GB2559804A

GB2559804A - Heatshield for a gas turbine

Info

Publication number: GB2559804A
Application number: GB1702759.0A
Authority: GB
Inventors: Walker Craig
Original assignee: Siemens AG
Current assignee: Siemens AG
Priority date: 2017-02-21
Filing date: 2017-02-21
Publication date: 2018-08-22
Also published as: GB201702759D0

Abstract

A turbine section for a gas turbine engine has a casing assembly comprising an annular array of circumferentially segmented heatshields 64a,b. At least two circumferentially adjacent heatshields 64a,b have opposing surfaces 98a,b defining a vent slot 104 therebetween, which is open to the working gas path. At least one of the heatshields 64a comprises at least one cooling hole 110, preferably extending from a cooled side (106, fig 3) of the heatshield, with an outlet 112 into the vent slot 104. The vent slot 104 is at least partly formed by a cut-out 105 in only one of the opposing surfaces 98a. The vent slot 104 may extend from a leading edge surface (96, fig 3) towards the trailing edge surface (94, fig 3). This configuration may maintain a path for cooling fluid when the intersegment gap 103 is closed due to transient movements or tolerance stack-up.

Description

(54) Title of the Invention: Heatshield for a gas turbine

Abstract Title: Vent slot for gas turbine heatshield segments (57) A turbine section for a gas turbine engine has a casing assembly comprising an annular array of circumferentially segmented heatshields 64a,b. At least two circumferentially adjacent heatshields 64a,b have opposing surfaces 98a,b defining a vent slot 104 therebetween, which is open to the working gas path. At least one of the heatshields 64a comprises at least one cooling hole 110, preferably extending from a cooled side (106, fig 3) of the heatshield, with an outlet 112 into the vent slot 104. The vent slot 104 is at least partly formed by a cut-out 105 in only one of the opposing surfaces 98a. The vent slot 104 may extend from a leading edge surface (96, fig 3) towards the trailing edge surface (94, fig 3). This configuration may maintain a path for cooling fluid when the intersegment gap 103 is closed due to transient movements or tolerance stack-up.

At least one drawing originally filed was informal and the print reproduced here is taken from a later filed formal copy.

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FIG 4

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HEATSHIELD FOR A GAS TURBINE

FIELD OF INVENTION

The present invention relates to circumferentially segmented heatshields that form part of the working gas flow path and in particular a cooling or venting arrangement of the heatshields.

BACKGROUND OF INVENTION

In a gas turbine engine the hot working gases from the combustor are channeled into the turbine section. The turbine section comprises at least a high pressure turbine which drives the high pressure compressor. The high pressure turbine I compressor can be referred to as the core turbine I compressor or the turbine can be referred to as the compressor-turbine. The turbine section can further comprise a low pressure turbine which drives a low pressure compressor, a fan or an air compressor or even an electrical generator. In addition, the turbine section can comprise an intermediate turbine, arranged between the high and low pressure turbines and which usually drives an intermediate compressor.

The high pressure or core turbine comprises a nozzle guide vane stage and at least one rotor stage. Radially inner and outer walls further define the working gas path. At least a part of the walls comprises an annular array of circumferential segments of heatshields. The heatshields are exposed to the hot working gases and require cooling with a flow of cooling air usually supplied from a compressor.

Between ambient and maximum operating temperatures, circumferentially adjacent heatshields can contact one another or be separated by a gap. To prevent ingress of hot gases into the gap and to provide a flow path for cooling air that cools the heatshields, the cooling air is vented into the gap. To ensure cooling air flow is maintained vent slots are formed by cut-outs in both circumferentially adjacent heatshields. The cut-outs are formed in the radially inner part of the adjacent heatshields’ opposing surfaces. However, the width of the vent slot is sized to accommodate manufacturing tolerances for both cut-outs and in order to prevent the ingress of hot gases the quantity of cooling air is substantial.

EP2138676 B1 discloses a method for cooling a component, such as a casing, of a turbine is provided, wherein a fluid with a pressure below 1 bar is guided away from the component. Moreover, a turbine is described comprising a component, a conduit which is connected to the component such that a fluid can be guided away from the component, and a fluid discharge which is connected to the conduit. The fluid discharge is constructed such that it removes a fluid with a pressure below 1 bar.

EP2702251A1 discloses a turbine including an inner casing to which at least a stator vane of a turbine section is mountable, and an outer casing arranged around the inner casing in such a way that an outer cooling channel is formed between the inner casing and the outer casing. The outer cooling channel includes a fluid inlet through which a cooling fluid is injectable from an outer volume of the turbine into the outer cooling channel. The cooling channel includes a fluid outlet such that the cooling fluid is exhausted into an inner volume of the turbine. The fluid inlet is located with respect to the fluid outlet such that the cooling fluid inside the outer cooling channel includes a flow direction which has a component that is orientated in opposite direction with respect to a main flow direction of a working fluid of the turbine.

STATEMENT OF INVENTION

To address the problem of the vent slot closing or reducing to an unacceptable width there is provided a turbine section for a gas turbine engine having an axis, the turbine section comprising at least one rotor stage and a casing assembly at least partly defining a working gas path. The casing assembly comprising an annular array of circumferentially segmented heatshields. At least two circumferentially adjacent heatshields have opposing surfaces which define a vent slot therebetween. The vent slot is open to the gas path. At least one of the heatshields comprising at least one cooling hole having an outlet into the vent slot. The vent slot is at least partly formed by a cut-out in only one of the opposing surfaces.

The heatshield may have a leading edge surface and a trailing edge surface and the vent slot extends between the leading edge surface and the trailing edge surface.

The vent slot may extend from the leading edge surface.

The heatshield may have a cooled side and a hot side and cooling holes are formed in the heatshield and extend from the cooled side to the vent slot.

The cooling holes may have an outlet, the outlet is in any one or both the opposing surfaces.

The cut-out may have a nominal width and a manufacturing tolerance, the actual cut-out width is within at least 12.5% of the nominal width and preferably at least 25% of the nominal width.

The vent slots exist to ensure that even if the circumferential gap between wedge faces closed to zero through transient movements, tolerance stack-ups or incorrect design assumptions a path for discharged coolant will always exist preventing overheating of components which would otherwise culminate in limitations of component life.

BRIEF DESCRIPTION OF THE DRAWINGS

The above mentioned attributes and other features and advantages of this invention and the manner of attaining them will become more apparent and the invention itself will be better understood by reference to the following description of embodiments of the invention taken in conjunction with the accompanying drawings, wherein

FIG. 1 shows part of a turbine engine in a sectional view and in which incorporates the presently described casing assembly having an annular array of circumferentially segmented heatshields,

FIG. 2 shows a close-up and part section of a turbine section having the casing assembly which comprises the annular array of circumferentially segmented heatshields,

FIG. 3 shows a perspective view of part of a heatshield, one of the array of circumferentially segmented heatshields, and

FIG. 4 shows a generally axial view on circumferentially adjacent heatshields of the array of circumferentially segmented heatshields.

DETAILED DESCRIPTION OF INVENTION

FIG. 1 shows an example of a gas turbine engine 10 in a sectional view. The gas turbine engine 10 comprises, in flow series, an inlet 12, a compressor section 14, a combustor section 16 and a turbine section 18 which are generally arranged in flow series and generally about and in the direction of a longitudinal or rotational axis

20. The gas turbine engine 10 further comprises a shaft 22 which is rotatable about the rotational axis 20 and which extends longitudinally through the gas turbine engine 10. The shaft 22 drivingly connects the turbine section 18 to the compressor section 14.

In operation of the gas turbine engine 10, air 24, which is taken in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or burner section 16. The burner section 16 comprises a burner plenum 26, one or more combustion chambers 28 and at least one burner 30 fixed to each combustion chamber 28. The combustion chambers 28 and the burners 30 are located inside the burner plenum 26. The compressed air passing through the compressor 14 enters a diffuser 32 and is discharged from the diffuser 32 into the burner plenum 26 from where a portion of the air enters the burner 30 and is mixed with a gaseous or liquid fuel. The air/fuel mixture is then burned and the combustion gas 34 or working gas from the combustion is channelled through the combustion chamber 28 to the turbine section 18 via a transition duct 17.

This exemplary gas turbine engine 10 has a cannular combustor section arrangement 16, which is constituted by an annular array of combustor cans 19 each having the burner 30 and the combustion chamber 28, the transition duct 17 has a generally circular inlet that interfaces with the combustor chamber 28 and an outlet in the form of an annular segment. An annular array of transition duct outlets form an annulus for channelling the combustion gases to the turbine 18.

The turbine section 18 comprises a number of blade carrying discs 36 attached to the shaft 22. In the present example, two discs 36 each carry an annular array of turbine blades 38. However, the number of blade carrying discs could be different, i.e. only one disc or more than two discs. In addition, guiding vanes 40, which are fixed to a stator 42 of the gas turbine engine 10, are disposed between the stages of annular arrays of turbine blades 38. Between the exit of the combustion chamber 28 and the leading turbine blades 38 inlet guiding vanes 44 are provided and turn the flow of working gas onto the turbine blades 38. Each annular array of vanes or blades can be referred to as a ‘stage’.

The combustion gas from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38 which in turn rotate the shaft 22. The guiding vanes 40, 44 serve to optimise the angle of the combustion or working gas on the turbine blades 38.

The turbine section 18 drives the compressor section 14. The compressor section 14 comprises an axial series of vane stages 46 and rotor blade stages 48. The rotor blade stages 48 comprise a rotor disc supporting an annular array of blades. The compressor section 14 also comprises a casing 50 that surrounds the rotor stages and supports the vane stages 48. The guide vane stages include an annular array of radially extending vanes that are mounted to the casing 50. The vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point. Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operations conditions.

The casing 50 defines a radially outer surface 52 of the passage 56 of the compressor 14. A radially inner surface 54 of the passage 56 is at least partly defined by a rotor drum 53 of the rotor which is partly defined by the annular array of blades 48.

The present invention is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine. However, it should be appreciated that the present invention is equally applicable to two or three shaft engines having two or three turbines sections and which can be used for industrial, aero or marine applications.

The terms upstream and downstream refer to the flow direction of the airflow and/or working gas flow through the engine unless otherwise stated. The terms forward and rearward refer to the general flow of gas through the engine. The terms axial, radial and circumferential are made with reference to the rotational axis 20 of the engine.

FIG. 2 shows an enlarged view on part of a turbine section 18 having a casing assembly 60 that defines part of a working gas path 62 for a flow of hot working gas 63 from the combustor 16. The turbine section 18 comprises the guide vane stages 40 and the rotor blade stages 38. The casing assembly 60 comprises an outer casing structure 66 and an inner casing structure 68. The outer casing structure 66 engages surrounding engine architecture and secures the inner casing structure 68. The inner casing structure 68 comprises an annular array of circumferentially segmented heatshields denoted by reference numbers 64. The inner casing structure 68 also partly supports the guide vane stages 40. The casing assembly 60 comprises cooling and/or sealing air passages 70, 72, 74 and an annular distribution chamber 76 through which cooling and/or sealing air 78 is channelled.

Generally, the cooling air 78 can be used in a number of ways to cool: the inner casing structure 68 by impingement and convection cooling, the vanes 40 by impingement, convection and effusion or film cooling, and to be channelled across the gas path 62 to cool and/or seal the components in a hub region denoted by 80. These components in the hub region 80 can include the discs 36, where the air cools the disc, and seals 82 where the air not only cools the seals but also provides a positive pressure across the seal to prevent the ingress of hot working gases 63.

Reference is now made to Figures 3 which shows a perspective view of part of a heatshield 64, one of the array of circumferential segments and Figure 4, which shows a generally axial view on circumferentially adjacent heatshields 64 (64a, 64b) of the array of circumferential segments.

Each heatshield 64 comprises a main body 100 that forms a gas washed surface 102, a leading edge surface 94, a trailing edge surface 96 and circumferential or opposing surfaces 98 only one of which is shown. The leading edge surface 94 is upstream of the trailing edge surface 96 with respect to the general flow direction of the working gas 63. Each heatshield 64 is located and supported via forward and rearward hooks 90, 92 which engage with the inner casing structure 68; further securing means can be used to secure the heatshield to the inner casing structure 68. In the turbine section the rotor blades 38 rotate radially inwardly of the heatshields 64 and a minimum gap exists between the gas washed surface 102 and the tips of the blades 38 to minimise over-tip leakage. The radially outer side of the heatshield 64 can be referred to as the cold side 106, by virtue of cooling air flowing over that surface, and the radially inner side or gas washed side can be referred to as the hot side by virtue of the hot working gases flowing over this surface 102.

Installed as an annular array of heatshields 64, at least two circumferentially adjacent heatshields, designated 64a and 64b in Figure 4, have their respective and opposing surfaces 98a and 98b defining a gap 103 therebetween. This gap 103 is provided to allow thermal expansion of the heatshields 64 in the circumferential direction. In operation of the gas turbine 10 and at maximum temperature this gap 103 can completely close such that the opposing surfaces 98a, 98b can be in contact. The annular array of heatshields 64 is designed such that a nominal and relatively small gap 103 is intentional throughout the operating window of the gas turbine; however, it remains a possibility that the gap 103 can completely close and adjacent heatshields can contact one another. When not in operation and the heatshields 64 are cold, the gap 103 or the spacing between immediately adjacent heatshields 64 can be at a maximum. Transient engine conditions can also cause maximum and minimum gap widths.

A vent slot 104 is partly formed by a cut-out 105 in only one of the opposing surfaces 98, in this case in heatshield 64a. The vent slot 104 is further formed by the flat and opposing surface 98b of the immediately adjacent heatshield 64b. The vent slot 104 is open to the gas path 62. Thus the vent slot 104 itself can include the gap 103 or at least the nominal spacing between opposing surfaces 98a, 98b at any point in the gas turbine’s operating envelope. Thus the vent slot width 114 is the cutout width 116 and the gap width 118, where the gap width can be zero and up to (and including) its maximum width.

The vent slot 104 is shown to extend in the axial direction from the leading edge surface 94 towards the trailing edge surface 96 and end before the trailing edge surface 96. In other embodiments the vent slot 104 may not need to ‘break out’ of the leading edge surface 94 and can extend to break out of the trailing edge surface 96. In some embodiments the vent slot 104 can extend between and not break out of the leading or trailing edge surfaces 94, 96. For any given implementation of the present heatshield 64, the location and extent to the vent slot 104 partly depends on the temperature and pressure characteristics of the working gas, cooling air flows to the heatshield 64 and relative location of the rotor blades 38 to the heatshield 64.

The heatshield 64a has at least one cooling hole 110 having an outlet 112 opening into the vent slot 104. In Figure 3 the heatshield 64 can be seen to have an array of cooling holes 110 spaced along the cut-out 105. The cooling hole or holes 110 generally channel cooling air 120 from the cooled side 106 of the heatshield 64 to the hot side 96. The cooling air 120 or coolant 120 can perform at least three tasks, one task is to provide a route to exhaust coolant passing through the heatshield 64 and another task is to seal the gap 103 to prevent the ingress of hot working gases. The present arrangement of heatshields 64 can assist in providing at least a part of a cooling film over the gas washed surface 102 of the heatshield. The quantity of coolant 120 passing through the cooling passages 110 is carefully designed such that sufficient is provided to perform these tasks, but also minimum amount to minimise losses or inefficiency in using air bled from the compressor. Although this coolant flow is necessary to cool hot components, it is parasitic and detrimental to the efficiency of the gas turbine engine. Thus it should be appreciated that efficient use of coolant and sufficient cooling of components is critical in gas turbine engines and in particular turbine sections.

The improved control of the vent slot width 114 reduces the ingress of hot gases and reduces the heat flux into the heatshield 64. In addition, the variation of vent slot widths around the annular array of heatshields is significantly reduced meaning that there is a more even distribution of coolant to each heatshield 64 and the thermal gradient around the annulus is reduced. This in turn leads to a better control of the over-tip leakage gap between the blades because thermal distortions of the heatshield are better controlled. Furthermore, improved engine efficiency can be realised.

Claims

1. A turbine section (18) for a gas turbine engine having an axis (20), the turbine section (18) comprising at least one rotor stage (48) and a casing assembly (50) at least partly defining a working gas path (62), the casing assembly (50) comprising an annular array of circumferentially segmented heatshields (64), wherein at least two circumferentially adjacent heatshields (64a, 64b) have opposing surfaces (98a, 98b) which define a vent slot (104) therebetween, the vent slot (104) is open to the gas path (62), at least one of the heatshields (64a, 64b) comprising at least one cooling hole (110) having an outlet (112) into the vent slot (64) characterised in that the vent slot (104) is at least partly formed by a cut-out (105) in only one of the opposing surfaces (98a, 98b).

2. A turbine section (18) as claimed in claim 1 wherein the heatshield (64) has a leading edge surface (94) and a trailing edge surface (96) and the vent slot (104) extends between the leading edge surface (94) and the trailing edge surface (96).

3. A turbine section (18) as claimed in claim 2 wherein the vent slot (104) extends from the leading edge surface (94).

4. A turbine section (18) as claimed in any one of claims 1-3 wherein the heatshield (64) has a cooled side (106) and a hot side (102) and cooling holes (110) are formed in the heatshield (64) and extend from the cooled side (106) to the vent slot (104).

5. A turbine section (18) as claimed in claim 4 wherein the cooling holes (110) have an outlet (112), the outlet (112) is in any one or both the opposing surfaces (98a, 98b).

6. A turbine section (18) as claimed in any one of claims 1-5 wherein the cut-out (105) having a nominal width and a manufacturing tolerance, the actual cut-out width (116) is within at least 12.5% of the nominal width and preferably at least 25% of the nominal width.

Intellectual

Property

Office

Application No: GB1702759.0 Examiner: Mr David Kirwin