CLEARANCE CONTROL

BACKGROUND

1. Field [0001] The present invention relates to a method of operating inlet heating systems for clearance control.

2. Description of the Related Art

[0002] In an axial flow industrial gas turbine engine, hot compressed gas is produced. The hot gas flow is passed through a turbine and expands to produce mechanical work used to drive an electric generator for power production. The turbine generally includes multiple stages of stator vanes and rotor blades to convert the energy from the hot gas flow into mechanical energy that drives the rotor shaft of the engine.

[0003] A combustion system receives air from a compressor and raises it to a high energy level by mixing in fuel and burning the mixture, after which products of the combustor are expanded through the turbine.

[0004] Gas turbines engines are becoming larger, more efficient, and more robust. Large blades and vanes are being utilized, especially in the hot section of the engine system. In view of high pressure ratios and high engine firing temperatures implemented in modern engines, certain components, such as airfoils, e.g., stationary vanes and rotating blades

[0005] In current assemblies, clearance between the rotating and stationary components in gas turbines are regions of impacting the engine performance significantly. There are several drivers of aerodynamic loss in the compressor- vane carrier and turbine-shroud cavity configuration, which lowers the gas turbine’s efficiency. One driver is the flow over the rotating components. The mixing losses that occur downstream of the clearance area are high and contribute to a reduction in stage efficiency and power. Additional mixing losses occur when the flow through the tip cavity combines with the main flow and the two streams have different velocities. Tip leakage is essentially lost opportunity for work extraction. The tip leakage also contributes towards aerodynamic secondary loss.

[0006] Compressor mid and rear stage clearance design for instance is dictated by hot restarts that are typically one to four hours after engine shutdown. After the engine powers down to a certain low turning gear speed, such as l20rpm for example, the compressor vane carriers cool a lot faster than the corresponding compressor disks. Hence, at the hot restart moment, the tip clearance prior to engine ignition is significantly tighter than the cold build. To overcome these clearance pinch points, the build clearance has to be increased accordingly.

SUMMARY

[0007] In an aspect of the present invention, a method for clearance control for at least one vane carrier comprises: having an inlet heating system connected to an engine, the inlet heating system connected along a component exit shell, piping to an inlet of the engine, wherein the inlet heating system extracts hot air from the component exit shell when activated and pipes the hot air back towards the inlet, to warm up the air approaching the inlet of the gas turbine, the inlet heating system controlled by control valves while the engine is running; activating the inlet heating system to recirculate air from the component exit shell back to the inlet as part of an engine shut down process, recovering some of the available energy to raise the temperature of the inlet heating system and gas path as the engine shuts down; keeping active the inlet heating system as the gas turbine cools down on turning gear speed, wherein the recirculating air from the component exit shell flows from the shell extraction, retaining heat in the engine and front component stages keeping at least one vane carrier warm; and keeping active the inlet heating system while the engine increases speed up to a predetermined load condition, as the engine prepares for restart, wherein the heating up of the at least one vane carrier is accelerated to increase clearances during restart.

[0008] These and other features, aspects and advantages of the present invention will become better understood with reference to the following drawings, description and claims.

BRIEF DESCRIPTION OF THE DRAWINGS

[0009] The invention is shown in more detail by help of figures. The figures show preferred configurations and do not limit the scope of the invention.

[0010] FIG. 1 is an elevational, cross sectional view of a combustion turbine engine, such as a gas turbine engine that of an exemplary embodiment of the present invention;

[0011] FIG. 2 illustrates an inlet heating system of an exemplary embodiment of the present invention;

[0012] FIG. 3 is a perspective view of a portion of a vane carrier of an exemplary embodiment of the present invention;

[0013] FIG. 4 is a cross-sectional view of a portion of the vane carrier of an exemplary embodiment of the present invention; and [0014] FIG. 5 illustrates an inlet heating system shell extraction of an exemplary embodiment of the present invention.

DETAILED DESCRIPTION

[0015] In the following detailed description of the preferred embodiment, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.

[0016] Broadly, an embodiment of the present invention provides a method for clearance control for at least one vane carrier. The method includes having an inlet heating system connected to an engine. The inlet heating system extracts hot air from a component exit shell when activated, piping the hot air back towards an inlet to warm up the air approaching the inlet of the gas turbine controlled by control valves. Activating the inlet heating system to recirculate air from the component exit shell back to the inlet as part of an engine shut down process. Keeping active the inlet heating system as the gas turbine cools down on turning gear speed, retaining heat in the engine and front component stages, keeping at least one vane carrier warm. The inlet heating system stays active up to a predetermined load condition, accelerating the heating up of the at least one vane carrier to increase clearances during restart.

[0017] A gas turbine engine typically includes a compressor section, a combustor section and a turbine section. The compressor section ingests ambient air and compresses it. The compressed air from the compressor section enters one or more combustors in the combustor section. The compressed air is mixed with fuel in the combustors, and an air-fuel mixture is combusted in the combustors to form a hot working gas. The hot working gas is routed to the turbine section where it is expanded through alternating rows of stationary airfoils and rotating airfoils and used to generate power that can drive a rotor. The expanded gas exiting the turbine section may then be exhausted from the engine via an exhaust section.

[0018] The compressor and turbine sections may include several locations in which there may be gaps or clearances between the rotating and stationary components. During engine operation, system loss may occur through fluid leakage through clearances in the compressor and turbine sections. This system loss decreases the operational efficiency of the system. An example of the flow leakage can occur across a clearance between the tips of rotating blades and a surrounding stationary structure or boundary, such as an outer shroud or a vane carrier.

[0019] Small clearances are desired to keep air leakage to a minimum; however, it is important to maintain at least some minimum clearance between the rotating and stationary components at all times. Rubbing of any of the rotating and stationary components can lead to substantial component damage, performance degradation, and extended outages. The size of the clearance can change during engine transient operation due to, as an example, differences in thermal inertia of rotor supporting the rotating blades compared to the thermal inertia of the stationary structure, such as the outer casing or the vane carrier. Because the thermal inertia of the vane carriers is substantially less than the thermal inertia of the rotor, the vane carrier has a faster thermal response time and can respond (through expansion or contraction) more quickly to a change in temperature than the rotor.

[0020] FIG. 1 shows a combustion turbine engine 10, such as a gas turbine. The engine 10 includes a compressor section 12 including an outer compressor casing 26 that encloses various compressor components, such as compressor vane carriers 28 supported from an interior structure defined on an inner side of the outer compressor casing 26. Stationary compressor vanes 30 are supported from the compressor vane carriers 28, and rotating blades 32 are supported on a rotor assembly 34 and may be located in alternating relation to the vanes 30 to form compressor stages. The vanes and blades extend radially across a flow path 36 extending from an inlet 38 at an upstream end of the compressor section 12 to an exhaust manifold 20.

[0021] FIG. 3 and FIG. 4 show a portion of the compressor blades 32 that include radially outer blade tips 32a that rotate proximate inner surfaces 28a of the vane carriers 28. The inner surfaces 28a of the vane carriers 28 define a radially outer boundary 29 for the flow path 36 within the compressor section 12. These figures, specifically FIG. 3, shows a more detailed view to see an example of clearance areas such as between the radially outer blade tip 32a and the radially inner surface 28a of the vane carrier 28.

[0022] The engine 10 further includes a combustor section 14 including a plurality of combustors 16, and a turbine section 18. Referring to FIG. 1, the combustor section 14 includes a combustor shell 44 defined within an outer combustor casing 46 that receives compressed air from the compressor section 12, referred herein as“shell air”. The shell air passes into the individual combustors 16 for combustion with a fuel to produce hot combustion gases. The hot combustion gases are conveyed through a transition duct 48 associated with each combustor 16 to the turbine section 18. The combustors 16 in the example shown are disposed about a longitudinal axis 24 of the engine 10 that defines an axial direction of the engine 10. [0023] Conceptually similar to the compressor section 12, the turbine section 18 includes vane carriers 50 supported within an outer turbine casing 52. Accordingly, based on the specific needs of a given application, such vane carriers can similarly benefit from the addition of heat from an inlet heating system 42, as described below.

[0024] The turbine section 18 includes blades and vanes as well. Stationary turbine vanes 54 are supported from the turbine vane carriers 50 and extend radially inward across the flow path 36. The turbine vane carriers 50 are supported within an outer turbine casing 52. The turbine vane carriers 50 also support outer shrouds or ring segments 55 located in an axially alternating arrangement with outer end walls of the vanes 54 to define a turbine portion of the radially outer boundary 29 of the flow path 36. Rotating turbine blades 56 are supported on respective turbine rotor discs 58 in an alternating arrangement with the vanes 54 to form stages of the turbine section 18. The rotating blades 56 extend radially outward across the flow path 36, and radially outer tips 56a of the blades 56 may be located adjacent to inner surfaces 55a of the ring segments 55. In certain embodiments, the outer compressor casing 26, the outer combustor casing 46, and the outer turbine casing 52 collectively define an outer casing 53 of the engine 10.

[0025] FIG. 2 shows an embodiment of a portion of an inlet heating system 42. The inlet heating system 42 is shown on the left side of the figure in the dashed boxed area. Hot air from the gas turbine/compressor enter from above in the schematic. Control valves 40 regulate the amount of hot air that enters through the inlet heating system 42. The hot air then can be sent to the inlet 38. Outside of the inlet heating system 42 the hot air may go through filters, and the like before moving towards the compressor section 12 at the far right of the figure.

[0026] FIG. 5 shows an inlet heating system 42 shell extraction 22 location from the outside of the engine 10. Hot air enters the piping leaving the engine 10 in FIG. 5 then proceeds to enter into the area shown in FIG. 2 shown as an arrow in the inlet heating system 42. The shell extraction 22 location shown in FIG. 5, the combustor shell 44, is an example of where the hot air can be extracted from the system. However, the extraction can occur at any location on the engine 10 where there is enough driving pressure for the flow and a temperature high enough the make a change in the temperature of the gas path 36.

[0027] In a first mode of operation, hot combustion gases are expanded through the stages of the turbine section 18 to extract energy, and at least a portion of the extracted energy from the combustion gases causes the rotor 34 to rotate and produce a work output during a power producing mode of operation of the engine 10.

[0028] In certain embodiments, the respective diameters of vane carriers 28, 50 and the respective lengths of blades 32, 56 are designed so that during engine startup, the tips 32a, 56a of the blades 32, 56 do not contact the inner surfaces 28a, 55a of the static structure defined by the vane carriers 28, 50 or equivalent structure such as the ring segments 55. The gap between the blade tips 32a, 56a and the static vane carriers 28, 50 can increase during transient operation due to the vane carrier temperature increasing.

[0029] During an initial engine startup (cold startup), the turbine blades 56 radially expand quickly due to a rapid increase in the temperature as a result of the hot working gases impinging on the blades 56 and centrifugal forces acting on the blades 56. Also, during start-up, the respective vane carriers 28 and 50 of the compressor 12 and turbine 18 expand radially outward away from the blade tips of the respective blades 32, 56 as the temperature of the vane carriers 28, 50 increases, typically creating a gap at the blade tips 32a, 56a that is larger than optimal for preventing or limiting secondary gas flows across the tips 32a, 56a.

[0030] During an engine startup that does not incorporate a pre-heating operation as described below, the respective vane carriers 28, 50 may expand at a slower rate than the radial outward expansion of the blades 32, 56 substantially reducing the gap or clearances between blades 32, 56 and the respective inner surfaces 28a, 55a of the vane carrier 28 and ring segment 55. Also, during a warm restart of the engine, the reduction in the blade to vane carrier clearances is exacerbated by the relatively high thermal inertia of the rotor assembly 34, with an associated higher temperature, in comparison to the vane carriers 28, 50 in that rotor assembly 34 can retain heat longer with an associated greater thermal expansion of the blades 32, 56 than the surrounding vane carriers 28, 50, causing the clearance gap to substantially decrease and be smaller than the cold gap. Hence, warm restarts represent a limiting transient clearance gap condition when the clearance gaps between blade tips 32a, 56a and inner surfaces of the outer flow boundary 29 are at a minimum. The clearance between the blade tips 32a, 56a and inner surfaces of the outer flow boundary 29 will hereinafter be referred to as“clearance gap”.

[0031] In certain embodiments, the inlet heating system 42 recirculates the shell air from the combustor shell 44 back to the inlet 38 of the compressor 12 section. Control valves 40 help to control the direction and output of the inlet heating system 42. Typically, during an engine shut down process, the inlet cools creating the issues mentioned above with clearance. The inlet heating system 42 can be activated during this engine shut down process as an improvement in order to recirculate hot shell air back to the inlet 38 to increase the temperature moving through either the compressor and/or the turbine sections.

[0032] In certain other embodiments, an air duct system 74 may be provided, that extends outside of the outer case 53 of the engine 10 between the compressor section 12 and the turbine section 18. One or more bleed air ducts may extend from the compressor section 12 to an axially downstream location on the engine 10, as shown in FIG. 1 by bleed air ducts 76a, 76b, and 76c. These air bleed ducts 76a, 76b, and 76c can also have lines extend upstream to the inlet 38. In these embodiments, the air duct system 74 may be operable during shutdown to recirculate the hot air from the air bleed ducts 76a, 76b, and 76c to provide higher temperatures at the inlet 38. Control valves 82a, 82b, and 82c can be used in the air duct system 74 as well, to be adjustable between fully open and fully closed positions, and can include a plurality of partially open positions between the fully open and fully closed positions, wherein valves 82a, 82b, and 82c may be configured to provide a range of continuously variable partially open positions to control the amount of flow through the respective bleed air ducts 76a, 76b, and 76c. In these embodiments, the air duct system 74 becomes part of the inlet heating system 42 with the shell air being provided from another location along the engine 10.

[0033] After a shutdown to turning gear operation (e.g., after engine has shut down), the inlet heating system can reduce the heat loss from the compressor section 12 and turbine section 18 of the engine 10. Additionally, use of the inlet heating system 42 can accelerate the heat up of the vane carriers 28, 50 during subsequent engine restarts.

[0034] In operation, the embodiments mentioned may be effective to transfer thermal energy directly to casings and/or casing components, e.g., vane carriers, which is conducive to a more efficient transfer of thermal energy to such components. Additionally, disclosed embodiments, may be effective to avoid or at least reduce casing ovalisation. That is, avoid or reduce the deviation of casing roundness (“ovalisation”) that otherwise would be caused by uneven temperature distribution within one or more stages of the compressor section 12 and/or turbine section 18, which can lead to increased and unequal radial clearances between rotating components and adjacent stationary components

[0035] In certain embodiments, the activation of the inlet heating system 42 can occur when a component exit temperature drops below a designated temperature, the activation recovering some of the available energy to raise the temperature of the inlet system and gas path as the engine shuts down. As the gas turbine continues to cool down while the speed decreases towards a turning gear speed, the recirculation of air through the inlet heating system 42 remains, in order to retain heat in the engine and front component stages keeping the vane carriers 28, 50 warm. The inlet heating system 42 remains active until a predetermined load condition occurs, ending the inlet heating system 42 activity. In certain embodiments, the inlet heating system 42 can remain activated until the engine reaches full speed with no load conditions. In other embodiments, the inlet heating system 42 can be active and engaged or partially engaged up to a specific percentage of full load above zero load. The percentage can be determined by design, such as the size of the engine, conditions, etc. At this point in the operation during restart, the heating up of the vane carriers 28, 50 is accelerated to increase clearances during a restart. The activation of the inlet heating system 42 allow for the respective van carriers 28, 50 or equivalent structure, e.g. ring segments 55, to be heated and avoid interference i.e. contact between adjacent rotating blade tips 32a, 56a and the boundary (e.g. structures) adjacent to the tips of the rotating blades. [0036] While specific embodiments have been described in detail, those with ordinary skill in the art will appreciate that various modifications and alternative to those details could be developed in light of the overall teachings of the disclosure. Accordingly, the particular arrangements disclosed are meant to be illustrative only and not limiting as to the scope of the invention, which is to be given the full breadth of the appended claims, and any and all equivalents thereof.