Patent Publication Number: US-10787923-B2

Title: Axially preloaded seal

Description:
FIELD 
     The present disclosure relates to gas turbine engines and, more particularly, to seals used to prevent leakage between gas paths within gas turbine engines. 
     BACKGROUND 
     Gas turbine engines, such as those used to power modern commercial and military aircraft, include a fan section to propel the aircraft, a compressor section to pressurize a supply of air from the fan section, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases in order to power the compressor and fan sections. 
     Various gas-flow streams or gas paths exist within gas turbine engines, including a core engine gas path and a bypass duct gas path. Typically, the various gas paths are kept separate from one another using various components, such as seals. Air flows within higher pressure gas paths, such as within high pressure compressor and turbine sections may, however, still tend to leak into air flows within lower pressure gas paths. Such leakages may be exacerbated by temperature extremes and other harsh environmental conditions existing within the internal engine environment and may affect the integrity of the components separating different gas-flow streams. Flow leakage from relatively high pressure gas paths into relatively low pressure gas paths may have a negative effect on engine fuel burn, performance, efficiency and component life. 
     SUMMARY 
     A seal for a gas turbine engine is disclosed. In various embodiments, the seal includes a first annular ring; a second annular ring; an annular brush having a first brush end disposed between the first annular ring and the second annular ring; and an axial ring seal having a first seal end configured for contact with the second annular ring. 
     In various embodiments, the annular brush has a second brush end configured for contact with a radially outer surface of a blade outer air seal. In various embodiments, the axial ring seal has a second seal end configured for contact with an axial face of a component disposed downstream of the annular brush. In various embodiments, the component is a hook connected to the blade outer air seal. In various embodiments, the component is an annular tab extending radially inward from an engine casing structure. In various embodiments, the second seal end of the axial ring seal is disposed radially inward of the first seal end. In various embodiments, the second seal end of the axial ring seal is disposed radially outward of the first seal end. 
     In various embodiments, the second annular ring includes a socket configured to receive the first seal end of the axial ring seal. In various embodiments, the socket is oriented in a radially inward direction and the second seal end of the axial ring seal is configured for positioning radially inward of the first seal end. In various embodiments, the socket is oriented in a radially outward direction and the second seal end of the axial ring seal is configured for positioning radially outward of the first seal end. 
     A gas turbine engine is disclosed. In various embodiments, the gas turbine engine includes a blade outer air seal disposed radially outward of a turbine rotor; an engine casing structure disposed radially outward of the blade outer air seal; and an annular seal configured to restrict intermixing of a core flow path and a cooling flow path. In various embodiments, the annular seal includes a first annular ring; a second annular ring; an annular brush having a first brush end disposed between the first annular ring and the second annular ring; and an axial ring seal having a first seal end configured for contact with the second annular ring. 
     In various embodiments, the annular brush has a second brush end configured for contact with a radially outer surface of the blade outer air seal and the axial ring seal has a second seal end configured for contact with a component disposed downstream of the annular brush. In various embodiments, the second seal end is configured for contact with an axial face of a hook connected to the blade outer air seal. In various embodiments, the second seal end is configured for contact with an annular tab extending radially inward from the engine casing structure. In various embodiments, the second seal end of the axial ring seal is disposed radially inward of the first seal end. In various embodiments, the second seal end of the axial ring seal is disposed radially outward of the first seal end. 
     In various embodiments, the second annular ring includes a socket configured to receive the first seal end of the axial ring seal. In various embodiments, the socket is oriented in a radially inward direction and the second seal end of the axial ring seal is configured for positioning radially inward of the first seal end. In various embodiments, the socket is oriented in a radially outward direction and the second seal end of the axial ring seal is configured for positioning radially outward of the first seal end. 
     A turbine section for a gas turbine engine is disclosed. In various embodiments, the turbine section includes a rotor having a plurality of blades; a blade outer air seal disposed radially outward of the rotor; an engine casing structure disposed radially outward of the blade outer air seal; and an annular seal. In various embodiments, the annular ring seal includes a first annular ring; a second annular ring; an annular brush having a first brush end configured for disposition between the first annular ring and the second annular ring and a second brush end configured for contact with the blade outer air seal; and an axial ring seal having a first seal end configured for contact with the second annular ring. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The subject matter of the present disclosure is particularly pointed out and distinctly claimed in the concluding portion of the specification. A more complete understanding of the present disclosure, however, may best be obtained by referring to the following detailed description and claims in connection with the following drawings. While the drawings illustrate various embodiments employing the principles described herein, the drawings do not limit the scope of the claims. 
         FIG. 1A  is a schematic view of a gas turbine engine, in accordance with various embodiments; 
         FIG. 1B  is a schematic side view of a rotor and vane assembly of a gas turbine engine, in accordance with various embodiments; 
         FIG. 2  is a schematic view of a seal used within a gas turbine engine, in accordance with various embodiments; and 
         FIG. 3  is a schematic view of a seal used within a gas turbine engine, in accordance with various embodiments. 
     
    
    
     DETAILED DESCRIPTION 
     The following detailed description of various embodiments herein makes reference to the accompanying drawings, which show various embodiments by way of illustration. While these various embodiments are described in sufficient detail to enable those skilled in the art to practice the disclosure, it should be understood that other embodiments may be realized and that changes may be made without departing from the scope of the disclosure. Thus, the detailed description herein is presented for purposes of illustration only and not of limitation. Furthermore, any reference to singular includes plural embodiments, and any reference to more than one component or step may include a singular embodiment or step. Also, any reference to attached, fixed, connected, or the like may include permanent, removable, temporary, partial, full or any other possible attachment option. Additionally, any reference to without contact (or similar phrases) may also include reduced contact or minimal contact. It should also be understood that unless specifically stated otherwise, references to “a,” “an” or “the” may include one or more than one and that reference to an item in the singular may also include the item in the plural. Further, all ranges may include upper and lower values and all ranges and ratio limits disclosed herein may be combined. 
     Referring now to the drawings,  FIG. 1A  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle  15 , while the compressor section  24  drives air along a core or primary flow path C for compression and communication into the combustor section  26  and then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines. 
     The gas turbine engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems at various locations may alternatively or additionally be provided and the location of the several bearing systems  38  may be varied as appropriate to the application. The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a low pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in this gas turbine engine  20  is illustrated as a fan drive gear system  48  configured to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  and a high pressure turbine  54 . A combustor  56  is arranged in the gas turbine engine  20  between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46  and may include airfoils  59  in the core flow path C for guiding the flow into the low pressure turbine  46 . The mid-turbine frame  57  further supports the several bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via the several bearing systems  38  about the engine central longitudinal axis A, which is collinear with longitudinal axes of the inner shaft  40  and the outer shaft  50 . 
     The air in the core flow path C is compressed by the low pressure compressor  44  and then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , and then expanded over the high pressure turbine  54  and low pressure turbine  46 . The low pressure turbine  46  and the high pressure turbine  54  rotationally drive the respective low speed spool  30  and the high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , the compressor section  24 , the combustor section  26 , the turbine section  28 , and the fan drive gear system  48  may be varied. For example, the fan drive gear system  48  may be located aft of the combustor section  26  or even aft of the turbine section  28 , and the fan section  22  may be positioned forward or aft of the location of the fan drive gear system  48 . 
     Referring now to  FIG. 1B , selected portions of a turbine section  100  of a gas turbine engine, such as, for example, the turbine section  28  described above with reference to  FIG. 1A , are illustrated. The turbine section  100  includes alternating rows of rotor assemblies  102  and stator assemblies  104 . Each of the rotor assemblies  102  carries one or more rotor blades  106  for rotation about a central axis A. Each of the rotor blades  106  includes a rotor platform  108  and an airfoil  110  extending in a radial direction R from the rotor platform  108  to a rotor tip  112 . The airfoil  110  generally extends in a chord-wise direction X between a leading edge  114  and a trailing edge  116 . A root section  118  of each of the rotor blades  106  is mounted to a rotor disk  103 . A blade outer air seal (BOAS)  120  is disposed radially outward of the rotor tip  112  of the airfoil  110 . The BOAS  120  includes a platform  121  configured to provide a seal to prevent hot gases from leaking outside the core airflow path C (see  FIG. 1 ). 
     Each of the stator assemblies  104  includes one or more vanes  122  positioned along the engine axis A and adjacent to one or more rotor blades  106 . Each of the vanes  122  includes an airfoil  124  extending between an inner vane platform  126  and an outer vane platform  128 . The stator assemblies  104  are connected to an engine casing structure  130 . The BOAS  120  and the stator assemblies  104  may be disposed radially inward of the engine casing structure  130 . In various embodiments, one or both of the BOAS  120  and the stator assemblies  104  may include full annular platforms or they may be segmented and include feather seals between segments to help prevent leakage of cooling fluid between the segments. In various embodiments, one or more of the vanes  122  may be configured to rotate about an axis extending between the inner vane platform  126  and the outer vane platform  128 . In various embodiments, and as described below, an annular seal  150  may be disposed between the BOAS  120  and the engine casing structure  130  to provide further assurance against leakage between separate gas paths. 
     Referring to  FIG. 2 , an annular seal  250  is illustrated, in accordance with various embodiments. The annular seal  250  is similar to the annular seal  150  described above with reference to  FIG. 1B . In various embodiments, the annular seal  250  is disposed between a BOAS  220  and an engine casing structure  230 . In various embodiments, the annular seal  250  includes a first annular ring  252  and a second annular ring  254 . The annular seal  250  further includes an annular brush  256 , having a first brush end  258  sandwiched between the first annular ring  252  and the second annular ring  254  and a second brush end  260  configured for sealing contact with a radially outer surface  262  of the BOAS  220 . In various embodiments, the first annular ring  252  is positioned adjacent an annular tab  253  that extends radially inward from the engine casing structure  230 . In various embodiments, the first annular ring  252  and the second annular ring  254  are connected to one another, thereby sandwiching the annular brush  256  there between, by rivets or welding or the like. 
     In various embodiments, the annular seal  250  further includes an annular ring seal  264  (sometimes referred to in the art as a dog-bone seal  266 ). In various embodiments, the annular ring seal  264  provides an axial seal that restricts intermixing of gas flow paths and operates as a mechanical spring, due to an elastic pre-load in the axial direction applied to the annular ring seal  264  during assembly of the engine. In various embodiments, the annular ring seal  264  provides an axial interference fit between, for example, a the second annular ring  254  and an axial face  268  of a BOAS hook  270  connected to the BOAS  220 . In various embodiments, the axial spring nature of the annular ring seal  264  enables an axial seal to be maintained in the presence of thermal expansion of the various engine components. For example, during instances where the axial face  268  of the BOAS hook  270  and the annular tab  253  of the engine casing structure  230  move axially apart from one another due to thermal expansion, a first seal end  272  of the annular ring seal  264  and a second seal end  274  of the annular ring seal  264  may be subject to a rolling type motion, where the first seal end  272  may be urged in an axially forward direction (i.e., toward the annular tab  253 ) and the second seal end  274  may be urged in an axially rearward direction (i.e., toward the axial face  268  of the BOAS hook  270 ). In various embodiments, a socket  276  is disposed within the second annular ring  254  and serves to maintain the first seal end  272  at a constant radial position during thermal expansion of the engine, while the second seal end  274  is free to move in both the axial and radial directions is response to such thermal expansion. In various embodiments, the socket  276  faces radially inward and the first seal end  272  is disposed radially outward of the second seal end  274 , enabling the second seal end  274  to move radially outward and axially rearward during thermal expansion of the engine. 
     Referring now to  FIG. 3 , an annular seal  350  is illustrated, in accordance with various embodiments. In various embodiments, the annular seal  350  is disposed between a BOAS  320  and an engine casing structure  330 . In various embodiments, the annular seal  350  includes a first annular ring  352  and a second annular ring  354 . The annular seal  350  further includes an annular brush  356 , having a first brush end  358  sandwiched between the first annular ring  352  and the second annular ring  354  and a second brush end  360  configured for sealing contact with a radially outer surface  362  of the BOAS  320 . In various embodiments, the first annular ring  352  is positioned adjacent an annular tab  353  that extends radially inward from the engine casing structure  330 . In various embodiments, the first annular ring  352  and the second annular ring  354  are connected to one another, thereby sandwiching the annular brush  256  there between, by rivets or welding or the like. 
     In various embodiments, the annular seal  350  further includes an annular ring seal  364  which, in various embodiments, comprises a dog-bone seal  366 . In various embodiments, the annular ring seal  364  provides an axial seal that restricts intermixing of gas flow paths and operates as a mechanical spring, due to an elastic pre-load in the axial direction applied to the annular ring seal  364  during assembly of the engine. In various embodiments, the annular ring seal  364  provides an axial interference fit between, for example, a the second annular ring  354  and an axial face  368  of a second annular tab  369  that extends radially inward from the engine casing structure  330 . In various embodiments, the axial spring nature of the annular ring seal  364  enables an axial seal to be maintained in the presence of thermal expansion of the various engine components. For example, during instances where the axial face  368  of the second annular tab  369  and the annular tab  353  of the engine casing structure  330  move axially apart from one another due to thermal expansion, a first seal end  372  of the annular ring seal  364  and a second seal end  374  of the annular ring seal  364  may be subject to a rolling type motion, where the first seal end  372  may be urged in an axially forward direction (i.e., toward the annular tab  353 ) and the second seal end  374  may be urged in an axially rearward direction (i.e., toward the axial face  368  of the second annular tab  369 ). In various embodiments, a socket  376  is disposed within the second annular ring  354  and serves to maintain the first seal end  372  at a constant radial position during thermal expansion of the engine, while the second seal end  374  is free to move in both the axial and radial directions is response to such thermal expansion. In various embodiments, the socket  376  faces radially outward and the first seal end  372  is disposed radially inward of the second seal end  374 , enabling the second seal end  374  to move radially inward and axially rearward during thermal expansion of the engine. 
     In various embodiments, both the annular seal  250  described with reference to  FIG. 2  and the annular seal  350  described with reference to  FIG. 3  provide a multiple point sealing configuration. For example, the annular brush  256  described with reference to  FIG. 2  provides a seal between a core gas path (e.g., air flowing through the turbine section defined by the rotor blades  106  and the vanes  122  and the various platforms described above with reference to  FIG. 1B ) and a cooling air gas path that may flow between the engine casing structure  230  and the BOAS  220 . In addition, the annular ring seal  264  provides a seal by the first seal end  272  and the second seal end  274  being maintained in contact with the corresponding faces of the annular tab  253  and the axial face  268 , thereby preventing the cooling air gas path from leaking past the annular ring seal  264 . Similarly, the annular brush  356  described with reference to  FIG. 3  provides a seal between a core gas path (e.g., air flowing through the turbine section defined by the rotor blades  106  and the vanes  122  and the various platforms described above with reference to  FIG. 1B ) and a cooling air gas path that may flow between the engine casing structure  330  and the BOAS  320 . In addition, the annular ring seal  364  provides a seal by the first seal end  372  and the second seal end  374  being maintained in contact with the corresponding faces of the annular tab  353  and the second annular tab  369 , thereby preventing the cooling air gas path from leaking past the annular ring seal  364 . 
     Benefits, other advantages, and solutions to problems have been described herein with regard to specific embodiments. Furthermore, the connecting lines shown in the various figures contained herein are intended to represent exemplary functional relationships and/or physical couplings between the various elements. It should be noted that many alternative or additional functional relationships or physical connections may be present in a practical system. However, the benefits, advantages, solutions to problems, and any elements that may cause any benefit, advantage, or solution to occur or become more pronounced are not to be construed as critical, required, or essential features or elements of the disclosure. The scope of the disclosure is accordingly to be limited by nothing other than the appended claims, in which reference to an element in the singular is not intended to mean “one and only one” unless explicitly so stated, but rather “one or more.” Moreover, where a phrase similar to “at least one of A, B, or C” is used in the claims, it is intended that the phrase be interpreted to mean that A alone may be present in an embodiment, B alone may be present in an embodiment, C alone may be present in an embodiment, or that any combination of the elements A, B and C may be present in a single embodiment; for example, A and B, A and C, B and C, or A and B and C. Different cross-hatching is used throughout the figures to denote different parts but not necessarily to denote the same or different materials. 
     Systems, methods and apparatus are provided herein. In the detailed description herein, references to “one embodiment”, “an embodiment”, “various embodiments”, etc., indicate that the embodiment described may include a particular feature, structure, or characteristic, but every embodiment may not necessarily include the particular feature, structure, or characteristic. Moreover, such phrases are not necessarily referring to the same embodiment. Further, when a particular feature, structure, or characteristic is described in connection with an embodiment, it is submitted that it is within the knowledge of one skilled in the art to affect such feature, structure, or characteristic in connection with other embodiments whether or not explicitly described. After reading the description, it will be apparent to one skilled in the relevant art(s) how to implement the disclosure in alternative embodiments. 
     Furthermore, no element, component, or method step in the present disclosure is intended to be dedicated to the public regardless of whether the element, component, or method step is explicitly recited in the claims. No claim element herein is to be construed under the provisions of 35 U.S.C. 112(f) unless the element is expressly recited using the phrase “means for.” As used herein, the terms “comprises”, “comprising”, or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus. 
     Finally, it should be understood that any of the above described concepts can be used alone or in combination with any or all of the other above described concepts. Although various embodiments have been disclosed and described, one of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. Accordingly, the description is not intended to be exhaustive or to limit the principles described or illustrated herein to any precise form. Many modifications and variations are possible in light of the above teaching.