Abstract: An exhaust nozzle of a gas turbine engine which includes a nozzle wall, the nozzle wall including an upstream, fixed structure and a downstream, translatable structure that is translatable relative to the fixed structure, a flow channel which is limited radially outwards by the nozzle wall, a centerbody arranged in the flow channel, at least one strut connecting the centerbody to the nozzle wall, a thrust reverser unit that comprises blocking doors, a first actuation system for deployment of the blocking doors into a deployed position for thrust reversal, and a second actuation system for translating the translatable structure between a stowed, upstream position and a fully deployed, downstream position. It is provided that the at least one strut is connected to the fixed structure, that the blocking doors are connected to the translatable structure, and that the first actuation system and the second actuation system are independent actuation systems.

Abstract: A gas turbine engine for an aircraft including a unit supplied with oil from a first oil circuit and a second oil circuit. The first oil circuit and the second oil circuit each are fluidly coupled with at least one inlet and with at least one outlet of the unit and with at least one inlet and with at least one outlet of an oil tank. The first oil circuit and the second oil circuit are configured to receive oil from the oil tank and to direct the received oil to the unit. The oil tank is incorporating offset outlets to each of the oil circuits. The offset outlet of the second oil circuit is positioned higher in the oil tank than the offset outlet of the first oil circuit.

Abstract: A gas turbine engine includes a compressor section, the compressor section including a variable inlet guide vane which is movable between distinct angles to control the airflow approaching the compressor section. A control is programmed to position the vane at startup of the engine to direct airflow across the compressor section. The engine includes a fan for delivering bypass air into a bypass duct positioned outwardly of a core engine including the compressor section. The position of the vane is configured to direct airflow across the compressor section while an aircraft associated with the gas turbine engine is in the air, and to increase a windmilling speed of the compressor section and the turbine rotors. A method and variable inlet vane are also disclosed.

Abstract: The present invention is a system, method and computer program for operating a land-or marine-based multi-spool gas turbine system for generating electrical power to supply a load, wherein the system comprises a multi-spool gas turbine engine, and the method comprises controlling the rotational speeds of the at least two generators independently of each other so as to directly control the rotational speeds of the shafts of the spools, the at last two generators being operable to generate electrical current to supply the load, and regulating the amount of heat generated the first heat generating equipment. A computer program to perform the method is also disclosed.

Abstract: A turbofan engine includes a geared architecture for driving a fan about an axis. The geared architecture includes a sun gear rotatable about an axis, a plurality of planet gears driven by the sun gear and a ring gear circumscribing the plurality of planet gears. A carrier supports the plurality of planet gears. The geared architecture includes a power transfer parameter (PTP) defined as power transferred through the geared architecture divided by gear volume multiplied by a gear reduction ratio and is between 219 and 328.

Abstract: The gas turbine includes a compressor to compress air introduced thereinto, a combustor to mix the compressed air with fuel for combustion, a main turbine having a plurality of turbine blades rotated by an energy produced by combustion gas in the combustor, a heat recovery boiler to produce steam by heat exchange with the combustion gas, and a fluid accelerator supplied with a first fluid compressed in the compressor to compress the first fluid and supply the compressed first fluid to the combustor, where the fluid accelerator includes a first inlet through which the first fluid is introduced, a second inlet through which a second fluid having a higher pressure than the first fluid is introduced, and an outlet through which the first and second fluids are mixed and discharged.

Abstract: A gas turbine engine for an aircraft has an engine core comprising turbine, compressor, and core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine of the engine, and having turbine blades, and the compressor being the lowest pressure compressor of the engine; fan located upstream of the engine core; and gearbox that receives an input from the core shaft and outputs drive to the fan. The engine core further has three bearings arranged to support the core shaft, the three bearings having two rearward bearings located downstream of the leading edge of the lowest pressure turbine blade of the turbine at the root of the blade, and/or, when the turbine comprises four sets of turbine blades, downstream of the trailing edge of a turbine blade of the third set of turbine blades from the front of the turbine, at the root of the blade.

Abstract: An aircraft fuel system can include a fuel line configured to transport fuel therein, an exposed aircraft structure in direct or indirect thermal communication with the fuel in the fuel line to receive heat from the fuel to provide a deicing or anti-icing heat to the exposed aircraft structure. The exposed aircraft structure can include at least one internal fuel channel in fluid communication with the fuel line for direct thermal communication with the fuel. The system can include a fuel/fluid heat exchanger in fluid communication with the fuel line to transfer heat from the fuel to a fluid to provide indirect thermal communication between the fuel and the exposed aircraft structure.

Abstract: An engine system for an aircraft includes a first gas turbine engine, a first core turning system, a second gas turbine engine, and a second core turning system. The engine system also includes a controller operable to shutdown the first gas turbine engine responsive to determining that the aircraft has landed and operate in the second gas turbine engine in a taxi mode while using the first core turning system to cool the first gas turbine engine. The controller is further operable to shutdown the second gas turbine engine and disable the first core turning system based on a power-down condition, restart the first gas turbine engine and use the second core turning system to cool the second gas turbine engine based on a restart condition, and complete cooling of the second gas turbine prior to restarting the second gas turbine engine.

Abstract: A nacelle comprising a fixed structure with a fixed cowling and a cowling movable in translation between closed and open positions, to open an opening between a duct and the exterior. A blocking door is rotatably movable on the movable cowling between a stowed position and a deployed position. A deployment system comprises an arm with a first end articulated on the blocking door and a second end that bears a stop, a slider fixed to the arm, a linear guide system is fixed on the fixed structure and guides the slide, and a cam. The stop is configured to lie against the cam when the movable cowling reaches an intermediate position from the closed position and to be guided by the cam to drive the rotation of the arm about the slider when the movable cowling moves from the intermediate position to an open position.

Abstract: A heat exchanger for a ducted fan gas turbine engine has a low temperature side and plural high temperature sides. The heat exchanger is configured such that heat is extracted from respective engine fluids flowing through the high temperature sides and is received by a portion of a bypass airflow of the engine which, on passing through the fan duct, is diverted through the low temperature side of the heat exchanger thereby cooling the engine fluids.

Abstract: A heat engine including a wall assembly is generally provided. The wall assembly includes a plurality of radial walls coupled together via a connecting member. The radial wall defines a flow opening therethrough. A flow cavity is defined between the plurality of radial walls and the connecting member.

Abstract: An apparatus is disclosed, including a first article, a second article, at least one interface structure, and a thermal break directly adjacent to the at least one interface structure. The first article includes a first material composition having a first thermal tolerance. The second article includes a second material composition having a second thermal tolerance greater than the first thermal tolerance. The first article and the second article are in contact with one another through the interface structure. The thermal break interrupts a thermal conduction path from the second article to the first article.

Abstract: An assembly for an aircraft propulsion system is provided, the assembly having an axial centerline. The assembly comprises a fixed structure, a first thrust reverser door, and a second thrust reverser door. The first thrust reverser door being pivotally attached to the fixed structure along a first pivot axis, and the second thrust reverser door being pivotally attached to the fixed structure along a second pivot axis. The first pivot axis and the second pivot axis are both radially located at a first distance from the assembly axial centerline. The first pivot axis is located at a first axial position, and the second pivot axis is located at a second axial position. The second axial position is displaced from the first axial position.

Abstract: A gas turbine engine generates noise during use, and one particularly important flight condition for noise generation is take-off. A gas turbine engine has high efficiency together with low noise, in particular from the bypass flow exiting the engine. The average velocity of the flow at the exit to the bypass duct is in the range of from 200 m/s to 275 m/s at a take-off lateral reference point, defined as the point on a line parallel to and 450 m from the runway centre line where the Effective Perceived Noise Level (EPNL) is a maximum during take-off.

Abstract: Provided is a two-shaft gas turbine power generation equipment that includes: an induction motor which receives/transmits power from/to a power system; and a speed reducer that makes a rotational speed of a motor rotor of the induction motor lower than a rotational speed of a compressor rotor of a two-shaft gas turbine. The speed reducer includes: a compressor-side shaft; a motor-side shaft; a first compressor-side helical gear and a second compressor-side helical gear which are mounted on the compressor-side shaft; and a first motor-side helical gear and a second motor-side helical gear which are mounted on the motor-side shaft. The first compressor-side helical gear meshes with the first motor-side helical gear, and the second compressor-side helical gear meshes with the second motor-side helical gear.

Abstract: A rotary machine includes at least one burner including a front panel having a front side and an opposing back side. The acoustic damper includes at least one wall, at least one cooling air inlet, at least one outlet, and at least one baffle. The wall extends from the back side of the front panel and defines a dampening chamber. The cooling air inlet is defined within the back side of the front panel and is configured to channel a flow of cooling air into the dampening chamber. The outlet is defined within the back side of the front panel and is configured to channel the flow of cooling air out of the dampening chamber. The baffle extends from the back side of the front panel and is configured to reduce a velocity of the flow of cooling air within the dampening chamber.

Abstract: The present disclosure generally relates to monolithic superstructures for supporting a rotating shaft coupled to a rotor relative to a stator. An integral superstructure supports the rotating component. The superstructure includes a bearing portion that contacts the shaft. A stator portion, is spaced a critical dimension radially outward, from the rotor. A first annular transfer portion extends axially forward from the bearing to the stator portion. A second annular transfer portion extends axially aft from the stator portion to a mounting flange. The mounting flange connects the superstructure to a frame. The superstructure maintains a critical dimension between the rotor and the stator as the temperature of the superstructure increases. The rotor may be a compressor impeller in a gas turbine engine and the stator may be an aero component that transfers air into a combustor.

Abstract: A turbofan engine includes a first spool including a first turbine, and a first tower shaft engaged to the first spool. A second spool includes a second turbine, and a second tower shaft is engaged to the second spool. A superposition gearbox includes a sun gear, a plurality of intermediate gears engaged to the sun gear, and is supported in a carrier and a ring gear circumscribing the intermediate gears. The first tower shaft or the second tower shaft drives one of the intermediate gears. A drive motor is engaged to drive the sun gear, an inner electric motor, a stator disposed radially outside of the inner electric motor, and an outer electric motor disposed radially outside the stator. A first load on the first spool and a second load on the second spool is adjusted by operation of at least one of the inner electric motor and the outer electric motor.

Abstract: The present disclosure is directed to a system for controlling an output of a gas generator via an operator manipulated input device. The system includes one or more sensors measuring one or more environmental conditions, a gas generator shaft speed, and a power turbine torque. The system further includes an operator manipulated input device and one or more controllers including one or more processors and one or more memory devices. The one or more memory devices stores instructions that when executed by the one or more processors cause the one or more processors to perform operations.