Turbine engine with exhaust system

Abstract

A turbine engine includes a high pressure (HP) compressor and a discharge system. The HP compressor includes an HP compressor flow path and a plurality of stages of HP compressor rotor blades and HP compressor stator vanes. The discharge system includes a plurality of discharge flow paths including at least three discharge flow paths in fluid communication with the HP compressor flow path. A plurality of discharge flow paths direct compressed air from the HP compressor flow path. The at least two discharge flow paths are at successive stages of the plurality of stages.

Description

Technical Field

[0001] The present disclosure generally relates to a turbine engine having a bleed system, such as for a high pressure compressor of the turbine engine. Background Art

[0002] For example, a turbine engine for an aircraft generally includes a fan and a core section arranged in flow communication with each other. The core section includes a high pressure compressor. Some turbine engines include a bleed system at the high pressure compressor to extract or "bleed" air from the high pressure compressor for use at various other locations in the turbine engine or aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0003] The foregoing and other features and advantages will be apparent from the following more particular description of various exemplary embodiments, as illustrated in the accompanying drawings, in which like reference numerals generally indicate identical, functionally similar, or structurally similar elements.

[0004] Figure 1 is a schematic cross-sectional view of a turbine engine according to the present disclosure, taken along a longitudinal centerline axis of the turbine engine.

[0005] Figure 2 According to the present disclosure Figure 1 Details of 2-2 intercepted Figure 1 Schematic cross-sectional view of a high-pressure compressor of a turbine engine.

[0006] Figure 3 According to the operation of the present disclosure Figure 1 and 2 A flow chart of a method for a turbine engine.

[0007] Figure 4 is an operation according to another embodiment Figure 1 and Figure 2 A flow chart of a method for a turbine engine.

[0008] Figure 5 The following diagram according to the present disclosure is shown in the form of a graph. Figure 1 The first mass flow rate W through the first exhaust flow path varies with the altitude of the turbine engine first_bleed and the second mass flow rate W through the second exhaust flow path second_bleed mass flow ratio. DETAILED DESCRIPTION

[0009] Features, advantages and embodiments of the present disclosure are set forth or apparent through consideration of the following detailed description, drawings and claims.Furthermore, the following detailed description is exemplary and is intended to provide further explanation without limiting the scope of the claimed disclosure.

[0010] Various embodiments of the present disclosure are discussed in detail below. Although specific embodiments are discussed, this is for illustrative purposes only. Those skilled in the relevant art will recognize that other components and configurations may be used without departing from the spirit and scope of the present disclosure.

[0011] As used herein, the terms “first,” “second,” “third,” and “fourth” may be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0012] The terms "upstream" and "downstream" refer to relative directions relative to the flow of a fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction toward which the fluid is flowing.

[0013] The terms "front" and "rear" refer to relative positions within a turbine engine or vehicle, and refer to the normal operating attitude of the turbine engine or vehicle. For example, with respect to a turbine engine, front refers to a position closer to the engine inlet, while rear refers to a position closer to the engine nozzle or exhaust.

[0014] Unless otherwise specified herein, the terms "coupled," "fixed," "attached," "connected," and the like refer to both direct coupling, fixing, attachment, or connection as well as indirect coupling, fixing, attachment, or connection through one or more intermediate components or features.

[0015] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0016] As used herein, the terms "axial" and "axially" refer to directions and orientations extending substantially parallel to the centerline of the turbine engine. Additionally, the terms "radial" and "radially" refer to directions and orientations extending substantially perpendicular to the centerline of the turbine engine. Additionally, as used herein, the terms "circumferential" and "circumferentially" refer to directions and orientations extending in an arc around the centerline of the turbine engine.

[0017] As used herein, "consecutive stages" of a plurality of stages of a high pressure (HP) compressor are two or more stages that proceed sequentially or are in an uninterrupted order. For example, the consecutive stages may include the first and second stages, the second and third stages, the third and fourth stages, and so on.

[0018] A turbine engine, such as a turbine engine for an aircraft, may include an exhaust system for exhausting air from a compressor section of the turbine engine. The exhaust is used to cool the turbine engine (e.g., at least one of the high-pressure turbine or the low-pressure turbine) and the components of the aircraft (e.g., at least one of the environmental control system or the wing anti-icing system). The exhaust is usually exhausted from each stage of the HP compressor. In some exhaust systems, the exhaust from one stage is used for both the aircraft system and the low-pressure turbine. In the exhaust system, air is usually exhausted from the upstream stage of the HP compressor to avoid sacrificing the performance of the HP compressor. For example, the more downstream the air is exhausted (e.g., the more downstream stages there are), the greater the reduction in the performance of the HP compressor by the exhaust system. Using the same exhaust for both the aircraft system and the low-pressure turbine requires exhaust from the downstream stage (e.g., the fourth stage) of the HP compressor to provide exhaust at sufficient pressure to both the aircraft system and the low-pressure turbine. Exhausting air from the upstream stage (e.g., the third stage) of the HP compressor will help reduce the associated fuel combustion losses (e.g., reduce the amount of HP compressor performance reduction). However, the pressure that may be provided by the three-stage exhaust air from the nine-stage HP compressor is not enough to purge the low-pressure turbine under all operating conditions. As a result, hot gas path ingestion of hot gases through the rotor disk and into the cavity of the low pressure turbine may occur. Hot gas path ingestion may lead to premature component durability issues or even failure of components of the turbine engine.

[0019] Therefore, the present disclosure provides an improved exhaust system for a turbine engine. The exhaust system includes three exhaust flow paths from three stages of an HP compressor. At least two exhaust flow paths are in consecutive stages (e.g., the third and fourth stages) of the HP compressor to provide efficient exhaust to an aircraft system (e.g., at least one of an environmental control system or a wing anti-icing system). Providing at least two exhaust flow paths at consecutive stages in this manner allows the exhaust system to meet the required return margin and low-pressure turbine pressure ratio requirements of a specific turbine engine.

[0020] The exhaust system of the present disclosure provides a flexible seal and a rigid support member for separating the exhaust air chamber of the exhaust system so that the exhaust flow paths are all separated from each other. The exhaust flow path includes a first exhaust flow path to the aircraft system, a second exhaust flow path to the low-pressure turbine, and a third exhaust flow path to the high-pressure turbine of the turbine engine. In an exemplary embodiment, the first exhaust flow path is in the third stage, the second exhaust flow path is in the fourth stage, and the third exhaust flow path is in the sixth stage of the HP compressor. In some embodiments, the exhaust system includes a fourth exhaust flow path downstream of the diffuser of the HP compressor and upstream of the combustor of the turbine engine. Exhausting air from the third stage to provide air to the aircraft environmental control system (ECS) or wing anti-icing system (WAI) of the aircraft allows lower performance losses or lower fuel burn losses. The present disclosure provides exhaust air from the fourth stage of the HP compressor to maintain a positive backflow margin in the low-pressure turbine, thereby purging the low-pressure turbine and preventing the hot gas path in the low-pressure turbine from being inhaled. Thus, the HP compressor bleed system of the present disclosure provides exhaust gas to the aircraft from the third stage of the HP compressor to provide lower fuel burn penalties (e.g., compared to a bleed system without the benefits of the present disclosure) while maintaining durability and safety by utilizing the positive return flow margin provided by the higher pressure air from the fourth stage of the HP compressor.

[0021] In some embodiments, the exhaust from the third stage is directed to the low-pressure turbine, and the exhaust from the fourth stage is directed to the ECS or WAI. Therefore, the present disclosure provides for separating the exhaust to the ECS or WAI and the low-pressure turbine by discharging air from different stages of the HP compressor. This configuration allows independent control of the exhaust to the ECS or WAI and to the low-pressure turbine. The present disclosure also provides a method for operating a turbine engine to change the amount of exhaust to the ECS or WAI based on the altitude of the turbine engine. Compared to a turbine engine without the benefits of the present disclosure, changing the amount of exhaust to the ECS or WAI based on the altitude of the turbine engine helps to achieve the required amount of exhaust to the ECS or WAI, while also reducing the amount of compressed air that needs to be discharged from the HP compressor due to the separate flow paths to the ECS or WAI and to the low-pressure turbine. Therefore, changing the amount of exhaust to the ECS or WAI based on the altitude of the turbine engine helps to reduce the performance of the HP compressor sacrificed by discharging air from the HP compressor.

[0022] Now referring to the accompanying drawings, Figure 1 is a schematic cross-sectional view of a turbine engine 10 taken along a longitudinal centerline axis 12 of the turbine engine 10 according to an embodiment of the present disclosure. Figure 1As shown, the turbine engine 10 defines an axial direction A (extending parallel to a longitudinal centerline axis 12 for reference) and a radial direction R perpendicular to the axial direction A. Generally speaking, the turbine engine 10 includes a fan section 14 and a core turbine engine 16 disposed downstream of the fan section 14 .

[0023] The depicted core turbine engine 16 generally includes a casing 18 that is substantially tubular and defines an annular inlet 20. Figure 1 As schematically shown in FIG. 1 , the casing 18 encloses in series flow relationship: a compressor section 21, which includes a supercharger or low pressure (LP) compressor 22, followed downstream by a high pressure (HP) compressor 24; a combustion section 26; a turbine section 27, which includes a high pressure (HP) turbine 28, followed downstream by a low pressure (LP) turbine 30; and an injection exhaust nozzle section 32. A high pressure (HP) shaft 34 or spool drivingly connects the HP turbine 28 to the HP compressor 24 so that the HP turbine 28 and the HP compressor 24 rotate in unison. A low pressure (LP) shaft 36 drivingly connects the LP turbine 30 to the LP compressor 22 so that the LP turbine 30 and the LP compressor 22 rotate in unison. The compressor section 21, the combustion section 26, the turbine section 27, and the injection exhaust nozzle section 32 together define a core air flow path.

[0024] for Figure 1 In the embodiment depicted in FIG. 1 , fan section 14 includes a fan 38 (eg, a variable pitch fan) having a plurality of fan blades 40 coupled to a disk 42 in a spaced-apart manner. Figure 1 , fan blades 40 extend outwardly from disk 42 generally in a radial direction R. Each fan blade 40 is rotatable relative to disk 42 about pitch axis P because fan blade 40 is operably coupled to actuation member 44, which is configured to collectively and uniformly change the pitch of fan blade 40, as described in further detail below. Fan blades 40, disk 42, and actuation member 44 are rotatable together about longitudinal centerline axis 12 via fan shaft 45, which is powered by LP shaft 36 across a power gearbox (also referred to as gearbox assembly 46). Gearbox assembly 46 is Figure 1 The gearbox assembly 46 includes a plurality of gears for adjusting the rotational speed of the fan shaft 45 , thereby adjusting the rotational speed of the fan 38 relative to the LP shaft 36 .

[0025] Still refer to Figure 1In an exemplary embodiment of the present invention, disk 42 is covered by a rotatable fan hub 48 having an aerodynamic profile to facilitate airflow through a plurality of fan blades 40. In addition, fan section 14 includes an annular fan case or nacelle 50 that circumferentially surrounds fan 38 and at least a portion of core turbine engine 16. Nacelle 50 is supported relative to core turbine engine 16 by a plurality of circumferentially spaced outlet guide vanes 52. In addition, a downstream section 54 of nacelle 50 extends over an outer portion of core turbine engine 16 to define a bypass airflow passage 56 therebetween.

[0026] During operation of the turbine engine 10, a quantity of air 58 enters the turbine engine 10 through the nacelle 50 or the inlet 60 of the fan section 14. As the quantity of air 58 passes through the fan blades 40, a first portion of air 62 is directed or directed into the bypass airflow passage 56, and a second portion of air 64 is directed or directed into the upstream section of the core air flow path, or more specifically, into the annular inlet 20 of the LP compressor 22. The ratio between the first portion of air 62 and the second portion of air 64 is generally referred to as the bypass ratio. The pressure of the second portion of air 64 is then increased to form compressed air 65, and the compressed air 65 is directed through the HP compressor 24 and into the combustion section 26, where the compressed air 65 is mixed with fuel and combusted to generate combustion gases 66.

[0027] The combustion gases 66 are directed into and expanded through the HP turbine 28, where a portion of the heat and kinetic energy from the combustion gases 66 is extracted via successive stages of HP turbine stator blades 68 coupled to the casing 18 and HP turbine rotor blades 70 coupled to the HP shaft 34, thereby causing the HP shaft 34 to rotate, thereby supporting the operation of the HP compressor 24. The combustion gases 66 are then directed into and expanded through the LP turbine 30. Here, a second portion of the heat and kinetic energy is extracted from the combustion gases 66 via successive stages of LP turbine stator blades 72 coupled to the casing 18 and LP turbine rotor blades 74 coupled to the LP shaft 36, thereby causing the LP shaft 36 to rotate, thereby supporting the operation of the LP compressor 22 and the rotation of the fan 38 via the gearbox assembly 46.

[0028] The combustion gases 66 are then directed through the jet exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. At the same time, the pressure of the first portion of air 62 is significantly increased, also providing propulsive thrust, as it is directed through the bypass airflow passage 56 before being discharged from the fan nozzle exhaust section 76 of the turbine engine 10. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 for directing the combustion gases 66 through the core turbine engine 16.

[0029] The controller 90 communicates with the turbine engine 10 for controlling various aspects of the turbine engine 10. For example, the controller 90 is in bidirectional communication with the turbine engine 10 for receiving signals from various sensors and control systems of the turbine engine 10, and for controlling components of the turbine engine 10, as described in further detail below. The controller 90 or components thereof may be onboard the turbine engine 10, onboard the aircraft, or may be remote from each of the turbine engine 10 and the aircraft. The controller 90 may be a full authority digital engine control (FADEC) that controls various aspects of the turbine engine 10.

[0030] The controller 90 may be a stand-alone controller or may be part of an engine controller to operate various systems of the turbine engine 10. In this embodiment, the controller 90 is a computing device having one or more processors and a memory. The one or more processors may be any suitable processing device, including but not limited to a microprocessor, a microcontroller, an integrated circuit, a logic device, a programmable logic controller (PLC), an application specific integrated circuit (ASIC), or a field programmable gate array (FPGA). The memory may include one or more computer-readable media, including but not limited to non-temporary computer-readable media, computer-readable non-volatile media (e.g., flash memory), RAM, ROM, a hard drive, a flash drive, or other memory devices.

[0031] The memory can store information that can be accessed by one or more processors, including computer-readable instructions that can be executed by one or more processors. The instructions can be any set of instructions or sequence of instructions that cause one or more processors and the controller 90 to operate when executed by one or more processors. The controller 90, and more specifically, one or more processors are programmed or configured to perform these operations, such as the operations discussed further below. In some embodiments, the instructions can be executed by one or more processors to cause one or more processors to complete any operations and functions that the controller 90 is configured for, as will be further described below. The instructions can be software written in any suitable programming language or can be implemented in hardware. Additionally or alternatively, the instructions can be executed in logically or virtually independent threads on the processor. The memory can further store data that can be accessed by one or more processors.

[0032] The techniques discussed herein refer to computer-based systems and the actions taken by computer-based systems and information sent to and from computer-based systems. It will be appreciated by those of ordinary skill in the art that the inherent flexibility of computer-based systems allows for a variety of possible configurations, combinations, and divisions of tasks and functions between and within components. For example, the processes discussed herein can be implemented using a single computing device or multiple computing devices working in combination. Databases, memory, instructions, and applications can be implemented on a single system, or can be distributed on multiple systems. Distributed components can operate sequentially or in parallel.

[0033] Figure 1 The turbine engine 10 depicted in the drawings is for example only. In other exemplary embodiments, the turbine engine 10 may have any other suitable configuration. For example, in other exemplary embodiments, the fan 38 may be configured in any other suitable manner (e.g., as a fixed pitch fan) and may also be supported using any other suitable fan frame configuration. In addition, in other exemplary embodiments, any other suitable number or configuration of compressors, turbines, shafts, or combinations thereof may be provided. In still other exemplary embodiments, aspects of the present disclosure may be incorporated into any other suitable turbine engine (e.g., such as a turbofan engine, a propfan engine, a turbojet engine, a turboprop engine, or a turboshaft engine).

[0034] Figure 2 According to the present disclosure Figure 1 Detail 2-2 of FIG. 2 shows a schematic cross-sectional view of a portion of an HP compressor 24 having an exhaust system 100. The HP compressor 24 includes an HP compressor flow path 102 and a plurality of stages 104 of HP compressor rotor blades 106 and HP compressor stator vanes 108. The HP compressor 24 includes at least three stages 104. For example, the HP compressor 24 includes nine stages 104 ( Figure 2 Only seven stages 104 are shown, and the two upstream stages are Figure 2 Each stage 104 is defined by a respective HP compressor rotor blade 106 and a respective HP compressor stator vane 108. The HP compressor 24 includes an outer HP compressor casing 110 (e.g., forming Figure 1 ) and an inner HP compressor shell 112 radially spaced apart on the inner side of the outer HP compressor shell 110.

[0035] HP compressor stator vanes 108 are coupled to the inner HP compressor casing 112 and are static components. HP compressor rotor blades 106 are coupled to the HP shaft 34 so as to rotate as the HP shaft 34 rotates. In this manner, as the compressed air 65 flows from the upstream portion of the HP compressor 24 to the downstream portion of the HP compressor 24 through the HP compressor flow path 102, the HP compressor 24 further compresses the compressed air 65. The compressed air 65 then passes through the HP compressor diffuser 109, which directs the compressed air 65 to the combustion section 26 ( Figure 1 )middle.

[0036] The exhaust system 100 includes one or more exhaust plenums 114, a flexible seal 116, and a rigid support member 118. The one or more exhaust plenums 114 are defined between the outer HP compressor shell 110 and the inner HP compressor shell 112. The one or more exhaust plenums 114 are circumferentially disposed about the longitudinal centerline axis 12 ( Figure 1 ) is annular. The flexible seal 116 is disposed in one or more exhaust plenums 114 and extends from the inner HP compressor shell 112 to the outer HP compressor shell 110. The flexible seal 116 is circumferentially disposed about the longitudinal centerline axis 12 ( Figure 1 ) is annular and divides one or more exhaust air chambers 114 axially and fluidically into multiple exhaust air chambers 114.

[0037] The flexible seal 116 accommodates relative movement between the two exhaust plenums 114 while maintaining a substantially airtight seal between the two exhaust plenums 114 such that the exhaust plenums 114 are fluidly separated volumes. The flexible seal 116 may be made of any metallic flexible sealing material that is resistant to temperatures up to at least 1000° F. (such as, by way of non-limiting example, titanium alloys (e.g., titanium 6-2 or titanium 17), nickel-chromium alloys (e.g., 718), steel, etc. In some embodiments, the flexible seal 116 includes a bellows or other spring-like features so that the flexible seal 116 has mechanical compliance.

[0038] The rigid support member 118 provides the primary mechanical support for the HP compressor flow path 102. The rigid support member 118 is disposed in one or more of the discharge plenums 114 and extends from the inner HP compressor shell 112 to the outer HP compressor shell 110. In this manner, the rigid support member 118 supports the inner HP compressor shell 112. The rigid support member 118 is disposed about a longitudinal centerline axis ( Figure 1 ) is annular and divides one or more exhaust air chambers 114 axially and fluidically into multiple exhaust air chambers 114.

[0039] The flexible seal 116 and the rigid support member 118 divide the one or more exhaust chambers 114 into three exhaust chambers 114, including a first exhaust chamber 114a, a second exhaust chamber 114b, and a third exhaust chamber 114c. The first exhaust chamber 114a is defined by the inner HP compressor shell 112, the outer HP compressor shell 110, the flexible seal 116, and the front support member 117. The second exhaust chamber 114b is defined by the inner HP compressor shell 112, the outer HP compressor shell 110, the flexible seal 116, and the rigid support member 118. The third exhaust chamber 114c is defined by the inner HP compressor shell 112, the outer HP compressor shell 110, the rigid support member 118, and the rear support member 119. The second exhaust chamber 114b is behind the first exhaust chamber 114a, and the third exhaust chamber 114c is behind the second exhaust chamber 114b. The one or more discharge plenums 114 also include a fourth discharge plenum 114 d , also referred to as a post-diffuser flow path, defined between the inner HP compressor shell 112 and the outer HP compressor shell 110 axially rearward of the HP compressor diffuser 109 .

[0040] The exhaust system 100 also includes a plurality of exhaust ports 120 and a plurality of exhaust outlets 122 in fluid communication with the one or more exhaust plenums 114. The plurality of exhaust ports 120 are disposed through the inner HP compressor casing 112 and provide fluid communication from the HP compressor flow path 102 to the one or more exhaust plenums 114. The plurality of exhaust outlets 122 are disposed through the outer HP compressor casing 110 and provide fluid communication from the one or more exhaust plenums 114 to at least one of the one or more turbine engine systems (e.g., LP turbine 30, HP turbine 28, etc.) or one or more aircraft systems 130 (e.g., such as an environmental control system (ECS) or wing anti-icing (WAI) system). In this manner, a portion of the compressed air 65 is exhausted from the HP compressor flow path 102, through the plurality of exhaust ports 120, through the one or more exhaust plenums 114, through the plurality of exhaust outlets 122, and is exhausted to at least one of the one or more turbine engine systems or one or more aircraft systems, as further described below.

[0041] The plurality of exhaust ports 120 include a first exhaust port 120a, a second exhaust port 120b, and a third exhaust port 120c. The plurality of exhaust outlets 122 include a first exhaust outlet 122a, a second exhaust outlet 122b, and a third exhaust outlet 122c. The first exhaust port 120a and the first exhaust outlet 122a are in fluid communication with the first exhaust air chamber 114a and define a first exhaust flow path 140. The second exhaust port 120b and the second exhaust outlet 122b are in fluid communication with the second exhaust air chamber 114b and define a second exhaust flow path 142. The third exhaust port 120c and the third exhaust outlet 122c are in fluid communication with the third exhaust air chamber 114c and define a third exhaust flow path 144. Therefore, the exhaust system 100 includes a plurality of exhaust flow paths 140, 142, and 144, and the plurality of exhaust flow paths 140, 142, and 144 include at least three exhaust flow paths 140, 142, and 144. In some embodiments, the plurality of exhaust ports 120 includes a fourth exhaust port 120d and the plurality of exhaust outlets 122 includes a fourth exhaust outlet 122d. The fourth exhaust port 120d and the fourth exhaust outlet 122d are in fluid communication with the fourth exhaust plenum 114d (eg, a post-diffuser flow path) and define a fourth exhaust flow path 146.

[0042] The first exhaust port 120a is in fluid communication with the third stage 104a of the HP compressor 24 (the first and second stages of the HP compressor 24 are not shown). Figure 2 ). The second exhaust port 120b is in fluid communication with the fourth stage 104b of the HP compressor 24. The third exhaust port 120c is in fluid communication with the sixth stage 104c of the HP compressor 24. In this manner, the first exhaust port 120a (e.g., the first exhaust flow path 140) and the second exhaust port 120b (e.g., the second exhaust flow path 142) are located at consecutive stages 104 of the HP compressor 24. Therefore, at least two of the exhaust flow paths 140, 142, and 144 are located at consecutive stages 104 of the HP compressor 24.

[0043] The sizes and positions (e.g., at a particular stage 104) of the plurality of exhaust ports 120 and the plurality of exhaust outlets 122 are selected so that the first exhaust flow path 140 recovers more than 10% of the dynamic pressure of the first exhaust flow path 140 to a static pressure within the first exhaust flow path 140. For example, the first exhaust port 120a is a slot that converts more than 10% of the dynamic pressure of the first exhaust flow path 140 into a static pressure within the first exhaust plenum 114a. In this manner, the first exhaust flow path 140 is an efficient exhaust flow path such that the first exhaust flow path 140 provides the highest possible pressure without having to utilize more fuel combustion by moving to a more downstream stage 104. Further, the sizes and positions of the plurality of exhaust ports 120 and the plurality of exhaust outlets 122 are selected so that the total pressure (P second_bleed ) and the total pressure (P first_bleed ) is less than the pressure ratio of the stage 104 in which the second exhaust flow path 142 is located. In this way, the first exhaust flow path 140 is more efficient than the second exhaust flow path 142 and provides a shorter length of the HP compressor 24. In some embodiments, the first exhaust flow path 140 is as efficient as or less efficient than the second exhaust flow path 142.

[0044] The sizes and positions (e.g., at a particular stage 104) of the plurality of discharge ports 120 and the plurality of discharge outlets 122 are selected so that the second discharge flow path 142 recovers less than 10% of the dynamic pressure of the second discharge flow path 142 to a static pressure within the second discharge flow path 142. For example, the second discharge port 120b is a hole that is aligned perpendicular to the HP compressor flow path 102 and converts less than 10% of the dynamic pressure of the second discharge flow path 142 to a static pressure within the second discharge plenum 114b. This configuration provides a second discharge flow path recovery ratio (e.g., a ratio of the total pressure at the outlet of the second discharge port 120b to the total pressure in the HP compressor flow path 102 at the fourth stage 104b) at the second discharge flow path 142 that is lower than a first discharge flow path recovery ratio (e.g., a ratio of the total pressure at the outlet of the first discharge port 120a to the total pressure in the HP compressor flow path 102 at the third stage 104a) of the first discharge flow path 140.

[0045] The first exhaust flow path 140 includes a first exhaust flow path valve 141, and the fourth exhaust flow path 146 includes a fourth exhaust flow path valve 147. The first exhaust flow path valve 141 is a regulating valve that can be fully open, fully closed, or positioned between fully open and fully closed. The first exhaust flow path valve 141 can be any type of regulating valve, such as, for example, a stop valve, a disk valve, a butterfly valve, a ball valve, etc. In this manner, the first exhaust flow path valve 141 is controlled to regulate (e.g., change) the amount of exhaust gas passing through the first exhaust flow path 140 to one or more aircraft systems 130, as described below with respect to Figure 4 The first exhaust flow path valve 141 is positioned in the first exhaust flow path 140, radially outward of the outer HP compressor shell 110. The first exhaust flow path valve 141 may be positioned in the first exhaust flow path 140, anywhere along the first exhaust flow path 140. For example, the first exhaust flow path valve 141 may be positioned within the first exhaust port 120a, within the first exhaust plenum 114a, or anywhere along the first exhaust flow path 140.

[0046] The fourth exhaust flow path valve 147 includes any type of valve that can be controlled to open to allow exhaust gas to flow through the fourth exhaust flow path 146 and to close to prevent exhaust gas from flowing through the fourth exhaust flow path 146, as described in further detail below. The fourth exhaust flow path valve 147 is positioned in the fourth exhaust flow path 146, radially outside the outer HP compressor shell 110. The fourth exhaust flow path valve 147 can be positioned in the fourth exhaust flow path 146 at any position along the fourth exhaust flow path 146. For example, the fourth exhaust flow path valve 147 can be positioned within the fourth exhaust port 120d, within the fourth exhaust plenum 114d (e.g., the rear diffuser flow path), or at any position along the fourth exhaust flow path 146. In some embodiments, the second exhaust flow path 142 includes a second exhaust flow path valve, which is a regulating valve. In some embodiments, the third exhaust flow path 144 includes a third exhaust flow path valve (not shown), which includes a regulating valve. In this way, the amount of the second portion of the compressed air 65 or the amount of the third portion of the compressed air 65 can also be changed.

[0047] Figure 3 is a flow chart of a method 300 of operating a turbine engine 10 according to the present disclosure. The turbine engine 10 is as shown above. Figure 1 In this manner, compressed air 65 is directed from LP compressor 22 to HP compressor 24 .

[0048] refer to Figure 2In step 305, method 300 includes directing compressed air 65 through HP compressor flow path 102. For example, HP compressor 24 directs compressed air 65 through HP compressor flow path 102 through rotation of HP compressor rotor blades 106. As compressed air 65 travels downstream within HP compressor flow path 102, one or more stages 104 further progressively compress compressed air 65. In this manner, as compressed air 65 travels through HP compressor 24, the pressure of compressed air 65 increases.

[0049] In step 310, method 300 includes directing a first portion of compressed air 65 through first exhaust flow path 140. For example, the first portion of compressed air is directed through first exhaust port 120a, through first exhaust plenum 114a, and through first exhaust outlet 122a. The first portion of compressed air 65 is then directed to one or more aircraft systems 130. For example, the first portion of compressed air 65 is used in at least one of an environmental control system or a wing anti-icing system of the aircraft.

[0050] In step 315, the method 300 includes directing a second portion of the compressed air 65 through the second exhaust flow path 142. For example, the second portion of the compressed air 65 is directed through the second exhaust port 120b, through the second exhaust plenum 114b, and through the second exhaust outlet 122b. The flexible seal 116 prevents the second portion of the compressed air 65 from entering the first exhaust plenum 114a, and prevents the first portion of the compressed air 65 from entering the second exhaust plenum 114b. The second portion of the compressed air 65 is then directed to the LP turbine 30 for cooling a portion of the LP turbine 30.

[0051] In step 320, the method 300 includes directing a third portion of the compressed air 65 through the third exhaust flow path 144. For example, the third portion of the compressed air 65 is directed through the third exhaust port 120c, through the third exhaust plenum 114c, and through the third exhaust outlet 122c. The rigid support member 118 prevents the third portion of the compressed air 65 from entering the second exhaust plenum 114b, and prevents the second portion of the compressed air 65 from entering the third exhaust plenum 114c. In this manner, the first exhaust plenum 114a, the second exhaust plenum 114b, and the third exhaust plenum 114c are fluidly separated, and the first exhaust flow path 140, the second exhaust flow path 142, and the third exhaust flow path 144 are fluidly separated. The third portion of the compressed air 65 is then directed to the HP turbine 28 for cooling a portion of the HP turbine 28.

[0052] Thus, the method 300 includes directing the compressed air 65 through at least three exhaust flow paths 140, 142, and 144. At least two of the exhaust flow paths 140, 142, and 144 are at consecutive stages 104 of the HP compressor 24. For example, the first exhaust flow path 140 is at the third stage 104a, and the second exhaust flow path 142 is at the fourth stage 104b. The third exhaust flow path 144 is at the sixth stage 104c. However, the exhaust flow paths 140, 142, and 144 may be at any stage 104 of the HP compressor 24, such that at least two of the exhaust flow paths 140, 142, and 144 are at consecutive stages 104. Further, the at least three exhaust flow paths 140, 142, 144 may be arranged in any order, as long as at least two of the exhaust flow paths 140, 142, 144 are at consecutive stages 104 of the HP compressor 24. For example, the first exhaust flow path 140 may be at the fourth stage 104 b and the second exhaust flow path 142 may be at the third stage 104 a. In this manner, the first exhaust flow path 140 directs a first portion of the compressed air 65 from the fourth stage 104 b to one or more aircraft systems 130 and the second exhaust flow path 142 directs a second portion of the compressed air 65 from the third stage 104 a to the low pressure turbine 30.

[0053] Providing at least two of the exhaust flow paths 140, 142, and 144 at consecutive stages 104 allows for improved fuel combustion by supplying higher pressure air (e.g., from the fourth stage 104b) to the LP turbine 30 to achieve a positive backflow margin to prevent hot gas path ingestion in the LP turbine 30. Further, providing at least two of the exhaust flow paths 140, 142, and 144 at consecutive stages 104 also allows for lower pressure air to be provided to the aircraft systems 130, which has a lower cost to fuel combustion because less work is done on the flow at the third stage 104a compared to the exhaust ports at the more downstream stages 104.

[0054] Separate exhaust flow paths 140, 142, and 144 allow exhaust to be directed to various systems at various pressures and temperatures. In this manner, exhaust is directed to LP turbine 30 separately (e.g., at a different pressure and at a different temperature) from exhaust directed to one or more aircraft systems 130. Thus, exhaust system 100 provides exhaust to various systems of turbine engine 10 or an aircraft without unduly sacrificing the performance of HP compressor 24.

[0055] In some embodiments, the method 300 includes directing a fourth portion of the compressed air 65 through a fourth exhaust flow path 146. For example, the fourth portion of the compressed air 65 is directed through a fourth exhaust port 120d, through a fourth exhaust plenum 114d (e.g., aft diffuser flow path), and through a fourth exhaust outlet 122d. The fourth portion of the compressed air 65 is then directed to the HP turbine 28 or to at least one of the one or more aircraft systems 130 (e.g., to an environmental control system).

[0056] Figure 4 is a flow chart of a method 400 of operating a turbine engine 10 according to another embodiment. In step 405, the method 400 includes determining the altitude of the turbine engine 10. For example, the aircraft may include one or more altitude sensors, such as, for example, an altimeter, an atmospheric pressure sensor, a temperature sensor, etc., for providing an indication of the altitude of the turbine engine 10. The controller 90 receives the indication of the altitude from the one or more altitude sensors and determines the altitude based on the indication of the altitude from the one or more altitude sensors.

[0057] In step 410, method 400 includes determining whether the altitude of turbine engine 10 is increasing. In step 415, if the altitude of turbine engine 10 is increasing (step 410: yes), method 400 includes increasing the amount of the first portion of compressed air 65 passing through first exhaust flow path 140. For example, controller 90 controls first exhaust flow path valve 141 to adjust first exhaust flow path valve 141 to allow more compressed air 65 to pass through first exhaust flow path 140 and reach one or more aircraft systems 130. In this manner, method 400 includes increasing the amount of the first portion of compressed air 65 passing through first exhaust flow path 140 as the altitude of turbine engine 10 increases. The amount of the first portion of compressed air 65 is determined by a first mass flow rate W of the first portion of compressed air 65 passing through first exhaust flow path 140. first_bleed Thus, method 400 includes increasing the first mass flow rate W of the first portion of compressed air 65 through first exhaust flow path 140 as the altitude of turbine engine 10 increases. first bleed Thus, as the altitude increases, the amount of the first portion of the compressed air 65 (eg, the first mass flow rate W ) directed to the one or more aircraft systems 130 increases. first_bleed )Increase.

[0058] In step 420, if the altitude of the turbine engine 10 is decreasing (step 410: No), the method 400 includes determining whether the altitude of the turbine engine 10 is decreasing. If the altitude of the turbine engine 10 is not decreasing (step 420: No), the method 400 continues to step 405 to determine the altitude of the turbine engine 10. In step 425, if the altitude of the turbine engine 10 is decreasing (step 420: Yes), the method 400 includes reducing the amount of the first portion of the compressed air 65 passing through the first exhaust flow path 140. For example, the controller 90 controls the first exhaust flow path valve 141 to adjust the first exhaust flow path valve 141 to allow less compressed air 65 to pass through the first exhaust flow path 140 and reach one or more aircraft systems 130. In this way, the method 400 includes reducing the amount of the first portion of the compressed air 65 passing through the first exhaust flow path 140 as the altitude of the turbine engine 10 decreases. Thus, method 400 includes reducing the first mass flow rate W of the first portion of compressed air 65 through first exhaust flow path 140 as the altitude of turbine engine 10 decreases. first_bleed Thus, as the altitude decreases, the amount of the first portion of the compressed air 65 (eg, the first mass flow rate W ) directed to the one or more aircraft systems 130 increases. first_bleed )reduce.

[0059] Thus, method 400 includes varying an amount (eg, first mass flow rate W) of a first portion of compressed air 65 passing through first exhaust flow path 140 based on an altitude of turbine engine 10. first_bleed For example, method 400 includes adjusting first exhaust flow path valve 141 to change the amount (eg, first mass flow rate W) of a first portion of compressed air 65 passing through first exhaust flow path 140 based on an altitude of turbine engine 10. first_bleed ).

[0060] As the altitude changes (e.g., increases or decreases), the amount of the second portion of the compressed air 65 passing through the second exhaust flow path 142 remains approximately constant. The amount of the second portion of the compressed air 65 is determined by the second mass flow rate W of the second portion of the compressed air 65 passing through the second exhaust flow path 142. second_bleed Thus, as the altitude changes (eg, increases or decreases), the second mass flow rate W as a fraction of the total amount of compressed air 65 passing through the HP compressor flow path 102 is second_bleed Thus, as the altitude changes (e.g., increases or decreases), the amount of the second portion of compressed air 65 (e.g., the second mass flow rate W ) directed to the low pressure turbine 30 as a portion of the total amount of compressed air 65 passing through the HP compressor flow path 102 increases.second_bleed ) remains approximately constant. Further, the first mass flow rate W first_bleed and the second mass flow W second_bleed The mass flow ratio of increases with increasing altitude and decreases with decreasing altitude, as described in further detail below.

[0061] In some embodiments, after the altitude increases to be greater than the altitude threshold, method 400 continues. In one non-limiting example, the altitude threshold is six thousand feet (6,000 ft.). When the altitude is less than the altitude threshold, the first mass flow rate W first_bleed Thus, when the altitude is greater than the altitude threshold, method 400 continues such that the amount of the first portion of compressed air 65 varies with altitude, as detailed above.

[0062] In one non-limiting embodiment, the fourth exhaust flow path 146 provides a fourth portion of the compressed air 65 to one or more aircraft systems 130 until the pressure at the first exhaust flow path 140 is greater than the pressure threshold. For example, when the pressure of the first exhaust flow path 140 is less than or equal to the pressure threshold, the fourth exhaust flow path valve 147 is opened, allowing the fourth portion of the compressed air 65 to flow through the fourth exhaust flow path 146 and to the one or more aircraft systems 130. When the pressure of the first exhaust flow path 140 is less than or equal to the pressure threshold, the first exhaust flow path valve 141 is closed, preventing the compressed air 65 from flowing through the first exhaust flow path 140 when the pressure is less than or equal to the pressure threshold. When the pressure of the first exhaust flow path 140 is greater than the pressure threshold, the fourth exhaust flow path valve 147 is closed and the first exhaust flow path valve 141 is opened. For example, the controller 90 controls the fourth exhaust flow path valve 147 to close the fourth exhaust flow path valve 147 and controls the first exhaust flow path valve 141 to open the first exhaust flow path valve 141. In this manner, compressed air 65 is prevented from flowing through fourth exhaust flow path 146, and a first portion of compressed air 65 flows through first exhaust flow path 140 and toward one or more aircraft systems 130. When turbine engine 10 transitions from an idle condition to a takeoff condition, the pressure of first exhaust flow path 140 increases to be greater than the pressure threshold. Thus, when turbine engine 10 transitions from an idle condition to a takeoff condition, a first portion of compressed air 65 begins to flow through first exhaust flow path 140 toward one or more aircraft systems 130.

[0063] Figure 5 The first mass flow rate W according to the present disclosure as a function of the altitude of the turbine engine 10 is shown in a graphical form. first_bleed and the second mass flow W second_bleedRegion 500 represents exemplary boundaries of mass flow ratio and altitude. In one non-limiting example, the altitude is in the range of six thousand feet to fifty-one thousand feet (6,000 ft. to 51,000 ft.) and the mass flow ratio is in the range of zero point four to one point six (0.40 to 1.60). As the altitude increases from the takeoff altitude (e.g., 0 feet) to the cruising altitude (e.g., 38,000 feet to 51,000 feet), the mass flow ratio increases. In a non-limiting exemplary embodiment, the mass flow ratio is in the range of 0.40 to 0.60 at zero feet (0 ft.), in the range of 0.60 to 0.72 at twelve thousand feet (12,000 ft.), in the range of 0.75 to 0.91 at nineteen thousand eight hundred feet (19,800 ft.), in the range of 1.17 to 1.44 at thirty-five thousand feet (35,000 ft.), in the range of 1.20 to 1.47 at thirty-eight thousand feet (38,000 ft.), and in the range of 1.47 to 1.60 at fifty-one thousand feet (51,000 ft.).

[0064] The ranges of mass flow ratios for various altitudes are exemplary only, and the ranges of mass flow ratios and values ​​of altitude may vary depending on the particular size of the turbine engine 10 or the particular size of the aircraft in which the turbine engine 10 is installed. For example, a larger aircraft requires more flow through the first exhaust flow path 140 to reach one or more aircraft systems 130. The ranges and values ​​of mass flow ratios and altitudes will also vary based on the number of stages 104 of the HP compressor 24 and based on the particular stages 104 to which the first exhaust flow path 140 and the second exhaust flow path 142 are coupled. For example, the various stages 104 will provide various pressures through the first exhaust flow path 140 and the second exhaust flow path 142, and will require various mass flow rates through the first exhaust flow path 140 and the second exhaust flow path 142 to achieve the desired flow to the one or more aircraft systems 130 and to the LP turbine 30, respectively. The values ​​and ranges of mass flow ratios and altitudes will also vary based on the number of stages of the LP turbine 30. For example, the flow required from the second exhaust flow path 142 to the LP turbine 30 will vary based on the number of stages of the LP turbine 30. In some embodiments, the values ​​and ranges of mass flow ratio and altitude will vary based on the specific configuration of the turbine engine 10 or the aircraft. For example, if the aircraft includes a separate compressor for cabin pressure, the values ​​and ranges of mass flow ratio and altitude will vary, and therefore, the mass flow through the first exhaust flow path will be lower than that of an aircraft without a separate compressor for cabin pressure. If the exhaust system 100 provides a portion of the compressed air to the HP turbine 28 and the LP turbine 30 from the same stage 104 of the HP compressor 24, the values ​​and ranges of mass flow ratio and altitude will also vary, and therefore, the mass flow through the second exhaust flow path 142 will be higher than that of a turbine engine with separate exhaust flow paths to the HP turbine 28 and the LP turbine 30.

[0065] As described above, the first exhaust flow path 140 and the second exhaust flow path 142 can be reversed so that the first exhaust flow path 140 is at a downstream stage of the second exhaust flow path 142. For example, the first exhaust flow path 140 can be at the fourth stage 104b, and the second exhaust flow path 142 can be at the third stage 104a. In this configuration, the first exhaust flow path 140 directs a first portion of the compressed air 65 from the fourth stage 104b to one or more aircraft systems 130, and the second exhaust flow path 142 directs a second portion of the compressed air 65 from the third stage 104a to the low pressure turbine 30. In this manner, the mass flow ratio is the second mass flow rate W second_bleed With the first mass flow W first_bleed For example, when the amount of the first portion of the compressed air 65 (eg, the first mass flow rate Wfirst_bleed ) increases, the mass flow ratio decreases. In this configuration, the mass flow ratio decreases as the altitude increases from the takeoff altitude (e.g., 0 feet) to the cruising altitude (e.g., 38,000 feet to 51,000 feet). For example, the mass flow ratio is in the range of 1.47 to 1.60 at zero feet (0 ft.), in the range of 1.20 to 1.47 at twelve thousand feet (12,000 ft.), in the range of 1.17 to 1.44 at nineteen thousand eight hundred feet (19,800 ft.), in the range of 0.75 to 0.91 at thirty-five thousand feet (35,000 ft.), in the range of 0.60 to 0.72 at thirty-eight thousand feet (38,000 ft.), and in the range of 0.40 to 0.60 at fifty-one thousand feet (51,000 ft.). The ranges of mass flow ratios for various altitudes are exemplary only, and the ranges of mass flow ratios and values ​​of altitude may vary, as detailed above.

[0066] Thus, the exhaust system 100 of the present disclosure allows independent control of exhaust to one or more aircraft systems 130, the LP turbine 30, and the HP turbine 28. The exhaust system 100 provides at least two of the exhaust flow paths 140, 142, and 144 at consecutive stages 104 (e.g., the third and fourth stages) of the HP compressor 24 to provide efficient exhaust to the one or more aircraft systems 130. Providing at least two of the exhaust flow paths 140, 142, and 144 at consecutive stages 104 in this manner allows the exhaust system 100 of the HP compressor 24 to meet the required return margin and low pressure turbine inductor pressure ratio requirements of a particular turbine engine 10. Further, regulating the amount of the first portion of the compressed air 65 based on the altitude of the turbine engine 10 helps achieve the desired amount of compressed air 65 to one or more aircraft systems 130, while also reducing the amount of compressed air 65 that needs to be bled from the HP compressor 24 due to separate flow paths to the one or more aircraft systems 130 and to the low-pressure turbine 30, as compared to a turbine engine without the benefit of the present disclosure.

[0067] Further aspects of the disclosure are provided by the subject matter of the following clauses.

[0068] A turbine engine comprises: a high-pressure compressor, the high-pressure compressor including a high-pressure compressor flow path and a plurality of stages of high-pressure compressor rotor blades and high-pressure compressor stator vanes; and a discharge system, the discharge system comprising a plurality of discharge flow paths, the plurality of discharge flow paths including at least three discharge flow paths in fluid communication with the high-pressure compressor flow path, the plurality of discharge flow paths directing compressed air from the high-pressure compressor flow path, at least two of the discharge flow paths being at consecutive stages of the plurality of stages.

[0069] According to the turbine engine of the preceding clause, the plurality of exhaust flow paths includes a first exhaust flow path at a third stage of the plurality of stages, a second exhaust flow path at a fourth stage of the plurality of stages, and a third exhaust flow path at a sixth stage of the plurality of stages.

[0070] A turbine engine according to any preceding clause, wherein the plurality of exhaust flow paths comprises a first exhaust flow path leading to one or more aircraft systems, a second exhaust flow path leading to a low pressure turbine of the turbine engine, and a third exhaust flow path leading to a HP turbine of the turbine engine.

[0071] A turbine engine according to any preceding clause, said plurality of exhaust flow paths comprising a fourth exhaust flow path at a high pressure compressor diffuser of said high pressure compressor, said fourth exhaust flow path being directed to at least one of said one or more aircraft systems or said low pressure turbine.

[0072] According to the turbine engine described in any preceding clause, the high-pressure compressor further includes an outer high-pressure compressor shell and an inner high-pressure compressor shell, and the exhaust system further includes one or more exhaust air chambers defined between the outer high-pressure compressor shell and the inner high-pressure compressor shell, and the one or more exhaust air chambers define a portion of the multiple exhaust flow paths.

[0073] A turbine engine as described in any preceding clause, the one or more exhaust plenums comprising a first exhaust plenum defining a first exhaust flow path, a second exhaust plenum defining a second exhaust flow path, and a third exhaust plenum defining a third exhaust flow path.

[0074] A turbine engine according to any preceding clause, further comprising a flexible seal separating the first exhaust plenum and the second exhaust plenum such that the first exhaust flow path is fluidly separated from the second exhaust flow path.

[0075] A turbine engine according to any preceding clause, further comprising a rigid support member separating the second exhaust plenum and the third exhaust plenum such that the second exhaust flow path is fluidly separated from the third exhaust flow path.

[0076] A turbine engine according to any preceding clause, further comprising a plurality of exhaust ports arranged through the inner high-pressure compressor casing and a plurality of exhaust outlets arranged through the outer high-pressure compressor casing, the plurality of exhaust ports and the plurality of exhaust outlets being in fluid communication with the one or more exhaust plenums.

[0077] A turbine engine according to any preceding clause, wherein the plurality of exhaust ports and the plurality of exhaust outlets are sized and located in the high pressure compressor such that the first exhaust flow path restores more than 10% of the dynamic pressure of the first exhaust flow path to a static pressure within the first exhaust flow path.

[0078] A turbine engine according to any preceding clause, wherein the plurality of exhaust ports and the plurality of exhaust outlets are sized and located in the high pressure compressor such that a total pressure (P second_bleed ) and the total pressure of the first exhaust flow path (P first_bleed ) is less than the pressure ratio of the stage in which the second exhaust flow path is located.

[0079] A turbine engine according to any preceding clause, wherein the plurality of exhaust ports and the plurality of exhaust outlets are sized and located in the high pressure compressor such that the second exhaust flow path restores less than 10% of the dynamic pressure of the second exhaust flow path to the static pressure of the second exhaust flow path.

[0080] The turbine engine of any preceding clause, wherein the one or more aircraft systems include at least one of an environmental control system or a wing anti-icing system.

[0081] Turbine engine according to any preceding clause, wherein said high pressure compressor comprises nine stages.

[0082] A turbine engine according to any preceding clause, the high pressure compressor rotor blades being coupled to a high pressure shaft and rotating with rotation of the high pressure shaft.

[0083] A turbine engine according to any preceding clause, the high pressure compressor stator blades being coupled to the inner high pressure compressor casing.

[0084] A turbine engine as described in any preceding clause, wherein at least one of said plurality of exhaust flow paths extends axially between said high pressure compressor stator blades of one of said plurality of stages.

[0085] A turbine engine as described in any preceding clause, wherein the flexible seal is made of a metallic material resistant to temperatures up to at least 1000°F.

[0086] A turbine engine according to any preceding clause, wherein the flexible seal is made of at least one of a titanium alloy, a nickel-chromium alloy or steel.

[0087] A bleed system for a turbine engine, the turbine engine comprising a high-pressure compressor having a high-pressure compressor flow path and a plurality of stages, the bleed system comprising a plurality of bleed flow paths, the plurality of bleed flow paths comprising at least three bleed flow paths in fluid communication with the high-pressure compressor flow path of the high-pressure compressor, the plurality of bleed flow paths directing compressed air from the high-pressure compressor flow path, at least two of the bleed flow paths being at consecutive stages of the plurality of stages of the high-pressure compressor.

[0088] According to the exhaust system described in the preceding clause, the plurality of exhaust flow paths include a first exhaust flow path at a third stage among the plurality of stages, a second exhaust flow path at a fourth stage among the plurality of stages, and a third exhaust flow path at a sixth stage among the plurality of stages.

[0089] An exhaust system according to any preceding clause, the plurality of exhaust flow paths comprising a first exhaust flow path leading to one or more aircraft systems, a second exhaust flow path leading to a low pressure turbine of the turbine engine, and a third exhaust flow path leading to a high pressure turbine of the turbine engine.

[0090] The exhaust system of any preceding clause, the plurality of exhaust flow paths comprising a fourth exhaust flow path at a high pressure compressor diffuser of the high pressure compressor, the fourth exhaust flow path being directed to at least one of the one or more aircraft systems or the low pressure turbine.

[0091] According to the exhaust system described in any of the preceding clauses, the high-pressure compressor includes an outer high-pressure compressor shell and an inner high-pressure compressor shell, and the exhaust system further includes one or more exhaust air chambers defined between the outer high-pressure compressor shell and the inner high-pressure compressor shell of the high-pressure compressor, and the one or more exhaust air chambers define a portion of the multiple exhaust flow paths.

[0092] The exhaust system of any preceding clause, the one or more exhaust plenums comprising a first exhaust plenum defining a first exhaust flow path, a second exhaust plenum defining a second exhaust flow path, and a third exhaust plenum defining a third exhaust flow path.

[0093] The exhaust system of any preceding clause, further comprising a flexible seal separating the first exhaust plenum and the second exhaust plenum such that the first exhaust flow path is fluidly isolated from the second exhaust flow path.

[0094] The exhaust system of any preceding clause, further comprising a rigid support member separating the second exhaust plenum and the third exhaust plenum such that the second exhaust flow path is fluidly isolated from the third exhaust flow path.

[0095] The discharge system according to any of the preceding clauses further comprises a plurality of discharge ports arranged through the inner high-pressure compressor shell and a plurality of discharge outlets arranged through the outer high-pressure compressor shell, the plurality of discharge ports and the plurality of discharge outlets being in fluid communication with the one or more discharge plenums.

[0096] According to the exhaust system described in any preceding clause, the multiple exhaust ports and the multiple exhaust outlets are sized and located in the high-pressure compressor so that the first exhaust flow path restores more than 10% of the dynamic pressure of the first exhaust flow path to the static pressure within the first exhaust flow path.

[0097] A discharge system according to any preceding clause, wherein the plurality of discharge ports and the plurality of discharge outlets are sized and located in the high pressure compressor such that a total pressure (P second_bleed ) and the total pressure of the first exhaust flow path (P first_bleed ) is less than the pressure ratio of the stage in which the second exhaust flow path is located.

[0098] According to the exhaust system described in any preceding clause, the plurality of exhaust ports and the plurality of exhaust outlets are sized and located in the high-pressure compressor so that the second exhaust flow path restores less than 10% of the dynamic pressure of the second exhaust flow path to the static pressure within the second exhaust flow path.

[0099] A drain system according to any preceding clause, wherein a second drain flow path recovery rate of the second drain flow path is lower than a first drain flow path recovery rate of the first drain flow path.

[0100] The ejection system of any preceding clause, the one or more aircraft systems comprising at least one of an environmental control system or a wing anti-icing system.

[0101] A discharge system according to any preceding clause, wherein the high pressure compressor comprises nine stages.

[0102] A discharge system as described in any preceding clause, wherein the high pressure compressor rotor blades are coupled to a high pressure shaft and rotate with rotation of the high pressure shaft.

[0103] A discharge system as described in any preceding clause, the high pressure compressor stator vanes being coupled to the inner high pressure compressor casing.

[0104] The exhaust system of any preceding clause, the plurality of stages of the high pressure compressor each comprising high pressure compressor rotor blades and high pressure compressor stator vanes.

[0105] The exhaust system of any preceding clause, at least one of the plurality of exhaust flow paths extending axially between the high pressure compressor stator vanes of one of the plurality of stages.

[0106] The exhaust system of any preceding clause, wherein the flexible seal is made of a metallic material resistant to temperatures up to at least 1000°F.

[0107] The exhaust system according to any preceding clause, wherein the flexible seal is made of at least one of a titanium alloy, a nickel-chromium alloy or steel.

[0108] A method of operating a turbine engine according to any of the preceding clauses, the method comprising: directing compressed air through the high pressure compressor flow path; directing a first portion of the compressed air through a first exhaust flow path of the plurality of exhaust flow paths; directing a second portion of the compressed air through a second exhaust flow path of the plurality of exhaust flow paths; and directing a third portion of the compressed air through a third exhaust flow path of the plurality of exhaust flow paths.

[0109] According to the method described in the preceding clause, the turbine engine includes an inner high-pressure compressor shell and an outer high-pressure compressor shell, and the method further includes: directing the first portion of the compressed air through a first discharge port in the inner high-pressure compressor shell of the high-pressure compressor, through a first discharge air chamber defined between the inner high-pressure compressor shell and the outer high-pressure compressor shell, and through a first discharge outlet in the outer high-pressure compressor shell of the high-pressure compressor; directing the second portion of the compressed air through a second discharge port in the inner high-pressure compressor shell, through a second discharge air chamber defined between the inner high-pressure compressor shell and the outer high-pressure compressor shell, and through a second discharge outlet in the outer high-pressure compressor shell; and directing the third portion of the compressed air through a third discharge port in the inner high-pressure compressor shell, through a third discharge air chamber defined between the inner high-pressure compressor shell and the outer high-pressure compressor shell, and through a third discharge outlet in the outer high-pressure compressor shell.

[0110] A method according to any preceding clause, the turbine engine comprising a third stage of the plurality of stages, a fourth stage of the plurality of stages, and a sixth stage of the plurality of stages, and the method further comprising directing the first portion of the compressed air from the third stage of the plurality of stages through the first exhaust flow path, directing the second portion of the compressed air from the fourth stage of the plurality of stages through the second exhaust flow path, and directing the third portion of the compressed air from the sixth stage of the plurality of stages through the third exhaust flow path.

[0111] A method as in any preceding clause, the first exhaust flow path restoring more than 10% of the dynamic pressure of the first exhaust flow path to a static pressure within the first exhaust flow path.

[0112] A method according to any preceding clause, wherein the total pressure (P second_bleed ) and the total pressure of the first exhaust flow path (P first_bleed ) is less than the pressure ratio of the stage in which the second exhaust flow path is located.

[0113] A method according to any preceding clause, wherein the second exhaust flow path restores less than 10% of the dynamic pressure of the second exhaust flow path to the static pressure of the second exhaust flow path.

[0114] The method of any preceding clause, further comprising directing the first portion of the compressed air to one or more aircraft systems, directing the second portion of the compressed air to a low-pressure turbine of the turbine engine, and directing the third portion of the compressed air to a high-pressure turbine of the turbine engine.

[0115] A method according to any preceding clause, wherein the high pressure compressor includes a high pressure compressor diffuser, and the method further comprises directing a fourth portion of the compressed air from the high pressure compressor diffuser of the high pressure compressor and directing the fourth portion of the compressed air to at least one of the one or more aircraft systems or the low pressure turbine.

[0116] A method as in any preceding clause, the one or more aircraft systems comprising at least one of an environmental control system or a wing anti-icing system.

[0117] A method according to any preceding clause, wherein the high pressure compressor comprises nine stages.

[0118] A method according to any preceding clause, the high pressure compressor rotor blades being coupled to a high pressure shaft and rotating with rotation of the high pressure shaft.

[0119] A method according to any preceding clause, the high pressure compressor stator vanes being coupled to the inner high pressure compressor casing.

[0120] A method as in any preceding clause, the exhaust system further comprising a flexible seal separating the first exhaust plenum and the second exhaust plenum such that the first exhaust flow path is fluidly isolated from the second exhaust flow path.

[0121] A method according to any preceding clause, wherein at least one of the plurality of exhaust flow paths extends axially between the high pressure compressor stator vanes of one of the plurality of stages.

[0122] A method as in any preceding clause, wherein the flexible seal is made of a metallic material resistant to temperatures up to at least 1000°F.

[0123] A method as in any preceding clause, wherein the flexible seal is made of at least one of a titanium alloy, a nickel-chromium alloy or steel.

[0124] A method as in any preceding clause, the exhaust system further comprising a rigid support member separating the second exhaust plenum and the third exhaust plenum such that the second exhaust flow path is fluidly isolated from the third exhaust flow path.

[0125] A method of operating a turbine engine having a high pressure compressor including a high pressure compressor flow path and a plurality of stages of high pressure compressor rotor blades and high pressure compressor stator vanes, and a bleed system including a plurality of bleed flow paths, the plurality of bleed flow paths including a first bleed flow path from one of the plurality of stages and a second bleed flow path from another of the plurality of stages. The method includes: directing compressed air through the high pressure compressor flow path; directing a first portion of the compressed air through the first bleed flow path, the first portion of the compressed air having a first mass flow rate; directing a second portion of the compressed air through the second bleed flow path, the second portion of the compressed air having a second mass flow rate; determining an altitude of the turbine engine; and varying the first mass flow rate of the first portion of the compressed air through the first bleed flow path based on the altitude of the turbine engine.

[0126] The method according to the preceding clause, wherein the second mass flow rate of the second portion of the compressed air being part of the compressed air passing through the high pressure compressor flow path is approximately constant as the altitude varies.

[0127] A method as in any preceding clause, wherein the first exhaust flow path and the second exhaust flow path are at consecutive stages of the plurality of stages.

[0128] The method of any preceding clause, further comprising directing the first portion of the compressed air to one or more aircraft systems.

[0129] The method according to any preceding clause, further comprising directing the second portion of the compressed air to a low-pressure turbine of the turbine engine.

[0130] A method as described in any preceding clause, wherein a mass flow ratio of the first mass flow rate to the second mass flow rate increases with increasing altitude.

[0131] A method as in any preceding clause, wherein the mass flow ratio of the first mass flow rate to the second mass flow rate decreases as the altitude increases.

[0132] A method according to any preceding clause, the first exhaust flow path comprising a first exhaust flow path valve, and the method further comprising controlling the first exhaust flow path valve to vary the first mass flow rate of the first portion of the compressed air through the first exhaust flow path.

[0133] A method as in any preceding clause, wherein the first exhaust flow path valve is a regulating valve such that the first exhaust flow path valve is adjustable between a fully open position and a fully closed position to control an amount of the first portion of the compressed air passing through the first exhaust flow path valve.

[0134] A method as described in any preceding clause, further comprising determining whether the altitude is increasing or decreasing.

[0135] The method of any preceding clause, varying the first mass flow rate comprises increasing the first mass flow rate of the first portion of the compressed air through the first exhaust flow path as the altitude increases.

[0136] A method according to any preceding clause, wherein varying the first mass flow rate comprises decreasing the first mass flow rate of the first portion of the compressed air through the first exhaust flow path as the altitude decreases.

[0137] A method according to any preceding clause, the exhaust system further comprising a third exhaust flow path from another stage of the plurality of stages, and the method further comprising directing a third portion of the compressed air through the third exhaust flow path.

[0138] A method according to any preceding clause, the turbine engine further comprising a high pressure turbine, and the method further comprising directing the third portion of the compressed air to the high pressure turbine.

[0139] A method according to any preceding clause, the high pressure compressor further comprising a high pressure compressor diffuser, the exhaust system further comprising a fourth exhaust flow path from the high pressure compressor diffuser, and the method further comprising directing a fourth portion of the compressed air through the fourth exhaust flow path.

[0140] The method of any preceding clause, further comprising directing the fourth portion of the compressed air to one or more aircraft systems.

[0141] A method according to any preceding clause, the fourth exhaust flow path comprising a fourth exhaust flow path valve, and the method further comprising controlling the fourth exhaust flow path valve to open the fourth exhaust flow path valve to direct the fourth portion of the compressed air through the fourth exhaust flow path.

[0142] A method according to any preceding clause, further comprising controlling the fourth exhaust flow path valve to close to prevent the fourth portion of the compressed air from flowing through the fourth exhaust flow path.

[0143] The method of any preceding clause, further comprising directing the fourth portion of the compressed air through the fourth exhaust flow path and preventing the first portion of the compressed air from flowing through the first exhaust flow path when the altitude is less than or equal to an altitude threshold.

[0144] The method of any preceding clause, further comprising directing the first portion of the compressed air through the first exhaust flow path and preventing the fourth portion of the compressed air from flowing through the fourth exhaust flow path when the altitude is greater than the altitude threshold.

[0145] A method according to any preceding clause, wherein a mass flow ratio of the first mass flow rate to the second mass flow rate is in the range of 0.40 to 1.60.

[0146] A method as in any preceding clause, wherein the altitude is in the range of 6,000 feet to 51,000 feet.

[0147] A method as claimed in any preceding clause, the turbine engine being a turbine engine as claimed in any preceding clause.

[0148] A method according to any preceding clause, wherein the exhaust system is an exhaust system according to any preceding clause.

[0149] Although the foregoing description is directed to the preferred embodiments of the present disclosure, other changes and modifications will be apparent to those skilled in the art and may be made without departing from the spirit or scope of the present disclosure. In addition, the features described in conjunction with one embodiment of the present disclosure may be used in conjunction with other embodiments, even if not explicitly stated above.

Claims

1. A turbine engine, characterized in that: include: a high pressure (HP) compressor including an HP compressor flow path and a plurality of stages of HP compressor rotor blades and HP compressor stator vanes; and a discharge system comprising a plurality of discharge flow paths including at least three discharge flow paths in fluid communication with the HP compressor flow path, the plurality of discharge flow paths directing compressed air from the HP compressor flow path, wherein at least two of the discharge flow paths are at consecutive stages of the plurality of stages.

2. The turbine engine according to claim 1, characterized in that wherein the plurality of exhaust flow paths include a first exhaust flow path at a third stage of the plurality of stages, a second exhaust flow path at a fourth stage of the plurality of stages, and a third exhaust flow path at a sixth stage of the plurality of stages.

3. The turbine engine according to claim 1, characterized in that: Wherein the plurality of exhaust flow paths include a first exhaust flow path leading to one or more aircraft systems, a second exhaust flow path leading to a low pressure (LP) turbine of the turbine engine, and a third exhaust flow path leading to a HP turbine of the turbine engine.

4. The turbine engine according to claim 3, characterized in that Wherein the plurality of exhaust flow paths includes a fourth exhaust flow path at an HP compressor diffuser of the HP compressor, the fourth exhaust flow path being directed to at least one of the one or more aircraft systems or the LP turbine.

5. The turbine engine according to claim 1, characterized in that Wherein the HP compressor further comprises an outer HP compressor shell and an inner HP compressor shell, and the exhaust system further comprises one or more exhaust plenums defined between the outer HP compressor shell and the inner HP compressor shell, the one or more exhaust plenums defining a portion of the plurality of exhaust flow paths.

6. The turbine engine according to claim 5, characterized in that The one or more exhaust plenums include a first exhaust plenum defining a first exhaust flow path, a second exhaust plenum defining a second exhaust flow path, and a third exhaust plenum defining a third exhaust flow path.

7. The turbine engine according to claim 6, characterized in that Further included is a flexible seal separating the first exhaust plenum and the second exhaust plenum such that the first exhaust flow path is fluidly isolated from the second exhaust flow path.

8. The turbine engine according to claim 6, characterized in that Further included is a rigid support member separating the second exhaust plenum and the third exhaust plenum such that the second exhaust flow path is fluidly isolated from the third exhaust flow path.

9. The turbine engine according to claim 6, characterized in that Further included are a plurality of discharge ports disposed through the inner HP compressor shell and a plurality of discharge outlets disposed through the outer HP compressor shell, the plurality of discharge ports and the plurality of discharge outlets being in fluid communication with the one or more discharge plenums.

10. The turbine engine according to claim 9, characterized in that Wherein the plurality of discharge ports and the plurality of discharge outlets are sized and located in the HP compressor such that the first discharge flow path restores more than 10% of a dynamic pressure of the first discharge flow path to a static pressure within the first discharge flow path.

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