WO2020131030A1

WO2020131030A1 - Gas turbine engine with a pre-swirl cavity

Info

Publication number: WO2020131030A1
Application number: PCT/US2018/066211
Authority: WO
Inventors: David May; Kok-Mun Tham
Original assignee: Siemens Aktiengesellschaft; Siemens Energy, Inc.
Priority date: 2018-12-18
Filing date: 2018-12-18
Publication date: 2020-06-25

Abstract

A gas turbine engine comprises a first passage, a second passage and a bypass channel. The first passage transmits a first leakage flow (335) from a bypass channel chamber to a pre-swirl supporting structure (310) and through an inner seal (330). The second passage transmits a second leakage flow (340) from a pre-swirl cavity (325) and through the inner seal (330). The bypass channel has a rim seal cavity at its end. The rim seal cavity has a plurality of ejection holes (315(1-M) to eject a flow back to a gas path. A combined leakage flow (345) of the first leakage flow (335) and the second leakage flow (340) is bypassed through the bypass channel to purge the rim seal cavity with the combined leakage flow (345) through the plurality of ejection holes (315(1-M)).

Description

GAS TURBINE ENGINE WITH A PRE-SWIRL CAVITY

BACKGROUND

1. Field

[0001] Aspects of the present invention generally relate, in a gas turbine pre-swirl system and a gas turbine blade cooling system, to exit holes placed at circumferential positions of gas path pressure peaks to form a non-axisymmetrie purge flow via a rim seal of a gas turbine engine to counter hot gas ingestion.

2. Description of the Related Art

[0002] In gas turbine engines, compressed air discharged from a compressor section and fuel introduced from a source of fuel are mixed together and burned in a combustion section, creating combustion products defining hot working gases. The working gases are directed through a hot gas path in a turbine section, where they expand to provide rotation of a turbine rotor. The turbine rotor may be linked to an electric generator, wherein the rotation of the turbine rotor can be used to produce electricity in the generator.

[0003] In view of high pressure ratios and high engine firing temperatures implemented in modern engines, certain components, such as rotating blade structures within the turbine section, must be cooled with cooling fluid, such as compressor discharge air, to prevent overheating of the components.

[0004] A high pressure turbine (HPT) cooling air delivery system traditionally includes a pre-swirl system. The system takes air from the high pressure compressor and injects it in front of the HPT disc through angled holes or nozzles (pre-swirlers). The angle is set so that the air tangential velocity ejected through the pre-swirls matches that of the disc speed. This yields minimum air temperature rise when washing the surface of the disc with incoming air. The air also feeds the HPT blade internal passage. For the same reason as the disc, it is desirable to feed air at the same tangential velocity as the blade/disc so the blade is cooled using the lowest possible feed air temperature.

[0005] The static and rotating surfaces that contain the swirled cooling air, i.e. the pre swirl cavity, are traditionally sealed by labyrinth seals. There is a certain amount of leakage (around 15% of total pre-swirl flow) into the pre-swirl cavity through an inner seal.

[0006] In a gas turbine blade cooling system, a cooling air delivery from a non rotating hardware to the rotating components has to be done as efficiently as possible. Modern turbine pre-swirl systems use a vane pack to accelerate and turn the cooling air originating from non-rotating components in the direction of rotation, to ensure efficient delivery to the rotating components. For example, cooled cooling air goes through pre-swirl vanes. Exiting vanes at near or slightly above local rotor speed, the intended blade coolant then enters a disk receiver hole with minimal pressure loss and rotor work input.

[0007] Key to a pre-swirl system efficiency is the effective swirl in a pre-swirl cavity. If the effective swirl matches the local rotor speed, the flow will enter the rotating system axially, with minimal pressure loss and parasitic work input. If pre-swirl cavity effective swirl was lower than the rotor speed at the disk receiver holes, then Euler work (Euler work relates to the rotational work done on/by the air, in the gas turbine. Say in the pre swirl flow context. If the flow is rotating slower than the turbine disc, then work has to be done on the air to change its angular momentum / accelerate the air in the circumferential direction, to match the rotational speed of the rotor (when the flow gets on board the rotor) has to be done to accelerate the coolant to rotor speed: pressure loss and also higher relative total temperature on board the rotor. Hence, a pre-swirl cavity inner leakage flow is deliberately made to flow in upstream engine direction and bypassed around the pre swirl cavity, to deter lower swirl leakage air from upstream shaft cover location to mix with vane exit air, and reduce the effective swirl. The leakage bypass holes may be angled in a direction of rotation. Pre-swirl cavity inner leakage: leakage flow needs to flow upstream and be bypassed around the pre-swirl cavity, to avoid flow contamination (high swirl vane exit flow reduced by low swirl incoming leakage).

[0008] Therefore, there is a need of a better cooling fluid structure in a gas turbine engine.

SUMMARY

[0009] Briefly described, aspects of the present invention relate to a cooling fluid structure in a gas turbine engine that provides cooling fluid from a turbine rim cavity via a plurality of exit holes near a gas path. In a gas turbine pre-swirl system and a gas turbine blade cooling system, a plurality of exit holes are placed at circumferential positions of gas path pressure peaks to form a non-axisymmetric purge flow via a rim seal of a gas turbine engine to counter hot gas ingestion. A radial position of the exit holes being close to the gas path enables this targeted approach to counter the hot gas ingestion. The exit holes itself are located as part of a seal assembly. The radial location of the exit holes helps ejecting the flow very close to the gas path which makes it effective to counter the high pressure in the gas path (which causes hot gas ingestion). If instead ejecting the flow is done at a more radially inward location, the ejected flow mixes out rapidly in circumferential direction and by the time it flows close to a rim seal area it loses the benefit of targeted hole positions to counter hot gas ingestion. Also the hole arrangement has a direct impact protecting against a 3D pressure profile in the gas path. Thus there is the 3D benefit from strategically placing the cooling holes to counter hot gas ingestion. The hot gas ingestion is caused by 3D pressure profile in the gas path.

[0010] In accordance with one illustrative embodiment of the present invention, a gas turbine engine comprises a first passage, a second passage and a bypass channel. The first passage is formed by a bypass channel chamber and a pre-swirl supporting structure to transmit fluid as a first leakage flow from the bypass channel chamber to the pre-swirl supporting structure and through an inner seal. The second passage is formed by a pre swirl cavity and the inner seal to transmit fluid as a second leakage flow from the pre- swirl cavity and through the inner seal. The bypass channel is formed by an inner wall of the gas turbine engine and a diaphragm and having a rim seal cavity at an end of the bypass channel. The rim seal cavity has a plurality of ejection holes to eject a flow back to a gas path. A combined leakage flow of the first leakage flow and the second leakage flow is bypassed through the bypass channel to the rim seal cavity to purge the rim seal cavity with the combined leakage flow through the plurality of ejection holes.

[0011] In accordance with another illustrative embodiment of the present invention, a gas turbine engine comprises a rim seal cavity having a plurality of ejection holes to eject a flow back to a gas path. A combined leakage flow of a first leakage flow and a second leakage flow is bypassed through a bypass channel to the rim seal cavity to purge the rim seal cavity with the combined leakage flow through the plurality of ejection holes.

BRIEF DESCRIPTION OF THE DRAWINGS

[0012] FIG. 1 illustrates a diagrammatic view of a gas turbine engine in accordance with an exemplary embodiment of the present invention. [0013] FIG. 2 illustrates a close up view of a pre-swirl cavity shown in FIG. 1 in accordance with an exemplary embodiment of the present invention.

[0014] FIG. 3 illustrates a pre- swirl inner leakage bypass system and reintroduction in accordance with an exemplary embodiment of the present invention.

[0015] FIG. 4 illustrates a pre- swirl inner leakage bypass system and reintroduction in accordance with an alternate embodiment of the present invention. [0016] FIG. 5 illustrates a view looking radially inwards to show circumferential orientation of a hole in accordance with an exemplary embodiment of the present invention.

[0017] FIG. 6 illustrates a cut-away view of a gas turbine engine in accordance with an exemplary embodiment of the present invention.

[0018] FIG. 7 illustrates a vane exit circumferential pressure asymmetry in accordance with an exemplary embodiment of the present invention.

[0019] FIG. 8 illustrates a flow ingress/egress at a turbine rim seal gap over one vane pitch in accordance with an exemplary embodiment of the present invention.

DETAILED DESCRIPTION

[0020] To facilitate an understanding of embodiments, principles, and features of the present invention, they are explained hereinafter with reference to implementation in illustrative embodiments. In particular, they are described in the context of a pre-swirl inner leakage bypass system and reintroduction. Circumferential orientation of a plurality of exit holes is provided. A better cooling fluid structure in a gas turbine engine is provided. In a gas turbine pre-swirl system and a gas turbine blade cooling system, a non-axisymmetric purge flow via a rim seal of a gas turbine engine counters hot gas ingestion. A pre-swirl leakage flow is re-used to protect against hot gas ingestion from the gas path, combining an inner seal leakage and a pre-swirl cavity inner leakage. The combined leakage flow is subsequently ejected back to the gas path, using the leakage flow ejection or exit holes located close to the gas path. The leakage flow ejection or exit holes have a preferred circumferential position relative to a vane to protect against high pressure areas in the gas path caused by the vane wakes (which drive hot gas ingestion). The leakage flow ejection or exit holes can also have a circumferential orientation to eject a flow with swirl in a direction of rotation (reduce aerodynamic mixing loss in the gas path). A leakage flow is re-used to purge a rim cavity and protect from over-temperature. The combined leakage flow (pre-swirl cavity inner leakage flow + inner seal leakage flow) is bypassed around pre-swirl nozzles and go radially outwards. Embodiments of the present invention, however, are not limited to use in the described devices or methods.

[0021] The components and materials described hereinafter as making up the various embodiments are intended to be illustrative and not restrictive. Many suitable components and materials that would perform the same or a similar function as the materials described herein are intended to be embraced within the scope of embodiments of the present invention.

[0022] Consistent with one embodiment of the present invention, FIG. 1 represents a representation of a diagrammatic view of a gas turbine engine 5 in accordance with an exemplary embodiment of the present invention. The gas turbine engine 5 comprises a first passage 7(1) formed by a bypass channel chamber 10 or a drive cone cavity and a pre-swirl supporting structure 12 to transmit fluid as a first leakage flow 15(1) from the bypass channel chamber 10 to the pre-swirl supporting structure 12 and through an inner seal 17. The gas turbine engine 5 further comprises a second passage 7(2) formed by a pre-swirl cavity 20 and the inner seal 17 to transmit fluid as a second leakage flow 15(2) from the pre-swirl cavity 20 and through the inner seal 17. The inner seal 17 is located in the pre-swirl cavity 20. The gas turbine engine 5 further comprises a bypass channel 7(3) formed by an inner wall 22 of the gas turbine engine 5 and a diaphragm 25 and having a rim seal cavity 27 at an end of the bypass channel 7(3). The pre-swirl cavity 20 is connected to the bypass channel 7(3). The bypass channel 7(3) is adapted to transmit fluid from the pre-swirl cavity 20 and the first passage 7(1). [0023] The first leakage flow 15(1) and the second leakage flow 15(2) are typically formed due to a leakage in an interface between different components since there is some physical gap that allows flow through. For example, a leakage can occur at an interface between non-rotating and non-rotating components and also between rotating and non- rotating components. In this case, both the first leakage flow 15(1) and the second leakage flow 15(2) are resulting from an interface between a rotating and a non-rotating component.

[0024] The rim seal cavity 27 has a plurality of ejection holes 30(1 -n) to eject a flow 32 back to a gas path 35. The plurality of ejection holes 30(1 -n) are located near the gas path 35. The plurality of ejection holes 30(1 -n) have a preferred circumferential position relative to a vane to protect against high pressure areas in the gas path 35 caused by the vane wakes (which drive hot gas ingestion). The plurality of ejection holes 30(1 -n) have a circumferential orientation to eject a flow with swirl in a direction of rotation (reduce aerodynamic mixing loss).

[0025] A combined leakage flow 15(3) of the first leakage flow 15(1) and the second leakage flow 15(2) is bypassed through the bypass channel 7(3) to the rim seal cavity 27 to purge the rim seal cavity 27 with the combined leakage flow 15(3) through the plurality of ejection holes 30(1 -n). The combined leakage flow 15(3) is re-used to protect against hot gas ingestion from the gas path 35. The combined leakage flow 15(3) is subsequently ejected back to the gas path 35, using the plurality of ejection holes 30(1 -n) located close to the gas path 35. A radial position of the plurality of ejection holes 30(1 - n) being close to the gas path 35 enables this targeted approach to counter the hot gas ingestion. The plurality of ejection holes 30(1 -n) itself are located as part of a seal assembly. The radial location of the plurality of ejection holes 30(1 -n) helps ejecting the combined leakage flow 15(3) to purge a rim seal 50 very close to the gas path 35 which makes it effective to counter the high pressure peaks in the gas path 35 (which causes hot gas ingestion). The combined leakage flow 15(3) is a non-axisymmetric purge flow via a rim seal of a gas turbine engine to counter hot gas ingestion. Non-axisymmetric relates to being non-uniform in a circumferential direction, i.e. it can vary circumferentially.

[0026] The gas turbine engine 5 further comprises a turbine disk 37 to support rotating turbine blades. The gas turbine engine 5 further comprises pre-swirl nozzles 40 configured to accelerate a cooling flow to match a rotational velocity of the turbine disk 37. The combined leakage flow 15(3) is bypassed around the pre-swirl nozzles 40. The gas turbine engine 5 further comprises a cooling flow supply 42 sourced from a compressor exit and used to cool the rotating turbine blades such that the cooling flow passes through the pre-swirl nozzles 40. The gas turbine engine 5 further comprises an outer seal 45 on a turbine disc outer seal arm located between the inner wall 22 of the gas turbine engine 5 and a rotating arm. The outer seal 45 is located adjacent to a rim seal being the outer discourager formed by vane/blade extensions. The outer seal 45 helps seal the bypass channel chamber 10 and allows dropping the pressure in the bypass channel chamber 10 to reverse a flow through the inner seal 17. A non-axisymmetrie purge flow via the seal 45 of the gas turbine engine 5 counters hot gas ingestion. The combined leakage flow 15(3) is re-used to purge the rim seal cavity 27 and protect from over-temperature. The combined leakage flow 15(3) (pre-swirl cavity inner leakage flow + inner seal leakage flow) is bypassed around the pre-swirl nozzles 40 and go radially outwards.

[0027] The gas turbine engine 5 may include a compressor section 51 for compressing air. The compressed air from the compressor section 51 is conveyed to a combustion section 52, which produces hot combustion gases by burning fuel in the presence of the compressed air from the compressor section 51. The combustion gases are conveyed through a plurality of transition ducts to a turbine section 54 of the engine 5. The turbine section 54 comprises alternating rows of rotating blades and stationary vanes 60.

[0028] In the turbine section 54, blade disc structures 55 are positioned adjacent to one another in an axial direction. The blade disc structures 55 define a rotor. Each of the blade disc structures 55 supports circumferentially spaced apart blades and each of a plurality of lower stator support structures support a plurality of circumferentially spaced apart vanes. The vanes direct the combustion gases from the transition ducts along a hot gas flow path to the blades such that the combustion gases cause rotation of the blades, which in turn causes corresponding rotation of the rotor.

[0029] A supply of fluid, can supply fluid within the gas turbine engine 5. The fluid may have a temperature of, for example, between about 1000-1200° F (this is referring to a flow temperature in the bypass channel chamber 10).

[0030] The fluid flows through passages formed between the inner wall 22 and an outer combustor wall. The fluid also flows between a rotating arm 56 and the inner wall 22. The rotating arm 56 may be connected to the compressor section 51 and the turbine section 54. The rotating arm 56 may also be referred to as a“drive cone” or“shaft.”

[0031] Referring to FIG. 2, it illustrates a close up view of the pre-swirl cavity 20 shown in FIG. 1 in accordance with an exemplary embodiment of the present invention. The pre-swirl cavity 20 may define a flow passage in which a pre-swirl structure exists. The pre-swirl structure may have swirl members 204. The swirl members 204 may include a leading edge and a circumferentially offset trailing edge. The pre-swirl members 204 turn the cooling fluid. Cooling fluid exits the pre-swirl cavity 20 with a velocity component in the circumferential direction of the gas turbine engine 5. A swirl ratio is defined as the velocity component in the circumferential direction of the cooling fluid as compared to a circumferential velocity component of a rotating shaft.

[0032] Leakage flow moves from the pre-swirl cavity 20 into the bypass channel chamber 10 that then enters the bypass channel 7(3). The combined leakage flow 15(3) (bypass channel flow) moves through the bypass channel 7(3). Prior to entering the bypass channel 7(3) the bypass channel flow may pass through the inner seal 17. The bypass channel 7(3) allows evacuating the mix of air from a passage and the inner seal 17, without contaminating the pre-swirl cavity 20. The combined leakage flow 15(3) (bypass channel flow) is dumped in the rim seal cavity 27 and it contributes to purging and cooling of the rim seal cavity 27.

[0033] Referring to FIG. 3, it illustrates a pre-swirl inner leakage bypass system 300 that enables reintroduction of a pre- swirl leakage flow in a gas path in accordance with an exemplary embodiment of the present invention. The pre-swirl imier leakage bypass system 300 comprises a cooling flow supply 302 (sourced from a compressor exit, used to cool rotating turbine blades). The pre- swirl inner leakage bypass system 300 further comprises pre-swirl nozzles 305 used to accelerate a cooling flow to match a rotational velocity of a turbine disk 307. The turbine disk 307 supports the rotating turbine blades. The pre-swirl inner leakage bypass system 300 further comprises a pre-swirl nozzles support structure 310 and a honeycomb 312 (as part of a labyrinth seal). The pre-swirl inner leakage bypass system 300 further comprises a plurality of purge flow ejection holes 315(l -m). The pre-swirl inner leakage bypass system 300 further comprises a cooling hole 320 to channel a cooling flow 322 to the rotating turbine blades.

[0034] The pre-swirl inner leakage bypass system 300 further comprises a pre swirl cavity 325 and an inner seal 330. An inner seal leakage flow is possible. A drive cone cavity leakage f1 ow 335 and a pre- swirl cavity inner leakage flow 340 provides a combined leakage flow 345 (pre- swirl cavity inner leakage flow + inner seal leakage flow) that is bypassed around the pre-swirl nozzles 305 and it goes radially outwards. The combined leakage flow 345 is re-used to purge a rim cavity 350 and protect from over- temperature. The plurality of purge flow ejection holes 315(l-m) are used to purge it. [0035] FIG. 4 illustrates a pre-swirl inner leakage bypass system 400 that enables reintroduction of a pre- swirl leakage flow in a gas path in accordance with an exemplary embodiment of the present invention. A drive cone cavity leakage flow 405 and a pre-swirl cavity inner leakage flow 410 provides a combined leakage flow 415 (pre-swirl cavity inner leakage flow + inner seal leakage flow ) that is bypassed around the pre- swirl nozzles 305 and it goes radially outwards. Here a first passage is instead formed by only the bypass channel chamber 10 to transmit fluid as a first leakage flow (the drive cone cavity leakage flow 405) directly from the bypass channel chamber 10 to the bypass channel 7(3).

[0036] As seen in FIG. 5, it illustrates a view looking radially inwards to show a circumferential orientation 500 of an exit or ejection hole 505(1) of a plurality of exit or ejection holes 505(1 -3) in accordance with an exemplary embodiment of the present invention. The plurality of exit or ejection holes 505(1-3) have the circumferential orientation 500 to eject a flow with swirl in a direction of rotation. Each exit or ejection hole 505 of the plurality of exit or ejection holes 505(1 -3) is aligned with a corresponding exit angle 510 of an associated vane 515. A plurality of blades 517(1-4) are shown.

[0037] The plurality of exit or ejection holes 505(1-3) have a preferred circumferential position 520 relative to the vane 515 to protect against high pressure areas in the gas path 35 (see FIG. 1) caused by vane wakes. One exit or ejection hole 505 of the plurality of exit or ejection holes 505(1-3) is provided per vane 515 and each exit or ejection hole 505 of the plurality of exit or ejection holes 505(1 -3) is longitudinally aligned with a corresponding trailing edge 525 of the associated vane 515.

[0038] As shown in FIG. 6, it illustrates a cut-away view of a gas turbine engine 600 in accordance with an exemplary embodiment of the present invention. The gas turbine engine 600 comprises a plurality of nozzle guide vanes 605(1-3) and a plurality of exit or ejection holes 610(1-3) to eject a flow such as the combined leakage flow 15(3). There is an opportunity to be deliberate with a pre-swirl inner cavity leakage that is bypassed around the pre-swirl cavity 20 (FIG. 2). This leakage flow reintroduction can be done to form a non-axisymmetric purge flow⁷ to counter hot gas ingestion, offering increased protection. The plurality of exit or ejection holes 610(1-3) for a bypass flow is relatively close to the gas path 35 so the flow re-introduction can serve to offer additional (targeted) protection against hot gas ingestion.

[0039] In FIG. 7, it illustrates a vane exit circumferential pressure asymmetry 700 in accordance with an exemplary embodiment of the present invention. Gas path-driven hot gas ingestion is driven by a mainstream flow circumferentially-varying static pressure distribution For example, the circumferential pressure distribution on a vane 705 is shown in FIG. 7, where there are pressure peaks 710 at trailing edge locations 715.

[0040] With regard to FIG. 8, it illustrates a flow ingress/egress at a turbine rim seal gap over one vane pitch in accordance with an exemplary embodiment of the present invention. These local pressure peaks drive ingestion into the rim cavity 350 is shown. The gas path pressure distribution due to the vane 705 is shown by a curve 800. The rim cavity pressure level is a horizontal line 805. At circumferential positions where local gas path pressure is higher than the rim cavity pressure, ingestion occurs (first-shaded area 810 on the graph). Conversely, purge outflow happens at circumferential positions where rim cavity pressure is higher than the gas path 35 (second-shaded area 815). Hence, the proposed locations of the plurality of exit or ejection holes 610(1-3) for the pre-swirl bypass flow such as the combined leakage flow 15(3) are to align with the gas path pressure peaks 710.

[0041] This gas turbine engine 5 deliberately re-uses an existing leakage flow for the purpose of downstream benefit of improved hot gas ingestion protection. It aligns the plurality of exit or ejection holes 610(1-3) for the pre-swirl bypass flow such as the combined leakage flow 15(3) to the circumferential positions of the gas path pressure peak, in particular the pressure peaks formed by an upstream vane row. Better protection against hot gas ingestion benefits turbine rotor mechanical integrity. [0042] While a pre-swirl inner seal leakage bypass flow is described here a range of other constructions of pre-swirl inner seal leakage bypass systems are also contemplated by the present invention. For example, other types of turbine pre-swirl systems may be implemented based on one or more features presented above without deviating from the spirit of the present invention. [0043] The techniques described herein can be particularly useful for inner seal leakage bypass flow exit holes placed at circu ferential positions of gas path pressure peaks. While particular embodiments are described in terms of the main inner seal leakage bypass, the techniques described herein are not limited to the inner seal leakage bypass but can also be used for other leakage bypass. [0044] While embodiments of the present invention have been disclosed in exemplary forms, it will be apparent to those skilled in the art that many modifications, additions, and deletions can be made therein without departing from the spirit and scope of the invention and its equivalents, as set forth in the following claims.

[0045] Embodiments and the various features and advantageous details thereof are explained more fully with reference to the non-limiting embodiments that are illustrated in the accompanying drawings and detailed in the following description. Descriptions of well-known starting materials, processing techniques, components and equipment are omitted so as not to unnecessarily obscure embodiments in detail. It should be understood, however, that the detailed description and the specific examples, while indicating preferred embodiments, are given by way of illustration only and not by way of limitation. Various substitutions, modifications, additions and/or rearrangements within the spirit and/or scope of the underlying inventive concept will become apparent to those skilled in the art from this disclosure.

[0046] As used herein, the terms“comprises,”“comprising,”“includes,”“including,” “has,”“having” or any other variation thereof, are intended to cover a non-exclusive inclusion. For example, a process, article, or apparatus that comprises a list of elements is not necessarily limited to only those elements but may include other elements not expressly listed or inherent to such process, article, or apparatus.

[0047] Additionally, any examples or illustrations given herein are not to be regarded in any way as restrictions on, limits to, or express definitions of, any term or terms with which they are utilized. Instead, these examples or illustrations are to be regarded as being described with respect to one particular embodiment and as illustrative only. Those of ordinary skill in the art will appreciate that any term or terms with which these examples or illustrations are utilized will encompass other embodiments which may or may not be given therewith or elsewhere in the specification and all such embodiments are intended to be included within the scope of that term or terms.

[0048] In the foregoing specification, the invention has been described with reference to specific embodiments. However, one of ordinary skill in the art appreciates that various modifications and changes can be made without departing from the scope of the invention. Accordingly, the specification and figures are to be regarded in an illustrative rather than a restrictive sense, and all such modifications are intended to be included within the scope of invention.

[0049] Although the invention has been described with respect to specific embodiments thereof, these embodiments are merely illustrative, and not restrictive of the invention. The description herein of illustrated embodiments of the invention is not intended to be exhaustive or to limit the invention to the precise forms disclosed herein (and in particular, the inclusion of any particular embodiment, feature or function is not intended to limit the scope of the invention to such embodiment, feature or function). Rather, the description is intended to describe illustrative embodiments, features and functions in order to provide a person of ordinary skill in the art context to understand the invention without limiting the invention to any particularly described embodiment, feature or function. While specific embodiments of, and examples for, the invention are described herein for illustrative purposes only, various equivalent modifications are possible within the spirit and scope of the invention, as those skilled in the relevant art will recognize and appreciate. As indicated, these modifications may be made to the invention in light of the foregoing description of illustrated embodiments of the invention and are to be included within the spirit and scope of the invention. Thus, while the invention has been described herein with reference to particular embodiments thereof, a latitude of modification, various changes and substitutions are intended in the foregoing disclosures, and it will be appreciated that in some instances some features of embodiments of the invention will be employed without a corresponding use of other features without departing from the scope and spirit of the invention as set forth. Therefore, many modifications may be made to adapt a particular situation or material to the essential scope and spirit of the invention.

[0050] Respective appearances of the phrases "in one embodiment," "in an embodiment," or "in a specific embodiment" or similar terminology in various places throughout this specification are not necessarily referring to the same embodiment. Furthermore, the particular features, structures, or characteristics of any particular embodiment may be combined in any suitable manner with one or more other embodiments. It is to be understood that other variations and modifications of the embodiments described and illustrated herein are possible in light of the teachings herein and are to be considered as part of the spirit and scope of the invention. [0051] In the description herein, numerous specific details are provided, such as examples of components and/or methods, to provide a thorough understanding of embodiments of the invention. One skilled in the relevant art will recognize, however, that an embodiment may be able to be practiced without one or more of the specific details, or with other apparatus, systems, assemblies, methods, components, materials, parts, and/or the like. In other instances, well-known structures, components, systems, materials, or operations are not specifically shown or described in detail to avoid obscuring aspects of embodiments of the invention. While the invention may be illustrated by using a particular embodiment, this is not and does not limit the invention to any particular embodiment and a person of ordinary skill in the art will recognize that additional embodiments are readily understandable and are a part of this invention.

[0052] It will also be appreciated that one or more of the elements depicted in the drawings/figures can also be implemented in a more separated or integrated manner, or even removed or rendered as inoperable in certain cases, as is useful in accordance with a particular application. [0053] Benefits, other advantages, and solutions to problems have been described above with regard to specific embodiments. However, the benefits, advantages, solutions to problems, and any component(s) that may cause any benefit, advantage, or solution to occur or become more pronounced are not to be construed as a critical, required, or essential feature or component.

Claims

What is claimed is:

Claim 1. A gas turbine engine (5) comprising:

a first passage (7(1)) formed by a bypass channel chamber (10) and a pre swirl supporting structure (12) to transmit fluid as a first leakage flow (15(1)) from the bypass channel chamber (10) to the pre-swirl supporting structure (12) and through an inner seal (17);

a second passage (7(2)) formed by a pre-swirl cavity (20) and the inner seal (17) to transmit fluid as a second leakage flow (15(2)) from the pre-swirl cavity (20) and through the inner seal (17); and

a bypass channel (7(3)) formed by an inner wall (22) of the gas turbine engine (5) and a diaphragm (25) and having a rim seal cavity (27) at an end of the bypass channel (7(3)),

wherein the rim seal cavity (27) having a plurality of ejection holes (30(1 - n)) to eject a flow (32) back to a gas path (35),

wherein a combined leakage flow (15(3)) of the first leakage flow (15(1)) and the second leakage flow (15(2)) is bypassed through the bypass channel (7(3)) to the rim seal cavity (27) to purge the rim seal cavity (27) with the combined leakage flow (15(3)) through the plurality of ejection holes (30(1 -n)).

Claim 2. The gas turbine engine (5) of claim 1, wherein the plurality of ejection holes (30(l-n)) are located near the gas path (35).

Claim 3. The gas turbine engine (5) of claim 1 or 2, wherein the combined leakage flow is re-used to protect against hot gas ingestion from the gas path.

Claim 4. The gas turbine engine (5) of claim 1 or 2 or 3, wherein the plurality of ejection holes (505(1-3)) have a preferred circumferential position (520) relative to a vane (515) to protect against high pressure areas in the gas path (35) caused by vane wakes.

Claim 5. The gas turbine engine (5) of claim 4, wherein one ejection hole (505) of the plurality of ejection holes (505(1 -3)) is provided per vane 515 and each ejection hole (505) of the plurality of ejection holes (505(1-3)) is longitudinally aligned with a corresponding trailing edge (525) of an associated vane (515).

Claim 6. The gas turbine engine (5) of claim 1, wherein the plurality of ejection holes (505(1-3)) have a circumferential orientation (500) to eject a flow with swirl in a direction of rotation.

Claim 7. The gas turbine engine (5) of claim 6, wherein each ejection hole (505) of the plurality of ejection holes (505(1-3)) is aligned with a corresponding exit angle 510 of an associated vane 515.

Claim 8. The gas turbine engine (5) of claim 1, further comprising:

a turbine disk (37) to support rotating turbine blades; and

pre-swirl nozzles (40) configured to accelerate a cooling flow (322) to match a rotational velocity of the turbine disk (37), wherein the combined leakage flow (15(3)) is bypassed around the pre-swirl nozzles (40).

Claim 9. The gas turbine engine (5) of claim 8, further comprising:

a cooling flow supply (42) sourced from a compressor exit and used to cool the rotating turbine blades such that the cooling flow (322) passes through the pre-swirl nozzles (40).

Claim 10. The gas turbine engine (5) of claim 1, wherein the first passage (7(1)) is instead formed by only the bypass channel chamber (10) to transmit fluid as the first leakage flow (15(1)) directly from the bypass channel chamber (10) to the bypass channel (7(3)).

Claim 11. The gas turbine engine (5) of claim 1 , further comprising:

an outer seal (45) on a turbine disc outer seal arm located between the inner wall (22) of the gas turbine engine (5) and a rotating arm, wherein the outer seal (45) helps seal the bypass channel chamber (10) and allows dropping the pressure in the bypass channel chamber (10) to reverse a flow through the inner seal (17).

Claim 12. The gas turbine engine (5) of claim 1 , wherein the pre-swirl cavity (20) is connected to the bypass channel (7(3)).

Claim 13. The gas turbine engine (5) of claim 1, wherein the bypass channel (7(3)) is adapted to transmit fluid from the pre-swirl cavity (20) and the first passage (7(1)).

Claim 14. The gas turbine engine (5) of claim 1, wherein the inner seal (17) is located in the pre-swirl cavity (20).

Claim 15. A gas turbine engine (5) comprising:

a rim seal cavity (27) having a plurality of ejection holes (30(l-n)) to eject a flow (32) back to a gas path (35),

wherein a combined leakage flow (15(3)) of a first leakage flow (15(1)) and a second leakage flow (15(2)) is bypassed through a bypass channel (7(3)) to the rim seal cavity (27) to purge the rim seal cavity (27) with the combined leakage flow (15(3)) through the plurality of ejection holes (30(1 -n)).

Claim 16. The gas turbine engine (5) of claim 15, further comprising:

a first passage (7(1)) formed by a bypass channel chamber (10) and a pre-swirl supporting structure (12) to transmit fluid as the first leakage flow (15(1)) from the bypass channel chamber (10) to the pre-swirl supporting structure (12) and through an inner seal (17).

Claim 17. The gas turbine engine (5) of claim 16, further comprising:

a second passage (7(2)) formed by a pre-swirl cavity (20) and the inner seal (17) to transmit fluid as the second leakage flow (15(2)) from the pre-swirl cavity (20) and through the inner seal (17).

Claim 18. The gas turbine engine (5) of claim 17, wherein the bypass channel (7(3)) is formed by an inner wall (22) of the gas turbine engine (5) and a diaphragm (25) and having the rim seal cavity (27) at an end of the bypass channel (7(3)).

Claim 19. The gas turbine engine (5) of claim 15, wherein the plurality of ejection holes (30(l-n)) have a preferred circumferential position (520) relative to a vane (515) to protect against high pressure areas in the gas path (35) caused by vane wakes.

Claim 20. The gas turbine engine (5) of claim 15, wherein the plurality of ejection holes (30(l-n)) have a circumferential orientation (500) to eject a flow with swirl in a direction of rotation.