Abstract:
A method of making a shroud assembly by: forming a shroud frame consisting of a one-piece hook and rail assembly which includes a first rail and a second rail disposed at lateral edges and a first support and a second support connected to the first rail and the second rail, the first and second rails and the first and second supports being integrally joined to one another and defining a central aperture therebetween; joining a metallic backsheet to the shroud frame covering the central aperture; and attaching a honeycomb rubs strip to the metallic backsheet via an upper edge of the honeycomb seal structure.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    This application is a divisional of U.S. Ser. No. 13/827,762, filed on Mar. 14, 2013, the entire disclosure of which is incorporated herein by reference. 
     
    
     BACKGROUND 
       [0002]    Present embodiments relate generally to a gas turbine engine. More specifically, the present embodiments relate, but are not limited to, reducing leakage at a in a low pressure turbine section of a gas turbine engine. 
         [0003]    A typical gas turbine engine generally possesses a forward end and an aft end with its several core or propulsion components positioned axially therebetween. An air inlet or intake is located at a forward end of the engine. Moving toward the aft end, in order, the intake is followed by a fan, a compressor, a combustion chamber, and a turbine. It will be readily apparent from those skilled in the art that additional components may also be included in the engine, such as, for example, low-pressure and high-pressure compressors, and low-pressure and high-pressure turbines. This, however, is not an exhaustive list. 
         [0004]    The compressor and turbine generally include rows of airfoils that are stacked axially in stages. Each stage includes a row of circumferentially spaced stator vanes and a row of rotor blades which rotate about a center shaft or axis of the turbine engine. A multi-stage low pressure turbine follows the multi-stage high pressure turbine and is typically joined by a second shaft to a fan disposed upstream from the compressor in a typical turbo fan aircraft engine configuration for powering an aircraft in flight. 
         [0005]    The stator is formed by a plurality of nozzle segments which are abutted at circumferential ends to form a complete ring about the axis of the gas turbine engine. Each nozzle segment may comprise a single vane, commonly referred to as a singlet. Alternatively, a nozzle segment may have two vanes per segment, which are generally referred to as doublets. In a third embodiment, additional numbers of vanes may be disposed on a single segment. In these embodiments, the vanes extend between an inner band and an outer band. 
         [0006]    A typical gas turbine engine utilizes a high pressure turbine and low pressure turbine to maximize extraction of energy from high temperature combustion gas. The turbine section typically has an internal shaft axially disposed along a center longitudinal axis of the engine. The blades are circumferentially distributed on a rotor causing rotation of the internal shaft. The internal shaft is connected to the rotor and the air compressor, such that the turbine provides a rotational input to the air compressor to drive the compressor blades. This powers the compressor during operation and subsequently drives the turbine. As the combustion gas flows downstream through the turbine stages, energy is extracted therefrom and the pressure of the combustion gas is reduced. 
         [0007]    In operation, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages. These turbine stages extract energy from the combustion gases. A high pressure turbine first receives the hot combustion gases from the combustor and includes a stator nozzle assembly directing the combustion gases downstream through a row of high pressure turbine rotor blades extending radially outwardly from a supporting rotor disk. The stator nozzles turn the hot combustion gas in a manner to maximize extraction at the adjacent downstream turbine blades. In a two stage turbine, a second stage stator nozzle assembly is positioned downstream of the first stage blades followed in turn by a row of second stage rotor blades extending radially outwardly from a second supporting rotor disk. The turbine converts the combustion gas energy to mechanical energy. 
         [0008]    During such operation of the gas turbine engine, it is desirable to minimize thermally induced deformation of the outer casing through the turbine section of the engine. This is accomplished, according to some embodiments, by isolating the outer casing from heat produced by the hot combustion gases flowing through the turbine. Turbine shrouds are connected to the engine casing to provide an outer boundary flow for the combustion gas limiting high temperature combustion gas from adversely affecting the casing. The shroud extends circumferentially to form a ring shape and may be formed of a plurality of circumferentially extending shroud segments. However, as combustion gas moves radially outward with rotation of the turbine blades, the combustion gas can pass through axial seams between the adjacent shroud segments. This is not optimal and results in energy losses. 
         [0009]    It would be desirable to overcome these and other deficiencies with turbine sections of gas turbine engines. More specifically it would be desirable to provide a restriction in at least a radial direction to flow of combustion gas between shroud segments. 
       SUMMARY 
       [0010]    According to some embodiments, disclosed herein is a method of making a shroud assembly by: forming a shroud frame consisting of a one-piece hook and rail assembly which includes a first rail and a second rail disposed at lateral edges and a first support and a second support connected to the first rail and the second rail, the first and second rails and the first and second supports being integrally joined to one another and defining a central aperture therebetween; joining a metallic backsheet to the shroud frame covering the central aperture; and attaching a honeycomb rubs strip to the metallic backsheet via an upper edge of the honeycomb seal structure. 
         [0011]    All of the above outlined features are to be understood as exemplary only and many more features and objectives of the turbine shroud with spline seal may be gleaned from the disclosure herein. Therefore, no limiting interpretation of this summary is to be understood without further reading of the entire specification, claims, and drawings included herewith. 
     
    
     
       BRIEF DESCRIPTION OF THE ILLUSTRATIONS 
         [0012]    The above-mentioned and other features and advantages of these exemplary embodiments, and the manner of attaining them, will become more apparent and the turbine shroud with spline seal feature will be better understood by reference to the following description of embodiments taken in conjunction with the accompanying drawings, wherein: 
           [0013]      FIG. 1  is a side section view of an exemplary gas turbine engine; 
           [0014]      FIG. 2  is an isometric view of a shroud assembly segment of instant embodiments; 
           [0015]      FIG. 3  is an exploded assembly view of the shroud assembly segment of  FIG. 2 ; 
           [0016]      FIG. 4  is a axial view of two shroud assembly segments; 
           [0017]      FIG. 5  is an isometric view of an alternate embodiment of a shroud assembly segment; 
           [0018]      FIG. 6  is a side view of an alternate shroud assembly segment. 
       
    
    
     DETAILED DESCRIPTION 
       [0019]    Reference now will be made in detail to embodiments provided, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation, not limitation of the disclosed embodiments. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present embodiments without departing from the scope or spirit of the disclosure. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to still yield further embodiments. Thus it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents. 
         [0020]    Referring to  FIGS. 1-6 , various embodiments of a gas turbine engine are depicted having a turbine shroud with spline seal. The shroud includes a spline for locating a spline seal between an adjacent spline. The shroud may also have a back sheet extending in an aft direction to limit leakage aft of the shroud. In addition to limiting weight, it is desirable to reduce weight. 
         [0021]    As used herein, the terms “axial” or “axially” refer to a dimension along a longitudinal axis of an engine. The term “forward” used in conjunction with “axial” or “axially” 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” used in conjunction with “axial” or “axially” refers to moving in a direction toward the engine nozzle, or a component being relatively closer to the engine nozzle as compared to another component. 
         [0022]    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. The use of the terms “proximal” or “proximally,” either by themselves or in conjunction with the terms “radial” or “radially,” refers to moving in a direction toward the center longitudinal axis, or a component being relatively closer to the center longitudinal axis as compared to another component. The use of the terms “distal” or “distally,” either by themselves or in conjunction with the terms “radial” or “radially,” refers to moving in a direction toward the outer engine circumference, or a component being relatively closer to the outer engine circumference as compared to another component. As used herein, the terms “lateral” or “laterally” refer to a dimension that is perpendicular to both the axial and radial dimensions. 
         [0023]    Referring initially to  FIG. 1 , a schematic side section view of a gas turbine engine  10  is shown. The function of the gas turbine engine is to extract energy from high pressure and temperature combustion gases and convert the energy into mechanical energy for work. The gas turbine engine  10  has an engine inlet end  12  wherein air enters the core or propulsor  13  which is defined generally by a compressor  14 , a combustor  16  and a multi-stage high pressure turbine  20 . Collectively, the propulsor  13  provides thrust or power during operation. The gas turbine  10  may be used for aviation, power generation, industrial, marine or the like. 
         [0024]    In operation air enters through the air inlet end  12  of the engine  10  and moves through at least one stage of compression where the air pressure is increased and directed to the combustor  16 . The compressed air is mixed with fuel and burned providing the hot combustion gas which exits the combustor  16  toward the high pressure turbine  20 . At the high pressure turbine  20 , energy is extracted from the hot combustion gas causing rotation of turbine blades which in turn cause rotation of the shaft  24 . The shaft  24  passes toward the front of the engine to continue rotation of the one or more compressor stages  14 , a turbofan  18  or inlet fan blades, depending on the turbine design. The turbofan  18  is connected by the shaft  28  to a low pressure turbine  21  and creates thrust for the turbine engine  10 . A low pressure turbine  21  may also be utilized to extract further energy and power additional compressor stages. 
         [0025]    Referring now to  FIG. 2 , an isometric view of a shroud assembly segment  30  is depicted. Instant embodiments of the shroud segment assembly  30  are located in the low pressure turbine  21  area of the engine  10  ( FIG. 1 ). The instant segment assembly  30  embodiment is a one-piece cast hook and rail assembly. The embodiment utilizes a frame  31  comprising a first rail  32  and a second rail  34  which extend in a generally axial direction parallel to the gas turbine engine axis  26  ( FIG. 1 ) or alternatively may be at an acute angle relative to the axis  26 . Extending laterally or in the circumferential direction between the first and second rails  32 ,  34  the frame  31  further comprises a first or forward support  40  and a second or aft support  42 . The shroud assembly segment  30  includes forward tips  44  and may include aft tips  46 ,  47  extending from the aft support  42 . As described further, the rails  32 ,  34  may be formed of various cross-sections and may be of various materials. Similarly, the supports  40 ,  42  may have various shapes and may be formed of various materials and formed by a variety of manufacturing processes. 
         [0026]    The rails  32 ,  34  include a slot or spline  39  disposed on circumferential ends or slash face surfaces  36 ,  38 . The splines  39  extend along the surfaces  36 ,  38  to receive a spline seal  37 . The spline seal  37  is positioned at one circumferential end to a first segment  30  and at a second circumferential end to a circumferentially adjacent segment (not shown). As turbine blades move radially beneath, the combustion gas moves both radially and axially. The spline seal  37  precludes combustion gas moving within the turbine from passing in a radial direction between segments  30  defining the shroud. 
         [0027]    Referring now to  FIG. 3 , an exploded assembly view of the shroud assembly segment  30  is depicted in isometric view. First, the shroud frame  31  includes a cast structure having the rails  32 ,  34  integrally joined to supports  40 ,  42 . According to alternative embodiments, the frame  31  may be forged, cast, direct metal laser sintered or as a further alternative may be formed of metallic plate or bar material. The rails  32 ,  34  may alternatively be formed of back sheet stock or other materials and of various cross-sectional shapes. 
         [0028]    The supports  40 ,  42  may take various shapes described further herein. According to the instant embodiment, the supports  40 ,  42  are hook-shaped which may include various cross-sections. For example, the depicted supports  40 ,  42  are generally inverted L-shaped structures extending vertically from, and between, the rails  32 ,  34 . According to the instant embodiment, the supports  40 ,  42  are integrally formed with the rails  32 ,  34 . As previously described, the rails  32 ,  34  have tips  44  at forward ends for aiding connection with an engine casing. A gap  45  is defined between the upper axial legs  41 ,  43  of the support  40 ,  42  and the rail tips  44 . In the instant embodiment, a similar gap is also defined between the upper leg of the aft support  42  and the rails  32 ,  34 . The gaps  45  receive a flange of the engine casing for mounting of the shroud assembly segment  30 . It should be understood by one skilled in the art that the supports  40 ,  42  are not limited to the L-shape shown but alternatively may be Z-shaped, C-shaped, straight or other shapes allowing the structure to be retained by the engine casing. Additionally, the supports  40 ,  42  and/or lower rail surfaces may be curved to approximate the curvature of the engine casing. 
         [0029]    The splines  39  are also positioned in the lateral or circumferentially outer faces  36 ,  38  of the rails  32 ,  34 . Each spline  39  allows for receiving a spline seal  37  to engage with an adjacent assembly segment  30 . The spline seal  37  inhibits radial leakage of air between the segments  30 . More specifically, since the segments  30  are circumferentially adjacent to one another, axial seams are formed between adjacent shroud assembly segments  30 . The spline seal  37  limits combustion gas from leaking through such seams. 
         [0030]    As depicted in broken line, the exemplary spline seal  37  is rectangular in shape, but may form a variety of shapes. For example, the seal structure  37  may be circular, square, rectangular, other polygons or geometries. The seal  37  may be formed of a singular material or may be a multi-material structure. The seal  37  may change shape at operating temperature as well. The seal  37  has a volumetric thermal expansion coefficient which is a thermodynamic property of the material. For example, the volumetric thermal expansion can be expressed as αv=(1/v)(ΔV/ΔT), where αv is the volumetric thermal expansion coefficient, V is the volume of the material and ΔV/ΔT with respect to the change in volume of the material with respect to the change in temperature of the material. Thus the volumetric thermal expansion coefficient measures the fractional change in volume per degree change in temperature at a constant temperature. 
         [0031]    When viewed in a forward looking aft direction, the adjacent shroud assembly segments  30  are positioned in their annular arrangement, the seals  37  are positioned in each adjacent spline  39  to block an air flow path which would otherwise allow flow between adjacent assembly segments  30 . 
         [0032]    Extending across the bottom surfaces of the rails  32 ,  34  is a backsheet  50  which may be, for example, metallic or various materials. The shield  50  is designed to extend in aft and circumferential directions of the shroud frame  31  so as to define a flow path along a radially inner side of the shroud frame  31 . The backsheet  50  is sized to extend circumferentially between lateral ends of the rails  32 ,  34  to the opposite circumferential end of rail  34 . The shield  50  also extends, in some embodiments, in an axial direction from forward end of the rails  32 ,  34  to aft ends of the rails  32 ,  34 . As depicted in the embodiment, the backsheet  50  may have a thickness which is less than prior art backsheet structures since the cast rails  32 ,  34  provide additional strength. The instant embodiment depicts the back sheet or shield  50  being of a constant thickness. However, according to some embodiments, the back sheet  50  may be formed of variable thickness. For example, areas which may be expected to receive impact from a detached rotor blade may have an increased thickness to dissipate energy of such ejected blade while areas adjacent the rails  32 ,  34  or supports  40 ,  42  are of thinner dimension radially. Similarly, while thicknesses of the rails  32 ,  34  are generally shown as constant, alternative embodiments may utilize rails of varying thickness. 
         [0033]    In addition to the first portion  52  of the sheet  50 , described above, the backsheet  50  may also include a second portion  54 . The second portion  54  of the shield  50  extends from an aft edge  53  of the first portion  52 . The first and second portions  52 ,  54  may be formed of a single sheet of metal as shown in the depicted view and bent or alternatively, may be joined from two separate pieces such as by welding or brazing. In a third alternative, the two pieces may be abutted against one another but not joined to one another. Instead, the first portion  52  may be joined with the frame  31  and the second portion  54  also joined with the frame  31  but the first and second portions  52 ,  54  closely abutting one another. 
         [0034]    The frame  31  includes aft tips  46 ,  47  extend from the aft side of support  42  and are formed at an angle to the rails  32 ,  34 . The angle of the tips  46 ,  47  approximate the angle of the second portion  54  relative to the first portion  52  of metallic sheet shield  50 . These tips  46 ,  47  may be formed integrally with the frame  31  or may be joined in a separate manufacturing step to extend from the frame  31 , for example welding or brazing. 
         [0035]    According to one embodiment, the tips  46 ,  47  may also include splines  49  within circumferential end surfaces of these structures. This allows for the additional spline seal  51  to be utilized in this area of the frame  31  inhibiting radial leakage between adjacent shroud assembly segments  30 . In an embodiment utilizing the spline  49 , the spline  49  may be formed continuously with spline  39  so that a single spline seal may be utilized. Alternatively, the spline  49  may be formed separately from but closely abut spline  39  and minimize any gap between these spline seal elements. In a further alternative, the spline  49  may be welded or brazed to spline  39  or may closely abut spline  39 . 
         [0036]    In still a further alternative, the second backsheet portions  54  may be widened in the circumferential direction so as to overlap second portions  54  of an adjacent backsheet  50 . This is shown in the embodiment as the optional back sheet portion in broken lines. This may eliminate the need or desire to have the spline seal  49  located in these tips  46 ,  47 . Thus in either embodiment, leakage aft of the second support  42  is limited. 
         [0037]    Referring still to  FIG. 3 , a honeycomb rub strip  60  is positioned beneath the sheet shield  50 . The honeycomb structure  60  is joined for example, mechanically, bonded, welded or brazed, directly to the back sheet  50  and is sized to extend circumferentially between the rails  32 ,  34  and axially from the tips  44  to the aft support  42 . The aft end of the honeycomb  60  may be cut on an angle to approximate the angle of the second portion  54  of the metallic back sheet  50  if such is utilized. The honeycomb rub strip  60  may take any of various conventional forms. The rub strip  60  may have a thickness in a radial direction so that it is radially inner surface spaced from a turbine tip to provide a minimal clearance gap therebetween. The honeycomb rub strip  60  may further include an abradable radially inner surface and define an outer boundary for the passage of hot combustion gas through the turbine section of the engine  10  ( FIG. 1 ). Additionally, the radially outward ends of the turbine blades  23  ( FIG. 5 ) may include sealing fins  25  ( FIG. 5 ) abutting the abradable surface  62  of the honeycomb rub strip  60 . The honeycomb rub strip  60  may be deformed by these sealing fins during rotation of the rotor blades  23  such that a nearly zero tolerance fit is defined between the honeycomb lower surface  62  and the sealing fins  25  of the rotor blades. This reduces the leakage of combustion gas through the turbine section of the engine  10 . 
         [0038]    Referring now to  FIG. 4 , an aft looking forward view of adjacent assembly segments is depicted. Each of the segments is joined by a spline seal  37  at slash face ends. Accordingly, the structure provides a circumferential design which lines the inner surface of the engine case to retain high temperature combustion gas on the lower side, as depicted, of the assembly segments  30  and inhibiting deformation of the engine casing along the outer perimeter of the assembly segments  30 . As depicted, the spline seal  37  inhibit high temperature combustion gas from escaping between the axially extending gaps between adjacent assembly segments  30 . The gaps in the depicted embodiment are exaggerated for ease of understanding, as one skilled in the art will understand. According to additional embodiments, the backsheets  50  may be extended in the circumferential direction to aid in reducing leakage near the aft end of the assembly segments  30 . These sheets  50  may overlap to aid in reducing leakage between the assembly segments  30 . 
         [0039]    Referring now to  FIG. 5 , an alternative embodiment is depicted for an exemplary shroud assembly segment  130 . According to this embodiment the frame  131 , defined by the rails  132 ,  134  and supports  140 ,  142  are not a one-piece structure. Instead, the rails  132 ,  134  are formed independently from the supports  140 ,  142 . The rails  132 ,  134  may be formed of cast rails, plate material or forged material in bar sheet stock form and may be formed of various cross-sections. Additionally, in order decrease weight, the supports  140 ,  142  are formed of sheet metal and are depicted to have an inverted L-shape although other cross sections may be utilized. In the instant embodiment, the sheet metal supports  140 ,  142  are welded or otherwise bonded to the rails  132 ,  134 . As an alternative, the rails  132 ,  134  may be formed of thickened sheet metal. In either embodiment, it is preferable that the sheet metal or the cast metal be thick enough to provide for formation of a spline  139  extending in the axial direction of the circumferential end faces of the rails  132 ,  134 . In the previous embodiment, the rails may or may not include a spline which is continuous or discontinuous from spline  149 . If a spline  149  is not used, it may be desirable to widen the sheet metal shield  150  so that adjacent sheets overlap aft of the second rail  142  to further limit leakage in these areas. 
         [0040]    Referring now to  FIG. 6 , a side view of an alternate shroud assembly segment  230  is depicted in an assembled view within a turbine section. The structure depicts views of a forward support  240  and a rear support  242  which is linear defining a shoulder, rather than the inverted L-shape previously described. These structures may be formed of a plurality of materials including but not limited to bar stock, plate stock, cast materials, forged materials and sheet metallics, including alloys. It should be understood from this description as well as the previous descriptions, that various cross sections may be utilized to define the support structures for any of shroud assembly segments. 
         [0041]    While multiple inventive embodiments have been described and illustrated herein, those of ordinary skill in the art will readily envision a variety of other means and/or structures for performing the function and/or obtaining the results and/or one or more of the advantages described herein, and each of such variations and/or modifications is deemed to be within the scope of the invent of embodiments described herein. More generally, those skilled in the art will readily appreciate that all parameters, dimensions, materials, and configurations described herein are meant to be exemplary and that the actual parameters, dimensions, materials, and/or configurations will depend upon the specific application or applications for which the inventive teachings is/are used. Those skilled in the art will recognize, or be able to ascertain using no more than routine experimentation, many equivalents to the specific inventive embodiments described herein. It is, therefore, to be understood that the foregoing embodiments are presented by way of example only and that, within the scope of the appended claims and equivalents thereto, inventive embodiments may be practiced otherwise than as specifically described and claimed. Inventive embodiments of the present disclosure are directed to each individual feature, system, article, material, kit, and/or method described herein. In addition, any combination of two or more such features, systems, articles, materials, kits, and/or methods, if such features, systems, articles, materials, kits, and/or methods are not mutually inconsistent, is included within the inventive scope of the present disclosure. 
         [0042]    Examples are used to disclose the embodiments, including the best mode, and also to enable any person skilled in the art to practice the apparatus and/or method, including making and using any devices or systems and performing any incorporated methods. These examples are not intended to be exhaustive or to limit the disclosure to the precise steps and/or forms disclosed, and many modifications and variations are possible in light of the above teaching. Features described herein may be combined in any combination. Steps of a method described herein may be performed in any sequence that is physically possible. 
         [0043]    All definitions, as defined and used herein, should be understood to control over dictionary definitions, definitions in documents incorporated by reference, and/or ordinary meanings of the defined terms. The indefinite articles “a” and “an,” as used herein in the specification and in the claims, unless clearly indicated to the contrary, should be understood to mean “at least one.” The phrase “and/or,” as used herein in the specification and in the claims, should be understood to mean “either or both” of the elements so conjoined, i.e., elements that are conjunctively present in some cases and disjunctively present in other cases. 
         [0044]    It should also be understood that, unless clearly indicated to the contrary, in any methods claimed herein that include more than one step or act, the order of the steps or acts of the method is not necessarily limited to the order in which the steps or acts of the method are recited. 
         [0045]    In the claims, as well as in the specification above, all transitional phrases such as “comprising,” “including,” “carrying,” “having,” “containing,” “involving,” “holding,” “composed of,” and the like are to be understood to be open-ended, i.e., to mean including but not limited to. Only the transitional phrases “consisting of” and “consisting essentially of” shall be closed or semi-closed transitional phrases, respectively, as set forth in the United States Patent Office Manual of Patent Examining Procedures, Section 2111.03.