Patent Publication Number: US-10774661-B2

Title: Shroud for a turbine engine

Description:
BACKGROUND OF THE INVENTION 
     Turbine engines, and particularly gas or combustion turbine engines, are rotary engines that extract energy from a flow of pressurized combusted gases passing through the engine onto a multitude of rotating and stationary turbine airfoils. The stationary turbine airfoils can be supported by shrouds that are interlocked to form a circumferential casing to the turbine. 
     BRIEF DESCRIPTION OF THE INVENTION 
     In one aspect, a shroud for a gas turbine engine comprises at least two shroud elements forming a ring and having confronting radial ends that define a split interface with axial fore and aft portions, where the fore portion defines a fore split surface interface forming a positive radial angle relative to a radial line, and the aft portion defines an aft split surface interface forming a negative radial angle relative to the radial line. 
     In another aspect, a shroud for a gas turbine engine comprises at least two shroud elements forming a ring and having confronting radial ends that define a split interface, where the radial ends have first complementary structures that impede relative radial movement of the at least two shroud elements, second complementary structures that impede relative axial movement of the at least two shroud elements, and third complementary structures that impede relative circumferential movement of the at least two shroud elements. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       In the drawings: 
         FIG. 1  is a schematic cross-sectional diagram of a turbine engine for an aircraft. 
         FIG. 2  illustrates a multi-element shroud in the turbine engine of  FIG. 1  viewed along the axial centerline of the engine. 
         FIG. 3  is a perspective view of a portion of the shroud in  FIG. 2  illustrating the interface between two of the shroud elements. 
         FIG. 4  is a perspective view of a portion of a first shroud element of the shroud in  FIG. 2 . 
         FIG. 5  is a perspective view of a portion of a second shroud element of the shroud in  FIG. 2 . 
         FIG. 6  is a circumferential view of the first shroud element of  FIG. 4 . 
         FIGS. 7A-7F  show various top views of the shroud in  FIG. 2 . 
     
    
    
     DESCRIPTION OF EMBODIMENTS OF THE INVENTION 
     The described embodiments of the present invention are directed to a shroud assembly for stationary airfoils. For purposes of illustration, the present invention will be described with respect to the turbine for an aircraft turbine engine. It will be understood, however, that the invention is not so limited and may have general applicability within an engine, including compressors, as well as in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications. 
     As used herein, the term “forward” or “upstream” refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component. The term “aft” or “downstream” used in conjunction with “forward” or “upstream” refers to a direction toward the rear or outlet of the engine or being relatively closer to the engine outlet as compared to another component. 
     Additionally, as used herein, the terms “radial” or “radially” refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference. 
     All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, aft, etc.) are only used for identification purposes to aid the reader&#39;s understanding of the present invention, and do not create limitations, particularly as to the position, orientation, or use of the invention. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and can include intermediate members between a collection of elements and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and in fixed relation to one another. The exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the drawings attached hereto can vary. 
       FIG. 1  is a schematic cross-sectional diagram of a gas turbine engine  10  for an aircraft. The engine  10  has a generally longitudinally extending axis or centerline  12  extending forward  14  to aft  16 . The engine  10  includes, in downstream serial flow relationship, a fan section  18  including a fan  20 , a compressor section  22  including a booster or low pressure (LP) compressor  24  and a high pressure (HP) compressor  26 , a combustion section  28  including a combustor  30 , a turbine section  32  including a HP turbine  34 , and a LP turbine  36 , and an exhaust section  38 . 
     The fan section  18  includes a fan casing  40  surrounding the fan  20 . The fan  20  includes a plurality of fan blades  42  disposed radially about the centerline  12 . The HP compressor  26 , the combustor  30 , and the HP turbine  34  form a core  44  of the engine  10 , which generates combustion gases. The core  44  is surrounded by core casing  46 , which can be coupled with the fan casing  40 . 
     A HP shaft or spool  48  disposed coaxially about the centerline  12  of the engine  10  drivingly connects the HP turbine  34  to the HP compressor  26 . A LP shaft or spool  50 , which is disposed coaxially about the centerline  12  of the engine  10  within the larger diameter annular HP spool  48 , drivingly connects the LP turbine  36  to the LP compressor  24  and fan  20 . The spools  48 ,  50  are rotatable about the engine centerline and couple to a plurality of rotatable elements, which can collectively define a rotor  51 . 
     The LP compressor  24  and the HP compressor  26  respectively include a plurality of compressor stages  52 ,  54 , in which a set of compressor blades  56 ,  58  rotate relative to a corresponding set of static compressor vanes  60 ,  62  (also called a nozzle) to compress or pressurize the stream of fluid passing through the stage. In a single compressor stage  52 ,  54 , multiple compressor blades  56 ,  58  can be provided in a ring and can extend radially outwardly relative to the centerline  12 , from a blade platform to a blade tip, while the corresponding static compressor vanes  60 ,  62  are positioned upstream of and adjacent to the rotating blades  56 ,  58 . It is noted that the number of blades, vanes, and compressor stages shown in  FIG. 1  were selected for illustrative purposes only, and that other numbers are possible. 
     The blades  56 ,  58  for a stage of the compressor can be mounted to a disk  61 , which is mounted to the corresponding one of the HP and LP spools  48 ,  50 , with each stage having its own disk  61 . The vanes  60 ,  62  for a stage of the compressor can be mounted to the core casing  46  in a circumferential arrangement. 
     The HP turbine  34  and the LP turbine  36  respectively include a plurality of turbine stages  64 ,  66 , in which a set of turbine blades  68 ,  70  are rotated relative to a corresponding set of static turbine vanes  72 ,  74  (also called a nozzle) to extract energy from the stream of fluid passing through the stage. In a single turbine stage  64 ,  66 , multiple turbine blades  68 ,  70  can be provided in a ring and can extend radially outwardly relative to the centerline  12  while the corresponding static turbine vanes  72 ,  74  are positioned upstream of and adjacent to the rotating blades  68 ,  70 . It is noted that the number of blades, vanes, and turbine stages shown in  FIG. 1  were selected for illustrative purposes only, and that other numbers are possible. 
     The blades  68 ,  70  for a stage of the turbine can be mounted to a disk  71 , which is mounted to the corresponding one of the HP and LP spools  48 ,  50 , with each stage having a dedicated disk  71 . The vanes  72 ,  74  for a stage of the compressor can be mounted to the core casing  46  in a circumferential arrangement. 
     Complementary to the rotor portion, the stationary portions of the engine  10 , such as the static vanes  60 ,  62 ,  72 ,  74  among the compressor and turbine section  22 ,  32  are also referred to individually or collectively as a stator  63 . As such, the stator  63  can refer to the combination of non-rotating elements throughout the engine  10 . 
     In operation, the airflow exiting the fan section  18  is split such that a portion of the airflow is channeled into the LP compressor  24 , which then supplies pressurized air  76  to the HP compressor  26 , which further pressurizes the air. The pressurized air  76  from the HP compressor  26  is mixed with fuel in the combustor  30  and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine  34 , which drives the HP compressor  26 . The combustion gases are discharged into the LP turbine  36 , which extracts additional work to drive the LP compressor  24 , and the exhaust gas is ultimately discharged from the engine  10  via the exhaust section  38 . The driving of the LP turbine  36  drives the LP spool  50  to rotate the fan  20  and the LP compressor  24 . 
     A portion of the pressurized airflow  76  can be drawn from the compressor section  22  as bleed air  77 . The bleed air  77  can be drawn from the pressurized airflow  76  and provided to engine components requiring cooling. The temperature of pressurized airflow  76  entering the combustor  30  is significantly increased. As such, cooling provided by the bleed air  77  is necessary for operating of such engine components in the heightened temperature environments. 
     A remaining portion of the airflow  78  bypasses the LP compressor  24  and engine core  44  and exits the engine assembly  10  through a stationary vane row, and more particularly an outlet guide vane assembly  80 , comprising a plurality of airfoil guide vanes  82 , at the fan exhaust side  84 . More specifically, a circumferential row of radially extending airfoil guide vanes  82  are utilized adjacent the fan section  18  to exert some directional control of the airflow  78 . 
     Some of the air supplied by the fan  20  can bypass the engine core  44  and be used for cooling of portions, especially hot portions, of the engine  10 , and/or used to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are normally downstream of the combustor  30 , especially the turbine section  32 , with the HP turbine  34  being the hottest portion as it is directly downstream of the combustion section  28 . Other sources of cooling fluid can be, but are not limited to, fluid discharged from the LP compressor  24  or the HP compressor  26 . 
       FIG. 2  illustrates an axial view of a shroud  100  in the turbine engine of  FIG. 1 . The shroud  100  comprises at least two shroud elements, illustrated as a first shroud element  101  and second shroud element  102  that together form a ring. The elements  101 ,  102  each have confronting radial ends  105  defining a split interface  120 . 
     Turning to  FIG. 3 , each radial end  105  can comprise an axial fore portion  111 , an axial aft portion  112 , and a circumferential portion  113 . Similarly, the split surface interface  120  can comprise a fore split surface interface  121 , an aft split surface interface  122 , and a circumferential interface  123  as shown. 
     The first shroud element  101  is shown in  FIG. 4  looking toward the aft direction, while the second shroud element  102  is shown in  FIG. 5  looking toward the fore direction, which is opposite the view in  FIG. 4 . For each element  101 ,  102 , the fore portion  111  can define the fore split surface interface  121 , the aft portion  112  can define the aft split surface interface  122 , and the circumferential portion  113  can define the circumferential interface  123 . Either or both of the fore and aft interfaces  121 ,  122  may be planar; for example, when viewed along the engine centerline a first plane can be defined by a fore surface plane  131  that forms a positive radial angle β relative to a radial line  150 , and a second plane can be defined by an aft surface plane  132  that forms a negative radial angle α relative to the radial line  150 . Further, the circumferential portion  113  can define the circumferential interface  123  which may form an angle (not shown) relative to the radial line  150 . 
     It is contemplated that the fore portions  111  of the radial ends  105  of the first and second elements  101 ,  102  comprise first complementary surfaces  171  when the elements  101 ,  102  are joined together; similarly, the aft portions  112  of the first and second elements  101 ,  102  comprise second complementary surfaces  172 . Either or both of the surfaces  171 ,  172  may be planar, where the first complementary surface  171  can form a positive radial angle β relative to the radial line  150  and the second complementary surface  172  can form a negative radial angle α relative to the radial line  150  as described above. 
     In  FIG. 6 , a circumferential view of the first shroud element  101  is shown. The circumferential portions  113  can comprise third complementary surfaces  173  which connect the first and second complementary surfaces  171 ,  172  and which may be planar. While illustrated in alignment with the radial line  150 , it is contemplated that the third complementary surfaces  173  of each shroud element  101 ,  102  may each form an angle relative to the radial line  150  in a manner similar to α and β wherein the surface  173  of the first shroud element  101  forms a positive angle, and the surface  173  of the second shroud element  102  forms a negative angle, with respect to the radial line  150 . 
     It can be appreciated that when the first and second elements  101 ,  102  are joined in a ring to form the shroud  100 , the first, second, and third complementary surfaces  171 ,  172 ,  173  on the radial ends  105  can be part of first, second, and third complementary structures  181 ,  182 , and  183 , respectively ( FIGS. 4 and 5 ). The first structure  181  can form a first angle α relative to the radial line  150 , and the second structure  182  can form a second angle β, which may be opposite the first angle α, relative to the radial line  150 . Further, the third structure  183  can form a third angle ( FIG. 6 ) which may be a compound angle relative to the radial line  150 ; for example, the third angle may be formed by a rotation in both the axial and circumferential directions with respect to the radial line  150 . 
     When joined, the first complementary structures  181  can impede relative radial movement, the second complementary structures  182  can impede relative axial movement, and the third complementary structures  183  can impede relative circumferential movement of the shroud elements  101 ,  102 . It is further contemplated that any of the structures  181 ,  182 ,  183  can impede relative movement of the shroud elements  101 ,  102  in the radial, axial, or circumferential direction. For example: in  FIG. 4 , the second structure  182  can impede relative movement in both radial and circumferential directions due to its angle α with respect to the radial line  150 , or the third structure  183  may impede relative movement in both axial and circumferential directions due to its compound third angle with the radial line  150 . 
     Turning to  FIGS. 7A-7F , top views of the shroud  100  illustrate various options for the split surface interface  120  where an axial centerline  160  is shown throughout for reference ( FIG. 7A ). The shroud  100  has been illustrated thus far with the fore and aft planes  131 ,  132  parallel to the axial centerline  160  and with the circumferential interface  123  perpendicular to the centerline  160  ( FIG. 7B ). It is contemplated that the fore plane  131  may form a first axial angle  191  with the centerline  160  ( FIG. 7C ), and the aft plane  132  may form a second axial angle  192  with the centerline  160  ( FIG. 7D ). It is also contemplated that the circumferential interface  123  may form a third axial angle  193  with the centerline  160  ( FIG. 7E ), and further, that any combination of angles  191 ,  192 ,  193  may be selected for use in the shroud  100 . For example, the first axial angle  191  may be positive while the second axial angle  192  may be negative with respect to the centerline  160  ( FIG. 7F ). It can be appreciated that any of the first, second, or third axial angles  191 ,  192 ,  193  can impede relative movement in both the axial and circumferential directions. 
     It can be further appreciated that preventing relative motion between the shroud elements  101 ,  102  can decrease the rate at which the walls of the shroud  100  are worn while the engine is in operation. In addition, the reduced relative motion can allow for the use of less rigid (and less expensive) materials when constructing the shroud  100 . 
     It should be understood that application of the disclosed design is not limited to turbine engines with fan and booster sections, but is applicable to turbojets and turbo engines as well. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.