COMBINED CONVECTION/EFFUSION COOLED ONE-PIECE CAN COMBUSTOR
An industrial turbine engine comprises a combustion section, an air discharge section downstream of the combustion section, a transition region between the combustion and air discharge section, a combustion transition piece and a sleeve. The transition piece defines an interior space for combusted gas flow. The sleeve surrounds the combustor transition piece so as to form a flow annulus between the sleeve and the transition piece. The sleeve includes a first set of apertures for directing cooling air from compressor discharge air into the flow annulus. The transition piece includes an outer surface bounding the flow annulus and an inner surface bounding the interior surface, and includes a second set of apertures for directing cooling air in the flow annulus to the interior space. Each of the second set of apertures extends from an entry portion on the outer surface to an exit portion on the inner surface.
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1. Field of the Invention
The present invention relates generally to means of cooling components of a gas turbine, and more particularly, to the cooling of a one-piece can combustor by a combination of convection cooling and effusion cooling.
2. Description of the Related Art
A gas turbine can operate with great efficiency if the turbine inlet temperature can be raised to a maximum. However, the combustion chamber, from which combusted gas originates before entering the turbine inlet, reaches operating temperatures well over 1500° F. and even most advanced alloys cannot withstand such temperatures for extended periods of use. Thus, the performance and longevity of a turbine is highly dependent on the degree of cooling that can be provided to the turbine components which are exposed to extreme heating conditions.
The general concept of using compressor discharge air to cool turbine components is known in the art. However, developments and variations in turbine designs are not necessarily accompanied by specific structures that are implemented with cooling mechanisms for the turbine components. Thus, there is a need to embody cooling mechanisms into newly developed turbine designs.
BRIEF DESCRIPTION OF THE INVENTIONAccordingly, it is an aspect of the present invention to enhance conventional gas turbines.
To achieve the foregoing and other aspects and in accordance with the present invention, an industrial turbine engine is provided that comprises a combustion section, an air discharge section downstream of the combustion section, a transition region between the combustion and air discharge section, a combustor transition piece defining the combustion section and transition region, and a sleeve. Said transition piece is adapted to carry combusted gas flow to a first stage of the turbine corresponding to the air discharge section. The transition piece defines an interior space for combusted gas flow. The sleeve surrounds the combustor transition piece so as to form a flow annulus between the sleeve and the transition piece. Said sleeve includes a first set of apertures for directing cooling air from compressor discharge air into the flow annulus. The transition piece includes an outer surface bounding the flow annulus and an inner surface bounding the interior surface. The transition piece includes a second set of apertures for directing cooling air in the flow annulus to the interior space. Each of the second set of apertures extends from an entry portion on the outer surface to an exit portion on the inner surface.
The foregoing and other aspects of the present invention will become apparent to those skilled in the art to which the present invention relates upon reading the following description with reference to the accompanying drawings, in which:
Example embodiments that incorporate one or more aspects of the present invention are described and illustrated in the drawings. These illustrated examples are not intended to be a limitation on the present invention. For example, one or more aspects of the present invention can be utilized in other embodiments and even other types of devices.
In
In
In conventional combustors, a combustor liner and a flow sleeve are generally found upstream of the transition piece and the sleeve respectively. However, in the one-piece can combustor of
The first apertures 400 are configured to be normal to the wall 500 such that air flow I is adapted to not strike or directly impinge an outer surface 300a of the transition piece 120 perpendicularly. The first apertures 400 may be formed directly above the second apertures 200 (
As shown in
Another variation of the apertures 200 is that the angular position of the entry portion 200a may be different from the angular position of the exit portion 200b on the circumference of the transition piece 120. Moreover, the exit portion 200b of the apertures 200 may be upstream or forward relative to the entry portion 200a of the apertures 200 thereby creating an obtuse angle between the longitudinal axes of the apertures 200 and the direction 202.
In
The first apertures 400 also have a substantially cylindrical geometry with a constant diameter. In one embodiment, the diameter may range from 0.1 inch to 1.0 inch. However, other dimensions for the apertures 400 are also contemplated.
Also, the apertures 200, 400 may gradually increase or decrease in diameter through the walls 300, 500 respectively.
The second apertures 200 may be formed on the wall 300 of the transition piece 120 by laser drilling or other machining methods selected based on factors such as cost and precision. The larger dimensions of the first apertures 400 allow for more tolerance and thus similar or more cost-effective machining methods may be used to form the apertures 400.
In
The invention has been described with reference to the example embodiments described above. Modifications and alterations will occur to others upon a reading and understanding of this specification. Example embodiments incorporating one or more aspects of the invention are intended to include all such modifications and alterations insofar as they come within the scope of the appended claims.
Claims
1. A turbine engine comprising:
- a combustion section;
- an air discharge section downstream of the combustion section;
- a transition region between the combustion section and air discharge section;
- a combustor transition piece defining the combustion section and transition region, said transition piece adapted to carry combusted gas flow to a first stage of the turbine engine corresponding to the air discharge section, the transition piece defining an interior space for combusted gas flow; and
- a sleeve surrounding the combustor transition piece so as to form a flow annulus between the sleeve and the transition piece, said sleeve including a first set of apertures for directing cooling air from compressor discharge air into the flow annulus,
- wherein the transition piece includes an outer surface bounding the flow annulus and an inner surface bounding the interior space, the transition piece includes a second set of apertures for directing cooling air in the flow annulus to the interior space, and each of the second set of apertures extends from an entry portion on the outer surface to an exit portion on the inner surface.
2. The turbine engine of claim 1, wherein the first set of apertures are normal to the sleeve.
3. The turbine engine of claim 1, wherein the first set of apertures has a constant diameter ranging from 0.1 inch to 1.0 inch.
4. The turbine engine of claim 1, wherein one of the entry portion and the exit portion is located further downstream than the other of the entry portion and the exit portion.
5. The turbine engine of claim 4, wherein the combustor transition piece is a can-annular, reverse-flow type such that combusted gas flow and compressor discharge air flow are configured to be in opposing directions such that longitudinal axes through the second set of apertures form an acute angle with a direction of combusted gas flow and an obtuse angle with a direction of compressor discharge air flow.
6. The turbine engine of claim 1, wherein longitudinal axes through the second set of apertures are oriented to form an acute angle with a downstream tangent to the outer surface.
7. The turbine engine of claim 6, wherein the acute angle ranges from 20° to 35°.
8. The turbine engine of claim 1, wherein the second set of apertures have a constant diameter from the entry portion to the exit portion ranging from 0.02 inch to 0.04 inch.
9. The turbine engine of claim 1, wherein the second set of apertures are substantially normal to the outer surface.
Type: Application
Filed: Apr 13, 2009
Publication Date: Oct 14, 2010
Applicant: General Electric Company (Schenectady, NY)
Inventors: Ronald James Chila (Greer, SC), Kevin Weston McMahan (Greer, SC)
Application Number: 12/422,536
International Classification: F23R 3/00 (20060101);