Patent Publication Number: US-8109098-B2

Title: Combustor liner for gas turbine engine

Description:
FIELD OF THE INVENTION 
     The invention generally relates to a gas turbine engine, and more particularly to the combustor liner of such an engine. 
     BACKGROUND OF THE INVENTION 
     In gas turbine engines, air is compressed at an initial stage, then is heated in combustors, and the hot gas so produced drives a turbine that does work, including rotating the air compressor. 
     A number of existing gas turbine engine designs utilize some of the air from the air compressor to cool specific components that are in need of cooling. In some designs air is passed along a surface to provide convective cooling, and the air then continues to an intake of a combustor, and into the combustor where the oxygen of the air is utilized in the combustion reaction with fuel. This approach generally is referred to as “closed cooling.” In other designs, generally referred to as “open cooling,” air for cooling is passed into the flow of hot gases downstream of the combustion intake. In the latter cases a percentage of oxygen in such air for cooling may not be utilized in combustion, and this represents a potential inefficiency in that a percentage of the work to rotate the compressor does not supply air to the combustor intake for combustion purposes. The ultimate determination of whether it is more cost-effective to provide open cooling depends on balancing a number of factors, including expected component life cycle, and the costs of alternative cooling. 
     Combustor liners help define a passage for combusting hot gases immediately downstream of swirler assemblies in a gas turbine engine combustor. The surfaces of combustor liners are subject to direct exposure to the combustion flames in a combustor, and are among the components that need cooling in various gas turbine engine designs. An effusion type of open cooling has been utilized to cool combustor liners. This generally is depicted in  FIG. 1A , which provides a cross-sectional of a prior art combustor  100 . A predominant air flow (shown by thick arrows) passes along the outside of combustor  100  and into an intake  102  of the combustor  100 . Centrally disposed in the combustor  100  is a pilot swirler assembly  104 , and disposed circumferentially about the pilot swirler assembly  104  are a plurality of main swirler assemblies  106 . Combustion generally takes place somewhat downstream of the pilot swirler assembly  104 , designated in  FIG. 1A  as combustion zone  108 . A transversely disposed base plate  110  receives downstream ends of the main swirler assemblies  106 , and provides a physical barrier to flames that may otherwise travel upstream. An outlet  111  at the downstream end passes combusting and combusted gases to a transition (not shown, see  FIG. 3 ). 
     Surrounding the combustion zone  108  is an annular effusion liner  112 , and further outboard is a cylindrical frame  114 . Welded to the frame  114  at its downstream end is an assembly of spring clips  116 , which contacts a transition ring  120  of a transition (not shown in  FIG. 1A ). A plurality of holes (not shown) in the frame  114  allows passage of a quantity of air (shown by narrow arrows) that may pass through spaced apart effusion holes (not shown in  FIG. 1A ) in the effusion liner  112 .  FIG. 1B  provides an enlarged view of the encircled section of  FIG. 1A , in which spaced apart effusion holes  122  are depicted. The passage of air through the effusion holes  122  provides for a cooling of the effusion liner  112 . 
     Referring to  FIG. 1B , passage of air also is designed to occur along a radial gap  125  between the respective downstream ends  113  and  115  of the effusion liner  112  and the frame  114 . The gap  125  is required to accommodate axial and radial differential expansion between the effusion liner  112  and the frame  114 , and air flowing through the gap  125  also provides a cooling effect for the end of the effusion liner  112  and the frame  114 . In certain embodiments a plurality of spaced apart protrusions  116  disposed at or near the end  113  of the effusion liner  112  establish the radial height of the gap  125 . 
     Based on observation and analysis of present systems, such as that described in  FIGS. 1A and 1B , and potential problems in some units of such systems, there is a need for an improved combustor liner that overcomes such problems. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Aspects of the invention are explained in following description in view of drawings that are briefly described below: 
         FIG. 1A  is a lateral cross-sectional view of a prior art combustor comprising an effusion-type combustor liner.  FIG. 1B  provides an enlarged view of an encircled portion of the prior art combustor depicted in  FIG. 1A . 
         FIG. 2A  provides a partial lateral cross-sectional view of one embodiment of a combustor liner of the present invention, with two components attached to the combustor liner.  FIG. 2B  provides a lateral cross-sectional view of a combustor comprising the combustor liner of  FIG. 2A .  FIG. 2C  is a cross-sectional view taken along the line  2 C- 2 C of  FIG. 2B , illustrating the flow control ring, inner liner wall and spring clips. 
         FIG. 3  is a schematic lateral cross-sectional depiction of a gas turbine showing major components, in which embodiments of the present invention may be utilized. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Embodiments of the present invention provide for uniformly controlled open cooling of a double-walled combustor liner that is effective to predictably and consistently provide cooling air currents to such liners. The present invention was created as a result of first identifying potential problems with presently used liner systems in gas turbine combustors. For example, referring to  FIG. 1B , it has been appreciated that the radial gap  125  may at times allow excessive air flow and/or provide an uneven air flow, either of which are hypothesized to have the potential to lead to lower gas turbine engine performance. Factors affecting the size and non-uniformity of the gap  125  may include: 1) in-tolerance ‘mismatches’ in which respective ends  113  and  115  of the effusion liner  112  and the frame  114  are within their respective tolerances, but at extreme ends of the respective in-tolerance ranges (i.e., end  113  at lower end, end  115  at upper end); 2) thermal expansion; 3) out of round condition of the effusion liner  112  and/or the frame  114 ; and 4) a permanent set in the effusion liner  112  and/or the frame  114 , such as due to creep or plastic deformation caused by thermally induced stresses. It is appreciated that the performance of individual units may vary depending on the effect of one or more of these factors, and this may lead to variability in performance among the different combustors in a particular gas turbine engine (such as a can-annular style). In addition to such potentially adverse performance, such variability is hypothesized make less clear the diagnosis of other issues. 
     Based on such appreciation of potential air leakage and unequal passage of cooling air with existing combustor liner designs, a new liner is developed. This development is directed to overcome gap variation and consequent performance imbalances hypothesized to affect some combustor units. The new liner comprises an inner annular wall the inside surface of which is directly exposed to the combustion zone, an outer annular wall, spaced from the inner annular wall, a cooling air flow channel formed there between, and a flow control ring to which are attached the downstream ends of the inner and outer annular walls. The flow control ring comprises a plurality of holes through which cooling air from the cooling air flow channel passes. As used with regard to the flow control ring and any other component of the present invention, the term “hole” is not meant to be limited to a round aperture through a body as is illustrated in the embodiment depicted in the figures. Rather, the term “hole” is taken to mean any defined aperture through a body, including but not limited to a slit, a slot, a gap, a groove, and a scoop. The liner structure eliminates the above-described gap between prior art liner and frame ends through which, it is hypothesized, air may flow unevenly and wastefully. In contrast, the present invention comprises a cooling air flow channel in fluid communication with spaced apart holes of the flow control ring which together may provide a desired level of cooling to the inner annular wall, the flow control ring and to components downstream of the flow control ring. Further as to temperature management, in certain embodiments a portion of the inner surface of the inner annular wall comprises a Thermal Barrier Coating (“TBC”), such as a ceramic coating, that provides enhanced thermal protection to this portion. Other aspects of the invention are disclosed during and after discussion of specific embodiments provided in the appended figures. 
       FIG. 2A  depicts an exemplary embodiment of a new liner  230 . Liner  230  comprises an inner wall  232 , an outer wall  238 , a cooling air flow channel  244  formed there between, and a flow control ring  246 . The inner wall  232  of liner  230  comprises an upstream end  233 , a downstream end  234 , welded to the flow control ring  246 , an inner surface  235 , and an outer surface  236 . The outer wall  238  comprises an upstream end  239 , a downstream end  240 , also welded to flow control ring  246 , an inner surface  241 , and an outer surface  242 . The flow channel  244  is annular and has a length defined from the upstream end  239  to the downstream end  240  of outer wall  238 , and a width defined as the distance between the inner wall  232  outer surface  236  and the opposing inner surface  241  of the outer wall  238 . 
     In the depicted embodiment, a major portion, meaning more than 50 percent, of the inner surface is coated with a thermal barrier coating  237 . Other embodiments may comprise no thermal barrier coating, a total coverage with a thermal barrier coating, or a smaller percentage coverage with a thermal barrier coating. 
     The downstream end  234  of inner wall  232  is welded to an inboard region  247  of flow control ring  246 , and the downstream end  240  of outer wall  238  is welded to flow control ring  246  along an outboard region  248  of flow control ring  246 . Thus, the flow control ring  246  may generally be considered to comprise an inboard region  247  lying inboard of a central region (identified as  249  in  FIG. 2C ) that comprises a plurality of holes  250 , and an outboard region  248  disposed outboard of the central region (identified as  249  in  FIG. 2C ). In  FIG. 2A  an inboard surface  251  of the inboard region  247  is shown as coated with thermal barrier coating  237 , and on an outboard surface  252  of the outboard region  248  there is an attachment of a spring clip assembly  255 . Neither the presence of the thermal barrier coating  237 , nor the attachment of the spring clip assembly  255  to flow control ring  246 , is meant to be limiting of the scope of the present invention. 
     An opening  228  allows for air to pass from the compressor (not shown) into the cooling air flow channel  244 . A protective barrier  229  covers the opening  228 , and may be constructed of screen, mesh, or sheet metal with holes  227  there through, having sufficient open area for passage of a desired amount of cooling air into cooling air flow channel  244 . The protective barrier  229  is provided when there is a concern that errant objects flowing with the compressor air flow may become entrapped in the cooling air flow channel  244  or the holes  250  of the flow control ring  246 . It is noted that some embodiments do not comprise protective barrier  229 . In various embodiments that do comprise a protective barrier such as protective barrier  229  in  FIGS. 2A and 2B , the protective barrier may be attached to either the inner or to the outer wall, that is, to at least one of the inner and the outer wall. Attachment to only one of the two walls allows differential movement of the two walls as a function of different thermal expansion of these two walls. Further, as one example of an alternative to the protective barrier  229 , the upstream end  239  of outer wall  238  may be bent downward, toward the outer surface  236  of inner wall  232 , and may have any types of holes through it, and/or grooves or cuts, etc. at its edge, that are of a desired size, so as to provide a variant of a protective barrier across the upstream end  239  of flow channel  244 . 
     The separation between the inner wall  232  and the outer wall  238  may be established by any spacing means (not shown) as is known to those skilled in the art. Structures generally known “stand-offs” may be provided at spaced intervals to establish a desired space between the inner wall  232  and outer wall  238 . One example of a stand-off, not to be limiting, is a rod of a desired length, having a broad head, that is inserted into a first wall so that the non-headed end of the rod contacts the inside surface of the opposing wall. While in such position the broad head is welded to the outside of the first wall. This provides a minimum distance between the walls. 
     While not meant to be limiting of the scope of the present invention, in the embodiment depicted in  FIG. 2A  a barrier structure  260  is attached, such as by welding, to the outside surface  242  of outer wall  238 . The barrier structure  260  limits movement of broken-off spring clips (not shown in  FIG. 2A ), and is described in greater detail in U.S. patent application Ser. No. 11/117,051, which is incorporated by reference herein for such teachings. More generally, this and all other patents, patent applications, patent publications, and other publications referenced herein are hereby incorporated by reference in this application in order to more fully describe the state of the art to which the present invention pertains, to provide such teachings as are generally known to those skilled in the art, and to provide specific teachings as may be noted herein. 
       FIG. 2B  depicts a combustor  200  in cross-section, comprising the liner  230  of  FIG. 2A . In addition to the liner  230 , combustor  200  comprises standard combustor components that include an intake  202 , a centrally disposed pilot fuel swirler assembly  204 , a plurality of main swirler assemblies  206 , a base plate  210 , and an outlet  211 . A combustion zone is indicated by  208 . 
     It is noted that for embodiment depicted in  FIGS. 2A and 2B , no component corresponds exactly to the cylindrical frame  114  in  FIG. 1A . As an alternative, the liner  230  may be constructed of sufficiently strong material to support the spring clip assembly  255  and forces transmitted through this structure. For example, not meant to be limiting, the thickness of the inner wall  232  may be 0.090 inches, rather than a more commonly used 0.060 inches thickness. As viewable in  FIG. 2B , the upstream end  233  of the inner wall  232  is shown welded to a curved section of base plate  210 . This provides for structural integrity and transfer of forces between the spring clip assembly  255  and the combustor  200 . However, this arrangement is not meant to be limiting. 
     Further to the thermal barrier coating  237 , as depicted in  FIGS. 2A and 2B , the thermal barrier coating  237  covers not only a major portion of the inner surface  235  of the inner wall  232 , but also covers most of the inboard surface  251  of the flow control ring  246 . A thermal barrier coating such as  237  may be comprised of any suitable composition recognized to provide an effective thermal barrier in the operating temperature range of the combustion zone  208 . A ceramic coating may be used, for example. This would be applied over the surface of the material of the inner wall  232  after suitable surface preparation. It is noted that the composition of the inner wall  232 , the outer wall  238 , and the flow control ring  246  may be a nickel-chromium-iron-molybdenum alloy (e.g. HASTELLOY® X alloy), an alloy known to those skilled in the art of gas turbine engine construction. Other metal alloys known to those skilled in the art, or other non-metallic materials, may alternatively be utilized. 
     Also, although not depicted in  FIG. 2A , a thermal barrier coating (such as  237 ) may be applied not only to the inner surface  235  of the inner wall  232 , and to the inboard surface  251  of the inboard region  247  of the flow control ring  246 , but also may be applied to cover the outboard surface  252  of the outboard region  248 , and the exposed downstream surfaces of the flow control ring  246  that are between the inboard surface  251  and the outboard surface  252 . 
       FIG. 2C  provides an upstream view from line  2 C- 2 C of  FIG. 2B , and depicts the inner wall  232  coated with thermal barrier coating  237 , the flow control ring  246 , and the spring clip assembly  255 . The flow control ring  246  is seen to be viewed as comprising the central region  249  that comprises a plurality of holes  250 , the inboard region  247  lying inboard of a central region  249 , and the outboard region  248  disposed outboard of the central region  249 . These regions are not meant to indicate that the flow control ring is comprised of three separate components annealed together; a typical method of construction is to form a unitary annular body and machine it to comprise desired features, such as the holes  250 . In various embodiments, the inboard region  247  and the outboard region  248  comprise respective weld preps (indicated as  253  and  254  in  FIG. 2A ) that may provide for stronger weld bonds with the adjoining regions of the inner wall  232  and the outer wall  238 . 
     In the embodiment depicted in  FIGS. 2A-2C , a cooling air flow supplied by the gas turbine engine compressor (not shown in these figures, see  FIG. 3 ) enters the flow channel  244  at the upstream end  239  of the outer wall  238  passing through the optional protective barrier  229 . The cooling air then travels toward the through the holes  250  of the flow control ring  246 . This flow of cooling air through the holes  250  is effective to control the cooling air flow, to provide convective cooling along the inner wall  232 , and to provide convective cooling of the flow control ring  246 . By control, as that term is used herein with regard to the holes  250  is not an active form of control. Rather the control of cooling air flow is a function of a predetermined cross-sectional flow area that does not change in order to effectuate the desired control. The predetermined cross-sectional flow area, and the size, shape, and distribution of holes  250  in a flow control ring  246  are determined as a function of the calculated or modeled flow to achieve a desired level of cooling under varying operating conditions, and may vary from embodiment to embodiment depending on factors that include the presence of a thermal barrier coating on the inner wall  232 , and the presence of optional effusion holes through the inner wall  232 . 
     Further, because the holes  250  of flow control ring  246  provide the only defined exits for such cooling air flow, when embodiments such as that depicted in  FIGS. 2A-2C  are installed in a plurality of combustors in a gas turbine engine, these embodiments are effective to provide a uniformly controlled open cooling of the combustor liner walls. This uniformity contrasts with the less controllable prior art embodiments that may be subject to the aforementioned sources of variability. It is appreciated that this provision of a uniformly controlled open cooling, or alternatively, the property of being effective to control a particular cooling air flow, is based on a passive control, related in part to the size, number and distribution of holes in a flow control ring (or more generally in a flow control regulator), rather than to an ‘active’ type of control. 
     The more general term ‘flow control regulator’ includes flow control rings such as described above, and a flow control regulator also may comprise a plurality of arcuate segments which together comprise an annular shape. However, a flow control regulator need not be annular shaped, nor an annular ring structure, and may be comprised of spacers (which may include weld beads) that are spaced apart to connect inner and outer liner walls proximate a combustor outlet, so that gaps, such as slits, between the spacers are the spaces through which a controlled cooling air flow flows. 
     Also, the plurality of holes in a flow control ring in embodiments such as that depicted in  FIGS. 2A-2C  may be effective to cool, to a determined maximum temperature, the inner wall without the use of effusion holes through the inner wall. However, this is not meant to be limiting. For instance, embodiments may comprise a combustor liner comprising an outer wall and an inner wall defining there between a flow channel, and a flow control ring sealingly connected to the inner and the outer walls proximate the combustor outlet, wherein spaced along the inner wall are a number of effusion holes that provide a supplemental flow of cooling air at desired locations along the inner wall. Such effusion holes are effective to supplement the cooling of an inner wall. A number of such optional effusion holes  270  are depicted in  FIG. 2B . Generally, these may be placed at appropriate locations along the inner wall  232  to achieve a desired supplemental cooling effect. 
     Additionally, the flow of cooling air entering the transition (not shown in  FIGS. 2A and 2B ) may cool the adjacent transition interior walls, an upstream portion  260  of which is depicted in  FIG. 2B . This may occur by providing a uniform and spaced flow of cooling air through the holes  250 . It is noted that the cooling air exiting the holes  250  are in fluid communication with the combustion zone  208 , albeit the holes  250  literally provide air into the transition at the juncture of the combustion zone  208  and the transition (not shown, see  FIG. 3 ). Also, although the inner wall  232  and the outer wall  238  are depicted in  FIGS. 2A and 2B  as parallel, this is not meant to be limiting. For instance, the spacing between an inner wall and an outer wall may decrease (or may increase) from upstream to downstream ends of a flow channel formed between such walls. 
     Embodiments of the present invention are used in gas turbine engines such as are represented by  FIG. 3 , which is a schematic lateral cross-sectional depiction of a prior art gas turbine  300  showing major components. Gas turbine engine  300  comprises a compressor  302  at a leading edge  303 , a turbine  310  at a trailing edge  311  connected by shaft  312  to compressor  302 , and a mid-frame section  305  disposed therebetween. The mid-frame section  305 , defined in part by a casing  307  that encloses a plenum  306 , comprises within the plenum  306  a combustor  308  (such as a can-annular combustor) and a transition  309 . During operation, in axial flow series, compressor  302  takes in air and provides compressed air to an annular diffuser  304 , which passes the compressed air to the plenum  306  through which the compressed air passes to the combustion chamber  308 , which mixes the compressed air with fuel (not shown), providing combusted gases via the transition  309  to the turbine  310 , whose rotation may be used to generate electricity. It is appreciated that the plenum  306  is an annular chamber that may hold a plurality of circumferentially spaced apart combustors  308 , each associated with a downstream transition  309 . Likewise the annular diffuser  304 , which connects to but is not part of the mid-frame section  305 , extends annularly about the shaft  312 . Embodiments of the present invention may be incorporated into each combustor (such as  308 ) of a gas turbine engine to provide a more uniform and controlled open cooling of the combustor liner walls. 
     Although the above embodiments provide for an outer wall that is distinguished from cylindrical frame  114  of  FIG. 1 , in some embodiments of the present invention the outer wall may comprise a cylindrical frame. For a gas turbine engine comprising a cylindrical frame as its outer wall, or another type of outer wall, it is appreciated that in a combustor in that engine, comprising an inner wall and such outer wall disposed about a combustion zone, a flow control ring comprising a plurality of holes may be attached to downstream ends of these walls. This would provide an alternative embodiment of the present invention that is effective to regulate and assure more uniformity in cooling fluid flow in this structure. 
     While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.