Abstract: A divergent flap includes a hinge joint to connect the divergent flap to a convergent flap, a plow structure, a first wall, a second wall, and two side walls. The plow structure is at a plow end of the flap opposite the hinge joint. The first wall and the second wall each extend from the hinge joint to the plow end. The second wall is spaced apart from the first wall by the two side walls to form at least one cooling channel extending between the hinge joint and the plow end. The divergent flap is integrally formed in layer-by-layer fashion as a single piece.

Abstract: A plurality of gas turbine combustors having a cross fire tube assembly that connects adjacent combustors. The combustors include combustion chambers having annular combustion air passages on outer peripheries thereof. The cross fire tube assembly has a dual pipe configuration including an inner tube that connects the combustion chambers of the adjacent combustors and an outer tube that covers therein the inner tube and connects the combustion air passages of the adjacent combustors. The cross fire tube assembly further has openings disposed between the inner tube and the outer tube of outer peripheral partition walls of the combustion air passages that are connected with the outer tube of the cross fire tube assembly centering on the inner tube. The openings allow combustion air to flow in areas upstream and downstream of the inner tube with respect to a flow of the combustion air flowing through the combustion air passages.

Abstract: A gas turbine engine for an aircraft includes a compressor, a combustion chamber, and a turbine having at least one stator, and at least one rotor. Each stator and rotor is formed by a plurality of blades, a fluid channel is formed between two consecutive blades, and each blade has two opposing surfaces. The compressor is in fluid communication with a first group of stator channels, and the combustion chamber is in fluid communication with a second group of stator channels, such that heat exchange can be performed through two opposing surfaces of at least one stator blade. The outer and the inner walls define a duct for the passage of the heated fluid through the rotor blades, and the outer wall is also arranged for directing the compressed air towards the combustion chamber.

Abstract: An overthrust protection system for an aircraft engine. The system comprises an engine overspeed protection unit comprising overspeed logic and an overspeed solenoid valve controlled by the overspeed logic, the overspeed logic configured to energize the overspeed solenoid valve for removing fuel flow to the engine upon detection of an overspeed condition of the aircraft engine; and an overthrust controller coupled to the overspeed protection unit and configured to measure engine thrust and to detect an overthrust condition when an engine thrust threshold has been exceeded, and configured to trigger energizing of the overspeed solenoid valve upon detection of the overthrust condition and an aircraft-on-ground condition.

Abstract: An air turbine starter device comprises a rotor arranged in a cavity of a housing, a first manifold having a cavity with a port operative to direct compressed air to the rotor, a second manifold having a cavity with a port operative to direct compressed air to the rotor, wherein the first manifold is larger than the second manifold.

Abstract: A combustor includes fuel nozzles that extend in an axial direction of a combustor main body and are capable of injecting fuel from injection holes on a combustion chamber side. The combustor includes a phase adjusting unit which partially changes a flow path cross-sectional area of at least one of the fuel nozzles in the axial direction such that phases of flow rate fluctuation of fuel do not match with respect to at least two of the fuel nozzles.

Abstract: Apparatus, systems, and methods store energy by liquefying a gas such as air, for example, and then recover the energy by regasifying the liquid and combusting or otherwise reacting the gas with a fuel to drive a heat engine. The process of liquefying the gas may be powered with electric power from the grid, for example, and the heat engine may be used to generate electricity. Hence, in effect these apparatus, systems, and methods may provide for storing electric power from the grid and then subsequently delivering it back to the grid.

Abstract: The present disclosure is directed to a fuel nozzle for a gas turbine engine. The fuel nozzle includes a sleeve defining a plurality of radially oriented air inlet ports. The sleeve defines a sleeve outlet at the downstream end of the fuel nozzle and a longitudinally extended annular inner wall inward of the sleeve in the radial direction. The inner wall defines a fluid passage and an inner wall outlet, in which the inner wall outlet is disposed toward the downstream end of the inner wall. At least a portion of the plurality of radially oriented air inlet ports is outward of the inner wall along the radial direction. The fuel nozzle further includes a plurality of fuel injectors surrounding the sleeve, in which the sleeve and each of the fuel injectors are connected to an aft body at the downstream end and a forward body at the upstream end.

Abstract: During a stage (E0) of starting the turbine engine, the method of the invention comprises: an open-loop generating step (E10) of generating a fuel flow rate command (WF_OL) from at least one pre-established relationship; and a closed-loop monitoring step (E20-E30) of monitoring at least one operating parameter of the turbine engine selected from: a rate of acceleration (dN2/dt) of a compressor of the turbine engine; and a temperature (EGT) at the outlet from a turbine of the turbine engine; this monitoring step comprising maintaining (E30) the operating parameter in a determined range of values by using at least one corrector network (R1, R2, R3) associated with the parameter and suitable for delivering a signal for correcting the open-loop generated fuel flow rate command so as to maintain the operating parameter in the determined range of values.

Abstract: A fuel nozzle includes an outer body extending parallel to a centerline axis, having a generally cylindrical exterior surface, forward and aft ends, and a plurality of openings through the exterior surface. The fuel nozzle further includes an inner body inside the outer body, cooperating with the outer body to define an annular space, and a main injection ring inside the annular space, the main injection ring including fuel posts extending therefrom. Each fuel post is aligned with one of the openings and separated from the opening by a perimeter gap which communicates with the annular space. There is a circumferential main fuel gallery in the main injection ring, and a plurality of main fuel orifices, wherein each orifice communicates with the main fuel gallery and extends through one of the fuel posts.

Abstract: An exhaust mixing system having an attachment flange configured to couple a lobed Oxide-CMC exhaust mixer to a turbine engine. The attachment flange includes an annulus and a plurality of projections extending from the annulus in a first direction that is substantially parallel to a central axis of the annulus. The annulus is configured to couple to the turbine engine and the plurality of projections is configured to couple to the lobed mixer.

Abstract: A fuel distribution manifold for a gas turbine engine is disclosed herein. The fuel distribution manifold includes a manifold body, a stage divider, and a plurality of manifold outlet ports. The stage divider subdivides an internal volume of the manifold body into a primary stage volume and a secondary stage volume. A majority of the plurality of manifold outlet ports extends from the manifold body between the primary inlet port and the stage divider, each being fluidly connected to the primary stage volume. A remainder of the plurality of manifold outlet ports extends from the manifold body between the secondary inlet port and the stage divider, each being fluidly connected to the secondary stage volume.

Abstract: A gas turbine engine has a fan rotor including at least one stage, with the at least one stage delivering a portion of air into a low pressure duct, and another portion of air into a compressor. The compressor is driven by a turbine rotor, and the fan rotor is driven by a fan drive turbine. A channel selectively communicates air from the low pressure duct across a boost compressor.

Abstract: A gas turbine equipped with a compressor, a combustion chamber and a turbine is further equipped with: a pressurizing device for taking a portion of the compressed air compressed by the compressor, and pressurizing the compressed air; a combustion chamber cooling line for cooling the combustion chamber using the pressurized compressed air; a temperature regulating line for regulating the temperature of a stationary member of a blade ring or the like of a turbine by using the pressurized compressed air; a combustion chamber supply line through which the compressed air flows from the pressurizing device to the combustion chamber cooling line; a turbine supply line through which the compressed air flows from the pressurizing device to the temperature regulating line; a heater for heating the compressed air, provided in the turbine supply line; and a control device capable of controlling the heater.

Abstract: An exhaust gas diffuser for a gas turbine engine whose inlet geometry can be selectively controlled to change the angular orientation of the diffuser at the location where the exhaust gas exits the last stage row of blades of the turbine section of the gas turbine engine. An end portion of the gas diffuser proximate the last stage row of blades can include one or more actuated sections that are independently controlled to change the angular orientation of the inlet geometry of the diffuser. In one embodiment, the angular orientation of the actuated sections is set at the manufacturing level for the service location of the engine. In another embodiment, the angular orientation of the actuated sections is selectively controlled based on the operating conditions of the engine. In another embodiment, the angular orientation of the actuated sections is controlled by pneumatic pressure from a compressor section of the engine.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a combustor wall, which includes a shell, a heat shield and an annular body. The body at least partially defines a first aperture through the shell and the heat shield. The body also defines one or more second apertures through which air is directed into the first aperture and provides non-uniform cooling to the body.

Abstract: A method for regulating a gas turbine wherein the fuel quantity supplied to the burners of the gas turbine is regulated using a target value for the corrected turbine outlet temperature. A stable operation of the gas turbine is to be allowed with a particularly high degree of efficiency and a high output at the same time. The target value for the corrected turbine outlet temperature is set using a value which characterizes the combustion stability in the burners, wherein the target value for the corrected turbine outlet temperature is set additionally using the surrounding temperature. Furthermore, the target value for the corrected turbine outlet temperature is set only below a specified surrounding temperature using the value which characterizes the combustion stability in the burners.

Abstract: A fluid intake including first and second ducts and a particle separation spinner defining an interface between the first and second ducts is disclosed. Spinner includes flow passages passing from first duct side of the spinner to second duct side of the spinner and splitter bodies separating the flow passages. Flow passages and splitter bodies are arranged in thread-like screw manner about spinner. In use, the spinner is spun about an axis of rotation, axis and direction of rotation being such that, from static frame of reference with respect to first duct, splitter bodies and flow passages have a component of movement towards a main travel direction of a fluid flow incident towards spinner from first duct. Splitter bodies are arranged such that they oblige fluid in the fluid flow to follow a convoluted path if it is to pass from first duct to second duct via the flow passages.

Abstract: A fuel nozzle apparatus has a centerline axis, and includes: a main fuel injector including an enclosed interior volume in fluid communication with a plurality of fuel ports configured to discharge fuel therefrom; a main fuel conduit disposed upstream of the main fuel injector and configured to supply liquid fuel to the interior volume; and a fluid lock disposed between the main fuel conduit and the main fuel injector, the fluid lock including a plurality of parallel capillary channels.

Abstract: A single, unitary thermal management system (TMS) manifold for a gas turbine engine is provided which comprises one or more interfaces for mounting various thermal management system components directly to the TMS manifold. The TMS manifold also defines fluid passages for transferring fuel, engine lubricant or generator oil from one component to another component. Packing numerous fuel system and lubricating system components within the TMS manifold reduces cost and weight and simplifies maintenance.