Abstract: The present disclosure is directed to a rotating detonation combustor that includes a forward wall, a radially inner wall, and a radially outer wall. The forward wall is disposed at an inlet end of the rotating detonation combustor. The radially inner wall surrounds a longitudinal axis and extends downstream from the forward wall to an outlet end of the rotating detonation combustor. The radially outer wall extends downstream from the forward wall to the outlet end and surrounds the radially inner wall to define at least one annular plenum between the radially inner wall and the radially outer wall. At least one partition is proximate to the inlet end and defines at least two mixing zones. A plurality of oxidizer inlets and a plurality of fuel inlets are disposed at the inlet end in fluid communication with the at least two mixing zones.

Abstract: A turbofan engine may include a variable pitch fan rotatable about an axis and operable in a forward thrust mode by generating flow opposite a direction of travel and in a reverse thrust mode by generating flow in the direction of travel. In reverse thrust mode, a nacelle assembly defines an intermediate opening disposed aft of the variable pitch fan. A flow guide assembly is configured to guide a first fluid flow opposite the direction of travel within a bypass passage from a forward opening to an aft opening of the nacelle assembly in forward thrust mode, and to redirect a second fluid flow from outside the bypass passage opposite the direction of travel to inside the bypass passage in the direction of travel through the intermediate opening in reverse thrust mode. One or more flow channels are defined for generating at least one fluidic injection pressurized by a core.

Abstract: A method for controlling a turbojet engine thrust reverser during an aborted aircraft takeoff, the thrust reverser including doors movable between a stowed position, an overstowed position and a deployed position; door actuators to move the doors between the stowed, overstowed and deployed positions; a device for locking the doors in the stowed position, moveable between a locking position and an unlocking position; and a lock actuator to move the locking device between the locking and unlocking positions. The method includes decreasing the engine speed of the turbojet engine by following a setpoint value below a first engine speed threshold value at which the aerodynamic forces being exerted on the doors are equal to the forces developed by the door actuators; controlling the door actuators to bring the doors into the overstowed position; controlling the lock actuator to bring the locking device into the unlocking position.

Abstract: A combustor nozzle including nozzle modules, a combustor, and a gas turbine including the same are proposed. The nozzle module includes a fuel supply pipe having an internal fuel flow path, manifolds communicating with the fuel supply pipe and arranged in a row in a radial direction so that air inlet flow paths are respectively formed at positions therebetween, and fuel mixers disposed along a circumferential direction at the rear side of the manifold to receive the fuel from the manifold and inject the fuel, each of the plurality of fuel mixers including a mixer body having one end communicating with the manifold, wherein a tip is formed at the other end of the mixer body, a shroud disposed to surround an outer circumferential surface of the mixer body to guide an airflow, and at least one fuel port formed through the mixer body to inject fuel.

Abstract: The method is for controlling a deflagration combustion process in a pistonless combustor. The method includes scavenging combustion products of the previous cycle, introducing air into the combustor thereby initiating a flow pattern having a first flow component within the combustor. Air is introduced into the pistonless combustor in a nonparallel angle in relation to the previous air input and thereby creating a second flow component to the flow pattern for increasing speed of combustion propagation. Fuel mixed into the air is introduced for creating a fuel-air mixture flowing within the flow pattern, and igniting the fuelair mixture within the pistonless combustor thereby increasing pressure within the pistonless combustor.

Abstract: A fuel system of an aircraft engine, has: a fuel manifold having a first manifold inlet and a second manifold inlet; a transfer tube assembly having a first tube slidably engaged to the fuel manifold and fluidly connected to the first manifold inlet, and a second tube slidably engaged to the fuel manifold and fluidly connected to the second manifold inlet; and an adaptor having: a body slidably engaged by the first tube and by the second tube, a first member defining a first fuel conduit fluidly connected to the first manifold inlet via the first tube, and a second member defining a second fuel conduit fluidly connected to the second manifold inlet via the second tube.

Abstract: A rotating detonation combustor (RDC) and an associated system is disclosed herein. The RDC can be operably coupled with a vehicle to provide additional power capability over the primary motive power source. The RDC can be operated on demand to provide power to a turbine connected to an electrical power generator in some forms. The RDC can be used in stationary systems as well as moving vehicle systems.

Abstract: A thrust reverser includes a fixed structure and a movable structure defining an air stream, a flap including a single wall perforated on the surface thereof, and a structure strengthening the flap and having an acoustic function. The flap being articulated between the fixed structure and the movable structure to allow in a direct jet position, disposing the wall along an acoustic section of the movable structure and allowing the circulation of an air flow through the air stream, the acoustic section forming with the structure of the wall an acoustic resonator, and in a reverse jet position, disposing the wall to deflect an air flow passing through the air stream and allow the passage through the wall and through the air stream of a portion of the air flow deflected by the flap.

Abstract: A rotating detonation combustor (RDC) and an associated system is disclosed herein. The RDC can be operably coupled with a vehicle to provide additional power capability over the primary motive power source. The RDC can be operated on demand to provide power to a turbine connected to an electrical power generator in some forms. The RDC can be used in stationary systems as well as moving vehicle systems.

Abstract: A combustion chamber of an aircraft turbomachine having a main axis includes: a body of revolution coaxial with the main axis having a plurality of combustion tubes extending mainly in the direction of the main axis and being distributed in a ring about the main axis; a first perforated rotary disc mounted at a first axial end of the body and rotatable about the main axis to selectively open or close a first end of each of the combustion tubes; and a second perforated rotary disc mounted at a second axial end of the body and rotatable about the main axis to selectively open or close a second end of each of the combustion tubes. The body includes a plurality of cooling segments extending mainly in the direction of the main axis, which are distributed in a ring about the main axis and around the combustion tubes.

Abstract: A constant volume combustion system for a turbomachine comprises a plurality of combustion chambers distributed annularly about an axis, each combustion chamber comprising an inlet orifice and an outlet orifice, a selective shut-off element movable in rotation relative to the combustion chambers, the selective shut-off element comprising a shroud facing the inlet and outlet orifices of the combustion chambers. The shroud including on a first annular portion at least one intake aperture intended to cooperate with the inlet orifice of each combustion chamber during the rotation of the selective shut-off element and on a second annular portion at least one exhaust aperture intended to cooperate with the outlet orifice of each combustion chamber during the rotation of the selective shut-off element. Each combustion chamber comprises a fuel injection device whose opening and closing are synchronized by the shut-off element.

Abstract: A maximum output creator is provided with: a temperature reception unit that receives an intake temperature, which is the temperature of air taken in by a compressor of a gas turbine; a change reception unit that receives the details of a change to a maximum opening degree in an intake quantity adjuster of the compressor; a basic maximum output computation unit that determines a basic maximum output of the gas turbine on the basis of the intake temperature received by the temperature reception unit; a coefficient creation unit that creates a maximum output correction coefficient for correcting the basic maximum output on the basis of the details of the change to the maximum opening degree received by the change reception unit and the intake temperature received by the temperature reception unit; and a maximum output correction unit that corrects the basic maximum output using the maximum output correction coefficient.

Abstract: A process for using an air input of a turboreactor nacelle of an aircraft, comprising an air input lip which comprises at least one fixed portion and at least one portion which can be moved between a first position, in which the air input lip has an aerodynamic profile so as to guide the internal air flow over the internal wall in order to promote a thrust phase, and a second position, in which the portion is displaced in relation to the fixed portion so that the air input lip has a second radial thickness in the second position which is less than the first radial thickness in the first position so as to promote a reverse thrust phase.

Abstract: A system for cooling a compartment for equipment in a nacelle having a member for drawing a cooling flow and a device for admitting the cooling flow into the compartment, having: a regulating member movably mounted with the ability to move in a first direction of movement limiting the admission when the internal pressure of the compartment is below the external pressure, and in a second direction of movement increasing the admission when the internal pressure is above the external pressure; a return member configured to keep the regulating member in a neutral open position and to damp the movement of the regulating member during modifications in the ratio of the pressures; a pressing member activated by a heat-sensitive element and configured to move the regulating number into a closed position so as to isolate the compartment in the event of fire.

Abstract: A propulsion system includes a turbine engine, and a ramjet engine arranged in parallel with the turbine engine, a fuel storage tank that stores a fuel, such as an ammonia-based fuel, a heat exchanger through which the fuel passes to obtain gaseous (or supercritical) fuel components of the fuel, and a separator that separates the gaseous fuel components into a first fuel component and a second fuel component. The first fuel component is provided to the ramjet engine for ramjet combustion and/or to the turbine engine for turbine combustion, in a high speed operating state of the propulsion system, while the second fuel component may be provided to an inlet of the turbine engine for the turbine combustion.

Abstract: A rotating detonation combustor includes a combustion chamber configured for a rotating detonation process to produce a flow of combustion gas and an air plenum configured to contain a volume of air. The rotating detonation combustor also includes a flow passage coupled in flow communication between the combustion chamber and the air plenum and configured to channel an airflow from the air plenum. The rotating detonation combustor also includes a fuel inlet coupled in flow communication with the flow passage and configured to channel a fuel flow into the flow passage. The flow passage includes a plurality of fuel mixing mechanisms configured to mix the airflow and the fuel flow within the combustion chamber.

Abstract: A position sensor for an aircraft nacelle thrust reverser door for an aircraft turbojet engine includes a target to be detected, a detector having a detection range and including a detection head configured to detect the target when it enters its detection range, and a target guide configured to guide the target in order to position it in the detection range of the detector. The detector is stationary relative to the target guide. In one form, the target is intended to be positioned on the door, and the detector and the target guide are intended to be positioned on a stationary structure of the thrust reverser. In another form, the target is intended to be positioned on a stationary structure of the thrust reverser, and the detector and the target guide are intended to be positioned on the door.

Abstract: A door for a thrust reverser of a propulsion assembly, the door being provided with a flexible baffle. The baffle is configured to change state by moving from a folded position to an unfolded position, and vice versa, depending on the configuration of the reverser. When the reverser is in reverse jet configuration, the door is open and the baffle is in an unfolded position in order to direct a flow of gas upstream of the propulsion assembly, thus generating a counter-thrust. When the reverser is in direct jet configuration, the door is closed and the baffle is in a folded position so as to cancel or minimise interference caused by the baffle in the flow of gas passing through the propulsion assembly and thus generating a thrust.

Abstract: A combustor nozzle including nozzle modules, a combustor, and a gas turbine including the same are proposed. The nozzle module includes a fuel supply pipe having an internal fuel flow path, manifolds communicating with the fuel supply pipe and arranged in a row in a radial direction, and fuel mixers disposed along a circumferential direction at the rear side of the manifold to receive the fuel from the manifold and inject the fuel, each of the plurality of fuel mixers including a mixer body having one end communicating with the manifold, the other end that is opened, and an mixing flow path through which air and the fuel flow, an air inlet port formed on a lateral side of the mixer body, a fuel port formed inside the mixer body and discharging the fuel supplied from the manifold to the mixing flow path, and an extension located at the other end of the mixer body.

Abstract: A stationary blade unit including a first engaging portion projecting from an outer tube toward a first side in an axial direction; and a second engaging portion projecting from the outer tube toward a second side in the axial direction. A first outer shell includes a first flange and a first engaged portion into which the first engaging portion is slidingly inserted in the axial direction. A second outer shell includes: a second flange fastened to the first flange by a fastener; and a second engaged portion into which the second engaging portion is slidingly inserted in the axial direction. The stationary blade unit includes a projection projecting outward in a radial direction from the outer tube. One of the first outer shell and the second outer shell includes an integrated stopper arranged on a rotation trajectory of the projection around an axis.