Abstract: In one embodiment, a plant control apparatus controls a power plant. The plant includes a gas turbine, a generator driven by the gas turbine, an exhaust heat recovering boiler to generate first steam by using heat of exhaust gas from the gas turbine, a first steam turbine driven by the first steam, a reheater provided in the boiler and configured to heat exhaust steam from the first steam turbine by the exhaust gas to generate reheat steam, and a second steam turbine driven by the reheat steam. The apparatus includes a first warming-up module to supply second steam from equipment different from the boiler to the first steam turbine to warm up this turbine, before this turbine is started. The apparatus further includes a second warming-up module to supply the second steam to the reheater to warm up the reheater, before the first steam turbine is started.

Abstract: A method for cooling a hybrid electric aircraft propulsion system comprises transporting a coolant through a two-phase pumped loop (TPPL) in thermal contact with an electrical machine and a plurality of power modules to be cooled, where the TPPL includes: a parallel arrangement of cold plates; an evaporator; a condenser; a first control valve; a liquid receiver; and a pump. A sensor positioned upstream of the cold plates, and in some cases upstream of the liquid receiver, measures pressure and/or temperature of a return stream of the coolant and transmits measurement data to a first controller electrically connected to the first control valve. The first controller regulates flow of a first liquid stream through the first control valve based on the pressure and/or temperature measured by the sensor, thereby keeping the return stream at a temperature within a predetermined temperature range.

Abstract: A nacelle of a turbine engine has a substantially tubular structure and defines a secondary channel. The nacelle includes a shroud forming a trailing edge on a downstream end of the nacelle, the shroud being defined by at least one outer face subjected to an outside air flow and an inner face subjected to an inside air flow circulating in the secondary channel of the nacelle. The nacelle further includes a cooling device having at least one heat-exchange channel that extends into the shroud.

Abstract: An acoustic device includes: a perforated plate that has a plurality of holes penetrating in a plate thickness direction of the perforated plate and in which a main flow is to flow on a first side of the perforated plate in the plate thickness direction; and a housing that is on a second side of the perforated plate in the plate thickness direction and partitions a space between the housing and the perforated plate, wherein a part of each of the plurality of holes on the first side in the thickness direction is inclined to at least one of the first side and a second side of a flow direction of the main flow.

Abstract: A gas turbine control device includes a first estimation unit configured to estimate a first temperature which is a first turbine inlet temperature estimation value based on a first model which is a physical model using a fuel flow rate to a gas turbine; a second estimation unit configured to estimate a second temperature which is a second turbine inlet temperature estimation value based on a second model which is a physical model using an exhaust gas temperature of the gas turbine; and a correction unit configured to correct the first temperature based on the second temperature to calculate a third turbine inlet temperature estimation value.

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: Systems and methods for conditioning a fluid using bleed air from a bypass duct of a turbofan engine are disclosed. The system comprises a heat exchanger configured to facilitate heat transfer between a flow of bleed air from the bypass duct of the turbofan engine and the fluid, and a fluid propeller configured to drive the bleed air through the heat exchanger. The fluid propeller is disposed downstream of the heat exchanger.

Abstract: The present disclosure relates to systems and methods that are useful for controlling one or more aspects of a power production plant. More particularly, the disclosure relates to power production plants and methods of carrying out a power production method utilizing different fuel chemistries. Combustion of the different fuel mixtures can be controlled so that a defined set of combustion characteristics remains substantially constant across a range of different fuel chemistries.

Abstract: A small gas turbine engine, such as is used to power a UAV, that includes at least one centrifugal compressor having an impeller with blades that form a gap between the blade tips and stationary shroud of the gas turbine engine, and where a resistance heating element is secured to or bonded to a compressor casing of the gas turbine engine in order to use heat to control the gap between the impeller blades and the stationary shroud. The resistance heating element is activated at cruise mode to move the shroud toward the impeller. Additionally or alternatively, the compressor casing is heated with bled-off compressed air to move the shroud toward the impeller. A capacitance tip clearance sensor can be mounted on the impeller shroud to monitor and control tip clearance in real time.

Abstract: A gas turbine engine coupled to an aircraft includes an engine core arranged axially along an axis, a bypass duct arranged circumferentially around the engine core to define a bypass channel, and a heat exchanger system. The bypass channel is arranged to conduct bypass air around the engine core to provide thrust for the gas turbine engine. The heat exchanger system is configured to provide cooling for the engine core.

Abstract: A high-efficiency power generation system includes: a combustor configured to generate a circulating fluid by burning a fuel; an expander configured to generate power by expanding the circulating fluid; a power generator configured to generate electricity using the power generated by the expander; a compressor configured to compress the expanded circulating fluid; a pump configured to circulate the compressed circulating fluid; a heat exchanger configured to allow the expanded circulating fluid passing through the expander and the compressed circulating fluid passing through the compressor to exchange heat with each other; and a power transmitter including a driving shaft, and configured to rotate a driven shaft, which includes shafts of the compressor and the pump, to transmit the power generated by the expander to the compressor and the pump.

Abstract: A gas turbine engine component assembly including: a first component having a first surface and a second surface opposite the first surface; and a second component having a first surface, a second surface opposite the first surface of the second component, and a plurality of pin fins extending away from the second surface of the second component, the first surface of the first component and the second surface of the second component defining a cooling channel therebetween, wherein the plurality of pin fins extend into the cooling channel, wherein each of the plurality of pin fins have a pointed ellipse shape.

Abstract: A method for reducing NOx emissions in a gas turbine in which a flow of primary air and a flow of fuel are fed into a dual annular counter rotating swirler, the primary air flow being fed into the inner and outer annular chambers, wherein the method comprises the step of injecting the flow of fuel into the inner annular chamber; another embodiment is a gas turbine air fuel mixer comprising a dual annular counter rotating swirler comprising a fuel supplying element adapted to supplying fuel inside the inner chamber of the swirler; another embodiment is a gas turbine provided by such air fuel mixer.

Abstract: An air-inlet duct includes an outer wall, an inner wall, and a splitter. The splitter cooperates with the outer wall and the inner wall to establish a particle separator which separates particles entrained in an inlet flow moving through the air-inlet duct to provide a clean flow of air to a compressor section of a gas turbine engine.

Abstract: A fan casing assembly for a turbine engine including a fan casing having a peripheral wall, a surface cooler having a first surface confronting the peripheral wall and a mounting assembly having a lower portion generally retained between the first surface of the annular surface cooler and the peripheral wall and having a floating body extending through the set of fan casing fastener openings and an upper portion located radially exterior of the annular fan casing and operably coupled to the lower portion.

Abstract: An exhaust device for a combined cycle power plant, includes a diffuser and a plenum. The diffuser includes a first wall, a second wall, a diffuser inlet, a diffuser outlet, and a diffuser flow path. The first and second walls extend circumferentially about a centerline axis of the exhaust device. The second wall is spaced from the first wall. The diffuser flow path is defined between the first and second walls, and extends from the diffuser inlet to outlet. The plenum includes an inlet wall portion and a non-circular plenum outlet, where the inlet wall portion is coupled to the diffuser outlet. The non-circular plenum outlet is spaced from the diffuser outlet along an axial direction of the exhaust device.

Abstract: The invention relates to an aircraft comprising a fuselage (1), which is propelled by a turbine engine with two coaxial fans, namely an upstream fan (7) and a downstream fan (8), driven by two contra-rotating rotors (5, 6) of a power turbine (3). The two fans (7, 8) and the turbine (3) are integrated into a nacelle (14) which projects downstream from the fuselage (1) and through which air flows. According to the invention, at least one of the fans (7, 8) of the aircraft and, in particular, the downstream fan (8) comprises variable-spacing blades, and at least one stator-forming variable-spacing blade ring (25) in the aircraft is placed upstream of the upstream fan (7). The variable-spacing stator blades (25) and the variable-spacing blades of the downstream fan (8) are mutually configured to direct the air flow in a first mode in which the air flows through the nacelle (14) from upstream to downstream and in a second mode in which the air is pushed back upstream through the nacelle (14).

Abstract: A fuel metering system for a combustion section of a turbo machine is provided. The turbo machine includes a main fuel line configured to provide a flow of fuel and a zone fuel line split from the main fuel line through which at least a portion of the flow of fuel is provided. A fuel valve is disposed at the zone fuel line and is configured to obtain and receive a present fuel valve area value and a present valve position value. A first pressure sensor is disposed upstream of the fuel valve, in which the first pressure sensor is configured to obtain a first pressure value. A second pressure sensor is disposed downstream of the fuel valve, in which the second pressure sensor is configured to obtain a second pressure value. A flow meter is disposed downstream of the fuel valve.

Abstract: Fuel-air heat exchange system and methods thereof for use in a gas turbine engine. The fuel-air heat exchanger allows heat transfer between a flow of cooling air used to cool components of the engine and a flow of fuel used to drive the engine.

Abstract: An injection apparatus for a rocket engine comprises an injection plate which delimits a combustion chamber upstream. The injection apparatus further comprises a plurality of coaxial injection elements distributed in the injection plate, each of which injection elements forms an inner outlet opening, delimited by a central sleeve body, for a first propellant component and an outer outlet opening for a second propellant component. The outer outlet opening is formed between the central sleeve body and a wall section which surrounds the central sleeve body in an annular manner. At least in a partial number of the injection elements, the central sleeve body projects from the wall section in the direction towards the combustion chamber. In particular, the central sleeve body projects relative to the injection plate into the combustion chamber.