Abstract: Systems and methods are provided for an air-driven augmentor fan equipped aircraft propulsor. The augmentor fan may increase the effective bypass ratio of the aircraft propulsor and reduce fuel consumption and carbon emissions of the aircraft. The augmentor fan may be driven by air energized by a ducted fan powered by the core engine of the aircraft propulsor. The energized air may be received by an inlet, flowed through a flow path, and exhausted out the outlet to drive the augmentor fan. The exhausted energized air may impart a torque on the augmentor fan or blades of the augmentor fan. One or more of the inlet, flow path, or outlet may be variable in size to control the volume of air flowed through the flow path.

Abstract: The invention relates to a device for supporting and centring a fuel injector in a turbine engine combustion chamber, which includes means for centring a fuel injector along an axis, which are movable in a plane that is radial to the centring axis (52) in supporting means intended for being attached to the bottom of an annular chamber (18). According to the invention, the centring means include at least two radially external tabs (54, 56) each inserted respectively in a circumferential recess (60, 64) of the supporting means, the device including circumferential abutment means (78, 80, 74, 76, 82, 86, 84, 88) of the radial tabs (54, 56) of the centring means in the circumferential recesses (60, 64), the circumferential abutment means being configured such as to enable a greater angular displacement of a first (54, 154) one of the radial tabs in a first circumferential recess (60) relative to a second (56, 156) one of the radial tabs in a second circumferential recess (64).

Abstract: A combustor arrangement for a gas turbine includes a first burner, a first combustion chamber, a mixer for admixing a dilution gas to the gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection. These elements of the combustor arrangement are arranged in a row to form a flow path extending between the first combustion chamber and the second burner. The combustor arrangement includes acentral lance body arranged inside the flow path and extending from the first burner through the first combustion chamber into the mixer and into the second burner, wherein the lance body includes a fuel duct for providing fuel for the first burner and/or for the second burner.

Abstract: The invention relates to a device for installing and removing a component on a gas turbine, having a rail system on which in particular a rotor can be displaced along a predetermined displacement axis, and having a terminal positioning part which is fixedly connected to the rail system. The positioning part has at least one guide portion which is designed to co-operate mechanically with a respective counterpart guide portion on the housing of the gas turbine, in order to achieve a suitable orientation of the device with respect to the housing of the gas turbine. The positioning part also has at least one connecting section which is designed in each case to form a fixed but releasable connection to a respective counterpart connecting section on the housing of the gas turbine construct.

Abstract: A gas turbine engine has a first source of air to be delivered into a core of the engine, and a second source of air, distinct from the first source of air and including separately controlled first and second fans, each delivering air into respective first and second conduits connected to distinct auxiliary applications.

Abstract: A fairing assembly for a gas turbine engine includes an inner fairing platform; an outer fairing platform located radially outward of the inner fairing platform; a plurality of vanes radially outward from the inner fairing platform between the inner fairing platform and outer fairing platform; a mount located on a radially outer side of the outer fairing platform, the mount attached to the outer fairing platform on the radially outer side; and a link that includes an elongated member with a first end and an opposite second end, the first end configured to fit within the mount and define a pivot axis of the link with respect to the mount.

Abstract: A combined cycle power plant is described, comprising a steam turbine, a first heat recovery steam generator and a second heat recovery steam generator. The first heat recovery steam generator is attached to the steam turbine by a first steam pipe system and the second heat recovery steam generator is attached to the steam turbine by a second steam pipe system, to allow steam to pass from the first and second heat recovery steam generators to the steam turbine. The steam turbine is arranged in between the first heat recovery steam generator and the second heat recovery steam generator.

Abstract: This invention provides a gas turbine power generation equipment adapted to prevent step out of an electric power generator due to a system trouble such as a momentary power failure in a local grid including intermittent renewable energy generation equipment and the gas turbine power generation equipment. The gas turbine power generation equipment that supplies electric power in cooperation with the intermittent renewable energy generation equipment in the local grid interconnected to a power grid includes: a fuel flow control valve that controls a flow rate of fuel supplied to a combustor; a bleed valve or inlet air flow control valve that controls a flow rate of air compressed by a compressor and supplied to the combustor; and a control unit configured so that if voltage on the power grid decreases below a threshold V(t), the control unit outputs a control signal to at least one of the fuel flow control valve, the bleed valve, and the inlet air flow control valve.

Abstract: A gas turbine engine blade containment system is disclosed. The blade containment system may include a generally cylindrical casing being made of a first material, and a generally cylindrical ring being made of a second material coaxially surrounding the casing, at least some portion of the ring metallurgically bonded to the casing.

Abstract: An aircraft with aturbofan engine assembly having at least one compressor, a nacelle surrounding the turbine engine and defimng an annular bypass duct between the nacelle and the turbine engine, a thrust reverser having at least one moveable control surface, a thrust reverser locking system configured to selectively lock the thrust reverserand an injection cooling system.

Abstract: A method to operate a combustor of a gas turbine is provided. The method includes monitoring the combustion gas temperature by temperature measurements downstream said combustor to measure a respective combustion gas temperature at different locations at respectively equal flow-distances to the burner of the combustion gas, comparing the temperature measurements, opening a valve or increasing the opening position of the valve to control the portion of oxygen containing gas to be tapped off when the comparison reveals that a difference between the temperature measurements exceeds a temperature difference threshold ?T1.

Abstract: An emergency starter that allows responsiveness within a few seconds, without having disadvantages associated with mass and size of a back-up hydraulic or pneumatic starter. An instantaneous gas thrust of pyrotechnic type is coupled with a positive displacement transmission generator in conjunction with automatic coupling to/uncoupling from a set that is to be started. An emergency start-up system includes at least one pyrotechnic gas generator connected to an electrical initiator itself connected to a computer, a positive displacement motor housing straight-cut gears, the pyrotechnic gas generator being coupled to the motor by an inlet in the casing. The motor includes a mechanism of connection capable of moving at one end of the drive shaft configured to couple the transmission shaft to a driven shaft of the set that is to be started via a centrifugal clutch.

Abstract: A flame-holder device for a reheat channel of a turbojet, the device including an arm in the form of a trough defining a cavity and a heat shield fastened in the cavity of the arm. The flame-holder device further includes a fastener plate including a first leg integrally formed with the fastener plate and a second leg removably mounted on the plate, the arm being fastened to the first and second legs via fastener members.

Abstract: A gas turbine engine has a fan drive turbine for driving a gear reduction. The gear reduction drives a fan rotor. A lubrication system supplies oil to the gear reduction, and includes a lubricant pump to supply an air/oil mixture to an inlet of a deaerator. The deaerator includes a separator for separating oil and air, delivering separated air to an air outlet, and delivering separated oil back into an oil tank. The separated oil is first delivered into a pipe outwardly of the oil tank, and then into a location beneath a minimum oil level in the tank. Air within the oil tank moves outwardly through an air exit into the deaerator. A method of designing a gas turbine engine is also disclosed.

Abstract: A gas turbine engine having a fan and a bypass duct for discharging airflow from the fan to generate engine thrust. The bypass duct includes a first component and a second component each having a gas washed surface, the first component being positioned adjacent the second component with a gap therebetween. A sealing strip is positioned in the gap between the first component and the second component distal to the gas washed surfaces of the first and second components. A sealant is positioned between the first and second component proximal to the gas washed surfaces of the first and second components, so as to seal the gap between the first component and the second component from airflow through the bypass duct.

Abstract: A transition duct defining an airflow pathway between a low pressure compressor and a high pressure compressor of a gas turbine engine is disclosed. The transition duct may comprise an inner wall and an outer wall located radially outward of the inner wall with respect to a central axis of the gas turbine engine. It may further comprise a first bend configured to turn the airflow radially inward with respect to the central axis, and a turning vane located at the first bend between the inner wall and the outer wall. The turning vane may be configured to assist the first bend in turning the airflow radially inward.

Abstract: An air intake structure for a turbojet engine nacelle includes an inner panel to be attached to a fan casing, and an external panel capable of a translational movement in a longitudinal direction of the nacelle. The external panel incorporates a portion of an air inlet lip able to provide a junction between the inner panel and the external panel, and each of the air inlet lip portion and the inner panel is equipped, at least in the region of a joining end, with an acoustic attenuation structure. In particular, at least one of the joining ends is equipped with a radial buffer able to come into contact with a corresponding joining flange exhibited by the other joining end when the external panel is in a closed position.

Abstract: Aspects of the disclosure are directed to a thrust reverser of an aircraft, comprising: a wall having a first surface that partially forms a flow channel associated with an air flow, a blocker door having a second surface that partially forms the flow channel, and a dielectric elastomeric device that is configured to selectively expand and contract within a cavity formed between the wall and the blocker door where the cavity is substantially radially adjacent to the flow channel when the thrust reverser is in a stowed state.

Abstract: A system for supplying a turbomachine with fluid, the supply system including a low-pressure pumping unit intended to increase the pressure of the fluid flowing toward a downstream circuit. The downstream circuit divides at an inlet node, situated between the low-pressure pumping unit and the high-pressure volumetric pump, into a circuit supplying an injection system and a variable geometries supply circuit. The circuit supplying the injection system includes a high-pressure volumetric pump. The variable geometries supply circuit is configured to convey the fluid toward variable geometry from the inlet node to an outlet node connecting the variable geometries supply circuit to the upstream circuit between two pumps of the low-pressure pumping unit.

Abstract: An engine for propelling an aircraft includes an annular fuel cell arrangement having at least one fuel cell, at least one electric motor couplable to the fuel cell arrangement, at least one fan couplable to the electric motor and a cowling having an inlet and an outlet nozzle. The fuel cell arrangement is arranged inside the cowling, the at least one fan is arranged between the inlet and the fuel cell arrangement inside the cowling, the electric motor is operable under consumption of electric power delivered by the fuel cell arrangement and the at least one fan is designed to produce a thrust force by creating an accelerated airflow at the outlet nozzle. The engine is extremely efficient and comprises a distinct low noise.