Turbine with axial thrust management

Abstract

A turbomachine includes a high pressure rotor shaft defining an outer surface and including a rotor seal disposed along the outer surface. A compressor exhaust chamber extends annularly around an outer surface of the high pressure rotor shaft and is at least partially formed by an inner housing assembly defining a first airflow opening. The rotor seal is configured to form a first seal against the first surface of the inner housing assembly. A pressure seal disc is coupled to the high pressure rotor shaft. The pressure seal disc is configured to form a second seal against the second surface of the inner housing assembly. The inner housing assembly, the high pressure rotor shaft, the rotor seal, and the pressure seal disc form a pressure chamber therebetween. The pressure chamber is in fluid communication with the compressor exhaust chamber via the first airflow opening.

Description

Technical Field

[0001] The present disclosure relates to a turbine. More particularly, the present disclosure relates to a turbine with axial thrust management. Background Art

[0002] Turbines, such as gas turbines, typically include a compressor section, a turbine section, and a rotor shaft or rotor shaft assembly that mechanically couples the compressor and the turbine. The rotor shaft is typically supported by various bearings that are designed to allow the rotor shaft to rotate while accommodating the radial and axial loads applied by the rotor shaft during various operating modes of the turbine. The axial load includes axial thrust. The axial thrust is typically absorbed by the rotor or thrust bearings. As the turbine cycles through various operating conditions, the axial thrust increases and decreases, leading to wear of the thrust bearings. Brief Description of the Drawings

[0003] A complete and enabling disclosure of the present disclosure, including its best mode, for a person of ordinary skill in the art, is set forth in the specification with reference to the accompanying drawings, in which:

[0004] Figure 1 is a perspective view of an exemplary aircraft in accordance with an exemplary aspect of the present disclosure.

[0005] Figure 2 is a cross-sectional view of an exemplary gas turbine in accordance with an exemplary aspect of the present disclosure.

[0006] Figure 3 is as shown in accordance with an exemplary embodiment of the present disclosure Figure 2 a schematic side view of the combustion section of the turbine.

[0007] Figure 4 is as shown in accordance with an exemplary embodiment of the present disclosure Figure 2 a schematic side view of the combustion section of the turbine.

[0008] Figure 5 is as shown in accordance with an exemplary embodiment of the present disclosure Figure 2 a schematic side view of the combustion section of the turbine.

[0009] Figure 6 is as shown in accordance with an exemplary embodiment of the present disclosure Figure 2 a schematic side view of the combustion section of the turbine.

[0010] Figure 7 is as shown in accordance with an exemplary embodiment of the present disclosure Figure 2 a schematic side view of the combustion section of the turbine.

[0011] Figure 8 is as shown in accordance with an exemplary embodiment of the present disclosureFigure 7 Schematic side view of the combustion section of the turbine shown. Detailed implementation

[0012] Reference will now be made in detail to the current embodiments of the present disclosure, one or more examples of which are shown in the accompanying drawings. The detailed description uses numerical and alphabetical identifiers to refer to features in the drawings. Similar or like identifiers in the drawings and the description have been used to refer to similar or like parts of the present disclosure.

[0013] As used herein, the term "exemplary" means "serving as an example, instance, or illustration". Any embodiment described herein as "exemplary" is not necessarily to be construed as superior or advantageous to other embodiments. Additionally, unless otherwise explicitly stated, all embodiments described herein are to be considered exemplary. Unless the context clearly dictates otherwise, the singular forms "a", "an", and "the" include plural referents. For example, in the context of "at least one of A, B, and C", the term "at least one" refers to only A, only B, only C, or any combination of A, B, and C.

[0014] As used herein, the terms "first", "second", and "third" may be used interchangeably to distinguish one component from another and are not intended to denote the position or importance of the respective components. Additionally, the terms "upstream" and "downstream" refer to the relative direction with respect to the fluid flow in a fluid path. For example, "upstream" refers to the direction from which the fluid flows, while "downstream" refers to the direction towards which the fluid flows. The terms "front" and "rear" refer to the position or orientation of a component or feature of a turbine or aircraft relative to another component or feature. The foremost position is typically associated with the inlet of the turbine.

[0015] The term "turbine" refers to a machine that includes one or more compressors, a heat generation section (such as a combustion section), and one or more turbines that together produce a torque output. The term "gas turbine" refers to an engine that has a turbine as all or part of its power source. Example gas turbines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, etc., as well as hybrid electric versions of one or more of these engines.

[0016] The present disclosure generally relates to managing the axial thrust loads generated during various operating modes of a turbine. Axial thrust loads include axial thrust, which is the force generated by aerodynamic and pressure forces, and axial thrust applies an axial force to the rotor shaft in the compressor and turbine, and may also include all pressure and vibratory forces acting on the rotor shaft in the axial direction. The present disclosure provides for creating a cavity or pressure chamber radially inward from the compressor exhaust chamber of the turbine to reduce or manage any negative impacts of the axial thrust loads.

[0017] A pressure chamber is formed between a rotor seal, a pressure seal disk coupled to a high-pressure rotor shaft, and a wall of an inner housing assembly that at least partially defines a compressor exhaust chamber. A portion of the compressed air from the compressor exhaust chamber is directed into the pressure chamber. The pressure within the pressure chamber exerts an axial force on a face or sidewall of the pressure seal disk. The axial force exerted on the face or sidewall resists or counteracts the axial thrust acting on the rotor shaft. This helps to relieve the axial thrust load on the axial bearing that partially supports the high-pressure rotor shaft, thereby reducing wear on the axial bearing and increasing the interval between required maintenance.

[0018] Now referring to the drawings, Figure 1 is a perspective view of an exemplary aircraft 10 that may incorporate at least one exemplary embodiment of the present disclosure. As Figure 1 shown, the aircraft 10 has a fuselage 12, wings 14 attached to the fuselage 12, and a tail 16. The aircraft 10 also includes a propulsion system 18 that generates propulsion thrust to propel the aircraft 10 during flight, taxi operations, and the like. Although the propulsion system 18 is shown attached to the wings 14, in other embodiments it may additionally or alternatively include one or more aspects coupled to other parts of the aircraft 10, such as the tail 16, the fuselage 12, or both. The propulsion system 18 includes at least one turbine. In the exemplary embodiment shown, the aircraft 10 includes a pair of gas turbine engines 20. Each gas turbine engine 20 is mounted to the aircraft 10 in an under-wing configuration. Each gas turbine engine 20 is capable of selectively generating propulsion thrust for the aircraft 10. The gas turbine engines 20 may be configured to combust various forms of fuel, including but not limited to jet fuel / aviation turbine fuel and hydrogen fuel, unless otherwise specified.

[0019] Figure 2 is a cross-sectional side view of a gas turbine engine 20 according to an exemplary embodiment of the present disclosure. More specifically, for Figure 2 the embodiment shown, the gas turbine engine 20 is a multi-spool high-bypass turbofan jet engine, sometimes also referred to as a "turbofan engine". As Figure 2 shown, the gas turbine engine 20 defines an axial direction "A" (extending parallel to a longitudinal centerline 22 provided for reference), a radial direction "R", and a circumferential direction "C" extending about the longitudinal centerline 22. Generally, the gas turbine engine 20 includes a fan section 24 and a turbine engine or turbine 26 disposed downstream of the fan section 24.

[0020] The turbine 26 shown generally includes an engine casing, housing, or core cowl 28 that defines an annular core inlet 30. The core cowl 28 at least partially surrounds, in a series flow relationship, a compressor section that includes a booster or low-pressure compressor 32 and a high-pressure compressor 34; a combustion section 35 that includes a burner 36; a turbine section that includes a high-pressure turbine 38 and a low-pressure turbine 40; and at least a portion of an exhaust jet nozzle 42. These components or sections together constitute the core engine of the turbine 26.

[0021] The turbine 26 includes a rotor shaft assembly 44 that includes at least one rotor shaft. In Figure 2 the exemplary embodiment shown, the rotor shaft assembly 44 includes a high-pressure rotor shaft 46 that drivingly connects the high-pressure turbine 38 to the high-pressure compressor 34. A low-pressure rotor shaft 48 drivingly connects the low-pressure turbine 40 to the low-pressure compressor 32. The compressor section, combustion section 35, turbine section, and exhaust jet nozzle 42 together define a working gas flow path 50 through the gas turbine engine 20.

[0022] For the embodiment shown, the fan section 24 includes a fan 52 that has a plurality of fan blades 54 that are circumferentially spaced and coupled to a disk 56. As shown, the fan blades 54 extend generally radially outward from the disk 56 along a radial direction R. Each fan blade 54 is operatively coupled to a pitch changing mechanism 58 such that each fan blade 54 can rotate with the disk 56 about a pitch axis P, and the pitch changing mechanism 58 is configured to commonly change the pitch of the fan blades 54, e.g., to change the pitch of the fan blades 54 uniformly. The fan blades 54, disk 56, and pitch changing mechanism 58 can together rotate about a longitudinal centerline 22 via the low-pressure rotor shaft 48.

[0023] In an exemplary embodiment, as Figure 2 shown, the gas turbine engine 20 further includes a power gearbox or gearbox 60. The gearbox 60 includes a plurality of gears for adjusting the rotational speed of the fan 52 relative to the low-pressure rotor shaft 48 such that the fan 52 and the low-pressure rotor shaft 48 can rotate at a more efficient relative speed. The gearbox 60 can be any type of gearbox suitable for facilitating the connection of the low-pressure rotor shaft 48 to the fan 52 while allowing each of the low-pressure turbine 40 and the fan 52 to operate at a desired relative speed. For example, in some embodiments, the gearbox 60 can be a reduction gearbox.

[0024] Still referring to Figure 2In an exemplary embodiment, the disk 56 is connected to the gearbox 60 via the fan shaft 62. The disk 56 is covered by the front hub 64 (sometimes also referred to as the "spinner") of the fan section 24. The front hub 64 has an aerodynamic profile to facilitate the flow of air through or across the plurality of fan blades 54. Additionally, the fan section 24 includes an annular fan casing or nacelle 66 that circumferentially surrounds at least a portion of the fan 52 and the turbine 26. In the illustrated embodiment, the nacelle 66 is supported relative to the turbine 26 by a plurality of circumferentially spaced struts or exit guide vanes 68. Further, a downstream section 70 of the nacelle 66 extends over an outer portion of the turbine 26 to define a bypass air flow passage 72 therebetween.

[0025] In an exemplary embodiment, the rotor shaft assembly 44 is at least partially supported by one or more bearings 74, 76. At least one of the one or more bearings 74, 76 is an axial thrust bearing or thrust bearing configured to accommodate the aerodynamic and pressure forces that exert axial forces on the high-pressure rotor shaft 46 and the low-pressure rotor shaft 48 when the turbine 26 cycles under various operating conditions. For example, the axial force may increase or decrease under certain operating conditions.

[0026] Figure 3 is according to an exemplary embodiment of the present disclosure as Figure 2 shown in a schematic side view of the combustion section 35 of the turbine 26. As Figure 3 shown, the combustion section 35 includes an inner casing assembly 78 and an outer casing assembly 80. The inner casing assembly 78 and the outer casing assembly 80 form a compressor exhaust chamber 82 therebetween. The burner 36 is typically disposed within the compressor exhaust chamber 82.

[0027] In operation, compressed air "CA" from the high-pressure compressor 34 ( Figure 2 ) flows through a flow regulator 84 disposed at the inlet 86 of the compressor exhaust chamber 82 and into the compressor exhaust chamber 82. A first portion "CA1" of the compressed air CA is directed around the fuel nozzle 88 and into the combustion chamber 90 of the burner 36 disposed within the compressor exhaust chamber 82, where the fuel is mixed with the first portion CA1 of the compressed air CA and burned to produce combustion gases "CG". The combustion gases CG are then directed across a row or stage of stationary or stator vanes 92 (only one is shown), and then across a row of turbine rotor blades 94 (only one is shown), thereby causing the high-pressure rotor shaft 46 to rotate about the longitudinal centerline 22. The turbine rotor blades 94 are fixed to the high-pressure rotor shaft 46 via a turbine rotor disk 96 (or other suitable means), and the turbine rotor disk 96 is coupled to the high-pressure rotor shaft 46 and at least partially defines the high-pressure rotor shaft 46.

[0028] In an exemplary embodiment, the high-pressure rotor shaft 46 includes or defines a radially outer surface or outer surface 98. The rotor seal 100 extends circumferentially along the outer surface 98. The compressor exhaust chamber 82, and more specifically, the inner housing assembly 78 and the outer housing assembly 80, extend annularly around the outer surface 98 of the high-pressure rotor shaft 46 relative to the longitudinal centerline 22. The wall 102 of the inner housing assembly 78 defines a first air flow opening 104. The rotor seal 100 is configured or formed to form a first seal "S1" against a first surface 106 of the inner housing assembly 78. The pressure seal disk 108 is coupled to the high-pressure rotor shaft 46 and rotates therewith. The pressure seal disk 108 is disposed rearward of the flow regulator 84, rearward of the rotor seal 100, and forward of the turbine rotor disk 96, and is radially inward from the wall 102 of the inner housing assembly 78 relative to the axial direction A, the radial direction R, and the longitudinal centerline 22.

[0029] The sealing or radially distal end or sealing surface 110 of the pressure seal disk 108 is configured or formed to form a second seal "S2" against a second surface 112 of the inner housing assembly 78. The inner housing assembly 78, the high-pressure rotor shaft 46, the rotor seal 100, and the pressure seal disk 108 together form a pressure chamber 114 therebetween. The pressure chamber 114 is in fluid communication with the compressor exhaust chamber 82 via the first air flow opening 104.

[0030] In operation, a second portion "CA2" of the compressed air CA from the compressor exhaust chamber 82 flows into the pressure chamber 114 via the first air flow opening 104 to pressurize or inflate the pressure chamber 114 at a first pressure "P1", thereby applying an axial load (as indicated by the arrow "AL") against the front side wall or front side surface 116 of the pressure seal disk 108 relative to the axial direction A. In an exemplary embodiment, when the turbine operates or transitions between various operating conditions, the axial load AL is used to reduce or manage the forward axial movement of the high-pressure rotor shaft 46. When the pressure "P" in the compressor exhaust chamber 82 drops below the first pressure P1 in the pressure chamber 114, the second portion CA2 of the compressed air CA flows out of the pressure chamber 114 via the first air flow opening 104 and returns to the compressor exhaust chamber 82.

[0031] Figure 4 is a schematic side view of the combustion section 35 of the turbine 26 according to an exemplary embodiment of the present disclosure. It should be understood that Figure 2 as shown. The combustion section shown is similar to Figure 4 the combustion section 35 shown in Figure 3 with similar components, numbers, and functions. As Figure 4As shown, the inner housing assembly 78 can define a second air flow opening 118 that is disposed rearward of a pressure seal disk 108 within the compressor discharge chamber 82 relative to the axial direction A and the longitudinal centerline 22. The second air flow opening 118 is in fluid communication with the compressor discharge chamber 82 to permit a third portion CA3 of the compressed air CA to flow into a rear pressure chamber 120 defined rearward of the pressure seal disk 108. The rear pressure chamber 120 is generally defined between a second or rear side 122 of the pressure seal disk 108 and the turbine rotor disk 96. During some or all operating conditions of the turbine 26, the pressure “PA” within the rear pressure chamber 120 can be lower or higher than a first pressure P1 of a second portion CA2 of the compressed air CA within the pressure chamber 114.

[0032] In a particular embodiment, as Figure 3 and Figure 4 collectively shown, at least one of the first air flow opening 104 and the second air flow opening 118 includes an insert 124 or an insert 126 disposed therein, respectively. The insert 124 can be configured to control or meter the flow of a second portion CA2 of the compressed air CA from the compressor discharge chamber 82 into and out of the pressure chamber 114. For example, the insert 124 can have a predetermined diameter for a fixed flow rate or can be made of a shape memory alloy for a variable flow rate that depends on the temperature of the compressed air CA, which can fluctuate depending on the operating mode of the turbine 26. Similarly, the insert 126 can be configured to control or meter the flow of a third portion CA3 of the compressed air CA from the compressor discharge chamber 82 into and out of the rear pressure chamber 120 ( Figure 4 ). For example, the insert 126 can have a predetermined diameter for a fixed flow rate or can be made of a shape memory alloy for a variable flow rate that depends on the temperature of the compressed air CA, which can fluctuate depending on the operating mode of the turbine 26.

[0033] Figure 5 is a schematic side view of a combustion section 35 of a turbine 26 according to an exemplary embodiment of the present disclosure. It should be understood that Figure 2 the combustion section shown therein is similar to Figure 5 the combustion section 35 shown in Figure 3 and Figure 4 with similar components, numbers, and functions. As Figure 5As shown, the high-pressure rotor shaft 46 defines a first bypass opening 128 that is defined relative to the axial direction A and the longitudinal centerline 22 and is located behind the flow regulator 84 and in front of the rotor seal 100. In an exemplary embodiment, the high-pressure rotor shaft 46 further defines a second bypass opening 130 that is defined relative to the axial direction A and the longitudinal centerline 22 and is located behind the pressure seal disk 108 and in front of the turbine rotor disk 96.

[0034] The inner wall 132 (radially inward relative to the radial direction R) of the high-pressure rotor shaft 46 at least partially forms a bypass air chamber 134 within the high-pressure rotor shaft 46. In these embodiments, another portion of the compressed air CA (designated as "CA4") flows out of the compressor discharge chamber 82, through an opening 136 defined in the inner housing assembly 78, through the first bypass opening 128 and into the bypass air chamber 134, passes through the pressure seal disk 108, flows out of the bypass air chamber 134 via the second bypass opening 130, and enters the rear pressure chamber 120.

[0035] In a particular embodiment, as Figure 5 shown, at least one of the first bypass opening 128 and the second bypass opening 130 includes inserts 138, 140 disposed therein, respectively. The insert 138 can be configured to control or meter the flow of the portion CA4 of the compressed air CA from the compressor discharge chamber 82 into the bypass air chamber 134. For example, the insert 138 can have a predetermined diameter for a fixed flow rate, or can be made of a shape memory alloy for a variable flow rate that depends on the temperature of the compressed air CA, which may fluctuate depending on the operating mode of the turbine 26. Similarly, the insert 140 can be configured to control or meter the flow of the portion CA4 of the compressed air CA from the bypass air chamber 134 into the rear pressure chamber 120. For example, the insert 140 can have a predetermined diameter for a fixed flow rate, or can be made of a shape memory alloy for a variable flow rate that depends on the temperature of the compressed air CA, which may fluctuate depending on the operating mode of the turbine 26. The portion CA4 of the compressed air CA can apply an axial load to the pressure seal disk 108 in the forward direction, the backward direction, or both the forward and backward directions relative to the axial direction A and the longitudinal centerline 22 to reduce the axial thrust on the axial thrust bearing during the care load.

[0036] Figure 6 is a schematic side view of the combustion section 35 of the turbine 26 according to an exemplary embodiment of the present disclosure. It should be understood that Figure 2 the combustion section shown in Figure 6 is similar to the combustion section 35 shown in Figure 3 , 4 and 5, having similar components, numbers, and functions. However, Figure 6Exemplary embodiments include a pump 142 fluidly coupled via various conduits, pipes, connectors, or other suitable connections to a compressor discharge chamber 82 and a first airflow opening 104. The pump 142 is configured to control or regulate a first pressure P1 of a second portion CA2 of compressed air CA within a pressure chamber 114 during operation of the turbine 26. The pump 142 may be configured to provide a constant first pressure P1 or regulate the first pressure P1 within the pressure chamber 114 to a desired pressure to achieve a desired axial load AL on the pressure seal disk 108, regardless of the operating conditions of the turbine or the pressure P within the compressor discharge chamber 82.

[0037] Figure 7 is a schematic side view of a combustion section of a turbine 26 according to an exemplary embodiment of the present disclosure. It should be understood that Figure 2 the combustion section 35 shown in Figure 7 is similar to the combustion section 35 shown in Figure 3 and Figure 4 having similar components, numbers, and functions. As Figure 7 shown, an inner housing assembly 78 defines a first airflow opening 104 and a second airflow opening 118. A pressure seal disk 108 is coupled to a high-pressure rotor shaft 46. The pressure seal disk 108 includes a front side 116 and a sealing surface 110. The pressure seal disk 108 also includes or defines a rear side 122.

[0038] The pressure chamber 114 is at least partially defined by the high-pressure rotor shaft 46, the inner housing assembly 78, and the front side 116 of the pressure seal disk 108. The pressure chamber 114 is in fluid communication with the compressor discharge chamber 82 via the first airflow opening 104. A second pressure chamber 144 is at least partially defined by the high-pressure rotor shaft 46, the inner housing assembly 78, and the rear side 122 of the pressure seal disk 108. The second pressure chamber 144 is in fluid communication with the compressor discharge chamber 82 via the second airflow opening 118.

[0039] In an exemplary embodiment, a rotor seal 100 is configured to form a first seal S1 against a first surface 106 of the inner housing assembly 78. The sealing surface 110 of the pressure seal disk 108 is configured to form a second seal S2 against a second surface 112 of the inner housing assembly 78. A second rotor seal 146 is configured to form a third seal "S3" against a third surface 150 of the inner housing assembly 78. The sizes of the first airflow opening 104 and the second airflow opening 118 may be designed to control a first pressure P1 in the first pressure chamber 114 and a second pressure P2 in the second pressure chamber 144. The first pressure P1 and the second pressure P2 may be predefined to provide a reaction axial load AL against the pressure seal disk 108 to reduce or prevent an axial thrust load on the axial thrust bearing.

[0040] In other embodiments, one or more of the first air flow opening 104 and the second air flow opening 118 may include an insert 124 or an insert 126 to control or meter the flow of a second portion CA2 of the compressed air CA from the compressor discharge chamber 82 into and out of the pressure chamber 114. For example, the insert 124 may have a predetermined diameter for a fixed flow rate or may be made of a shape memory alloy for a variable flow rate that depends on the temperature of the compressed air CA, which may fluctuate according to the operating mode of the turbine 26. Similarly, the insert 126 may be configured to control or meter the flow of a third portion CA3 of the compressed air CA from the compressor discharge chamber 82 into and out of the second pressure chamber 144. For example, the insert 126 may have a predetermined diameter for a fixed flow rate or may be made of a shape memory alloy for a variable flow rate that depends on the temperature of the compressed air CA, which may fluctuate according to the operating mode of the turbine 26.

[0041] Figure 8 is a schematic side view of a combustion section of a turbine 26 as Figure 7 shown in accordance with an exemplary embodiment of the present disclosure. It should be understood that Figure 8 the combustion section 35 shown in Figure 7 is similar to the combustion section 35 shown in

[0042] and has similar components, numbers, and functions. In an exemplary embodiment, the flow control valve 152 is fluidly coupled to the compressor discharge chamber 82 via one or more fluid conduits or pipes coupled to an inlet 154 of the flow control valve 152. The flow control valve 152 also includes a first outlet 156 and a second outlet 158. The first outlet 156 is fluidly coupled to and in fluid communication with the first pressure chamber 114 via the first air flow opening 104, and the second outlet 158 is in fluid communication with the second pressure chamber 144 via the second air flow opening 118. In an exemplary embodiment, the flow control valve 152 is configured to control a first pressure P1 of a second portion CA2 of the compressed air CA in the first pressure chamber 114. Additionally or alternatively, the flow control valve 152 may be configured to control a second pressure P2 of a third portion CA3 of the compressed air CA in the second pressure chamber 144. The flow control valve 152 may be configured to adjust the first pressure P1 and the second pressure P2 to provide a desired axial pressure or axial load AL, either axially forward or axially backward, on the pressure seal disk 108 and the high-pressure rotor shaft 46 according to the operating conditions of the turbine 26 to reduce or prevent an axial thrust load on the axial thrust bearing.

[0042] Under certain operating conditions, vibrations may occur on the high-pressure rotor shaft 46. The most severe vibrations are typically observed when the rotor thrust is minimal, zero, or during thrust crossover conditions. The present disclosure provides an effective method of changing or cushioning the rotor thrust to mitigate these high vibrations in the high-pressure rotor shaft 46. The components described in the present disclosure avoid high N2 vibration problems without installing additional hardware to the turbine, thereby applying additional axial forward or backward thrust, thus protecting against possible thrust reversal / N2 vibrations at the bearing locations.

[0043] A further aspect is provided by the subject matter of the following clauses:

[0044] A turbine, comprising: a high-pressure rotor shaft extending along a longitudinal centerline of the turbine, the high-pressure rotor shaft having an outer surface and a rotor seal disposed along the outer surface; a compressor exhaust chamber extending annularly around the outer surface of the high-pressure rotor shaft, wherein the compressor exhaust chamber is at least partially formed by an inner housing assembly, wherein the inner housing assembly defines a first airflow opening, and wherein the rotor seal is configured to form a first seal against a first surface of the inner housing assembly; and a pressure seal disk coupled to the high-pressure rotor shaft, wherein the pressure seal disk is configured to form a second seal against a second surface of the inner housing assembly, wherein the inner housing assembly, the high-pressure rotor shaft, the rotor seal, and the pressure seal disk form a pressure chamber therebetween, wherein the pressure chamber is in fluid communication with the compressor exhaust chamber via the first airflow opening, and wherein compressed air from the compressor exhaust chamber inflates the pressure chamber and applies an axial load to the face of the pressure seal disk during operation of the turbine.

[0045] A turbine, comprising: a high-pressure rotor shaft defining an outer surface and including a rotor seal disposed along the outer surface; a compressor exhaust chamber extending annularly around the outer surface of the high-pressure rotor shaft; an inner housing assembly defining a first airflow opening, wherein the rotor seal is configured to form a first seal against a first surface of the inner housing assembly, wherein the compressor exhaust chamber is at least partially formed by the inner housing; and a pressure seal disk coupled to the high-pressure rotor shaft, wherein the pressure seal disk is configured to form a second seal against a second surface of the inner housing assembly, wherein the inner housing assembly, the high-pressure rotor shaft, the rotor seal, and the pressure seal disk form a pressure chamber therebetween, wherein the pressure chamber is in fluid communication with the compressor exhaust chamber via the first airflow opening.

[0046] The turbine according to the foregoing or any of the following clauses, further comprising a first insert disposed within the first airflow opening, wherein the insert is configured to control the flow of compressed air from the compressor exhaust chamber into and out of the pressure chamber.

[0047] A turbine according to any of the foregoing or following clauses, wherein the inner housing assembly defines a second air flow opening disposed within the compressor exhaust chamber, behind the pressure seal disk, and wherein the second air flow opening is in fluid communication with the compressor exhaust chamber.

[0048] A turbine according to any of the foregoing or following clauses, further comprising a pump, wherein the pump is fluidly coupled to the compressor exhaust chamber and the first air flow opening, and wherein the pump is configured to control the pressure of the compressed air from the compressor exhaust chamber inside the pressure chamber.

[0049] A turbine according to any of the foregoing or following clauses, further comprising a turbine rotor disk coupled to the high-pressure rotor shaft, wherein the pressure seal disk is disposed in front of the turbine rotor disk with respect to the axial centerline.

[0050] A turbine according to any of the foregoing or following clauses, further comprising a compressor flow regulator disposed at an inlet of the compressor exhaust chamber, and a turbine rotor disk disposed downstream of the compressor exhaust chamber, wherein the pressure seal disk is disposed between the compressor flow regulator and the turbine rotor disk.

[0051] A turbine according to any of the foregoing or following clauses, further comprising a burner disposed within the compressor exhaust chamber.

[0052] A turbine according to any of the foregoing or following clauses, wherein the rotor shaft defines a first bypass opening, and wherein the first bypass opening is defined in front of the rotor seal.

[0053] A turbine according to any of the foregoing or following clauses, wherein the rotor shaft defines a second bypass opening, and wherein the second bypass opening is defined behind the pressure seal disk and in front of the turbine rotor disk of the turbine.

[0054] A turbine according to any of the foregoing or following clauses, further comprising at least one of a first bypass insert disposed within the first bypass opening and a second bypass insert disposed within the second bypass opening.

[0055] A turbine, comprising: a high-pressure rotor shaft defining an outer surface; an inner housing assembly; a compressor exhaust chamber at least partially formed by the inner housing assembly, wherein the inner housing assembly defines a first air flow opening and a second air flow opening; a pressure seal disk coupled to the high-pressure rotor shaft, the pressure seal disk having a front side, a sealing surface, and a rear side; a first pressure chamber at least partially defined by the high-pressure rotor shaft, the inner housing assembly, and the front side of the pressure seal disk, wherein the first pressure chamber is in fluid communication with the compressor exhaust chamber via the first air flow opening; and a second pressure chamber at least partially defined by the high-pressure rotor shaft, the inner housing assembly, and the rear side of the pressure seal disk, wherein the second pressure chamber is in fluid communication with the compressor exhaust chamber via the second air flow opening.

[0056] The turbine according to any of the foregoing or following clauses further includes a first rotor seal and a second rotor seal, wherein the first rotor seal is configured to form a first seal against a first surface of the inner housing assembly, the sealing surface of the pressure seal disc is configured to form a second seal against a second surface of the inner housing assembly, and wherein the second rotor seal is configured to form a third seal against a third surface of the inner housing assembly.

[0057] The turbine according to any of the foregoing or following clauses further includes a flow control valve having an inlet in fluid communication with the compressor discharge chamber, a first outlet in fluid communication with a first pressure chamber via a first air flow opening, and a second outlet in fluid communication with a second pressure chamber via a second air flow opening.

[0058] The turbine according to any of the foregoing or following clauses, wherein the flow control valve is configured to control or regulate a first pressure in the first pressure chamber and a second pressure in the second pressure chamber.

[0059] An aircraft includes: a fuselage; and a turbine. The turbine includes: a rotor shaft defining an axial centerline, the rotor shaft having an outer surface and a rotor seal disposed along the outer surface; a compressor discharge chamber annularly extending around the outer surface of the rotor shaft, wherein the compressor discharge chamber is at least partially formed by an inner housing assembly, wherein the inner housing assembly defines a first air flow opening, and wherein the rotor seal is configured to form a first seal against a first surface of the inner housing assembly; and a pressure seal disc coupled to the rotor shaft, wherein the pressure seal disc is configured to form a second seal against a second surface of the inner housing assembly, wherein the inner housing assembly, the rotor shaft, the rotor seal, and the pressure seal disc form a pressure chamber therebetween, and wherein the pressure chamber is in fluid communication with the compressor discharge chamber via the first air flow opening.

[0060] An aircraft includes a fuselage; and a turbine including a high-pressure rotor shaft having an outer surface and a rotor seal disposed along the outer surface; an inner housing assembly; a compressor discharge chamber annularly extending around the outer surface of the high-pressure rotor shaft, wherein the inner housing assembly defines a first air flow opening, and wherein the rotor seal is configured to form a first seal against a first surface of the inner housing assembly; and a pressure seal disc coupled to the high-pressure rotor shaft, wherein the pressure seal disc is configured to form a second seal against a second surface of the inner housing assembly, wherein the inner housing assembly, the high-pressure rotor shaft, the rotor seal, and the pressure seal disc form a pressure chamber therebetween, and wherein the pressure chamber is in fluid communication with the compressor discharge chamber via the first air flow opening.

[0061] An aircraft according to any of the foregoing or following clauses, wherein the turbine further includes a first insert disposed within the first air flow opening, and the insert is configured to control the flow of compressed air from the compressor exhaust chamber into and out of the pressure chamber.

[0062] An aircraft according to any of the foregoing or following clauses, wherein the inner housing assembly defines a second air flow opening disposed within the compressor exhaust chamber and behind the pressure seal disc, and the second air flow opening is in fluid communication with the compressor exhaust chamber.

[0063] An aircraft according to any of the foregoing or following clauses, wherein the turbine further includes a pump, and the pump is fluidly coupled to the compressor exhaust chamber and the first air flow opening, and the pump is configured to control the pressure of the compressed air from the compressor exhaust chamber on the inner side of the pressure chamber.

[0064] An aircraft according to any of the foregoing or following clauses, wherein the turbine further includes a turbine rotor disc coupled to the rotor shaft, and the pressure seal disc is disposed in front of the turbine rotor disc with respect to the axial centerline.

[0065] An aircraft according to any of the foregoing or following clauses, wherein the turbine further includes a compressor flow regulator disposed at the inlet of the compressor exhaust chamber, and a turbine rotor disc disposed downstream of the compressor exhaust chamber, and the pressure seal disc is disposed between the compressor flow regulator and the turbine rotor disc.

[0066] An aircraft according to any of the foregoing or following clauses, wherein the turbine further includes a burner disposed within the compressor exhaust chamber.

[0067] An aircraft according to any of the foregoing or following clauses, wherein the rotor shaft defines a first bypass opening, and the first bypass opening is defined in front of the rotor seal.

[0068] An aircraft according to any of the foregoing or following clauses, wherein the rotor shaft defines a second bypass opening, and the second bypass opening is defined behind the pressure seal disc and in front of the turbine rotor disc of the turbine.

[0069] An aircraft according to any of the foregoing or following clauses, wherein the turbine further includes at least one of a first bypass insert disposed within the first bypass opening and a second bypass insert disposed within the second bypass opening.

[0070] This written description uses examples to disclose the present disclosure, including the best mode, and also enables any person skilled in the art to practice the present disclosure, including making and using any device or system and performing any incorporated method. The patentable scope of the present disclosure is defined by the claims and may include other examples that occur to those skilled in the art. If these other examples include structural elements that are identical to the literal language of the claims or if they include equivalent structural elements that are not materially different from the literal language of the claims, then these other examples are intended to fall within the scope of the claims.

Claims

1. A turbine, characterized in that: include: a high pressure rotor shaft defining an outer surface and including a rotor seal disposed along the outer surface; a compressor discharge chamber extending annularly around the outer surface of the high-pressure rotor shaft; an inner casing assembly defining a first air flow opening, wherein the rotor seal is configured to form a first seal against a first surface of the inner casing assembly, wherein the compressor discharge plenum is at least partially formed by the inner casing; as well as a pressure sealing disk coupled to the high-pressure rotor shaft, wherein the pressure sealing disk is configured to form a second seal against a second surface of the inner casing assembly, wherein the inner casing assembly, the high-pressure rotor shaft, the rotor seal, and the pressure sealing disk form a pressure chamber therebetween, wherein the pressure chamber is in fluid communication with the compressor discharge chamber via the first airflow opening.

2. The turbine according to claim 1, characterized in that Further included is a first insert disposed within the first air flow opening, wherein the first insert is configured to control the flow of compressed air from the compressor discharge chamber into and out of the pressure chamber.

3. The turbine according to claim 2, characterized in that The inner housing assembly defines a second air flow opening disposed within the compressor discharge chamber and rearward of the pressure sealing disk, wherein the second air flow opening is in fluid communication with the compressor discharge chamber.

4. The turbine according to claim 1, characterized in that Further included is a pump, wherein the pump is fluidly coupled to the compressor discharge plenum and the first air flow opening, wherein the pump is configured to control a pressure of compressed air from the compressor discharge plenum within the pressure chamber.

5. The turbine according to claim 1, characterized in that Further included is a turbine rotor disk coupled to the high pressure rotor shaft, wherein the pressure sealing disk is disposed forward of the turbine rotor disk.

6. The turbine according to claim 1, characterized in that The invention further includes a compressor flow conditioner disposed at an inlet of the compressor discharge chamber, and a turbine rotor disk disposed downstream of the compressor discharge chamber, wherein the pressure sealing disk is disposed between the compressor flow conditioner and the turbine rotor disk.

7. The turbine according to claim 1, characterized in that Further comprising a combustor disposed within the compressor exhaust chamber.

8. The turbine according to claim 1, characterized in that The high pressure rotor shaft defines a first bypass opening, wherein the first bypass opening is defined in front of the rotor seal.

9. The turbine according to claim 8, characterized in that The high pressure rotor shaft defines a second bypass opening, wherein the second bypass opening is defined behind the pressure sealing disk and in front of a turbine rotor disk of the turbine.

10. The turbine according to claim 9, characterized in that Further included is at least one of a first bypass insert disposed within the first bypass opening and a second bypass insert disposed within the second bypass opening.

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