A vane arm has: a first end and a second end; a shank extending between the first end and the second end and having a first face and a second face; a first hole in the shank proximate the first end; a second hole in the shank proximate the second end; and a transverse groove in the first face intersecting the second hole. The groove has: a base; a first open end and a second open end; a first side face and a second side face. An asymmetry between the groove first side face and second side face provides for uniquely angularly registering the vane arm to a shaft.
An assembly for an aircraft propulsion system includes a first open propulsor rotor, a turbine engine, a pylon structure, a support structure, a second open propulsor rotor and a drive system. The turbine engine includes a flowpath, a compressor section, a combustor section, a turbine section and a rotating structure. The rotating structure includes a turbine rotor in the turbine section. The rotating structure is configured to drive rotation of the first open propulsor rotor. The pylon structure is mounted to the turbine engine. The support structure is arranged with the pylon structure independent of the turbine engine. The second open propulsor rotor is rotationally supported by the support structure. The drive system is configured to drive rotation of the second open propulsor rotor.
B64C 11/48 - Units of two or more coaxial propellers
B64D 27/10 - Aircraft characterised by the type or position of power plants of gas-turbine type
B64D 35/06 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions characterised by the transmission driving a plurality of propellers or rotors the propellers or rotors being counter-rotating
B64D 27/00 - Arrangement or mounting of power plants in aircraftAircraft characterised by the type or position of power plants
B64D 27/40 - Arrangements for mounting power plants in aircraft
A vane arm has: a first end and a second end. A shank extends between the first end and the second end and has a first face and a second face. A first hole is in the shank proximate the first end and a second hole is in the shank proximate the second end. A tab projects distally from the shank proximate the second end and has: a bend; and a distal portion distally of the bend and projecting back toward the first end and less than fully overlapping the second hole, if at all.
A vane arm has first and second holes in its shank proximate first and second ends. First and second tabs each project laterally opposite each other from the shank proximate the second end and have: a bend; and a distal portion distally of the bend. The second tab distal portion partially overlaps the first tab distal portion. A third hole is in the second tab distal portion overlapping a distal end edge of the first tab and having a proximal edge and a distal edge. At least along a region across an axis of the second hole, the third hole distal edge and proximal edge diverge from each other in a direction toward the first end and/or the first tab distal end edge and the third hole proximal edge diverge from each other in a direction toward the first end.
A vane arm has: a first end and a second end; a shank extending between the first end and the second end and having a first face and a second face; a first hole in the shank proximate the first end; a second hole in the shank proximate the second end; and a transverse groove in the first face intersecting the second hole. The transverse groove has: a base; a first open end and a second open end; a first side face and a second side face. The transverse groove first side face and second side face converge toward each other from the first open end to the second open end.
An aircraft propulsion system includes a core engine, a propulsive fan, and a bottoming cycle that includes a working fluid mixture within a closed circuit that is heated and expanded through a bottom turbine to generate shaft power. The working fluid mixture includes a proportion of at least two fluids having different properties that is adjustable to change a capability of the working fluid mixture at different bottoming cycle operating points. The bottom cycle includes a reactor assembly configured to generate a reaction with one of the at least two fluids for adjusting the proportion of the at least two fluids of the working fluid mixture.
F02C 6/18 - Plural gas-turbine plantsCombinations of gas-turbine plants with other apparatusAdaptations of gas-turbine plants for special use using the waste heat of gas-turbine plants outside the plants themselves, e.g. gas-turbine power heat plants
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
A method of manufacture is provided which includes steps of: providing a component body of an engine component for a turbine engine; and applying a coating onto the component body using a cold spray process. The applying of the coating includes impinging coating feedstock material against the component body to form the coating. The feedstock material includes a mixture of metal powder and grit particles. The grit particles include solid metal oxide particles and/or solid carbide particles.
A process of fabricating at least one channel in a ceramic-containing material comprising providing at least one ceramic-containing material having a first major surface, a second major surface opposite the first major surface, at least one first side surface, and at least one second side surface adjacent the at least one first side surface; removing at least a portion of material from at least one of the first major surface or the second major surface of the at least one ceramic-containing material and form one or more channels comprising at least one heat augmentation feature; providing at least one metal cover plate; disposing the at least one metal cover plate adjacent the at least one ceramic-containing material and covering the one or more channels and the at least one heat augmentation feature; and securing together the at least one ceramic-containing material and the at least one metal cover plate.
C04B 37/02 - Joining burned ceramic articles with other burned ceramic articles or other articles by heating with metallic articles
B23K 26/364 - Laser etching for making a groove or trench, e.g. for scribing a break initiation groove
B32B 9/00 - Layered products essentially comprising a particular substance not covered by groups
B32B 9/04 - Layered products essentially comprising a particular substance not covered by groups comprising such substance as the main or only constituent of a layer, next to another layer of a specific substance
B32B 15/04 - Layered products essentially comprising metal comprising metal as the main or only constituent of a layer, next to another layer of a specific substance
B32B 37/18 - Methods or apparatus for laminating, e.g. by curing or by ultrasonic bonding characterised by the properties of the layers with all layers existing as coherent layers before laminating involving the assembly of discrete sheets or panels only
B32B 38/00 - Ancillary operations in connection with laminating processes
B32B 38/10 - Removing layers, or parts of layers, mechanically or chemically
C04B 41/00 - After-treatment of mortars, concrete, artificial stone or ceramicsTreatment of natural stone
9.
CERAMIC MATRIX COMPOSITE BLADE OUTER AIR SEAL FEATHER SEAL CONFIGURATION AND METHOD
A feather seal arrangement for a ceramic matrix composite (CMC) blade outer air seal (BOAS) of a turbine engine includes a CMC BOAS segment having a specific geometry and a first feather seal. The CMC BOAS segment includes a shoe having a flowpath side, an opposing non-flowpath side, and first and second matefaces extending axially between the flowpath side and the non-flowpath side from an upstream leading edge to a downstream trailing edge. One or more support flanges extend substantially transverse from the non-flowpath side. A geometry of the CMC BOAS segment allows a substantially flat first feather seal surface to be disposed on the non-flowpath side adjacent and substantially transverse to the first mateface. A first feather seal is disposed atop the non-flowpath side on the first feather seal surface and extends to an adjacent feather seal surface of an adjacent CMC BOAS segment.
A process for restoring a ceramic containing material comprising the steps of cleaning at least a portion of a location on a surface of a ceramic containing material to form a cleaned location, the ceramic containing material comprising at least one of the following: a coated preform, a partially densified ceramic matrix composite, a ceramic matrix composite component, and combinations thereof; masking the surface, except for the cleaned location, of the ceramic matrix composite with a maskant to form an aperture exposing the cleaned location; shaping the aperture to form a shaped, exposed cleaned location; depositing a material onto the shaped exposed location to form a location having a deposited material; densifying the deposited material to form a densified location; and, reshaping the densified location to form a reshaped densified location.
C04B 41/53 - After-treatment of mortars, concrete, artificial stone or ceramicsTreatment of natural stone involving the removal of part of the materials of the treated article
11.
PROPULSION SYSTEM WITH ADJUSTABLE CORE INLET GUIDE VANE
An aircraft assembly includes an open propulsor rotor and a turbine engine configured to drive rotation of the open propulsor rotor about an axis. The turbine engine includes a flowpath, a compressor section, a combustor section, a turbine section, an inlet vane structure and an inlet vane actuation system. The flowpath extends longitudinally through the compressor section, the combustor section and the turbine section from a flowpath inlet into the turbine engine to a flowpath exhaust from the turbine engine. The inlet vane structure is located at the flowpath inlet. The inlet vane structure includes a plurality of inlet guide vanes arranged in an annular array. Each of the inlet guide vanes extend across the flowpath. The inlet guide vanes include a first inlet guide vane. The inlet vane actuation system is configured to change at least a parameter of the first inlet guide vane.
F01D 9/04 - NozzlesNozzle boxesStator bladesGuide conduits forming ring or sector
F02K 3/02 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
12.
EXTERNALLY SOURCED COOLING FOR PROBE AND LEAD AND METHOD OF DELIVERY
A probe assembly includes a cup including at least one cooling outlet, and a cap including an inlet, the cap affixed to the cup. The probe assembly also includes hypo tubing affixed to the inlet, and a probe disposed within an interior volume of the cup and the cap, the probe including a lead at least partially disposed within the hypo tubing.
F01D 21/00 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for
G01D 3/028 - Measuring arrangements with provision for the special purposes referred to in the subgroups of this group mitigating undesired influences, e.g. temperature, pressure
A system for monitoring operation of a gas turbine engine comprises a plurality of expert networks each configured to generate a separate gas path parameter responsive to a separate actuator position. A router network configured to generate a weighting vector responsive to at least one ambient condition parameter. The weighting vector includes a plurality of weighting values each associated with one of the plurality of expert networks. Summing circuitry configured to apply the plurality of weighting values to each of the associated separate gas path parameters from the plurality of expert networks and sum each of the plurality of weighted separate gas path parameters from the plurality of expert networks to a weighted sum value.
F02C 9/18 - Control of working fluid flow by bleeding, by-passing or acting on variable working fluid interconnections between turbines or compressors or their stages
14.
DUAL PASSAGE HEAT EXCHANGER FOR REDUCING COLD WALL STRESS AND METHOD THEREOF
A ceramic matrix composite (CMC) component of a gas turbine engine forming a portion of a substantially ring-shaped gas turbine engine stage includes an inner surface configured for exposure to a hot gas path of the gas turbine engine and an outer surface on a non-gas path side. The inner surface can be cooled with cold air fed to an inner heat exchange passage disposed adjacent the inner surface of the CMC component at a first radial distance relative to an axis of the gas turbine engine. Thermal stress at the cold outer surface can be reduced by feeding hot air to an outer heat exchange passage disposed adjacent the outer surface at a second radial distance greater than the first radial distance.
A method for forming a cooling hole diffuser in a component includes forming a cooling passage precursor through the component, the cooling passage precursor having a meter center; applying a coating to the component including portions of the cooling passage precursor to create an at least partially coated cooling passage having a meter remnant center that is offset relative to the meter center; and removing material from the at least partially coated cooling passage to form a finished cooling passage, wherein the step of removing material is guided optically based upon the meter remnant center and a correction from the meter remnant center to the meter center.
A repaired ceramic vane of a gas turbine engine includes at least one platform and an airfoil section that extends from the at least one platform. The platform includes at least one mounting surface for supporting the ceramic vane and the airfoil section includes a damaged region. There is at least one over-wrap fiber ply wrapped around the airfoil section. The over-wrap fiber ply covers the damaged region. The over-wrap fiber ply is densified with a ceramic matrix material. There is at least one build-up layer on the mounting surface. The build-up layer has an adjusted thickness and contour that provides a desired stagger angle to the ceramic vane.
F02C 3/04 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
F01D 5/00 - BladesBlade-carrying membersHeating, heat-insulating, cooling, or antivibration means on the blades or the members
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
F02C 7/00 - Features, component parts, details or accessories, not provided for in, or of interest apart from, groups Air intakes for jet-propulsion plants
17.
REPAIR OF CERAMIC MATRIX COMPOSITE GAS TURBINE ENGINE ARTICLE WITH MONOLITHIC CERAMIC REPAIR ELEMENT
A gas turbine engine article includes a ceramic matrix composite (CMC) wall that has an exterior core gaspath surface and a blind hole. A monolithic ceramic repair element is attached with the CMC wall. The repair element includes a repair element surface that is substantially flush with the at least one exterior core gaspath surface. The ceramic repair element includes a through-hole aligned with the blind hole. A fastener extends through the through-hole and into the blind hole and secures the repair element with the CMC wall.
A method includes obtaining, from a first plurality of sources, production data relating to a plurality of stages of production of a part of a machine. The production data includes at least one of test data, simulation data, validation data, verification data, and reliability data for the part. The method also includes obtaining, from a plurality of second plurality of sources that differs from the first plurality of sources, manufacturing metadata that is separate from the production data. The manufacturing metadata data including data that describes machinery used to manufacture the part. The method also includes storing the production data and the manufacturing metadata in one or more blockchains. A system is also disclosed.
G05B 19/418 - Total factory control, i.e. centrally controlling a plurality of machines, e.g. direct or distributed numerical control [DNC], flexible manufacturing systems [FMS], integrated manufacturing systems [IMS] or computer integrated manufacturing [CIM]
A method of manufacture is provided that includes steps of: providing an airfoil body of a fan blade for a turbine engine; and cold spraying a coating onto the airfoil body at a body tip. The airfoil body includes a body leading edge, a body trailing edge and the body tip. The airfoil body extends longitudinally between the body leading edge and the body trailing edge. The airfoil body projects spanwise out to the body tip.
An aircraft propulsion system includes an open propulsor rotor, a turbine engine and an auxiliary flowpath. The turbine engine includes an engine flowpath, a compressor section, a combustor section and a turbine section. The engine flowpath extends through the compressor section, the combustor section and the turbine section from an engine flowpath inlet to an engine flowpath exhaust. The auxiliary flowpath extends longitudinally along a trajectory from an auxiliary flowpath inlet to an auxiliary flowpath outlet. The auxiliary flowpath inlet is disposed along an exterior of the propulsion system that borders an environment external to the propulsion system. The auxiliary flowpath outlet is disposed along the engine flowpath. The turbine engine is axially aft of the open propulsor rotor along an axis. The trajectory extends axially forward along the axis as the auxiliary flowpath extends longitudinally from the auxiliary flowpath inlet.
A method for machining a thin-walled part includes selecting a tool for machining a part; obtaining tool-based dynamics including tool-based frequency response; developing a preliminary tool path using the tool-based dynamics; simulating a machining process of the part with the preliminary tool path to produce an in-process model; developing a cut stock frequency response from the in-process model; determining stability lobes for machining the part with the tool, the stability lobes being determined from the tool-based frequency response and the cut stock frequency response; choosing operating parameters for a stable machining step from the stability lobes; and machining the part with the tool at the operating parameters.
G06F 30/20 - Design optimisation, verification or simulation
G05B 19/18 - Numerical control [NC], i.e. automatically operating machines, in particular machine tools, e.g. in a manufacturing environment, so as to execute positioning, movement or co-ordinated operations by means of programme data in numerical form
G05B 19/4069 - Simulating machining process on screen
B23P 15/02 - Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from one piece
G06F 119/18 - Manufacturability analysis or optimisation for manufacturability
22.
AIRCRAFT PROPULSION SYSTEM WITH ACOUSTIC TREATMENT ALONG BYPASS FLOWPATH
An apparatus is provided for an aircraft propulsion system. This apparatus includes an inner housing structure extending axially along and circumferentially about an axis. A radial outer side of the inner housing structure forms an inner peripheral boundary of a flowpath. The inner housing structure includes an inner case, an inner platform of a guide vane structure, a flowpath wall and a bulkhead. The flowpath wall is axially adjacent the inner platform along the inner peripheral boundary of the flowpath. The flowpath wall is spaced radially outboard from the inner case. The flowpath wall is configured with a flowpath wall acoustic treatment extending along the inner peripheral boundary of the flowpath. The bulkhead extends radially between and is connected to the inner case and the flowpath wall.
A static impact test fixture includes a load plate, a broach, and an under-root padding. The load plate is configured to support weight of a fan blade when a blade root of the fan blade is placed onto the load plate. The broach includes a notch, a flange portion, and a throat opening. The notch is configured to fit over the load plate and to allow the broach to slide along a first axis relative to the load plate. The flange portion extends in a first direction over the notch and is configured to overlap the load plate. The throat opening is configured for a blade neck to pass through. The under-root padding is composed of a compliant material having a specified thickness and compliance that enable the under-root padding to deform to an initial deformation caused by a clamping load between the load plate and the broach.
An apparatus for detecting proximity within a gas turbine engine comprising a proximity probe configured to monitor proximity between engine components within areas of the gas turbine engine. An articulated housing contains the proximity probe. The articulated housing enables the proximity probe to monitor the proximity between the engine components in a 360° range about the apparatus.
G01B 11/14 - Measuring arrangements characterised by the use of optical techniques for measuring distance or clearance between spaced objects or spaced apertures
G01B 7/14 - Measuring arrangements characterised by the use of electric or magnetic techniques for measuring distance or clearance between spaced objects or spaced apertures
25.
OPEN ROTOR GAS TURBINE ENGINE WITH ADAPTIVE TURBINE CLEARANCE CONTROL SYSTEM AND METHOD
A gas turbine engine disposed within a nacelle is provided that includes an open rotor propulsion system, compressor, combustion, and turbine sections, a lubrication system, and a turbine active clearance control system. The turbine active clearance control system includes first and second valves, a heat exchanger, and a nacelle air inlet device. The heat exchanger receives air from the air inlet device, accepts a lubricant flow therethrough, and permits heat transfer between the air flow and the lubricant to produce a conditioned air flow. The first valve receives conditioned air flow from the heat exchanger and may pass the conditioned air flow to the second valve. The second valve receives a second air flow from the air inlet device. The turbine active clearance control system may operate in a heating mode in which conditioned air flow, or a combined air flow is provided to the turbine case clearance control segment.
A fan exit guide vane with a load member including a leading edge and a trailing edge opposite chordwise from the leading edge; a radially inner attachment region opposite spanwise from a radially outer attachment region; a span dimension extending between the radially inner attachment region and the radially outer attachment region; a chord dimension extending between the leading edge and the trailing edge; a pressure side opposite a suction side of the fan exit guide vane; a load member cavity formed within the fan exit guide vane extending spanwise through the fan exit guide vane from the radially inner attachment region to the radially outer attachment region; and the load member extending through the load member cavity beyond each of the radially inner attachment region and the radially outer attachment region of the fan exit guide vane.
A machine (800) has a shaft (98) having: a metallic substrate with an outer diameter surface (101) having a groove (100); a seal ring (20); and a counterface (110) to the seal ring. A coating (310) is on the shaft substrate and at least partially along at least one of the groove first (302) and second (304) axial end faces and base surface (106). The coating is atop a textured surface portion of the substrate. The seal ring is accommodated in the groove and has a seal ring substrate softer than the coating.
A process for fabricating a partially densified ceramic matrix composite, comprising the steps of fabricating a ceramic fiber preform comprising a fiber, a fiber tow, or both; optionally depositing an interface coating on the fiber, the fiber tow, or both; depositing a structural support material on the fiber, the fiber tow, or both to form a partially densified ceramic matrix composite comprising pores, a porous network or both; preparing at least one slurry composition comprising particulate particles; infiltrating the slurry composition into the partially densified ceramic matrix composite and depositing the particulate particles into the pores, the porous network or both; rapidly solidifying one or more liquid phases of the slurry composition within a slurry infiltrated partially densified ceramic matrix composite; drying a solidified slurry infiltrated partially densified ceramic matrix composite.
A gas turbine engine includes a seal support that has first and second spaced-apart support flanges and a ceramic matrix composite (CMC) seal that is affixable to the seal support. The support flanges have at least one set of support flange pin holes. The CMC seal includes a radially inwardly-facing side and a radially outwardly-facing side. the radially outwardly-facing side has first and second spaced-apart seal flanges that include two independent sets of support flange pin holes. The CMC seal is affixable to the seal support by alignment of either one of the two independent sets of support flange pin holes with the support flange pin holes and receipt of support pins there through.
A combustor system for an aerial vehicle includes an outer liner with a first end and a second end, an inner combustor case within the outer liner, defining an outer periphery of a central combustor chamber, and having a lumen extending into the central combustor chamber. The system includes an air flow path extending between the outer liner and the inner combustor case, an integrated air scoop which directs air from the air flow path into the central combustor chamber, a fuel manifold with a fuel orifice within the integrated air scoop, and a lattice splash plate with a first face, second face, first end, and a second end. The lattice splash plate extends through the first lumen into the central combustor chamber, and includes lattice holes which are graded into the lattice splash plate from the first end to the second end of the lattice splash plate.
A combustor liner for a gas turbine engine, the combustor liner defining a combustion chamber, includes a curved body and a cooling ring. The curved body includes a closed forward end, an open aft end opposite the forward end, an inner wall, and an outer wall radially outward from the inner wall, the outer wall defining an outer liner wall of the combustor liner. The cooling ring includes a fin that has a space filling curve that is modified such that an outer wall of the cooling ring matches a shape of the inner wall of the curved body. An inner wall of the cooling ring defines an inner liner wall of the combustor liner.
An apparatus is provided for an aircraft that includes a turbine engine. The turbine engine includes a compressor section, a combustor section, a turbine section, a support structure, a first rotor system and a second rotor system. The support structure extends axially along an axis. Each rotor system may include an electric machine and a bladed rotor. The electric machine includes an electric machine stator and an electric machine rotor. The electric machine stator is mounted to the support structure. The electric machine rotor circumscribes the electric machine stator and is operatively coupled to the bladed rotor. The electric machine is configured to generate an electromagnetic field with the electric machine stator and the electric machine rotor. The bladed rotor is configured to rotate about the axis.
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
F02K 3/02 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
A gas turbine engine is provided that includes a fan section, a fan bypass air path, a compressor section, a combustion section, a turbine section, an annular compressor inlet, and an inlet panel structure actuation system. The annular compressor inlet is defined by an inner radial surface and a plurality of inlet panel structures. Each inlet panel structure includes an inlet side segment and a bypass side segment. The inlet side segments collectively define an outer radial surface of the compressor inlet. The inlet side segment and the bypass side segment of each inlet panel structure intersect at an inlet panel leading edge. The inlet panel structures are disposable in a radial inner most configuration and in a radial outer most configuration. The inlet panel structure actuation system is configured to selectively actuate the inlet panel structures between the radial inner most configuration and the radial outer most configuration.
A gas turbine engine disposed within a nacelle is provided that includes an open rotor propulsion system, a compressor section, an airflow inlet, and an inlet panel structure actuation system. The airflow inlet is in fluid communication with the compressor section. The airflow inlet is defined by a plurality of inlet panel structures. Each inlet panel structure includes an inlet side segment and an outer panel segment. The inlet side segments collectively define an outer radial surface of the airflow inlet. The inlet side segment and the outer panel segment of each inlet panel structure intersect at an inlet panel leading edge. The inlet panel structure actuation system is configured to selectively actuate the inlet panel leading edge of each inlet panel structure in a radial inward direction or in a radial outward direction.
An aerospace piston seal ring including a body shaped as an annular ring, the body defines an axis centered within the body; a sealing surface extending radially relative to the axis between an inner diameter and an outer diameter; and a surface feature formed along the sealing surface, the surface feature configured as an interruption in the sealing surface, wherein the surface feature is configured to contain a lubricant.
An aircraft propulsion system, a propulsor section and a core engine section configured to generate an exhaust gas flow that is expanded through a turbine to drive a shaft for driving the fan. A propulsor mount assembly is attached to the propulsor section and supports the propulsor section on an aircraft structure and isolates loads generated in the propulsor section from the core engine section.
F02C 7/20 - Mounting or supporting of plantAccommodating heat expansion or creep
B64D 27/10 - Aircraft characterised by the type or position of power plants of gas-turbine type
B64D 27/40 - Arrangements for mounting power plants in aircraft
F02C 3/06 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages
A CMC component is provided with at least one cooling cavity, defined by side walls and a cavity bottom wall, and a cover plate the covers the at least one cooling cavity opening. The cover plate has one or more cooling air inlets to permit cooling air to flow from a region above the cover plate into the cooling cavity. The cover plate further has one or more walls extending downward from an inner surface of the cover plate to the cavity bottom wall thereby dividing the cooling cavity into a plurality of cooling channels or cooling subcavities, and/or creating a circuitous cooling pathway within the cooling cavity.
The disclosure describes methods and devices for directing/controlling cooling air flow for cooling CMC components. A contoured cover plate is positioned within a cooling cavity in the CMC component, the cooling cavity being defined by side walls and a cavity bottom wall. The cover plate comprises an edge region around the perimeter of the cover plate and a central recessed region which extends into the cooling cavity, the central recessed region is defined by side walls and a bottom wall wherein each side wall of the central recessed region is spaced from a corresponding cavity side wall and the bottom wall of the central recessed region is spaced from the cavity bottom wall. The central recessed region has a plurality of cooling air inlets to permit cooling air to flow from a region above the cover plate into the cooling cavity.
A process for fabricating a melt-infiltrated ceramic matrix composite, comprising the steps of fabricating a ceramic fiber preform comprising fibers, fiber tows, or both fibers and fiber tows; depositing a structural support material on the fibers, the fiber tows, or the fibers and fiber tows to form a partially densified ceramic matrix composite comprising pores, a porous network or both pores and a porous network; preparing a slurry composition comprising sacrificial particles; infiltrating the slurry composition into the partially densified ceramic matrix composite and depositing the sacrificial particles into the pores, the porous network or both the pores and porous network; freeze-drying a slurry infiltrated partially densified ceramic matrix composite; sublimating a frozen slurry infiltrated partially densified ceramic matrix composite; and, melt-infiltrating a sublimated partially densified ceramic matrix composite to form a melt-infiltrated ceramic matrix composite.
The disclosure describes methods and devices for directing/controlling cooling air flow for cooling CMC components. A cover plate is positioned within a cooling cavity in the CMC component wherein the cover plate is angled to vary the depth of the cooling cavity between the cover plate and a bottom wall of the cavity. The angled cooling plate can be used to vary the flux of cooling air within the cooling cavity and thereby achieve a desired cooling performance within the cooling cavity.
A method is disclosed for a design geometry of a ceramic matrix composite (CMC) vane arc segment having a platform and an airfoil section extending from the platform. The platform defines an axially trailing face, an axially leading face circumferentially offset from the axially trailing face, and first and second circumferential faces, and a modal excitation response of the platform under cruise conditions is determined. If the modal excitation response exceeds a target modal excitation response, the design geometry of the platform is adjusted to be detuned, by a predetermined margin, from an external engine operation excitation frequency of turbine blades rotating past the vane arc segment by reducing an overhang distance of the platform such that the modal excitation response at the cruise condition is equal to or lower than the target modal excitation response.
A gas turbine engine case includes: a case wall having an inner surface and an outer surface; a fuel injector protruding inward from the case wall and having an outlet; and a splash plate having a first face and a second face. The fuel injector outlet faces the second face. A support connects the splash plate to the case wall and the fuel injector. The support directs air to a gap between the splash plate and the injector.
An aircraft propulsion system includes a gas turbine engine and a power transfer assembly. The gas turbine engine includes a high-pressure (HP) spool and a low-pressure (LP) spool. The power transfer assembly includes a first motor-generator, a second motor-generator, a motor control assembly, and a common assembly housing. The first motor-generator is operably coupled with the HP spool. The second motor-generator is operably coupled with the LP spool. The first motor-generator is electrically connected to the second motor-generator. The motor control assembly is electrically connected to the first motor-generator and the second motor-generator. The motor control assembly is configured to operate the first motor-generator to apply a first rotational force to the HP spool or to operate the second motor-generator to apply a second rotational force to the LP spool. The common assembly housing houses the first motor-generator and the second motor-generator.
An aircraft propulsion system includes a gas turbine engine, a power transfer assembly, and a controller. The gas turbine engine includes a first spool and a second spool. The power transfer assembly includes a first motor-generator, a second motor-generator, and a motor control assembly. The first motor-generator is operably coupled with the first spool. The second motor-generator is operably coupled with the second spool. The motor control assembly is electrically connected to the first motor-generator and the second motor-generator. The motor control assembly is configured to control an operating voltage of the power transfer assembly. The controller is connected in signal communication with the motor control assembly. The is controller is configured to identify a target operating voltage of the power transfer assembly selected in response to an aircraft altitude input to the controller and control the motor control assembly to control the operating voltage at the target operating voltage.
A gas turbine engine is provided that includes compressor, combustor, and turbine sections, an engine shaft, an accessory gearbox, and first and second starters. The engine shaft is in communication with the compressor section and the turbine section. The accessory gearbox has first and second accessory mounting pads, and is in communication with the engine shaft. The first starter is mounted on the first accessory mounting pad, and the second starter is mounted on the second accessory mounting pad. The first starter is configured to selectively provide rotational drive to the accessory gearbox and the accessory gearbox in turn provides rotational drive to the engine shaft. The second starter is configured to selectively provide rotational drive to the accessory gearbox and the accessory gearbox in turn provides rotational drive to the engine shaft.
A method for detecting and removing at least one atmospheric pollutant from a surface of an aircraft engine component includes detecting the at least one atmospheric pollutant on the surface of the aircraft engine component using a test probe and treating the at least one aircraft engine component to generate a layer of at least one of a soot or other carbonaceous substance on the at least one aircraft engine component to at least partially remove the at least one atmospheric pollutant from the surface of the aircraft engine component.
F02C 7/30 - Preventing corrosion in gas-swept spaces
B08B 3/08 - Cleaning involving contact with liquid the liquid having chemical or dissolving effect
B08B 3/10 - Cleaning involving contact with liquid with additional treatment of the liquid or of the object being cleaned, e.g. by heat, by electricity or by vibration
G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
G01N 21/33 - Investigating relative effect of material at wavelengths characteristic of specific elements or molecules, e.g. atomic absorption spectrometry using ultraviolet light
47.
REACTION CONTROLLED MELT-INFILTRATED FABRICATION OF CERAMIC MATRIX COMPOSITES
A process for fabricating a melt-infiltrated ceramic matrix composite comprises fabricating a ceramic fiber preform comprising at least one fiber, fiber tow, or both; depositing at least one structural support material coating on the fiber, fiber tow, or both to form a partially densified ceramic matrix composite comprising at least one pore, porous network or both; depositing at least one fiber protection material coating on the fiber, fiber tow, or both; infiltrating a slurry composition comprising at least one sacrificial particle into the pore, porous network, or both; depositing at least one reaction control material coating on the fiber, the fiber tow, or the fiber, fiber tow and sacrificial particle of the partially densified ceramic matrix composite; melt-infiltrating at least one molten infiltrant into the partially densified ceramic matrix composite to form a melt-infiltrated ceramic matrix composite.
A multi-mode blade tip timing sensor includes a housing, a sensing coil provided within the housing and a powered exciter coil. The multi-mode blade tip timing sensor can operate in a first, sensing mode, and a second, de-icing mode. When operating in the first, sensing mode, a first current is provided to the powered exciter coil to create a magnetic field such that a movement of one or more moving metallic fan blades through the magnetic field creates a signal in the sensing coil. When operating in the second, de-icing mode, a second current is provided to the powered exciter coil to induce an eddy current in a tip of the one or more moving metallic turbine fan blades, wherein the eddy current raises a temperature of the one or more moving metallic turbine fan blades above an icing temperature.
B64D 15/12 - De-icing or preventing icing on exterior surfaces of aircraft by electric heating
B64D 15/22 - Automatic initiation by icing detector
G01V 3/10 - Electric or magnetic prospecting or detectingMeasuring magnetic field characteristics of the earth, e.g. declination or deviation operating with magnetic or electric fields produced or modified by objects or geological structures or by detecting devices using induction coils
A method of determining a fan liner blade tip clearance includes providing a measurement instrument. The measurement instrument includes a set screw, and a base including threaded through hole configured to accept the set screw. The method also includes placing a bottom surface of the base against the fan liner wherein the through hole aligns with a wear hole of the fan liner, and after the placement, turning the set screw within the through hole until a bottom surface of the set screw contacts a lower surface of the wear hole. The method also includes measuring an exposed portion of the set screw extending below the bottom surface of the base after contact with the lower surface of the wear hole is made, and determining a blade tip clearance with the fan liner based on a value of the measurement.
G01B 5/14 - Measuring arrangements characterised by the use of mechanical techniques for measuring distance or clearance between spaced objects or spaced apertures
A nose cone for debris deflection comprising a deflector formed in the nose cone, and an actuator in operative communication with the deflector, the actuator configured to engage the deflector configured to create a deflection air stream; the deflection air stream being configured to manipulate a fan inlet air flow such that any debris entrained in the fan inlet air flow has a trajectory line directed away from a core flow of a gas turbine engine and toward a bypass flow of the gas turbine engine.
F02C 7/05 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
B64D 33/02 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
F02C 7/042 - Air intakes for gas-turbine plants or jet-propulsion plants having variable geometry
F02C 7/052 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles with dust-separation devices
F04D 29/32 - Rotors specially adapted for elastic fluids for axial-flow pumps
A combustor assembly for a gas turbine engine includes a combustor liner disposed about an axis and defining a combustion chamber and an outer case disposed about the combustor liner and defining an air flow duct therebetween. The combustor liner includes an outer wall having a plurality of mounting bosses and a plurality of first holes open to the combustion chamber. The outer case includes a plurality of inlet scoops extending radially inward of the outer case and radially aligned with the plurality of first holes to guide an air flow into first holes. A plurality of pins is fixed to the outer case and slidingly received in openings in the plurality of mounting bosses.
A test rig with a multi-specimen fixture system including an actuator including a central shaft; a moving component in operative communication with the central shaft; a stationary component located proximate the moving component arrayed around the central shaft; a first connector coupled to the stationary component, the first connector configured to secure a test specimen; a second connector coupled to the moving component, the second connector configured to secure the test specimen; a load cell in operative communication with at least one of the first connector or the second connector; and at least one compliant element in operative communication with the moving component, wherein the at least one compliant element is configured to produce an independent load to the test specimen.
G01N 3/34 - Investigating strength properties of solid materials by application of mechanical stress by applying repeated or pulsating forces generated by mechanical means, e.g. hammer blows
G01N 3/06 - Special adaptations of indicating or recording means
53.
Aircraft Propulsion System With Internal Compartment Ventilation
An assembly is provided for an aircraft. This assembly includes an engine core, an engine case, a nacelle wall and an air circuit. The engine core includes a compressor section, a combustor section and a turbine section. The engine case houses the engine core. The nacelle wall is radially outboard of and covers the engine case. A compartment is formed by and extends radially between the engine case and the nacelle wall. The air circuit includes a manifold disposed within the compartment. The air circuit is configured to direct air from an air source into the manifold. The manifold includes a plurality of outlets. The manifold is configured to direct the air into the compartment through the outlets.
A system for traceability of manufacturing data may include one or more processors coupled to memory. The one or more processors may be collectively operable to execute a mapping environment. The mapping environment may be operable to access a production set of manufacturing instructions associated with a component design. The mapping environment may be operable to access real manufacturing data associated with execution of the production set of manufacturing instructions. The mapping environment may be operable to generate an evaluation set of manufacturing instructions associated with respective unique identifiers. The unique identifiers may be assigned to respective geometric features of the component design. The mapping environment may be operable to generate a mapped set of manufacturing instructions including the unique identifiers assigned to respective portions of the production set of manufacturing instructions based on the evaluation set of manufacturing instructions.
G05B 19/4097 - Numerical control [NC], i.e. automatically operating machines, in particular machine tools, e.g. in a manufacturing environment, so as to execute positioning, movement or co-ordinated operations by means of programme data in numerical form characterised by using design data to control NC machines, e.g. CAD/CAM
G05B 19/4069 - Simulating machining process on screen
G06F 30/17 - Mechanical parametric or variational design
Environmental barrier coatings (EBCs) are disclosed having a topcoat layer having a primary hafnon phase and a dispersed hafnia phase. The hafnia can be homogeneously dispersed or the layer can exhibit a concentration gradient in which the hafnia concentration increases continuously or stepwise. The hafnon primary phase provides a desirable CTE match with the underlying CMC substrate, while the dispersed hafnia provides desirable corrosion resistance.
A brush seal of a turbine engine may provide a cooling flow to a ceramic matrix composite (CMC) component that has a hot side exposed to a hot gas path and an opposing cold side. The brush seal includes an inner diameter (ID) backing plate providing a radial face seal across a cavity in the CMC component. The brush seal also includes an outer diameter (OD) backing plate, a plurality of brush seal bristles disposed between the ID and OD backing plates, and a hole extending through the brush seal to provide the cooling flow to the cavity of the CMC component. A cross-flow channel is disposed on an inner surface of the ID backing plate such that the cross-flow channel extends from an edge of the ID backing plate to the hole to pass leakage air to mix with the cooling flow and reduce an impingement cooling effect thereof.
A gas turbine engine component, comprising a first surface having at least one first aperture, and exposed to a first plenum comprising a mixed fluid flow and a first pressure; at least one second surface having at least one second aperture, and exposed to a gas path fluid flow and at least one second pressure; at least one channel, at least one internal plenum or both at least one channel and at least one internal plenum disposed between and in fluid communication with each the at least one first aperture and the at least second aperture.
A method of operating an aircraft may include determining, for an engine of an aircraft, a high delta T5 (MDT5) between a first channel and a second channel over a flight of the aircraft. The method may also include updating a multilevel perception neural network with the MDT5, and determining, based on the neural network, a fuel coking factor for the engine.
The disclosure describes methods and devices for directing cooling air to cool CMC components. A baffle member or splash plate is used to direct the flow of cooling air towards an outer radial surface of a CMC component into a one or more cooling cavities provided within the interior of the CMC component. The splash plate provides an efficient distribution of cooling air into the one or more cooling cavities and aids in the cooling of the component to reduce the formation of thermal stresses induced by the thermal differential between the outer radial surface (cold side) and inner radial surface (hot side) of the component.
The disclosure describes methods and devices for directing/controlling cooling air flow for cooling CMC components. A cover plate is positioned within a cooling cavity in the CMC component wherein the cover plate is angled to vary the depth of the cooling cavity between the cover plate and a bottom wall of the cavity. The angled cooling plate can be used to vary the flux of cooling air within the cooling cavity and thereby achieve a desired cooling performance within the cooling cavity.
An artificial-intelligence-assisted (AI-assisted) certification system includes an argumentation processor and an assurance case processor. The argumentation processor is configured to generate an argumentation pattern. The assurance case processor is configured to obtain the argumentation pattern from the argumentation processor, to automatically generate an assurance case based on one or more argumentation patterns, to determine evidence indicative of premises in the argumentation pattern, and to automatically assess the assurance case based on the evidence.
G06F 21/57 - Certifying or maintaining trusted computer platforms, e.g. secure boots or power-downs, version controls, system software checks, secure updates or assessing vulnerabilities
A brush seal for use with a ceramic matrix composite (CMC) component of a gas turbine engine may feed a cooling flow to the CMC component, which has a hot side configured for exposure to a hot gas path of the gas turbine engine and an opposing cold side having a cavity for feeding a cooling flow to at least one film cooling hole of the CMC component. The brush seal includes: an outer diameter (OD) backing plate; an inner diameter (ID) backing plate; a plurality of brush seal bristles sandwiched between the OD backing plate and the ID backing plate; at least one hole forming a passage extending through the OD backing plate, the brush seal bristles, and the ID backing plate; and a cover plate attached to an inner surface of the ID backing plate and configured to extend into the cavity of the CMC component.
An apparatus with a communication system includes a method of operating the communication system. The communication system includes first device at a first location, a radio frequency transceiver at the first device, a waveguide extending between the first device to a second device at a second location, and a processor. The processor is configured to transmit a first radio frequency signal through the waveguide toward the second location, receive a second radio frequency signal in response to the first radio frequency signal, determine a presence of a fault in the waveguide from the second radio frequency signal, and transmit a third radio frequency signal via the radio frequency transceiver outside of the waveguide when the presence of the fault is determined.
An apparatus is provided for an aircraft that includes an open rotor propulsion system. The open rotor propulsion system includes an open propulsor rotor, an open guide vane structure and a turbine engine. The open guide vane structure is axially next to the open propulsor rotor. The turbine engine is configured to drive rotation of the open propulsor rotor about an axis. An exterior surface of a component of the open rotor propulsion system is exposed to and borders an environment external to the open rotor propulsion system. The component is configured with an acoustic treatment extending axially and circumferentially along the exterior surface.
An apparatus is provided for an aircraft. This apparatus includes an electric machine and an air cooling circuit. The electric machine includes an electric machine rotor and an electric machine stator. The electric machine is configured to generate an electromagnetic field with the electric machine rotor and the electric machine stator. The electric machine rotor is configured to rotate about an axis. The air cooling circuit includes a cooling boot. The air cooling circuit is configured to direct air from an air source into the cooling boot. The cooling boot forms an air plenum with an exterior surface of the electric machine. The cooling boot includes a plurality of air outlets. The cooling boot is configured to direct the air through the air outlets and into the air plenum to air cool the electric machine.
An apparatus is provided for an aircraft that includes a propulsion system. The propulsion system includes a propulsor rotor and an engine core configured to power rotation of the propulsor rotor about an axis. The propulsor rotor includes a plurality of propulsor blades and an outer platform. The propulsor blades are arranged circumferentially about the axis. Each of the propulsor blades projects radially out from the outer platform to a respective propulsor blade tip. The outer platform is configured with a platform acoustic treatment. The engine core includes a flowpath, a compressor section, a combustor section and a turbine section. The flowpath extends through the compressor section, the combustor section and the turbine section.
An aircraft assembly includes first and second propulsion systems. Each propulsion system includes an open propulsor rotor and a turbine engine configured to drive rotation of the open propulsor rotor. The turbine engine includes a compressor section, a combustor section, a turbine section, a first rotating structure, a second rotating structure and a flowpath. The flowpath extends through the compressor section, the combustor section and the turbine section with the first bladed rotor disposed between the second bladed rotor and the combustor section along the flowpath. The open propulsor rotor of the first propulsion system is configured to rotate a first rotational direction. The open propulsor rotor of the second propulsion system is configured to rotate a second rotational direction that is opposite the first rotational direction.
B64D 27/20 - Aircraft characterised by the type or position of power plants of jet type within, or attached to, fuselages
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
F02K 3/02 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
68.
Symmetric Open Propulsor Rotating Patterns for Aircraft
An aircraft assembly includes first and second propulsion systems. Each of the propulsion systems includes an open propulsor rotor and a turbine engine configured to drive rotation of the open propulsor rotor. The turbine engine includes a compressor section, a combustor section, a turbine section, a first rotating structure, a second rotating structure and a flowpath. The first rotating structure includes a first bladed rotor. The second rotating structure includes a second bladed rotor and is operable to rotate independent of the first rotating structure. The open propulsor rotor of the first propulsion system is configured to rotate a first rotational direction. The open propulsor rotor of the second propulsion system is configured to rotate a second rotational direction that is opposite the first rotational direction.
B64D 27/10 - Aircraft characterised by the type or position of power plants of gas-turbine type
F02C 3/06 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages
F02K 3/072 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with counter-rotating rotors
A gas turbine engine stator stage has: a case; a plurality of vane clusters; and a plurality of vane singlets. The plurality of vane clusters each have: an inner diameter first platform; an outer diameter first shroud mounted to the case; and a plurality of first airfoils extending between the first platform and first shroud. The plurality of vane singlets each have: an inner diameter second platform; an outer diameter second shroud; and a single second airfoil extending between the second platform and second shroud.
An under-root fan blade spacer system including a fan hub including a fan hub receiver, the fan hub receiver having a hub receiver floor, the fan hub receiver having an opening configured to receive a fan blade root, the fan blade root includes a base end; a gap formed between the base end of the fan blade root and a hub receiver floor; and a spacer disposable within the gap, the spacer comprising a shape memory alloy material.
Attachment pins are used to attach a ceramic matrix composite (CMC) component to a support structure, the attachment pins passing through openings in flange structures extending from a surface of the CMC component. The attachment pins are provided with a cooling air passageway extending in an axial direction, the cooling passageway having an inlet opening in a first end of the attachment pin for introduction of cooling air, and each attachment pin having one or more cooling air outlets extending radially from the cooling air passageway to provide for discharge of cooling air in a desired direction.
A brush seal of a turbine engine may provide a cooling flow to a ceramic matrix composite (CMC) component that has a hot side configured for exposure to a hot gas path of the turbine engine and an opposing cold side. The brush seal includes an inner diameter (ID) backing plate providing a radial face seal across a cavity in the CMC component, with the cavity having an entrance disposed on the cold side to receive the cooling flow. The brush seal also has an outer diameter (OD) backing plate, a plurality of brush seal bristles disposed between the ID backing plate and OD backing plate and attached thereto in a welding zone, and a hole extending through the OD backing plate, the brush seal bristles, and the ID backing plate in the welding zone to provide the cooling flow to the cavity of the CMC component.
A special tool in intermediate form, comprising: at least one aft core module having at least one aft alignment aperture and at least one baffle; a forward core having at least one forward alignment aperture, the at least one aft core module and the forward core disposed adjacent to and aligned with one another; at least one alignment tool disposed within each at least one aft alignment aperture, and aligned with and also disposed within each at least one forward alignment aperture; and a preform material disposed about the adjacent, aligned aft core and forward core.
A ceramic matrix composite (CMC) component segment forming a portion of a substantially ring-shaped gas turbine engine stage includes a mateface extending axially between an upstream end and a downstream end of the CMC component and configured to form an inter-segment gap with an adjacent CMC component, and a plurality of film cooling holes having diffused openings at the mateface to provide film cooling of the mateface. A method of providing an inter-segment purge flow to such a CMC component includes disposing a first mateface of the CMC component segment next to a second mateface of an adjacent CMC component segment to form an inter-segment gap, and providing first film cooling of the first mateface via the plurality of film cooling holes having diffused openings at the first mateface.
A piston seal ring for sealing between an inner and an outer radial component includes a core having a shell, wherein the core includes a material selected from the group consisting of titanium alloy, titanium-based metal matrix, carbon-carbon composite, electro-graphitic carbon, ceramic matrix composite and combinations thereof, and wherein the shell includes a material selected from the group consisting of nickel alloy, cobalt alloy and combinations thereof.
A method for oxide formation on the surface of a ring includes: mounting a ring in a fixture; driving rotation of the ring about a first axis; contacting an outer diameter surface of the ring with a roller; and laser heating the outer diameter surface of the ring to form an oxide layer. The rotation causes compaction of the oxide by the roller.
A thermo-cell includes a first electrode and a second electrode defining a space therebetween with an electrolyte disposed within the space. At least one p-type membrane is positioned in the space and facing the first electrode and at least one n-type membrane is positioned in the space and facing the second electrode. The p-type membrane and the n-type membrane are physically separated and create an electric field gradient through the space for the selective transport of ions within the electrolyte. A system including a plurality of such thermo-cells arranged to be electrically coupled is also provided.
H10N 10/17 - Thermoelectric devices comprising a junction of dissimilar materials, i.e. devices exhibiting Seebeck or Peltier effects operating with only the Peltier or Seebeck effects characterised by the structure or configuration of the cell or thermocouple forming the device
H01M 10/0525 - Rocking-chair batteries, i.e. batteries with lithium insertion or intercalation in both electrodesLithium-ion batteries
H01M 10/0585 - Construction or manufacture of accumulators having only flat construction elements, i.e. flat positive electrodes, flat negative electrodes and flat separators
78.
Turbine blade for a gas turbine engine and method for forming same
A blade for a gas turbine engine includes an airfoil body. The airfoil body includes a pressure side wall, a suction side wall, and a tip end wall. The tip end wall forms a blade tip. The airfoil body forms a plurality of suction side wall passages and a main body cavity. The suction side wall includes an exterior wall segment and an interior wall segment. The suction side wall forms the plurality of suction side wall passages between the exterior wall segment and the interior wall segment. The interior wall segment and the exterior wall segment extend to the tip end wall. The plurality of suction side wall passages extend through the tip end wall to the blade tip. The interior wall segment and the pressure side wall form the main body cavity.
A start mechanism ignitor for a propulsion system combustor including a nozzle comprising a nozzle wall forming a nozzle interior; and a torch ignitor attached to the nozzle wall and fluidly coupled with the nozzle interior.
An aircraft propulsion system includes a dual exhaust path assembly that includes actuatable blocker doors moveable between a bypass configuration for directing the exhaust gas flow through an outer passage and a heat exchange configuration for directing the exhaust gas flow through a heat exchanger assembly.
F02C 9/52 - Control of fuel supply conjointly with another control of the plant with control of working fluid flow by bleeding or by-passing the working fluid
F02C 7/08 - Heating air supply before combustion, e.g. by exhaust gases
An aircraft assembly includes a turbine engine, a first electric machine, a first controller, an electric machine fluid circuit and a controller fluid circuit. The turbine engine includes a compressor section, a combustor section, a turbine section, a flowpath and a first rotating structure. The first rotating structure includes a first bladed rotor disposed in one of the compressor section or the turbine section. The first electric machine includes a first electric machine rotor. The first electric machine rotor is operatively coupled to the first rotating structure. The first controller is configured to electrically couple the first electric machine to an electrical system. The electric machine fluid circuit is configured to circulate a first liquid and service the first electric machine. The controller fluid circuit is configured to circulate a second liquid and service the first controller. The controller fluid circuit is fluidly discrete from the electric machine fluid circuit.
An aircraft apparatus includes a turbine engine, a first electric machine, a first controller, a first fluid circuit, a second electric machine, a second controller and a second fluid circuit. The turbine engine includes a first rotating structure and a second rotating structure. The first electric machine is operatively coupled to the first rotating structure. The first controller is configured to control operation of the first electric machine. The first fluid circuit is configured to circulate a first liquid and service the first electric machine and the first controller. The second electric machine is operatively coupled to the second rotating structure. The second controller is configured to control operation of the second electric machine. The second fluid circuit is configured to circulate a second liquid and service the second electric machine and the second controller. The second fluid circuit is fluidly discrete from the first fluid circuit.
B64D 31/18 - Power plant control systemsArrangement of power plant control systems in aircraft for electric power plants for hybrid-electric power plants
F02C 6/00 - Plural gas-turbine plantsCombinations of gas-turbine plants with other apparatusAdaptations of gas-turbine plants for special use
A heat exchanger for providing thermal energy transfer between a first flow along a first flowpath and a second flow along a second flowpath has a plate bank having a plurality of plates, each plate having: a first face and a second face opposite the first face; a leading edge along the second flowpath and a trailing edge along the second flowpath; a proximal edge having at least one inlet port along the first flowpath and at least one outlet port along the first flowpath; and at least one passageway along the first flowpath. An inlet manifold has at least one inlet port and at least one outlet port. An outlet manifold has at least one outlet port and at least one inlet port. The first flowpath passes from the at least one inlet port of the inlet manifold, through the at least one passageway of each of the plurality of plates, and through the at least one outlet port of the outlet manifold and means linking distal portions of the plates.
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
F02C 7/141 - Cooling of plants of fluids in the plant of working fluid
F28D 1/03 - Heat-exchange apparatus having stationary conduit assemblies for one heat-exchange medium only, the media being in contact with different sides of the conduit wall, in which the other heat-exchange medium is a large body of fluid, e.g. domestic or motor car radiators with the heat-exchange conduits immersed in the body of fluid with plate-like or laminated conduits
F28F 9/013 - Auxiliary supports for elements for tubes or tube-assemblies
A fan drive gear system for a turbine engine includes a sun gear configured to be driven by an engine shaft that is rotatable about an axis, a plurality of intermediate gears coupled to a first gear portion of the sun gear, a ring gear that is coupled to the plurality of intermediate gears, and a carrier supporting rotation of the plurality of intermediate gears. A coupling shaft is engaged to drive the sun gear and includes a gear portion. An accessory component is coupled to and driven by the gear portion of the coupling shaft.
An aircraft propulsion system includes at least two gas turbine engines and a bottoming cycle system where a working fluid is circulated within a closed circuit that includes a bottoming compressor section and a bottoming turbine section. Each of the at least two gas turbine engines include a primary heat exchanger for communicating thermal energy into the working fluid of the bottoming cycle.
A method of repairing a coating on an article according to an exemplary embodiment of this disclosure, among other possible things includes defining a work area surrounding a discontinuity in the coating, the coating including undisturbed bond coat and undisturbed top coat; removing coating material within the work area; applying a slurry containing bond coat constituents in a carrier fluid to the work area; curing the slurry to form a repaired bond coat; applying a slurry containing top coat constituents in a carrier fluid to the work area; and curing the slurry to form a repaired top coat. A method of repairing a coating on an article and an article are also disclosed.
B05D 5/00 - Processes for applying liquids or other fluent materials to surfaces to obtain special surface effects, finishes or structures
B05D 3/12 - Pretreatment of surfaces to which liquids or other fluent materials are to be appliedAfter-treatment of applied coatings, e.g. intermediate treating of an applied coating preparatory to subsequent applications of liquids or other fluent materials by mechanical means
C04B 35/80 - Fibres, filaments, whiskers, platelets, or the like
An apparatus is provided for an aircraft. This apparatus includes a turbine engine, a first electric machine, a first controller, a second electric machine, a second controller and a first fluid circuit. The first electric machine is operatively coupled to the turbine engine. The first controller is configured to control operation of the first electric machine. The second electric machine is operatively coupled to the turbine engine. The second controller is configured to control operation of the second electric machine. The first fluid circuit is configured to circulate a first liquid to cool and/or lubricate the first electric machine, the first controller, the second electric machine and the second controller.
An apparatus is provided for an aircraft that includes an open rotor propulsion system. The open rotor propulsion system includes a turbine engine, a first electric machine, a second electric machine, a first controller, a second controller and a first fluid circuit. The first controller is configured to control operation of the first electric machine. The second controller is configured to control operation of the second electric machine. The first fluid circuit is configured to circulate a first liquid to cool and/or lubricate the first electric machine, the first controller, the second electric machine and the second controller.
An aircraft assembly includes a propulsor rotor, a turbine engine core, an inner case, an outer case, a first gearbox, a second gearbox, a first electric machine and a second electric machine. The turbine engine core is configured to drive rotation of the propulsor rotor about an axis. The turbine engine core includes a first rotating structure and a second rotating structure. The first rotating structure includes a first bladed rotor. The second rotating structure includes a second bladed rotor. The inner case houses the turbine engine core. The outer case houses the propulsor rotor. The first gearbox is mounted with the outer case. The second gearbox is mounted with the outer case. The first electric machine is operatively coupled to the first rotating structure through the first gearbox. The second electric machine is operatively coupled to the second rotating structure through the second gearbox.
B64D 27/10 - Aircraft characterised by the type or position of power plants of gas-turbine type
B64D 35/08 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions characterised by the transmission being driven by a plurality of power plants
H02K 7/116 - Structural association with clutches, brakes, gears, pulleys or mechanical starters with gears
H02K 7/18 - Structural association of electric generators with mechanical driving motors, e.g.with turbines
90.
ELECTRIC MACHINE AND ACCESSORY GEARBOX ARRANGEMENT FOR AIRCRAFT PROPULSION SYSTEM
An aircraft assembly is provided that includes a propulsor rotor, a turbine engine core, an inner case, a first gearbox, a second gearbox, a first electric machine and a second electric machine. The turbine engine core includes a first rotating structure and a second rotating structure. The first gearbox is mounted with the inner case. The second gearbox is located remote from the inner case. The first electric machine is operatively coupled to the first rotating structure through the first gearbox. The second electric machine is operatively coupled to the second rotating structure through the second gearbox. The first electric machine may be configurable as a first electric motor and/or a first electric generator. In addition or alternatively, the second electric machine may be configurable as a second electric motor and/or a second electric generator.
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
91.
Powder Bed Fusion Build Plates Restoration with Friction Surfacing Additive Manufacturing (FSAM)
A build plate for a powder bed fusion-laser (PBF-L) additive manufacturing system has a support region and a top region. The top region is formed on the support region by a friction surfacing additive manufacturing (FSAM) process, such that the top region is under a compressive stress. The build plate can be prepared by preparing the build plate support region to receive the top region and depositing, using a FSAM process, a layer of metal on the support region. The layer of metal is formed with a compressive stress to form the top region. The top region is then machined to provide a desired surface roughness.
A system includes a turbine engine, a low spool gearbox, and a high spool gearbox. The system also includes a first lubrication system configured to serve the low spool gearbox, and a second lubrication system, independent from the first lubrication system, configured to serve the high spool gearbox. The turbine engine includes a low-pressure spool and a high-pressure spool. The low spool gearbox is disposed within a core compartment of the turbine engine. The low spool gearbox is configured to transfer mechanical power between the low-pressure spool and a first accessory. The high spool gearbox is disposed within the core compartment of the turbine engine. The high spool gearbox is configured to transfer mechanical power between the high-pressure spool and a second accessory.
A turbine engine including one or more conduits is disclosed herein. At least one conduit of the turbine engine includes a first end coupled to the turbine engine, a second end, and a body extending from the first end to the second end, the body including an interior surface and an exterior surface, wherein at least one of the interior surface or the exterior surface includes a protrusion configuration.
A machine comprising a rotor having: an inner member; an outer member encircling the inner member; and a groove in one of the inner member and the outer member. The groove has a first side wall, a second side wall and a base. A split ring seal is accommodated in the groove and contacts a surface of the other of the inner member and the outer member The first side wall has a plurality of open radial first channels and the second side wall has a plurality of open radial second channels.
A mid-turbine frame of a gas turbine engine includes a fairing and a damping system. The fairing includes an outer fairing wall, an inner fairing wall, and a plurality of struts. Each of the outer fairing wall and the inner fairing wall extends between and to an inner surface and an outer surface. The inner surface forms a core flow path through the fairing. One of the outer fairing wall or the inner fairing wall forms a damping wall of the fairing. The damping system includes at least one damping ring and a plurality of clamp assemblies. The at least one damping ring is disposed at the outer surface of the damping wall. The plurality of clamp assemblies is circumferentially arrayed on the damping wall about the axis. Each of the plurality of clamp assemblies clamps the at least one damping ring on the damping wall.
A gas turbine engine includes propulsor means including a plurality of propulsor blades rotatable about an engine longitudinal axis, and an inlet including an inlet leading edge plane. The gas turbine engine further includes reduction means, first compression means, second compression means, first expansion means, and second expansion means. A hub mounts the propulsor blades and mounts a spinner forward of the propulsor blades. The propulsor means include a propulsor leading edge forward-most point spaced apart from the inlet leading edge plane by an inlet length. The spinner includes a spinner length from a spinner forward-most point to the propulsor leading edge forward-most point. A ratio of the spinner length to the inlet length is greater than or equal to 0.5. A normalized internal area distribution defined between the inlet leading edge plane and the propulsor leading edge forward-most point is monotonically convergent.
F01D 25/24 - CasingsCasing parts, e.g. diaphragms, casing fastenings
F02C 3/04 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
An assembly for a gas turbine engine includes a primary combustor disposed within a core flow path of the gas turbine engine; a high-pressure turbine positioned downstream of the primary combustor; a low-pressure turbine positioned downstream of the high-pressure turbine; and a supplemental combustor positioned within the core flow path, downstream of the high-pressure turbine and upstream of the low-pressure turbine.
F02C 7/00 - Features, component parts, details or accessories, not provided for in, or of interest apart from, groups Air intakes for jet-propulsion plants
G01S 17/66 - Tracking systems using electromagnetic waves other than radio waves
An assembly is provided for a turbine engine with a flowpath. This assembly includes a fuel source and an engine component. The engine component forms a peripheral boundary of the flowpath. The engine component includes a component internal passage. The engine component is configured to receive fuel from the fuel source. The engine component is configured to crack at least some of the fuel within the component internal passage thereby cooling the engine component and providing at least partially cracked fuel. The assembly is configured to direct the at least partially cracked fuel into the flowpath for combustion.
A gas turbine engine includes a propulsor having a rotor with blades, and an adapter connecting the rotor to a drive line at a connection. The drive line is connected to an accessory. The drive line and the accessory are received within a static center body. Bolts secure the adapter to the rotor. The center body has a rear end closely spaced from the connection, such that the center body limits access to the connection of the adapter to the rotor. The center body is formed with a forward portion and a separate rear portion defining the rear end. The forward portion and the rear portion are secured together. The rear portion is selectively slidable relative to the forward portion such that the rear portion can be moved away from the connection to provide access to change balance members. A method is also disclosed.
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.
