Patent Publication Number: US-11028712-B2

Title: Seal support feature for brush seals

Description:
FIELD 
     The present disclosure relates to gas turbine engines, and, more specifically, to a seal between a blade outer air seal of a turbine section or a compressor section. 
     BACKGROUND 
     A gas turbine engine typically includes a fan section, a compressor section, a combustor section, and a turbine section. Air moves into the engine through the fan section. Airfoil arrays in the compressor section rotate to compress the air, which is then mixed with fuel and combusted in the combustor section. The products of combustion are expanded to rotatably drive airfoil arrays in the turbine section. Rotating the airfoil arrays in the turbine section drives rotation of the fan and compressor sections. 
     Gas turbine engines may have various gas-flow streams that may be kept separate from one another. The gas-flow streams may be separated by various components. Higher pressure gaspath air may tend to leak between airfoil arrays into lower pressure gaspaths. The internal engine environment is exposed to temperature extremes and other harsh environmental conditions, which may affect the integrity of the components separating different gas-flow streams. The loss of secondary flow air into the gas-path of a turbine has a negative effect on engine fuel burn, performance/efficiency, and component life. 
     SUMMARY 
     A seal assembly for a gas turbine engine is described herein, in accordance with various embodiments. The seal assembly may include a first engine component including a flange, the first engine component disposed between a first airflow path and a second airflow path, a first seal disposed between the first airflow path and the flange, and a second seal disposed between the first airflow path and the second airflow path, wherein the second seal is disposed in a series arrangement with the first seal, and the second seal mechanically supports the first seal. 
     In various embodiments, the first airflow path is at a pressure greater than a pressure of the second airflow path. 
     In various embodiments, the first seal and the second seal are disposed in a third airflow path between the first airflow path and the second airflow path. 
     In various embodiments, the first seal comprises a “W” seal. 
     In various embodiments, the second seal comprises a brush seal. 
     In various embodiments, the seal assembly further comprises a second engine component disposed between the first airflow path and the second airflow path, wherein the third airflow path is disposed between the first engine component and the second engine component. 
     In various embodiments, the first engine component comprises a blade outer air seal and the second engine component comprises a stator vane. 
     A turbine section of a gas turbine engine is disclosed, comprising a stator vane, a blade outer air seal adjacent to the stator vane, a first seal disposed between the stator vane and the blade outer air seal, and a second seal disposed between the stator vane and the blade outer air seal, wherein the second seal is disposed in a series arrangement with the first seal, and the second seal mechanically supports the first seal during engine operation. 
     In various embodiments, the blade outer air seal comprises a flange extending toward the stator vane. 
     In various embodiments, the second seal is disposed on the flange. 
     In various embodiments, the blade outer air seal comprises a seal support extending toward the stator vane. 
     In various embodiments, the first seal is disposed on the seal support. 
     In various embodiments, the second seal comprises a second seal support extending radially from the second seal, a radially outer surface of the second seal support is substantially flush with a radially outer surface of the seal support. 
     In various embodiments, the first seal comprises a “W” seal. 
     In various embodiments, the second seal includes a brush seal. 
     In various embodiments, the first seal is mechanically supported by a radially outer surface of the second seal. 
     A gas turbine engine is disclosed, comprising a turbine section or a compressor section including a stator vane, a blade outer air seal adjacent to the stator vane, a first seal disposed between the stator vane and the blade outer air seal, and a second seal disposed between the stator vane and the blade outer air seal, wherein the second seal is disposed in a series arrangement with the first seal, and the second seal mechanically supports the first seal during engine operation. 
     In various embodiments, the first seal and the second seal are disposed between a first airflow path and a second airflow path and wherein the first airflow path is at a pressure greater than a pressure of the second airflow path. 
     In various embodiments, the first seal comprises a “W” seal and the second seal includes a brush seal. 
     In various embodiments, the blade outer air seal comprises a seal support extending toward the stator vane, the first seal disposed on the seal support, and the second seal comprises a second seal support extending radially from the second seal, a radially outer surface of the second seal support is substantially flush with a radially outer surface of the seal support. 
     The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, the following description and drawings are intended to be exemplary in nature and non-limiting. 
    
    
     
       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 detailed description and claims when considered in connection with the figures, wherein like numerals denote like elements. 
         FIG. 1  illustrates a cross-sectional view of an exemplary gas turbine engine, in accordance with various embodiments; 
         FIG. 2A  illustrates a cross-sectional view of a turbine section of a gas turbine engine including rotor-stator assembly, in accordance with various embodiments; 
         FIG. 2B  illustrates a cross-sectional view of a blade outer air seal and a stator vane, in accordance with various embodiments; 
         FIG. 3  illustrates cross-sectional views of a seal assembly with a support structure for a first seal, in accordance with various embodiments; 
         FIG. 4  illustrates a perspective cross-sectional view of a brush seal, in accordance with various embodiment; 
         FIG. 5  illustrates a flow chart for a method for manufacturing a brush seal, in accordance with various embodiments; and 
         FIG. 6  illustrates a flow chart for a method for installing a brush seal arrangement, in accordance with various embodiments. 
     
    
    
     DETAILED DESCRIPTION 
     All ranges and ratio limits disclosed herein may be combined. It is to be understood that unless specifically stated otherwise, references to “a,” “an,” and/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. 
     The 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 logical, chemical, and mechanical changes may be made without departing from the spirit and scope of the disclosure. Thus, the detailed description herein is presented for purposes of illustration only and not of limitation. For example, the steps recited in any of the method or process descriptions may be executed in any order and are not necessarily limited to the order presented. 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, and/or any other possible attachment option. Additionally, any reference to without contact (or similar phrases) may also include reduced contact or minimal contact. Cross hatching lines may be used throughout the figures to denote different parts but not necessarily to denote the same or different materials. 
     As used herein, “aft” refers to the direction associated with the tail (e.g., the back end) of an aircraft, or generally, to the direction of exhaust of the gas turbine engine. As used herein, “forward” refers to the direction associated with the nose (e.g., the front end) of an aircraft, or generally, to the direction of flight or motion. 
     As used herein, “distal” refers to the direction radially outward, or generally, away from the axis of rotation of a turbine engine. As used herein, “proximal” refers to a direction radially inward, or generally, towards the axis of rotation of a turbine engine. 
     In various embodiments and with reference to  FIG. 1 , a gas turbine engine  20  is provided. Gas turbine engine  20  may be a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines may include other systems or features. In operation, fan section  22  can drive coolant (e.g., air) along a bypass flow-path B while compressor section  24  can drive coolant along a core flow-path C for compression and communication into combustor section  26  then expansion through turbine section  28 . Although depicted as a turbofan gas turbine engine  20  herein, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     Gas turbine engine  20  may generally comprise a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A-A′ relative to an engine static structure  36  or engine case via several bearing systems  38 ,  38 - 1 , and  38 - 2 . Engine central longitudinal axis A-A′ is oriented in the z direction on the provided xyz axis. It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided, including for example, bearing system  38 , bearing system  38 - 1 , and bearing system  38 - 2 . 
     Low speed spool  30  may generally comprise an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a low pressure turbine  46 . Inner shaft  40  may be connected to fan  42  through a geared architecture  48  that can drive fan  42  at a lower speed than low speed spool  30 . Geared architecture  48  may comprise a gear assembly  60  enclosed within a gear housing  62 . Gear assembly  60  couples inner shaft  40  to a rotating fan structure. High speed spool  32  may comprise an outer shaft  50  that interconnects a high pressure compressor  52  and high pressure turbine  54 . A combustor  56  may be located between high pressure compressor  52  and high pressure turbine  54 . A mid-turbine frame  57  of engine static structure  36  may be located generally between high pressure turbine  54  and low pressure turbine  46 . Mid-turbine frame  57  may support one or more bearing systems  38  in turbine section  28 . Inner shaft  40  and outer shaft  50  may be concentric and rotate via bearing systems  38  about the engine central longitudinal axis A-A′, which is collinear with their longitudinal axes. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
     The core airflow C may be compressed by low pressure compressor  44  then high pressure compressor  52 , mixed and burned with fuel in combustor  56 , then expanded over high pressure turbine  54  and low pressure turbine  46 . Turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
     Gas turbine engine  20  may be, for example, a high-bypass ratio geared aircraft engine. In various embodiments, the bypass ratio of gas turbine engine  20  may be greater than about six (6). In various embodiments, the bypass ratio of gas turbine engine  20  may be greater than ten (10). In various embodiments, geared architecture  48  may be an epicyclic gear train, such as a star gear system (sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear) or other gear system. Geared architecture  48  may have a gear reduction ratio of greater than about 2.3 and low pressure turbine  46  may have a pressure ratio that is greater than about five (5). In various embodiments, the bypass ratio of gas turbine engine  20  is greater than about ten (10:1). In various embodiments, the diameter of fan  42  may be significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  may have a pressure ratio that is greater than about five (5:1). Low pressure turbine  46  pressure ratio may be measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of low pressure turbine  46  prior to an exhaust nozzle. It should be understood, however, that the above parameters are exemplary of various embodiments of a suitable geared architecture engine and that the present disclosure contemplates other gas turbine engines including direct drive turbofans. A gas turbine engine may comprise an industrial gas turbine (IGT) or a geared aircraft engine, such as a geared turbofan, or non-geared aircraft engine, such as a turbofan, or may comprise any gas turbine engine as desired. 
     Still referring to  FIG. 1  and now to  FIG. 2A , according to various embodiments, each of low pressure compressor  44 , high pressure compressor  52 , low pressure turbine  46 , and high pressure turbine  54  in gas turbine engine  20  may comprise one or more stages or sets of rotating blades (“rotors blades”) and one or more stages or sets of stationary vanes (“stator vanes”) axially interspersed with the associated blade stages but non-rotating about engine central longitudinal axis A-A′. The low pressure compressor  44  and high pressure compressor  52  may each comprise one or more compressor stages. The low pressure turbine  46  and high pressure turbine  54  may each comprise one or more turbine stages. Each compressor stage and turbine stage may comprise multiple interspersed stages of rotor blades  70  and stator vanes  72 . The rotor blades  70  rotate about engine central longitudinal axis A-A′ with the associated inner shaft  40  or outer shaft  50  while the stator vanes  72  remain stationary about engine central longitudinal axis A-A′. For example,  FIG. 2A  schematically shows, by example, a turbine stage of turbine section  28  of gas turbine engine  20 . Unless otherwise indicated, the term “blade stage” refers to at least one of a turbine stage or a compressor stage. The compressor and turbine sections  24 ,  28  may comprise rotor-stator assemblies. 
     With reference to  FIG. 2A , a portion of turbine section  28  is illustrated, in accordance with various embodiments. Rotor blade  70  may be, for example, a turbine rotor including a circumferential array of blades configured to be connected to and rotate with a rotor disc about engine central longitudinal axis A-A′. Upstream (forward) and downstream (aft) of rotor blade  70  are stator vanes  72 , which may be, for example, turbine stators including circumferential arrays of vanes configured to guide core airflow C flow through successive turbine stages, such as through rotor blade  70 . A radially outer portion  74  of stator vane  72  may be coupled to engine case structure  36 . 
     According to various embodiments, compressor and turbine rotors may include stationary annular fluid seals, referred to as a blade outer air seal (BOAS) assembly, circumscribing the rotor blades to contain and direct core airflow C. A BOAS assembly may include one or more BOAS segments  80  circumferentially arranged to form a ring about engine central longitudinal axis A-A′ radially outward of rotor blades  70 . Although only one BOAS segment  80  is shown in  FIG. 2 , turbine section  28  comprises an associated array of BOAS segments  80 . A BOAS segment  80  may be disposed radially outward of a rotor blade  70  and/or a plurality of rotor blades  70  relative to engine central longitudinal axis A-A′. Each BOAS segment  80  may couple to an adjacent BOAS segment  80  to form the annular BOAS assembly. Each BOAS segment  80  may further couple to engine case structure  36 . A BOAS assembly may thus comprise an annular structure comprising a plurality of BOAS segments  80 , with each BOAS segment  80  disposed radially about one or more of a plurality of rotors blades  70 . According to various embodiments, BOAS segments  80  may, in various embodiments, be formed as a unitary BOAS structure, with the same features described herein. 
     In various embodiments, BOAS segments  80  are coupled to axially adjacent stator vanes  72 .  FIG. 2A  shows an area within turbine section  28  that includes BOAS segment or segments  80  disposed between forward and aft stator vanes  72 . A BOAS segment  80  locally defines a radially outward boundary of core airflow path C through turbine section  28 . During engine operation, stator vanes  72  and BOAS segments  80  are subjected to different thermal loads and environmental conditions. Cooling air may be provided to BOAS segment  80  and stator vanes  72  to enable operation of the turbine during exposure to hot combustion gasses produced within the combustion area, as described above. Referring momentarily to  FIG. 1 , pressurized air may be diverted from combustor section  26  and/or compressor section  24  and used to cool components within the turbine section  28 . The BOAS assembly and stator vanes  72  may be in fluid communication with a secondary airflow source, such as an upstream compressor in the compressor section  24  or other source, which provides cooling airflow, such as bleed compressor air. A BOAS segment  80  and stator vanes  72  may be coupled to engine case structure  36  and may define a secondary airflow path S between engine case structure  36  and BOAS segment  80 . A secondary airflow S is shown flowing axially downstream between engine case structure  36  and radially outer portion  74  of stator vanes  72 . Secondary airflow S provides varying levels of cooling to different areas of BOAS segment  80 , where a plurality of segmented seals, BOAS segments  80 , may be disposed circumferentially around blades  70 . A BOAS assembly may thus comprise an annular structure comprising a plurality of BOAS segments  80 , with each BOAS segment  80  disposed radially about one or more of a plurality of rotors blades  70  and stator vanes  72 . Secondary airflow S may have a higher pressure than core airflow C. A difference in pressure of secondary airflow S may typically be about 100-200 PSI (689-1379 kPa) greater than core airflow C at BOAS segment  80 , creating a tendency for secondary airflow S to leak past BOAS segment  80  into core airflow C. As rotor blade  70  and stator vanes  72  are heated by exhaust gasses exiting the combustion chamber in core airflow C, the cooling air of secondary airflow S may undesirably escape into the turbine through a variety of gaps and other pathways core airflow C. Effective seals may therefore be incorporated between stages of BOAS segment  80  and stator vanes  72  to prevent, control, or reduce leakage of cooling air from secondary airflow S cavities into core airflow path C. 
     Referring to  FIG. 2B , BOAS segment  80  may be coupled to an axially adjacent stator vane  72 , in accordance with various embodiments. An axial separation may exist between BOAS segment  80  and stator vane  72 . For example, as shown, stator vane  72  may be axially separated from the BOAS segment  80  by a distance or gap  88 . Gap  88  may expand and contract (axially and/or radially) in response to the thermal and/or mechanical environment. In addition, gap  88  may expand and/or contract (axially and/or radially) as a result of thermal, mechanical, and pressure loading imparted in BOAS segment  80 , stator vane  72 , and/or supporting structure during various transient and steady state engine operating conditions. 
     Cooling air from secondary airflow S may tend to leak through the distance or gap  88  between BOAS segment  80  and stator vane  72  in response to a pressure differential. Thus, a seal assembly  90  may be disposed between BOAS segment  80  and stator vane  72  to prevent, reduce, and/or control leakage of secondary airflow S through gap  88  into core airflow path C. Seal assembly  90  may include a plurality of annular seals, as described herein, and may be placed between BOAS segment  80  and stator vane  72  to limit leakage of secondary airflow S between BOAS segment  80  and stator vane  72  and into core airflow path C. Leakage of secondary airflow S into core airflow path C is illustrated as leakage airflow path L. Seal assembly  90  may form a partial seal or a complete seal between a BOAS segment  80  of stator vane  72 , thereby reducing or eliminating leakage airflow L. 
     In various embodiments, a first engine component  180  and a second engine component  172  may be disposed between a first airflow path, such as secondary airflow path S, and a second airflow path, such as core airflow path C. The first engine component  180  and second engine component  172  may comprise any stationary or rotating engine component. The first engine component  180  may include BOAS segment  80  and the second engine component  172  may include stator vane  72 . A third airflow path, such as leakage airflow L, may be disposed between the first engine component  180  and the second engine component  172 . 
     According to various embodiments, the BOAS segment  80  may include a main body  82  that extends generally axially from a forward portion (not shown) to an aft portion  84 . BOAS segment  80  may also include at least one attachment portion or flange  86  disposed at or near the aft portion  84 . Flange  86  may extend in an axially aft direction from main body  82  toward stator vane  72 . Flange  86  may alternatively extend in an axially forward direction, or in an x direction or y direction. Each axially extending flange  86  of BOAS segment  80  may correspond to a receiving portion  76  of stator vane  72  to support and attach BOAS segment  80 . Stator vane  72  may also include at least one attachment portion or flange  78  disposed at or near a forward edge portion  79  of stator vane  72 . BOAS segment  80  may further be configured to receive flange  78  of stator vane  72 . Aft portion  84  of BOAS segment  80  and forward edge portion  79  of stator vane  72  interface to form a cavity  92  or a plurality of cavities  92 . Cavity  92  may be configured to house a seal assembly  90 . 
     In various embodiments, seal assembly  90  includes a first seal  102  and a second seal  104 . First seal  102  and second seal  104  may be disposed within cavity  92  between BOAS segment  80  and stator vane  72 . First seal  102  may be located distally with respect to second seal  104 . In various embodiments, first seal  102  may be disposed in a series arrangement with second seal  104  with respect to a leakage airflow path L. As secondary airflow S enters gap  88 , between BOAS segment  80  and stator vane  72 , the leakage airflow L contacts first seal  102  prior to reaching second seal  104  and if leakage airflow L is blocked by first seal  102 , the leakage airflow L does not reach second seal  104 . Seal assembly  90  may thus prevent or reduce the passage of cooling and/or secondary airflow S as leakage airflow L through the gap  88  defined between BOAS segment  80  and stator vane  72 . 
     In various embodiments, first seal  102  may include a “W” seal (e.g., a seal having a “W”-shaped cross-section or that forms a “W” shape), an “S” seal (e.g., a seal having a “S”-shaped cross-section or that forms a “S” shape), or other similarly shaped seal. Second seal  104  may include a “W” seal, a dog bone seal, a brush seal, a rope seal, a “C” seal, a crush seal, a flap seal, a feather seal, or other suitable seal. Second seal  104  may comprise a different seal type than first seal  102 . 
     With reference to  FIG. 3 , additional detail of a seal assembly  90  is shown in accordance with various embodiments. First seal  102  and second seal  104  may be configured in a series arrangement to operate under high temperature and high pressure conditions, thereby providing an improved seal within cavity  92  to reduce leakage airflow L. 
     In various embodiments, first seal  102  may include a “W” seal (e.g., seals having a “W”-shaped cross-section or that form a “W” shape). First seal  102  may include a metal, such as titanium, titanium-based alloy, nickel, nickel-based alloy, aluminum, aluminum-based alloy, steel, or stainless steel, or other materials. A “W” seal may be flexible enough axially to expand and seal against the walls of BOAS segment  80  and stator vane  72 . First seal  102  may be biased outwardly along engine central longitudinal axis A-A′ to contact aft portion  84  (see  FIG. 2B ) of BOAS segment  80  and forward edge portion  79  (see  FIG. 2B ) of stator vane  72 . The “W” shape may enable first seal  102  to compress and extend as stator vane  72  moves relative to BOAS segment  80  in response to thermally driven deformations and pressure loads. Thus, first seal  102  prevents or greatly reduces leakage airflow L passing through or around first seal  102 . 
     In various embodiments, second seal  104  may be made of materials that are capable of enduring and/or surviving in environments with relatively high temperatures associated with the various thermal loads and/or heat loads from core airflow C. Components near core airflow C may be exposed to and/or reach temperature of more than 2000° F. (approximately 1093° C.). Seal materials that are capable of surviving in environments with relatively high temperatures may generally have lower strength properties making the seals more susceptible to permanent deformation, failure, and/or liberation. In various embodiments, second seal  104  may include a brush seal. A brush seal, such as brush seal  110  in  FIG. 5 , may be mounted axially to the BOAS segment  80  to extend axially between BOAS segment  80  and stator vane  72 . Second seal  104  may extend axially beyond an aft end  106  of flange  86  to interface with stator vane  72 . 
     Referring now to  FIG. 4 , a brush seal is shown, in accordance with various embodiments. A brush seal  110  may be used for second seal  104 . A brush seal may include a bristle-pack  112  that is compressed, or otherwise engaged between two metallic sheets as a top plate  114  and a backing plate  116 . Bristle-pack  112  may include one or more sizes of bristle wire  112   a - 112   b . Each wire of the bristle-pack  112 , as well as top plate  114  and backing plate  116 , may comprise a metal, high temperature metal, cobalt alloy, high performance nickel-chromium alloy such as an austenitic nickel-chromium-based alloy (e.g., an alloy having a nominal composition of nickel fifty-eight percent (58%), chromium 20% to 23%, iron up to 5%, molybdenum between 8% to 10%, niobium (plus tantalum) between 3.15% to 4.15% that is available under the trade name INCONEL 625™, available from Special Metals Corporation of New Hartford, N.Y., USA), or any suitable metallic, non-metallic material, and/or any other composite or alloy material. Brush seal  110  may be a continuous annular ring or may have at least one circumferential split or splice that maintains a pressure loaded, radial, contact regardless of relative thermal displacements. 
     In various embodiments, brush seal  110  includes a seal support  200  (also referred to herein as a second seal support) extending from top plate  114 . Seal support  200  may extending radially from top plate  114 . Seal support  200  may include a flange extending radially from top plate  114 . Seal support  200  may be disposed at an edge  212  of top plate  114 . In various embodiments, edge  212  is an aft edge of top plate  114 . Top plate  114  and seal support  200  may be manufactured as a monolithic piece. 
     Returning to  FIG. 3 , BOAS segment  80  may further include a seal support  120  (also referred to herein as a first seal support) extending axially aft toward receiving portion  76  of stator vane  72 , similarly to flange  86 . Seal support  120  may comprise a mechanical support base for first seal  102 . The radially outer surface  202  of seal support  200  may comprise a mechanical support base for a portion of first seal  102 . The radially outer (i.e., cylindrical) surface  122  of seal support  120  may be substantially flush with the radially outer surface  202  of seal support  200 . First seal  102  and second seal  104  may generally be free-floating within cavity  92 . A pressure differential between secondary airflow path S and core airflow path C may load first seal  102  against the radially outer surface  122  of seal support  120  and the radially outer surface  202  of seal support  200 , and may load second seal  104  against the radially outer (i.e., cylindrical) surface  108  of flange  86 . As the pressure differential between secondary airflow path S and core airflow path C loads first seal  102  against radially outer surface  122 , the end  103  of first seal  102  may tend to bend or flex radially inwards, which may damage first seal  102  and/or reduce the efficiency of first seal  102 . In this regard, seal support  200  may provide a mechanical support base to prevent the end  103  of first seal  102  from bending or flexing radially inward during engine operation. As BOAS segment  80  and stator vane  72  move axially and radially relative to each other, the brush ends of bristle-pack  112  deflect against and slide along the mating surfaces to maintain sealing. 
     In various embodiments, first seal  102  may be configured to withstand higher pressures than second seal  104  while maintaining an effective seal, while second seal  104  may be configured to withstand higher temperatures than first seal  102 . Second seal  104  may be located between first seal  102  and core airflow path C, thereby protecting first seal  102  from exposure to hot gas and reducing thermal fatigue. First seal  102  may be located between second seal  104  and secondary airflow path S, thereby protecting second seal  104  from high pressures. Therefore, seal assembly  90  including first seal  102  and second seal  104  may optimize the sealing effectiveness and operating capabilities with a series seal arrangement. 
     Additionally, during standard operation of the gas turbine engine, in response to receiving a flow of fluid, such as secondary airflow path S and core airflow path C, stator vane  72  may move axially, radially, and/or circumferentially relative to BOAS segment  80 . By providing seal support  200  for supporting seal  102 , first seal  102  experiences less radial deflection, thereby extending the operational life of first seal  102 . Additionally, by resting first seal  102  directly on seal support  200 , the axial length (measured in the Z-direction) of seal support  120  may be minimized to stiffen seal support  120 . In various embodiments, the aspect ratio (i.e., the axial length) of seal support  120  may be limited due to adverse vibratory responses in response to lengthening seal support  120 . 
     Thus, well-sealed gaps between components along the radially inward direction (i.e., along they axis), such as BOAS segment  80  and stator vane  72 , reduce leakage airflow L into airflow path C and reduce negative performance impacts (such as efficiency). In addition, while a turbine stage is depicted in  FIG. 2A , it is to be understood that various embodiments may be utilized for static gas turbine engine components in any turbine stage of the high pressure turbine or low pressure turbine and in any compressor stage of the high pressure compressor or the low-pressure compressor. While a BOAS segment having specially configured seal assembly for deflecting core airflow C air has been described in accordance with various embodiments, it is to be understood that other gas turbine engine components may benefit from an optimized seal assembly according to various embodiments. 
     In addition to providing support for first seal  102 , seal support  200  may tend to prevent second seal  104  from being installed backwards over flange  86 . In this regard, seal support  200  provide mistake-proofing for installation of second seal  104 . 
     With reference to  FIG. 5 , a method  500  for manufacturing a brush seal is provided, in accordance with various embodiments. Method  500  includes forming a radial extension into an annular top plate (step  510 ). Method  500  includes coupling a bristle wire between the annular top plate and an annular backing plate (step  520 ). 
     With combined reference to  FIG. 4  and  FIG. 5 , step  510  may include forming radial extension  203  (i.e., seal support  200 ) into annular top plate  114 . In various embodiments, step  520  is performed using an additive manufacturing process, including but not limited to 3D printing and/or welding. In various embodiments, step  520  is performed using a subtractive manufacturing process, including but not limited to machining, milling, and/or cutting. Step  520  may include coupling bristle pack  112  between the annular top plate  114  and annular backing plate  116 . In various embodiments, step  520  is performed using a metal joining process such as welding, brazing, and/or soldering. 
     With reference to  FIG. 6 , a method  600  for installing a brush seal is provided, in accordance with various embodiments. Method  600  includes disposing a brush seal into a cavity between a first engine component and a second engine component (step  610 ). Method  600  includes disposing a radially outer surface of a seal support extending from the brush seal to be substantially flush with a radially outer surface of a seal support (step  620 ). Method  600  includes disposing a second seal radially from the radially outer surface (step  630 ). 
     With combined reference to  FIG. 3  and  FIG. 6 , step  610  may include disposing second seal  104  into cavity  92  between first engine component  180  and a second engine component  172 . Step  620  may include disposing radially outer surface  202  of seal support  200  extending from second seal  104  to be substantially flush with radially outer surface  122  of seal support  120 . Step  630  may include disposing first seal  102  radially (e.g., radially outward) from radially outer surface  122 . 
     Benefits and other advantages 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, and any elements that may cause any benefit or advantage 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. 
     Systems, methods and apparatus are provided herein. In the detailed description herein, references to “various embodiments”, “one embodiment”, “an embodiment”, “an example embodiment”, 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 is intended to invoke 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.