Patent Description:
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. Combustor panels are often employed in modern annular gas turbine combustors to form the inner flow path. The panels are part of a two-wall liner and are exposed to a thermally challenging environment.

In typical combustor chamber designs, combustor Impingement Film-Cooled Floatwall (IFF) liner panels typically include 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 contact the inner surface of the liner shells.

The wall assemblies are segmented to accommodate growth of the panels in operation and for other considerations. Combustor panels typically have a quadrilateral projection (i.e. rectangular or trapezoid) when viewed from the hot surface. The panels have a straight edge that forms the front or upstream edge of the panel and a second straight edge that forms the back or downstream edge of the combustor. The panels also have side edges that are linear in profile.

The liner panels extend over an arc in a conical or cylindrical fashion in a plane and terminate in regions where the combustor geometry transitions, diverges, or converges. This may contribute to durability and flow path concerns where forward and aft panels merge or form interfaces. These areas can be prone to steps between panels, dead regions, cooling challenges and adverse local aerodynamics.

<CIT> discloses a prior art liner panel of a combustor.

<CIT> discloses a prior art bypass heat shield element of a combustor, as set forth in the preamble of claim <NUM>.

<CIT> discloses a prior art combustion chamber.

<CIT> discloses a prior art support for a gas turbine combustion liner segment.

<CIT> discloses prior art combustor panel T-junction cooling.

From a first aspect, the invention provides a combustor for a gas turbine engine as recited in claim <NUM>.

The gas turbine engine <NUM> is disclosed herein as a two-spool turbo fan that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. Alternative engine architectures might include an augmentor section among other systems or features. The fan section <NUM> drives air along a bypass flowpath and into the compressor section <NUM>. The compressor section <NUM> drives air along a core flowpath for compression and communication into the combustor section <NUM>, which then expands and directs the air through the turbine section <NUM>. 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 <NUM> generally includes a low spool <NUM> and a high spool <NUM> mounted for rotation about an engine central longitudinal axis A relative to an engine static structure <NUM> via several bearing structures <NUM>. The low spool <NUM> generally includes an inner shaft <NUM> that interconnects a fan <NUM>, a Low Pressure Compressor ("LPC") <NUM> and a Low Pressure Turbine ("LPT") <NUM>. The inner shaft <NUM> drives the fan <NUM> directly or through a geared architecture <NUM> to drive the fan <NUM> at a lower speed than the low spool <NUM>. An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system.

The high spool <NUM> includes an outer shaft <NUM> that interconnects a High Pressure Compressor ("HPC") <NUM> and High Pressure Turbine ("HPT") <NUM>. A combustor <NUM> is arranged between the HPC <NUM> and the HPT <NUM>. The inner shaft <NUM> and the outer shaft <NUM> 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 <NUM>, then the HPC <NUM>, mixed with the fuel and burned in the combustor <NUM>, then expanded over the HPT <NUM> and the LPT <NUM>. The LPT <NUM> and HPT <NUM> rotationally drive the respective low spool <NUM> and high spool <NUM> in response to the expansion. The main engine shafts <NUM>, <NUM> are supported at a plurality of points by bearing systems <NUM> within the static structure <NUM>.

In one non-limiting example, the gas turbine engine <NUM> is a high-bypass geared aircraft engine. In a further example, the gas turbine engine <NUM> bypass ratio is greater than about six. The geared architecture <NUM> 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 <NUM>, and in another example is greater than about <NUM>:<NUM>. The geared turbofan enables operation of the low spool <NUM> at higher speeds which can increase the operational efficiency of the LPC <NUM> and LPT <NUM> and render increased pressure in a fewer number of stages.

A pressure ratio associated with the LPT <NUM> is pressure measured prior to the inlet of the LPT <NUM> as related to the pressure at the outlet of the LPT <NUM> prior to an exhaust nozzle of the gas turbine engine <NUM>. In one non-limiting embodiment, the bypass ratio of the gas turbine engine <NUM> is greater than about ten, the fan diameter is significantly larger than that of the LPC <NUM>, and the LPT <NUM> has a pressure ratio that is greater than about five. 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 <NUM> of the gas turbine engine <NUM> is designed for a particular flight condition - typically cruise at about <NUM> Mach and about <NUM>,<NUM> feet (<NUM>). This flight condition, with the gas turbine engine <NUM> 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 <NUM> 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 <NUM> is less than <NUM>. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of ("Tram" °R / <NUM>°R)<NUM> (where °R = K x <NUM>/<NUM>). The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine <NUM> is less than about <NUM> fps (<NUM>/s).

With reference to <FIG>, the combustor section <NUM> generally includes a combustor <NUM> with an outer combustor wall assembly <NUM>, an inner combustor wall assembly <NUM>, and a diffuser case module <NUM>. The outer combustor wall assembly <NUM> and the inner combustor wall assembly <NUM> are spaced apart such that a combustion chamber <NUM> is defined therebetween. The combustion chamber <NUM> is generally annular in shape to surround the engine central longitudinal axis A.

The outer combustor liner assembly <NUM> is spaced radially inward from an outer diffuser case 64A of the diffuser case module <NUM> to define an outer annular plenum <NUM>. The inner combustor liner assembly <NUM> is spaced radially outward from an inner diffuser case 64B of the diffuser case module <NUM> to define an inner annular plenum <NUM>. 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 <NUM>, <NUM> contain the combustion products for direction toward the turbine section <NUM>. Each combustor wall assembly <NUM>, <NUM> generally includes a respective support shell <NUM>, <NUM> which supports one or more liner panels <NUM>, <NUM> mounted thereto arranged to form a liner array. The support shells <NUM>, <NUM> may be manufactured by, for example, the hydroforming of a sheet metal alloy to provide the generally cylindrical outer shell <NUM> and inner shell <NUM>. Each of the liner panels <NUM>, <NUM> may be generally rectilinear with a circumferential arc. The liner panels <NUM>, <NUM> may 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 panels 72A and a multiple of aft liner panels 74A that are circumferentially staggered to line the outer shell <NUM>. A multiple of forward liner panels 72B and a multiple of aft liner panels 74B are circumferentially staggered to line the inner shell <NUM>.

The combustor <NUM> further includes a forward assembly <NUM> immediately downstream of the compressor section <NUM> to receive compressed airflow therefrom. The forward assembly <NUM> generally includes a cowl <NUM>, a bulkhead assembly <NUM>, and a multiple of swirlers <NUM> (one shown). Each of the swirlers <NUM> is circumferentially aligned with one of a multiple of fuel nozzles <NUM> (one shown) and the respective hood ports <NUM> to project through the bulkhead assembly <NUM>.

The bulkhead assembly <NUM> includes a bulkhead support shell <NUM> secured to the combustor walls <NUM>, <NUM>, and a multiple of circumferentially distributed bulkhead liner panels <NUM> secured to the bulkhead support shell <NUM> around the swirler opening. The bulkhead support shell <NUM> is generally annular and the multiple of circumferentially distributed bulkhead liner panels <NUM> are segmented, typically one to each fuel nozzle <NUM> and swirler <NUM>.

The cowl <NUM> extends radially between, and is secured to, the forwardmost ends of the combustor walls <NUM>, <NUM>. The cowl <NUM> includes a multiple of circumferentially distributed hood ports <NUM> that receive one of the respective multiple of fuel nozzles <NUM> and facilitates the direction of compressed air into the forward end of the combustion chamber <NUM> through a swirler opening <NUM>. Each fuel nozzle <NUM> may be secured to the diffuser case module <NUM> and project through one of the hood ports <NUM> and through the swirler opening <NUM> within the respective swirler <NUM>.

The forward assembly <NUM> introduces core combustion air into the forward section of the combustion chamber <NUM> while the remainder enters the outer annular plenum <NUM> and the inner annular plenum <NUM>. The multiple of fuel nozzles <NUM> and adjacent structure generate a blended fuel-air mixture that supports stable combustion in the combustion chamber <NUM>.

Opposite the forward assembly <NUM>, the outer and inner support shells <NUM>, <NUM> are mounted to a first row of Nozzle Guide Vanes (NGVs) 54A in the HPT <NUM>. The NGVs 54A are static engine components which direct core airflow combustion gases onto the turbine blades of the first turbine rotor in the turbine section <NUM> to facilitate the conversion of pressure energy into kinetic energy. The core airflow combustion gases are also accelerated by the NGVs 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>, a multiple of studs <NUM> extend from each of the liner panels <NUM>, <NUM> so as to permit a liner array (partially shown in <FIG>) of the liner panels <NUM>, <NUM> to be mounted to their respective support shells <NUM>, <NUM> with fasteners <NUM> such as nuts. That is, the studs <NUM> project rigidly from the liner panels <NUM>, <NUM> to extend through the respective support shells <NUM>, <NUM> and receive the fasteners <NUM> on a threaded section thereof (<FIG>).

A multiple of cooling impingement passages <NUM> penetrate through the support shells <NUM>, <NUM> to allow air from the respective annular plenums <NUM>, <NUM> to enter cavities <NUM> formed in the combustor walls <NUM>, <NUM> between the respective support shells <NUM>, <NUM> and liner panels <NUM>, <NUM>. The impingement passages <NUM> are generally normal to the surface of the liner panels <NUM>, <NUM>. The air in the cavities <NUM> provides cold side impingement cooling of the liner panels <NUM>, <NUM> that is generally defined herein as heat removal via internal convection.

A multiple of effusion passages <NUM> penetrate through each of the liner panels <NUM>, <NUM>. The geometry of the passages, e.g., diameter, shape, density, surface arcuate surface, incidence arcuate surface, 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 <NUM> allow the air to pass from the cavities <NUM> defined in part by a cold side <NUM> of the liner panels <NUM>, <NUM> to a hot side <NUM> of the liner panels <NUM>, <NUM> and thereby facilitate the formation of a thin, relatively cool, film of cooling air along the hot side <NUM>.

In one disclosed non-limiting embodiment, each of the multiple of effusion passages <NUM> are typically <NUM>" (<NUM>) in diameter and define a surface arcuate surface section of about thirty (<NUM>) degrees with respect to the cold side <NUM> of the liner panels <NUM>, <NUM>. The effusion passages <NUM> are generally more numerous than the impingement passages <NUM> and promote film cooling along the hot side <NUM> to sheath the liner panels <NUM>, <NUM> (<FIG>). 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 <NUM> and effusion passages <NUM> may be referred to as an Impingement Film Floatwall (IFF) assembly. A multiple of dilution passages <NUM> are located in the liner panels <NUM>, <NUM> each along a common axis D. For example only, the dilution passages <NUM> are located in a circumferential line W (shown partially in <FIG>). Although the dilution passages <NUM> are illustrated in the disclosed non-limiting embodiment as within the aft liner panels 74A, 74B, the dilution passages may alternatively be located in the forward liner panels 72A, 72B or in a single liner panel which replaces the fore/aft liner panel array. Further, the dilution passages <NUM> although illustrated in the disclosed non-limiting embodiment as integrally formed in the liner panels, it should be appreciated that the dilution passages <NUM> may be separate components. Whether integrally formed or separate components, the dilution passages <NUM> may be referred to as grommets.

With reference to <FIG>, in one disclosed non-limiting embodiment, each of the forward liner panels 72A, 72B, and the aft liner panels 74A, 74B in the liner panel array includes a perimeter rail 120a, 120b formed by a forward circumferential rail 122a, 122b, an aft circumferential rail 124a, 124b, and axial rails 126Aa 126Ab, 126Ba, 126Bb, that interconnect the forward and aft circumferential rail 122a, 122b, 124a, 124b. The perimeter rail <NUM> seals each liner panel with respect to the respective support shell <NUM>, <NUM> to form the impingement cavity <NUM> therebetween. That is, the forward and aft circumferential rail 122a, 122b, 124a, 124b are located at relatively constant curvature shell interfaces while the axial rails 126Aa 126Ab, 126Ba, 126Bb, extend across an axial length of the respective support shell <NUM>, <NUM> to complete the perimeter rail 120a, 120b that seals the forward liner panels 72A, 72B, and the aft liner panels 74A, 74B to the respective support shell <NUM>, <NUM>.

A multiple of studs <NUM> are located adjacent to the respective forward and aft circumferential rail 122a, 122b, 124a, 124b. Each of the studs <NUM> may be at least partially surrounded by posts <NUM> to at least partially support the fastener <NUM> and provide a stand-off between each forward liner panels 72A, 72B, and the aft liner panels 74A, 74B and respective support shell <NUM>, <NUM>.

The dilution passages <NUM> are located downstream of the forward circumferential rail 122a, 122b in the aft liner panels 74A, 74B to quench the hot combustion gases within the combustion chamber <NUM> by direct supply of cooling air from the respective annular plenums <NUM>, <NUM>. That is, the dilution passages <NUM> pass air at the pressure outside the combustion chamber <NUM> directly into the combustion chamber <NUM>.

This dilution air is not primarily used for cooling of the metal surfaces of the combustor shells or panels, but to condition the combustion products within the combustion chamber <NUM>. In this disclosed non-limiting embodiment, the dilution passages <NUM> include at least one set of circumferentially alternating major dilution passages 116A and minor dilution passages 116B (also shown in <FIG>). That is, in some circumferentially offset locations, two major dilution passages 116A are separated by one minor dilution passage 116B. Here, every two major dilution passages 116A are separated by one minor dilution passage 116B but may still be considered "circumferentially alternating" as described herein.

With reference to <FIG>, each of the forward liner panels 72A, 72B includes a forward section <NUM>, and an aft section <NUM> that defines an arcuate surface section <NUM> therebetween. That is, there is a smooth arcuate transition in the axial profile between the forward section <NUM> and the aft section <NUM> of the forward liner panels 72A, 72B profile to form a converging or diverging geometry in the inner flow path of the combustor. Combustor liners with such a gradual radius can eliminate interfaces that may result in steps, dead regions, cooling challenges, and/or adverse local aerodynamics.

The forward liner panels 72A, 72B include an arcuate surface section <NUM> that extends over an angle of about <NUM> to <NUM> degrees. The combustor liner extends across two segments of the combustor liner support shell <NUM>, <NUM> with the arcuate surface section <NUM> in the region where the combustor liner support shell <NUM>, <NUM> is formed with a complementary arcuate surface section <NUM>. It should be appreciated that in some embodiments, the complementary arcuate surface section <NUM> is conventional in that no modification need be performed to the combustor liner support shell <NUM>, <NUM> to utilize the forward liner panels 72A, 72B with the arcuate surface section <NUM>. That is, the aft circumferential rail 124a, 124b of the forward liner panel 72A, 72B is adjacent to the forward circumferential rail 122a, 122b of the aft liner panel 74A, 74B downstream of the combustor liner support shell <NUM>, <NUM>.

In this embodiment, the forward liner panel 72A, 72B defines about <NUM>% the length of the combustor and the aft liner panel 74A, 74B defines about <NUM>% the length of the combustor. That is, the forward liner panel 72A, 72B is longer than the aft liner panel 74A, 74B.

The non-linear axial profile of the forward liner panels 72A, 72B increases combustor durability and the ability to optimize the combustor design and performance. Combustor liners with a kink or bend can eliminate interfaces that result in steps, dead regions, cooling challenges and adverse local aerodynamics. Panels of this geometry edges are readily employed in cast and machined panel designs and incorporated in dual wall liners.

Claim 1:
A combustor (<NUM>) for a gas turbine engine (<NUM>) comprising:
a support shell (<NUM>, <NUM>) having a bend; and
a combustor liner including a forward liner panel (72A, 72B) mounted to the support shell (<NUM>, <NUM>) via a first multiple of studs (<NUM>), the forward liner panel (72A, 72B) comprising a forward section (<NUM>) and an aft section (<NUM>) that defines a
profile internal to the combustor (<NUM>) with an arcuate surface section (<NUM>) between the forward section (<NUM>) and the aft section (<NUM>), the aft section (<NUM>) being adjacent to the bend; and
the combustor liner further comprising an aft liner panel (74A, 74B) mounted to the support shell (<NUM>, <NUM>) via a second multiple of studs (<NUM>) downstream of the forward liner panel (72A, 72B);
characterised in that:
the arcuate surface section (<NUM>) extends over an angle between <NUM> to <NUM> degrees with respect to a cold side (<NUM>) of the forward liner panel (72A, 72B);
the support shell (<NUM>, <NUM>) is formed with an arcuate surface section (<NUM>) that is complementary to the arcuate surface section (<NUM>) of the forward liner panel (72A, 72B), wherein the arcuate surface section (<NUM>) of the forward liner panel (72A, 72B) is in the region of the arcuate surface section (<NUM>) of the support shell (<NUM>, <NUM>);
the combustor liner extends across two segments of the support shell (<NUM>, <NUM>).