Abstract: A system and method for optimizing performance of an aircraft or boat through detection and trending of engine deterioration based on the performance of the vehicle's gas turbine. The system and method detects declines in engine power due to either of in-transit events or over the extended lifetime of the engine, the declines being due to routine engine part aging or an event. As engine power gradually or suddenly deteriorates, the system and method lowers a maximum operating line which defines the safe limits for peak engine power consumption during flight. For in-transit events, the system and method detects when actual power consumption is approaching the current maximum operating line. The controller may then automate changes to operations of entirely separate aircraft systems, such as rebalancing electrical energy consumption by various non-engine elements of the aircraft.

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: Input is received. A conformance is determined. When conformance is determined, a raw torque is measured from a torque sensor at an engine; the raw torque is calibrated using the torque sensor parameters to produce a calibrated torque value, non-torque parameters associated with the engine are measured and the non-torque parameters are applied to a lookup structure to obtain an expected torque value. A separation of the calibrated torque value and the expected torque value is determined and based upon the separation, an operation of the torque sensor is controlled.

Abstract: A lubricant system for supplying lubrication to a component in a turbine engine includes a lubricant reservoir, a supply line fluidly coupling the lubricant reservoir to the component in the turbine engine, a scavenge line fluidly coupling the component to the lubricant reservoir, and a bypass line fluidly coupling the supply line to the scavenge line and bypassing the component.

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: A gas turbine engine includes a fan located at a forward portion of the gas turbine engine, and a compressor section and a turbine section arranged in serial flow order. The compressor section and the turbine section together define a core airflow path. A rotary member is rotatable with the fan and with a low pressure turbine of the turbine section. The low pressure turbine includes a rotating drum to which a first airfoil structure is connected and extends radially inward toward the rotary member. A torque frame connects the rotating drum to the rotary member and transfers torque from the first airfoil structure mounted to the rotating drum to the rotary member. The torque frame includes an inner disk mounted to the rotary member, an outer ring and a second airfoil structure formed separately from the outer ring and connected thereto by a releasable connecting structure. The second airfoil structure extends radially inward from the outer ring toward the inner disk.

Abstract: A planetary gear system includes a carrier including a first side having a first oil manifold opening therethrough, and a second side having a second oil manifold opening therethrough, a first oil manifold and a second oil manifold. The first oil manifold has a first flange portion floatingly engaging with the first oil manifold opening, and a second flange portion engaging with the second oil manifold opening of the carrier, the second flange portion including a first oil manifold mounting flange. The second oil manifold has a second oil manifold flange portion engaging with the second oil manifold opening of the carrier, and including a second oil manifold mounting flange. At least one connecting member connects the first oil manifold mounting flange and the second oil manifold mounting flange so as to mount the first oil manifold and the second oil manifold to the carrier.

Abstract: Flow-metering fuel systems and related methods are disclosed. An example apparatus includes a pipe defining a flow path for fuel, the pipe fluidly coupled to a combustor, a first portion of the pipe having a first cross-sectional area, a second portion of the pipe having a second cross-sectional area smaller than the first cross-sectional area, the second portion downstream of the first portion, and an actuator to adjust a flow rate of the fuel in the pipe based on a first pressure of the fuel in the first portion of the pipe, a second pressure of the fuel in the second portion of the pipe, and a temperature of the fuel.

Abstract: A turbomachinery apparatus (10) includes: a turbine (22), including: a turbine component (36) defining at least one arcuate flowpath surface (40); an array of axial-flow turbine airfoils (46) extending from the flowpath surface, the turbine airfoils defining spaces therebetween; and a plurality of splitter airfoils (146) extending from the at least one flowpath surface, in the spaces between the turbine airfoils, each splitter airfoil having opposed pressure and suction sides extending between a leading edge and a trailing edge, wherein the splitter airfoils have a thickness ratio which is less than a thickness ratio of the turbine airfoils.

Abstract: A build table assembly for supporting a component, the build table assembly including a base having an upper surface, a frame extending from the upper surface of the base, the frame defining an upper wall; a rim formed on the upper wall of the frame, and a plurality of fixture segments positioned on the rim opposite the upper wall of the frame, the plurality of fixture segments defining an arcuate build surface on which the component is built, the plurality of fixture segments spaced apart from one another.

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.

Abstract: A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows: ( T O ? U ? T A D ? T ? E ? x ? i ? t ) 2 * E ? G ? T A H ? P ? C ? E ? x ? i ? t * 1 ? 0 - 1 ? 1 ; wherein CSP is greater than 0.0001194×EGT2?0.103×EGT+22.

Abstract: A gearbox assembly includes a plurality of planet gears. At least one planet gear of the plurality of planet gears includes a layshaft that supports a first stage planet gear and a second stage planet gear. The at least one planet gear is characterized by a Radial Stage Distance greater than or equal to 30 millimeters and less than or equal to 100 millimeters. The Radial Stage Distance is Rrm1?Rrt2. Rrm1 is a first stage rim inner radius of the first stage planet gear and Rrt2 is a second stage root radius of the second stage planet gear.

Abstract: A turbomachine engine can include a fan assembly, a vane assembly, a core engine, and a gearbox. The fan assembly can include a plurality of fan blades. The vane assembly can include a plurality of vanes, and the vanes can be disposed aft of the fan blades. The core engine can include one or more compressor sections and one or more turbine sections. The gearbox includes an input and an output. The input is coupled to the one or more turbine sections of the core engine and comprises a first rotational speed, the output is coupled to the fan assembly and has a second rotational speed. A gear ratio of the first rotational speed to the second rotational speed is within a range of 4.1-14.0.

Abstract: An air/oil separator is provided. The air/oil separator includes an oil manifold having a first air/oil inlet and a second air/oil inlet; a first separation chamber in communication with the first air/oil inlet; and a second separation chamber separate from the first separation chamber, the second separation chamber in communication with the second air/oil inlet.

Abstract: A turbomachine engine can include a fan assembly, a vane assembly, a core engine, a gearbox, and an overall engine efficiency rating. The fan assembly can include a plurality of fan blades. The vane assembly can include a plurality of vanes, and the vanes can, in some instances, be disposed aft of the fan blades. The core engine can include a low-pressure turbine. The gearbox includes an input and an output. The input of the gearbox is coupled to the low-pressure turbine of the core engine and comprises a first rotational speed, the output of the gearbox is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 2.0-4.0.

Abstract: Propeller control systems and methods for controlling the pitch of a plurality of propeller blades of a variable pitch propeller assembly operatively coupled with an engine is provided. In one exemplary aspect, the propeller control system includes features for combining overspeed and feathering protective functions in a protective control valve communicatively coupled with a controller. In such an event the controller controls the protective control valve to selectively allow a controlled amount of hydraulic fluid to flow to or from a pitch actuation assembly such that the pitch of the propeller blades can be adjusted based at least in part on the condition of the engine.

Abstract: A rotating blade assembly for a turbine engine having a drive shaft, the rotating blade assembly comprising a disc, at least one blade assembly, and a retainer. The disc being operably coupled to the drive shaft and including a seat having at least a portion of a first through hole. The at least one blade assembly having an upper platform, a lower platform, a dovetail extending from the lower platform, and a blade extending between the upper platform and the lower platform. The retainer assembly securing the disc to the at least one blade assembly and comprising a hollow tubular element, a pin, and a fastener.

Abstract: A turbomachine engine includes a fan section having a fan shaft, and a core engine having one or more compressor sections, one or more turbine sections that includes a power turbine, and a combustion chamber in flow communication with the compressor sections and turbine sections. The turbomachine engine includes a low-speed shaft coupled to the power turbine and having a midshaft that extends from a forward bearing to an aft bearing. The low-speed shaft is characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-speed shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine includes a gearbox assembly that couples the fan shaft to the low-speed shaft and characterized by a gearbox assembly mode less than 95% of a midshaft mode of the midshaft or greater than 105% of the midshaft mode.

Abstract: A gearbox assembly includes a plurality of planet gears and a planet carrier including a plurality of planet gear pockets. Each planet gear is mounted in a respective planet gear pocket about a pin. The planet carrier is characterized by a planet carrier envelope in a range of 2.7 and 4.50. The planet carrier envelope is equal to a product of a planet gear pocket aspect ratio multiplied by a planet carrier radial ratio. The planet gear pocket aspect ratio is Dp/L. The planet carrier radial ratio is D/aw. Dp is a planet gear pocket diameter of the plurality of planet gear pockets, L is a planet gear pocket length of the plurality of planet gear pockets, D is a planet carrier diameter of the planet carrier, and aw is a pin center distance from a longitudinal axis of the planet carrier to an axis of the pin.