Abstract: A method implemented using at least one processor includes receiving a plurality of images acquired from a plurality of image sensors disposed on a vehicle configured to engage an aircraft for ground operations. The method further includes determining at least one parameter about a potential obstacle based on the plurality of images and a machine vision algorithm. The method also includes generating an alert signal based on the at least one parameter, useful for avoiding collision of the aircraft.

Abstract: The present disclosure generally relates to variable cellular structures, methods of making such cellular structures, and variable cellular flow discouragers for turbine engines for jet aircraft.

Abstract: A fuel nozzle assembly includes a centerbody including an outer wall. The outer wall defines a plurality of fuel injection apertures. The fuel injection apertures include a first portion of the plurality of fuel injection apertures configured to induce a first fuel flow rate. The fuel injection apertures also include a second portion of the plurality of fuel injection apertures configured to induce a second fuel flow rate. The second fuel flow rate is less than the first fuel flow rate.

Abstract: A fuel injection system for use in a combustor of a turbine assembly includes a substantially annular fuel injection housing at least partially defining an annular cavity. The fuel injection system also includes a fuel manifold and an air manifold. The fuel manifold includes a plurality of fuel injection ports extending through the fuel injection housing. The air manifold includes a first plurality of air injection ports and a second plurality of air injection ports that each extends through the fuel injection housing.

Abstract: A tuned cavity rotating detonation combustion system includes a an annular chamber having an inlet and an outlet; a valve plate at the inlet of the annular chamber and comprising a plurality of openings spaced circumferentially around the inlet; a plurality of tubes each having an open end in communication with a corresponding opening of the valve plate and a closed end forming a tuned cavity, and a first opening between the open end and the closed end for injection of air; and a plurality of fuel injectors corresponding to the plurality of tubes, each fuel injector being configured to inject fuel into the tube between the first opening and the open end. Each of the tuned cavities has a length sized to resonate at a same frequency as a continuous detonation frequency of at least one detonation wave in the annular chamber.

Abstract: A gas turbine engine system and method of operating gas turbine engines are provided. The gas turbine engine assembly includes a gas turbine engine includes a power shaft configured to rotate about an axis of rotation. The gas turbine engine assembly also includes a first fan and a second fan coupled to the power shaft coaxially with the gas turbine engine. The gas turbine engine assembly also includes a first fan duct configured to direct a first stream of air to the first fan. The gas turbine engine assembly also includes a second fan duct configured to direct a second stream of air to the second fan. The gas turbine engine assembly also includes an exhaust duct configured to direct a stream of exhaust gases of the gas turbine engine in a direction of the axis of rotation.

Abstract: A turbine engine airfoil apparatus, including an airfoil defined by a plurality of airfoil sections arrayed along a stacking axis that extends between a root and a tip, wherein at least two of the airfoil sections spaced apart from each other have differing airfoil section thermal expansion properties.

Abstract: Embodiments of a propulsion system are provided herein. In some embodiments, a propulsion system for an aircraft may include an electrical power supply; a motor coupled to the electrical power supply, wherein the electrical power supply provides power to the motor; and a fan disposed proximate a rear portion of an aircraft and rotatably coupled to the motor, wherein the fan is driven by the motor.

Abstract: A fuel injection system for use in a combustor of a turbine assembly includes a substantially annular fuel injection housing at least partially defining an annular cavity. The fuel injection system also includes a fuel manifold and an air manifold. The fuel manifold includes a plurality of fuel injection ports extending through the fuel injection housing. The air manifold includes a first plurality of air injection ports and a second plurality of air injection ports that each extends through the fuel injection housing.

Abstract: A turbine assembly is provided. The turbine assembly includes a gas turbine engine including at least one hot gas path component formed at least partially from a ceramic matrix composite material. The turbine assembly also includes a treatment system positioned to receive a flow of exhaust gas from the gas turbine engine. The treatment system is configured to remove water from the flow of exhaust gas to form a flow of treated exhaust gas, and to channel the flow of treated exhaust gas towards the at least one hot gas path component. The at least one hot gas path component includes a plurality of cooling holes for channeling the flow of treated exhaust gas therethrough, such that a protective film is formed over the at least one hot gas path component.

Abstract: A turbine system and method of operating is provided. The system includes a compressor configured to generate a compressed low-oxygen air stream and a combustor configured to receive the compressed low-oxygen air stream and to combust a fuel stream to generate a post combustion gas stream. The turbine system also includes a turbine for receiving the post combustion gas stream to generate a low-NOx exhaust gas stream, a heat recovery system configured to receive the low-NOx exhaust gas stream and generate a cooled air stream and an auxiliary compressor configured to generate an oxygen and water vapor deficient cooled and compressed air stream. A portion of the oxygen and water vapor deficient cooled and compressed air stream is directed to the combustor to generate an Oxygen and H2O deficient film on exposed portions of the combustor, and another portion is directed to the turbine to provide a cooling flow.

Abstract: A gas turbine engine system includes a compressor, a combustor, and a turbine. The combustor is coupled to the compressor and disposed downstream of the compressor. The combustor includes a secondary combustor section coupled to a primary combustor section and disposed downstream of the primary combustor section. The combustor also includes a transition nozzle coupled to the secondary combustor section and disposed downstream of the secondary combustor section. The combustor further includes an injector coupled to the secondary combustor section, for injecting an air-fuel mixture to the secondary combustor section. The turbine is coupled to the combustor and disposed downstream of the transition nozzle; wherein the transition nozzle is oriented substantially tangential to the turbine.

Abstract: A combustor assembly for use in a gas turbine engine is provided. The combustor assembly includes a dome assembly, a fuel nozzle coupled to the dome assembly, and a cowl assembly. The cowl assembly includes a radially inner cowl coupled to the dome assembly, and a radially outer cowl including a first cowl member and a second cowl member each positioned about the fuel nozzle. The first and second cowl members substantially align to define a substantially continuous flow path configured to direct a flow of air about the dome assembly.

Abstract: A combustor assembly including a combustor liner defining therein a combustion chamber for the downstream flow of a main fluid. At least two annular trapped vortex cavities are located on the combustor liner and staged axially spaced apart. A cavity opening is located at a radially inner end of each of the at least two annular trapped vortex cavities spaced apart from a radially outer wall and extending between an aft wall and a forward wall of each cavity. A plurality of injectors are configured tangentially relative to the circular radially outer wall to provide for an injection of air and fuel to form an annular rotating trapped vortex of a fuel and air mixture within a respective annular trapped vortex cavity. The annular rotating trapped vortex of the fuel and air mixture at the cavity openings is substantially perpendicular to the downstream flow of the main fluid. A gas turbine engine including the combustor assembly is disclosed.

Abstract: A system for premixing fuel and air prior to combustion in a gas turbine engine includes a mixing duct, a centerbody fuel injector located along a central axis of the mixing duct an outer annular swirler located adjacent an upstream end of the mixing duct for swirling air flowing therethrough in a first swirl direction and an inner annular swirler located adjacent of the mixing duct upstream end for swirling air flowing therethrough in a second swirl direction. The system includes a hub separating the inner and outer annular swirlers to permit independent rotation of an air stream therethrough and multiple hollow paths located radially outward around the centerbody fuel injector and at a radially inward side of the inner annular swirler for allowing a flow of sweeping air over the surface of the centerbody fuel injector.

Abstract: A system includes an intercooler configured to receive an airflow from a first compressor, to transfer heat from the airflow to a working fluid, and to provide the airflow to a second compressor. The system also includes an evaporative chiller configured to receive the working fluid from the intercooler, to chill the working fluid via evaporative cooling within an ambient air environment, and to provide the working fluid to the intercooler. In addition, the system includes a desiccant system configured to reduce a humidity of ambient air within the evaporative chiller.

Abstract: A shrouded fan assembly for use with a gas turbine engine includes a central hub and a plurality of variable-pitch fan blades coupled to and extending radially outward from the central hub. Each variable-pitch fan blade includes a radially outer tip is rotatable about a longitudinal axis extending therethrough. A shroud is coupled to each radially outer tip.

Abstract: An air-shielded fuel injection assembly for use in a combustion chamber of a turbine assembly. The air-shielded fuel injection assembly generally includes a fuel manifold including a plurality of fuel injection ports and an air manifold including a plurality of air injection ports. Each of the plurality of fuel injection ports is configured to introduce a fuel column into an annular cavity of a mixer assembly. Each of the plurality of air injection ports is configured to introduce an air curtain about an associated fuel injection column to minimize recirculation upstream of the fuel injection column and increase penetration of the fuel injection column into the cavity. Also disclosed are a mixer assembly and a turbine assembly including the air-shielded fuel injection assembly.

Abstract: A gas detector and method are presented. The gas detector includes a launcher unit for coupling and merging light beams in mid-infrared and infrared wavelength ranges into a single light beam and directing the merged single light beam towards a gas flow path; a receiver unit for generating at least one photo detector current signal based on the light beam transmitted through the gas flow path; and a control unit for processing at least one photo detector current signal to measure concentration of the at least two gases present in the gas flow path.

Abstract: A method of manufacturing an internally cooled blisk by additive manufacturing processes to provide weight reduction and power-to-weight ratio improvement, leading to improved efficiency, a decrease in fuel burn and lifecycle costs in an engine employing the internally cooled blisk. The method of manufacturing including defining a configuration for the internally cooled blisk, the configuration comprising a disk, an annular array of angularly spaced blades extending about a periphery of the disk and one or more internal cooling features defined within the internally cooled blisk. The method further including the step of programming the configuration into an additive manufacturing system. A powder is deposited into a chamber and an energy source is applied to the deposited powder to consolidate the powder into a cross-sectional shape corresponding to the defined configuration. Additionally provided is an internally cooled blisk.