operating a turbomachine having a core flow therethrough. The turbomachine used in this method can be the example turbomachine 102 or can be constructed in various other manners. At 902, a first bleed stream from upstream of a combustion section is received with a first flow tap. At 904, a second bleed stream from downstream of the combustion section is received with a second flow tap. It will be appreciated that the first and second flow taps can correspond to any of those described above, e.g., any of the flow taps 112, 212, 312, 313, 412, 413, 512, 612; or can be constructed as any other suitable flow tap.

[0112] Still referring to Figure 10 the method of 803, at 903, the first bleed stream is directed from the first flow tap with a first flow outlet. At 905, the second bleed stream is directed from the second flow tap with a second flow outlet. The flow outlets at 903 and 905 can be any of the flow outlets 345, 347, 445, 447 as described above, or can be constructed in another manner, e.g., combining features of multiple those aircraft flow assemblies.

[0113] Still referring to Figure 10the method of 906, the method provides receiving a first exhaust stream and a second exhaust stream with 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 of those aircraft flow assemblies. In various embodiments, the at least one aircraft flow assembly at 906 can include a first aircraft flow assembly and a second aircraft flow assembly as described in greater detail above, for example with reference to Figure 9 the method of 906.

[0114] Reference is now made to Figure 11 , a flow diagram of a method for operating a gas turbine engine is provided in accordance with a fourth example aspect of the present disclosure. Figure 11 the method of 906 can be used with one or more example gas turbine engines described herein (e.g., the example gas turbine engines of Figure 6 or Figure 7 However, in other example aspects, Figure 11 the method of 906 can instead be used with any other suitable gas turbine engine.

[0115] Figure 11 The method shown includes, at 1001, receiving a portion of a core flow, an exhaust stream, with a flow tap, such as from the example turbine 102 described above or other suitable turbine or pressure source. It should be appreciated that the flow tap at 1001 can correspond to one or more of those described above, such as any of the flow taps 112, 212, 312, 313, 412, 413, 512, 612; or can be constructed as any other suitable flow tap.

[0116] Still referring to Figure 11 , at 1002, the method provides driving a machine load with the exhaust output through an exhaust flow machine. The machine load at 1002 can be any of the machine loads 126, 226, 526, 626 as described above, or can be constructed in another manner, such as combining features of multiple of those machine loads. Likewise, the exhaust flow machine used in the method can be any of the exhaust flow machines 124, 224, 524, 624, or can be constructed in another manner, such as combining features of multiple of those exhaust flow machines. As described above, it should be appreciated that the exhaust flow machine at 1002 is constructed to receive the exhaust stream from 1001.

[0117] Still referring to Figure 11 , at 1003, the method provides adjusting the exhaust output driving the machine load at 1002 with an exhaust regulator. It should be appreciated that the exhaust regulator at 1003 can be one of the exhaust regulators 550, 650; or can be constructed as described with reference toFigure 6 and Figure 7 The control bleed flow machine can be configured in other ways as described in greater detail. As such, the capture rate of the bleed flow by the control bleed flow machine can include any of the above configurations, such as the variable adjustment of the stator or the variable adjustment of the exit area of the bleed flow machine's bleed air expansion turbine configuration at 1003.

[0118] This written description uses examples to disclose the disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the disclosure is defined by the claims, and can include other examples that occur to those skilled in the art. Such other examples are intended to fall within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

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

[0120] A gas turbine engine comprising: a turbomachine comprising, in serial flow order, a compressor section, a combustion section, and a turbine section, the turbomachine defining a core flow through therewith during operation, the core flow defining a core mass flow rate; a bleed assembly comprising a bleed flow machine and a machine load, the bleed flow machine being in fluid communication with the compressor section of the turbomachine and being configured to drive the machine load; and a machine outlet in fluid communication with the bleed assembly, the machine outlet defining a bleed flow through therewith during operation of the gas turbine engine, the bleed flow defining a bleed mass flow rate, wherein the compressor section is configured to provide the bleed flow to an aircraft flow assembly through the bleed flow machine and the machine outlet, wherein the bleed mass flow rate is at least twelve percent (12%) of the core mass flow rate.

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

[0122] The gas turbine engine of any of the preceding clauses, wherein the machine outlet defines the bleed flow therethrough during an aircraft wing ice accretion operation, wherein during the aircraft wing ice accretion operation the bleed mass flow rate is at least twelve percent (12%) of the core mass flow rate, and wherein the aircraft flow assembly comprises a wing ice protection assembly.

[0123] The gas turbine engine of any of the preceding clauses, wherein the bleed mass flow rate is at least twenty percent (20%) of the core mass flow rate.

[0124] The gas turbine engine of any of the preceding clauses, wherein the aircraft flow assembly comprises an aircraft environmental control system, a wing ice protection assembly, or both.

[0125] The gas turbine engine of any of the preceding clauses, wherein the machine load comprises a drivable mechanical connection to the turbine.

[0126] The gas turbine engine of any of the preceding clauses, wherein the machine load comprises a generator configured to transmit electrical power to an electric machine configured to drive the turbine.

[0127] The gas turbine engine of any of the preceding clauses, wherein the gas turbine engine further comprises a first heat exchange assembly disposed in serial flow order between the turbine and the exhaust flow machine.

[0128] The gas turbine engine of any of the preceding clauses, wherein the gas turbine engine further comprises a second heat exchange assembly disposed in serial flow order between the exhaust flow machine and the aircraft flow assembly.

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

[0130] The gas turbine engine of any of the preceding clauses, wherein the gas turbine engine further comprises an exhaust regulator configured to bypass the exhaust flow machine with at least a portion of the exhaust flow from the turbine to the machine outlet.

[0131] A method of operating a gas turbine engine, comprising: operating a turbine to provide a core flow through the turbine, the core flow defining a core mass flow rate, and the turbine comprising, in serial flow order, a compressor section, a combustion section, and a turbine section; exhausting 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 comprising 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 with 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 an aircraft flow assembly, wherein the exhaust mass flow rate is at least twelve percent (12%) of the core mass flow rate.

[0132] The method of any of the preceding clauses, wherein the method further comprises driving the turbine with the machine load.

[0133] The method of any of the preceding clauses, wherein the method further comprises: receiving data indicative of an aircraft anti-icing condition; and in response to receiving the data indicative of the aircraft anti-icing condition, directing at least twelve percent (12%) of the discharge mass flow rate of the discharge flow having the core mass flow rate to the aircraft flow assembly, wherein the aircraft flow assembly comprises a wing anti-icing assembly.

[0134] The method of any of the preceding clauses, wherein the discharge mass flow rate is at least twenty percent (20%) of the core mass flow rate.

[0135] The method of any of the preceding clauses, wherein the method further comprises: driving the turbine with the machine load through a mechanical connection.

[0136] The method of any of the preceding clauses, wherein the method further comprises: generating electrical power with a generator of the machine load; transmitting electrical power from the generator to an electric machine; and driving the turbine with the electric machine.

[0137] The method of any of the preceding clauses, wherein the method further comprises: cooling the discharge flow with a first heat exchange assembly disposed in serial flow order between the turbine and the discharge flow machine.

[0138] The method of any of the preceding clauses, wherein the machine outlet is configured to provide at least a portion of the discharge flow to an air starter assembly.

[0139] The method of any of the preceding clauses, wherein the method further comprises: using a discharge regulator to bypass the discharge flow machine with at least a portion of the discharge flow from the turbine to the machine outlet.

[0140] A gas turbine engine comprising: a turbomachine comprising, in serial flow order, a compressor section, a combustion section, and a turbine section, the turbomachine defining a core flow therethrough during operation; a first heat exchange assembly in fluid communication with the turbomachine for receiving a first discharge flow from the turbomachine; a first flow outlet in fluid communication with the first heat exchange assembly for receiving the first discharge flow from the first heat exchange assembly and providing the first discharge flow to a first aircraft flow assembly; a second heat exchange assembly in fluid communication with the turbomachine for receiving a second discharge flow from the turbomachine; and a second flow outlet in fluid communication with the second heat exchange assembly for receiving the second discharge flow from the second heat exchange assembly and providing the second discharge flow to a second aircraft flow assembly.

[0141] The gas turbine engine of any of the preceding clauses, wherein the first aircraft flow assembly comprises a cabin environmental control assembly and the second aircraft flow assembly comprises an ice protection assembly.

[0142] The gas turbine engine of any of the preceding clauses, wherein the ice protection assembly comprises a wing ice protection assembly.

[0143] The gas turbine engine of any of the preceding clauses, wherein the ice protection assembly comprises a nacelle ice protection assembly.

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

[0145] The gas turbine engine of any of the preceding clauses, wherein the first heat exchange assembly is configured to receive the first discharge flow from a high pressure compressor of the compressor section of the turbomachine.

[0146] The gas turbine engine of any of the preceding clauses, wherein the second heat exchange assembly is configured to receive the second discharge flow from a low pressure turbine of the turbine section of the turbomachine.

[0147] The gas turbine engine of any of the preceding clauses, wherein the first discharge flow defines a first mass flow rate during operation of the gas turbine engine, and the second discharge 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] The gas turbine engine of any of the preceding clauses, wherein the second mass flow rate is at least twice the first mass flow rate.

[0149] The gas turbine engine of any of the preceding clauses, wherein the first flow outlet comprises a first cross-sectional area and the second flow outlet comprises a second cross-sectional area, wherein the first cross-sectional area is greater than the second cross-sectional area.

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

[0151] The gas turbine engine of any of the preceding clauses, wherein the first heat exchange assembly comprises a first heat exchange volume and the second heat exchange assembly comprises a second heat exchange volume, the second heat exchange volume being greater than the first heat exchange volume.

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

[0153] The method of any of the preceding clauses, wherein the first aircraft flow assembly comprises a cabin environmental control assembly and the second aircraft flow assembly comprises an ice protection assembly.

[0154] The method of any of the preceding clauses, wherein the method further comprises: receiving, with the first heat exchange assembly, the first exhaust flow from upstream of the combustion section of the turbomachine; and receiving, with the second heat exchange assembly, the second exhaust flow from downstream of the combustion section of the turbomachine.

[0155] The method of any of the preceding clauses, wherein the method further comprises: receiving, with the first heat exchange assembly, the first exhaust flow from a high pressure compressor of the compressor section of the turbomachine.

[0156] The method of any of the preceding clauses, wherein the method further comprises: receiving, with the second heat exchange assembly, the second exhaust flow from a low pressure turbine of the turbine section of the turbomachine.

[0157] The method of any of the preceding clauses, wherein the first discharge stream comprises a first mass flow rate and the second discharge stream comprises a second mass flow rate, wherein the second mass flow rate is greater than the first mass flow rate.

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

[0159] The method of any of the preceding clauses, wherein the first heat exchange assembly comprises a first heat exchange volume and the second heat exchange assembly comprises a second heat exchange volume, the second heat exchange volume being greater than the first heat exchange volume.

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

[0161] The gas turbine engine of any of the preceding clauses, wherein the first flow outlet is configured to direct the first discharge stream to a first aircraft flow assembly and the second flow outlet is configured to direct the second discharge stream to a second aircraft flow assembly.

[0162] The gas turbine engine of any of the preceding clauses, wherein the first aircraft flow assembly comprises a cabin environmental control assembly, and wherein the second aircraft flow assembly comprises an anti-ice assembly.

[0163] The gas turbine engine of any of the preceding clauses, wherein the anti-ice assembly is a wing anti-ice assembly.

[0164] The gas turbine engine of any of the preceding clauses, wherein the anti-ice assembly is a nacelle anti-ice assembly.

[0165] The gas turbine engine of any of the preceding clauses, wherein the gas turbine engine further comprises: a first heat exchange assembly configured to receive the first discharge stream from the first flow header; and a second heat exchange assembly configured to receive the second discharge stream from the second flow header.

[0166] The gas turbine engine of any of the preceding clauses, wherein the first flow tap is configured to receive the first discharge flow from a high pressure compressor of the compressor section of the turbine engine.

[0167] The gas turbine engine of any of the preceding clauses, wherein the second flow tap is configured to receive the second discharge flow from a low pressure turbine of the turbine section of the turbine engine.

[0168] The gas turbine engine of any of the preceding clauses, wherein the first discharge flow from upstream of the combustion section defines a first mass flow rate during operation, and the second discharge flow 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] The gas turbine engine of any of the preceding clauses, wherein the second mass flow rate is at least twice the first mass flow rate.

[0170] The gas turbine engine of any of the preceding clauses, wherein the first flow outlet includes a first cross-sectional area, and the second flow outlet includes a second cross-sectional area, wherein the first cross-sectional area is greater than the second cross-sectional area.

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

[0172] The gas turbine engine of any of the preceding 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 greater than the first heat exchange volume.

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

[0174] The method of any of the preceding clauses, wherein the method further comprises: directing, with the first flow outlet, the first discharge flow from the first flow tap to a first aircraft flow assembly; directing, with the second flow outlet, the second discharge flow from the second flow tap to a second aircraft flow assembly.

[0175] The method of any of the preceding clauses, wherein the method further comprises: receiving, with the first flow tap, the first discharge flow from a high pressure compressor of the compressor section of the turbine engine.

[0176] The method of any of the preceding clauses, wherein the method further comprises: receiving, with the second flow tap, the second discharge flow from a low pressure turbine of the turbine section of the turbine engine.

[0177] The method of any of the preceding clauses, wherein the first aircraft flow assembly comprises a cabin environmental control assembly and the second aircraft flow assembly comprises an anti-ice assembly.

[0178] The method of any of the preceding clauses, wherein the anti-ice assembly is a wing anti-ice assembly.

[0179] The method of any of the preceding clauses, wherein the anti-ice assembly is a nacelle anti-ice assembly.

[0180] A gas turbine engine comprising: a turbine engine comprising, in serial flow order, a compressor section, a combustion section, and a turbine section, the turbine engine defining a core flow therethrough during operation; a flow tap in fluid communication with the turbine engine, the flow tap configured to receive a portion of the core flow therethrough as a discharge flow; and a discharge assembly comprising: a machine load; a discharge flow machine disposed in fluid communication with the turbine engine through the flow tap, the discharge flow machine configured to drive the machine load; and a discharge regulator configured to regulate a discharge output provided to the discharge flow machine by controlling a capture rate of the discharge flow by the discharge flow machine.

[0181] The gas turbine engine of any of the preceding clauses, wherein the gas turbine engine further comprises a flow outlet downstream of the discharge flow machine, wherein the discharge regulator is further configured to control the capture rate of the discharge flow by the discharge flow machine while maintaining a mass flow rate of the discharge flow through the flow outlet.

[0182] The gas turbine engine of any of the preceding clauses, wherein the bleed regulator comprises: a split flow path fluidly connecting the flow tap and the flow outlet and bypassing the bleed flow machine; and a split flow valve disposed upstream of the bleed flow machine, the split flow valve configured to control a split flow through the split flow path.

[0183] The gas turbine engine of any of the preceding clauses, wherein the bleed regulator comprises at least one variable bleed feature configured to control the capture rate of the bleed flow machine on the bleed flow.

[0184] The gas turbine engine of any of the preceding clauses, wherein the at least one variable bleed feature is configured as a component of the bleed flow machine, the component configured to regulate a flow out of the bleed flow machine.

[0185] The gas turbine engine of any of the preceding clauses, wherein the bleed flow comprises a bleed flow mass flow rate and the core flow comprises a core mass flow rate, wherein the bleed mass flow rate is at least twelve percent (12%) of the core mass flow rate.

[0186] The gas turbine engine of any of the preceding clauses, wherein the gas turbine engine further comprises an aircraft flow assembly in fluid communication with the bleed assembly, wherein the flow assembly comprises at least one of: a wing ice protection assembly; a nacelle ice protection assembly; or a cockpit environmental control assembly.

[0187] The gas turbine engine of any of the preceding clauses, wherein the machine load comprises a drivable mechanical connection with the turbomachine.

[0188] The gas turbine engine of any of the preceding clauses, wherein the machine load comprises a generator configured to transmit electrical power to an electric machine configured to drive the turbomachine.

[0189] The gas turbine engine of any of the preceding clauses, wherein the gas turbine engine further comprises a first heat exchange assembly disposed in serial flow order between the turbomachine and the bleed flow machine.

[0190] The gas turbine engine of any of the preceding clauses, wherein the gas turbine engine further comprises a second heat exchange assembly disposed in serial flow order between the bleed flow machine and the flow outlet.

[0191] The gas turbine engine of any of the preceding clauses, wherein the flow outlet is configured to provide at least a portion of the discharge flow to an air starter assembly.

[0192] A method of operating a gas turbine engine including a turbomachinery having a core flow therethrough, the turbomachinery including, in serial flow order, a compressor section, a combustion section, and a turbine section, the method including: receiving, with a flow tap in fluid communication with the turbomachinery, a portion of the core flow defining a discharge flow; driving a machine load with a discharge output of a discharge flow machine disposed in fluid communication with the turbomachinery through the flow tap; and regulating the discharge output with a discharge regulator by controlling a capture rate of the discharge flow machine on the discharge flow.

[0193] The method of any of the preceding clauses, wherein the method further includes controlling, with the discharge regulator, the capture rate of the discharge flow machine on the discharge flow while maintaining a discharge mass flow rate of the discharge flow through a flow outlet disposed downstream of the discharge flow machine.

[0194] The method of any of the preceding clauses, wherein the method further includes controlling, with a split flow valve disposed upstream of the discharge flow machine, a split flow through a split flow path fluidly connecting the flow tap and the flow outlet and bypassing the discharge flow machine.

[0195] The method of any of the preceding clauses, wherein the method further includes controlling, with at least one variable discharge feature of the discharge regulator, the capture rate of the discharge flow machine on the discharge flow.

[0196] The method of any of the preceding clauses, wherein the at least one variable discharge feature is configured as a component of the discharge flow machine.

[0197] The method of any of the preceding clauses, wherein the method further includes regulating, with the at least one variable discharge feature, a fluid flow out of the discharge flow machine.

[0198] The method of any of the preceding clauses, wherein the discharge flow includes a discharge flow mass flow rate and the core flow includes a core mass flow rate, wherein the discharge mass flow rate is at least twelve percent (12%) of the core mass flow rate.

[0199] The method of any of the preceding clauses, wherein the method further includes generating electrical power with a generator of the machine load; transmitting the electrical power to an electric machine; and driving the turbomachinery with the electric machine.

Claims

1. A gas turbine engine characterized by, Comprising: a turbine comprising, in serial 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 and a bleed flow path, the flow tap configured to receive a portion of the core flow therethrough as a bleed; and a secondary pressure source fluidically coupled to the flow tap via a secondary tap valve, wherein the secondary tap valve is configured to facilitate flow from the secondary pressure source into the bleed flow path and from the bleed flow path into the secondary pressure source; a bleed assembly comprising: a machine load; a bleed flow machine disposed in fluid communication with the turbine and the secondary pressure source through the flow tap, the bleed flow machine configured to drive the machine load; and a bleed regulator configured to regulate a bleed output provided to the bleed flow machine by controlling a capture rate of the bleed flow by the bleed flow machine.

2. The gas turbine engine of claim 1, wherein, further comprising a flow outlet downstream of the bleed flow machine, wherein the bleed regulator is further configured to control the capture rate of the bleed flow by the bleed flow machine while maintaining a mass flow rate of the bleed flow through the flow outlet.

3. The gas turbine engine of claim 2, wherein, wherein the flow outlet is configured to provide at least a portion of the bleed flow to an air starter assembly.

4. The gas turbine engine of claim 2, wherein, wherein the bleed regulator comprises: a split flow path fluidically connecting the flow tap and the flow outlet and bypassing the bleed flow machine; and a split flow valve disposed upstream of the bleed flow machine, the split flow valve configured to control a split flow through the split flow path.

5. The gas turbine engine of claim 2, wherein, wherein the bleed regulator comprises at least one variable bleed feature configured to control the capture rate of the bleed flow by the bleed flow machine.

6. The gas turbine engine of claim 5, wherein, wherein the at least one variable bleed feature is configured as a component of the bleed flow machine, the component configured to regulate a flow out of the bleed flow machine.

7. The gas turbine engine of claim 1, wherein, wherein the bleed flow comprises a bleed flow mass flow rate and the core flow comprises a core mass flow rate, wherein the bleed mass flow rate is at least twelve percent (12%) of the core mass flow rate.

8. The gas turbine engine of claim 1, wherein, further comprising an aircraft flow assembly in fluid communication with the bleed assembly, wherein the flow assembly comprises at least one of: a wing ice protection assembly; a nacelle ice protection assembly; or a cockpit environmental control assembly.

9. The gas turbine engine of claim 1, wherein, wherein the machine load comprises a drivable mechanical connection to the turbine.

10. The gas turbine engine of claim 1, wherein, wherein the machine load comprises a generator configured to transmit electrical power to an electric machine configured to drive the turbine.

11. The gas turbine engine of claim 1, wherein, further comprising a first heat exchange assembly disposed between the turbine and the bleed flow machine in serial flow order.

12. The gas turbine engine of claim 11, wherein, further comprising a second heat exchange assembly disposed between the bleed flow machine and a flow outlet downstream of the bleed flow machine in serial flow order.

13. A method of operating a gas turbine engine including a turbomachinery having a core flow therethrough, the turbomachinery including, in serial flow order, a compressor section, a combustion section, and a turbine section, characterized by, the method comprising: receiving a bleed stream from a flow tap in fluid communication with the turbine and the core flow, wherein the flow tap is fluidly coupled to a secondary pressure source, wherein the secondary pressure source is fluidly coupled to the flow tap via a secondary tap valve, wherein the secondary tap valve is configured to facilitate flow from the secondary pressure source into the bleed flow path and to facilitate flow from the bleed flow path into the secondary pressure source; driving a machine load with a bleed output of a bleed flow machine, the bleed flow machine disposed in fluid communication with the turbine through the flow tap; and regulating the bleed output with a bleed regulator by controlling a capture rate of the bleed flow by the bleed flow machine.

14. The method of claim 13, wherein, further comprising controlling the capture rate of the bleed flow by the bleed flow machine with the bleed regulator while maintaining a bleed mass flow rate of the bleed flow through a flow outlet, the flow outlet disposed downstream of the bleed flow machine.

15. The method of claim 14, wherein, further comprising controlling a bleed flow through a bleed flow path with a bleed flow valve disposed upstream of the bleed flow machine, the bleed flow path fluidly connecting the flow tap and the flow outlet and bypassing the bleed flow machine.

16. The method of claim 14, wherein, further comprising controlling the capture rate of the bleed flow by the bleed flow machine with at least one variable bleed feature of the bleed regulator.

17. The method of claim 16, wherein, wherein the at least one variable bleed feature is configured as a component of the bleed flow machine.

18. The method of claim 17, wherein, further comprising: regulating a fluid flow out of the bleed flow machine with the at least one variable bleed feature.

19. The method of claim 13, wherein, wherein the bleed flow comprises a bleed mass flow rate and the core flow comprises a core mass flow rate, wherein the bleed mass flow rate is at least twelve percent (12%) of the core mass flow rate.

20. The method of claim 13, wherein, further comprising: generating electrical power with a generator of the machine load; transmitting the electrical power to an electric machine; and driving the turbine with the electric machine.

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