Gas turbine engine wall assembly with circumferential rail stud architecture

A liner panel for use in a gas turbine engine. The liner panel has a hot side and a cold side. The liner panel includes a rail extending from the cold side and a multiple of studs extending from the cold side. At least one of the multiple of studs extends from, in part, the rail.

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 modern 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.

Among the engine components, relatively high temperatures are observed in the combustor section such that cooling airflow is provided to meet desired service life requirements. The combustor section typically includes a combustion chamber fondled by an inner and outer wall assembly. Each wall assembly includes a support shell lined with heat shields often referred to as liner panels. In certain combustion architectures, dilution passages direct airflow to condition air within the combustion chamber. In addition to the dilution passages, the shells may have relatively small air impingement passages to direct cooling air to impingement cavities between the support shell and the liner panels. This cooling air exits numerous effusion passages through the liner panels to effusion cool the passages and film cool a hot side of the liner panels to reduce direct exposure to the combustion gases.

With lower emission requirements and higher combustor operational temperatures, effective sealing between the shell and liner panels may be of increased significance. However, relatively large tolerances between the cast liner panels and sheet metal shell may complicate such effective sealing.

SUMMARY

A liner panel for use in a gas turbine engine, according to one disclosed non-limiting embodiment of the present disclosure, has a hot side and a cold side. This liner panel includes a rail extending from the cold side, and a multiple of studs extending from the cold side. At least one of the multiple of studs extend from, in part, the rail.

In a further embodiment of the present disclosure, each of the multiple of studs extends from the rail.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the rail is an intermediate circumferential rail.

In a further embodiment of any of the foregoing embodiments of the present disclosure, a perimeter rail is included that extends around the liner panel to at least partially define an impingement cavity therein. The intermediate rail is within the perimeter rail.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the perimeter rail forms a generally rectilinear boundary around the cold side.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the liner panel is rectilinear.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the intermediate rail is circumferentially arranged.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the intermediate rail follows an are formed by the cold side.

In a further embodiment of any of the foregoing embodiments of the present disclosure, a multiple of pedestals are included that extend from the cold side around the at least one of the multiple of studs.

A liner panel for use in a gas turbine engine, according to another disclosed non-limiting embodiment of the present disclosure, includes a forward circumferential rail; an aft circumferential rail axially aft of the forward circumferential rail; a first axial rail and a second axial rail that interconnects the aft circumferential rail and the forward circumferential rail; an intermediate circumferential rail that extends between the first axial rail and the second axial rail; and a multiple of studs that extend, in part, from the intermediate rail.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the intermediate circumferential rail is parallel to the forward circumferential rail and the aft circumferential rail.

In a further embodiment of any of the foregoing embodiments of the present disclosure, a second intermediate circumferential rail is included that extends between the first axial rail and the second axial rail. A second multiple of studs are also included that extend from the second intermediate circumferential rail.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the second intermediate circumferential rail is parallel to the intermediate circumferential rail.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the liner panel is mountable within a combustor of the gas turbine engine.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the liner panel includes a multiple of dilution passages therethrough.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the multiple of dilution passages are located along a line generally parallel to the intermediate circumferential rail.

A method of assembling a wall assembly within a gas turbine engine, according to another disclosed non-limiting embodiment of the present disclosure, includes a locating a stud that extends in part from a rail of a liner panel through a support shell; and attaching a fastener onto the stud to retain the liner panel to the support shell.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the method includes deflecting the support shell toward the liner panel to seal the rail with the support shell.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the method includes defining an impingement cavity between the support shell and the liner panel.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the method includes forming a dilution passage through the impingement cavity.

DETAILED DESCRIPTION

FIG. 1schematically illustrates a gas turbine engine20. The gas turbine engine20is disclosed herein as a two-spool turbo fan that generally incorporates a fan section22, a compressor section24, a combustor section26and a turbine section28. Alternative engine architectures200might include an augmentor section12, an exhaust duct section14and a nozzle section16in addition to the fan section22′, compressor section24′, combustor section26′ and turbine section28′ (seeFIG. 2) among other systems or features. The fan section22drives air along a bypass flowpath and into the compressor section24. The compressor section24drives air along a core flowpath for compression and communication into the combustor section26, which then expands and directs the air through the turbine section28. Although depicted as a turbofan in the disclosed non-limiting embodiment, 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 such as a turbojets, turboshafts, and three-spool (plus fan) turbofans wherein an intermediate spool includes an Intermediate Pressure Compressor (“IPC”) between a Low Pressure Compressor (“LPC”) and a High Pressure Compressor (“HPC”), and an Intermediate Pressure Turbine (“IPT”) between a High Pressure Turbine (“HPT”) and a Low Pressure Turbine (“LPT”).

The engine20generally includes a low spool30and a high spool32mounted for rotation about an engine central longitudinal axis A relative to an engine static structure36via several bearing structures38. The low spool30generally includes an inner shaft40that interconnects a fan42, a low pressure compressor (“LPC”)44and a low pressure turbine (“LPT”)46. The inner shaft40may drive the fan42directly or through a geared architecture48as shown inFIG. 1to drive the fan42at a lower speed than the low spool30. An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system.

The high spool32includes an outer shaft50that interconnects a high pressure compressor (“HPC”)52and a high pressure turbine (“HPT”)54. A combustor56is arranged between the HPC52and the HPT54. The inner shaft40and the outer shaft50are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes.

Core airflow is compressed by the LPC44then the HPC52, mixed with the fuel and burned in the combustor56, then expanded over the HPT54and the LPT46. The LPT46and the HPT54rotationally drive the respective low spool30and the high spool32in response to the expansion. The main engine shafts40,50are supported at a plurality of points by the bearing systems38within the static structure36.

In one non-limiting example, the gas turbine engine20is a high-bypass geared aircraft engine. In a further example, the gas turbine engine20bypass ratio is greater than about six (6:1). The geared architecture48can 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, and in another example is greater than about 2.5:1. The geared turbofan enables operation of the low spool30at higher speeds which can increase the operational efficiency of the LPC44and the LPT46and render increased pressure in a fewer number of stages.

A pressure ratio associated with the LPT46is pressure measured prior to the inlet of the LPT46as related to the pressure at the outlet of the LPT46prior to an exhaust nozzle of the gas turbine engine20. In one non-limiting embodiment, the bypass ratio of the gas turbine engine20is greater than about ten (10:1), the fan diameter is significantly larger than that of the LPC44, and the LPT46has a pressure ratio that is greater than about five (5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.

In one embodiment, a significant amount of thrust is provided by the bypass flow path due to the high bypass ratio. The fan section22of the gas turbine engine20is 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 engine20at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fan section22without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine20is 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 one non-limiting embodiment of the example gas turbine engine20is less than about 1150 fps (351 m/s).

With reference toFIG. 3, the combustor section26generally includes a combustor56with an outer combustor wall assembly60, an inner combustor wall assembly62and a diffuser case module64. The outer combustor wall assembly60and the inner combustor wall assembly62are spaced apart such that a combustion chamber66is defined therebetween. The combustion chamber66is generally annular in shape to surround the engine central longitudinal axis A.

The outer combustor liner assembly60is spaced radially inward from an outer diffuser case64A of the diffuser case module64to define an outer annular plenum76. The inner combustor liner assembly62is spaced radially outward from an inner diffuser case64B of the diffuser case module64to define an inner annular plenum78. It should be understood that although a particular combustor is illustrated, other combustor types with various combustor liner arrangements will also benefit herefrom. It should be further understood that the disclosed cooling flow paths are but an illustrated embodiment and should not be limited only thereto.

The combustor wall assemblies60,62contain the combustion products for direction toward the turbine section28. Each combustor wall assembly60,62generally includes a respective support shell68,70which supports one or more liner panels72A,72B (generally referred to as “72”) and74A,74B (generally referred to as “74”) mounted thereto that are arranged to form a liner array. The support shells68,70may be manufactured by, for example, the hydroforming of a sheet metal alloy to provide the generally cylindrical outer shell68and inner shell70. Each of the liner panels72,74may be generally rectilinear with a circumferential arc. The liner panels72,74may be manufactured of, for example, a nickel based super alloy, ceramic or other temperature resistant material. In one disclosed non-limiting embodiment, the liner array includes a multiple of forward liner panels72A and a multiple of aft liner panels72B that are circumferentially staggered to line the outer shell68. A multiple of forward liner panels74A and a multiple of aft liner panels74B are circumferentially staggered to line the inner shell70.

The combustor56further includes a forward assembly80immediately downstream of the compressor section24to receive compressed airflow therefrom. The forward assembly80generally includes a cowl82, a bulkhead assembly84, and a multiple of swirlers90(one shown). Each of the swirlers90is circumferentially aligned with one of a multiple of fuel nozzles86(one shown) and the respective hood ports94to project through the bulkhead assembly84.

The bulkhead assembly84includes a bulkhead support shell96secured to the combustor walls60,62, and a multiple of circumferentially distributed bulkhead liner panels98secured to the bulkhead support shell96around the swirler opening. The bulkhead support shell96is generally annular and the multiple of circumferentially distributed bulkhead liner panels98are segmented, typically one to each fuel nozzle86and swirler90.

The cowl82extends radially between, and is secured to, the forward most ends of the combustor walls60,62. The cowl82includes a multiple of circumferentially distributed hood ports94that receive one of the respective multiple of fuel nozzles86and facilitates the direction of compressed air into the forward end of the combustion chamber66through a swirler opening92. Each fuel nozzle86may be secured to the diffuser case module64and project through one of the hood ports94and through the swirler opening92within the respective swirler90.

The forward assembly80introduces core combustion air into the forward section of the combustion chamber66while the remainder enters the outer annular plenum76and the inner annular plenum78. The multiple of fuel nozzles86and adjacent structure generate a blended fuel-air mixture that supports stable combustion in the combustion chamber66.

Opposite the forward assembly80, the outer and inner support shells68,70are mounted to a first row of Nozzle Guide Vanes (NGVs)54A in the HPT54. The NGVs54A are static engine components which direct core airflow combustion gases onto the turbine blades of the first turbine rotor in the turbine section28to facilitate the conversion of pressure energy into kinetic energy. The core airflow combustion gases are also accelerated by the NGVs54A because of their convergent shape and are typically given a “spin” or a “swirl” in the direction of turbine rotor rotation. The turbine rotor blades absorb this energy to drive the turbine rotor at high speed.

With reference toFIG. 4, a multiple of studs100extend from each of the liner panels72,74so as to permit an array (seeFIG. 5) of the liner panels72,74to be mounted to their respective support shells68,70with fasteners102(seeFIG. 6) such as nuts. That is, the studs100project rigidly from the liner panels72,74and through the respective support shells68,70to receive the fasteners102at a threaded section thereof.

A multiple of cooling impingement passages104penetrate through the support shells68,70to allow air from the respective annular plenums76,78to enter cavities106formed in the combustor walls60,62between the respective support shells68,70and liner panels72,74. The cooling impingement passages104are generally normal to the surface of the liner panels72,74. The air in the cavities106provide cold side impingement cooling of the liner panels72,74that is generally defined herein as heat removal via internal convection.

A multiple of effusion passages108penetrate through each of the liner panels72,74. The geometry of the passages (e.g., diameter, shape, density, surface angle, incidence angle, etc.) as well as the location of the passages with respect to the high temperature main flow also contributes to effusion film cooling. The combination of impingement passages104and effusion passages108may be referred to as an Impingement Film Floatwall (IFF) assembly.

The effusion passages108allow the air to pass from the cavities106defined in part by a cold side110of the liner panels72,74to a hot side112of the liner panels72,74and thereby facilitate the formation of a thin, cool, insulating blanket or film of cooling air along the hot side112. The effusion passages108are generally more numerous than the impingement passages104to promote the development of film cooling along the hot side112to sheath the liner panels72,74. Film cooling as defined herein is the introduction of a relatively cooler air at one or more discrete locations along a surface exposed to a high temperature environment to protect that surface in the region of the air injection as well as downstream thereof.

A multiple of dilution passages116may penetrate through both the respective support shells68,70and liner panels72,74along a common axis. For example only, in a Rich-Quench-Lean (R-Q-L) type combustor, the dilution passages116are located downstream of the forward assembly80to quench the hot combustion gases within the combustion chamber66by direct supply of cooling air from the respective annular plenums76,78.

With reference toFIG. 6, in one disclosed non-limiting embodiment, a portion of each of the respective support shells68,70is non-parallel with respect to the forward liner panels72A,74A to form a convergent passage120therebetween along a contoured region124. That is, the contoured region124is a radially displaced profile section of the respective support shells68,70with respect to the engine axis A. Although the forward liner panels72A, and the respective forward portion of the outer support shell68will be specifically described and illustrated in some of the disclosed non-limiting embodiments, it should be appreciated that the inner support shell70as well as various other wall assemblies within a gas turbine engine such as within the walls of the augmentor section12, the exhaust duct section14and the nozzle section16(seeFIG. 2) may alternatively or additionally benefit herefrom. That is, the contoured region124may be alternatively or additionally located within engine sections other than the combustor section26which utilize a support shell, liner panel type wall arrangement. The various contoured regions124form one or more convergent passages120for panel cooling air by varying the profile of the combustor shell adjacent to the respective liner panels. Various contours and configurations are possible to tailor the location of the effusion air exit, and optimize heat transfer, pressure loss, manufacturability, NOx reduction, etc. Beneficially, the contoured regions do not require additional hardware over conventional float wall combustor panels to create the convergence and are readily produced with current manufacturing methods.

In this disclosed non-limiting embodiment, the contoured region124of the support shell68defines a hyperbolic cosine (COSH) profile in longitudinal cross-section that extends away from the forward liner panels72A. That is, the forward liner panels72A are generally linear in longitudinal cross-section, while the contoured region124is non-linear in longitudinal cross-section. For perspective, in this disclosed non-limiting embodiment, each of the forward liner panels72A defines an axial length of about 1.5 inches (38 mm) and each may extend over a circumferential arc of about forty (40) degrees (one shown inFIG. 7).

The contoured region124is located adjacent to a row of studs100A, and an intermediate circumferential rail126is located between a forward circumferential rail128and an aft circumferential rail130. Each of the studs100A may be at least partially surrounded by posts132to at least partially support the fastener102and operate as stand-offs between the support shell68and the forward liner panels72A.

Each of the forward liner panels72A, in one disclosed non-limiting embodiment, includes a single row of studs100A (Five shown inFIG. 7) that extend through respective stud apertures134in the support shell68. A center or “king” stud100Aa is received within a central circular stud aperture134awhile the remainder of the studs100Ab are received within elongated apertures134bto facilitate operational thermal growth relative to the center or “king” stud100Aa (FIG. 7).

With continued reference toFIG. 6, the contoured regions124form respective boundaries of impingement cavities106Aa,106Ab that respectively converge toward the forward circumferential rail128and the aft circumferential rail130. As shown, the cavity106A is subdivided by the intermediate circumferential rail126into the forward cavity106Aa and the aft cavity106Ab. The forward cavity106Aa and the aft cavity106Ab thereby accelerate and direct impingement airflow from impingement passages104on each respective side of the intermediate circumferential rail126toward forward effusion apertures108aand aft effusion apertures108b. The forward effusion apertures108aand the aft effusion apertures108bmay define respective angles through the forward liner panels72A to direct effusion airflow generally forward and aft into the combustion chamber66. It should be appreciated that various contours and configurations are possible to tailor the location of the effusion air passages to optimize heat transfer, pressure loss, manufacturability, etc., without need for additional hardware between the respective support shell68and the forward liner panels72A.

Referring toFIGS. 8-11, axial end rails136circumferentially close-out each forward liner panels72A with respect to the support shell68. That is, the forward circumferential rail128and the aft circumferential rail130are located at relatively constant curvature axial interfaces while the axial end rails136extend across an axial length of the support shell68to complete a perimeter rail138that seals the periphery of each forward liner panels72A with respect to the respective support shell68.

With reference toFIG. 8, each axial end rail136and the contoured region124define a respective rail profile140and support shell profile142contoured to form a gap G (e.g., a preassembly gap). That is, the rail profile140of the axial end rails136and the support shell profile142of the contoured region124are of a slightly different profile such that the gap G is located in an area150generally axially adjacent to the intermediate rail126to form respective axial reacting areas152A,152B adjacent to the forward circumferential rail128and the aft circumferential rail130. In one disclosed non-limiting embodiment, the typical gap G is about 0.005-0.020 inches (0.1-0.5 mm).

The gap G is closed in response to deflection of the contoured region124as the fasteners102are tightened onto the studs100(seeFIG. 9). That is, the support shell profile142slides along the respective axial engagement areas152A,152B as the fasteners102are engaged.

As the fasteners102are tightened onto the studs100, the attachment at least partially elastically deflects the support shell68adjacent to the intermediate rail126and produces a tight seal between the perimeter rail138and the support shell68to assure an effective seal therebetween. It should be further appreciated that the intermediate rail126may be of the same height from the cold side110as the forward circumferential rail128and the aft circumferential rail130or alternatively, of a lesser height to further facilitate formation of the gap G.

The elastic deformation of the support shell68to seal with the intermediate rail126reduces—or eliminates—leakage to facilitate formation of relatively large pressure drops across the liner panels72,74and thereby increase cooling effectiveness. That is, tightening of the fasteners102closes the gap G such that the contoured region124closely follows the axial end rails136to circumferentially close-out and form an interference fit between each of the forward liner panels72A with respect to the associated support shell68. It should be appreciated that fasteners such as clips and mechanisms other than threads may alternatively or additionally be utilized to minimize leakage from both the neighboring rail and adjacent rails.

With reference toFIG. 10, in one disclosed non-limiting embodiment, the multiple of studs100A extend from the intermediate rail126. That is, the studs100A are aligned with, and may at least partially form, the intermediate rail126. It should be appreciated that the studs100A may also at least partially extend from the cold side110. That is, the relatively cylindrical studs100A extend beyond and form a footprint greater than the intermediate rail126.

In this disclosed non-limiting embodiment, each of the studs100A may be at least partially surrounded by posts132which extend from the cold side110to at least partially support the fastener102. In another disclosed non-limiting embodiment, the studs100A are not surrounded by the posts132as the intermediate rail126provides adequate support for the respective fastener102(seeFIG. 11).

With reference toFIG. 12, in another disclosed non-limiting embodiment, multiple intermediate rails126A,126B (two shown) with associated studs100B,100C extend from a single liner panel74—here shown as an aft outer liner panel74B. In this disclosed non-limiting embodiment, the dilution passages116are located axially between the intermediate rails126A,126B.

Location of studs100on the intermediate rails126facilitates an effective seal between the liner panels72,74and respective support shell68,70that reduces leakage to facilitate formation of a relatively larger pressure drops across the liner panels72,74and thereby increase cooling effectiveness.