Abstract: The invention relates to a method for starting a turbine engine in cold weather, including a starting system intended for rotating a drive shaft of the turbine engine, the method comprising the following steps: —a pre-starting step in which a first starting signal is generated to control the drive shaft in a first direction of rotation about a longitudinal axis (X) and in a second opposite direction of rotation in an alternating manner; and —a starting step in which a second starting signal is transmitted to the starting system in order for the latter to drive the drive shaft of the turbine engine in a normal direction of rotation and in which the drive shaft is rotated until a rotation speed that causes the turbine engine to start.

Abstract: A lance for a burner includes an innermost conduit defining a first fluid passage and a plurality of first fuel injection channels, each first fuel injection channel terminating at a first outlet; an intermediate conduit circumferentially surrounding the innermost conduit, the intermediate conduit defining a second fluid passage and a plurality of second fuel injection channels, each second fuel injection channel terminating at a second outlet; an outermost conduit circumferentially surrounding the intermediate conduit, the outermost conduit defining a third fluid passage, a plurality of third air outlets through the outermost conduit and surrounding the first outlets, a plurality of fourth air outlets through the outermost conduit and surrounding the second outlets, and a plurality of cooling microchannels; wherein each cooling microchannel includes and extends between a microchannel inlet in fluid communication with the third fluid passage and a microchannel outlet on an outer surface of the outermost condu

Abstract: A gas turbine engine that includes an inlet volume flow control appliance and methods of operating the same are provided. The method includes operating the gas turbine engine with the inlet volume flow control appliance supplying a compressor inlet volume flow that is below a maximum compressor inlet volume flow. A mass flow of a liquid agent is added to a compressor gas mass flow while the gas turbine engine is operated with a compressor inlet volume flow below a maximum compressor inlet volume flow. The mass flow of a liquid agent may be controlled as a function of the pitch of variable inlet guide vanes. The method further comprises adjusting the volume flow control appliance to increase the compressor inlet volume flow and increasing the mass flow of liquid agent added to the compressor gas mass flow while the inlet volume flow control appliance increases the compressor inlet volume.

Abstract: A fuel injector providing a flow of fuel having a variable swirl and/or variable effective area is disclosed. The fuel injector may have a nozzle defining a mixing chamber having an outlet proximate a centerline of the mixing chamber, a first fuel line in fluid communication with the mixing chamber and a second fuel line in fluid communication with the mixing chamber. The first fuel line may terminate in a tangential fuel inlet positioned tangential to the centerline of the nozzle, while the second fuel line may terminate in a radial fuel inlet positioned radial to the centerline of the nozzle.

Abstract: In a main flow passage, a first heat exchanger, a first heat storage unit, a second heat exchanger, and a second heat storage unit are connected by a heating medium flow passage. The main flow passage allows a heating medium to be circulated. A sub flow passage includes a shortened flow passage which is a part of the heating medium flow passage and branches from the heating medium flow passage between the second heat exchanger and the second heat storage unit and extends to the first heat storage unit. The sub flow passage allows circulation of the heating medium between the first heat storage unit and the second heat exchanger. A first heating means in a middle of the shortened flow passage, the first heating means heating a passing heat medium, and a switching means conducting switching between the main flow passage and the sub flow passage are provided.

Abstract: A ceramic heat shield for a gas turbine. The ceramic heat shield has a ceramic body containing aluminium oxide and has a surface layer of the ceramic body which contains yttrium aluminium garnet as reaction coating material. A gas turbine includes such a ceramic heat shield and a method produces such a ceramic heat shield.

Abstract: The invention relates to an aircraft turbine engine module (10), comprising a degassing tube (14) and an ejection cone (12). The tube and the cone comprise centring means engaging together. The invention also relates to a locating and adjusting tool for assembling said module.

Abstract: Systems and methods for controlling fuel flow to an engine during start are provided. Fuel is caused to be injected into a combustor of the engine according to a first fuel schedule defining a minimum fuel flow limit required to achieve light-off of the engine, the minimum fuel flow limit set at an initial value. Following light-off of the engine, at least one operating parameter of the engine is monitored. Based on the at least one operating parameter, occurrence of flameout in the engine is detected. In response to detecting occurrence of flameout in the engine, the minimum fuel flow limit is increased from the initial value to a first value to obtain an adjusted fuel schedule, and fuel is caused to be injected into the combustor according to the adjusted fuel schedule.

Abstract: A gas turbine engine includes a compressor section and a combustion section with a scroll, a scroll baffle, a combustor, and a combustor case. The scroll defines an interior scroll flow path. The scroll baffle surrounds the scroll to define a scroll cooling passage. The combustor case surrounds the combustor and the scroll baffle to define a collector space. Moreover, the engine includes a turbine section with a turbine rotor and a turbine rotor blade shroud that includes a shroud cooling passage. The compressor flow path is fluidly connected to the scroll for cooling the scroll. Also, the scroll cooling passage is fluidly connected to the shroud cooling passage for cooling the turbine rotor blade shroud. Furthermore, the shroud cooling passage is fluidly connected to the collector space. Flow from the collector space flows into the combustor, along the interior scroll flow path, toward the turbine rotor.

Abstract: A gas turbine engine configured to combust, by means of a combustor, a compressed air obtained through compression by a compressor, and drive a turbine with use of high-temperature and high-pressure combustion gas generated by the combustion, the gas turbine engine including: a plurality of air passages through which different portions in the gas turbine engine communicate with one another; and a switching device configured to switch air flow paths between the plurality of air passages, wherein the plurality of air passages are each formed by an air pipe disposed outside a compressor casing, and the switching device is mounted to an outer surface of the compressor casing via a heat insulation member.

Abstract: A distributed fuel module includes a fuel pressure vessel with a gas port and a fuel port, a hydraulic circuit breaker connected to the fuel port, and a gaseous circuit breaker. The gaseous circuit breaker is connected to the gas port, is fluidly coupled to the hydraulic circuit breaker through the fuel pressure vessel, and is cooperatively associated with the gaseous circuit breaker to isolate the fuel pressure vessel from a compressed gas header and a fuel header according to pressure differential within the hydraulic circuit breaker and pressure differential within the gaseous circuit breaker. Power modules and methods of controlling fuel flow in fuel modules are also described.

Abstract: A gas turbine engine component includes an exterior pressure side with a plurality of cooling holes located in the exterior pressure side. A relief cut surrounds at least one of the plurality of cooling holes.

Abstract: A combustor adapted for use in a gas turbine engine is disclosed. The combustor includes a metallic shell forming a cavity and a ceramic liner arranged in the cavity of the metallic shell. The ceramic liner defines a combustion chamber in which fuel is burned during operation of a gas turbine engine. The ceramic liner includes a plurality of ceramic tiles mounted to the metallic shell and arranged to shield the metallic shell from heat generated in the combustion chamber.

Abstract: An article includes a substrate, a bond coat layer disposed on the substrate, and a layer of networked ceramic nanofibers disposed on the bond coat layer.

Abstract: A fuel injection nozzle for a combustion turbine engine has thermal stress-relief vanes, which accommodate and relieve localized thermal stresses within its monolithic, three-dimensional nozzle structure, imparted by heat transfer during engine combustion. At least one first vane is coupled to opposing, spaced nozzle sleeves at both ends. At least one cantilever-like second vane is coupled to one of the opposing sleeves on one end, while the other free or floating end is spaced by a second vane gap from the other opposing sleeve. Some embodiments include a plurality of second vanes, which have locally varying orientation, and/or structure, and/or second vane gaps, for normalizing spatially and/or temporally thermal stresses within the nozzle structure. The monolithic structure is fabricated, in some nozzle embodiments, by additive manufacturing.

Abstract: An aircraft turbine engine assembly including a nacelle structure with a strut attachment zone, a first cover forming an air intake lip, a hinged cover arranged on the nacelle upper portion, between the first cover and the attachment zone, a hinge system of the cover comprising, on either side of a mid-plane of the nacelle, a front rod and a rear rod that are one in front of the other. The two ends of each rod are co-planar, parallel to the mid-plane. A first end of each rod is hinged on the cover. A second end of each rod is hinged on the structure. A locking system is actuatable from the outside of the nacelle between a locked and an unlocked positon. The nacelle comprises at least one cylinder with a first end mounted hinged on the structure and the other end mounted hinged on the cover.

Abstract: Systems and methods for preventing fuel leakage in a gas turbine engine are provided. A fuel accumulation system includes a control valve section fluidly coupled to a fuel manifold passage and an accumulator valve section fluidly coupled at a first side to the control valve section. The control valve section is configured to control expansion of a fluid flowing in the fuel manifold passage. The accumulator valve section is configured to receive fluid expanded in the fuel manifold passage via the control valve section.

Abstract: A turbomachine assembly for an aircraft comprising a nacelle with a nacelle structure, cowls attached to the structure to create an aerodynamic surface, an articulated cowl, comprising four edges, in the upper portion of the nacelle, an articulation at a first edge of the articulated cowl attached between the articulated cowl and the nacelle structure allowing the articulated cowl to move between a closed and an open position, a locking system at a second edge of the articulated cowl, opposite the first edge. The locking system comprises a handle actuable from the exterior of the nacelle, between a first position and a second position, a locking device movable between a locking position and an unlocking position, and a transmission system that moves the locking device from the locking position to the unlocking position when the handle passes from the first position to the second position, and vice versa.

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.