Patent Publication Number: US-2022213797-A1

Title: Turbomachine with low leakage seal assembly for combustor-turbine interface

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This Application is a continuation of, and claims the benefit of co-pending U.S. patent application Ser. No. 17/142,460 filed Jan. 6, 2021. The disclosure of the above referenced application is incorporated herein by reference. 
    
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     This invention was made with Government support under contract number FA8650-15-D-2504 awarded by the United States Air Force Research Lab. The Government has certain rights in the invention. 
    
    
     TECHNICAL FIELD 
     Embodiments of the subject matter described herein generally relate to improved sealing for efficiency and operational enhancement in turbomachinery. More particularly, embodiments of the subject matter relate to low leakage seal assemblies for a combustor-turbine interface area. 
     BACKGROUND 
     A turbomachine such as a gas turbine engine may be used to power various types of vehicles and/or systems. Gas turbine engines typically include a compressor that receives and compresses incoming gas such as air; a combustor in which the compressed gas is mixed with fuel and burned to produce high-pressure and high-velocity exhaust gas; and one or more turbines that extract energy from the exhaust gas exiting the combustor. 
     Interfaces between the combustor and the turbine may be prone to leakage 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 engine&#39;s air plenums. As a result of thermal growth, leakage paths may increase between mating components even if such components fit closely in a normal, pre-combustion state. Uncontrolled air leakage across interfaces has an impact on gas turbine engine performance and therefore, is undesirable. 
     In gas turbine engines, relatively cool air from the compressor may be routed around the combustor prior to entry into the combustion chamber and to cool components. Leaks between the combustion gases and the cool incoming air may occur through gaps between components. The need to assemble a number of individual components in fabricating the engine inherently results in potential gaps between parts, such as due to tolerance variations. In addition, operational ranges of a gas turbine may result in significant thermal expansion, increasing challenges in maintaining a tight seal at interfaces. Leakage flow at the interface of the combustor and the turbine, in particular, interacts with the flow field entering the turbine resulting in less than optimal combustor exit temperatures and temperature distributions. As designed turbine inlet temperatures are increasing to accomplish new performance objectives such as engine specific fuel consumption improvements, there is an increasing need to reduce the possibility of leakage. Therefore, an increasing need exists to provide sealing arrangements that minimize leakage across the combustor-turbine interface without compromising ease of assembly, while simultaneously addressing the differential thermal growth that occurs. 
     Accordingly, it is desirable to provide improved sealing approaches for interfaces in turbomachines. Furthermore, other desirable features and characteristics of the inventive subject matter will become apparent from the subsequent detailed description of the inventive subject matter and the appended claims, taken in conjunction with the accompanying drawings and this background of the inventive subject matter. 
     BRIEF SUMMARY 
     This summary is provided to describe select concepts in a simplified form that are further described in the Detailed Description section of this disclosure. This summary is not intended to identify key or essential features of the claimed subject matter, nor is it intended to be used as an aid in determining the scope of the claimed subject matter. 
     In a number of embodiments, a turbomachine includes a combustion chamber configured to receive air for combustion. The combustion chamber is defined by a combustion liner terminating in a seal ring that extends from the combustion liner to a terminal end. The combustion liner includes a head at the terminal end that has a pair of sealing surfaces. A turbine receives combusted gases from the combustor, and a transition liner directs the combustion gases to the turbine. The transition liner has walls on three sides of the head forming a cavity with an open end. The seal ring extends through the open end and the head is disposed in the cavity, with each of the sealing surfaces facing one of the walls. 
     In other embodiments, a turbomachine includes a compressor configured to generate compressed air. A combustion chamber receives the compressed air and is defined by a combustion liner terminating in a seal ring. The seal ring has an enlarged a head that is thicker than the combustion liner. A combustor casing surrounds the combustor and defines an air plenum around the combustor that is configured to receive the compressed air. A turbine is configured to receive combusted gases from the combustor. A transition liner directs the combustion gases generated in the combustion chamber toward the turbine. The transition liner has three walls forming a cavity with an open end. The seal ring extends through the open end, and the head nests in the cavity. The transition liner and the seal ring comprise a seal assembly that is exposed on one side to the compressed air and that is exposed on another side to the combustion gases. 
     In additional embodiments, a turbomachine includes a compressor configured to generate compressed air and a combustion chamber that receives the compressed air for combustion. The combustion chamber is defined by a combustion liner, with a seal ring extending from the combustion liner to a terminal end. The seal ring has an enlarged head that is thicker than the combustion liner. A turbine is configured to receive combusted gases from the combustor and a transition liner directs the combustion gases generated in the combustion chamber to the turbine. The transition liner has an inner wall on one side of the head, an outer wall on another side of the head, and an end wall extending between the first and second walls adjacent the terminal end. The walls form a cavity with an open end. The seal ring extends through the open end and the head nests in the cavity. Variable gaps are defined between the walls and the head. The head is disposed with a bias toward one of the inner and outer walls when the turbomachine is assembled. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The present invention will hereinafter be described in conjunction with the following drawing figures, wherein like numerals denote like elements, and wherein: 
         FIG. 1  is a schematic cross-sectional illustration of a turbomachine in the form of a gas turbine engine, according to an exemplary embodiment; 
         FIG. 2  is a fragmentary, cross-sectional illustration of part of the engine of  FIG. 1 , according to an exemplary embodiment; 
         FIG. 3  is a fragmentary, cross-sectional illustration of a seal assembly area for the engine of  FIG. 1  in a first state, according to an exemplary embodiment; 
         FIG. 4  is a fragmentary, cross-sectional illustration of the seal assembly area of  FIG. 3  in a second state, according to an exemplary embodiment; 
         FIG. 5  is a graph demonstrating gap influence on profile and pattern factors of a gas turbine engine shown as impact on temperature distribution in percent; and 
         FIG. 6  is a fragmentary, cross-sectional illustration of an alternate seal assembly area for the engine of  FIG. 2 , according to 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. As used herein, the word “exemplary” means “serving as an example, instance, or illustration.” Thus, any embodiment described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other embodiments. All of the embodiments described herein are exemplary embodiments provided to enable persons skilled in the art to make or use the invention and not to limit the scope of the invention which is defined by the claims. Furthermore, there is no intention to be bound by any expressed or implied theory presented in the preceding technical field, background, brief summary, or the following detailed description. 
     In a number of embodiments, turbomachine gap leakage is minimized by an active seal assembly that compensates for differential thermal growth and provides ease of assembly. An exemplary turbomachine may include a compressor configured to generate compressed air. A combustion chamber receives the compressed air for combustion, and is defined by a combustion liner. The combustion liner, such as at its inner combustion liner, terminates in an extending seal ring that may comprise a material that is thicker than the rest of the liner. Due to the increased thickness, the seal ring presents an enlarged head with at least two sealing surfaces. A transition liner directs the combustion gases from the combustion chamber to the turbine. The transition liner has an inner wall, an outer wall and an end wall forming an annular cavity with an open end. The head extends through the open end and nests in the cavity. During thermal expansion and contraction, the head is moveable in the cavity between the walls to maintain a minimal gap between the sealing surfaces of the head and the outer/inner walls. In a number of embodiments, sealing may also be provided between the head and the end wall. In some embodiments a compressible seal may be disposed between the head and the end wall. Providing multiple sealing points in one structural combination ensures adequate sealing engagement under all conditions between the combustor liner sealing ring and the inner transition liner. 
     In embodiments and examples described herein, applications such as turbomachines with seal assemblies may be described in association with an aircraft turboprop engine, but the disclosure is not limited in utility to such an application. In the example of a turboprop engine, the transition between the combustor and the turbine has a complex shape, and pressurized air delivered from the compressor may reside in a plenum located outside the combustor and the transition. Any uncontrolled gaps would allow leakage between the plenum and the transition. Uncontrolled air leakage from the plenum would impact gas turbine engine performance and so controlling leakage is beneficial in a turboprop engine. In addition, the embodiments disclosed herein have applicability where leakage control at interfacing components is similarly desirable. For example, various other engine environments, as well as different types of rotating or otherwise moving machinery will benefit from the features described herein. Thus, no particular feature or characteristic is constrained to an aircraft, an aircraft engine, or a turboprop engine, and the principles disclosed herein may be embodied in other vehicles, and/or in other turbomachinery or equipment. 
     A schematic, partially sectioned view of the engine assembly  20  is shown in  FIG. 1  according to an exemplary embodiment. The engine assembly  20  in general, includes an inlet section  22 , a gearbox  26 , a compressor section  28 , a combustion section  30 , a turbine section  32 , and an exhaust section  34 , all of which may be disposed within, or defined by, a cowling  24 . The compressor section  28 , the combustion section  30 , the turbine section  32 , and the exhaust section  34  may collectively be referred to as the engine core  38 . During operation, air enters the inlet section  22  from atmosphere and is directed into the compressor section  28 . The compressor section  28  may include a series of compressor impellers that increase the pressure of the air, which is then directed toward the combustion section  30 , such as by ducting (not shown). In this embodiment, the compressor section  28  includes a rotor with a two-stage axial compressor rotating about an axis  42 . In other embodiments, any number of stages or compressor types, such as axial or centrifugal, including a single stage may be employed. In the combustion section  30 , the high-pressure air from the compressor section  28  is mixed with fuel and combusted. The gases from the combusted fuel and air are then directed into the turbine section  32 . The turbine section  32  includes a rotor with a series of turbines, which may be disposed in axial flow series or in other arrangements and which also rotate about the axis  42 . The combustion gas from the combustion section  30  expands through, and rotates, the rotor of the turbine section  32 , from which power is derived. From the turbine section  32 , the air is then exhausted from the engine core  38  through the exhaust section  34  to the atmosphere. 
     In the exemplary embodiment, the rotor of the turbine section  32  is coupled to one or more shafts  36  to drive equipment in the engine assembly  20 . Specifically, the turbines may drive the rotor of the compressor section  28 . The shaft  36  may additionally be coupled to a hub  40  via the gearbox  26 . A propeller (not shown) may be mounted on the hub  40  and may also be driven by the turbine section  32 . Operation of the engine assembly  20  may be conducted over a wide range of ambient conditions and in response to a various operational demands. As described herein, exemplary embodiments ensure adequate leakage control is provided between areas that contain the compressed air and the combustion gases, including as temperature and resulting thermal expansion fluctuate. 
     Referring additionally to  FIG. 2 , illustrated is a fragmentary, cross-sectional view of an embodiment of a combustor-turbine transition area  44 , such as of the engine assembly  20  of  FIG. 1 . A reverse flow combustor  46  is included which turns combustion gas flow 180° prior to entry to the turbine section  32 . In this embodiment, the combustor  46  is generally disposed radially outward from a turbine plenum  48  of the turbine section  32  and is configured to contain the turbine rotor. The combustor  46  includes a combustion chamber  50  that is generally annular in shape and is defined by a liner assembly  52 . The combustor  46  also includes a plurality of shrouded injectors  54  for delivering fuel and air to the combustion chamber  50 . The liner assembly  52  generally includes a dome assembly  56 , an outer combustor liner  58  and an inner combustor liner  60 . The turbine plenum  48  is defined by a turbine shroud assembly  62  that surrounds the turbine rotor ( FIG. 1 ), and that channels the combustion gases therethrough. The turbine shroud assembly  62  is generally of a stepped cylindrical shape of various diameters to contain different sized blade sets and includes a turbine shroud  67  that extends from an inlet end  64  to an outlet end  66 . A shroud ring  68  is disposed at the inlet end  64  and engages an inner transition liner  70  that is disposed between the inner combustor liner  60  and the turbine shroud assembly  62 . An outer transition liner  72  extends from the outer combustor liner  58  to the shroud ring  68 . The shroud ring  68  defines an annular nozzle opening  74  through which combustion gases are channeled to the turbine section  32 . 
     The combustion chamber  50  is fluidly coupled to receive compressed air supplied from the compressor section  28  and more particularly through the injectors  54 , through the dome  56 , and in a number of embodiments, through a number of openings (not shown) in the outer and inner combustor liners  58 ,  60 . It should be appreciated that the openings may be provided at multiple locations to permit controlled flow through the liner assembly  52 , while uncontrolled flow at other locations, such as at component interfaces, is not desired. Fuel and air is supplied to the combustion chamber is ignited within the combustion chamber  50  by one or more igniters (not shown), generating combustion gas. The combustion gas flowstream  79  then flows through a transition liner passageway  76  which directs it into the turbine section  32 . 
     The combustor  46  is mounted within a combustor casing  75 , which is disposed radially outward of, and at least partially surrounds, the outer combustor liner  58 . Together, the combustor casing  75  and the outer combustor liner  58  at least partially define a compressed air passageway  80  for the flow of compressed air from the compressor section  28  to the combustor  46 . The passageway  80  delivers compressed air which pressurizes an air plenum  82  that surrounds the combustor  46  and is disposed both radially outside the outer combustor liner  58  and radially inside the inner combustor liner  60 . As such, the compressed air is delivered to the area outside the transition liner passageway  76 , including at the inner transition liner  70 . The compressed air in the air plenum  82 , in addition to being delivered to the combustion chamber  50 , may advantageously provide cooling for the components exposed to the hot combustion gases. 
     For various purposes including to aid in assembly, the liner structure is fabricated from various individual liner components requiring effective interfaces to provide a controlled space. For enhanced control of the combustion gas space, a seal assembly  78  is disposed at the interface  84  between the inner combustor liner  60  and the inner transition liner  70  to control leakage through the joint between the components. The interface  84  is positioned at a relatively tight bend of the liner structure near the area where the combustion gases are turned and directed into the turbine section  32 . The interface  84  is also located at an area where temperatures reach very high levels. The seal assembly  78  includes the terminal end  86  of the inner combustion liner  60  and includes the inner transition liner  70 . The seal assembly  78  is configured to actively compensate for thermal expansion of the components, which may occur at different rates in the individual components. For example, as a result of combustion in the combustion chamber  50 , the terminal end  86  of the inner combustor liner  60  may move in the axial direction  88  relative to the inner transition liner  70  and may move in the radial direction  90  relative to the inner transition liner  70 . As a result, differential thermal expansion between the combustor liner  60  and the inner transition liner  70  is accommodated by allowing relative movement, while the sealing properties are maintained as further described below. 
     Referring to  FIGS. 3 and 4 , fragmentary, detail cross sections of the area at the seal assembly  78  are shown. In general, the inner combustor liner  60 , the inner transition liner  70 , the shroud ring  68 , and the turbine shroud  67  come together in the area of the seal assembly  78 . As noted, uncontrolled leakage across interfaces such as at the interface  84 , impacts gas turbine engine performance. The combustor-turbine interface  84  is a very sensitive region to leakage. Component tolerances are included for these surfaces to accommodate engine assembly which results in gap variability from engine to engine. Further, during hot conditions differential thermal growth may further open gaps, increasing leakage. To control leakage at the inner combustor-turbine interface  84 , the seal assembly  78  includes features to compensate for differential thermal expansion. In other words, the seal assembly  78  actively controls leakage to avoid conditions where temperature distribution in the gas flow stream  79  entering the turbine section  32  is negatively impacted, and to maintain temperature distribution both radially and circumferentially in the flowstream  79 . 
     Uncontrolled leakage across this interface  84  would negatively impact performance and durability, and may disturb the designed flow field. In addition, uncontrolled leakage influences combustor exit temperature distribution and may affect turbine durability. Referring to  FIG. 5 , performance factor values are indicated on the vertical axis and gap size is indicated on the horizontal axis for both radial profile factors  96  and pattern factors  98  shown in a normalized fashion as impact on temperature distribution in percent. From a radial temperature distribution perspective, the chart shows that uncontrolled leakage increases lead to an increasing radial profile factor  96 , which may tend to increase temperatures midstream. The radial profile factor describes the radial temperature distribution in the flow stream  79 , with a higher factor indicating undesirable greater temperature differences. From a circumferential temperature distribution perspective, uncontrolled leakage leads to an increasing pattern factor  98  as also shown in the chart, which represents undesirable higher circumferential temperature distribution differences. Increased profile and pattern factors lead to less than optimal conditions from efficiency, operational and durability perspectives and therefore, the seal assembly  78  minimizes leakage. 
     As shown in  FIG. 3 , the inner combustor liner  60  includes an inner seal ring  100 , that, in this embodiment, extends from a combustor curl section  102  in the axial direction  88 . The inner seal ring  100  includes a segment  104  that connects a head  106  to the combustor curl section  102  and that projects in the axial direction  88 . The head  106  extends from the segment  104  to the terminal end  86 . As such, the head  106  presents a radially outward facing annular surface  108  and a radially inward facing annular surface  110 . The surfaces  108 ,  110  are sealing surfaces configured to inhibit leakage. As such, in the current embodiment the surfaces  108 ,  110  are machined during fabrication to present a precise sealing surface with good circularity. The head  106  is enlarged in thickness relative to the segment  104  to provide rigidity for fabrication processing and to maintain circularity. 
     The inner transition liner  70  includes an opening  112  that is defined by a ring-shaped base wall  114 . The base wall  114  may also be referred to as an inner wall in this embodiment due to its location. The opening  112  receives the shroud ring  68  and the base wall  114  abuts the turbine shroud  67 . A liner section  116  extends from the base wall  114  and curls radially inward around the end  118  of the shroud ring  68 . An opening  115  extends through the base wall  114  allowing cooling air from the air plenum  82  to into a cavity  117  to cool the liner section  116 , which is exposed directly to the combustion gases passing through the transition liner passageway  76 . A cavity  120  is defined by the inner transition liner  70  and presents a receptacle in which the head  106  of the inner seal ring  100  may nest. The cavity  120  is annular in shape, and includes three closed sides defined by the base wall  114  which comprises an inner wall, a radially extending end wall  122  and an axially extending outer wall  124 . The end wall  122  connects the outer wall  124  with the base wall  114 . The inner wall (base wall  114 ) and the outer wall  124  oppose each other across the cavity  120  and present sealing surfaces  126 ,  128 , respectively. The sealing surfaces  126 ,  128  may be machined during fabrication to present precise sealing surfaces with good circularity. The sealing surface  126  is disposed adjacent to, and faces, the annular surface  110  of the head  106 . The sealing surface  128  is disposed adjacent to, and faces, the annular surface  108  of the head  106 . The inner/base wall  114 , the end wall  122 , and the outer wall  124  define three sides of the cavity  120  and present an open end  129  through which the head  106  is received during assembly. 
     The engine assembly  20  is fabricated with tolerances for the dimensions of each component. With a possibility for up to a minimum-maximum tolerance range combination between a pair of mating components fabricated according to specifications, the potential for gaps exists. A larger gap leads to an increased potential for leakage. In the current embodiment, the tolerances are selected to bias the head  106  to initially be positioned closer to one of the walls, in this embodiment the base wall  114 . As the inner combustor liner  60  is subjected to heat loads, thermal expansion has an effect on position of the head  106  relative to the inner transition liner  70 . In the current embodiment, it has been determined that the head  106  has a tendency to move in a variety of ways, including radially outward relative to the inner transition liner  70 . Therefore, the head  106  is designed with tolerances biased inward for cold positioning adjacent the inner/base wall  114 . Accordingly, when cold the annular surface  110  is disposed very close to, or against, the sealing surface  126  with a minimal or zero gap  130 . Concurrently, the annular surface  108  is disposed away from the sealing surface  128  defining a slightly larger gap  132  that facilitates assembly when the parts are mated. When the combustion process generates heat, the head  106  moves relative to the inner transition liner  70  such that the gap  130  opens and the gap  132  simultaneously closes. Leakage is controlled by the two alternate engaging sealing surface pairs, wherein when the gap  132  is open, the gap  130  is closed ( FIG. 3 ) and when the gap  130  is open, the gap  132  is closed ( FIG. 4 ). Accordingly, one of the gaps  130 ,  132  may, at a given point in time, comprise a zero gap when completely closed by its respective mating surfaces. As used herein, the term gap includes a variable gap that may, at times, close but is are still referred to as a gap. It will be appreciated that in other embodiments, the head  106  and the inner transition liner  70  may respond to thermal loads in different ways. For example, the head  106  may move radially inward relative to the inner transition liner  70  when heated. In addition, the head  106  may move axially relative to the inner transition liner  70  when heated. 
     An additional element is shown in  FIG. 6  where the seal assembly  78  includes a seal  136  disposed in the cavity  120 . The seal  136  is compressed between the head  106  and the end wall  122 . In a number of embodiments, a W-seal is included as the seal  136 , which is a compression seal that relies on force created by “W” configuration to apply continuous pressure against the head  106  and the end wall  122  to maintain the sealed path in an airtight condition. In other embodiments, other seal types may be used. The seal  136  ensures substantially no leakage because it remains in contact with both the head  106  and the end wall  122  even in cases where they may move relative to one another in the axial direction  88 . Advantageously, the seal  136  is contained in the cavity  120  by being completely surrounded by the head  106 , and the walls  114 ,  122 ,  124 . 
     Through the foregoing embodiments, a low-leakage seal assembly, such as for a combustor-turbine interface, is provided that may reduce leakage by over fifty-percent compared to other designs. Gaps are actively controlled when the engine is operating by allowing differential movement, axial and radial, between the combustor inner liner seal ring and the inner transition liner. The seal assembly accommodates radial growth to ensure adequate sealing engagement at all conditions. In addition, the seal assembly design incorporates tolerances that accommodate assembly considerations for use without changes to engine assembly sequences. 
     While at least one exemplary embodiment has been presented in the foregoing detailed description of the inventive subject matter, 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 inventive subject matter 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 inventive subject matter. 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 inventive subject matter as set forth in the appended claims.