US20180128485A1 - Stud arrangement for gas turbine engine combustor - Google Patents
Stud arrangement for gas turbine engine combustor Download PDFInfo
- Publication number
- US20180128485A1 US20180128485A1 US15/343,988 US201615343988A US2018128485A1 US 20180128485 A1 US20180128485 A1 US 20180128485A1 US 201615343988 A US201615343988 A US 201615343988A US 2018128485 A1 US2018128485 A1 US 2018128485A1
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- Prior art keywords
- liner panel
- free zone
- combustor
- recited
- stud free
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- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/002—Wall structures
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/00017—Assembling combustion chamber liners or subparts
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/03042—Film cooled combustion chamber walls or domes
Definitions
- 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.
- the combustor section typically includes a combustion chamber formed by an inner and outer wall assembly.
- Each wall assembly includes a support shell lined with heat shields often referred to as liner panels.
- combustor Impingement Film-Cooled Floatwall (IFF) liner panels are typically a curved flat surface on a hot side exposed to the gas path.
- the opposite, or cold side has features such as cast in threaded studs to mount the liner panel and a full perimeter rail that contacts the inner surface of the respective liner shell. These features may result in durability issues.
- a liner panel for use in a combustor of a gas turbine engine can include a stud free zone downstream of a combustor swirler.
- a further embodiment of the present disclosure may include, wherein the stud free zone is trapezoidal in shape.
- a further embodiment of the present disclosure may include, wherein the stud free zone is defined by a forward liner panel.
- a further embodiment of the present disclosure may include an aft liner panel aft of the forward liner panel.
- a further embodiment of the present disclosure may include an aft stud free zone downstream of the forward liner panel stud free zone.
- a further embodiment of the present disclosure may include at least one major diffusion aperture an aft stud free zone downstream of the forward liner panel stud free zone.
- a further embodiment of the present disclosure may include, wherein the stud free zone is trapezoidal in shape and defined by a forward liner panel.
- a further embodiment of the present disclosure may include, wherein the stud free zone is located toward an aft edge of the forward liner panel.
- a further embodiment of the present disclosure may include, wherein the stud free zone is defined by a truncated triangle with a truncated apex located at combustor swirler.
- a further embodiment of the present disclosure may include, wherein the stud free zone includes a multiple of film cooling holes.
- a combustor for a gas turbine engine can include a support shell; and a liner panel mounted to the support shell via a multiple of studs, the liner panel including a stud free zone downstream of each respective combustor swirler, the stud free zone including a multiple of film cooling holes.
- a further embodiment of the present disclosure may include a forward assembly including a bulkhead support shell, a bulkhead assembly mounted to the bulkhead support shell, and a multiple of the combustor swirlers mounted at least partially through the bulkhead assembly.
- a further embodiment of the present disclosure may include, wherein the forward assembly is mounted to the support shell.
- a further embodiment of the present disclosure may include a multiple of circumferentially distributed bulkhead liner panels secured to the bulkhead support shell around the swirler opening.
- a further embodiment of the present disclosure may include, wherein the stud free zone is defined by a forward liner panel.
- a further embodiment of the present disclosure may include an aft liner panel downstream of the forward liner panel, an aft stud free zone downstream of the forward liner panel stud free zone.
- a method of directing airflow through a wall assembly within a combustor of a gas turbine engine can include providing a stud free zone in a forward liner panel downstream of a combustor swirler, the stud free zone including a multiple of film cooling holes.
- a further embodiment of the present disclosure may include locating a dilution passage within an aft stud free zone in an aft liner panel, the aft liner panel downstream of the forward liner panel.
- a further embodiment of the present disclosure may include defining the stud free zone in the forward liner panel as a trapezoidal shape.
- a further embodiment of the present disclosure may include defining the stud free zone in the forward liner panel as a truncated triangle with a truncated apex located adjacent to the combustion swirler.
- FIG. 1 is a schematic cross-section of an example gas turbine engine architecture
- FIG. 2 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;
- FIG. 3 is an exploded partial sectional view of a portion of a combustor wall assembly
- FIG. 4 is a perspective cold side view of a portion of a liner panel array
- FIG. 5 is a perspective partial sectional view of a combustor
- FIG. 6 is a sectional view of a portion of a combustor wall assembly.
- FIG. 7 is a perspective view of a liner panel array.
- 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 .
- Alternative engine architectures might include an augmentor section among other systems or features.
- 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 .
- 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 the high pressure turbine (“HPT”) and the Low pressure Turbine (“LPT”).
- IPC intermediate pressure compressor
- LPC Low Pressure Compressor
- HPC High Pressure Compressor
- IPT intermediate pressure turbine
- 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 40 that interconnects a fan 42 , a low pressure compressor (“LPC”) 44 and a low pressure turbine (“LPT”) 46 .
- the inner shaft 40 drives the fan 42 directly or through a geared architecture 48 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 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.
- the main engine shafts 40 , 50 are supported at a plurality of points by bearing systems 38 within the static structure 36 .
- the gas turbine engine 20 is a high-bypass geared aircraft engine.
- the gas turbine engine 20 bypass ratio is 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, 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 and 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 .
- 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
- the LPT 46 has a pressure ratio that is greater than about five (5:1). It should be appreciated, 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.
- a significant amount of thrust is provided by the bypass flow path due to the high bypass ratio.
- the fan section 22 of the gas turbine engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10668 m). 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).
- TSFC Thrust Specific Fuel Consumption
- Fan Pressure Ratio is the pressure ratio across a blade of the fan section 22 without 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 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 one non-limiting embodiment of the example gas turbine engine 20 is less than about 1150 fps (351 m/s).
- 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 a combustion chamber 66 is defined therebetween.
- the combustion chamber 66 is generally annular in shape to surround the engine central longitudinal axis A.
- the outer combustor liner 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 liner 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 appreciated that although a particular combustor is illustrated, other combustor types with various combustor liner arrangements will also benefit herefrom. It should be further appreciated that the disclosed cooling flow paths are but an illustrated embodiment and should not be limited only thereto.
- 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 support shell 68 , 70 which supports one or more liner panels 72 , 74 mounted thereto arranged to form a liner array.
- the support shells 68 , 70 may be manufactured by, for example, the hydroforming of a sheet metal alloy to provide the generally cylindrical outer shell 68 and inner shell 70 .
- Each of the liner panels 72 , 74 may be generally rectilinear with a circumferential arc.
- the liner panels 72 , 74 may be manufactured of, for example, a nickel based super alloy, ceramic or other temperature resistant material.
- the liner array includes a multiple of forward liner panels 72 A and a multiple of aft liner panels 72 B that are circumferentially staggered to line the outer shell 68 .
- a multiple of forward liner panels 74 A and a multiple of aft liner panels 74 B are circumferentially staggered to line the inner shell 70 .
- 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 a cowl 82 , a bulkhead assembly 84 , and a multiple of swirlers 90 (one shown). Each of the swirlers 90 is circumferentially aligned with one of a multiple of fuel nozzles 86 (one shown) and the respective hood ports 94 to project through the bulkhead assembly 84 .
- the bulkhead assembly 84 includes a bulkhead support shell 96 secured to the combustor walls 60 , 62 , and a multiple of circumferentially distributed bulkhead liner panels 98 secured to the bulkhead support shell 96 around the swirler opening.
- the bulkhead support shell 96 is generally annular and the multiple of circumferentially distributed bulkhead liner panels 98 are segmented, typically one to each fuel nozzle 86 and swirler 90 .
- the cowl 82 extends radially between, and is secured to, the forwardmost ends of the combustor walls 60 , 62 .
- the cowl 82 includes a multiple of circumferentially distributed hood ports 94 that receive one of the respective multiple of fuel nozzles 86 and facilitates the direction of compressed air into the forward end of the combustion chamber 66 through a swirler opening 92 .
- Each fuel nozzle 86 may be secured to the diffuser case module 64 and project through one of the hood ports 94 and through the swirler opening 92 within the respective swirler 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 blended fuel-air mixture that supports stable combustion in the combustion chamber 66 .
- the outer and inner support shells 68 , 70 are mounted 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 the turbine blades of the first turbine rotor 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.
- the turbine rotor blades absorb this energy to drive the turbine rotor at high speed.
- a multiple of studs 100 extend from each of the liner panels 72 , 74 so as to permit a liner array (partially shown in FIG. 4 ) of the liner panels 72 , 74 to be mounted to their respective support shells 68 , 70 with fasteners 102 such as nuts. That is, the studs 100 project rigidly from the liner panels 72 , 74 to extend through the respective support shells 68 , 70 and receive the fasteners 102 on a threaded section thereof ( FIG. 5 ).
- a multiple of cooling impingement passages 104 penetrate through the support shells 68 , 70 to allow air from the respective annular plenums 76 , 78 to enter cavities 106 formed in the combustor walls 60 , 62 between the respective support shells 68 , 70 and liner panels 72 , 74 .
- the impingement passages 104 are generally normal to the surface of the liner panels 72 , 74 .
- the air in the cavities 106 provides cold side impingement cooling of the liner panels 72 , 74 that is generally defined herein as heat removal via internal convection.
- a multiple of effusion passages 108 penetrate through each of the liner panels 72 , 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 combustion flow also contributes to effusion cooling.
- the effusion passages 108 allow the air to pass from the cavities 106 defined in part by a cold side 110 of the liner panels 72 , 74 to a hot side 112 of the liner panels 72 , 74 and thereby facilitate the formation of a thin, relatively cool, film of cooling air along the hot side 112 .
- each of the multiple of effusion passages 108 are typically 0.025′′ (0.635 mm) in diameter and define a surface angle of about thirty (30) degrees with respect to the cold side 110 of the liner panels 72 , 74 .
- the effusion passages 108 are generally more numerous than the impingement passages 104 and promote film cooling along the hot side 112 to sheath the liner panels 72 , 74 ( FIG. 6 ).
- 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.
- impingement passages 104 and effusion passages 108 may be referred to as an Impingement Film Floatwall (IFF) assembly.
- IFF Impingement Film Floatwall
- a multiple of dilution passages 116 are located in the liner panels 72 , 74 each along a common axis D.
- the dilution passages 116 are located in a circumferential line W (shown partially in FIG. 4 ).
- the dilution passages 116 are illustrated in the disclosed non-limiting embodiment as within the aft liner panels 72 B, 74 B, the dilution passages may alternatively be located in the forward liner panels 72 A, 72 B or in a single liner panel which replaces the fore/aft liner panel array.
- the dilution passages 116 although illustrated in the disclosed non-limiting embodiment as integrally formed in the liner panels, it should be appreciated that the dilution passages 116 may be separate components. Whether integrally formed or separate components, the dilution passages 116 may be referred to as grommets.
- each of the liner panels 72 A, 72 B, 74 A, 74 B in the liner panel array includes a perimeter rail 120 formed by a forward circumferential rail 122 , an aft circumferential rail 124 , and axial rails 126 A, 126 B, that interconnect the forward and aft circumferential rail 122 , 124 .
- the perimeter rail 120 seals each liner panel with respect to the respective support shell 68 , 70 to form the impingement cavity 106 therebetween.
- the forward and aft circumferential rail 122 , 124 are located at relatively constant curvature shell interfaces while the axial rails 126 extend across an axial length of the respective support shell 68 , 70 to complete the perimeter rail 120 that seals the liner panels 72 , 74 to the respective support shell 68 , 70 .
- a multiple of studs 100 are located adjacent to the respective forward circumferential rail 122 and the aft circumferential rail 124 .
- Each of the studs 100 may be at least partially surrounded by posts 130 to at least partially support the fastener 102 and provide a stand-off between each liner panels 72 B, 74 B and respective support shell 68 , 70 .
- the quantity and location of the multiple of studs 100 is typically based on structural analysis and symmetry of the studs 100 relative to the liner to facilitate proper sealing of the panel rail to the inner combustor shell.
- the conventional position would often locate one or more of the multiple of studs 100 downstream of the combustor swirlers 90 . As this area may have relatively high metal temperatures, durability issues may result from the lack of effusion cooling as this issue may be more significant at the aft section of the forward row of liners.
- the multiple of studs 100 of the forward liner panels 72 A, 74 A are not located within a stud free zone 200 defined downstream of each of the combustion swirlers 90 .
- Each stud free zone 200 is defined as an essentially truncated triangular shape.
- the stud free zone 200 is defined by a truncated triangle with a truncated apex 201 located at the combustor swirler 90 .
- the stud free zone 200 is a trapezoidal shaped zone located at the aft edge of the forward liner panels 72 A, 74 A. That is, to increase durability, the studs 100 are specifically moved away from each zone 200 directly aft of the respective combustor swirlers 90 in the aft section of the forward liner panels 72 A, 74 A as this area has the relatively hottest surface metal temperatures.
- the stud free zone 200 facilitates a more efficient distribution of film cooling holes in the hottest areas of the segment as the studs no longer hinder location of film cooling holes 108 ( FIG. 3 ).
- the stud free zone 200 extends from about 0.75 inches (19 mm) to 1.7 inches (43 mm) from the respective combustor swirler 90 and the sides between the fore and aft lines are at about 20 degrees.
- the forward most row of studs 100 A are also intentionally moved out away from an aft liner panel stud free zone 202 downstream of the stud free zone 200 and the respective combustor swirler 90 .
- this permits a more advantageous distribution of cooling holes around the area of the liner segments which typically have the hottest metal temperatures.
- at least one dilution passages 116 may be located within the aft liner panel stud free zone 202 .
- the stud free zone 200 in the aft liner panels 72 B, 74 B defines a rectangle of about 1.5 inches (38 mm) by 2.8 inches (71 mm) and is located 1.8 inches (45 mm) behind the respective combustor swirlers 90 .
Abstract
Description
- 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 formed 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 typical combustor chamber designs, combustor Impingement Film-Cooled Floatwall (IFF) liner panels are typically a curved flat surface on a hot side exposed to the gas path. The opposite, or cold side, has features such as cast in threaded studs to mount the liner panel and a full perimeter rail that contacts the inner surface of the respective liner shell. These features may result in durability issues.
- A liner panel for use in a combustor of a gas turbine engine, the liner panel according to one disclosed non-limiting embodiment of the present disclosure can include a stud free zone downstream of a combustor swirler.
- A further embodiment of the present disclosure may include, wherein the stud free zone is trapezoidal in shape.
- A further embodiment of the present disclosure may include, wherein the stud free zone is defined by a forward liner panel.
- A further embodiment of the present disclosure may include an aft liner panel aft of the forward liner panel.
- A further embodiment of the present disclosure may include an aft stud free zone downstream of the forward liner panel stud free zone.
- A further embodiment of the present disclosure may include at least one major diffusion aperture an aft stud free zone downstream of the forward liner panel stud free zone.
- A further embodiment of the present disclosure may include, wherein the stud free zone is trapezoidal in shape and defined by a forward liner panel.
- A further embodiment of the present disclosure may include, wherein the stud free zone is located toward an aft edge of the forward liner panel.
- A further embodiment of the present disclosure may include, wherein the stud free zone is defined by a truncated triangle with a truncated apex located at combustor swirler.
- A further embodiment of the present disclosure may include, wherein the stud free zone includes a multiple of film cooling holes.
- A combustor for a gas turbine engine according to another disclosed non-limiting embodiment of the present disclosure can include a support shell; and a liner panel mounted to the support shell via a multiple of studs, the liner panel including a stud free zone downstream of each respective combustor swirler, the stud free zone including a multiple of film cooling holes.
- A further embodiment of the present disclosure may include a forward assembly including a bulkhead support shell, a bulkhead assembly mounted to the bulkhead support shell, and a multiple of the combustor swirlers mounted at least partially through the bulkhead assembly.
- A further embodiment of the present disclosure may include, wherein the forward assembly is mounted to the support shell.
- A further embodiment of the present disclosure may include a multiple of circumferentially distributed bulkhead liner panels secured to the bulkhead support shell around the swirler opening.
- A further embodiment of the present disclosure may include, wherein the stud free zone is defined by a forward liner panel.
- A further embodiment of the present disclosure may include an aft liner panel downstream of the forward liner panel, an aft stud free zone downstream of the forward liner panel stud free zone.
- A method of directing airflow through a wall assembly within a combustor of a gas turbine engine according to another disclosed non-limiting embodiment of the present disclosure can include providing a stud free zone in a forward liner panel downstream of a combustor swirler, the stud free zone including a multiple of film cooling holes.
- A further embodiment of the present disclosure may include locating a dilution passage within an aft stud free zone in an aft liner panel, the aft liner panel downstream of the forward liner panel.
- A further embodiment of the present disclosure may include defining the stud free zone in the forward liner panel as a trapezoidal shape.
- A further embodiment of the present disclosure may include defining the stud free zone in the forward liner panel as a truncated triangle with a truncated apex located adjacent to the combustion swirler.
- 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 appreciated, however, the following description and drawings are intended to be exemplary in nature and non-limiting.
- 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:
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FIG. 1 is a schematic cross-section of an example gas turbine engine architecture; -
FIG. 2 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; -
FIG. 3 is an exploded partial sectional view of a portion of a combustor wall assembly; -
FIG. 4 is a perspective cold side view of a portion of a liner panel array; -
FIG. 5 is a perspective partial sectional view of a combustor; -
FIG. 6 is a sectional view of a portion of a combustor wall assembly; and -
FIG. 7 is a perspective view of a liner panel array. -
FIG. 1 schematically illustrates agas turbine engine 20. Thegas turbine engine 20 is disclosed herein as a two-spool turbo fan that generally incorporates afan section 22, acompressor section 24, acombustor section 26 and aturbine section 28. Alternative engine architectures might include an augmentor section among other systems or features. Thefan section 22 drives air along a bypass flowpath and into thecompressor section 24. Thecompressor section 24 drives air along a core flowpath for compression and communication into thecombustor section 26, which then expands and directs the air through theturbine section 28. Although depicted as a turbofan in the disclosed non-limiting embodiment, it should be appreciated 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 the high pressure turbine (“HPT”) and the Low pressure Turbine (“LPT”). - The
engine 20 generally includes alow spool 30 and ahigh spool 32 mounted for rotation about an engine central longitudinal axis A relative to an enginestatic structure 36 viaseveral bearing structures 38. Thelow spool 30 generally includes aninner shaft 40 that interconnects afan 42, a low pressure compressor (“LPC”) 44 and a low pressure turbine (“LPT”) 46. Theinner shaft 40 drives thefan 42 directly or through a gearedarchitecture 48 to drive thefan 42 at a lower speed than thelow spool 30. An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system. - The
high spool 32 includes anouter shaft 50 that interconnects a high pressure compressor (“HPC”) 52 and high pressure turbine (“HPT”) 54. Acombustor 56 is arranged between the HPC 52 and the HPT 54. Theinner shaft 40 and theouter 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 thecombustor 56, then expanded over the HPT 54 and theLPT 46. TheLPT 46 and HPT 54 rotationally drive the respectivelow spool 30 andhigh spool 32 in response to the expansion. Themain engine shafts systems 38 within thestatic structure 36. - In one non-limiting example, the
gas turbine engine 20 is a high-bypass geared aircraft engine. In a further example, thegas turbine engine 20 bypass ratio is greater than about six (6:1). The gearedarchitecture 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, and in another example is greater than about 2.5:1. The geared turbofan enables operation of thelow spool 30 at higher speeds which can increase the operational efficiency of theLPC 44 andLPT 46 and 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 theLPT 46 as related to the pressure at the outlet of theLPT 46 prior to an exhaust nozzle of thegas turbine engine 20. In one non-limiting embodiment, the bypass ratio of thegas turbine engine 20 is greater than about ten (10:1), the fan diameter is significantly larger than that of theLPC 44, and theLPT 46 has a pressure ratio that is greater than about five (5:1). It should be appreciated, 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 section 22 of thegas turbine engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10668 m). This flight condition, with thegas turbine engine 20 at 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 section 22 without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the examplegas 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 one non-limiting embodiment of the examplegas turbine engine 20 is less than about 1150 fps (351 m/s). - With reference to
FIG. 2 , thecombustor section 26 generally includes acombustor 56 with an outercombustor wall assembly 60, an inner combustor wall assembly 62, and adiffuser case module 64. The outercombustor wall assembly 60 and the inner combustor wall assembly 62 are spaced apart such that acombustion chamber 66 is defined therebetween. Thecombustion chamber 66 is generally annular in shape to surround the engine central longitudinal axis A. - The outer
combustor liner assembly 60 is spaced radially inward from anouter diffuser case 64A of thediffuser case module 64 to define an outerannular plenum 76. The inner combustor liner assembly 62 is spaced radially outward from aninner diffuser case 64B of thediffuser case module 64 to define an innerannular plenum 78. It should be appreciated that although a particular combustor is illustrated, other combustor types with various combustor liner arrangements will also benefit herefrom. It should be further appreciated that the disclosed cooling flow paths are but an illustrated embodiment and should not be limited only thereto. - The
combustor wall assemblies 60, 62 contain the combustion products for direction toward theturbine section 28. Eachcombustor wall assembly 60, 62 generally includes arespective support shell more liner panels support shells outer shell 68 andinner shell 70. Each of theliner panels liner panels forward liner panels 72A and a multiple ofaft liner panels 72B that are circumferentially staggered to line theouter shell 68. A multiple offorward liner panels 74A and a multiple ofaft liner panels 74B are circumferentially staggered to line theinner shell 70. - The
combustor 56 further includes aforward assembly 80 immediately downstream of thecompressor section 24 to receive compressed airflow therefrom. Theforward assembly 80 generally includes acowl 82, abulkhead assembly 84, and a multiple of swirlers 90 (one shown). Each of theswirlers 90 is circumferentially aligned with one of a multiple of fuel nozzles 86 (one shown) and therespective hood ports 94 to project through thebulkhead assembly 84. - The
bulkhead assembly 84 includes abulkhead support shell 96 secured to thecombustor walls 60, 62, and a multiple of circumferentially distributed bulkhead liner panels 98 secured to thebulkhead support shell 96 around the swirler opening. Thebulkhead support shell 96 is generally annular and the multiple of circumferentially distributed bulkhead liner panels 98 are segmented, typically one to eachfuel nozzle 86 andswirler 90. - The
cowl 82 extends radially between, and is secured to, the forwardmost ends of thecombustor walls 60, 62. Thecowl 82 includes a multiple of circumferentially distributedhood ports 94 that receive one of the respective multiple offuel nozzles 86 and facilitates the direction of compressed air into the forward end of thecombustion chamber 66 through aswirler opening 92. Eachfuel nozzle 86 may be secured to thediffuser case module 64 and project through one of thehood ports 94 and through theswirler opening 92 within therespective swirler 90. - The
forward assembly 80 introduces core combustion air into the forward section of thecombustion chamber 66 while the remainder enters the outerannular plenum 76 and the innerannular plenum 78. The multiple offuel nozzles 86 and adjacent structure generate a blended fuel-air mixture that supports stable combustion in thecombustion chamber 66. - Opposite the
forward assembly 80, the outer andinner support shells HPT 54. TheNGVs 54A are static engine components which direct core airflow combustion gases onto the turbine blades of the first turbine rotor in theturbine section 28 to facilitate the conversion of pressure energy into kinetic energy. The core airflow combustion gases are also accelerated by theNGVs 54A 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 to
FIG. 3 , a multiple ofstuds 100 extend from each of theliner panels FIG. 4 ) of theliner panels respective support shells fasteners 102 such as nuts. That is, thestuds 100 project rigidly from theliner panels respective support shells fasteners 102 on a threaded section thereof (FIG. 5 ). - A multiple of cooling
impingement passages 104 penetrate through thesupport shells annular plenums combustor walls 60, 62 between therespective support shells liner panels impingement passages 104 are generally normal to the surface of theliner panels liner panels - A multiple of
effusion passages 108 penetrate through each of theliner panels effusion passages 108 allow the air to pass from the cavities 106 defined in part by acold side 110 of theliner panels hot side 112 of theliner panels hot side 112. - In one disclosed non-limiting embodiment, each of the multiple of
effusion passages 108 are typically 0.025″ (0.635 mm) in diameter and define a surface angle of about thirty (30) degrees with respect to thecold side 110 of theliner panels effusion passages 108 are generally more numerous than theimpingement passages 104 and promote film cooling along thehot side 112 to sheath theliner panels 72, 74 (FIG. 6 ). 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. - The combination of
impingement passages 104 andeffusion passages 108 may be referred to as an Impingement Film Floatwall (IFF) assembly. A multiple ofdilution passages 116 are located in theliner panels dilution passages 116 are located in a circumferential line W (shown partially inFIG. 4 ). Although thedilution passages 116 are illustrated in the disclosed non-limiting embodiment as within theaft liner panels forward liner panels dilution passages 116 although illustrated in the disclosed non-limiting embodiment as integrally formed in the liner panels, it should be appreciated that thedilution passages 116 may be separate components. Whether integrally formed or separate components, thedilution passages 116 may be referred to as grommets. - With reference to
FIG. 4 , in one disclosed non-limiting embodiment, each of theliner panels perimeter rail 120 formed by a forwardcircumferential rail 122, an aftcircumferential rail 124, andaxial rails circumferential rail perimeter rail 120 seals each liner panel with respect to therespective support shell circumferential rail respective support shell perimeter rail 120 that seals theliner panels respective support shell - A multiple of
studs 100 are located adjacent to the respective forwardcircumferential rail 122 and the aftcircumferential rail 124. Each of thestuds 100 may be at least partially surrounded byposts 130 to at least partially support thefastener 102 and provide a stand-off between eachliner panels respective support shell - With reference to
FIG. 7 , the quantity and location of the multiple ofstuds 100 is typically based on structural analysis and symmetry of thestuds 100 relative to the liner to facilitate proper sealing of the panel rail to the inner combustor shell. The conventional position would often locate one or more of the multiple ofstuds 100 downstream of thecombustor swirlers 90. As this area may have relatively high metal temperatures, durability issues may result from the lack of effusion cooling as this issue may be more significant at the aft section of the forward row of liners. - In one embodiment, the multiple of
studs 100 of theforward liner panels free zone 200 defined downstream of each of thecombustion swirlers 90. Each studfree zone 200 is defined as an essentially truncated triangular shape. - In one embodiment, the stud
free zone 200 is defined by a truncated triangle with atruncated apex 201 located at thecombustor swirler 90. In other words, the studfree zone 200 is a trapezoidal shaped zone located at the aft edge of theforward liner panels studs 100 are specifically moved away from eachzone 200 directly aft of therespective combustor swirlers 90 in the aft section of theforward liner panels free zone 200 facilitates a more efficient distribution of film cooling holes in the hottest areas of the segment as the studs no longer hinder location of film cooling holes 108 (FIG. 3 ). In one example, the studfree zone 200 extends from about 0.75 inches (19 mm) to 1.7 inches (43 mm) from therespective combustor swirler 90 and the sides between the fore and aft lines are at about 20 degrees. - For the
aft liner panels studs 100A are also intentionally moved out away from an aft liner panel studfree zone 202 downstream of the studfree zone 200 and therespective combustor swirler 90. As with theforward liner panels dilution passages 116 may be located within the aft liner panel studfree zone 202. The studfree zone 200 in theaft liner panels respective combustor swirlers 90. - 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 appreciated 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 limitations 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.
Claims (20)
Priority Applications (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US15/343,988 US20180128485A1 (en) | 2016-11-04 | 2016-11-04 | Stud arrangement for gas turbine engine combustor |
EP17200015.0A EP3318803B1 (en) | 2016-11-04 | 2017-11-03 | Stud arrangement for gas turbine engine combustor |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
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US15/343,988 US20180128485A1 (en) | 2016-11-04 | 2016-11-04 | Stud arrangement for gas turbine engine combustor |
Publications (1)
Publication Number | Publication Date |
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US20180128485A1 true US20180128485A1 (en) | 2018-05-10 |
Family
ID=60245009
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
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US15/343,988 Abandoned US20180128485A1 (en) | 2016-11-04 | 2016-11-04 | Stud arrangement for gas turbine engine combustor |
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US (1) | US20180128485A1 (en) |
EP (1) | EP3318803B1 (en) |
Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5782294A (en) * | 1995-12-18 | 1998-07-21 | United Technologies Corporation | Cooled liner apparatus |
US5799491A (en) * | 1995-02-23 | 1998-09-01 | Rolls-Royce Plc | Arrangement of heat resistant tiles for a gas turbine engine combustor |
US20150036219A1 (en) * | 2013-08-02 | 2015-02-05 | Forward Optics Co., Ltd. | Lens array module |
US20150260404A1 (en) * | 2012-09-30 | 2015-09-17 | United Technologies Corporation | Interface heat shield for a combustor of a gas turbine engine |
US20150260399A1 (en) * | 2012-09-28 | 2015-09-17 | United Technologies Corporation | Combustor section of a gas turbine engine |
Family Cites Families (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4614082A (en) * | 1972-12-19 | 1986-09-30 | General Electric Company | Combustion chamber construction |
US7509809B2 (en) * | 2005-06-10 | 2009-03-31 | Pratt & Whitney Canada Corp. | Gas turbine engine combustor with improved cooling |
US20100095680A1 (en) * | 2008-10-22 | 2010-04-22 | Honeywell International Inc. | Dual wall structure for use in a combustor of a gas turbine engine |
US20100095679A1 (en) * | 2008-10-22 | 2010-04-22 | Honeywell International Inc. | Dual wall structure for use in a combustor of a gas turbine engine |
EP2946092B1 (en) * | 2013-01-17 | 2019-04-17 | United Technologies Corporation | Gas turbine engine combustor liner assembly with convergent hyperbolic profile |
-
2016
- 2016-11-04 US US15/343,988 patent/US20180128485A1/en not_active Abandoned
-
2017
- 2017-11-03 EP EP17200015.0A patent/EP3318803B1/en active Active
Patent Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5799491A (en) * | 1995-02-23 | 1998-09-01 | Rolls-Royce Plc | Arrangement of heat resistant tiles for a gas turbine engine combustor |
US5782294A (en) * | 1995-12-18 | 1998-07-21 | United Technologies Corporation | Cooled liner apparatus |
US20150260399A1 (en) * | 2012-09-28 | 2015-09-17 | United Technologies Corporation | Combustor section of a gas turbine engine |
US20150260404A1 (en) * | 2012-09-30 | 2015-09-17 | United Technologies Corporation | Interface heat shield for a combustor of a gas turbine engine |
US20150036219A1 (en) * | 2013-08-02 | 2015-02-05 | Forward Optics Co., Ltd. | Lens array module |
Also Published As
Publication number | Publication date |
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EP3318803B1 (en) | 2021-05-05 |
EP3318803A1 (en) | 2018-05-09 |
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