Patent Publication Number: US-2016245518-A1

Title: Combustor panel with multiple attachments

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     This application claims priority to U.S. Patent Application No. 61/886,987 filed Oct. 4, 2013, which is hereby incorporated herein by reference in its entirety. 
    
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     This disclosure was made with Government support under FA8650-09-D-2923 0021 awarded by the United States Air Force. The Government may have certain rights in this disclosure. 
    
    
     BACKGROUND 
     The present disclosure relates to a gas turbine engine and, more particularly, to a combustor section therefor. 
     Gas turbine engines, such as those that power modem commercial and military aircraft, generally include a compressor section to pressurize an airflow, 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. 
     The combustor section typically includes an inner and outer double wall assembly each with an outer shell lined with heat shields to define an annular combustion chamber. The floatwall liner panels are attached to the outer shell with studs and nuts. These attachments have been sufficient and achievable due to the spatial availability of legacy engine diffuser cases. In certain engine architectures, however, these attachments may pose packing issues with regard to the engine case structure. 
     SUMMARY 
     A liner panel for a gas turbine engine, according to one disclosed non-limiting embodiment of the present disclosure, includes a first attachment and a second attachment that extends from a cold side of the liner panel. The first attachment is different than the second attachment. 
     In a further embodiment of the present disclosure, the first attachment includes a stud and a clip engageable with the stud. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the first attachment is upstream of the second attachment. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the second attachment includes a threaded stud and a nut. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the first attachment includes a stud and a clip engageable with the clip. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the second attachment includes a trapped interface. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the first attachment includes a trapped interface. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the second attachment includes a stud and a clip engageable with the stud. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the first attachment is upstream of the second attachment. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the liner panel is manufactured of a ceramic. 
     A combustor of a gas turbine engine, according to another disclosed non-limiting embodiment of the present disclosure, includes an inner combustor wall assembly with an inner support shell and a multiple of inner liner panels. At least one of the multiple of inner liner panels includes a first inner attachment to the inner support shell and a second inner attachment to the inner support shell. The first inner attachment different than, and upstream of, the second inner attachment. An outer combustor wall assembly is spaced from the inner combustor wall assembly to define an annular combustion cavity therebetween. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the first inner attachment includes a stud and a clip engageable with the stud. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the second inner attachment includes a threaded stud and a nut. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the second inner attachment includes a trapped interface. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, a bulkhead assembly is mounted to the inner combustor wall assembly and the outer combustor wall assembly. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the first inner attachment includes a trapped interface with a bulkhead panel of the bulkhead assembly. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the multiple of inner liner panels are manufactured of a ceramic. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the bulkhead panel is manufactured of a metallic alloy. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the outer combustor wall assembly includes an outer support shell and a multiple of outer liner panels. At least one of the multiple of outer liner panels includes a first outer attachment to the outer support shell and a second outer attachment to the outer support shell. The first outer attachment equivalent to the second outer attachment. 
     In a further embodiment of any of the foregoing embodiments of the present disclosure, the first outer attachment and the second outer attachment each includes a threaded stud and a nut. 
     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 
       Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG. 1  is a schematic cross-section of an example gas turbine engine architecture; 
         FIG. 2  is a schematic cross-section of another example gas turbine engine architecture; 
         FIG. 3  is an expanded longitudinal schematic sectional view of a combustor section according to one non-limiting embodiment that may be used with the example gas turbine engine architectures shown in  FIGS. 1 and 2 ; 
         FIG. 4  is an expanded longitudinal schematic sectional view of a combustor section with an inner wall assembly having different first and second attachments according to one disclosed non-limiting embodiment; 
         FIG. 5  is a perspective view of an annular combustor with annular inner and outer wall assemblies; 
         FIG. 6  is a sectional view of the annular combustor of  FIG. 5  along an engine axis; 
         FIG. 7  is an expanded exploded view of one liner panel attachment according to one disclosed non-limiting embodiment; 
         FIG. 8  is an expanded exploded view of one liner panel attachment according to one disclosed non-limiting embodiment; 
         FIG. 9  is an expanded longitudinal schematic sectional view of a combustor section with an inner wall assembly having different first and second attachments according to another disclosed non-limiting embodiment; 
         FIG. 10  is a perspective view of one liner panel attachment according to one disclosed non-limiting embodiment; 
         FIG. 11  is a perspective view of one liner panel attachment according to one disclosed non-limiting embodiment; 
         FIG. 12  is a longitudinal sectional view of one liner panel attachment according to one disclosed non-limiting embodiment; 
         FIG. 13  is a longitudinal sectional view of one liner panel attachment according to one disclosed non-limiting embodiment; and 
         FIG. 14  is an expanded longitudinal schematic sectional view of a combustor section with an inner wall assembly having different first and second attachments according to another disclosed non-limiting embodiment. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbo fan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Referring to  FIG. 2 , another alternative engine architecture  200  might include an augmentor section  12 , an exhaust duct section  14  and a nozzle section  16  in addition to the fan section  22 ′, compressor section  24 ′, combustor section  26 ′ and turbine section  28 ′. Referring again to  FIG. 1 , although depicted as an aero engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not so limited and the teachings may be applied to other types of turbine engines such as a turbojets, turboshafts, and three-spool (plus fan) turbofans with an intermediate spool as well as industrial gas turbines. 
     The fan section  22  drives air along a bypass flowpath and into the compressor section  24 . The compressor section  24  drives air along a core flowpath for compression and communication into the combustor section  26 , which then expands and directs the air through the turbine section  28 . The engine  20  generally includes a low spool  30  and a high spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing structures  38 . The low spool  30  generally includes an inner shaft an inner shaft  40  that interconnects a fan  42 , a low pressure compressor (“LPC”)  44  and a low pressure turbine (“LPT”)  46 . The inner shaft  40  may drive the fan  42  directly or through a geared architecture  48  as illustrated in  FIG. 1  to drive the fan  42  at a lower speed than the low spool  30 . An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system. 
     The high spool  32  includes an outer shaft  50  that interconnects a high pressure compressor (“HPC”)  52  and a high pressure turbine (“HPT”)  54 . A combustor  56  is arranged between the HPC  52  and the HPT  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     Core airflow is compressed by the LPC  44  then the HPC  52 , mixed with the fuel and burned in the combustor  56 , then expanded over the HPT  54  and the LPT  46 . The LPT  46  and HPT  54  rotationally drive the respective low spool  30  and high spool  32  in response to the expansion. 
     In one non-limiting example, the gas turbine engine  20  is a high-bypass geared aircraft engine with a bypass ratio greater than about six (6:1). The geared architecture  48  can include an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.3:1, and in another example, is greater than about 2.5:1. The geared turbofan enables operation of the low spool  30  at higher speeds which can increase the operational efficiency of the LPC  44  and LPT  46  to render increased pressure in a fewer number of stages. 
     A pressure ratio associated with the LPT  46  is pressure measured prior to the inlet of the LPT  46  as related to the pressure at the outlet of the LPT  46  prior to an exhaust nozzle of the gas turbine engine  20 . In another non-limiting example, the bypass ratio of the gas turbine engine  20  is greater than about ten (10:1), the fan diameter is significantly larger than that of the LPC  44 , and the LPT  46  has a pressure ratio greater than about five (5:1). It should be appreciated, however, that the above parameters are only exemplary of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans. 
     In an example high-bypass turbofan, significant thrust is provided by the high bypass ratio as the fan section  22  may be designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine  20  at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC) which is an industry standard parameter of fuel consumption per unit of thrust. 
     Fan Pressure Ratio is the pressure ratio across a fan blade of the fan section  22  without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one example gas turbine engine  20  is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of (“Tram”/518.7) 0.5 . The Low Corrected Fan Tip Speed according to the example gas turbine engine  20  is less than about 1150 fps (351 m/s). 
     With reference to  FIG. 3 , the combustor section  26  generally includes a combustor  56  with an outer combustor wall assembly  60 , an inner combustor wall assembly  62  and a diffuser case module  64 . The outer combustor wall assembly  60  and the inner combustor wall assembly  62  are spaced apart such that an annular combustion chamber  66  is defined therebetween. 
     The outer combustor wall assembly  60  is spaced radially inward from an outer diffuser case  64 A of the diffuser case module  64  to define an outer annular plenum  76 . The inner combustor wall assembly  62  is spaced radially outward from an inner diffuser case  64 B of the diffuser case module  64  to define an inner annular plenum  78 . It should be understood that although a particular combustor is illustrated, other combustor types with various combustor wall and diffuser case module arrangements will also benefit herefrom. 
     The combustor wall assemblies  60 ,  62  contain the combustion products for direction toward the turbine section  28 . Each combustor wall assembly  60 ,  62  generally includes a respective outer and inner support shell  68 ,  70  which supports one or more liner panels  72 ,  74  mounted within the respective support shell  68 ,  70 . Each of the liner panels  72 ,  74  may be generally rectilinear with a circumferential arc and manufactured of, for example, a nickel based super alloy, ceramic or other temperature resistant material and are arranged to form a liner array of circumferentially and/or radially staggered liner panels  72 ,  74 . 
     The combustor  56  further includes a forward assembly  80  immediately downstream of the compressor section  24  to receive compressed airflow therefrom. The forward assembly  80  generally includes an annular hood  82  and a bulkhead assembly  84  that support a multiple of fuel nozzles  86  (one shown) and a multiple of swirlers  90  (one shown) along an axis F. The annular hood  82  extends radially between, and is secured to, the forwardmost ends of the combustor wall assemblies  60 ,  62 . The annular hood  82  includes a multiple of circumferentially distributed hood ports  94  that accommodate the respective fuel nozzle  86  and introduce air into the forward end of the combustion chamber  66  through a respective swirler  90 . The bulkhead assembly  84  includes a bulkhead support shell  96  secured to the combustor wall assemblies  60 ,  62 , and a multiple of circumferentially distributed bulkhead liner panels  98  secured to the bulkhead support shell  96 . Each fuel nozzle  86  may be secured to the diffuser case module  64  and project through one of the hood ports  94  and respective swirlers  90 . 
     The forward assembly  80  introduces core combustion air into the forward section of the combustion chamber  66  while the remainder enters the outer annular plenum  76  and the inner annular plenum  78 . The multiple of fuel nozzles  86  and adjacent structure generate a fuel-air mixture that supports stable combustion in the combustion chamber  66 . 
     Opposite the forward assembly  80 , the outer and inner support shells  68 ,  70  are mounted adjacent to a first row of Nozzle Guide Vanes (NGVs)  54 A in the HPT  54 . The NGVs  54 A are static engine components which direct core airflow combustion gases onto turbine blades in the turbine section  28  to facilitate the conversion of pressure energy into kinetic energy. The core airflow combustion gases are also accelerated by the NGVs  54 A because of their convergent shape and are typically given a “spin” or a “swirl” in the direction of turbine rotor rotation. 
     With reference to  FIG. 4 , the inner liner panels  74  (only one shown) include a first attachment  100  and a second attachment  102 , where the first attachment  100  is upstream of the second attachment  102 . That is, the first attachment  100  is closer to the forward assembly  80  than the second attachment  102 . The first attachment  100  and the second attachment  102  attach each of the respective inner liner panel  74  to the inner support shell  70 . 
     The combustor wall assemblies  60 ,  62  of the annular combustor  56  ( FIG. 5 ) are canted outward from the forward assembly  80  with respect to the interface NGVs  54 A such that the first attachment  100  and the second attachment  102  are sized to permit passage of the annular combustor  56  through the diffuser case module  64  for assembly and maintenance. That is, the diffuser case module  64  defines an opening  104  with an inner boundary W (illustrated schematically by a phantom line) through which the annular combustor  56  with the first attachment  100  and the second attachment  102  are sized to pass (also shown in  FIG. 6 ). That is, the upstream first attachment  100  is sized to be radially smaller than the second attachment  102  so that the first attachment  100  will not extend beyond inner boundary W to permit axial aft passage of the outward canted inner combustor wall assembly  62 . 
     With continued reference to  FIG. 4 , according to one disclosed non-limiting embodiment, the first attachment  100  includes a multiple of studs  110  each with an aperture  112  to receive respective clip  114  ( FIG. 7 ) such as an “R-clip” or other such retainer. The second attachment  102  includes a multiple of studs  116  each with a threaded section  118  to receive respective fastener  120  such as a nut ( FIG. 8 ). The studs  110 ,  116  project rigidly from a cold side  122  of the liner panels  74  opposite a hot side which is directly exposed to combustion gases. The studs  110 ,  116  extend through the support shell  70  to receive the respective clips  114  and fasteners  120  to secure the liner panels  74  thereto. 
     A rail  124  typically extends at least partially around the periphery of the cold side  122  to interface with the support shell  70  when mounted thereto and to define one or more impingement cavities  126 . That is, the rails  120  at least extend around the cold side  122  periphery and may include further internal rails to define additional compartments. 
     The studs  110  of the first attachment  100  extends a shorter distance from the cold side  122  than the studs  116  of the second attachment  102  so that the first attachment  100  will not extend beyond inner boundary W and permit axial passage of the outward canted inner combustor wall assembly  62 . That is, the first attachment  100  and the second attachment  102  permit passage through opening  104  ( FIG. 6 ). 
     As the annular combustor  56  is canted outward, the outer wall assembly  60 , in this disclosed non-limiting embodiment, may include a third attachment  130  that is equivalent to the downstream fourth attachment  132 . The third attachment  130  and the fourth attachment  132  may, for example, include relatively conventional studs  134  each with a threaded section  136  to receive respective fasteners  138  such as a nut. It should be appreciated that various attachments may be utilized for the outer wall assembly  60  either similar or different than that of the inner wall assembly  62 . 
     With reference to  FIG. 9 , according to another disclosed non-limiting embodiment, the second attachment  102 A includes a trapped interface  140 . The trapped interface  140  includes a projection  142  that extends from the support shell  70  into which an aft edge  144  of the liner panel  74  is received. That is, the trapped interface  140  retains the aft end  144  of the liner panel  74 , which is secured by the first attachment  100 A. In this disclosed non-limiting embodiment, the liner panel  74  is manufactured of a ceramic material and the support shell  70  is manufactured of a metal alloy. The support shell  70  permits thermal expansion yet retains the ceramic liner panel  74 . The trapped interface  140  also facilitates elimination of hard bolting that may otherwise potentially lead to ceramic cracking during thermal expansion of the support shell  70 . 
     In this disclosed non-limiting embodiment, the projection  142  includes a radial portion  146  and an axial portion  148  the projection  142  may be segmented ( FIG. 10 ) or continuous ( FIG. 11 ) about the full circumference of the support shell  70 . Further, the axial portion  144  may be coplanar ( FIG. 12 ) or recessed ( FIG. 13 ) with respect to a hot side of the liner panel  74  such that the axial portion  144  need no be directly exposed to combustion gases. 
     With reference to  FIG. 14 , according to another disclosed non-limiting embodiment, the first attachment  100 B includes a trapped interface  150 . The trapped interface  150  includes a forward end  152  of the liner panel  74  that is trapped by the transverse bulkhead liner panels  98 A. That is, the trapped interface  150  retains the forward end  152  of the liner panel  74 , which is then secured by the second attachment  102 B. In this disclosed non-limiting embodiment, the liner panel  74  is manufactured of a ceramic material and the transverse bulkhead liner panels  98 A are manufactured of a metal alloy. As the metal alloy has a higher coefficient to of thermal expansion, the bulkhead liner panels  98 A expands more than the liner panel  74  which is retained by the bulkhead liner panels  98 A. 
     The different attachments arrangements facilitate a relatively lighter weight combustor that can be located in a relatively smaller package space. In addition, threaded studs are minimized which may lead to panel hot spots around rails. 
     The use of the terms “a” and “an” and “the” and similar references in the context of description (especially in the context of the following claims) are to be construed to cover both the singular and the plural, unless otherwise indicated herein or specifically contradicted by context. The modifier “about” used in connection with a quantity is inclusive of the stated value and has the meaning dictated by the context (e.g., it includes the degree of error associated with measurement of the particular quantity). All ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other. It should be appreciated that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting. 
     Although the different non-limiting embodiments have specific illustrated components, the embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments. 
     It should be appreciated that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be appreciated that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. 
     Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure. 
     The foregoing description is exemplary rather than defined by the features within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be appreciated that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.