GB2102558A - Combustor or combustion turbine - Google Patents

Abstract

A combustor for a combustion turbine comprises a plurality of ring segments (30), each with an enhanced wall cooling mechanism. Each ring segment (30) comprises a shell member (38) and a skin member (40) with coolant channels (48) spiraling longitutudinally therebetween for conduction of cooling air from the ring segment exterior at one end to the segment interior at the opposite end. <IMAGE>

Description

SPECIFICATION Combustor for combustion turbine The present invention relates to combustion turbines as may be employed in a variety of uses, such as-industrial processes, electric power generation, or aircraft engines. More particularly, the present invention is directed to combustors employed in combustion turbines to heat motive gases which drive the turbine. The description hereafter details the structure of a combustion turbine used for power generation, but applies as well to any type of combustion turbine.

A combustion turbine used for electric power generation typically has some eight to sixteen combustors peripherally arranged about the turbine longitudinal axis. Combustion of fuel within each combustor heats pressurized air supplied from a compressor section of the turbine. The hot motive gas generated within each combustorflows by means of a transition duct to a turbine section of the combustion turbine. Impingement of the hot motive gases on turbine blades causes rotation of the turbine rotor and generation of power.

Continuing efforts to obtain higher power generation at greater efficiency have resulted in a continuous increase in gas operating temperatures.

Even with the use of improved stainless steel or nickel-based alloys such as Hastalloy or Inco 617, it has become increasingly difficult to provide adequate cooling of combustors so as to assure long combustor life. At present, the temperature of operating gases is typically 2300"F while the temperature operating limit for special metals used in construction of combustors is about 1500 F. It is expected that turbine gas operating temperatures will increase even further in the future.

Typical prior art combustors are presently cooled by introduction of air in the form of a thin film along an interior surface of a combustor wall. The primary purpose of the film of air is to prevent impingement of hot motive gas upon the interior surface of the combustorwall. Such interiorfilm cooling has generally been adequate. However, it is becoming less so with increasing turbine gas operating tem peraturesforthree reasons. First, depletion of the cooling film with hot motive gas tends to defeat the insulating properties of the cooling film. Second, inability of the cooling air to persist as a jet film due to the force of impingement by the hot motive gases destroys the film effect and reduces the effectiveness of the cooling technique.Finally, the use of fuels which generate a highly radiative flame increases the heat load on the combustorwall.

A second prior art combustor cooling technique involves use of a porous, laminated liner within the combustor, both to cool the liner as well as to provide a film of cooling air between the liner and the hot motive gases. The laminated liner comprises a plurality layers, each having channels formed therein, the channels of adjacent layers oriented generally perpendicular to one another. The channels of adjacent layers oriented generally perpendicularto one another. The channels of adjacent layers are interconnected by means of holes in the layers to provide flow communication through the liner to the interior of the combustor.

Transpiration cooling has also generally been adequate. However, increasing gas operating temperatures has resulted in the same type of problems for a transpiration-cooled combustor as were described above for a film-cooled combustor. In addition, transpiration cooling requires significant volumes of compressor discharge air to provide adequate cooling, depleting the air which would normally be used to drive the turbine end thereby reducing turbine efficiency.

Hence, it would be advantageous to develop a cooling apparatus which provides improved cooling efficiency for a combustor wall, thereby permitting higher gas operating temperatures and resulting in longer combustor life, and reducing the volume of cooling air drawn from the compressor section, permitting hig her turbine operating efficiency.

It is an object of the present invention to provide an improved combustor for a combustion turbine with a view to overcoming the deficiencies of the prior art.

The invention resides in a combustor for a combustion turbine comprising a plurality of ring members coupled to form an elongated chamber wherein compressed gases are heated for driving a turbine, characterized in that each of said ring members comprises an outer shell member and an inner skin member having a plurality of coolant channels defined therebetween, said coolant channels being arranged peripherally about said shell and skin members to direct coolant generally longitudinally of said ring member between said shell and skin members, said shell member having inlet holes for directing compressor discharge coolant air from the combustor exterior to and through said coolant channels, and said shell and skin members having an exit channel for discharging the coolant from said channels to a combustor internal gas flow.

The invention will become readily apparent from the following description of an exemplaryembodi- ment thereof when taken in conjunction with the accompanying drawings, in which: Figure 1 shows a longitudinal section of a combustion turbine in which a combustor embodying the invention is arranged to be cooled; Figure 2 shows an elevational view of an upper portion of the combustor along a longitudinal section thereof; Figure 3 shows a ring subassembly of the combustor in section; Figure 4 shows an upstream end view of a cross-section of the ring subassembly at a coolant inlet hole; Figure 5 shows an upstream end view of a cross-section of the ring subassembly at a point downstream of the coolant inlet hole; Figure 6 shows a top view of a cross section of the ring subassembly depicting spiral channels within a skin member;; Figure 7 shows an enlarged view of a portion of Figure 5 with an alternative configuration of channels; Figure 8 shows an alternative embodiment of the ring subassembly in section.

More particularly, there is shown in Figure 1 a combustion turbine 10 having a plurality of generally cylindrical combustors 12. Fuel is supplied to the combustors 12 through a nozzle structure 14 and air is supplied to the combustors 12 by a compressor 16 through air flow space 18 within a combustion casing 20.

Hot gaseous products of combustion are directed from each combustor 12 through a transition duct 22 where they are discharged into the annular space through which turbine blades 24,26 rotate underthe driving force of the expanding gases.

To provide high turbine operating efficiency and extended combustor operating life, a combustor is structured with enhanced wall cooling. The preferred structure for the combustor 12 is shown in Figures 2-8.

The combustor 12 (Figure 2) comprises a plurality of cylindrical ring subassemblies 30 and a conical ring subassembly 32 all joined by appropriate means such as welding. The length of a single ring subassembly 30, 32 is shown as extending between reference numbers 34 and 36. Cylindrical ring subassemblies 30 are joined end-to-end to form a cylindrical combustor assembly 12. The conical ring subassembly 32 is attached to the combustor assembly 12 at a fuel intake end of the assembly 12. Each ring subassembly 30, 32 comprises a cooling mechanism complete within itself but cooperating with the cooling mechanism of every other ring subassembly 30, 32 to provide an efficiently cooled combustor assembly 12.

Figure 3 shows a ring subassembly 30 of the combustor 12 in section. Structure of a conical subassembly 32 is substantially like that described below for a cylindrical subassembly 30. Each ring subassembly 30 comprises a cylindrical exterior shell 38 surrounding and adjoined to a cylindrical interior skin 40. The exterior shell 38 is continuous from one end of the ring subassembly 30 to the other end about the entire exterior surface of the subassembly 30 except for a plurality of coolant inlet holes 42 positioned around the circumference of the subassembly 30 at the end of each subassembly 30 nearest the fuel intake end of the combustor assembly 12.

The inlet holes 42 provide an entrance for cooling air 44 into a cross channel 46 defined by an inward facing groove in the shell 40 and extending around the circumference of the ring subassembly 30.

Figure 4 depicts the inlet holes 42 and the cross channel 46 from a different view. The cross channel 46 is in flow communication with a plurality of coolant channels 48 between the shell 38 and the skin 40. The coolant channels 48 shown in greater detail in Figures 4 through 7, extend from one end of the subassembly 30 to the opposite end.

Figure 6 depicts spiral grooves 50 which are formed in an outer surface of the skin 40 by an appropriate technique such as machining or etching.

The grooved surface of the skin 40 is attached to the inner surface of the shell 38 to define parallel coolant channels 48 which spiral longitudinally about the circumference of the subassembly 30 from the cross channel 46 to the exit point 52. The coolant channels 48 can, for example, be .03 inches wide by .03 inches depp and spaced from each other by .03 inches.

An alternative to the rectangular cross sectional structure of coolant channels 48 is depicted in Figure 7. In this arrangement, the coolant channels 48 are formed by grooving the interior surface of the shell 38. The walls 54 of the channels 48 are canted to provide a decrease in actual surface contact between the shell 38 and the skin 40 and a resultant increase in surface contact between the skin 38 and cooling air within the coolant channels 48. Canting the channel walls 54 also increases surface contact between the channel walls 54 and the cooling air.

The canted channel arrangement thus provides for more efficient transfer of heat from the interior skin and exterior shell 38 to the cooling air within the channel 48.

The spiral arrangement of coolant channels 48 is preferred over a strictly longitudinal arrangement.

Spiraling the channels 48 increases the effective length of the channels 48 without increasing the length of the ring subassembly 30. The length of a spiral channel 48 may be controlled by the angle chosen for the spiral. Spiraling the channels 48 also improves the distribution of coolant channels 48 within the subassembly 30. For example, the spiral channels 48 more effectively protect the combustor assembly 12 from a streak of hot motive gas flowing in a straight line through the combustor assembly 12 by providing the cooling effect of a plurality of cooling channels 48 as opposed to the cooling effect of a few longitudinal channels.

Control of cooling air flow within a ring subassembly 30 may be accomplished by variation of five structural features: (1) the cross-sectional dimensions of a coolant channel 48; (2) the number of coolant channels 48; (3) the angle of spiral of the channels 48; (4) the length of the ring subassembly 30; and (5) the dimensions of the inlet holes. Each ring subassembly 30 may be individually designed to achieve the cooling characteristics required for its position within the combustor assembly.

An alternative structure of the ring subassembly 30 is depicted in Figure 8. Each ring subassembly 30 comprises an exterior shell 38, an interior skin 40 and an interior thermal barrier 60 bonded to the inner surface of the skin 40. The relatively short length of the ring subassembly 30 and the nonporous nature of the skin 40 permits easy application of the thermal barrier 60 to the interior of the combustor 12. The thermal barrier 60, typically 0.015 to 0.025 inches in depth, may be any ceramic or other thermal barrier coating, such as yttriumstabilized zirconium oxide, which effectively insulates the skin 40 and shell 38 of the ring subassembly 30 from the harsh interior temperatures of the combustor 12. The thermal barrier 60 provides a feature with structural characteristics such as material type, grade and thickness which may be varied wth each subassembly to achieve desired insulating characteristics.

Claims (7)

1. A combustor for a combustion turbine comprising a plurality of ring members coupled to form an elongated chamber wherein compressed gases are heated for driving a turbine, each of said ring members comprising an outer shell member and an inner skin member having a plurality of coolant channels defined therebetween, said coolant channels being arranged peripherally about said shell and skin members to direct coolant generally longitudinally of said ring member between said shell and skin members, said shell member having inlet holes for directing compressor discharge coolant air from the combustor exterior to and through said coolant channels, and said shell and skin members having an exit channel for discharging the coolant from said channels to a combustor internal gas flow.

2. A combustor as set forth in claim 1 wherein said channels spiral around the periphery of said ring member.

3. A combustor as set forth in claim 1 or 2 wherein said inlet holes are disposed adjacent an end of said ring member in flow communication through an annular cross channel with said coolant channels, said cross channel comprising a circumferential inwardly facing groove in said shell mem bersubadjacentthe inlet holes.

4. A combustor as set forth in claim 1 or 2 wherein said coolant channels comprises a plurality of generally parallel channels spiraling about the longitudinal axis of said ring member within said combination of said shell and skin members, the channels defined by outwardly facing grooves on the surface of said skin member.

5. A combustor as set forth in claim 1,2,3 or 4 wherein said exit channel provides flow communication between said coolant channels and the combustor interior, the exit channel defined by an extension of said shell member beyond the end of said skin member leaving said coolant channels open to the combustor interior.

6. A combustor as set forth in any one of the preceding claims wherein said coolant channels are defined by inwardly facing grooves in the surface of said shell member, each such groove having walls canted to define a groove base narrower than a groove mouth.

7. A combustor as set forth in any one of the preceding claims wherein a thermal barrier member is bonded to and continuous with the inner side of the skin member.