Abstract:
A gas turbine engine assembly includes a combustor configured to combust an air-fuel mixture to produce combustion gases in a first direction; a transition liner coupled to the combustor and adapted to receive the combustion gases from the combustor and to redirect the combustion gases in a second direction; and a turbine coupled to the transition liner and adapted to receive the combustion gases from the transition liner. The transition liner has a plurality of effusion holes that include a first group that extend at least partially in a tangential direction.

Description:
TECHNICAL FIELD 
       [0001]    The present invention generally relates to gas turbine engines, and more particularly relates to systems and methods for cooling gas turbine engine transition liners. 
       BACKGROUND 
       [0002]    A gas turbine engine may be used to power aircraft or various other types of vehicles and systems. The engine typically includes a compressor that receives and compresses incoming gas such as air; a combustion chamber in which the compressed gas is mixed with fuel and burned to produce exhaust gas; and one or more turbines that extract energy from the high-pressure, high-velocity exhaust gas exiting the combustion chamber. 
         [0003]    The arrangement and configuration of these sections impact many characteristics of the gas turbine engine, including overall engine length and weight, as well as the materials to construct the turbine engine. The overall length of the turbine engine may be shortened, thereby saving on materials, weight and length, by the use of a reverse flow annular combustion chamber. This type of combustion chamber is so named because the general direction of flow within and out of the chamber is opposite to the general direction of air flow that subsequently enters the turbine. Typically, a transition liner is fitted to the downstream portion of the annular combustion chamber and serves to redirect the flow of combustion gas into the turbine section, thereby resulting in a gas flow aligned with the turbine and the general direction of overall flow through the engine. The transition liner is typically configured as an annular ring with a concave hot side facing the combustion chamber. 
         [0004]    The engine is subject to extreme temperatures, particularly at the transition liner that receives and redirects the combustion products. The high temperatures may cause thermal stresses and other problems. Conventional system and methods for cooling the transition liner, such as heat shields, louvers and impingement cooling, have been met with mixed success, at best. 
         [0005]    Accordingly, it is desirable to provide improved systems and methods for cooling the transition liner. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention. 
       BRIEF SUMMARY 
       [0006]    In accordance with one exemplary embodiment, a gas turbine engine assembly includes a combustor configured to combust an air-fuel mixture to produce combustion gases in a first direction; a transition liner coupled to the combustor and adapted to receive the combustion gases from the combustor and to redirect the combustion gases in a second direction; and a turbine coupled to the transition liner and adapted to receive the combustion gases from the transition liner. The transition liner has a plurality of effusion holes that include a first group that extend at least partially in a tangential direction. 
         [0007]    In accordance with another exemplary embodiment, a reverse-flow combustor assembly includes a first liner; and a second liner circumscribed by the first liner to form a combustion chamber therebetween. The combustion chamber is configured to combust an air-fuel mixture to produce combustion gases exiting form the combustion chamber in a first direction. A transition liner is coupled to the combustion chamber and configured to receive the exiting combustion gases and to redirect the combustion gases in a second direction, generally opposite to the first direction. The transition liner has a plurality of effusion holes that include a first group that extend at least partially in a tangential direction. 
         [0008]    In accordance with yet another exemplary embodiment, a gas turbine engine assembly includes a combustor configured to combust an air-fuel mixture to produce combustion gases in a first direction; a transition liner coupled to the combustor and adapted to receive the combustion gases from the combustor and redirect the combustion gases in a second direction; and a turbine coupled to the transition liner and adapted to receive the combustion gases from the transition liner. The transition liner defines a plurality of effusion holes that include a first group having a tangential orientation and with a compound angle, a second group downstream of the first group and having a radial orientation, and a third group radially between the first and second group and transitioning in orientation between approximately tangential and approximately radial. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0009]    The present invention will hereinafter be described in conjunction with the following drawing figures, wherein like numerals denote like elements, and 
           [0010]      FIG. 1  is a cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment; 
           [0011]      FIG. 2  is a more detailed cross-sectional view of a portion of the engine of  FIG. 1 ; 
           [0012]      FIG. 3  is an isometric view of an exemplary transition liner; 
           [0013]      FIG. 4  is a partial plan view of the transition liner in a tangential-radial plane; 
           [0014]      FIG. 5  is a partial cross-sectional view of the transition liner in a radial-axial plane; and 
           [0015]      FIG. 6  is a partial cross-sectional view of the transition liner in an axial-tangential plane. 
       
    
    
     DETAILED DESCRIPTION 
       [0016]    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. 
         [0017]    Broadly, the exemplary embodiments discussed herein provide cooling schemes for transition liners in gas turbine engines. More particularly, the transition liners are provided with effusion holes for supplying a continuous film of cooling air to the liner surface. The effusion holes can have at least one group oriented in a tangential direction and one group in a radial direction, and may additionally include effusion holes with compound angles. Embodiments discussed herein may find beneficial use in many industries and applications, including aerospace, automotive, and electricity generation. 
         [0018]      FIG. 1  is a cross-sectional view of an engine  100  in accordance with an exemplary embodiment. In one embodiment, the engine  100  is a multi-spool gas turbine main propulsion engine. The engine  100  includes an intake section  110 , a compressor section  120 , a combustion section  130 , a turbine section  140 , and an exhaust section  150 . 
         [0019]    The intake section  110  includes a fan that draws air into the engine  100  and accelerates it into the compressor section  120 . The compressor section  120  may include one or more compressors that raise the pressure of the air directed into it, and directs the compressed air into the combustion section  130 . In the depicted embodiment, a two-stage compressor is shown, although it will be appreciated that one or more additional compressors could be used. 
         [0020]    The combustion section  130 , which is discussed in greater detail below, includes a combustor unit  160  that mixes the compressed air with fuel supplied from a fuel source (not shown). The fuel/air mixture is combusted to generate high energy combusted gas that is then directed into the turbine section  140 . The combustor unit  160  may be implemented as any one of numerous types of combustor units. However, as will be discussed in more detail further below in one embodiment, the combustor unit  160  is implemented as a reverse flow combustor unit. The turbine section  140  includes one or more turbines in which the combusted gas from the combustion section  130  expands and causes the turbines to rotate. The gas is then exhausted through the exhaust section  150 . 
         [0021]      FIG. 2  is a more detailed cross-sectional view of a portion of the engine  100  of  FIG. 1 , and particularly illustrates section  2  in  FIG. 1 . The reverse flow combustor unit  160  and portions of the turbine section  140  are depicted and will now be described in more detail. 
         [0022]    In one embodiment, the combustor unit  160  includes an outer liner  202  and an inner liner  204  circumscribed by the outer liner  202 . The outer liner  202  and the inner liner  204  form a combustion chamber  206  therebetween. In the combustion chamber  206 , the compressed air is mixed with fuel and combusted to generate combustion gas  208 . 
         [0023]    The combustion gas  208  then flows into a transition liner  210 , which receives the combustion gas  208  and diverts it in an opposite direction into the turbine section  140 . More specifically, in this particular engine  100 , the direction of the combustion gas flow  208  is reversed with respect to the overall orientation of the turbine engine  100 . The combustion gas  208  is directed from an upstream section  212  of the transition liner  210  to a downstream section  214  of the transition liner  210 . In the depicted embodiment, the upstream section  212  of the transition liner  210  is coupled to the outer liner  202 , while the downstream section  214  is coupled to the turbine section  140 . 
         [0024]    As also shown in  FIG. 2  and discussed in greater detail below, the transition liner  210  has a number of effusion holes  216  to permit compressed air to pass through for cooling the interior surface of the transition liner  210 . In particular, the effusion holes  216  allow a buffering layer  218  of cool air to pass from the exterior surface to the interior surface of the transition liner  210 , and then in a generally downstream direction with the hot combustion gasses  208  from the combustion chamber  206  into the turbine section  140 . This layer  218  of cooler air reduces the direct contact of the hot combustion gasses  208  with interior surface of transition liner  210  as well as convectively cools the wall of the transition liner  210  as the air passes through the holes  216 . 
         [0025]    Unlike the prior art systems and methods that require heat shields and/or louvers, the effusion holes  216  may simplify cooling in that no additional components need be attached to the transition liner in some embodiment; such components may be provided in addition, however, in embodiments where more cooling is desired. Manufacturing costs may be reduced due to a decrease in part count and an overall simplified design. The durability of the transition liner  210  may be extended by a reduction in temperature gradients along the surface. 
         [0026]      FIG. 3  is an isometric view of the transition liner  210  removed from the engine  100 . As noted above, the transition liner  210  is configured as ring with a concave hot surface  302  that faces the combustion chamber  206  ( FIG. 2 ). The transition liner  210  includes a first edge  304  adjacent the upstream section  212  and a second edge  306  adjacent the downstream section  214 . 
         [0027]    Characteristics of the transition liner  210  can be considered in three dimensions, as indicted by the legend  350  and discussed further in  FIGS. 4-6 . A radial direction  308  extends between the first edge  304  and the second edge  306  along the surface of the transition liner  210 , i.e. radially inward and outward within the circular configuration of the transition liner  210 . The radial direction  308  also corresponds to the downstream direction of the combustion gases during operation from the first edge  304  to the second edge  306 . An axial direction  310  extends outwardly from the surface of the transition liner  210 . A tangential direction  312  extends around the surface of the transition liner  210  and around the center axis  314 . Although  FIG. 3  shows a representation of the effusion holes  216 , the orientation and arrangement are discussed in greater detail with reference to  FIGS. 4-6 . Particularly, as discussed in  FIGS. 4-6 , the effusion holes  216  can be oriented in one or more of the radial direction  308 , axial direction  310 , and tangential direction  312  to result in a compound angle for at least some of the effusion holes  216 , thereby resulting in improved cooling characteristics. 
         [0028]      FIG. 4  is a partial plan view of the transition liner in a tangential-radial plane and more clearly show the effusion holes  216 . As discussed above, the effusion holes  216  are generally relatively small, closely spaced holes serving to direct a flow of cooling air onto the walls of transition liner  210 . The cooling holes are generally 0.01 to 0.04 inches in diameter, although the diameter may vary with application and may depend on factors such as the dimensions of the transition liner  210 , the temperature of the combustion gases  208  ( FIG. 2 ), and the velocity of the cooling flow  218  ( FIG. 2 ). Individual hole shape is generally cylindrical or oval, with minor deviations due to manufacturing method i.e. edge rounding, tapers, out-of-round or oblong, etc. Other embodiments could use holes shaped other than circular or oval. 
         [0029]    The effusion holes  216  can be patterned to improve cooling. Particularly, the effusion holes  216  can be arranged in groups  402 ,  404 ,  406  having tangentially staggered rows. A first group  402  of effusion holes is adjacent the first edge  304  of the transition liner  210  and has an orientation that is approximately completely tangential. In other words, the first group  402  of effusion holes has angles of approximately 0° relative to a tangential axis and can direct cooling air around the transition liner  210 , as indicated by arrow  452 . In one embodiment, the first group  402  has between 2 and 10 rows of effusion holes. 
         [0030]    A second group  404  of effusion holes is downstream of the first group  402  and transitions between an approximately tangential direction and an approximately radial direction, i.e., between approximately 0° relative to a tangential axis to approximately 90° relative to a tangential axis, as indicated by arrow  454 . In one embodiment, the second group  404  has between 2 and 10 rows of effusion holes. 
         [0031]    A third group  406  of effusion holes is downstream of the second group  404  and adjacent the second edge  306 . The third group  406  has an orientation that is approximately completely radial. In other words, the third group  406  of effusion holes  216  has angles approximately 90° relative to the tangential axis, and can direct cooling air downstream the transition liner  210 , as indicated by arrow  456 . In one embodiment, the third group  406  has between 2 and 10 rows of effusion holes. 
         [0032]    As a result of this arrangement, the cooling air passing through the effusion holes  216  of the first and second groups  402 ,  404  have at least some tangential component. As such, the cooling air may linger for a longer period of time on the surface of the transition liner  210  to provide improved cooling. However, the effusion holes  216  the second and third groups  404 ,  406  transition from a more tangential direction  452  to a radial direction  456  such that the cooling air is transitioned into the direction of the combustion gases exiting the combustor and entering the turbine. As a result, the cooling air does not interfere with the aerodynamics of the combustion gases. 
         [0033]      FIG. 5  is a partial cross-sectional view of the transition liner  210  in a radial-axial plane. In this embodiment, the angle of the effusion holes  216  in the radial-axial plane can vary from about 35° to about 60° relative to an axial axis. Generally, the effusion holes  216  have a greater angle adjacent the first edge  304  and transition to a lesser angle adjacent the second edge  306 . 
         [0034]      FIG. 6  is a partial cross-sectional view of the transition liner  210  in an axial-tangential plane. The angle  602  of these effusion holes  216  can vary from about 0° to about 65° from a tangential axis in the axial-tangential plane. In one embodiment, the angle along the axial-tangential plane can increase from the second edge  306  (e.g.,  FIG. 3 ) to the first edge  304  (e.g.  FIG. 3 ). As with the other components of the compound angle of the effusion holes  216 , the angle  602  in the axial-tangential plane increases the length of the hole  216  through the transition liner  210 , thereby increasing the surface area from which the cooling flow can extract heat from the transition liner  210 . 
         [0035]    The effusion holes  216  may be formed by drilling techniques such as electrical-discharge machining (EDM), stationary percussion laser machining and percussion on-the-fly laser drilling or with complex casting techniques. The density of the effusion holes  216  may vary with application and may depend on factors including the dimensions of the transition liner  210 , the material of manufacture of the transition liner  210 , the velocity of the cooling flow, and the temperature of the combustion gases. For some applications, the effusion holes  216  may be uniformly spaced. Alternatively, the effusion holes  216  may be unevenly spaced to provide more cooling flow to “hot spots” on the transition liner  210 . 
         [0036]    While at least one exemplary embodiment has been presented in the foregoing detailed description of the invention, 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.