Patent Publication Number: US-8534076-B2

Title: Combustor-turbine seal interface for gas turbine engine

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     This invention was made with Government support under Contract No. W911W6-08-2-0001 awarded by U.S. Army. The Government has certain rights in this invention. 
    
    
     TECHNICAL FIELD 
     The present invention relates generally to gas turbine engines and, more particularly, to a combustor-turbine seal interface having improved leakage, cooling, and compliancy characteristics. 
     BACKGROUND 
     A generalized gas turbine engine (GTE) includes an intake section, a compressor section, a combustion section, a turbine section, and an exhaust section disposed in axial flow series. The compressor section includes one or more compressor stages, and the turbine section includes one or more air turbine stages each joined to a different compressor stage via a rotatable shaft or spool. During operation, the compressor stages rotate to compress air received from the intake section of the GTE. A first portion of the compressed air is directed into an annular combustor mounted within the combustion section, and a second portion of the air is directed through cooling flow passages that flow over and around the combustor. Within the combustion chamber, the compressed air is mixed with fuel and ignited. The air heats rapidly and exits each combustor chamber via an outlet provided through the combustor&#39;s downstream end. The air is received by at least one turbine nozzle, which is sealingly coupled to the combustor&#39;s downstream end. The turbine nozzle directs the air through the air turbines to drive the rotation of the air turbines, as well as the rotation of the spools and compressor stages coupled thereto. Finally, the air is expelled from the GTE&#39;s exhaust section. The power output of the GTE may be utilized in a variety of different manners, depending upon whether the GTE assumes the form of a turbofan, turboprop, turboshaft, or turbojet engine. 
     The sealing interface between the turbine nozzle and the combustor preferably maximizes the operational lifespan of the GTE while simultaneously minimizing leakage between the turbine nozzle and the combustor. It has, however, proven difficult to design a durable, low leakage combustor-turbine seal interface largely due to the extreme thermal gradients that result from temperature fluctuations in the air exhausted from the combustor, as well as the temperature differentials between the air exhausted from the combustor and the cooler air bypassing the conductor. Such thermal gradients cause thermal distortion and relative movement between the various components of the combustor-turbine seal interface; e.g., between the liner walls and the turbine nozzle, which become relatively hot during combustion, and the engine casing, which remains relatively cool during combustion and which may be fabricated from a low thermal growth material, such as a titanium-based alloy. As a result of thermal distortion, leakage paths may form between mating components even if such components fit closely in a non-distorted, pre-combustion state. Compression seals (e.g., metallic W-seals) may be employed to minimize the formation of such leakage paths; however, such compression seals may also be heated to undesirably high temperatures by the hot air exhausted from the combustor, and the sealing characteristics and strength of the compliant seals can be compromised. Furthermore, if the components of the combustor-turbine seal interface are unable to adequately accommodate such thermal distortion, the combustor-turbine seal interface may experience relatively rapid thermomechanical fatigue and decreases in performance. The GTE may consequently require premature removal from service and repair, resulting in economic loss due to the non-availability of the GTE, as well as direct maintenance costs. 
     There thus exists an ongoing need to provide a combustor-turbine seal interface that significantly reduces or eliminates leakage between a combustor and a turbine nozzle (or nozzles). Ideally, embodiments of such a combustor-turbine seal interface would include one or more compliant structures that accommodate relative movement between the combustor, the turbine nozzle, and the engine casing to reduce thermomechanical fatigue and increase operational lifespan of combustor-turbine seal interface. It would also be desirable for embodiments of such a combustor-turbine seal interface to promote efficient cooling of the combustor and, perhaps, of the leading edge portion of the turbine nozzle. Lastly, it would be desirable for embodiments of the combustor-turbine seal interface to provide aerodynamically efficient flow paths for the heated air exhausted from the combustor, as well as for the cooler air bypassing the combustor. Other desirable features and characteristics of the present invention will become apparent from the subsequent Detailed Description and the appended Claims, taken in conjunction with the accompanying Drawings and this Background. 
     BRIEF SUMMARY 
     A combustor-turbine seal interface is provided for deployment within a gas turbine engine. In one embodiment, the combustor-turbine seal interface comprising combustor, a turbine nozzle downstream of the combustor, and a first compliant dual seal assembly. The first compliant dual seal assembly includes a compliant seal wall sealingly coupled between the combustor and the turbine nozzle, a first compression seal sealingly disposed between the compliant seal wall and the turbine nozzle, and a first bearing seal generally defined by the compliant seal wall and the turbine nozzle. The first bearing seal is sealingly disposed in series with the first compression seal. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       At least one example of the present invention will hereinafter be described in conjunction with the following figures, wherein like numerals denote like elements, and: 
         FIG. 1  is a generalized cross-sectional view of the upper portion of an exemplary gas turbine engine; 
         FIG. 2  is a generalized cross-sectional view of an exemplary combustor-turbine seal interface deployed within the gas turbine engine shown in  FIG. 1 ; 
         FIG. 3  is a cross-sectional view illustrating a first compliant dual seal assembly employed by the combustor-turbine seal interface shown in  FIG. 2  in accordance with an exemplary embodiment; 
         FIG. 4  is an isometric cross-sectional view of an upper portion of the compliant seal wall and the seal retainer included within the first compliant dual seal assembly shown in  FIG. 3 ; and 
         FIG. 5  is a cross-sectional view illustrating a second compliant dual seal assembly employed by the combustor-turbine seal interface shown in  FIG. 2  in accordance with an exemplary embodiment. 
     
    
    
     DETAILED DESCRIPTION 
     The following Detailed Description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. Furthermore, there is no intention to be bound by any theory presented in the preceding Background or the following Detailed Description. 
       FIG. 1  is a generalized cross-sectional view of the upper portion of an exemplary gas turbine engine (GTE)  20 . In the exemplary embodiment illustrated in  FIG. 1 , GTE  20  assumes the form of a three spool turbofan engine including an intake section  24 , a compressor section  26 , a combustion section  28 , a turbine section  30 , and an exhaust section  32 . Intake section  24  includes a fan  34 , which may be mounted within an outer fan case  36  Compressor section  26  includes an intermediate pressure (IP) compressor  38  and a high pressure (HP) compressor  40 ; and turbine section  30  includes an HP turbine  42 , an IP turbine  44 , and a low pressure (LP) turbine  46 . IP compressor  38 , HP compressor  40 , HP turbine  42 , IP turbine  44 , and LP turbine  46  are disposed within a main engine casing  48  in axial flow series. HP compressor  40  and HP turbine  42  are mounted on opposing ends of an HP shaft or spool  50 ; IP compressor  38  and IP turbine  44  are mounted on opposing ends of an IP spool  52 ; and fan  34  and LP turbine  46  are mounted on opposing ends of a LP spool  54 . LP spool  54 , IP spool  52 , and HP spool  50  are substantially co-axial. That is, LP spool  54  extends through a longitudinal channel provided through IP spool  52 , and IP spool  52  extends through a longitudinal channel provided through HP spool  50 . Combustion section  28  and turbine section  30  further include an annular combustor  56  and an annular turbine nozzle  58 , which sealingly mates with annular combustor  56  as described more fully below. 
     As illustrated in  FIG. 1  and described herein, GTE  20  is offered by way of example only. It will be readily appreciated that embodiments of the present invention are equally applicable to various other types of gas turbine engine including, but not limited to, other types of turbofan, turboprop, turboshaft, and turbojet engines. Furthermore, the particular structure of GTE  20  will inevitably vary amongst different embodiments. For example, in certain embodiments, an open rotor configuration may be employed wherein fan  34  is not mounted within an outer fan case. In other embodiments, the GTE may employ radially disposed (centrifugal) compressors instead of axial compressors. In still further embodiments, GTE  20  may not include a single, annular turbine nozzle and may instead include a number of turbine nozzles, which are circumferentially arranged around the longitudinal axis of GTE  20  (represented in  FIG. 1  by dashed line  55 ) and each sealingly coupled to annular combustor  56 . 
       FIG. 2  is a generalized cross-sectional view of combustion section  28  and turbine nozzle  58  illustrating combustor-turbine seal interface  60  in accordance with an exemplary embodiment. Combustor  56  is mounted within a cavity  59  provided within engine casing  48 . Combustor  56  includes an inner liner wall  61  and an outer liner wall  63 . Inner liner wall  61  and outer liner wall  63  each have a generally conical shape and collectively define an annular combustion chamber  64  within combustor  56 . As is conventionally known, liner walls  61  and  63  may be formed from a temperature-resistant material (e.g., a ceramic, a metal, or an alloy, such as a nickel-based super alloy), and the interior of liner walls  61  and  63  may each be coated with a thermal barrier coating (TBC) material, such as a friable grade insulation. Additionally, a number of small apertures  65  may be formed through liner walls  61  and  63  (e.g., via a laser drilling process) for effusion cooling or aerodynamic purposes (only two effusion cooling apertures  65  are shown in  FIG. 2  and exaggerated for clarity). 
     Combustor  56  further includes a combustor dome inlet  66  and a combustor outlet  68  formed through leading and trailing end portions of combustor  56 , respectively. Combustor dome inlet  66  and effusion apertures  65  fluidly couple cavity  59  to combustion chamber  64 , and combustor outlet  68  fluidly couples combustion chamber  64  to turbine nozzle  58 . A combustor dome shroud  70  is mounted to liner wall  61  and to liner wall  63  proximate the leading end portion of combustion chamber  64  and partially encloses combustor dome inlet  66 . A carburetor assembly  72  is mounted within combustion chamber  64  proximate the leading end portion of combustor  56 . Carburetor assembly  72  receives the distal end of a fuel injector  74 , which extends radially inward from an outer portion of engine casing  48  as generally shown in  FIG. 2 . 
     A diffuser  78  is mounted within engine casing  48  upstream of combustor  56 . During operation of GTE  20  ( FIG. 1 ), diffuser  78  directs compressed air received from compressor section  26  ( FIG. 1 ) into cavity  59 . A portion of the compressed air supplied by diffuser  78  flows through combustor dome shroud  70  and into carburetor assembly  72 . Carburetor assembly  72  mixes this air with fuel and air received from fuel injector  74  and introduces the resulting fuel-air mixture into combustion chamber  64 . Within combustion chamber  64 , the fuel-air mixture is ignited by an igniter  76  mounted through liner wall  63 . The air heats rapidly, exits combustion chamber  64  via outlet  66 , and flows into turbine nozzle  58 . Turbine nozzle  58  then directs the air through the sequential series of air turbines mounted within turbine section  30  (i.e., turbines  42 ,  44 , and  46  shown in  FIG. 1 ) to drive the rotation of the air turbines and, therefore, the rotation of the fan and compressor stages mechanically coupled thereto. In the embodiments wherein GTE  20  assumes the form of a turbojet, the air is subsequently exhausted (e.g., via a nozzle  80  provided in exhaust section  32  shown in  FIG. 1 ) to produce forward thrust. 
     A certain volume of the air supplied by diffuser  78  into cavity  59  is directed over and around combustor  56 . As indicated in  FIG. 2  by arrows  82 , a first portion of this air flows along a first cooling flow path  84  generally defined by outer portion of liner wall  63  and an inner portion of engine casing  48 . Similarly, as indicated in  FIG. 2  by arrows  86 , a second portion of the compressed air flows along a second cooling path  88  generally defined by an inner portion of liner wall  61  and an internal mounting structure  90  provided within engine casing  48 . The air flowing along cooling flow paths  84  and  88  is considerably cooler than the air exhausted from combustion chamber  64 . Airflow along cooling flow paths  84  and  88  is utilized to convectively cool combustor  56 , turbine nozzle  58 , and the other components of combustor-turbine seal interface  60 . With respect to combustor  56 , in particular, airflow along cooling flow paths  84  and  88  may convectively cool the exterior of liner walls  61  and  63  through direct convection. Furthermore, in embodiments wherein liner walls  61  and  63  are provided with effusion apertures  65 , the air conducted along cooling flow paths  84  and  88  may also cool liner walls  61  and  63  via convection cooling through effusion apertures  65 . Effusion apertures  65  may also help create a cool barrier air film along the inner surface of liner walls  61  and  63  defining combustion chamber  64 . The combustion process (through radiation heat transfer) and flow of exhaust from combustor  56  (through convection), in concert with airflow along cooling flow paths  84  and  88 , results in thermal gradients between the various components of combustor-turbine seal interface  60 . Due to such thermal gradients, turbine nozzle  58 , liner wall  61 , and liner wall  63  will typically become relatively hot during combustion, while engine casing  48  and other surrounding components remain relatively cool. 
     As a point of emphasis, embodiments of the combustor-turbine seal interface employ at least one compliant dual seal assembly to sealingly couple the combustor to the turbine nozzle (or nozzles). In the exemplary embodiment illustrated in  FIG. 2 , combustor  56  is sealingly coupled to turbine nozzle  58  utilizing two compliant dual seal assemblies, namely, a first compliant dual seal assembly  92  and a second compliant dual seal assembly  94 . First and second compliant dual seal assemblies  92  and  94  are each sealingly coupled between a downstream or trailing end portion of combustor  56  and an upstream or leading end portion of turbine nozzle  58 . In addition, first compliant dual seal assembly  92  is coupled between an outer portion of liner wall  63  and an outer portion of turbine nozzle  58 ; and second compliant dual seal assembly  94  is coupled between an inner portion of liner wall  61  and an inner portion of turbine nozzle  58 . First compliant dual seal assembly  92  resides further from the longitudinal axis of GTE  20  ( FIG. 1 ) than does second compliant dual seal assembly  94 . 
       FIG. 3  is a cross-sectional view illustrating first compliant dual seal assembly  92  in greater detail. In the example shown in  FIG. 3 , compliant dual seal assembly  92  includes four main components: (i) a compliant seal wall  96 , (ii) a seal retainer  98 , (iii) a compression seal  100 , and (iv) a bearing seal  124 . Compliant seal wall  96  and seal retainer  98  are also shown in  FIG. 4  in isometric cross-section. As can be most easily appreciated in  FIG. 4 , seal retainer  98  comprises a generally annular body  102  having a plurality of axially-elongated flanges  104  extending therefrom in a downstream direction. Axially-elongated flanges  104  are radially spaced to define a plurality of airflow channels  105  ( FIG. 4 ) through seal retainer  98 . Airflow channels  105  are radially interspersed between axially-elongated flanges  104  and permit airflow through seal retainer  98 , and therefore around first compliant dual seal assembly  92 , as indicated in  FIG. 3  by arrows  82 . Airflow channels  105  also increase the flexibility of seal retainer  98  along axially-elongated flanges  104  and, consequently, permit seal retainer  98  to better accommodate thermal displacement that may occur between the various components of seal assembly  92  and engine casing  48  as described more fully below. As shown most clearly in  FIG. 3 , each flange  104  may be mounted to engine casing  48  utilizing, for example, a bolt  106 , a rivet, or other fastener (only one flange  104  and one bolt  106  is shown in  FIG. 3  for clarity). When mounted to engine casing  48  in this manner, generally annular body  102  engages a first nozzle wall  111  (e.g., a radial flange) projecting from the main body of turbine nozzle  58  to physically capture turbine nozzle  58  and help maintain the radial position thereof. 
     With continued reference to  FIGS. 3 and 4 , an annulus  108  is provided within generally annular body  102  and receives compression seal  100  therein. When compliant dual seal assembly  92  is assembled, compression seal  100  is sealingly compressed between an inner surface of seal retainer  98  and first nozzle wall  111 . When sealingly compressed in this manner, compression seal  100  eliminates or minimizes leakage between combustor  56  and turbine nozzle  58 . In the illustrated example, compression seal  100  assumes the form of a metallic W-seal; however, in alternative embodiments, compression seal  100  may assume various other geometries (e.g., that of a C-seal, a V-seal, various other convolute seals, or an elastic gasket configuration) and may be formed from other suitable materials. In addition to carrying compression seal  100 , seal retainer  98  also serves as a pilot to ensure precise radial alignment between the various components of combustor-turbine seal interface  60 . First nozzle wall  111  may be directly affixed to or integrally formed with the main body of turbine nozzle  58 . In embodiments wherein turbine nozzle  58  comprises a plurality of circumferentially-spaced turbine nozzles or turbine nozzle segments, each turbine nozzle may be individually mounted to first nozzle wall  111  utilizing bolts, rivets, or other mechanical fastening means. 
     With continued reference to  FIG. 3 , compliant seal wall  96  has a first end portion  116 , a second end portion  118  substantially opposite first end portion  116 , and an axially-overlapping intermediate portion  120  between first end portion  116  and second end portion  118 . First end portion  116  of compliant seal wall  96  is fixedly coupled to seal retainer  98 , and second end portion  118  of compliant seal wall  96  is fixedly coupled to a downstream end portion of combustor  56 . In one embodiment, first end portion  116  is fabricated from sheet metal and/or machined from a forging and subsequently brazed or welded (e.g., e-beam structure welded, seam welded, etc.) to an outer circumferential portion of seal retainer  98 . Second end portion  118  of compliant seal wall  96  may also be formed as a separate piece and subsequently affixed (e.g., brazed or welded) to intermediate portion  118  of compliant seal wall  96  and to a downstream end portion of combustor  56 . In a preferred group of embodiments, axially-overlapping intermediate portion  120  has a generally conical geometry that accommodates the conical shape of combustor  56  while providing radial and axial compliancy as described more fully below. 
     Second end portion  118  of compliant seal wall  96  abuts turbine nozzle  58 , and specifically a leading edge portion  122  of turbine nozzle  58 , to form a bearing seal  124  between combustor  56  and turbine nozzle  58 . As may be appreciated by referring to  FIG. 4 , compliant seal wall  96  is a substantially solid structure sealingly coupled between seal retainer  98  and liner wall  63  of combustor  56 . Compliant seal wall  96  thus serves to generally prevent airflow from bypassing compression seal  100 . As may be appreciated by referring to  FIG. 3 , compression seal  100  and bearing seal  124  are coupled in flow series and, in combination with compliant seal wall  96 , significantly reduce leakage between combustor  56  and turbine nozzle  58 . This, in turn, improves the overall efficiency of GTE  20  ( FIG. 1 ). Additionally, the air saved from minimizing leakage between combustor  56  and turbine nozzle  58  can be utilized to cool the combustor or turbine components and/or utilized to tailor combustor aerodynamics. Although not shown in  FIG. 3  for clarity, an aperture may be provided in a lower portion of complaint seal wall  96  (e.g., the bottom dead center of GTE  20 ) to allow residual fuel to drain from the cavity formed by compliant seal  96  and seal retainer  98 . 
     Although compression seal  100  and bearing seal  124  significantly reduce the development of leakage paths between combustor  56  and turbine nozzle  58 , a minimal amount of leakage may still occur between combustor  56  and turbine nozzle  58 . If a leakage path should develop, leakage will generally flow from the exterior of combustor  56  and turbine nozzle  58  into the interior of combustor  56  and turbine nozzle  58  (indicated in  FIG. 3  by leakage arrow  126 ). For this reason, it may be stated that compression seal  100  resides upstream of bearing seal  124  as taken along a combustor leakage path. In the illustrated exemplary embodiment, bearing seal  124  generally resides between compression seal  100  and outlet  68  of combustor  56 . 
     As shown most clearly in  FIG. 3 , the outer portion of liner wall  63  and compliant seal wall  96  (in particular, the innermost segment of axially-overlapping intermediate portion  120 ) are radially spaced apart along their lengths. Collectively, compliant seal wall  96  and liner wall  63  define an effusion cooling path  128  along an outer surface of combustor  56  that extends to the downstream end of combustor  56 . As indicated in  FIG. 3  by arrows  130 , the effusion cooling path  128  permits the cooler air flowing along cooling flow path  84  (also indicated by arrows  82  in  FIG. 3 ) to flow substantially unimpeded over the downstream end of combustor  56 . Thus, in contrast to certain known combustor-turbine sealing interfaces that block or restrict airflow to the downstream exterior of the combustor, compliant dual seal assembly  92  permits the entire body of combustor  56  to be effusively cooled. 
     To provide improved cooling of turbine nozzle  58 , one or more cooling channels may be provided through second end portion of compliant seal wall  96  to direct a cooling jet against the leading portion of turbine nozzle  58  as shown in  FIG. 3  at  132 . Furthermore, as indicated in  FIG. 3  at  113 , the innermost circumferential edge of seal retainer  98  is radially offset from the neighboring portion of compliant seal wall  96 . This radial offset or gap permits liner wall  63 , which becomes relatively hot during combustion, to grow radially outward relative to compliant seal wall  96 , which remains relatively cool during combustion. In a preferred embodiment, the radial clearance between seal retainer  98  and compliant seal wall  96  is such that complaint seal wall  96  seats on seal retainer  98  prior to the outlet of cooling channel  132  being obstructed by leading portion  122  of turbine nozzle  58 . Stated differently, the innermost edge of generally annular body  104  of seal retainer  98  serves as a hard stop that physically prevents complaint seal wall  96  from growing radially outward to a positional extreme wherein cooling channel  132  is obstructed by the leading edge of turbine nozzle  58 . In certain embodiments, second end portion  118  of seal wall  96  may not directly contact seal retainer  98  to provide a hard stop; instead, second end portion  118  may be formed to include one or more projections (e.g., a raised bump) that abut seal retainer  98  to provide a hard stop that prevents the obstruction of cooling channel  132 . 
     In contrast to certain known combustor-turbine seal interfaces, combustor-turbine seal interface  60  is designed such that compression seal  100  is radially offset or spaced apart from the outlet of combustor  56 . This radial offset results in an improved thermal isolation of compression seal  100  from the heated air exhausted from combustor  56  and the leading edge portion  122  of turbine nozzle  58 , which becomes relatively hot during combustion. Excessive heating of compression seal  100  is thus avoided, and the sealing characteristics and structural integrity of compression seal  100  are maintained during operation of GTE  20  ( FIG. 1 ). 
     As previously noted, compliant seal wall  96 , and specifically axially-overlapping intermediate portion  120 , provides a radial compliance between the hot downstream end portion of combustor  56  and the cooler seal retainer  98 . This radial compliance permits compliant seal wall  96  to flex radially and thereby accommodate relative movement between combustor  56  and seal retainer  98 . Furthermore, bearing seal  124  permits turbine nozzle  58  to slide radially relative second end portion  118  of compliant seal wall  96  while generally maintaining an airtight seal. Compliant seal wall  96  and bearing seal  124  thus cooperate to permit compliant dual seal assembly  92  to accommodate relative movement between the various components of combustor-turbine seal interface  60  that may occur as a result of thermal deflection. In this manner, thermomechanical fatigue within combustor-turbine seal interface  60  is reduced, and the operational lifespan of interface  60  is increased. Compliant seal wall  96  also provides an axial compliancy between engine casing  48  and the core components of GTE  20  ( FIG. 1 ), which further helps to accommodate relative movement and to maintain a substantially constant axial load through compression seal  100  and bearing seal  124  to maintain the sealing characteristics thereof. Similarly, axially-elongated flanges  104  of seal retainer  98  provide a radial compliance between the main body of seal retainer  98 , which undergoes considerable thermal expansion during combustion, and engine casing  48 , which experiences relatively limited thermal expansion during combustion, and which maybe formed from a low thermal growth material, such as a titanium-based alloy. This again results in a reduction in thermomechanical stress, and an increase in operational lifespan. 
       FIG. 5  is a cross-sectional view illustrating second compliant dual seal assembly  94  in greater detail. Second compliant dual seal assembly  94  includes an outer beam structure  138  and an inner beam structure  134 . The downstream end of outer beam structure  138  is fixedly coupled (e.g., welded or brazed) to liner wall  61  of combustor  56 . The downstream end portion of inner beam structure  134  abuts and is captured by a radial lip  146  provided around turbine nozzle  58 . Outer beam structure  138  axially overlaps with inner beam structure  134  to form a radial spring member that provides radial compliance between combustor  56  and internal mounting structure  90 . Outer beam structure  138  is retained by a flange  142 , which may be mounted to internal mounting structure  90  utilizing, for example, a plurality of bolts  147  (only one of which is shown in  FIG. 5 ), rivets, or other such fasteners. Collectively, beam structures  138  and  142  provided a radial compliance to accommodate relative movement that may occur between combustor  56  and structure  90  during combustion. In so doing, beam structures  138  and  142  minimizes mechanical stressors within second compliant dual seal assembly  94  and thereby increase the operational lifespan of GTE  20  ( FIG. 1 ). 
     As was the case with first compliant dual seal assembly  92 , second compliant dual seal assembly  94  includes a compression seal  136  and a bearing seal  144 . Compression seal  136  (e.g., a metallic W-seal) is sealingly compressed between the upstream end portion of outer beam structure  138  and the upstream end portion of inner beam structure  134  (e.g., a radial flange), which is attached to turbine nozzle  58 . Bearing seal  144  is generally defined by the downstream end of outer beam structure  138  and the leading edge portion of turbine nozzle  58 . Bearing seal  144  and compression seal  136  are coupled in series, and bearing seal  144  generally resides between compression seal  136  and the downstream outlet of combustor  56 . Bearing seal  144  and compression seal  136  cooperate to significantly reduce or eliminate leakage between combustor  56  and turbine nozzle  58  and thereby improve the efficiency of GTE  20  ( FIG. 1 ). Notably, beam structures  138  and  134  position compression seal  136  at a location that is axially offset from the leading edge portion of turbine nozzle  58 , which becomes relatively hot during combustion. By offsetting compression seal  136  from turbine nozzle  58  in this manner, compression seal  136  may be maintained in a cooler state and the sealing characteristics of compression seal  136  may be better preserved during operation of GTE  20  ( FIG. 1 ). 
     One or more cooling channels  148  may be provided through the downstream end portion of outer beam structure  138  to form cooling jets that cool turbine nozzle  58  during operation of GTE  20 . More specifically, cooling channels  148  direct the relatively cool air flowing between liner wall  61  and outer beam structure  138  (represented in  FIG. 5  by arrow  86 ) against the leading edge portion of turbine nozzle  58  to convectively cool turbine nozzle  58 . As further shown in  FIG. 5 , a radial gap  150  may be provided between the downstream end of outer beam structure  138  and the downstream end of inner beam structure  134 . Radial gap  150  generally accommodates the transient inward growth of liner wall  61  and outer beam structure  138  relative to inner beam structure  134 . Inner beam structure  134  may cool more slowly during a deceleration transient than liner wall  61  and outer beam structure  138 , which would result in an interference unless gap  150  is provided. At the same time, the radial width of radial gap  150  is preferably such that outer beam structure  138  contacts inner beam structure  134  as the leading edge portion of turbine nozzle  58  and flange  146  grow radially outward, to provide a hard stop before cooling channels  148  are obstructed by the leading edge portion of turbine nozzle  58 . 
     The foregoing has thus provided an exemplary embodiment of a combustor-turbine nozzle-case assembly that significantly reduces or eliminates leakage between the combustor and the turbine nozzle. In foregoing example, the combustor-turbine nozzle-case assembly employed at least one compliant dual seal assembly having a radial compliance that accommodates relative movement between the combustor, the turbine nozzle, and the engine casing to reduce thermomechanical fatigue and thus increase operational lifespan of combustor-turbine seal interface. It should be appreciated that, in the above-described exemplary embodiment, the combustor-turbine seal interface promoted efficient cooling of the combustor and the leading edge portion of the turbine nozzle. It should also be appreciated that the above-described combustor-turbine seal interface provided aerodynamically efficient flow paths for the heated air exhausted from the combustor and for the cooler air bypassing the combustor. 
     While at least one exemplary embodiment has been presented in the foregoing Detailed Description, it should be appreciated that a vast number of variations exist. It should also be appreciated that the exemplary embodiment or exemplary embodiments are only examples, and are not intended to limit the scope, applicability, or configuration of the invention in any way. Rather, the foregoing Detailed Description will provide those skilled in the art with a convenient road map for implementing an exemplary embodiment of the invention. It being understood that various changes may be made in the function and arrangement of elements described in an exemplary embodiment without departing from the scope of the invention as set-forth in the appended Claims.