A fully compounding rotorcraft includes a fuselage having first and second wings extending therefrom and configured to provide lift compounding responsive to forward airspeed. A twin boom includes first and second tail boom members that extend aftward from the first and second wings. An empennage is coupled between the aft ends of the tail boom members. An anti-torque system includes a tail rotor that is rotatably coupled to the empennage. An engine is disposed within the fuselage and is configured to provide torque to a main rotor assembly via an output shaft and a main rotor gearbox. An auxiliary propulsive system is coupled to the fuselage and is configured to generate a propulsive thrust to offload at least a portion of a thrust requirement from the main rotor during forward flight, thereby providing propulsion compounding to increase the forward airspeed of the rotorcraft.

A distributed control system for a turbomachine and method of operating the distributed control system are provided. In one aspect, a distributed control system includes a central controller and a distributed controller communicatively coupled thereto. The distributed controller has one or more associated local actuators and one or more associated local sensors. The actuators and the sensors are communicatively coupled with the distributed controller. If a communication link between the central controller and the distributed controller becomes faulty, the distributed controller enters an autonomous safety mode. In this mode, the distributed controller uses a combination of its own associated local sensors and past commands received from the central controller to autonomously govern its associated local actuators to maintain safe operation of the turbomachine.

A method of modifying an existing gas turbine to create a storage engine, the gas turbine having a combustor, a compressor section, and a turbine section, the method comprising modifying the compressor section of the gas turbine to form the storage engine such that air supplied to the combustor of the storage engine is heated by exhaust of the storage engine and is supplied from a remote source.

The present disclosure relates to a novel gas turbine system having applications, for example, in thermal power generation in an environmentally friendly manner. The multiloop gas turbine system may have multiple functional units each comprising a compressor, a regenerator, a combustion unit, and a turbine. Typically, exhaust flow of a turbine of a preceding loop may be routed to the combustion unit of the next loop, allowing mixing of exhaust flow with hot compressed air of the next loop, and the expanded exhaust from the turbine of the ultimate loop is fed back into the regenerators of each loop to recover exhaust heat.

An aircraft turbine engine includes at least one tubular element, such as a rotor shaft, and at least one rotor wheel extending around the tubular element and having a disk carrying at an external periphery of same, an annular row of blades, the disk extending at a radial distance h from the tubular element in such a way as to define an annular flow space for a cooling gas stream (Fr) during operation, wherein the tubular element includes at least one annular wall extending radially outwards and configured to divert the gas stream (Fr) in order for it to pass substantially radially between the disk and this annular wall.

An aircraft propulsion assembly comprising at least a main turbine mounted along a longitudinal axis, at least one main fan arranged upstream of the main turbine along the longitudinal axis and driven in rotation by the said main turbine, the said main fan being ducted by a main fan casing, an auxiliary turbine mounted along the longitudinal axis, the auxiliary turbine being independent of the main turbine, an auxiliary fan of axis offset with respect to the longitudinal axis and driven by the auxiliary turbine, the auxiliary fan being ducted by an auxiliary fan casing, the main casing being separate and distinct from the auxiliary casing so as respectively to generate a main secondary flow and an auxiliary secondary flow which remain independent of one another until they are discharged into the atmosphere.

Blade assemblies for gas turbine engines are described. The blade assemblies include a fan blade having a first tab extending from a leading edge and a second tab extending from a trailing edge. A first platform is affixed to a first side of the fan blade at the first and second tabs and a second platform is affixed to a second side of the fan blade at the first and second tabs. At least one first fastener is arranged to connect the first platform, the first tab, and the second platform at a first end and at least one second fastener is arranged to connect the first platform, the second tab, and the second platform at a second end.

A gas turbine engine for an aircraft may include a shaft fixed to a compressor of the gas turbine engine, at least one of a turboprop and a turbofan, a power gearbox having a power output shaft fixed to the at least one of the turboprop and the turbofan, wherein the power gearbox receives an input rotation from the shaft, an auxiliary gearbox having an auxiliary output shaft powering at least one auxiliary component of the gas turbine engine, a first lubrication system, and a second lubrication system that is separated from the first lubrication system. The first lubrication systems may circulates a first lubricant through the power gearbox, and the second lubrication system may circulate a second lubricant through the auxiliary gearbox.

An engine device, in particular based on a gas turbine, which can preferably be used for a hybrid-electrically driven aircraft. A drive section of the engine device produces an accelerated gas stream, which is further processed in a gas turbine of the engine device in order to generate thrust. The engine device further includes a power turbine section having multiple power turbines for providing drive power for multiple electric generators. The power turbines are designed in such a way that they can be driven solely on account of direct interaction with the accelerated gas stream leaving the gas turbine, i.e. solely by the gas stream itself and in particular not with the aid of a mechanical coupling to one of the moving components of the drive section.

An injector for a liquid rocket engine includes an array of injector elements. Each injector element includes a central passage and a plurality of peripheral transverse passages. The central passages are configured to provide axial injection flow and the peripheral transverse passages are configured to provide swirl injection flow about the axial injection flow. A portion of the injector elements are configured to provide the swirl injection flow in a clockwise direction and another portion of the injector elements are configured to provide the swirl injection flow in a counter-clockwise direction. The injector elements are arranged to form a plurality of circumferential rows. The injector elements of each individual circumferential row are either all of the clockwise direction or all of the counter-clockwise direction. At least one of the circumferential rows is of the clockwise direction and at least one of the circumferential rows is of the counter-clockwise direction.