Patent Application
20140245746
The subject matter disclosed herein relates to gas turbine engines, and more particularly to a combustion arrangement, as well as a method of reducing pressure fluctuations of the combustion arrangement.
A combustor section of a gas turbine system typically includes a combustor chamber disposed relatively adjacent a transition region that routes hot gas from the combustor chamber to a turbine section. Traditionally, the combustor chamber is defined by a combustor liner that is surrounded by a flow sleeve, with the transition region defined by a transition liner that is surrounded by an impingement sleeve. More recently, combustor sections have included the combustor chamber and the transition region within a single liner. An aft end of the combustor section may experience large pressure fluctuations. Such pressure fluctuations may reduce the life of the liner, as well as buckets within the turbine section, due to a continuous imposition of high overall dynamics amplitudes for pressure tones exhibited at the aft end of the combustor section.
According to one aspect of the invention, a combustion arrangement includes a combustion section. Also included is an air discharge section downstream of the combustion section. Further included is a transition region disposed between the combustion section and the air discharge section. Yet further included is a transition piece defining the combustion section and the transition region, wherein the transition piece is configured to carry a combusted gas flow from the combustion section to the air discharge section. Also included is a damping device operatively coupled to the transition piece proximate the air discharge section.
According to another aspect of the invention, a damped exit of a transition piece includes an air discharge section located at a downstream end of a liner of the transition piece. Also included is a resonator operatively coupled to the liner proximate the air discharge section, wherein the resonator is configured to damp pressure fluctuations within the transition piece.
According to yet another aspect of the invention, a method of reducing pressure fluctuations of a combustion arrangement is provided. The method includes flowing a combusted gas flow through a transition region of a transition piece from a combustion section to an air discharge section. Also included is damping pressure fluctuations within the transition piece with a damping device operatively coupled to the transition piece proximate the air discharge section.
These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings.
The subject matter, which is regarded as the invention, is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which:
The detailed description explains embodiments of the invention, together with advantages and features, by way of example with reference to the drawings.
Referring to
In operation, air flows into the compressor 12 and is compressed into a high pressure gas. The high pressure gas is supplied to the combustion arrangement 14 and mixed with fuel, for example natural gas, fuel oil, process gas and/or synthetic gas (syngas), in the combustion section 18. The fuel/air or combustible mixture ignites to form a high pressure, high temperature combustion gas stream. In any event, the combustion arrangement 14 channels the combustion gas stream to the turbine 24 which converts thermal energy to mechanical, rotational energy.
Referring now to
In one embodiment, a transition piece 42 is included and comprises a liner 44 that transitions as a single component directly from the head end 32, which may be of a substantially circular geometry, to the air discharge section 40, which may be of an oval cross-sectional geometric configuration that corresponds to an annular exit for the combusted gas flow 34 to a segment of the turbine 24. The liner 44 may be formed from two halves or several components welded or joined together for ease of assembly or manufacture. A sleeve 46 at least partially surrounds and is disposed radially outwardly of the liner 44. The sleeve 46 also transitions directly from the head end 32 to the air discharge section 40 as a single component. Similar to the liner 44, the sleeve 46 may be formed from two halves and welded or joined together for ease of assembly or manufacture. It is to be understood that reference to a “single” component, as employed above with respect to the liner 44 and the sleeve 46, may refer to multiple pieces joined together to form a single overall structure, where the joining is by any suitable process to join elements.
The air discharge section 40 may be shaped in various configurations to achieve desirable exit conditions of the combusted gas flow 34 at the point of expulsion to the turbine 24, and more specifically for delivery of the combusted gas flow 34 to a first stage 48 of the turbine 24. The first stage 48 typically includes a plurality of airfoils 50, such as a row of circumferentially spaced nozzles or buckets. In one embodiment, the transition piece 42 is shaped in what is referred to as a choked flow region 52 proximate the air discharge section 40. The choked flow region 52 refers to a region that imposes a restriction on the combusted gas flow 34 by decreasing the cross-sectional area through which the combusted gas flow 34 passes through. Due to the restriction, as well as a lower pressure environment of the turbine 24 disposed downstream of the choked flow region 52, the fluid velocity of the combusted gas flow 34 increases. The choked flow region 52 may be employed to mimic a first stage nozzle of the turbine 24, such that inclusion of the first stage nozzle is optional.
One effect of routing the combusted gas flow 34 through the choked flow region 52 is a large magnitude of pressure fluctuation proximate the air discharge section 40. To dampen the pressure fluctuations experienced, a damping device 60 is operatively coupled to the transition piece 42 proximate the air discharge section 40. In one embodiment, the damping device 60 is coupled to an outer surface 62 of the liner 44, and between the liner 44 and the sleeve 46 for embodiments including the sleeve 46. Alternatively, the damping device 60 may be coupled to an outer portion of the sleeve 46. Irrespective of the precise location of coupling of the damping device 60 to the transition piece 42, the damping device 60 may partially surround or fully surround the transition piece 42 along an axial segment of the transition piece 42.
In an exemplary embodiment, the damping device 60 comprises a resonator, which may be of an electromagnetic or mechanical type. The resonator exhibits resonance or resonant behavior, that is, it naturally oscillates at some frequencies, called its resonant frequencies, with greater amplitude than at others. The resonant frequencies may be configured to dampen the overall pressure fluctuations exhibited proximate the air discharge section 40 by reducing the amplitude of pressure anti-nodes.
The damping device 60 may also include at least one, but typically a plurality of cooling holes 70 for routing of a cooling flow 64 to the outer surface 62 of the liner 44 (
As illustrated in the flow diagram of
While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.
This application was made with U.S. Government support under Agreement No. DE-FC26-05NT42643 awarded by the Department of Energy. The U.S. Government may have certain rights in this invention.
