WO2019168590A1

WO2019168590A1 - Gas turbine engine with turbine cooling air delivery system

Info

Publication number: WO2019168590A1
Application number: PCT/US2019/012452
Authority: WO
Inventors: David May; Constant CHARRETON
Original assignee: Siemens AG; Siemens Corp; Siemens Energy Inc
Current assignee: Siemens AG; Siemens Corp; Siemens Energy Inc
Priority date: 2018-02-27
Filing date: 2019-01-07
Publication date: 2019-09-06
Anticipated expiration: 2020-08-27
Also published as: WO2019168501A1

Abstract

A gas turbine engine includes a first passage 19 formed by an exterior wall 18 of a combustor and an inner wall 17 within the gas turbine engine and a second passage 20 formed by the inner wall 17 and a rotating arm 16. A pre-swirl cavity 30 is located at a downstream end of the first passage 19. A first self-adjusting seal 40 is located in the second passage 20. A second self-adjusting seal 41 is located radially outward of the pre-swirl cavity 30.

Description

GAS TURBINE ENGINE WITH TURBINE COOLING AIR DELIVERY SYSTEM

BACKGROUND

[0001] 1. Field

[0002] Disclosed embodiments are generally related to turbine engines, and in particular to cooling systems of the turbine engine.

[0003] 2. Description of the Related Art

[0004] A high pressure turbine (HPT) cooling air delivery system typically 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 into the pre-swirl cavity through the lower seal. This leakage is hotter than the pre-swirled air for at least two reasons. First, windage heating of the leakage, due to friction on the multiple teeth. Second, the leakage swirls at a lower tangential velocity (around half) of that of the pre-swirled air which creates an offset between the mixed air swirl and the disc speed.

[0006] Another variation of arrangement of the gas turbine engine comprises dropping the pressure upstream of the lower labyrinth seal by means of another sealing element so as to reverse the flow direction. The result is that higher temperature air is no longer leaking into the pre-swirl cavity, i.e. only pre-swirled air is feeding the disc and the blade internal passages. This reversed air flow is directed to the rim cavity through a bypass in the pre-swirlers support structure.

SUMMARY

[0007] Briefly described, aspects of the present disclosure relate to gas turbine engines and more particularly to sealing portions of the gas turbine engine. [0008] An aspect of the present disclosure is directed to a gas turbine engine. The gas turbine engine includes a first passage formed by an exterior wall of a combustor and an inner wall within the gas turbine engine and a second passage formed by the inner wall and a rotating arm. A pre-swirl cavity is located at a downstream end of the first passage. A first self-adjusting seal is located in the second passage. A second self- adjusting seal located radially outward of the pre-swirl cavity.

BRIEF DESCRIPTION OF THE DRAWINGS

[0009] Fig. 1 is diagrammatic view of a gas turbine engine.

[0010] Fig. 2 is a cut-away view of a self-adjusting seal.

[0011] Fig. 3 is a diagrammatic view of a gas turbine engine with self-adjusting seals.

[0012] Fig. 4 is a close up view of the pre-swirl cavity shown in Fig. 3.

[0013] Fig. 5 is a close up view of the pre-swirl support structure shown in Fig. 3.

DETAILED DESCRIPTION

[0014] To facilitate an understanding of embodiments, principles, and features of the present disclosure, they are disclosed hereinafter with reference to implementation in illustrative embodiments. Embodiments of the present disclosure, however, are not limited to use in the described systems or methods and may be utilized in other systems and methods as will be understood by those skilled in the art.

[0015] The components described hereinafter as making up the view of various embodiments are intended to be illustrative and not restrictive. Many suitable components that would perform the same or a similar function as the components described herein are intended to be embraced within the scope of embodiments of the present disclosure.

[0016] Through use of self-adjusting seals, the leakage into a pre-swirl cavity can be reduced or stopped. Referring now to Fig. 1, a gas turbine engine 10 is shown. The gas turbine engine 10 may include a compressor section 11 for compressing air. The compressed air from the compressor section 11 is conveyed to a combustion section 12, which produces hot combustion gases by burning fuel in the presence of the compressed air from the compressor section 11. The combustion gases are conveyed through a plurality of transition ducts to a turbine section 14 of the engine 10. The turbine section 14 comprises alternating rows of stationary vanes 8 and rotating blades 9.

[0017] In the turbine section 14, blade disc structures 15 are positioned adjacent to one another in an axial direction. The blade disc structures 15 define a rotor. Each of the blade disc structures 15 supports circumferentially spaced apart blades 9 and each of a plurality of lower stator support structures support a plurality of circumferentially spaced apart vanes 8. The vanes 8 direct the combustion gases from the transition ducts along a hot gas flow path to the blades 9 such that the combustion gases cause rotation of the blades 9, which in turn causes corresponding rotation of the blade disc structures 15 of the rotor.

[0018] A supply of fluid, can supply fluid within the gas turbine engine 10. The fluid may have a temperature of, for example, between about 1000-1200° F.

[0019] The fluid flows through passages formed between an inner wall 17 and an outer combustor wall 18. The fluid also flows between a rotating arm 16 and inner wall 17. The rotating arm 16 may be connected to the compressor section 11 and the turbine section 14. The rotating arm 16 may also be referred to as a“drive cone” or“shaft.”

[0020] The first passage 19 is formed between the outer combustor wall 18 and the inner wall 17. The second passage 20 is formed between the rotating arm 16 and the inner wall 17. A pre-swirl structure 33 is located downstream of, and in fluid communication with the first passage 19. The pre-swirl structure 33 comprises pre- swirlers 34, which may be configured, for example, as angled holes or nozzles, which are configured so that the tangential velocity of fluid ejected through the pre-swirl ers

34 matches that of the disc speed. The fluid from the first passage 19 flows through pre- swirl ers 34 into a pre-swirl cavity 30 formed between the pre-swirl er structure 33 and the blade disc structure 15. The fluid flow from the second passage 20 may leak into the pre-swirl cavity 30 through a lower seal 35. In the shown example, the lower seal

35 is a labyrinth seal defined by knife edges formed on rotating blade disc structures 15 which interface with stationary honeycombed structures provided on the lower extension of the pre-swirl structure 33. In Fig. 1, the area of leakage is shown at position 31. [0021] It is desirable to prevent or minimize the leakage that occurs into the pre- swirl cavity 30 to improve the efficiency of the gas turbine engine 10. As illustrated in Fig. 3, this can be accomplished by using a first self-adjusting seal 40 in the second passage 20 and a second self-adjusting seal 41 surrounding (i.e., positioned radially outward of) the pre-swirl cavity 30. An example of a self-adjusting seal 40, 41 is shown in Fig. 2. Self-adjusting seals 40, 41 of the type that may be employed in the instant invention may be found in U.S. Patent nos. 8,002,285; 8,172,232 and 8,919,781, the contents of which are hereby incorporated by reference..

[0022] As shown in Fig. 2, each self-adjusting seal 40, 41 comprises a seal shoe 43 supported by flexible beams 42. The beams 42, which may be, for example, configured as leaf springs, permit the radial position of the seal shoe 43 to adjust by being responsive to the pressures in the surrounding environment. In order to be adjustable, the beams 42 can be made of a material that is both durable and flexible. For example the beams 42 may be made of stainless steel. The beams 42 may also be made of inconel for high temperature application.

[0023] The pressure at the location of the self-adjusting seal 40, 41 determines the movement of the self-adjusting seal 40, 41. Too little leakage will result in a small pressure drop which may not be sufficient to activate the self-adjusting seal 40, 41, so the beams 42 will stay at their cold build position. Too much leakage and pressure drop across the self-adjusting seal 40, 41 may result in the seal shoe 43 being pulled into place by the suction pressure, whereby the position of the seal shoe 43 self-adjusts to reduce the leakage.

[0024] Referring to Fig. 3, similar to the configuration of Fig. 1, the pre-swirl structure 33 is located downstream of and in fluid communication with the first passage 19. The pre-swirl structure 33 comprises pre-swirlers 34, which may be configured, for example, as angled holes or nozzles, which are configured so that the tangential velocity of the fluid ejected through the pre-swirlers 34 matches that of the disc speed. The fluid from the first passage 19 flows through pre-swirlers 34 into a pre-swirl cavity 30 formed between the pre-swirler structure 33 and the blade disc structure 15. A lower seal 35 is positioned radially inward of the pre-swirlers 34, which may be defined by knife edges formed on rotating blade disc structures 15 and honeycombed structures on the lower extension of the stationary pre-swirl structure 33. Cooling fluid exits the pre-swirl cavity 30 with a velocity component in a direction tangential to the circumferential direction of the gas turbine engine 10. A swirl ratio is defined as the velocity component in the direction tangential to the circumferential direction of the cooling fluid as compared to a velocity component of a rotating shaft in the direction tangential to the circumferential direction.

[0025] Referring jointly to Fig. 3 and 4, the second self-adjusting seal 41 is oriented so that the seal shoe 43 is located proximate to the transition duct and the first passage 19 within the pre-swirl cavity 30. The beams 42 of the second self-adjusting seal 41 move in response to the pressures within the pre-swirl cavity 30 and surrounding environment. The second self-adjusting seal 41 is located radially outward of the pre- swirl cavity 30 and above the lower seal 35. The second self-adjusting seal 41 may be secured to the inner wall 17 using a bolt 3.

[0026] Referring back to Fig. 3, the first self-adjusting seal 40 is located in the second passage 20 between the inner wall 17 and the rotating arm 16. The beams 42 of the first self-adjusting seal 40 are located proximate to the inner wall 17. The shoe 43 is located proximate to the rotating arm 16. The first self-adjusting seal 40 may be secured to the inner wall 17 using bolt 4.

[0027] The first self-adjusting seal 40 seals a leakage flow between the second passage 20 and a collection chamber 21 located downstream of the first self-adjusting seal 40. The first self-adjusting seal 40 thus acts to produce a pressure drop at the collection chamber 21. The reduced pressure in the collection chamber 21 causes fluid to move from the pre-swirl cavity 30 into the collection chamber 21 via a reverse flow 37 across the lower seal 35. Fluid from the collection chamber 21 is bypassed, via a bypass channel 23 into a rim cavity 46, where it contributes to purging and cooling of the rim cavity. The bypass channel 23 allows evacuating the mix of air from the second passage20 and the lower seal 35, without contaminating the pre-swirl cavity 30.

[0028] Referring to Fig. 5, the bypass channel 23 may be composed of drillings and pockets in the pre-swirl support structure 27. Fluid from the first passage 19 enters through the pockets 28 into the pre-swirl cavity 30. The flow in the bypass channel 23 enters through the pockets 29 in the pre- swirl support structure 27.

[0029] Forming part of the bypass channel 23 is a diaphragm 48. The diaphragm 48 isolates the bypass air flow 23 from the leakage flow 47 through the self-adjusting seal 40 that comes from pre-swirl cavity 30.

[0030] A blade cooling passage 50 is defined through the blade disc structure 15, located downstream of the pre-swirl cavity 30. The second self-adjusting seal 41 prevents or minimizes a leakage flow from the pre-swirl cavity 30 to the rim cavity 46, causing a fluid flow 51 to move from the pre-swirl structure 33 through the blade cooling flow passage 50 into the turbine section 14 in order to provide cooling for the turbine discs 15.

[0031] As discussed above, the first self-adjusting seal 40 causes the fluid flow to reverse through the lower seal 35 that is located in the pre-swirl cavity 30, causing a reverse flow 37 The lower seal 35 may be a labyrinth seal, as illustrated above, or a brush seal. The first self-adjusting seals 40 and the second self-adjusting seal 41 react to the pressure drop across them to maintain a tight running clearance over a wide range of conditions. Unlike labyrinth seals, the first self-adjusting seal 40 and the second self- adjusting seal 41 are effectively insensitive to thermal movements and instead respond to changes in pressure.

[0032] A gas turbine engine 10 having an air system architecture with a collection chamber 21, bypass channel 23 combined with the first self-adjusting seals 40 and the second self-adjusting seal 41 has several advantages. The first self-adjusting seal 40 and the second self-adjusting seal 41 provide a uniform performance of the gas turbine engine 10 over a wide range of operating conditions. The first self-adjusting seals 40 and the second self-adjusting seal 41 also enable a robust and consistent reverse flow 37 through the lower seal 35, which results in reduced blade cooling air temperature. The resulting temperature control can provide a potential for power increase and/or gain in the life of the gas turbine components.

[0033] The first self-adjusting seals 40 and the second self-adjusting seal 41 have a tight clearance with respect to traditional labyrinth seals. Wear of the honeycomb or line facing the lower seal 35 can contribute to the deterioration of the system performance. The low leakages and flow through the lower seal 35 improve the aerodynamics in the pre-swirl cavity 30. Better cooling is equivalent to additional power margins to maintain the same blade metal temperature.

[0034] While embodiments of the present disclosure 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.

Claims

CLAIMS What is claimed is:

1. A gas turbine engine comprising:

a first passage formed by an exterior wall of a combustor and an inner wall within the gas turbine engine;

a second passage formed by the inner wall and a rotating arm;

a pre-swirl cavity located at a downstream end of the first passage;

a first self-adjusting seal located in the second passage; and

a second self-adjusting seal located radially outward of the pre-swirl cavity.

2. The gas turbine engine of claim 1, further comprising a bypass channel connected to the pre-swirl cavity.

3. The gas turbine engine of claim 2, wherein the bypass channel is adapted to transmit fluid from the pre-swirl cavity and the second passage.

4. The gas turbine engine of claim 1, wherein the first self-adjusting seal comprises a seal shoe supported on a plurality of flexible beams.

5. The gas turbine engine of claim 1, wherein the second self-adjusting seal comprises a seal shoe supported on a plurality of flexible beams.

6. The gas turbine engine of claim 1, further comprising a lower seal located in the pre-swirl cavity.

7. The gas turbine engine of claim 6, wherein the first self-adjusting seal causes a fluid flow through the lower seal into a bypass channel.