Abstract: The present disclosure relates to systems and methods that provide a low pressure liquid CO2 stream. In particular, the present disclosure provides systems and methods wherein a high pressure CO2 stream, such as a recycle CO2 stream from a power production process using predominately CO2 as a working fluid, can be divided such that a portion thereof can be expanded and used as a cooling stream in a heat exchanger to cool the remaining portion of the high pressure CO2 stream, which can then be expanded to form a low pressure CO2 stream, which may be in a mixed form with CO2 vapor. The systems and methods can be utilized to provide net CO2 from combustion in a liquid form that is easily transportable.

Abstract: An example hybrid aircraft propulsion system includes one or more parallel propulsion units, each of the parallel propulsion units comprising: a first propulsor; a gas turbine engine configured to drive the first propulsor; and an electrical machine selectively configurable to: generate, for output via one or more electrical busses, electrical energy using mechanical energy derived from the first propulsor or the gas turbine engine; and drive the first propulsor using electrical energy received via the one or more electrical busses; and one or more series propulsion units, each of the series propulsion units comprising: a second propulsor; and an electrical machine selectively configurable to: generate, for output via the one or more electrical busses, electrical energy using mechanical energy derived from the second propulsor or the gas turbine engine; and drive the second propulsor using electrical energy received from one or more electrical busses.

Abstract: A turbomachine comprising an engine body and an annular nacelle. The nacelle has an air intake comprising an outer skin, an inner skin, an air intake lip connecting the outer and inner skins together, and an annular frame connecting the outer and inner skins together. The turbomachine comprises a system for de-icing the air intake having an air bleeding port for bleeding a flow of hot air from the engine body, and two interconnected heating systems. An inner skin heating system receives the hot air flow to heat, by means of a gas-liquid heat exchanger, a heat transfer liquid contained in a hydraulic circuit arranged in part in the heat exchanger and in the inner skin. An air intake lip heating system diffuses the flow of hot air, once it has passed through the inner skin heating system, through the annular frame to heat the air intake lip.

Abstract: An anti-icing system is disclosed. In various embodiments, the anti-icing system includes an inner duct having a first end configured to deliver heated gas to a plenum and a second end spaced from the first end; an outer duct circumferentially encompassing at least a portion of the inner duct; and a seal system disposed proximate the second end, the seal system including a first annular seal having a radially inner end positioned proximate a flange disposed on the inner duct.

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: 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: 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 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: A heat shield panel for use in a combustor of a gas turbine engine is disclosed. In various embodiments, the heat shield panel includes an inner base layer, an outer base layer, and a grommet having a flange disposed between the inner base layer and the outer base layer.

Abstract: A cooling system configured to reduce a temperature within a gas turbine engine in a shutdown mode of operation includes a first gas turbine engine including a compressor having a bleed port. In a first operating mode of the gas turbine engine, the compressor bleed port is configured to channel a high pressure flow of air from the compressor. During a shutdown mode of operation, the compressor bleed port is configured to channel an external flow of cooling air into the compressor. The cooling system also includes a source of cooling air and a conduit coupled in flow communication between the compressor bleed port and the source of cooling air. The source of cooling air configured to deliver a flow of cooling air into the compressor through the compressor bleed port.

Abstract: A vibration isolating bushing assembly may comprise a first T-bushing configured to receive a fastener, a second T-bushing configured to couple to the first T-bushing and the fastener, a first isolating washer, and a second isolating washer, wherein the first isolating washer and the second isolating washer are configured to couple between the first T-bushing and the second T-bushing.

Abstract: The present invention discloses a system for providing mobile power, in which all equipment required for the mobile power are integrated on three transport vehicles, which can be combined for transport as a whole, greatly reducing the connection time required for field operations, with flexible combinations and wider applicability. An exhaust system can be flexibly connected to the side or the top. An air intake system, a gas turbine engine and a generator are connected at the same axial direction, avoiding unnecessary elbows, with small intake pressure losses and a stable flow direction.

Abstract: A nozzle for a combustion chamber of an engine for providing a fuel-air mixture at a nozzle exit opening of the nozzle includes a nozzle main body including the nozzle exit opening and extends along a nozzle longitudinal axis. The nozzle main body has a first air guiding channel, a fuel guiding channel and a further, radially outwardly second air guiding channel. One end of the fuel guiding channel is positioned in front of the end of the second air guiding channel. A tapering channel portion with a shell surface inclined with respect to a nozzle longitudinal axis in the axial direction and connects to a radially outer shell surface of the fuel guiding channel is formed between the end of the fuel guiding channel and the end of the second air guiding channel at the nozzle.

Abstract: A cooling arrangement provides cooling around a dilution hole defined in a liner circumscribing a combustion chamber of a gas turbine engine. The cooling arrangement comprises a hollow boss projecting from an outer surface of the liner about the dilution hole. The hollow boss defines an internal cavity extending circumferentially around the dilution hole. The internal cavity has an inlet in fluid flow communication with an air plenum surrounding the liner and an outlet in fluid flow communication with the combustion chamber.

Abstract: A gas turbine engine includes an inner annular combustor radially inboard of an outer annular combustor. An outer variable turbine vane array is downstream of the outer annular combustor and an inner variable turbine vane array downstream of the inner annular combustor.

Abstract: The present invention discloses a method of a mobile power generation system. A power generation apparatus is respectively quickly connected, through expansion joints, to an intake assembly and an exhaust duct which are separately transported, to implement the quick installation and connection of the power generation system at the fracturing operation site. Two conveyances are respectively provided for the intake assembly and the exhaust duct to achieve more flexible adjustment during connection. While the position of the power generation apparatus is fixed, the intake assembly is moved to connect to an intake chamber of the power generation apparatus, and the exhaust duct is moved to connect to an exhaust collector of the power generation apparatus.

Abstract: A method of controlling at least a part of a start-up or re-light process of a gas turbine engine, the method comprising: controlling rotation of a low pressure compressor using a first electrical machine to increase angular velocity of the low pressure compressor; and controlling rotation of a high pressure compressor using a second electrical machine to restrict angular velocity of the high pressure compressor while the angular velocity of the low pressure compressor is increased.

Abstract: An aircraft engine system has a turbofan engine with a lubricating oil system. An oil pump is connected to pump oil from the oil tank through a cooling circuit to the turbofan engine. The cooling circuit has a bleed air boosted engine oil cooler assembly with a liquid/air heat exchanger (LAHEX) connected to an oil inlet conduit and receiving fan air from a high bypass fan of the turbofan engine as the cooling working fluid. The LAHEX is connected to an oil exit conduit. An ejector downstream of the LAHEX receives bleed air from a compressor section of the turbofan engine. The ejector draws the fan air through the LAHEX.