EP3106613A1

EP3106613A1 - Cooling system for gas turbine engines

Info

Publication number: EP3106613A1
Application number: EP15001689.7A
Authority: EP
Inventors: Gabriel L. Suciu; Brian D. Merry; James D. Hill
Original assignee: United Technologies Corp
Current assignee: Raytheon Technologies Corp
Priority date: 2015-06-06
Filing date: 2015-06-08
Publication date: 2016-12-21

Abstract

A cooling system for gas turbine engines (100) includes a turbine rotor compartment (106) defining a cooling air plenum (120). A plurality of turbine discs (130,132) are rotatably housed within the rotor compartment (106). A cooling air inlet (122) is in fluid communication with the plenum (120). Each turbine disc includes a cooling outlet (138) in fluid communication with the plenum for cooling the rotor compartment.

Description

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present disclosure relates to turbine engines, and more particularly to turbine engines having improved high pressure turbine cooling.

2. Description of Related Art

A gas turbine engine typically includes a multi-stage axial compressor, a combustor, and a turbine. Airflow entering the compressor is compressed and directed to the combustor where it is mixed with fuel and ignited, producing hot combustion gases used to drive the turbine. To control the heat transfer induced by the hot combustion gases entering the turbine, typically cooling air is channeled through turbine cooling circuits and is used to cool various turbine components.
Maintaining sufficient cooling air within the gas turbine engine is critical to proper engine performance and component longevity. The flow of cooling air across the turbine rotor and through the interior of the blades removes heat so as to prevent excessive reduction of the mechanical strength properties of the turbine blades and turbine rotor. Typical cooling methods include directing cooling air from a variety of sources having different pressures and temperatures.
Such conventional methods and systems have generally been considered satisfactory for their intended purpose. However, there is still a need in the art for improved turbine cooling systems. The present disclosure provides a solution for this need.

SUMMARY OF THE INVENTION

A cooling system for gas turbine engines includes a turbine rotor compartment defining a cooling air plenum. A plurality of turbine discs are rotatably housed within the rotor compartment. A cooling air inlet is in fluid communication with the plenum. Each turbine disc includes a cooling outlet in fluid communication with the plenum for cooling the rotor compartment. The cooling air inlet can be in fluid communication with a cooling air conduit aligned to deliver fluid output from a high pressure compressor.
The rotor compartment can include a plurality of seals proximate a hub of each turbine disc to fluidly seal the plenum. The rotor compartment can have first and second turbine discs. Each turbine disc can be sealingly engaged to the rotor compartment. A segmented seal can be disposed between the first and second turbine discs to isolate the plenum from a gas path outboard of the rotor compartment.
Each turbine disc can include a plurality of turbine blades mounted thereabout, wherein the turbine blades provide a cooling air outlet of the plenum. Each turbine disc can also have at least one passage through the hub thereof to allow cooling air from the inlet to circulate through the plenum. Further, each turbine blade can have an axial passage therethrough to place the cooling air inlet in fluid communication with a portion of the plenum between the turbine discs. The passage can be disposed proximate of the turbine blades of each respective turbine disc.
A method for providing cooling air to a rotor stage includes delivering fluid output from a high pressure compressor through a cooling air inlet into a cooling air plenum defined by a rotor compartment. The method can further include sealing the cooling air plenum with a plurality of seals disposed near a hub of each turbine disc. In addition, a segmented seal can be disposed between first and second turbine discs to isolate the cooling air plenum from a gas path outboard of the rotor compartment. The method can further include cooling the plenum by flowing the fluid output through at least one passage of a hub of each turbine disc. Each turbine disc may include a plurality of turbine blades, each turbine disc having at least one passage proximate the turbine blades to place the cooling air inlet in fluid communication with a portion of the plenum between the turbine discs.
These and other features of the systems and methods of the subject disclosure will become more readily apparent to those skilled in the art from the following detailed description of the preferred embodiments taken in conjunction with the drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

So that those skilled in the art to which the subject disclosure appertains will readily understand how to make and use the devices and methods of the subject disclosure without undue experimentation, preferred embodiments thereof will be described in detail herein below by way of example only and with reference to certain figures, wherein:

Fig. 1 is a side view of an exemplary embodiment of an incremental cooling system constructed in accordance with the present disclosure, showing a rotor compartment of a gas turbine engine being cooled from a single source of cooled air; and
Fig. 2 is a flow chart showing the method steps for providing cooling air to a rotor stage.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Reference will now be made to the drawings wherein like reference numerals identify similar structural features or aspects of the subject disclosure. For purposes of explanation and illustration, and not limitation, a partial view of an exemplary embodiment of a cooling system for gas turbine engines in accordance with the disclosure is shown in Fig. 1 and is designated generally by reference character 100. Other embodiments of the incremental cooling system in accordance with the disclosure, or aspects thereof, are provided in Fig. 2, as will be described.
A typical prior art method of cooling a gas turbine engine includes using cooling air to cool first and second rotors from a variety of sources. For example, one cooling line is typically provided from cooled air external to the system whereas another cooling line is typically for cooled compressor discharge air. Each of the different cooling lines being directed to cool different parts of the rotors. Typically, turbine discs in turbine engines of the prior art do not allow for cooled air to reach between or around each disc. Moreover, the sources of cooling air used in the typical configurations are at different levels of pressure and temperature. This variation leads to compromises and increased stress on the turbine components due to the thermal gradient.
Fig. 1 illustrates an exemplary embodiment of a turbine engine 100 of the present disclosure where first and second turbine discs 130, 132, respectively, of a rotor compartment 106 are cooled entirely from the same source of cooling air S1. Uncombusted high pressure air, typically referred to as T3 air, for example, is passed through a tangential on-board injector ("TOB1") nozzle. The TOB1 nozzle reduces the relative total temperature of the T3 air. The reduced temperature T3 air is directed to the rotor compartment 106 to cool the entire rotor system. In addition, the flow paths within the rotor compartment and a plurality of seals around the rotor compartment are arranged to provide cooling air to the entire rotor compartment from the single source, without the need for additional sources of cooling air. This ensures that the entire rotor system is cooled from a single source with one temperature and pressure thereby improving the longevity of motor life compared to a traditional system.
With continued reference to Fig. 1, the rotor compartment 106 defines a cooling air plenum 120 in fluid communication with a cooling air inlet 122. T3 cooling air is directed to the cooling air inlet via a cooling air conduit 124. A plurality of turbine discs 130, 132 are rotatably housed within the rotor compartment 106. First and second turbine discs 130,132 are shown in Fig. 2, however, the rotor compartment may contain any suitable number of turbine discs.
The cooling air plenum 120 is fluidly sealed by a plurality of seals 140, 142, 144 disposed around the turbine discs 130, 132 to cool the entire rotor compartment 106. For example, seals 140 and 142 are located near hubs 136B, 136A of the turbine discs 130, 132. In addition, a segmented seal 144 is disposed between the first and second turbine discs 130, 132. The segmented seal 144 isolates the cooling air plenum 120 from a gas path 126 outboard of the rotor compartment 106. This allows the cooling air to flow inboard of the segmented seal 144 around the first and second turbine discs 130, 132.
To direct the flow of cool air within the entire rotor compartment 106, each turbine disc 130, 132 of the present disclosure includes passages in fluid communication with the plenum for cooling the rotor compartment. More specifically, a plurality of passages 150, 152, 154, 156 are disposed within turbine discs 130, 132 to allow the cooling air to flow therethrough and reach the entire plenum. At least one passage 150, 152 is provided through hubs 136B, 136A of each of the first and second turbine discs 130, 132. This allows cooling air to flow through the hub and between the first and second turbine discs 130, 132 as indicated by the flow arrows in Fig. 1. At least one axial passage 154, 156 is also provided within each turbine disc 130, 132. This allows cooling air to flow between each turbine disc 130, 132 and around the segmented seal 144. Each turbine disc includes a plurality of turbine blades 137, 139 mounted thereabout. The turbine blades 137, 139 provide cooling air outlets 138 to the plenum 120.
The seals and passages are positioned to fluidly seal the rotor compartment while allowing the cooling air to circulate within the entire plenum to cool components of the turbine engine. Additional seals and/or passages may be included. Further, the positioning of the seals and/or passages may be adjusted, changed or modified without departing from the scope of the present disclosure.
A method for providing cooling air to a rotor stage is also disclosed. The method 200 comprises, at box 202, delivering fluid output from a high pressure compressor through a cooling air inlet, e.g., air inlet 122, into a cooling air plenum, e.g., plenum 120, defined by a rotor compartment, e.g., rotor compartment 106. The method further includes sealing the plenum with a plurality of seals, e.g., seals 140, 142, 144, at box 204. At least one seal disposed near a hub of each turbine disc and a segmented seal between first and second turbine discs. The segmented seal isolating the plenum from a gas path outboard the rotor compartment, as shown in box 206. At described in boxed 208 and 210, to allow fluid to flow through each turbine disc and cool the plenum with fluid output, at least one passage, e.g., passages 150, 152, 154, 156, is positioned through the hub of each turbine disc. Further, placing the cooling air inlet in fluid communication with a portion of the plenum between the turbine discs is constructed by providing at least one passage proximate turbine blades, e.g., blades 137, 139.
While shown and described in the exemplary context of using T3 air as cooling fluid, any other suitable cooling fluid from any other suitable source can be used without departing from the scope of this disclosure. Moreover, while shown and described in the exemplary context of cooling turbine components, any other suitable components, e.g., compressor or combustor components, can also be cooled with the methods and delivery described herein.
The methods and systems of the present disclosure, as described above and shown in the drawings, provide for an incremental cooling system for turbine engines with superior properties including cooling a turbine engine with cooled air from a single source. While the apparatus and methods of the subject disclosure have been shown and described with reference to preferred embodiments, those skilled in the art will readily appreciate that changes and/or modifications may be made thereto without departing from the scope of the subject disclosure as defined by the claims.

Claims

A cooling system for gas turbine engines (100), comprising:
a turbine rotor compartment (106) defining a cooling air plenum (120);

a plurality of turbine discs (130,132) rotatably housed within the rotor compartment; and

a cooling air inlet (122) in fluid communication with the plenum, wherein each turbine disc includes a cooling outlet (138) in fluid communication with the plenum for cooling the rotor compartment.
The cooling system of claim 1, wherein the cooling air inlet is in fluid communication with a cooling air conduit (124) aligned to deliver fluid output from a high pressure compressor.
The cooling system of claim 1 or 2, wherein each turbine disc has a hub (136A, 136B) having at least one passage (150,152) therethrough constructed to allow the fluid output to cool the plenum.
The cooling system of claim 3, wherein the rotor compartment includes a plurality of seals (140,142,144) near each hub to fluidly seal the plenum.
The cooling system of any preceding claim, wherein the rotor compartment has first (130) and second (132) turbine discs with a segmented seal (144) disposed therebetween to isolate the plenum from a gas path (126) outboard of the rotor compartment.
The cooling system of any preceding claim, wherein each turbine disc includes a plurality of turbine blades (137,139), each turbine blade having at least one cooling outlet (138) in fluid communication with the plenum (120).
The cooling system of claim 6, wherein at least one of the turbine discs has an axial passage (154,156) therethrough to place the cooling air inlet in fluid communication with a portion of the plenum between the turbine discs.
The cooling system of claim 7, wherein the axial passage of each turbine disc is disposed proximate of turbine blades thereof.
A method (200) of providing cooling air to a rotor stage, comprising:
delivering (202) fluid output from a high pressure compressor through a cooling air inlet (122) into a cooling air plenum (120) defined by a rotor compartment (106) having a set of turbine discs (130,132) rotatably housed therein, each turbine disc including a cooling outlet (138) in fluid communication with the cooling air plenum for cooling the rotor compartment.
The method of claim 9, further including sealing (204) the cooling air plenum with a plurality of seals (140,142,144) disposed near a hub (136A, 136B) of each turbine disc.
The method of claim 9 or 10, further including isolating (206) the cooling air plenum from a gas path (126) outboard of the rotor compartment by disposing a segmented seal (144) between first and second turbine discs.
The method of claim 9, 10 or 11 further including cooling (208) the plenum by flowing (210) the fluid output through at least one passage (150,152) of a hub (136A,136B) of each turbine disc.
The method of any of claims 9 to 12, wherein each turbine disc includes a plurality of turbine blades (137,139), each turbine blade having at least one cooling outlet (138) in fluid communication with the cooling plenum (120).
The method of claim 13, further including placing the cooling air inlet in fluid communication with a portion of the plenum between the turbine discs by flowing the fluid output through at least one axial passage (154, 156) proximate the turbine blades.