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Title:
GAS TURBINE CLEARANCE CONTROL SYSTEM INCLUDING ELECTRIC RADIANT INFRARED HEATER AND CORRESPONDING METHOD OF OPERATING A GAS TURBINE ENGINE
Document Type and Number:
WIPO Patent Application WO/2016/064389
Kind Code:
A1
Abstract:
A clearance control system includes an electrically powered radiant heater (62a, 62b, 62c, 62d) supported to a radially inner surface of an outer casing (53) for a turbine engine (10), and an absorption surface (28s, 50s) on a radially outward facing side of an outer flow path boundary (29) extending through the engine (10). A controller (92) energizes the radiant heater (62a, 62b, 62c, 62d) to produce an infrared emission that passes through a plenum (40b, 40c, 68a, 68b) and is absorbed at the absorption surface (28s, 50s) to effect a heating of the outer flow path boundary (29) to increase a clearance gap between the outer flow path boundary (29) and the tip of blades (32, 56) supported for rotation on a rotor (34) within the engine (10). A corresponding method of operating a gas turbine engine to control clearances is also provided.

Inventors:
THAM KOK-MUN (US)
LAURELLO VINCENT P (US)
LEE CHING-PANG (US)
Application Number:
PCT/US2014/061880
Publication Date:
April 28, 2016
Filing Date:
October 23, 2014
Export Citation:
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Assignee:
SIEMENS AG (DE)
SIEMENS ENERGY INC (US)
International Classes:
F01D11/24; F01D25/10; F02C7/26
Foreign References:
EP2754859A12014-07-16
EP2548974A12013-01-23
Other References:
None
Attorney, Agent or Firm:
SWANSON, Erik C. (3501 Quadrangle Blvd Ste 230Orlando, Florida, US)
Download PDF:
Claims:
CLAIMS

What is claimed is: 1 . In a gas turbine engine having a compressor, a combustor and a turbine, a clearance control system controlling a clearance between rotating blades and an outer flow path boundary adjacent to tips of the rotating blades, the clearance control system comprising:

the outer flow path boundary mounted to a radially inner side of an outer casing for the engine to define an outer boundary for gas flow through the engine; at least one plenum defined between the outer casing and the outer flow path boundary;

an electrically powered radiant heater supported to a radially inner surface of the outer casing;

an absorption surface defined on a radially outward facing side of the outer flow path boundary, the absorption surface having a direct line of sight to the radiant heater and receiving radiative heat transfer from the radiant heater; and

a controller operable to energize the radiant heater to produce an infrared emission that passes through the plenum and is absorbed at the outer flow path boundary to effect a heating of the outer flow path boundary.

2. The clearance control system of claim 1 , wherein the outer flow path boundary includes a vane carrier. 3. The clearance control system of claim 2, wherein the radiant heater has a power output of 100 kW to raise the temperature of the vane carrier by 240°C during a heating duration of one hour.

4. The clearance control system of claim 1 , wherein the radiant heater is operated at a temperature of at least about 500°C to produce the infrared emission.

5. The clearance control system of claim 1 , wherein a radiation view factor for the radiative heat transfer is 0.6 or greater.

6. The clearance control system of claim 1 , including a reflector structure located adjacent to and extending radially inward from at least one end of the radiant heater to reflect the infrared emission from the radiant heater toward the outer flow path boundary.

7. The clearance control system of claim 1 , wherein the absorption surface is a non-reflective surface.

8. The clearance control system of claim 1 , including an inlet flap located at an inlet to the compressor, the inlet flap being movable between an open position permitting flow of air into the compressor during operation of the gas turbine engine and a closed position restricting flow of air into the compressor during operation of the radiant heater.

9. The clearance control system of claim 1 , wherein the radiant heater includes plural radiant heaters positioned in both the compressor and turbine.

10. The clearance control system of claim 9, wherein the plural radiant heaters include a radiant heater located within a combustor shell of the combustor and producing an infrared emission to effect heating of a vane carrier supporting first row vanes of the turbine.

1 1 . A method of operating a gas turbine engine having a compressor, a combustor and a turbine to control a clearance between outer tips of rotating blades and an outer flow path boundary for a gas flow path in the engine, the method comprising: providing an electrically powered radiant heater supported on an inner surface of an outer casing for the engine, the radiant heater providing an infrared emission;

prior to starting the engine, energizing the radiant heater to provide radiative heat transfer to a portion of the outer flow path boundary mounted to a radially inner side of the outer casing; and

starting the engine subsequent to heating the outer flow path boundary.

12. The method of claim 1 1 , wherein providing radiative heat transfer to a portion of the outer flow path boundary includes energizing a radiative heater to heat a vane carrier in the compressor for a predetermined period of time and subsequently energizing a further radiative heater to heat a vane carrier in the turbine.

13. The method of claim 12, wherein the predetermined period of time is greater than one hour.

14. The method of claim 12, wherein at a predetermined time after the radiative heater for the turbine is energized, the radiative heater for the compressor is turned off while the radiative heater for the turbine remains energized.

15. The method of claim 1 1 , wherein providing radiative heat transfer to a portion of the outer flow path boundary includes energizing the radiative heater to heat a vane carrier in the compressor for a predetermined period of time with a bleed air line between the compressor and turbine closed and subsequently opening the bleed air line to provide a flow of warmed air from the compressor to the turbine.

16. The method of claim 1 1 , wherein the radiative heater is energized

subsequent to a turbine shut down when a temperature of a rotor supporting the blades is greater than a temperature of the outer flow path boundary.

17. The method of claim 1 1 , wherein an inlet flap located at an inlet to the compressor is moved from an open position to a close position restricting flow of air into the compressor when the radiant heater is energized. 18. The method of claim 1 1 , wherein a radiation view factor for the radiative heat transfer is 0.6 or greater.

19. The method of claim 1 1 , wherein a portion of the infrared emission from the radiant heater is reflected toward the outer flow path boundary by a reflector structure located adjacent to and extending radially inward from at least one end of the radiant heater.

20. The method of claim 1 1 , wherein the radiant heater has a power output of 100 kW and is operated at a temperature of at least about 500°C to raise the temperature of the vane carrier by 240°C during a heating duration of one hour.

Description:
GAS TURBINE CLEARANCE CONTROL SYSTEM INCLUDING ELECTRIC RADIANT INFRARED HEATER AND CORRESPONDING METHOD OF OPERATING A GAS TURBINE ENGINE

FIELD OF THE INVENTION

The present invention relates to gas turbine engines and, more particularly, 5 to a system and method for controlling clearances between stationary and rotating components in a gas turbine engine.

BACKGROUND OF THE INVENTION

A gas turbine engine generally includes a compressor section, a combustor

10 section, a turbine section and an exhaust section. In operation, the compressor

section may induct ambient air and compress it. The compressed air from the compressor section enters one or more combustors in the combustor section. The compressed air is mixed with the fuel in the combustors, and the air-fuel mixture can be burned in the combustors to form a hot working gas. The hot working gas is

15 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 the exhaust section.

The compressor and turbine sections contain several areas in which there is

20 a gap or clearance between the rotating and stationary components. During engine operation, fluid leakage through clearances in the compressor and turbine sections can contribute to system losses, making the operational efficiency of a turbine engine less than the theoretical maximum. For example, flow leakage can occur across a clearance between the tips of rotating blades and the surrounding

25 stationary structure, such as an outer shroud or a vane carrier. Small clearances

are desired to keep air leakage to a minimum; however, it is critical to maintain a 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

30 clearances can change during engine transient operation due to the difference in

the thermal inertia of the 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 significantly 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.

SUMMARY OF THE INVENTION

In accordance with an aspect of the invention, a clearance control system for a gas turbine engine is described. The engine has a compressor, a combustor and a turbine, and the clearance control system controls a clearance between rotating blades and an outer flow path boundary adjacent to tips of the rotating blades. The clearance control system comprises the outer flow path boundary mounted to a radially inner side of an outer casing for the engine to define an outer boundary for gas flow through the engine, and at least one plenum defined between the outer casing and the outer flow path boundary. An electrically powered radiant heater is supported to a radially inner surface of the outer casing, and an absorption surface is defined on a radially outward facing side of the outer flow path boundary. The absorption surface has a direct line of sight to the radiant heater and receives radiative heat transfer from the radiant heater. A controller is operable to energize the radiant heater to produce an infrared emission that passes through the plenum and is absorbed at the outer flow path boundary to effect a heating of the outer flow path boundary.

The outer flow path boundary can include a vane carrier.

The radiant heater can have a power output of 100 kW to raise the temperature of the vane carrier by 240°C during a heating duration of one hour.

The radiant heater can be operated at a temperature of at least about 500°C to produce the infrared emission.

A radiation view factor for the radiative heat transfer can be 0.6 or greater. A reflector structure can be located adjacent to and extends radially inward from at least one end of the radiant heater to reflect the infrared emission from the radiant heater toward the outer flow path boundary. The absorption surface can be a non-reflective surface.

An inlet flap can be located at an inlet to the compressor, the inlet flap being movable between an open position permitting flow of air into the compressor during operation of the gas turbine engine and a closed position restricting flow of air into the compressor during operation of the radiant heater.

The radiant heater can include plural radiant heaters positioned in both the compressor and turbine.

The plural radiant heaters can include a radiant heater located within a combustor shell of the combustor and producing an infrared emission to effect heating of a vane carrier supporting first row vanes of the turbine.

In accordance with another aspect of the invention, a method is provided for operating a gas turbine engine having a compressor, a combustor and a turbine to control a clearance between outer tips of rotating blades and an outer flow path boundary for a gas flow path in the engine. The method comprises providing an electrically powered radiant heater supported on an inner surface of an outer casing for the engine, the radiant heater providing an infrared emission; prior to starting the engine, energizing the radiant heater to provide radiative heat transfer to a portion of the outer flow path boundary mounted to a radially inner side of the outer casing; and subsequent to heating the outer flow path boundary, starting the engine.

The step of providing radiative heat transfer to a portion of the outer flow path boundary can include energizing a radiative heater to heat a vane carrier in the compressor for a predetermined period of time and subsequently energizing a further radiative heater to heat a vane carrier in the turbine.

The predetermined period of time can be greater than one hour.

At a predetermined time after the radiative heater for the turbine is energized, the radiative heater for the compressor can be turned off while the radiative heater for the turbine remains energized.

The step of providing radiative heat transfer to a portion of the outer flow path boundary can include energizing the radiative heater to heat a vane carrier in the compressor for a predetermined period of time with a bleed air line between the compressor and turbine closed and subsequently opening the bleed air line to provide a flow of warmed air from the compressor to the turbine.

The radiative heater can be energized subsequent to a turbine shut down when a temperature of a rotor supporting the blades is greater than a temperature of the outer flow path boundary.

An inlet flap located at an inlet to the compressor can be moved from an open position to a close position restricting flow of air into the compressor when the radiant heater is energized.

A radiation view factor for the radiative heat transfer can be 0.6 or greater. A portion of the infrared emission from the radiant heater can be reflected toward the outer flow path boundary by a reflector structure located adjacent to and extending radially inward from at least one end of the radiant heater.

The radiant heater can have a power output of 100 kW and can be operated at a temperature of at least about 500°C to raise the temperature of the vane carrier by 240°C during a heating duration of one hour.

BRIEF DESCRIPTION OF THE DRAWINGS

While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the

accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein:

Fig. 1 is an elevational cross-section view of a gas turbine engine illustrating aspects of the present invention;

Fig. 2A is an enlarged cross-section view similar to Fig. 1 illustrating an infrared heater in a compressor section of the engine in accordance with aspects of the invention;

Fig. 2B is an enlarged cross-section view similar to Fig. 1 illustrating an infrared heater in a turbine section of the engine in accordance with aspects of the invention; Fig. 3A is a graph illustrating variation of a clearance gap in the engine without a pre-heat operation;

Fig. 3B is a graph illustrating variation of a clearance gap in the engine, as controlled through a pre-heat operation in accordance with aspects of the present invention; and

Fig. 4 is a diagrammatic illustration of gas turbine engine including a control for a clearance control system provided in accordance with aspects of the invention.

DETAILED DESCRIPTION OF THE INVENTION

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 preferred 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.

Referring to Fig. 1 , a gas turbine engine 10 is shown illustrating aspects of the present invention. The engine includes a compressor section 12, a combustor section 14 including a plurality of combustors 16 (only one shown), and a turbine section 18. It is noted that the engine 10 illustrated herein comprises an annular array of combustors 16 that are disposed about a longitudinal axis 24 of the engine 10 that defines an axial direction of the engine 10. Such a configuration is typically referred to as a "can-annular combustion system."

Referring additionally to Fig. 2, the compressor section 12 comprises an outer compressor casing 26 enclosing various compressor components including vane carriers 28 supported from an interior structure defined on an inner side of the outer casing 26. Stationary vanes 30 are supported from the vane carriers 28, and rotating blades 32 are supported on a rotor assembly 34 and are located in alternating relation to the vanes 30 to form compressor stages. The vanes 30 and blades 32 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. As is best seen in Fig. 2, the blades 32 include radially outer blade tips 32a that rotate in close proximity to 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. Further, bleed air cavities 40 are defined between at least some of the vane carriers 28 and the outer casing 26, and comprise annular plenum or cavities extending circumferentially within the outer casing 26. In the illustrated embodiment, three bleed air cavities are particularly identified as 40a, 40b and 40c, and are located at axially downstream locations within the compressor section 12. Respective bleed air passages 42a, 42b and 42c connect the bleed air cavities 40a, 40b and 40c in fluid communication with the flow path 36. The bleed air passages 42a, 42b, 42c may be defined by radially extending gaps formed between adjacent vane carriers 28 for bleeding off a portion of the compressed air from the flow path 36 into the bleed air cavities 40a, 40b, 40c.

Referring to Fig. 1 , the combustor section 14 includes a combustor shell 44 defined within a combustor casing 46 that receives compressed air from the compressor section 12, referred to 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 turbine section 18 includes vane carriers 50 supported within an outer turbine casing 52. The outer compressor casing 26, outer combustor casing 46, and outer turbine casing 52 collectively define an outer casing 53 of the engine 10.

Stationary turbine vanes 54 are supported on the vane carriers 50 and extend radially inward across the flow path 36. The vane carriers 50 additionally support outer shrouds or ring segments 55 located in an axially alternating arrangement with outer endwalls 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 disks 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 are located adjacent to inner surfaces 55a of the ring segments 55. The 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, referred to herein as a "first mode of operation".

It should be noted that the vane carriers 28 of the compressor section 12 may comprise multiple pieces, such as two semi-cylindrical halves defining a ring around the path of the blade tips 32a. Similarly, the vane carriers 50 of the turbine section 18 may comprise multiple segments defining a ring around the path of the blade tips 56a. In addition, aspects of the invention may be applied to any structure that can comprise the vane carriers 28, 50 or equivalent structure that either defines or supports an outer boundary forming a static structure located in close proximity to the tips of rotating blades 32, 56 extending in the flow path 36. For example, it may also be understood that the present invention is not intended to be limited by the particular terminology used to describe the illustrated embodiment, in that the terminology "vane carrier" may be understood to encompass "blade segment" or "blade ring" and that such structure may be incorporated as a support for "ring segments", "shrouds", "shroud segments", and similar structure.

The diameter of the vane carriers 28, 50 and the length of the 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, e.g., the ring segments 55. However, as is described in greater detail below, the gap between the blade tips 32a, 56a and the static vane carrier 28, 50 can increase during transient operation due to the vane carrier temperature increasing.

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 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. During an engine startup, that does not incorporate a pre-heat as described below, the 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 the 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. 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".

In accordance with an aspect of the invention, one or more portions of the structure defining the outer boundary 29 of the flow path 36 may be heated by a source of radiant heat energy to quickly adjust the temperature of the outer boundary 29 and avoid interference, i.e., contact, with adjacent rotating blade tips 32a, 56a. Referring to Fig. 2, an auxiliary heating unit 60 can be provided for supplying heat to the radially outer boundary 29 wherein the heat supplied from the auxiliary heating unit 60 can be supplied independently of the operating state of the engine 10.

The auxiliary heating unit 60 comprises one or more infrared heaters, illustrated as first and second compressor section heaters 62a, 62b and first and second turbine section heaters 62c, 60d (Fig. 1 ) and are collectively referred to herein as "infrared heaters 62". Further, it should be understood that, although the particular heaters 62a, 62b, 62c, 62d are described for the presently illustrated embodiment, the invention is not limited to this number of heaters and more or fewer heaters may be provided. Referring to Fig. 2A, the infrared heaters 62 can be supported to an interior surface of the respective outer casings 26, 52 of the compressor section 12 and turbine section 18. In particular, the first compressor section heater 62a can be located on the interior surface of the outer casing 26 and faces a radially outward facing absorption surface 28 s of the vane carrier 28 within the plenum 40b, and the second compressor section heater 62b can be located on the interior surface of the outer casing 26 and faces the absorption surface 28 s on the vane carrier 28 within the plenum 40c.

Referring to Fig. 2B, the first turbine section heater 62c can be supported on an interior surface of the combustor casing 46 and faces a radially outward facing absorption surface 50 s of the vane carrier 50 within the combustor shell 44 adjacent to a first stage of the turbine section 18. The second turbine section heater 62d can be supported on an interior surface of the turbine casing 52 and faces the

absorption surface 50 s of the vane carrier 50 within a cooling air plenum 68a adjacent to a second stage of the turbine section 18.

The infrared heaters 62 are provided in spaced relation to the surface of the radially outer boundary 29 of the flow path 36 that is located to receive heat energy directly from a respective one of the infrared heaters 62. That is, the vane carrier surfaces 28 s are located spaced along a direct line-of-sight from the first and second compressor section heaters 62a, 62b within the respective bleed air cavities 40b, 40c, and the vane carrier surfaces 50 s are located along a direct line-of-sight from the first and second turbine section heaters 62c, 62d within the combustor shell 44 and the cooling air cavity 68a, respectively.

The infrared heaters 62 are preferably electrically powered radiant heaters configured to produce a sufficiently high heat output in the form of radiated energy to quickly heat the vane carriers 28, 50, such as to maintain a minimum clearance gap during a warm restart of the engine 10. The amount of radiated heat energy absorbed by the vane carriers 28, 50 is dependent on the following factors: a) the amount of power radiated by the infrared heaters 62, b) the radiation view factor, i.e., the percentage of heat energy radiated by the infrared heaters 62 that is actually received by the vane carriers 28, 50, and c) the emissivity of the vane carriers 28, 50.

In general, the higher the temperature of the infrared heaters, the greater the amount of energy that is radiated in a given time period. The Stefan-Boltzman equation predicts that the amount of power radiated by a blackbody surface will be proportional to the absolute temperature of the surface to the fourth power, such that small increases in temperature can result in exponentially larger increases in radiated power. In accordance with an aspect of the present invention, the infrared heaters 62 can be operated at a temperature of about 500°C or greater. By "about 500°C" it is meant that the temperature of the infrared heaters 62 could vary below 500°C, such as by 50°C; however, it is believed that optimal operation of the infrared heaters 62 to heat the vane carriers 28, 50 can be obtained at temperatures of 500°C and above.

The radiation view factor accounts for the geometric relationship for the radiative heat transfer between the infrared heaters 62 and the respective

absorption surfaces 28 s , 50 s . It should be noted that the infrared heaters 62 at each location in the engine 10 can extend circumferentially around the outer boundary 29 and has an axial dimension in the direction of the longitudinal axis 24 of the engine 10. Hence, the geometric relationship of the infrared heaters 62 to the respective absorption surfaces 28 s , 50 s can be generally described by a view factor between two finite concentric cylinders. It may be understood that the view factor between two finite concentric cylinders can be determined from known geometric

relationships based on the radii of the absorption surfaces 28 s , 50 s , the radii defined by the infrared heaters 62 at the respective locations on the interior surface of the engine casing 53, and the length of the absorption surfaces 28 s , 50 s and of the infrared heaters 62. For example, in the present exemplary embodiment it is believed that a radiation view factor of between 0.6 and 0.8 can be provided , with it being preferred to provide a view factor as close as possible to 0.8 for all locations of the infrared heaters 62.

In accordance with an aspect of the invention, the radiative view factor can be improved or optimized by providing a reflector structure, as is illustrated in Fig. 2B by a reflector structure 70 including a reflective surface 70s, such as may be provided by a highly polished metal surface. The reflector structure 70 extends circumferentially and is shown located adjacent to and extending radially inward from an axial end of the first turbine section heater 62c. The reflector structure 70 is oriented to direct incident infrared light waves radially inward toward the associated vane carrier absorption surface 50 s . It may be understood that a similar reflector structure may be located at one or both ends of each of the infrared heaters 62 to optimize the view factor for infrared light waves that are transmitted from the infrared heaters 62 to the absorption surfaces 28 s , 50 s .

The emissivity of the absorption surfaces 28 s , 50 s can affect the rate at which the outer boundary 29 is heated by the infrared heaters 62. It may be understood that emissivity describes the ability of the absorption surfaces 28 s , 50 s to absorb radiant energy. Typically, the outer boundary 29 may be formed of a metal that forms the absorption surfaces 28 s , 50 s and which may normally have a relatively low emissivity. However, in accordance with an aspect of the invention, the emissivity of the absorption surfaces 28 s , 50 s can be improved or optimized by providing a non- reflective coating, as is depicted by non-reflective coating 72 on absorption surfaces 28s, 50s (Figs. 2A and 2B). The non-reflective coating 72 may be, for example, a black non-reflective surface treatment capable of withstanding the high

temperatures of the outer boundary 29. Further, the non-reflective coating 72 may include a rough surface treatment designed to counteract reflection and facilitate absorption of radiant energy.

The described infrared heaters 62 can operate to input a large amount of radiant heat energy to the outer boundary in a short amount of time to substantially shorten the warm-up time of the engine 10, and enable construction of the engine 10 with smaller clearance gaps to improve the efficiency of the engine. For example, the estimated power requirement for heating the turbine section vane carrier 50 is about 100 kW, which is based on the design goal of raising the turbine section vane carrier temperature by 240°C to achieve about 3 mm of thermal growth, i.e., 3 mm of movement in the radial outward direction, over a time period of one hour. Providing 3 mm thermal growth of the vane carrier 50 is expected to maintain a desired clearance gap in that the transient blade tip clearance closure during rapid engine acceleration is about 3 mm, such as may occur during a high ramp rate of 30 MW/min for a turbine.

In accordance with a further aspect of the invention, additional control of the clearance gap between the blade tips 32a, 56a and the outer boundary 29 during a warm restart can be implemented by controlling the air flow through the engine 10 during the turning gear operation. As may be understood from the following description, heating of the vane carriers 28, 56 can be performed independently of the availability of warm air or the air flow conditions in the engine 10.

Referring to Fig. 1 , an air duct system 74 is provided extending outside of the outer casing 53 of the engine 10 between the compressor section 12 and the turbine section 18. The air duct system 74 can include one or more bleed air ducts extending from the compressor section 12 to an axially downstream location on the engine 10, as is illustrated in Fig. 1 by bleed air ducts 76a, 76b, 76c. The bleed air ducts 76a, 76b, 76c extend axially between first ends 78a, 78b, 78c (Fig. 2A) connected to respective bleed air ports extending through the compressor outer casing 26 and associated with the bleed air cavities 40a, 40b, 40c. The bleed air ducts 76a, 76b, 76c include respective second ends 80a, 80b, 80c (Fig. 2B) connected to ports associated with respective turbine cooling air plenum or cavities 68c, 68b, 68a defined between the turbine casing 52 and the vane carriers 50. The air duct system 74 is operable in the first mode of operation, i.e., powered turbine engine operation, to provide cooling air from the compressor section 12 to the turbine cooling air cavities 68a, 68b, 68c.

The air duct system 74 can also include a valve structure that can comprise control valves 82a, 82b, 82c located in the bleed air ducts 76a, 76b, 76c,

respectively. The valves 82a, 82b, 82c are 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 the valves 82a, 82b, 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, 76c. The valves 82a, 82b, 82c can be operated during a non-power producing mode of operation of the engine 10, referred to herein as a "second mode of operation", as will be described further below. The positions of the valves 82a, 82b, 82c may be controlled by a controller 92, which may also comprise a controller for controlling other operations of the engine 10 including operation of the infrared heaters 62 (see Fig. 4).

Referring to Fig. 1 , the inlet 38 to the compressor 12 can be provided with an inlet flap 84. The inlet flap 84 is movable between a first, open position, depict by dotted line 84, and a second, closed position, depicted by solid line 84. The inlet flap 84 can be positioned in the first position, such as under control of the controller 92, to permit unrestricted air flow during the first mode of operation comprising steady state powered operation of the engine 10, and the inlet flap 84 can be positioned to the second position to restrict air flow during the second mode of operation. In addition, conventional variable inlet guide vanes and variable stator vanes, collectively identified by 86, can be provided at the initial stages of the compressor section 12 for controlling air flow into the engine 10, such as during part load operation, and may also be moved toward a closed position to restrict flow during the second mode of operation.

The infrared heaters 62 can be used prior to a cold start of the engine 10 to expand either or both the compressor and turbine vane carriers 28, 50 in order to ensure that an adequate clearance gap exists. The air duct system 74 and inlet flap 84 may be operated in combination with the infrared heaters 62 during a warm restart to reduce the amount of heat dissipated by convection from the vane carriers 28, 50. In particular, during a turning gear operation in the second mode of operation, such as during a warm restart when the engine is still warm after a shutdown, the inlet flap 84 can be located to its second position to restrict air flow into the compressor section 12 and thereby reduce passage of air that would otherwise cool the vane carriers 28, 50. At generally the same time, the first and second compressor section heaters 62a, 62b can be energized to heat compressor vane carriers 28. Also during this time, the turbine section heaters 62c, 62d may remain initially deactivated and the valves 82a, 82b, 82c for the air duct system 74 can be closed to prevent air from flowing to the turbine section 18 and potentially cooling the turbine vane carriers 50. Hence, the reduced air flow reduces

convective cooling and facilitates retention of the heat produced by the infrared heaters 62.

The initial time period for warming the compressor section 12 may be about one hour or more, after which at least valves 82b and 82c can be opened to permit warmed air produced in the bleed air cavities 40b and 40c by radiation from the compressor section heaters 62a, 62b to pass to the turbine cooling air cavities 68a and 68b. At this time, the turbine section heaters 62c, 62d can be energized to heat the turbine section vane carriers 50. Also at this time, or at a predetermined time thereafter, the compressor section heaters 62a, 62b may be deactivated while the turbine section heaters 62c, 62d continue to heat the turbine vane carriers 50. In this regard, it should be noted that the compression section clearance gaps are typically smaller than the clearance gaps in the turbine section 18, and initially heating the compressor vane carriers 28 provides an initial increase in the critical compressor clearance gap, while a sufficient clearance gap is maintained in the turbine section 18 to avoid interference without creating an excessive gap that could adversely affect engine efficiency during startup.

The turbine section heaters 62c, 62d may remain energized after the engine 10 is started, such as throughout any part of the transient cycle as the engine warms to a steady state temperature. It may be understood that although the turbine section heaters 62c, 62d are described as having a delayed activation and the compressor section heaters 62a, 62b are described as being deactivated when the turbine section heaters 62c, 62d are activated, the particular timing for activation and deactivation of the infrared heaters 62 may be determined by the requirements for maintaining minimum clearance gaps. An advantage of the present invention is the provision of a heater configuration capable of providing a consistently available source of high energy heat to control clearance gaps regardless of the operating status of the engine, and regardless of availability of heat energy, e.g., warmed air, from the engine.

Additionally, by providing improved control over the clearance gaps in the engine 10 during transient cycle operation, the steady state clearance gaps can be minimized. That is, the engine 10 can be designed to operate with smaller clearance gaps between the blade tips 32a, 56a and the outer boundary 29 in that less differential expansion between the outer boundary 29 and the blades 32, 56 will occur with operation of the infrared heaters 62. An improvement in the control of the clearance gap associated with the transient cycle operation of the engine 10 is illustrated in Figs. 3A and 3B. Fig. 3A depicts a variation in the clearance gap that can occur when an engine is operated in a warm restart without preheating provided by the infrared heaters, and in which it can be seen that the transient cycle results in a substantial closure of the clearance gap as the engine ramps up to full rotational speed at steady state operation; after which the clearance gap increases, resulting in a less than optimal or desired clearance gap at steady state operation. Fig. 3B depicts a variation in the clearance gap that can occur with operation of the infrared heaters 62 in combination with control of the air flow through the engine during a warm restart. As seen in the graph of Fig. 3B, the clearance gap can initially increase with the heat input from the infrared heaters 62, such as during turning gear operation but prior to ignition of the turbine, and the clearance gap can gradually and monotonically close to the final steady state operation clearance gap. Hence, the present invention can provide control over the expansion of the ring carriers 28, 50 that maintains a clearance gap greater than the steady state clearance gap during engine ramp-up, and which permits a reduced clearance gap with greater engine efficiency at steady state operation.

In accordance with another aspect of the invention, the infrared heaters 62 can comprise a plurality of individually controlled heaters, such as may be controlled by controller 92, located at different circumferential positions within one or more stages of the compressor section 12 and/or turbine section 18. The infrared heaters 62 located at different circumferential positions can be controlled to selectively provide heat to particular circumferential locations within a stage, such as to heat a circumferential section of the stage to selectively increase the clearance gap at that section. For example, more heating may be applied to the top of the vane carrier 28, 50 as opposed to the bottom of the vane carrier 28, 50, such as to achieve any desired circumferentially-asymmetric tip gap clearance or to offset a circumferentially-asymmetric heat distribution within the stage and maintain an equalized clearance gap around the circumference of the vane carrier 28, 50.

In accordance with a further operational aspect of the invention, the operation of the infrared heaters 62 can be configured to meet specific transient conditions of the engine 10. For example, when the engine 10 enters a low load carbon monoxide (LLCO) mode for part load operation, the infrared heaters 62 may be operated to maintain a minimum clearance gap during the transition to LLCO. In particular, during LLCO mode, a significant amount of air is diverted away from the combustors, e.g., by increasing a bleed air flow, resulting in reduced air flow into the combustors and a higher combustion temperature to minimize CO emissions. The compressor and turbine clearance gap can tighten during the transition from standard to LLCO operation, and the infrared heaters 62 can be used to mitigate any detrimental clearance gap effects occurring during this transition. Hence, the infrared heaters 62 can used to heat the ring carriers 28, 50 independently of air flow conditions associated with any particular operation of the engine 10.

While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.