Abstract: A seal assembly for a rotary machine. The seal assembly includes a rotor and a stator. The rotor is rotatable about a rotational axis and has a rotor seal face. The stator has a stator seal face. The stator seal face is positioned opposite the rotor seal face and faces the rotor seal face with a gap therebetween. A portion of one of the rotor and the stator is formed of (i) a shape memory alloy or (ii) a first metal and a second metal with the second metal having a coefficient of thermal expansion different from the first metal. The seal assembly is characterized by a seal clearance compliance ratio (SCCR) from 20% to 90%.

Abstract: Separators for use with gas turbine engines are disclosed herein. An example cabin bleed system disclosed herein includes a bleed line coupled to a gas turbine engine, a separator defining a cavity fluidly coupled to the bleed line, the separator including a first ring including a first opening fluidly coupled to the cavity. a second ring surrounding the first ring, the second ring including a second opening fluidly coupled to a cabin of an aircraft, and a filter disposed in a fluid pathway extending between the first opening and the second opening.

Abstract: Systems, apparatus, articles of manufacture, and methods are disclosed to improve fan operability control using smart materials. An engine comprising an engine surface in an airflow path, a sensor positioned on the engine surface, and a smart-material-based feature positioned on the engine surface, the smart-material-based feature triggered to modify the airflow path when the sensor outputs an indication of a detected deviation from a reference value of an operating parameter of the engine.

Abstract: A mounting assembly for a gearbox assembly of a gas turbine engine includes at least one mounting member configured to mount a gear of the gearbox assembly to a component of the gas turbine engine, the at least one mounting member characterized by a lateral impedance parameter, a bending impedance parameter, and a torsional impedance parameter. A gas turbine engine includes the mounting assembly. The at least one mounting member may be a flex mount, a fan frame, or a flex coupling. The gas turbine engine includes an electric power system including at least one electric machine. The electric power system includes a plurality of power converters and a plurality of power distribution management units. At least two of the plurality of power converters or the plurality of power distribution management units are integrated together in a single housing.

Abstract: Disclosed herein are example variable flowpath casings for blade tip clearance control. An example casing for a turbine engine includes a first annular substrate extending along an axial direction; a second annular substrate positioned radially inward relative to the first annular substrate, the second annular substrate movably coupled to the first annular substrate; and an actuator coupled to the second annular substrate such that a force applied by the actuator moves the second annular substrate relative to the first annular substrate to adjust a tip clearance.

Abstract: Example bearing lubrication system and methods of operating the same are disclosed herein. An example closed loop system to provide a lubricant to a fluid pump includes a lubrication flow network disposed within the fluid pump; a sensor fluidly coupled to the fluid pump to measure a condition of a fluid that is to enter the lubrication flow network; a first transport bus fluidly coupled to the lubrication flow network, the first transport bus to transport an inert gas; a controller to actuate a valve fluidly coupled to the first transport bus, the controller to transmit signals to the valve based on the condition of the fluid to cause the valve to open or close; and a separator fluidly coupled between an outlet of the fluid pump and the first transport bus, the separator to separate the fluid and the inert gas.

Abstract: Seals for a gas turbine engine are disclosed herein. An example sealing apparatus for a gas turbine engine includes a first end having a first contact surface to contact a first component of a rotating machine, a second end having a second contact surface to contact a second component of the rotating machine, the second component positioned in proximity to the first component such that a cavity is formed therebetween, the apparatus dividing the cavity into a first cavity portion and a second cavity portion, and a central portion extending the first end from the second end, the central portion defining a first chamber and a second chamber, the first chamber in fluid communication with the first cavity portion, the second chamber in fluid communication with the second cavity portion, the first chamber is fluidly isolated from the second chamber.

Abstract: Disclosed herein are example variable flowpath casings for blade tip clearance control. An example casing for a turbine engine includes an annular substrate extending along an axial direction, the annular substrate including a first surface and a second surface that is radially inward relative to the first surface; an actuator structure coupled to the second surface of the annular substrate; and a smart structure cantilevered from the second surface of the annular substrate, the smart structure including: a support structure, a first region of the support structure coupled to the second surface of the annular substrate, a second region of the support structure coupled to the actuator structure; and a radially inward surface defining a variable surface, the support structure to move the variable surface in a radial direction.

Abstract: A dome-deflector assembly for a combustor of a gas turbine includes a dome portion having a dome-side swirler opening therethrough, and a deflector portion having a deflector-side swirler opening therethrough. The dome portion and the deflector portion are connected together to form a dome-deflector cavity therebetween. The deflector portion includes a first-side surface and a second-side surface, the second-side surface being arranged within the dome-deflector cavity, and the second-side surface including a plurality of stress-relief contours arranged in a pattern about the deflector-side swirler opening.

Abstract: This disclosure is directed to seal assemblies for a turbomachine. The seal assemblies include stationary and rotating components and at least one interface between the stationary and rotating components. The seal assembly further includes a self-lubricating lattice element made of carbon or a carbon-based material. The self-lubricating lattice element may have a porous and compressible microstructure capable of retaining a liquid substance. During engine operation, the self-lubricating lattice element can compress and expel some portion of the liquid substance to form a liquid-containing layer between the rotating and stationary components of the seal assembly. Also described herein are self-lubricating wear sleeves for use in other portions of the turbomachine.

Abstract: A turbine engine with a compressor section, a combustion section, and a turbine section in serial flow arrangement along an engine centerline. The combustion section includes a primary combustor liner including an inner liner and an outer liner. A dome wall and a primary dome inlet are located in the dome wall. The outer liner defines at least one opening downstream from the primary dome inlet. A primary combustion chamber and a set of secondary combustors are fluidly coupled to the primary combustion chamber at the at least one opening.

Abstract: A gearbox assembly for a gas turbine engine includes a hollow gear. The hollow gear has one or more reservoirs configured to store a lubricant and one or more passages configured to supply the lubricant to one or more gear meshes in the gearbox assembly. The hollow gear provides a dampening system configured to absorb vibrations and loading experienced in the gearbox assembly and configured to provide impact resistance for the gearbox assembly.

Abstract: A gearbox assembly includes a first gear and a second gear. The first gear includes a plurality of first gear teeth. The second gear includes a plurality of second gear teeth. The plurality of first gear teeth and the plurality of second gear teeth mesh with each other as the first gear and the second gear rotate. A profile shape of at least one first gear tooth of the first gear is characterized by a total profile modification between 66 micrometers and 120 micrometers.

Abstract: Aero-acoustically damped bleed valves are disclosed. An example variable bleed valve apparatus comprises a variable bleed valve door to actuate the variable bleed valve apparatus, and a variable bleed valve port including an upstream edge and a downstream edge, the VBV port to define a secondary flowpath, the VBV door to cover the VBV port in a closed position, and a vortex device at the upstream edge of the variable bleed valve port, the vortex device including a vorticity generating feature along the upstream edge of the variable bleed valve port.

Abstract: This disclosure is directed to cooling systems for turbomachine seal assemblies. The seal assemblies include stationary and rotating components and at least one interface between the stationary and rotating components. During operation of the turbomachine, the components of the seal assemblies generate heat, which is removed from the seal assembly by the cooling systems of this disclosure. In some examples, the cooling system includes a reservoir of lubricant that is distributed to a face of one of the components of the seal assembly. In other examples, the cooling system includes a lattice structure capable of retaining lubricant against the components of the seal assembly, or a pressurized jet of lubricant directed against the components of the seal assembly. The cooling systems can further include a combination of deflectors, hairpin members, and channels for distributing the lubricant to the components of the seal assemblies.

Abstract: A seal assembly for a gas turbine includes a seal case, and a seal shoe arranged within the seal case. The seal shoe includes a main seal portion that is configured to engage with a radial bearing engagement portion of a rotor shaft, a radial bearing surface that engages with the radial bearing engagement portion of the rotor shaft, at least one hydrodynamic lift cavity extending through the radial bearing surface, and at least one airflow passage extending from a first side of the seal shoe through the at least one hydrodynamic lift cavity and providing airflow communication between the first side of the seal shoe and the at least one hydrodynamic lift cavity.

Abstract: A turbine engine is provided. The turbine engine defines an axial direction and a radial direction, and includes: a rotor; a stator comprising a carrier; a seal assembly disposed between the rotor and the stator, the seal assembly defining a high-pressure side and a low-pressure side and comprising a plurality of seal segments, the plurality of seal segments including a seal segment having a seal face forming a fluid bearing with the rotor, a lip, and a body, the lip extending from the body along the axial direction of the turbine engine on the high-pressure side and including an outer pressurization surface along the radial direction of the turbine engine; and a seal support assembly, the seal support assembly comprising a spring arrangement extending between the carrier and the first seal segment to counter a pressure on the outer pressurization surface during operation of the turbine engine.

Abstract: A heat shield for a combustor of a gas turbine engine. The heat shield includes an annular ring having an axial direction, a radial direction, and a circumferential direction. The annular ring includes a plurality of circumferential segments. Each circumferential segment of the plurality of circumferential segments is disconnected from an adjacent circumferential segment of the plurality of circumferential segments to allow for thermal growth of each circumferential segment during operation of the combustor.

Abstract: A gas turbine engine is provided having a plurality of outlet guide vanes, each defining an internal thermal fluid passageway. The engine defining an Outlet Guide Vane Cooling Capacity greater than 0.01 and less than 13, wherein OGVCC equals: [ HTSA OGV × BPR ( BPR + 1 ) × C a ? i ? r × ( T i ? n ? l ? e ? t - T a ? i ? r ) × v f ? l ? i ? g ? h ? t × D f ? a ? n F ? n T ? o ? t ? a ? l × v tip ? speed × ? ? H ] 1 / 3 , and wherein HTSA OGV = N v ? a ? n ? e × D fan 2 × ( 1 - R OGV ? _ ? ratio 2 ) 2 × sin ? ( 180 / N v ? a ? n ? e ) × sin ? ? OGV × f OGV .

Abstract: A liner assembly for a combustor. The liner assembly includes a liner defining a combustion chamber of the combustor. The liner includes a looped section at a forward end of the liner. The looped section includes one or more bends such that the looped section forms a compliant joint and vibrations are dampened through the liner downstream of the looped section.