Abstract: Structures for reducing forward loads in compressors are described. A compressor includes inner and outer circumferential support structures positioned concentrically around a central axis, and an aft-most stage including a vane extending radially inward from the outer circumferential support structure. An axial length of the aft-most stage is defined by a spacer arm of the inner circumferential support structure. The vane includes a root, a tip, and a trailing edge extending between the root and the tip. A ratio of a first radial distance between a first point located at an intersection of the tip of the vane and the trailing edge to a radially inner wall of the spacer arm of the inner circumferential support structure to a second radial distance between the first point and a second point located at an intersection of the root of the vane and the trailing edge is between 0.95 and 4.5.

Abstract: A method of repairing a multi-layer component of an engine in-module includes identifying a damage location on the multi-layer component. The damage location extends at least partially into a ceramic-based top layer of the multi-layer component. The method further includes depositing a material composition onto the damage location at a first temperature range so as to cover the damage location. The material composition includes one or more sintering additives. Further, the method includes applying localized curing to the material composition deposited at the damage location at a second temperature range, the second temperature range being higher than the first temperature range.

Abstract: Structures for reducing forward loads in compressors are described. compressor includes inner and outer circumferential support structures positioned concentrically around a central axis, and an aft-most stage including a vane extending radially inward from the outer circumferential support structure. An axial length of the aft-most stage is defined by a spacer arm of the inner circumferential support structure. The vane includes a root, a tip, and a trailing edge extending between the root and the tip. A ratio of a first radial distance between a first point located at an intersection of the tip of the vane and the trailing edge to a radially inner wall of the spacer arm of the inner circumferential support structure to a second radial distance between the first point and a second point located at an intersection of the root of the vane and the trailing edge is between 0.95 and 4.5.

Abstract: A gas turbine engine including a compressor section including a compressor and a turbine section located downstream of the compressor section. The turbine section including a high pressure turbine, a low pressure turbine, a sump positioned between the high pressure turbine and the low pressure turbine, and a rotating cross flow arrangement defining a plurality of guiding passages fluidly coupling the compressor to the sump.

Abstract: A mixer assembly for a gas turbine engine. The mixer assembly includes a housing and a fuel injection port. The housing has a passage formed therein, and the housing includes a passage wall facing the passage. The fuel injection port is fluidly connected to a fuel source and is configured to inject a hydrocarbon fuel into the passage. At least a portion of the passage wall is a coated passage wall. The coated passage wall is (i) coated with a layer of catalyst particles comprising a rare earth oxide component and (ii) located downstream of the fuel injection port.

Abstract: Methods for repairing a defect in a multilayered coating on a silicon-containing substrate are provided, along with the resulting compared component. The method includes: applying a first patch slurry into the defect; drying the first patch slurry to form a dried first patch layer; applying a second patch slurry into the defect and on the dried first patch layer; and drying the second patch slurry to form a second dried patch layer.

Abstract: Methods for repairing a thermal barrier coating deposited on a component with localized spallation of the thermal barrier coating includes depositing a primer slurry on a thermally grown oxide of the component exposed by the localized spallation, depositing a ceramic slurry on the primer slurry, and heating the primer slurry and the ceramic slurry. The primer slurry includes a primer that includes at least one of a metal and a metal oxide. The ceramic slurry includes a ceramic material, a ceramic slurry binder material, and a ceramic slurry fluid carrier. Heating the primer slurry and the ceramic slurry forms a first chemical bond between the primer and the thermally grown oxide and a second chemical bond between the primer and the ceramic material.

Abstract: A gas turbine engine having a compressor section, a combustion section, and a turbine section in serial flow arrangement to define a diffuser cavity between the compressor section and the combustor section, and an aft cavity. A compressor discharge pressure duct fluidly draws air from the diffuser cavity, passes it through a heat exchanger to cool the air, and then supplies the cooled air to the aft cavity.

Abstract: A gas turbine engine including a compressor section including a compressor and a turbine section located downstream of the compressor section. The turbine section including a high pressure turbine, a low pressure turbine, a sump positioned between the high pressure turbine and the low pressure turbine, and a rotating cross flow arrangement defining a plurality of guiding passages fluidly coupling the compressor to the sump.

Abstract: A gas turbine engine includes a casing, a rotor assembly rotatably mounted within the casing, the rotor assembly comprising a forward rotor portion defining a forward rotor cavity and an aft rotor portion defining an aft rotor cavity, the aft rotor cavity being fluidly isolated from the forward rotor cavity, and a cooling system. The cooling system includes a forward cooling fluid supply, a forward supply line providing fluid communication between the forward cooling fluid supply and the forward rotor cavity, an aft cooling fluid supply, an aft supply line providing fluid communication between the aft cooling fluid supply and the aft rotor cavity, and a flow control system fluidly coupled to the forward supply line and the aft supply line for regulating a flow of forward cooling fluid and a flow of aft cooling fluid.

Abstract: A turbine engine having a drive shaft rotatable about an axis, a multi-stage compressor, a turbine section, a thrust bearing, and a balance cavity. The thrust bearing being provided between the drive shaft and at least a portion of the multi-stage compressor section and rotationally supporting the drive shaft. During operation of the turbine engine, a first axial force is applied to the thrust bearing by the drive shaft and a second axial force is applied to the thrust bearing in an opposite direction of the first axial force by the balance cavity.

Abstract: A coating method is provided. The coating method includes applying, via electrophoretic deposition or slurry deposition, an overcoat composition on an outer surface of a thermal barrier coating system on a substrate. The overcoat composition includes a coating material comprising a plurality of particles having a particle size of less than 1000 nm. The method includes sintering the overcoat composition in the presence of one or more sintering aids to form an overcoat layer having a surface roughness of less than 1 micrometer.

Abstract: In one aspect, the present disclosure is directed a gas turbine engine including a compressor, a combustor, and a turbine in a serial flow arrangement. The compressor includes a casing having an inner wall and an outer wall. The inner wall defines a passageway for airflow through the compressor. The casing defines a bleed cavity between the inner wall and the outer wall. The inner wall has an opening. The casing defining a channel between the opening and the bleed cavity to direct airflow into the bleed cavity. The channel is defined by a first wall and a second wall of the casing. The second wall being downstream of the first wall. A surface of the first wall has a pattern to reduce flow separation in the channel.

Abstract: In one aspect, the present disclosure is directed a gas turbine engine including a compressor, a combustor, and a turbine in a serial flow arrangement. The compressor includes a casing having an inner wall and an outer wall. The inner wall defines a passageway for airflow through the compressor. The casing defines a bleed cavity between the inner wall and the outer wall. The inner wall has an opening. The casing defining a channel between the opening and the bleed cavity to direct airflow into the bleed cavity. The channel is defined by a first wall and a second wall of the casing. The second wall being downstream of the first wall. A surface of the first wall has a pattern to reduce flow separation in the channel.

Abstract: A gas turbine engine includes a casing, a rotor assembly rotatably mounted within the casing, the rotor assembly comprising a forward rotor portion defining a forward rotor cavity and an aft rotor portion defining an aft rotor cavity, the aft rotor cavity being fluidly isolated from the forward rotor cavity, and a cooling system. The cooling system includes a forward cooling fluid supply, a forward supply line providing fluid communication between the forward cooling fluid supply and the forward rotor cavity, an aft cooling fluid supply, an aft supply line providing fluid communication between the aft cooling fluid supply and the aft rotor cavity, and a flow control system fluidly coupled to the forward supply line and the aft supply line for regulating a flow of forward cooling fluid and a flow of aft cooling fluid.

Abstract: Coated components, along with methods of their formation, are provided. The coating may include a component having a surface and a coating over the surface of the component. The coating may include a first plurality of coarse micrograins having an coarse median grain size of greater than 10 ?m to 20 ?m, a second plurality of medium micrograins having a medium median grain size of 4 ?m to 10 ?m, a third plurality of fine micrograins having a fine median grain size of 0.001 ?m to less than 4 ?m. Each of the coarse micrograins, medium micrograins, and fine micrograins may include, independently, a yttria-stabilized zirconia (YSZ) constituent having a chemical formula of (ZrO2)(1?x)(Y2O3)x with x is from greater than 0 to less than 1.

Abstract: A turbine engine having a drive shaft rotatable about an axis, a multi-stage compressor, a turbine section, a thrust bearing, and a balance cavity. The thrust bearing being provided between the drive shaft and at least a portion of the multi-stage compressor section and rotationally supporting the drive shaft. During operation of the turbine engine, a first axial force is applied to the thrust bearing by the drive shaft and a second axial force is applied to the thrust bearing in an opposite direction of the first axial force by the balance cavity.

Abstract: An example compressor bleed slot apparatus includes an annular compressor casing, a blade row mounted for rotation about a centerline axis inside the compressor casing, a bleed slot passing through the forward section of the compressor casing, wherein the bleed slot is bounded by inboard and outboard walls defined within the compressor casing, the bleed slot extending along a slot axis, at least a portion of the bleed slot lying within an axial extent of the blade row, an array of struts interconnecting the inboard and outboard walls, and an annular supplemental flange extending radially outward from the forward section of the compressor casing, wherein at least a portion of the supplemental flange is axially positioned within an axial extent of the bleed slot, wherein the supplemental flange includes a necked-down portion adjacent the forward section of the compressor casing, the necked-down portion positioned axially between an inlet of the bleed slot and an outlet of the bleed slot.

Abstract: A sprayable thermal barrier coating powder mixture for a gas turbine engine includes: a dry composition having a low surface area ceramic powder having a median particle size distribution greater than 5 microns and less than 50 microns, and a high surface area ceramic powder having a median particle size distribution smaller than 5 microns, wherein the low surface area ceramic powder makes up at least 50% by weight of the dry composition of the sprayable thermal barrier coating powder mixture.

Abstract: A corrosion resistant coating including a matrix, corrosion resistant particles dispersed throughout the matrix, and a glass-forming additive is disclosed. The glass-forming additive and one or more materials in the matrix form a glassy-phase when cured. Coated gas turbine engine components and methods for coating components are also disclosed.