Abstract: A method for determining wear mode and surface characteristics of wear on a gas turbine engine component, comprising the steps of obtaining images of a wear scar on a gas turbine engine component; feeding the images to an image model; identifying geometry of the wear scar from the images and the image model; identifying a wear mode from the geometry of the wear scar; obtaining surface characterization of the wear scar; analyzing the surface characterization to identify tribo-layer details; identifying constituents of the tribo-layer from the tribo-layer details; and characterizing wear on the gas turbine engine component based upon the wear mode and the constituents of the tribo-layer. A system is also disclosed.

Abstract: A system includes a laser drilling apparatus having a laser to drill into coating layers of a component. The component includes cooling holes through a substrate and an existing layer. The existing coating layer is stripped from the component. The locations of the cooling holes are mapped using a coordinate measurement machine that generates a macro file of the locations. The macro file is loaded onto the laser drilling apparatus. A new coating layer is applied that covers the cooling holes. The laser drilling apparatus collects data at the locations mapped to the macro file of the component having a new coating layer. A modification is made to the locations of the at least one cooling holes based on a parameter in the collected data. The laser drilling apparatus drills using the laser at the modified locations into the new coating layer.

Abstract: A process of treating a carbon-carbon composite material including modifying a surface of the carbon-carbon composite material to increase a hydrophilic surface functionality; applying a plasma surface treatment to the carbon-carbon composite material; infiltrating the carbon-carbon composite with an oxidation inhibiting composition; applying filler particles to the surface of the carbon-carbon composite material to make the oxidation inhibiting composition resistant to shear flow and heating the carbon-carbon composite at a predetermined temperature sufficient to form a deposit from the oxidation inhibiting composition in at least some pores of the carbon-carbon composite.

Abstract: A composite double flange full-hoop ring has a central section that extends axially between first and second fillets with first and second axial flanges, respectively, with the second fillet having a locally increased thickness. A plurality of full-length woven plies extends a full-length of the first flange, first fillet, central section, second fillet, and second flange to define an inner surface and an outer surface of the double flange full-hoop ring. Disposed therebetween, at least one first partial-length woven ply extends a first length consisting of the first flange, first fillet, central section, and a first portion of the second fillet, and at least one second partial-length woven ply extends a second length consisting of the second flange and a second portion of the second fillet that at least partially overlaps the first portion of the second fillet to provide the locally increased thickness of the second fillet.

Abstract: An apparatus for adapting a build stand includes a plate structure having a lower side configured to mount to a first build stand configuration, and an upper side configured to present a different build stand configuration. The adapter convert a build stand that is suitable for the mounting structure of one engine assembly to receive a different mounting structure of a different engine assembly that that is not otherwise compatible with the build stand.

Abstract: A method of forming a ceramic matrix composite includes preparing a plurality of fibrous ceramic plies by applying a binder solution loaded with ceramic particles to each of a first subset of plies, and applying the binder solution without ceramic particles to each of a second subset of plies. The method further includes arranging the plurality of plies to form a preform with a gradient along a thickness direction of the preform by stacking the first subset of plies to form a first zone of the preform, and stacking individual ones of the second subset of the plies on both sides of the first zone to form a second zone on each side of the first zone. Loading of the ceramic particles is higher in the first zone than in the second zones. The method further includes densifying the preform with a ceramic matrix.

Abstract: An aircraft propulsion system includes a water injection system that includes a plurality of nozzles configured to inject water onto at least one of a plurality of compressor stages and turbine stages, and a controller that is configured to operate the water injection system and a drive system to rotate the compressor and turbine section while injecting water through the plurality of nozzles for cooling a portion of one of the compressor section, the turbine section and a case structure.

Abstract: An operating method is provided during which a plurality of start parameters for a gas turbine engine are determined. A first of the start parameters is indicative of a temperature of air at an inlet into the gas turbine engine. A second of the start parameters is indicative of a pressure of the air at the inlet. Rotation of a rotating assembly of the gas turbine engine is driven. The rotating assembly includes a compressor rotor and a turbine rotor. The compressor rotor is within a compressor section of the gas turbine engine. The turbine rotor is within a turbine section of the gas turbine engine. Fuel is directed into a combustion chamber of the gas turbine engine based on the start parameters and a speed parameter. The speed parameter is indicative of a speed of the rotation of the rotating assembly. A mixture of the fuel and compressed air within the combustion chamber is ignited to start the gas turbine engine.

Abstract: An assembly is provided for an aircraft propulsion system. This assembly includes an open propulsor rotor and an open guide vane structure. The open propulsor rotor is configured to rotate about an axis. The open propulsor rotor includes a plurality of open propulsor blades arranged circumferentially about the axis in a blade array. Each of the open propulsor blades projects spanwise into an environment external to the aircraft propulsion system to a respective blade tip. The open propulsor blades include a first span propulsor blade and a second span propulsor blade. The first span propulsor blade has a first span length. The second span propulsor blade has a second span length that is less than the first span length. The open guide vane structure is axially next to the open propulsor rotor.

Abstract: An assembly for an aircraft includes a first propulsion system and a second propulsion system. A first open propulsor rotor is configured to rotate in a first rotational direction about a first axis. A first engine core includes a first rotating structure with a first turbine rotor. A first rotating structure is configured to rotate in the first rotational direction about the first axis and drive rotation of the first open propulsor rotor through a first geartrain. A second open propulsor rotor is configured to rotate in a second rotational direction about a second axis that is opposite the first rotational direction. A second engine core includes a second rotating structure with a second turbine rotor. A second rotating structure is configured to rotate in the first rotational direction about the second axis and drive rotation of the second open propulsor rotor through a second geartrain.

Abstract: A compressor diffuser vane and turbine nozzle vane flow network including at least one suction point formed in a suction side of the compressor diffuser vane; a suction flow passage fluidly coupled to the at least one suction point; a main artery formed within an outer platform, the outer platform in operative communication with the compressor diffuser vane, the main artery formed within the turbine nozzle vane, the turbine nozzle vane in operative communication with the outer platform, the main artery formed within an inner platform, the inner platform in operative communication with the turbine nozzle vane; film hole passages fluidly coupled to the main artery; and at least one film hole formed in the inner platform, the at least one film hole fluidly coupled to the main artery.

Abstract: An aircraft propulsion system includes a heat exchanger assembly that is operable in a heat exchanger configuration where the exhaust gas flow is in thermal communication with a working fluid flow, a bypass configuration where the exhaust gas flow passes through the heat exchanger assembly without thermal communication with the working fluid flow, and an intermediate configuration where a first portion of the exhaust gas flow passes through the heat exchanger and a second portion of the exhaust gas flow is bypassed around the heat exchanger.

Abstract: A method for operating a rotating detonation engine, having a radially outer wall extending along an axis; a radially inner wall extending along the axis, wherein the radially inner wall is positioned within the radially outer wall to define an annular detonation chamber having an inlet and an outlet, wherein the method includes flowing liquid phase fuel along at least one wall of the radially inner wall and the radially outer wall in a direction from the outlet toward the inlet to cool the at least one wall and heat the liquid fuel to provide a heated liquid fuel; flowing the heated liquid fuel to a mixer at the inlet to reduce pressure of the heated liquid fuel, flash vaporize the heated liquid fuel and mix flash vaporized fuel with oxidant to produce a vaporized fuel-oxidant mixture; and detonating the mixture in the annular detonation chamber.

Abstract: An aircraft powerplant includes a turbine engine, a first fluid system, an electric machine system, a second fluid system, a heat exchanger and a control system. The first fluid system directs a first fluid through the first fluid circuit using a first fluid pump to service engine components during a warmup mode. The electric machine system includes an electric machine configured to drive operation of the first fluid pump during the warmup mode. The second fluid system directs a second fluid through a second fluid circuit such that electronics heat energy generated by power electronics is transferred into the second fluid and cools the power electronics during the warmup mode. The heat exchanger is configured to transfer the electronics heat energy from the second fluid in the second fluid circuit into the first fluid in the first fluid circuit to heat the first fluid during the warmup mode.

Abstract: A vane structure includes a structure passage that extends internally from a passage inlet to a passage outlet. The passage inlet fluidly couples the structure passage to an external environment outside of the vane structure. The first heat exchanger is disposed within the vane structure along the structure passage between the passage inlet and the passage outlet. The second heat exchanger is disposed within the vane structure along the structure passage between the passage inlet and the passage outlet. The electric machine system includes a plurality of electric components including a first electric machine and a first electric machine controller for the first electric machine. The first electric machine is configurable as an electric motor and/or an electric generator. The first fluid circuit is configured to cool and/or lubricate a first of the electric components. The first fluid circuit includes the first heat exchanger and/or the second heat exchanger.

Abstract: Disclosed herein is a method of reinforcing a composite comprising determining a location of a first cooling hole in a plurality of plies; where a cooling gas is transported through the cooling hole; disposing a z-fiber in the plurality of plies at a location proximate to where the first cooling hole will be located; where the z-fiber enters the plurality of plies at either an upper surface or a lower surface; and where the z-fiber traverses a portion of the plurality of plies in the z-direction proximate to the first cooling hole; and traverses the plurality of plies in an x or y direction further away from the first cooling hole; where the z-direction is in the thickness direction of the plurality of plies and where the x and y-direction are perpendicular to the z-direction.

Abstract: A system is provided for an aircraft. This system includes a turbine engine and a cooling system. The turbine engine extends axially along an axis. The turbine engine includes a core flowpath, a compressor section, a combustor section and a turbine section. The core flowpath extends through the compressor section, the combustor section and the turbine section from an airflow inlet into the core flowpath to a combustion products exhaust from the core flowpath. The cooling system is configured to air cool the turbine engine following shutdown of the turbine engine. The cooling system includes a plurality of electric air movers. The electric air movers are arranged axially along the turbine engine.

Abstract: An assembly is provided for an aircraft powerplant. This assembly includes a differential geartrain, a first rotating assembly, a second rotating assembly, a first actuator and a second actuator. The differential geartrain includes a sun gear, a ring gear, a plurality of intermediate gears and a carrier. The first rotating assembly is coupled to the differential geartrain through the carrier. The first rotating assembly includes a first turbine rotor. The second rotating assembly is coupled to the differential geartrain through the ring gear. The second rotating assembly includes a second turbine rotor. The first actuator is coupled to the differential geartrain through the ring gear. The second actuator is coupled to the differential geartrain through the sun gear.

Abstract: A method can comprise: performing a finite element static analysis of an inspected blade of an inspected bladed rotor, the inspected blade having a repair blend profile modeled thereon, the repair blend profile exceeding a threshold repair size; performing a finite element modal analysis of the inspected blade having the repair blend profile; performing a fatigue assessment based on results from the finite element static analysis and the finite element modal analysis, the fatigue assessment including limits based on material properties of the inspected blade, the material properties based on test results at a threshold significance level; and repairing the inspected bladed rotor with the repair blend profile in response to the fatigue assessment meeting a deterministic criteria.

Abstract: An ultrasonic impact grinder apparatus configured to generate an ultrasonic displacement along a central axis is provided with a non-line-of-sight tool having a tool body axially aligned with the central axis, a tool tip, a narrowing horn disposed between the tool body and the tool tip, and a curved or cutout portion disposed on the non-line-of-sight tool to permit ultrasonic impact grinding at the tool tip without a line-of-sight of the tool tip along the central axis. The non-line-of-sight tool of the ultrasonic impact grinder apparatus can be positioned relative to the component such that the tip of the non-line-of-sight tool is disposed to drill non-line-of-sight holes into the component without interfering with the component in order to drill the non-line-of-sight hole in the component.