Mixed flow turbine core

Abstract

A gas turbine engine is provided having a compressor section, a combustion section downstream of the compressor section, and a turbine section downstream of the combustion section. A structural component extends from the compressor section to the turbine section to reinforce one or more components of the gas turbine engine. The structural component also defines a flow path extending between an inlet in air flow communication with the compressor section and an outlet in air flow communication with the turbine section. The flow path is configured to provide bleed air from the compressor section to the turbine section to capture at least a portion of the energy in the bleed air.

Description

Technical Field

[0001] The present subject matter generally relates to gas turbine engines having a mixed flow core. Background Art

[0002] The gas turbine engine generally comprises a fan and a core arranged in flow communication with each other. A first portion of the air over the fan may flow through the core via a bypass air flow (defined between the core and the outer nacelle), and a second portion of the air over the fan may be provided to the core.

[0003] The core of a gas turbine engine generally includes a compressor section, a combustion section, a turbine section, and an exhaust section, which are in a continuous flow sequence. In operation, air supplied to the core flows through the compressor section, where one or more axial compressors progressively compress the air until the air reaches the combustion section. Fuel is mixed with the compressed air and combusted in the combustion section to provide combustion gases. The combustion gases are sent from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then sent through the exhaust section, for example, to the atmosphere.

[0004] A typical gas turbine engine also includes a plurality of bleed air flow paths in air flow communication with the compressor section. For example, if the compressor section includes a low-pressure compressor and a high-pressure compressor, the plurality of bleed air flow paths may be in air flow communication with the low-pressure compressor. Depending on certain operating conditions of the gas turbine engine, at least a portion of the air flow through the low-pressure compressor may be diverted to, for example, a bypass air flow or the atmosphere through the bleed air flow path. Providing at least a portion of the air flow from the low-pressure compressor to, for example, a bypass air flow or the atmosphere through the bleed air flow path may assist in controlling certain parameters of the gas turbine engine (e.g., reducing the overall pressure ratio of the compressor section). Reducing the overall pressure ratio may increase the stall margin of the compressor.

[0005] However, providing a portion of the air flow from the low-pressure compressor through the bleed air flow path may result in reduced efficiency of the gas turbine engine. For example, any energy in the portion of the air flow from the low-pressure compressor provided to the atmosphere through the bleed air flow path may be lost due to this configuration. Therefore, a gas turbine engine that can reduce unnecessary energy losses while bleeding air from the compressor section would be beneficial. More specifically, a gas turbine engine that can capture energy within the bleed air flow would be particularly useful. Summary of the invention

[0006] Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.

[0007] In an exemplary embodiment of the present disclosure, a gas turbine is provided. The gas turbine engine includes a compressor section, a combustion section located downstream of the compressor section, and a turbine section located downstream of the combustion section. The gas turbine engine also includes a structural component extending from the compressor section to the turbine section to strengthen the gas turbine engine. The structural component defines a flow path extending between an inlet in air flow communication with the compressor section and an outlet in air flow communication with the turbine section. The flow path is configured to provide bleed air from the compressor section to the turbine section.

[0008] In another exemplary embodiment of the present disclosure, a gas turbine engine is provided. The gas turbine engine includes a compressor section, a combustion section located downstream of the compressor section, and a turbine section located downstream of the combustion section. The gas turbine engine also includes a casing that at least partially encloses the compressor section, the combustion section, and the turbine section. The gas turbine engine also includes a structural component that extends within the casing and adds structural support to the casing. The structural component further defines a flow path that extends between an inlet in air flow communication with the compressor section and an outlet in air flow communication with the turbine section.

[0009] Technical Solution 1. A gas turbine engine, comprising:

[0010] compressor section;

[0011] a combustion section downstream of the compressor section;

[0012] a turbine section downstream from the combustion section; and

[0013] A structural component extends from the compressor section to the turbine section, the structural component defining a flow path extending between an inlet in air flow communication with the compressor section and an outlet in air flow communication with the turbine section, the flow path being configured to provide bleed air from the compressor section to the turbine section.

[0014] Technical Solution 2. The gas turbine engine according to Technical Solution 1 is characterized in that the turbine section includes a high-pressure turbine, a low-pressure turbine and a staged air flow injection nozzle, and wherein the outlet of the flow path defined by the structural component is in air flow communication with the staged air flow injection nozzle.

[0015] Technical Solution 3. A gas turbine engine according to Technical Solution 2, characterized in that the structural component extends through the staged air flow injection nozzle, and wherein the flow path is configured to provide a portion of the bleed air from the compressor section to the turbine section at a position downstream of the air flow injection nozzle.

[0016] Technical Solution 4. The gas turbine engine according to Technical Solution 1 is characterized in that the compressor section includes a low-pressure compressor, a high-pressure compressor, and a frame component positioned between the low-pressure compressor and the high-pressure compressor, and wherein the inlet of the flow path defined by the structural component is in air flow communication with the frame component.

[0017] Technical Solution 5. The gas turbine engine according to Technical Solution 1 is characterized in that the gas turbine engine further comprises

[0018] A valve is positioned in a flow path defined by the structural component, wherein the valve is configured to vary an amount of air flow permitted through the flow path defined by the structural component.

[0019] Technical Solution 6. The gas turbine engine according to Technical Solution 1 is characterized in that the gas turbine engine further comprises

[0020] A heat exchanger is in air flow communication with the flow path defined by the structural member.

[0021] Technical Solution 7. The gas turbine engine according to Technical Solution 1 is characterized in that the gas turbine engine further comprises

[0022] An auxiliary compressor is in air flow communication with the flow path defined by the structural member.

[0023] Technical Solution 8. The gas turbine engine according to Technical Solution 7 is characterized in that the auxiliary compressor is a core-driven compressor that is mechanically connected to the high-pressure compressor of the compressor section.

[0024] Technical Solution 9. The gas turbine engine according to Technical Solution 7 is characterized in that the auxiliary compressor includes a plurality of variable inlet guide vanes.

[0025] Technical Solution 10. A gas turbine engine according to Technical Solution 1, characterized in that the compressor section limits a maximum mass flow rate therethrough, wherein a flow path defined by the structural component limits the maximum mass flow rate therethrough, and wherein the maximum flow rate through the flow path defined by the structural component is at least 10% of the maximum flow rate through the compressor section.

[0026] Technical Solution 11. A gas turbine engine according to Technical Solution 10, characterized in that the maximum flow rate through the flow path defined by the structural component is at least 5% of the maximum flow rate through the compressor section.

[0027] Technical Solution 12. The gas turbine engine according to Technical Solution 1 is characterized in that the compressor section, the combustion section and the turbine section are each at least partially enclosed in an outer casing, and wherein the structural component extends within the outer casing and increases the structural rigidity of the outer casing.

[0028] Technical Solution 13. A gas turbine engine, comprising:

[0029] compressor section;

[0030] a combustion section downstream of the compressor section;

[0031] a turbine section downstream of the combustion section;

[0032] a casing at least partially enclosing the compressor section, the combustion section, and the turbine section; and

[0033] A structural member extends within the outer casing and adds structural support to the outer casing, the structural member additionally defining a flow path extending between an inlet in air flow communication with the compressor section and an outlet in air flow communication with the turbine section.

[0034] Technical Solution 14. A gas turbine engine according to Technical Solution 13, characterized in that the turbine section includes a high-pressure turbine, a low-pressure turbine and a staged air flow injection nozzle, and wherein the outlet of the flow path defined by the structural component is in air flow communication with the staged air flow injection nozzle.

[0035] Technical Solution 15. A gas turbine engine according to Technical Solution 14, characterized in that the structural component extends through the staged air flow injection nozzle, and wherein the flow path is configured to provide a portion of the bleed air from the compressor section to the turbine section at a position downstream of the air flow injection nozzle.

[0036] Technical Solution 16. A gas turbine engine according to Technical Solution 13, characterized in that the compressor section includes a low-pressure compressor, a high-pressure compressor, and a frame component positioned between the low-pressure compressor and the high-pressure compressor, and wherein the inlet of the flow path defined by the structural component is in air flow communication with the frame component.

[0037] Technical Solution 17. The gas turbine engine according to Technical Solution 13 is characterized in that the gas turbine engine further comprises

[0038] A valve is positioned in a flow path defined by the structural component, wherein the valve is configured to vary an amount of air flow permitted through the flow path defined by the structural component.

[0039] Technical Solution 18. The gas turbine engine according to Technical Solution 13 is characterized in that the gas turbine engine further comprises

[0040] A heat exchanger is in air flow communication with the flow path defined by the structural member.

[0041] Technical Solution 19. The gas turbine engine according to Technical Solution 13 is characterized in that the gas turbine engine further comprises

[0042] An auxiliary compressor is in air flow communication with the flow path defined by the structural member.

[0043] Technical Solution 20. A gas turbine engine according to Technical Solution 13, characterized in that the compressor section limits a maximum mass flow rate therethrough, wherein a flow path defined by the structural component limits the maximum mass flow rate therethrough, and wherein the maximum flow rate through the flow path defined by the structural component is at least 5% of the maximum flow rate through the compressor section.

[0044] These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims.The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the present invention and, together with the description, serve to explain the principles of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0045] A complete and enabling disclosure of the present invention, including the best mode thereof, is set forth in the specification for those skilled in the art, and the specification refers to the accompanying drawings, in which:

[0046] Figure 1 is a schematic cross-sectional illustration of an exemplary gas turbine engine according to various embodiments of the present subject matter.

[0047] Figure 2 is a simplified schematic diagram of the core of a gas turbine engine according to an exemplary embodiment of the present disclosure.

[0048] Figure 3 is a simplified schematic diagram of the core of a gas turbine engine according to another exemplary embodiment of the present disclosure.

[0049] Figure 4 is a schematic diagram of an auxiliary compressor according to an exemplary embodiment of the present disclosure.

[0050] Figure 5 is a simplified schematic diagram of the core of a gas turbine engine according to another exemplary embodiment of the present disclosure.

[0051] Figure 6 is a simplified schematic diagram of the core of a gas turbine engine according to yet another exemplary embodiment of the present disclosure.

[0052] Figure 7 is a simplified schematic diagram of the core of a gas turbine engine according to yet another exemplary embodiment of the present disclosure.

[0053] Figure 8 is an enlarged schematic diagram of an air flow injection nozzle in a staged air flow injection nozzle according to an exemplary embodiment of the present disclosure.

[0054] Fig. 9 is a cross-sectional view of a pair of air flow injection nozzles according to an exemplary embodiment of the present disclosure.

[0055] Fig.10 is a cross-sectional view of a pair of air flow injection nozzles according to another exemplary embodiment of the present disclosure.

[0056] Fig.11 yes Fig.10 Another cross-sectional view of a pair of exemplary air flow injection nozzles.

[0057] Fig.12 is a cross-sectional view of a pair of air flow injection nozzles according to yet another exemplary embodiment of the present disclosure.

[0058] Fig.13 yes Fig.12 Another cross-sectional view of a pair of exemplary air jet nozzles

[0059] Fig.14 is an enlarged schematic diagram of a turbine rear frame component according to an exemplary embodiment of the present disclosure.

[0060] Fig.15 yes Fig.14 Another enlarged schematic diagram of an exemplary turbine rear frame component.

[0061] Parts List

[0062] 10 turbofan jet engines

[0063] 12 Longitudinal or axial centerline

[0064] 14 fan sections

[0065] 16-core turbo engine

[0066] 18 Shell

[0067] 20Entrance

[0068] 22 Low pressure compressor

[0069] 24 High pressure compressor

[0070] 26 Combustion section

[0071] 28 High-pressure turbine

[0072] 30 Low-pressure turbine

[0073] 32 Jet exhaust section

[0074] 34 High pressure shaft / shaft

[0075] 36 low pressure shaft / shaft

[0076] 37 core air flow path

[0077] 38 Fan

[0078] 40 blades

[0079] 42 plates

[0080] 44 Actuating components

[0081] 46 Power Gear Box

[0082] 48 Cabin

[0083] 50 Fan case or nacelle

[0084] 52 outlet guide vane

[0085] 54 downstream section

[0086] 56 Bypass air flow channel

[0087] 58 Air

[0088] 60 entrances

[0089] 62 First air part

[0090] 64 Second air part

[0091] 66 Combustion gas

[0092] 68 stator guide vane

[0093] 70 turbine rotor blades

[0094] 72 stator guide vanes

[0095] 74 Turbine rotor blades

[0096] 76 Fan nozzle exhaust section

[0097] 78 Hot Gas Path

[0098] 80 compressor frame parts

[0099] 82-stage bypass mixing nozzle

[0100] 84 Bypass mixing nozzle

[0101] 86LP Turbine Front Module

[0102] 88LP Turbo Rear Module

[0103] 90 Turbine Rear Frame

[0104] 92 bearings

[0105] 94Structural components

[0106] 96 Flow Path

[0107] 98 Entrance

[0108] 100 Exit

[0109] 102 valve

[0110] 104 Turbine frame parts

[0111] 106 Auxiliary compressor

[0112] 108 Auxiliary compressor blades

[0113] 109OG

[0114] 110 auxiliary compressor IGV

[0115] 111 rings

[0116] 112 Heat exchanger

[0117] 114 Flow Path

[0118] 116 Entrance

[0119] Exit 118

[0120] 120 auxiliary turbine

[0121] 121 First heat exchanger

[0122] 122Central axis of auxiliary turbine

[0123] 123 Second heat exchanger

[0124] 124 gear box

[0125] 126 Turbine section bushing

[0126] 128 Turbine frame parts

[0127] 130 Chamber

[0128] 132 cavity

[0129] 134 Opening

[0130] 136 Pressure side

[0131] 138 Suction side

[0132] 140 front end

[0133] 142 Backend

[0134] 144 Air flow direction

[0135] 146 Opening direction

[0136] 148 petals

[0137] 150 Pivot axis

[0138] 152 Maximum nozzle throat area

[0139] 154Minimum nozzle throat area

[0140] 156 cavity

[0141] 158 doors

[0142] 160 outer bushing

[0143] 162 flow paths. DETAILED DESCRIPTION

[0144] Reference will now be made in detail to the present embodiments of the invention, one or more examples of which are shown in the accompanying drawings. The detailed description uses numerical and letter references to the features in the drawings. The same or similar reference numerals in the drawings and description have been used to reference the same or similar parts of the present invention. As used herein, the terms "first", "second" and "third" are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of individual components. The terms "upstream" and "downstream" refer to the relative direction of a fluid flow in a fluid path. For example, "upstream" refers to the direction from which a fluid flows, while "downstream" refers to the direction to which a fluid flows.

[0145] Referring now to the drawings, wherein like reference numerals refer to like elements throughout the several views, Figure 1 is a schematic cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure. More particularly, Figure 1 In the embodiment of the present invention, the gas turbine engine is a high bypass turbofan jet engine 10, referred to herein as a "turbofan engine 10". Figure 1 , the turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference) and a radial direction R. The turbofan engine 10 may also define a circumferential direction C (not shown) extending circumferentially around the axial direction A. Generally, the turbofan 10 includes a fan section 14 and a core turbine engine 16 disposed downstream of the fan section 14 .

[0146] The depicted exemplary core turbine engine 16 is generally enclosed within a generally tubular casing 18 that defines an annular inlet 20. The casing 18 encloses, in serial flow relationship: a compressor section including a supercharger 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 air jet exhaust nozzle section 32. A high pressure (HP) shaft or shaft 34 drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft or shaft 36 drivingly connects the LP turbine 30 to the LP compressor 22. The compressor section, combustion section 26, turbine section, and nozzle section 32 collectively define a core air flow path 37 therethrough.

[0147] For the depicted embodiment, the fan section 14 includes a variable pitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced apart manner. As depicted, the fan blades 40 extend outwardly from the disk 42 generally along a radial direction R. Each fan blade 40 can be rotated relative to the disk 42 about a pitch axis P by operatively coupling the fan blade 40 to an appropriate actuation member 44, the actuation member 44 being configured to collectively and in unison change the pitch of the fan blade 40. The fan blades 40, disk 42, and actuation member 44 can be rotated together about the longitudinal axis 12 by passing through the LP shaft 36 of a power gearbox 46. The power gearbox 46 includes a plurality of gears to step down the rotational speed of the LP shaft 36 to a more efficient rotational fan speed.

[0148] Still refer to Figure 1 In an exemplary embodiment of the present invention, the disk 42 is covered by a rotatable front hub 48, the profile of the front hub 48 being aerodynamically configured to facilitate air flow through the plurality of fan blades 40. In addition, the exemplary fan section 14 includes an annular fan case or outer nacelle 50 that circumferentially surrounds the fan 38 and / or at least a portion of the core turbine engine 16. It should be understood that the nacelle 50 may be configured to be supported relative to the core turbine engine 16 by a plurality of circumferentially spaced outlet guide vanes 52. In addition, a downstream section 54 of the nacelle 50 may extend over an outer portion of the core turbine engine 16 so as to define a bypass air flow path 56 therebetween.

[0149] During operation of the turbofan engine 10, a certain amount of air 58 enters the turbofan 10 through the associated inlet 60 of the nacelle 50 and / or the fan section 14. As the certain amount of air 58 passes through the fan blades 40, a first portion of the air 58 is directed or sent to the bypass air flow passage 56 as indicated by arrow 62, and a second portion of the air 58 is directed or sent to the core air flow path 37 as indicated by arrow 64, or more particularly directed or sent to the LP compressor 22. The ratio between the first air portion 62 and the second air portion 64 is generally referred to as the bypass ratio. Then as the second air portion 64 is sent through the high pressure (HP) compressor 24, the pressure of the second air portion 64 is increased. The second air portion 64 then flows into the combustion section 26, where it is mixed with fuel and combusted to provide combustion gases 66.

[0150] The combustion gases 66 are routed through the HP turbine 28, wherein a portion of the thermal and / or kinetic energy from the combustion gases 66 is extracted by successive stages of HP turbine stator vanes 68 coupled to the casing 18 and HP turbine rotor blades 70 coupled to the HP shaft or shaft 34, thereby causing the HP shaft or shaft 34 to rotate, thereby supporting operation of the HP compressor 24. The combustion gases 66 are then routed through the LP turbine 30, wherein a second portion of the thermal and kinetic energy is extracted from the combustion gases 66 by successive stages of LP turbine stator vanes 72 coupled to the casing 18 and LP turbine rotor blades 74 coupled to the LP shaft or shaft 36, thereby causing the LP shaft or shaft 36 to rotate, thereby supporting operation of the LP compressor 22 and / or rotation of the fan 38.

[0151] The combustion gases 66 are then routed through the jet exhaust nozzle section 32 of the core turbine engine 16 to provide thrust. At the same time, the pressure of the first air portion 62 is significantly increased as the first air portion 62 is routed through the bypass air flow path 56 before it is discharged from the fan nozzle exhaust section 76 of the turbofan 10, thereby also providing propulsion. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 to route the combustion gases 66 through the core turbine engine 16.

[0152] But it should be understood that Figure 1The exemplary turbofan engine 10 depicted in the drawings is by way of example only, and in other exemplary embodiments, the turbofan engine 10 may have any other suitable configuration, including, for example, any other suitable number of shafts or axles. Additionally, or alternatively, aspects of the present disclosure may be incorporated into any other suitable gas turbine engine. For example, in other exemplary embodiments, aspects of the present disclosure may be incorporated into a turboshaft engine, a turbocore engine, a turboprop aircraft engine, a turbojet aircraft engine, etc.

[0153] Now refer to Figure 2 , provides a simplified schematic diagram of the core 16 of a turbofan engine 10 according to another exemplary embodiment. In certain embodiments, Figure 2 Various aspects of the core 16 of the exemplary turbofan engine 10 depicted in FIG. 1 may be described with reference to the above. Figure 1 The exemplary turbofan engine 10 depicted is constructed in a similar manner. Accordingly, like reference numerals are used to indicate like or similar components.

[0154] As shown, the core 16 of the turbofan engine 10 includes a compressor section, a combustion section 26 located downstream of the compressor section, and a turbine section located downstream of the combustion section 26. The compressor section generally includes the LP compressor 22, the HP compressor 24, and a compressor frame member 80 positioned between the LP compressor 22 and the HP compressor 24. Additionally, the turbine section generally includes the HP turbine 28, the LP turbine 30, and, as will be discussed in greater detail below, air flow injection nozzles 84 (see FIG. 1 ) in stages 82. Figure 8 ). For the depicted embodiment, the LP turbine 30 is divided into an LP turbine front module 86 and an LP turbine rear module 88, and the air flow injection nozzles 84 of the stage 82 are positioned between the LP turbine front module 86 and the LP turbine rear module 88. The turbine section additionally includes a turbine rear frame 90, which is located behind or downstream of the LP turbine 30.

[0155] As also discussed above, turbofan engine 10 includes one or more shafts that mechanically couple the compressor section and the turbine section. In particular, for the depicted embodiment, LP compressor 22 is mechanically coupled to LP turbine 30 via LP shaft 36, and HP compressor 24 is mechanically coupled to HP turbine 28 via HP shaft 34. A plurality of bearings 92 are provided to facilitate rotation of these various components.

[0156] Additionally, the compressor section, combustion section 26, and turbine section are at least partially enclosed in outer casing 18. Casing 18 may be a generally tubular casing that encloses core 16 of turbofan engine 10 and at least partially defines a bypass air flow path 56 with fan case / nacelle 50.

[0157] In addition, the core 16 of the exemplary turbofan engine 10 includes a structural member 94 that extends from the compressor section to the turbine section to strengthen the turbofan engine 10. In particular, for the depicted embodiment, the structural member 94 is positioned inboard of the outer casing 18 of the core 16 along the radial direction R and extends at least from the compressor frame member 80 located between the LP compressor 22 and the HP compressor 24 to the air flow injection nozzle 84 of the stage 82 located between the front module 86 of the LP turbine 30 and the rear module 88 of the LP turbine 30. The structural member 94 may add structural rigidity and support to the core 16 of the turbofan engine 10. Although not depicted, the structural member 94 may include one or more struts extending between the structural member 94 and, for example, the outer casing 18. The structural member 94 may be formed from a single piece of rigid material, such as a suitable metal material, and may define a substantially annular shape that encloses at least a portion of the compressor section, the combustion section 26, and the turbine section (see also FIG. 1 ). Figure 4 ). However, alternatively, the structural member 94 may be formed from a plurality of separate components attached in any suitable manner to extend continuously between the compressor section and the turbine section. It should be understood that, as used herein, the continuously extending structural member 94 may include one or more ports to allow at least a portion of the bleed air flow passing through the bleed air flow path to be diverted to one or more auxiliary systems of the engine or aircraft to which the engine is mounted.

[0158] In addition, the exemplary structural component 94 defines a flow path 96 that extends between an inlet 98 and an outlet 100. In particular, the depicted exemplary flow path 96 is defined by the radially outer portions of the compressor section, the combustion section 26, and the turbine section and the structural component 94. However, alternatively, in certain exemplary embodiments, the core 16 may additionally include an inner liner positioned between the stationary component 94 and the outer portions of one or more of the compressor section, the combustion section 26, and the turbine section. Due to this configuration, the flow path 96 may be defined by the structural component 94 and at least partially defined by the inner liner.

[0159] The flow path 96 defined by the structural member 94 serves as a bleed air flow path, and thus an inlet 98 of the flow path 96 is in air flow communication with the compressor section to receive the bleed air flow from the compressor section. In particular, for the depicted embodiment, the inlet 98 of the flow path 96 is in air flow communication with a portion of the core air flow path 37 extending through the compressor frame member 80 through an opening.

[0160] In addition, the flow path 96 defined by the structural member 94 is configured to provide the bleed air flow therethrough to the turbine section. Therefore, the outlet 100 of the flow path 96 is in air flow communication with the turbine section. In particular, the outlet 100 is in air flow communication with a portion of the core air flow path 37 extending through the LP turbine 30 through the air flow injection nozzle 84 of the stage 82. The flow path 96 defined by the structural member 94 can thus provide or reintroduce the bleed air flow pressurized by the LP compressor 22 to the portion of the core air flow path 37 extending through the LP turbine 30 through, for example, the air flow injection nozzle 84 of the stage 82. Due to this configuration, energy in the bleed air flow can be extracted therefrom. For example, the energy in the bleed air flow can be used to assist in driving the LP turbine 30 and to provide rotational power to the LP shaft 36.

[0161] Furthermore, for the exemplary embodiment depicted, a portion of the bleed air flow provided through the flow path 96 may additionally be directed downstream of the bypass injection nozzle 86 of the stage 82. For example, as shown, a portion of the bleed air flow through the flow path 96 is additionally provided to the aft module 88 of the LP turbine 30, such as through one or more openings 101. Furthermore, although not depicted, a portion of the bleed air flow provided through the flow path 96 may additionally be provided to, for example, a nozzle portion (not shown) of the turbine aft frame 90 or the core 16.

[0162] Still refer to Figure 2 In order to adjust the amount of air flow provided through the flow path 96 defined by the structural component 94, a valve 102 is provided. The depicted exemplary valve 102 is positioned in the flow path 96 defined by the structural component 94 and is configured to change the amount of air flow allowed through the flow path 96 defined by the structural component 94. For example, the valve 102 may be a variable flux valve that is movable between an open position that allows air flow to pass completely through the flow path 96 defined by the structural component 94 and a closed position that blocks all air flow through the flow path 96 defined by the structural component 94. In addition, the valve 102 may be movable to a plurality of positions between the open position and the closed position to adjust the amount of air flow allowed through the flow path 96 defined by the structural component 94. The valve 102 may be in operative communication with, for example, a controller of the turbofan engine 10 and / or an aircraft including the turbofan engine 10.

[0163] It should be understood that the flow path 96 defined by the structural component 94 can be configured to allow a relatively large amount of bleed air flow to pass therethrough. For example, the compressor section can define a maximum flow rate therethrough, and the flow path 96 can also define a maximum flow rate therethrough. As used herein, the term "maximum flow rate" refers to the maximum air flow rate that the corresponding component (one or more) can accommodate during normal operation of the turbofan engine 10. In at least some exemplary embodiments, the maximum flow rate through the flow path 96 defined by the structural component 94 can be at least about 5% of the maximum flow rate through the compressor section. However, alternatively, in other exemplary embodiments, the maximum flow rate through the flow path 96 defined by the structural component 94 can be at least about 10% of the maximum flow rate through the compressor section, at least about 25% of the maximum flow rate through the compressor section, at least about 40% of the maximum flow rate through the compressor section, at least about 50% of the maximum flow rate through the compressor section, at least about 75% of the maximum flow rate to the compressor section, or at least about 100% of the maximum flow rate to the compressor section. Thus, in certain exemplary embodiments, the core 16 of the turbofan engine 10 may define a bleed air ratio (i.e., a ratio of air flow through flow path 96 to air flow through the compressor section) of up to about 1:1. It should be understood that as used herein, approximate terms such as "approximately" and "substantially" mean within a 10% margin of error.

[0164] Thus, in one or more of the above embodiments, the inlet 98 of the flow path 96 may be configured to receive at least about 5% of the air flow through the compressor section, or more specifically, the core air flow path 37 of the LP compressor 22 extending through the compressor section. Alternatively, but in other embodiments, the inlet 98 of the flow path 96 may additionally be configured to receive up to at least about 10%, up to at least about 20%, up to at least about 30%, up to at least about 40%, or up to at least about 50% of the air flow through the core air flow path 37 of the LP compressor 22 extending through the compressor section.

[0165] Notably, providing bleed air flow from a compressor section (e.g., a portion of core air flow path 37 located just downstream of LP compressor 22) may assist in controlling certain parameters of turbofan engine 10. For example, bleeding at least a portion of the air flow from LP compressor 22 through flow path 96 defined by structural member 94 may allow for a reduction in the overall pressure ratio of the compressor section (depending on the operating conditions of turbofan engine 10), which may improve the stall margin of turbofan engine 10.

[0166] However, it should be understood that the Figure 2, and described above. For example, the inlet 98 of the flow path 96 defined by the structural member 94 may alternatively be directly in air flow communication with the LP compressor 22 (e.g., at the aft end LP of the compressor 22). Similarly, in other exemplary embodiments, the outlet 100 of the flow path 96 defined by the structural member 94 may instead be in air flow communication with a turbine frame member 104 located upstream of the LP turbine 30. Thus, in certain exemplary embodiments, the core 16 of the turbofan engine 10 may not include the compressor frame member 80, and / or may not include the air flow injection nozzles 84 of the stage 82 located between the forward module 86 of the LP turbine 30 and the aft module 88 of the LP turbine 30. For example, in certain exemplary embodiments, the LP turbine 30 may be a single unit, not divided into the forward module 86 and the aft module 88, and the air flow injection nozzles 84 of the stage 82 may be located upstream of the LP turbine 30 (e.g., Figure 2 ). Additionally, in still other embodiments, the core 16 may not include the turbine aft frame 90 located aft of the LP turbine 30 .

[0167] Now refer to Figure 3 , a simplified schematic diagram of the core 16 of a turbofan engine 10 is provided in accordance with another exemplary embodiment of the present disclosure. Figure 3 The core 16 of the exemplary embodiment may be as described above and Figure 2 The exemplary core 16 depicted in FIG. 1 is constructed in a substantially identical manner. Therefore, the same or similar reference numerals may indicate the same or similar components.

[0168] For example, Figure 3 The exemplary core 16 depicted in FIG. 1 includes a compressor section, a combustion section 26 located downstream of the compressor section, and a turbine section located downstream of the combustion section 26. The compressor section includes one or more compressors (i.e., LP compressor 22 and HP compressor 24), and the turbine section includes one or more turbines (i.e., HP turbine 28 and LP turbine 30). The LP compressor 22 and HP compressor 24 of the compressor section, and the HP turbine 28 and LP turbine 30 of the turbine section each rotate about the longitudinal centerline 12 of the turbofan engine 10.

[0169] in addition, Figure 3 The exemplary core 16 depicted in FIG. 1 includes a structural member 94 extending from the compressor section to the turbine section to strengthen the core 16 of the turbofan engine 10. The exemplary structural member 94 also at least partially defines a flow path 96 extending between an inlet 98 in air flow communication with the compressor section and an outlet 100 in air flow communication with the turbine section. However, for Figure 3In the embodiment depicted in FIG. 8 , core 16 of turbofan engine 10 includes additional components to utilize and / or enhance bleed air flow from the compressor section to the turbine section through flow path 96 defined by structural member 94 .

[0170] For example, now also refer to Figure 4 , Figure 3 The exemplary core 16 depicted in FIG. 1 further includes an auxiliary compressor 106 in air flow communication with the flow path 96 defined by the structural member 94. For the depicted embodiment, the auxiliary compressor 106 is a fan mechanically coupled to the HP shaft 34 by the HP compressor 24 (i.e., a core driven fan). Figure 3 As schematically depicted in FIG. 1 , the auxiliary compressor 106 includes a stepped auxiliary compressor blade 108, a stepped inlet guide vane 110 located upstream of the blade 108, and a stepped outlet guide vane 109 located downstream of the blade 108. In addition, a plurality of auxiliary compressor blades 108 are attached to a solid ring 111, which is in turn attached to the radially outer ends of the stepped HP compressor rotor blades 112 and can rotate with the stepped HP compressor rotor blades 112 ( Figure 4 The auxiliary compressor 106 may be configured to increase the pressure of the bleed air flow through the flow path 96 during operation.

[0171] In addition, the exemplary auxiliary compressor 106 includes the ability to change the air flow rate allowed therethrough. In particular, as stated, the exemplary auxiliary compressor 106 includes a plurality of inlet guide vanes 110 above the inlet of the auxiliary compressor 106. The inlet guide vanes 110 can be moved between a fully open position that allows the flow to pass through the auxiliary compressor 106 in its entirety and a fully closed position that limits at least a portion of the flow to pass through the auxiliary compressor 106. Thus, the inlet guide vanes 110 can be configured as variable inlet guide vanes and can cooperate with the valve 102 to adjust the bleed air flow through the flow path 96 defined by the structural member 94. Alternatively, in certain exemplary embodiments, including the variable inlet guide vanes 110 in the auxiliary compressor 106 can completely eliminate the need for the valve 102 (and thus such exemplary embodiments may not include the valve 102). Alternatively, still in other exemplary embodiments, when the inlet guide vanes 110 are in the fully closed position, the flow path 96 can bypass the auxiliary compressor 106. For example, in certain exemplary embodiments, the core 16 may define a bypass line (not shown) capable of bypassing operation of the auxiliary compressor 106 .

[0172] Including the auxiliary compressor 106 may allow additional energy to be extracted from the bleed air flow through the flow path 96 defined by the structural member 94. Additionally, where, for example, the auxiliary compressor 106 is driven by the HP shaft 34, the auxiliary compressor 106 may be a means for certain high pressure components of the core 16 (e.g., the HP compressor 24) to drive certain low pressure components of the core 16 (e.g., the LP turbine 30). Notably, the above benefits may occur while still allowing the core 16 to bleed air from a low pressure operating line (e.g., from the LP compressor 22).

[0173] Still refer to Figure 3 , the turbofan engine 10, or more suitably the core 16 of the turbofan engine 10, further includes a heat exchanger 113 in air flow communication with the flow path 96 defined by the structural member 94. For the depicted embodiment, the heat exchanger 113 is located downstream of the auxiliary compressor 106, at least partially within the flow path 96 defined by the structural member 94. However, in other exemplary embodiments, the heat exchanger 113 may alternatively be positioned upstream of the auxiliary compressor 106, at least partially within the flow path 96 defined by the structural member 94.

[0174] The heat exchanger 113 may be configured to transfer heat from one or more systems of the turbofan engine 10 to the bleed air flow in the flow path 96 defined by the structural member 94. For example, the heat exchanger 113 may be configured to transfer heat from a primary lubrication system of the core 16 of the turbofan engine 10 to the bleed air flow in the flow path 96 defined by the structural member 94. Additionally, although not depicted, in other embodiments, the core 16 of the turbofan engine 10 may additionally include a plurality of heat exchangers 113 in air flow communication with the flow path 96 defined by the structural member 94, spaced apart along the flow path 96 defined by the structural member 94 in any suitable manner.

[0175] Including one or more heat exchangers 113 in the flow path 96 defined by the structural member 94 may allow for better thermal control of the thermal management system of the turbofan engine 10 while also adding energy to the bleed air flow.

[0176] Now refer to Figure 5 , a simplified schematic diagram of the core 16 of a turbofan engine 10 is provided in accordance with yet another exemplary embodiment of the present disclosure. Figure 5 The core 16 of the exemplary embodiment may be as described above with reference to Figure 2 The exemplary core 16 described is constructed in substantially the same manner. Therefore, the same or similar reference numerals may indicate the same or similar components.

[0177] For example, Figure 5The exemplary core 16 depicted in FIG. 1 includes a compressor section, a combustion section 26 located downstream of the compressor section, and a turbine section located downstream of the combustion section 26. The compressor section includes one or more compressors (i.e., LP compressor 22 and HP compressor 24), and the turbine section includes one or more turbines (i.e., HP turbine 28 and LP turbine 30). The LP compressor 22 and HP compressor 24 of the compressor section, and the HP turbine 28 and LP turbine 30 of the turbine section each rotate about the longitudinal centerline 12 of the turbofan engine 10.

[0178] in addition, Figure 5 The exemplary core 16 depicted in FIG. 1 includes a flow path 114 extending between an inlet 116 and an outlet 118. Similar to the exemplary flow path 96 defined by the structural member 94 (see FIG. Figure 2 ), the exemplary flow path 114 is configured as a bleed air flow path 114 such that an inlet 116 is in air flow communication with the compressor section. Specifically, the inlet 116 of the exemplary flow path 114 is in air flow communication with the rear end LP of the compressor 22. Additionally, for the exemplary embodiment depicted, an outlet 118 of the flow path 114 is in air flow communication with the turbine section, or more specifically, with the turbine aft frame 90 of the turbine section.

[0179] It is worth noting that Figure 5 The exemplary embodiment depicted in FIG. 1 does not show the structural components defining the flow path 114. Although in certain exemplary embodiments, Figure 5 The exemplary core 16 depicted in FIG. 1 may include structural components defining a flow path 114 (and in a manner similar to that described above with reference to FIG. Figure 2 and / or Figure 3 The exemplary structural member 94 is constructed in substantially the same manner as described above, but in other exemplary embodiments, Figure 5 The exemplary core 16 depicted in FIG. 1 may not include structural components that define the flow path 114 . In such an exemplary embodiment, the flow path 114 may alternatively be defined by one or more suitable air flow conduits, which may or may not add structural rigidity to the core 16 .

[0180] The auxiliary turbine 120 is positioned in air flow communication with the flow path 114 to extract energy from the bleed air flow passing therethrough. As depicted, the auxiliary turbine 120 is offset relative to the longitudinal centerline 12 of the turbofan engine 10. In particular, the auxiliary turbine 120 defines a central axis 122 that is offset relative to the longitudinal centerline 12. However, it should be understood that in other exemplary embodiments, the auxiliary turbine 120 may instead be configured as a "tip turbine" having one or more of the HP compressor 24 or the HP turbine 28. As used herein, the term tip turbine refers to a turbine that is configured in an annular configuration around the HP compressor 24 or HP turbine 28 so as to rotate coaxially with the HP compressor 24 or HP turbine 28. For example, the tip turbine may be configured as described above with reference to Figure 3 and 4 The auxiliary compressor 106 described above is constructed in substantially the same manner. In such an embodiment, the tip turbine may include a solid ring fixed to the radially outer ends of the rotating rotor blades of the stages in the HP compressor 24 or HP turbine 28. In addition, the tip turbine may include a plurality of circumferentially spaced blades attached to the ring such that air flow over the blades of the tip turbine rotates the ring and the HP compressor blades or HP turbine blades of the corresponding stage.

[0181] Still refer to Figure 5 In an embodiment of the invention, the exemplary flow path 114 extends through the auxiliary turbine 120, such that the auxiliary turbine 120 is positioned in direct air flow communication with the flow path 114. The auxiliary turbine 120 may allow energy to be extracted from the bleed air flow passing through the flow path 114, and such extracted energy to be provided to one or more shafts of the core 16 of the turbofan engine 10. In particular, the auxiliary turbine 120 may be constructed in a manner similar to one or both of the HP turbine 28 or the LP turbine 30. For example, the auxiliary turbine 120 may include a series of rotor blades coupled to an auxiliary turbine shaft that may rotate about a central axis 122 of the auxiliary turbine 120. As schematically depicted, the turbine shaft may be mechanically coupled to the HP shaft 34 via a gearbox 124 to assist in rotating the HP shaft 34. The gearbox 124 may be a fixed ratio gearbox, or alternatively may be a variable ratio gearbox.

[0182] Including the auxiliary turbine 120 may allow energy to be extracted from the bleed air flow through the flow path 114. For example, the auxiliary turbine 120 may allow energy to be extracted from such bleed air and provided directly to the HP spool 34. Thus, such a configuration may allow the core 16 of the turbofan engine 10 to operate more efficiently. For example, such a configuration may allow the bleed air flow through the flow path 114 to provide power to the HP spool 34 during, for example, idle conditions, such that less fuel is required to operate the core 16 during idle conditions.

[0183] However, it should be understood that this is provided only by way of example. Figure 5 1 and 2. The exemplary core 16 and auxiliary turbine 120 depicted in FIG. 1. In other exemplary embodiments, for example, the auxiliary turbine 120 may instead be configured to provide energy extracted from the bleed air flow to the HP shaft 34 through an auxiliary gearbox. For example, the auxiliary turbine may be mounted to the auxiliary gearbox. With this configuration, the auxiliary turbine 120 may provide dual functionality: also configured as an air-driven starter / generator for the auxiliary gearbox, which may be configured to start the turbofan engine 10. The rotation of the auxiliary turbine may transfer the extracted energy to the HP shaft 34.

[0184] Furthermore, in still other exemplary embodiments, Figure 5 The core 16 depicted in the figure may additionally include the above referenced Figures 2 to 4 One or more aspects of the exemplary core 16 described. For example, now referring to Figure 6 , provides a simplified schematic diagram of the core 16 of a turbofan engine 10 according to another exemplary embodiment of the present disclosure, the exemplary core 16 including an auxiliary turbine 120 and an auxiliary compressor 106, which is also in air flow communication with the flow path 114. Figure 6 In the embodiment of the invention, the auxiliary compressor 106 defines a central axis 108 offset relative to the longitudinal centerline 12 of the turbofan engine 10 and is driven by one or more shafts of the core 16 of the turbofan engine 10. Alternatively, however, as in the embodiment described above, the auxiliary compressor 106 may be configured as a core-driven auxiliary compressor with the HP compressor 24. In addition, for Figure 6 In the exemplary embodiment, the inlet 116 of the flow path 114 is in air flow communication with the compressor frame member 80 downstream of the LP compressor 22 and upstream of the HP compressor 24, and the outlet 118 is in air flow communication with the turbine rear frame 90. However, in other exemplary embodiments, the outlet 118 of the flow path 114 may instead be in air flow communication with the atmosphere or the bypass air flow passage 56, for example (see Figure 1 For example, the depicted auxiliary turbine 120 and flow path 114 may be enclosed by the outer shell 18 of the core 16 (as in Figure 2 and 3 ), and an outlet 118 of the flow path 114 may be defined in the housing 18.

[0185] In addition, now refer to Figure 7 , another exemplary embodiment of the present disclosure is provided. In particular, Figure 7 A simplified schematic diagram of a core 16 of a turbofan engine 10 is provided in accordance with yet another exemplary embodiment of the present disclosure. Figure 7 The core 16 of the exemplary embodiment may be as described above with reference to Figure 5 The exemplary core 16 described is constructed in substantially the same manner. Therefore, the same or similar reference numerals may indicate the same or similar components.

[0186] For example, Figure 7 The exemplary core 16 of the turbofan engine 10 includes a compressor section, a combustion section 26 located downstream of the compressor section, and a turbine section located downstream of the combustion section 26. The compressor section includes one or more compressors (i.e., LP compressor 22 and HP compressor 24) and the turbine section includes one or more turbines (i.e., HP turbine 28 and LP turbine 30). The LP compressor 22 and HP compressor 24 of the compressor section, and the HP turbine 28 and LP turbine 30 of the turbine section each rotate about the longitudinal centerline 12 of the turbofan engine 10.

[0187] in addition, Figure 5 The exemplary core 16 depicted in FIG. 1 includes a flow path 114 extending between an inlet 116 and an outlet 118, with an auxiliary turbine 120 positioned in the flow path 114 between the inlet 116 and the outlet 118. The flow path 114 and the auxiliary turbine 120 may be arranged as described above with reference to FIG. Figure 5 The exemplary flow path 114 and auxiliary turbine 120 depicted are configured in the same manner.

[0188] in addition, Figure 7 The exemplary core 16 of the embodiment includes: a first heat exchanger 121, which is positioned to be in thermal communication with the flow path 114 at a position upstream of the auxiliary turbine 120; and a second heat exchanger 123, which is in thermal communication with the flow path at a position downstream of the auxiliary turbine 120. For the depicted embodiment, the first heat exchanger 121 is configured to remove heat from the bleed air flow through the flow path 114. The first heat exchanger 121 can be configured as an "air-to-air" heat exchanger that also heats the bypass air flow of the turbofan engine 10 (similar to the above reference Figure 1 For example, the first heat exchanger 121 may be positioned within the bypass passage (e.g., Figure 7 ), or alternatively may be located within the housing 18 of the core 16. With respect to such an embodiment, a portion of the bypass air flow may instead be directed through the first heat exchanger 121 again from the bypass passage.

[0189] The cooled and compressed bleed air may then flow through flow path 114 to auxiliary turbine 120, where the bleed air may expand as energy is extracted from such bleed air flow. Expansion of the bleed air through auxiliary turbine 120 may further reduce the temperature of the bleed air. The bleed air may then flow through flow path 114 to second heat exchanger 123 (at a location downstream of auxiliary turbine 120), where the temperature of the bleed air in flow path 114 is increased as second heat exchanger 123 removes heat from one or more systems of turbofan engine 10 and transfers such heat to the bleed air in flow path 114. For example, in certain exemplary embodiments, second heat exchanger 123 may be configured as a "liquid-to-air" heat exchanger configured to transfer heat from, for example, a main lubrication system or a fuel system of turbofan engine 10 to the bleed air in flow path 114.

[0190] The bleed air in flow path 114 is then provided to the rear turbine module 90. However, in other embodiments, the bleed air may instead be provided at any other suitable location in the turbine section. For example, in certain exemplary embodiments, the core 16 may further include an auxiliary compressor in air flow communication with the bleed air flow in flow path 114 to increase the pressure of such flow. With respect to configuration, the air flow may instead be provided as described above with reference to Figures 2 to 4 The same approach as described for one or more embodiments of the present invention is provided to, for example, the LP turbine 30 .

[0191] Now refer to Figure 8 , provides an enlarged schematic diagram of an air flow injection nozzle of stage 82 according to an exemplary embodiment of the present disclosure. Figure 8 The air flow injection nozzle 84 of the exemplary stage 82 depicted in FIG is located within the turbine section, downstream of the first turbine, ie, the HP turbine 28. In particular, for the depicted embodiment, the air flow injection nozzle 84 of the exemplary stage 82 is located between a front module 86 of the LP turbine 30 and an aft module 88 of the LP turbine 30.

[0192] In addition, as shown, the air flow injection nozzle 84 of the stage 82 is in air flow communication with the outlet of the flow path. For the depicted embodiment, the flow path is a flow path 96 defined by a structural member 94 of the core 16 of the turbofan engine 10. Thus, the air flow injection nozzle 84 of the exemplary stage 82 may be incorporated into the above referenced Figure 2 3 and / or one or more of the exemplary turbofan engines 10 described herein. Alternatively, however, in other embodiments, the air flow injection nozzles 84 of the stage 82 may be located at any other suitable location within the turbine section, such as in the turbine aft frame 90 of the turbine section. Thus, with respect to such exemplary embodiments, the air flow injection nozzles 84 of the stage 82 may instead be incorporated into the above referenced Figure 4and / or one or more exemplary turbofan engines 10 described in conjunction with FIG. 5 .

[0193] Still refer to Figure 8 For the depicted exemplary embodiment, each of the air flow injection nozzles 84 of the stage 82 is configured as a structural guide vane of the turbine section. In particular, each of the air flow injection nozzles 84 of the stage 82 generally extends through the core air flow path 37 along the radial direction R, thereby providing a structural connection between the outer liner 126 and the turbine frame member 128 of the turbine section. However, in other embodiments, one or more of the air flow injection nozzles 84 of the stage 82 may be additionally or alternatively configured as a non-structural nozzle. In such an embodiment, the nozzle 84 may or may not extend completely through the core air flow path 37 along the radial direction R.

[0194] The air flow injection nozzle 84 of the stage 82 may include a plurality of air flow injection nozzles 84 that are spaced apart along the circumferential direction of the turbofan engine 10. For example, the air flow injection nozzle 84 of the stage 82 may include 6 or more nozzles 84, 14 or more nozzles 84, 50 or more nozzles 84, 80 or more nozzles 84, 100 or more nozzles 84, or 150 or more nozzles 84. In addition, the air flow injection nozzle 84 of the stage 82 generally includes a chamber 130 at the radially outer end. The chamber 130 may be an annular chamber that extends substantially along the circumferential direction of the turbofan engine 10. The annular chamber 130 may allow each of the air flow injection nozzles 84 of the stage 82 to be in air flow communication with the outlet 100 of the flow path 96.

[0195] In addition, each of the air flow injection nozzles 84 of the stage 82 generally includes a cavity 132 and an opening 134 (depicted in phantom), the cavity 132 extending generally along the radial direction R through the respective nozzle 84. Air from the annular chamber 130 may flow into the hollow cavity 132 and out through the opening 134 into the turbine section, or more particularly, into a portion of the core air flow path 37 extending through the turbine section. Thus, the opening 134 of each of the air flow injection nozzles 84 may be configured to inject the bleed air flow extracted from the compressor section into the core air flow path 37 extending through the turbine section. For the depicted embodiment, each opening 134 defined by the respective nozzle 84 generally extends along the radial direction R, substantially along the length of the respective nozzle 84. However, in other embodiments, each nozzle 84 may instead include a plurality of openings 134 spaced generally along the radial direction R.

[0196] Notably, the air flow injection nozzles 84 of the stage 82 are configured to reintroduce bleed air from the flow path 96. In at least some exemplary embodiments, the air flow injection nozzles 84 of the stage 82 may define a maximum mass flow rate therethrough that is equal to the maximum mass flow rate of the flow path 96, such as described above with reference to FIG. Figure 2 The maximum mass flow rate of flow path 96 is depicted.

[0197] Now, with particular reference to Fig. 9 , provides a cross-sectional view of a pair of nozzles 84 taken generally along the radial direction R according to an exemplary embodiment of the present disclosure. For example, Fig. 9 The nozzle 84 can be combined with the above reference Figure 8 The air stream of the depicted stage 82 is injected into a nozzle 84 .

[0198] for Fig. 9 , the nozzles 84 are fixed flux nozzles 84 that each define an opening 134. The depicted nozzles 84 define a pressure side 136 and a suction side 138, as well as a forward end 140 and a rearward end 142. The forward end 140 is located upstream of the rearward end 142 in the air flow through the core air flow path 37. In addition, the opening 134 of each respective nozzle 84 is located on the pressure side 136 of the respective nozzle 84 and is generally oriented along the flow direction 144 of the air flow through the turbine section (i.e., through the portion of the core air flow path 37 extending through the turbine section). This configuration may allow the bleed air flow extracted / bleeded from the compressor section to be injected into the core air flow path 37 extending through the turbine section in a direction that flows with the air flow through the turbine section (as indicated by arrow 146) so as to minimize disturbance of such air flow.

[0199] However, it should be understood that in other exemplary embodiments, the spray nozzle 84 may have any other suitable configuration. For example, in other exemplary embodiments, the nozzle 84 may instead be a variable flux nozzle 84. In particular, now referring to Fig.10 and 11 , provides a cross-sectional view of a pair of nozzles 84 also generally taken along the radial direction R according to another exemplary embodiment of the present disclosure. The exemplary air flow injection nozzle 84 can be arranged as described above with reference to Fig. 9 The exemplary air flow injection nozzles 84 depicted are constructed in substantially the same manner, and thus, the same or similar reference numbers may indicate the same or similar components.

[0200] but, Fig.10 and 11 The exemplary nozzle 84 is configured to allow for varying the amount of air flow permitted therethrough. In particular, Fig.10 and 11The opening 134 of the exemplary nozzle 84 is a variable flux opening 134 that is positioned near the rear end 142 of the corresponding nozzle 84. As shown, Fig.10 and 11 Each of the exemplary nozzles 84 includes a flap 148 positioned at the rear end 142 of the respective nozzle. Each flap 148 can generally be pivoted about a pivot axis 150 (which can be parallel to the radial direction R) in a first position or open position (see Fig.10 ) and a second position or closed position (see Fig.11 ) between the first and second positions. When the flaps are in the open position, a maximum amount of air (such as bleed air) can flow therethrough. Conversely, when in the closed position, the flaps 148 are positioned over and block the air flow through the opening 134 of the corresponding nozzle 84. Each flap 148 is also movable to a plurality of positions between the first and second positions to adjust the amount of air allowed through the air flow injection nozzle 84 of the stage 82. Fig.10 and 11 The variable flux opening 134 of the nozzle 84 depicted in FIG. 8 may be used in conjunction with or in place of a valve in the flow path, such as the valve 102 in the flow path 96 , to provide bleed air thereto.

[0201] In addition, the flaps 148 of the nozzles 84 in the air flow injection nozzles 84 of the stage 82 can also control the air flow through the core air flow path 37 of the turbine section. In particular, each pair of adjacent nozzles 84 defines a nozzle throat area, which can affect the capacity of the turbine section. When the flaps 148 are in the first open position, each pair of adjacent nozzles 84 defines a maximum nozzle throat area 152, and when the flaps 148 are in the second closed position, each pair of adjacent nozzles 84 defines a minimum nozzle throat area 154. Therefore, the flaps 148 are configured to reduce the nozzle throat area when moved to the second closed position, which can reduce the capacity of the turbine section. When the HP shaft 34 of the turbofan engine 10 is operating at a lower speed, the air flow injection nozzles 84 can operate to increase the nozzle throat area to increase the capacity of the LP turbine 30. Still alternatively, now referring to Fig.12 and 13 , provides a cross-sectional view of a pair of nozzles 84 also generally taken along the radial direction R according to another exemplary embodiment of the present disclosure. The exemplary air flow injection nozzle 84 can be arranged as described above with reference to Fig. 9 The exemplary air flow injection nozzles 84 depicted are constructed in substantially the same manner, and thus, the same or similar reference numbers may indicate the same or similar components.

[0202] as Fig.10 and 11 Like the exemplary nozzle 84, Fig.12 and 13The exemplary nozzles 84 are configured as variable flux nozzles, such that the nozzles 84 are configured to allow the amount of air flow allowed therethrough to be varied. In particular, each exemplary nozzle 84 includes a variable flux opening 134 positioned near a rear end 142 of the respective nozzle 84, a blocking door 135 that can be in a closed position ( Fig.12 ) and open position ( Fig.13 ) and slide between multiple positions therebetween. Therefore, the nozzle 84 can adjust the amount of air allowed to pass through the air flow injection nozzle 84 of the stage 82 by moving the blocking door 135 to the desired position. Fig.12 and 13 The variable flux opening 134 in the nozzle 84 depicted in FIG. 8 may be used in conjunction with or in place of a valve in the flow path (such as the valve 102 in the flow path 96 ).

[0203] Now refer to Fig.14 and 15 , an alternative embodiment is provided to inject bleed air flow extracted from the compressor section back into the core air flow path 37 extending through the turbine section. Fig.14 A simplified schematic diagram of a turbine frame member 104 is provided that is located downstream of the HP turbine 28 and upstream of the LP turbine 30. The turbine frame member 104 defines a cavity 156 that is in air flow communication with the outlet of the flow path. For the exemplary embodiment depicted, the cavity 156 is in air flow communication with the outlet 100 of the flow path 96 defined by the structural member 94, such as described above with reference to FIG. Figure 2 However, alternatively, the cavity 156 may be similar to that described above with reference to Figure 4 and 5 The bleed air flow path 114 is depicted in air flow communication.

[0204] The exemplary turbine frame member 104 generally includes a door 158 on an outer liner 160 that at least partially defines the core air flow path 37 through the turbine section. The door 158 can be generally in an axial direction A of the turbofan engine 10 in a forward open position (see Fig.14 ) and rear closed position (see Fig.15). When the door 158 is in the forward open position, the door 158 defines an opening 162 from the flow path 96 in the cavity 156 to the core air flow path 37 extending through the turbine frame member 104 to allow injection / re-introduction of bleed air from the flow path 96. Conversely, when the door 158 is in the rearward closed position, the door 158 forms a seal with the turbine outer liner 160 such that air flow from the cavity 156 or through the flow path 96 therein is not permitted to enter the core air flow path 37 through the turbine frame member 104. This can be a simple and economical configuration for injecting bleed air from the flow path 96 into the core air flow path 37 extending through the turbine frame member 104.

[0205] In certain exemplary embodiments, the turbine frame member 104 may include a plurality of doors 158, each of which may be movable between an open position and a closed position generally along the axial direction A. The plurality of doors 158 may be spaced apart along a circumferential direction of the turbofan engine 10. Additionally, while the turbine frame member 104 (including the movable doors 158) is located between the HP turbine 28 and the LP turbine 30, in other exemplary embodiments, the turbine frame member 104 (including the movable doors 158) may instead be located, for example, between a forward module of the LP turbine and an aft module of the LP turbine.

[0206] This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems, and performing any combined methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that do not differ substantially from the literal language of the claims.

Claims

1. A gas turbine engine comprising: a fan section (14) comprising a fan case (50); as well as A core turbine engine (16) is disposed downstream of the fan section and the fan case surrounds at least a portion of the core turbine engine, the core turbine engine comprising: compressor section; a combustion section downstream of the compressor section; a turbine section downstream of the combustion section; a casing (18) at least partially enclosing the compressor section, the combustion section, and the turbine section; and a structural component (94) extending from the compressor section to the turbine section and disposed radially inwardly of the casing, the structural component defining a flow path extending between an inlet in air flow communication with the compressor section and an outlet in air flow communication with the turbine section, the flow path being configured to provide a first portion of bleed air from the compressor section to the turbine section, wherein the structural member (94) extends through the outlet (100), wherein the flow path is configured to provide a second portion of bleed air from the compressor section to the turbine section through an opening (101) downstream of the outlet (100), Therein, the structural member defines a substantially annular shape enclosing at least a portion of the compressor section.

2. The gas turbine engine according to claim 1, characterized in that The turbine section includes a high pressure turbine, a low pressure turbine, and a staged air flow injection nozzle, and wherein an outlet of the flow path defined by the structural component is in air flow communication with the staged air flow injection nozzle.

3. The gas turbine engine according to claim 2, characterized in that The structural component extends through the staged air flow injection nozzles, and wherein the flow path is configured to provide a portion of bleed air from the compressor section to the turbine section at a location downstream of the air flow injection nozzles.

4. The gas turbine engine according to claim 1, characterized in that The compressor section includes a low-pressure compressor, a high-pressure compressor, and a frame member positioned between the low-pressure compressor and the high-pressure compressor, and wherein an inlet of a flow path defined by the structural member is in air flow communication with the frame member.

5. The gas turbine engine according to claim 1, characterized in that The gas turbine engine further includes a valve positioned in a flow path defined by the structural component, wherein the valve is configured to vary an amount of air flow permitted through the flow path defined by the structural component.

6. The gas turbine engine according to claim 1, characterized in that The gas turbine engine further includes a heat exchanger in air flow communication with the flow path defined by the structural component.

7. The gas turbine engine according to claim 1, characterized in that The gas turbine engine further includes an auxiliary compressor in air flow communication with the flow path defined by the structural component, the auxiliary compressor disposed inside the casing and within radially outer portions of the compressor section, the combustion section, and the turbine section.

8. The gas turbine engine according to claim 7, characterized in that The auxiliary compressor is a core-driven compressor that is mechanically coupled to the high-pressure compressor of the compressor section.

9. The gas turbine engine according to claim 7, characterized in that The auxiliary compressor includes a plurality of variable inlet guide vanes.

10. The gas turbine engine according to claim 1, characterized in that The compressor section defines a maximum mass flow rate therethrough, wherein a flow path defined by the structural component defines a maximum mass flow rate therethrough, and wherein the maximum flow rate through the flow path defined by the structural component is at least 10% of the maximum flow rate through the compressor section.

11. The gas turbine engine according to claim 10, characterized in that A maximum flow rate through a flow path defined by the structural component is at least 5% of a maximum flow rate through the compressor section.

12. The gas turbine engine according to claim 1, characterized in that The structural member extends within the housing and adds structural rigidity to the housing.

13. The gas turbine engine according to claim 1, characterized in that The turbine section includes a low-pressure turbine, a high-pressure turbine and a staged air flow injection nozzle, the low-pressure turbine includes a rotating front module and a rotating tail module, the staged air flow injection nozzle includes a plurality of air flow injection nozzles spaced apart in the circumferential direction, each air flow injection nozzle extends generally radially from an outer liner of the turbine section and toward a turbine frame member, wherein the compressor section, the combustion section and the turbine section are in a continuous flow sequence and define a core air flow path, wherein an outlet of the flow path defined by the structural member is in air contact with the staged air flow injection nozzle air flow communication, the staged air flow injection nozzle is located within the core air flow path, and wherein the structural component extends through the staged air flow injection nozzle and through a rotating tail module of the low pressure turbine, wherein a portion of the flow path defined by the structural component is configured to provide a portion of bleed air from the compressor section to the turbine section through an opening separate from an outlet of the flow path, and wherein the opening injects bleed air into the core air flow path near the rotating tail module and downstream of the staged air flow injection nozzle.

14. The gas turbine engine according to claim 1, characterized in that The compressor section includes a low-pressure compressor and a high-pressure compressor, the compressor section defines a maximum mass flow rate therethrough, wherein the turbine section includes a high-pressure turbine, a low-pressure turbine and staged air flow injection nozzles, the low-pressure turbine includes a rotating front module and a rotating tail module, the staged air flow injection nozzles include a plurality of air flow injection nozzles spaced circumferentially, each air flow injection nozzle extending generally radially from an outer liner of the turbine section and toward a turbine frame member, wherein the compressor section, the combustion section and the turbine section are in a continuous flow sequence and define a core air flow path, wherein the inlet is in air flow communication with the compressor section at a position downstream of the low-pressure compressor and upstream of the high-pressure compressor, the outlet of the flow path is in air flow communication with the turbine section at a position downstream of the high-pressure turbine, and the air flow defined by the structural components The flow path defines a maximum mass flow rate therethrough, the maximum mass flow rate through the flow path defined by the structural component being at least twenty-five percent of the maximum mass flow rate through the compressor section, wherein an outlet of the flow path defined by the structural component is in air flow communication with the staged air flow injection nozzle, the staged air flow injection nozzle is located within the core air flow path, and wherein the structural component extends through the staged air flow injection nozzle and through a rotating tail module of the low pressure turbine, wherein a portion of the flow path defined by the structural component is configured to provide a portion of bleed air from the compressor section to the turbine section through an opening separate from the outlet of the flow path, and wherein the opening injects bleed air into the core air flow path near the rotating tail module and downstream of the staged air flow injection nozzle.

15. A gas turbine engine comprising: a fan section (14) comprising a fan case (50); as well as A core turbine engine (16) is disposed downstream of the fan section and the fan case surrounds at least a portion of the core turbine engine, the core turbine engine comprising: compressor section; a combustion section downstream of the compressor section; a turbine section downstream of the combustion section; a casing at least partially enclosing the compressor section, the combustion section, and the turbine section; and a structural member extending inboard of the outer casing and adding structural support to the outer casing, the structural member further defining a flow path extending between an inlet in air flow communication with the compressor section and an outlet in air flow communication with the turbine section, wherein the structural component extends through the outlet, wherein the flow path is configured to provide a portion of bleed air from the compressor section to the turbine section through an opening downstream of the outlet, Therein, the structural member defines a substantially annular shape enclosing at least a portion of the compressor section.

16. The gas turbine engine according to claim 15, characterized in that The turbine section includes a high pressure turbine, a low pressure turbine, and a staged air flow injection nozzle, and wherein an outlet of a flow path defined by the structural component is in air flow communication with the staged air flow injection nozzle.

17. The gas turbine engine according to claim 16, characterized in that The structural component extends through the staged air flow injection nozzles, and wherein the flow path is configured to provide a portion of bleed air from the compressor section to the turbine section at a location downstream of the air flow injection nozzles.

18. The gas turbine engine according to claim 15, characterized in that The compressor section includes a low-pressure compressor, a high-pressure compressor, and a frame member positioned between the low-pressure compressor and the high-pressure compressor, and wherein an inlet of a flow path defined by the structural member is in air flow communication with the frame member.

19. The gas turbine engine according to claim 15, characterized in that The gas turbine engine further comprises A valve is positioned in a flow path defined by the structural component, wherein the valve is configured to vary an amount of air flow permitted through the flow path defined by the structural component.

20. The gas turbine engine according to claim 15, characterized in that The gas turbine engine further comprises A heat exchanger is in air flow communication with the flow path defined by the structural member.

21. The gas turbine engine according to claim 15, characterized in that The gas turbine engine further comprises An auxiliary compressor is in air flow communication with the flow path defined by the structural component, the auxiliary compressor being disposed inside the casing and within radially outer portions of the compressor section, the combustion section, and the turbine section.

22. The gas turbine engine according to claim 15, characterized in that The compressor section defines a maximum mass flow rate therethrough, wherein a flow path defined by the structural component defines a maximum mass flow rate therethrough, and wherein the maximum flow rate through the flow path defined by the structural component is at least 5% of the maximum flow rate through the compressor section.

23. The gas turbine engine according to claim 15, characterized in that The turbine section includes a low-pressure turbine, a high-pressure turbine and a staged air flow injection nozzle, the low-pressure turbine includes a rotating front module and a rotating tail module, the staged air flow injection nozzle includes a plurality of air flow injection nozzles spaced apart in the circumferential direction, each air flow injection nozzle extends generally radially from an outer liner of the turbine section and toward a turbine frame member, wherein the compressor section, the combustion section and the turbine section are in a continuous flow sequence and define a core air flow path, wherein an outlet of the flow path defined by the structural member is in air contact with the staged air flow injection nozzle air flow communication, the staged air flow injection nozzle is located within the core air flow path, and wherein the structural component extends through the staged air flow injection nozzle and through a rotating tail module of the low pressure turbine, wherein a portion of the flow path defined by the structural component is configured to provide a portion of bleed air from the compressor section to the turbine section through an opening separate from an outlet of the flow path, and wherein the opening injects bleed air into the core air flow path near the rotating tail module and downstream of the staged air flow injection nozzle.

24. The gas turbine engine according to claim 15, characterized in that The compressor section includes a low-pressure compressor and a high-pressure compressor, the compressor section defines a maximum mass flow rate therethrough, wherein the turbine section includes a high-pressure turbine, a low-pressure turbine and staged air flow injection nozzles, the low-pressure turbine includes a rotating front module and a rotating tail module, the staged air flow injection nozzles include a plurality of air flow injection nozzles spaced circumferentially, each air flow injection nozzle extending generally radially from an outer liner of the turbine section and toward a turbine frame member, wherein the compressor section, the combustion section and the turbine section are in a continuous flow sequence and define a core air flow path, wherein the inlet is in air flow communication with the compressor section at a position downstream of the low-pressure compressor and upstream of the high-pressure compressor, the outlet of the flow path is in air flow communication with the turbine section at a position downstream of the high-pressure turbine, and the air flow defined by the structural components The flow path defines a maximum mass flow rate therethrough, the maximum mass flow rate through the flow path defined by the structural component being at least twenty-five percent of the maximum mass flow rate through the compressor section, wherein an outlet of the flow path defined by the structural component is in air flow communication with the staged air flow injection nozzle, the staged air flow injection nozzle is located within the core air flow path, and wherein the structural component extends through the staged air flow injection nozzle and through a rotating tail module of the low pressure turbine, wherein a portion of the flow path defined by the structural component is configured to provide a portion of bleed air from the compressor section to the turbine section through an opening separate from the outlet of the flow path, and wherein the opening injects bleed air into the core air flow path near the rotating tail module and downstream of the staged air flow injection nozzle.

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