Patent Application
20140050558
The subject matter disclosed herein relates to turbine systems, and more particularly to a temperature gradient management arrangement for such turbine systems, as well as a method of managing temperature gradients.
Turbine systems typically include a rotor having a plurality of stacked wheels. An outer radial region of the stacked wheels is known as a rim portion while a central radial region of the stacked wheels is known as a bore portion. Typical operation of turbine systems entails high temperatures which may subject various components of the turbine to relatively extreme thermal loads. The high temperatures of the main flow path are the result of the compression process and the combustion process within the turbine system. As a result, large temperature gradients in the stacked wheels are often present, particularly during startup of the turbine system while main flow path temperatures are significantly higher in temperature than the rotor. Specifically, the large temperature gradients are caused by the rim portion being significantly hotter than the bore portion.
The large temperature gradients described above impart thermal stresses that are superimposed on the mechanical stresses due to centrifugal forces and surface pressures. The stress and temperature history experienced by the rotor components determines damage accumulated over each operating cycle and therefore the life expectancy of the rotor.
According to one aspect of the invention, a temperature gradient management arrangement for a turbine system includes a rotor comprising a rotor bore extending axially along the rotor. Also included is a secondary flow path comprising an inlet for a secondary airflow to flow to the rotor bore and an outlet disposed axially upstream of the inlet, relative to a main flow direction of the turbine system. Further included is a flow rate manipulator disposed proximate the outlet and configured to increase a flow rate of the secondary airflow during a first turbine system operating condition and to decrease the flow rate of the secondary airflow during a second turbine system operating condition.
According to another aspect of the invention, a temperature gradient management arrangement for a turbine system includes a rotor comprising a plurality of stacked wheels and a rotor bore. Also included is a plurality of compressor stages. Further included is at least one secondary flow path comprising at least one inlet for a secondary airflow to flow to the rotor bore, wherein the at least one secondary flow path extends from the at least one inlet in an upstream direction relative to a main flow direction toward at least one outlet disposed proximate at least one of the plurality of compressor stages. Yet further included is a flow rate manipulator disposed proximate the at least one outlet and configured to increase a flow rate of the secondary airflow during a first turbine system operating condition and to decrease the flow rate of the secondary airflow during a second turbine system operating condition.
According to yet another aspect of the invention, a method of managing a temperature gradient of a turbine system is provided. The method includes routing a secondary airflow through an inlet to a rotor bore disposed within a rotor along a secondary flow path to an outlet. Also included is increasing a flow path clearance at the outlet to increase a flow rate of the secondary airflow through the secondary flow path during a first turbine system operating condition. Further included is decreasing the flow path clearance at the outlet to decrease the flow rate of the secondary airflow through the secondary flow path during a second turbine system operating condition.
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.
The terms “axial” and “axially” as used in this application refer to directions and orientations extending substantially parallel to a center longitudinal axis of a turbine system. The terms “radial” and “radially” as used in this application refer to directions and orientations extending substantially orthogonally to the center longitudinal axis of the turbine system. The terms “upstream” and “downstream” as used in this application refer to directions and orientations relative to an axial flow direction with respect to the center longitudinal axis of the turbine system.
Referring to
The combustor section 14 uses a combustible liquid and/or gas fuel, such as natural gas or a hydrogen rich synthetic gas, to run the gas turbine system 10. For example, fuel nozzles 20 are in fluid communication with a main flow path 26 exiting the compressor 12 and a fuel supply 22. The fuel nozzles 20 create an air-fuel mixture, and discharge the air-fuel mixture into the combustor section 14, thereby causing a combustion that creates a hot pressurized exhaust gas 28. The combustor section 14 directs the hot pressurized gas through a transition piece into a turbine nozzle (or “stage one nozzle”), and other stages of buckets and nozzles causing rotation of turbine blades within an outer casing 24 of the turbine section 16.
Referring now to
Each of the plurality of solid wheels 30 and the plurality of annular wheels 32 includes a rotor blade 38 projecting radially outwardly from the rotor 18, while a plurality of stator vanes 40 are mounted on a stator (not illustrated). Each of the plurality of stator vanes 40 is typically positioned alternately between the rotor blades 38 and for illustration simplicity, only two of the plurality of stator vanes 40 are referenced. The rotor blades 38 and the plurality of stator vanes 40 form a passage through which the main flow path 26 in the compressor section 12 flows. The temperature of the main flow path 26 increases progressively as its pressure increases through the compressor section 12. Consequently, the rim portion 34 of the rotor 18 is exposed to the progressively hotter main flow path 26, while the rotor bore 36 remains shielded from the main flow path 26. The rotor bore 36 forms the root of the rotor 18, thereby extending proximate a rotor centerline 42. Based on exposure of the rim portion 34 to the main flow path 26, a temperature gradient often exists between the rim portion 34 and the rotor bore 36 of the rotor 18. To mitigate the temperature gradient noted above, the rotor bore 36 may be heated during startup of the gas turbine system 10 and cooled during shutdown, which are the two operating periods that typically result in particularly high temperature gradients.
As illustrated, the compressor section 12 includes a plurality of compressor stages 44, with each stage comprising one or more circumferentially spaced stator vanes aligned in a row at a similar axial location, along with an axially preceding or succeeding row of circumferentially spaced rotor blades disposed at a similar axial location. The plurality of compressor stages 44 include a middle stage 46 disposed at a relatively axial mid-point of the plurality of compressor stages 44. A plurality of forward stages 48 are positioned upstream of the middle stage 46, with respect to a direction of the main flow path 26 flowing through the compressor section 12. Additionally, a plurality of aft stages 50 are positioned downstream of the middle stage 46, also with respect to a direction of the main flow path 26.
As described above, heating or cooling of the rotor bore 36 may beneficially reduce the temperature gradient present between radially inner portions of the rotor 18, such as the rotor bore 36 itself, and radially outer portions of the rotor 18, such as the rim portion 34. Accordingly, a secondary airflow 52 is provided from the main flow path 26 flowing throughout the compressor section 12 to the rotor bore 36. The remainder of the main flow path 26 typically flows to the combustor section 14 and as a cool/purge flow 53 for the turbine section 16. The secondary airflow 52 is routed to the rotor bore 36 through at least one inlet 54 disposed proximate the middle stage 46 and/or at least one of the plurality of aft stages 50, with the at least one inlet 54 part of a secondary flow path 56. Subsequent to routing of the secondary airflow 52 through the at least one inlet 54, the secondary airflow 52 is directed upstream (relative to the main flow path 26) along the secondary flow path 56 that is defined by the rotor bore 36. As the secondary airflow 52 travels upstream, the rotor bore 36, and therefore the radially inner portion of the rotor 18 are heated during startup to reduce the temperature gradient between the rotor bore 36 and the rim portion 34. It is to be appreciated that the secondary flow path 56 that the secondary airflow 52 is routed along may be of various path dimensions and shapes. For example, the secondary flow path 56 may extend around the annular wheels 32 and through the solid wheels 30, thereby forming a curved flow path referred to as a serpentine flow path.
Irrespective of the precise dimensions and shape of the secondary flow path 56, it is to be understood that the secondary flow path 56 typically extends upstream from the at least one inlet 54 to at least one outlet 58 disposed proximate at least one of the plurality of forward stages 48. Although the at least one outlet 58 may extend through a variety of components proximate at least one of the plurality of forward stages 48, it is contemplated that the at least one outlet 58 is disposed at a stator vane diaphragm 60 proximate the rim portion 34 of the rotor 18.
A flow rate manipulator 62 is located within or proximate the at least one outlet 58 to control a clearance 64 that facilitates expulsion of the secondary airflow 52 from the secondary flow path 56. In one embodiment, the flow rate manipulator 62 comprises an adjustable seal that alters the flow rate of the secondary airflow 52 traveling throughout the secondary flow path 56. Typically, during startup and shutdown of the gas turbine system 10, the rotor components may undergo an axial deflection with respect to the stator parts due to a combination of temperature differential, associated thermal expansion and load stresses. The transiently varying relative deflection reaches a maximum during the full load operating time and remains constant during the entire duration of steady state, or near steady state, operation of the gas turbine system 10. Thus, prior to and during startup, the flow rate manipulator 62 may be positioned such that it does not block the at least one outlet 58, thereby inducing the secondary airflow 52 throughout the secondary flow path 56 due to a pressure differential between the rotor bore 36 and the main flow path 26. As the stator vane components to which the flow rate manipulator 62 is operably coupled to expand axially during transient state operation (referred to as a first operating condition) with respect to the rotor components, the flow rate manipulator may relatively move and begin to cover and block at least a portion of the at least one outlet 58, thereby restricting the flow of the secondary airflow 52 by decreasing the clearance 64, which in turn decreases the flow rate of the secondary airflow 52 within the secondary flow path 56 for heating therealong. As the gas turbine system 10 reaches steady state, or near steady state, operation, referred to as a second operating condition, the flow rate manipulator 62 slows or stops movement, with respect to the rotor components and may continue to cover all or part of the at least one outlet 58.
It is contemplated that various alternative arrangements may be employed to suitably increase and decrease the flow rate of the secondary airflow 52 during transient state operation and steady state operation, respectively.
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.
