Abstract: Adhesion of particulate matters to the burner accompanying combustion in a lean-combustion gas turbine combustor is suppressed, and the structural reliability is improved. In a gas turbine combustor including: a tubular liner that forms a combustion chamber; and a burner including an air hole plate that is arranged at an inlet of the liner and includes a plurality of air holes for guiding compressed air to the combustion chamber, and a plurality of fuel nozzles that are arranged on a side opposite to the combustion chamber with the air hole plate being sandwiched therebetween, the plurality of fuel nozzles each injecting a fuel toward a corresponding air hole, the air holes and the fuel nozzles forming a plurality of concentric annular lines, a plurality of small holes having opening diameters smaller than those of the air holes are provided through the air hole plate such that the plurality of small holes are positioned in an inner area of an innermost annular line of the air holes.

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 fuel injector including: a plurality of air swirler passages; at least one fuel supply passage arranged to supply fuel into at least one of the air swirler passages; and at least one cavity separating an exterior of the fuel supply passage from a body of the fuel injector; wherein the cavity is at least partially filled with a thermally insulating material.

Abstract: An inertial particle separator, having: an inlet duct defining an intake; an intermediate duct extending from the inlet duct to an engine inlet; a bypass duct in fluid communication with and extending downstream from the inlet duct, the bypass duct defining an outlet communicating with the environment of the aircraft engine, a splitter defined at an intersection of a wall of the bypass duct and a wall of the intermediate duct; a splitter vane within the intermediate duct and having a leading edge located upstream of the splitter relative to a flow circulating through the separator, the splitter vane and the wall of the intermediate duct defining a channel therebetween; and a porous plate extending across the channel and defining openings sized so as to aggregate ice and be blocked by ice under icing conditions.

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 combustor assembly for a heat engine is generally provided. The combustor assembly includes a liner wall defining a combustion chamber, and a deflector assembly. The deflector assembly includes a radially extended first wall disposed adjacent to the combustion chamber, and further an axially extended second wall disposed forward of the first wall and adjacent thereto. The second wall is coupled to the liner wall.

Abstract: A combustion chamber for a gas turbomachine. The combustion chamber comprising inner and outer walls, a chamber bottom, and a heat shield arranged downstream of the chamber bottom, to protect it thermally. The inner and outer walls and the heat shield form a one-piece unit.

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: A resonant sound absorbing device of a gas turbine combustor includes a plurality of resonance chambers independently disposed side by side in an axial direction of the gas turbine combustor so as to communicate with a gas passage of the gas turbine combustor via acoustic holes. The plurality of resonance chambers include n related resonance chambers each satisfying: 0.9 × ? i = 1 n ? F i n ? F i ? 1.1 × ? i = 1 n ? F i n ( A ) where n is an integer of 2 or more, and Fi is a peak frequency corresponding to a maximum sound absorbing ratio of the ith related resonance chamber of the n related resonance chambers.

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: The invention refers to a method and a device for flame stabilization in a burner system of a stationary combustion engine, preferably a stationary gas turbine, in which a flow of an air/fuel mixture is produced and being swirled to form a vortex flow to which a swirl number is assignable before entering a combustion zone in which the vortex flow of the air/fuel mixture is ignited to form a flame within a reverse flow zone caused by vortex breakdown. The swirl number perturbation driven by thermoacoustic oscillation inside the burner system is controlled by affecting the vortex flow actively before entering the combustion zone on basis of changing a flame transfer function assigned to the burner system with the proviso of minimizing pulsation amplitudes of the flame transfer function.

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: A system and method for adaptively controlling a cooldown cycle of an auxiliary power unit (APU) that is operating and rotating at a rotational speed includes reducing the rotational speed of the APU to a predetermined cooldown speed magnitude that ensures combustor inlet temperature has reached a predetermined temperature value, determining, based on one or more of operational parameters of the APU, when a lean blowout of the APU is either imminent or has occurred, and when a lean blowout is imminent or has occurred, varying one or more parameters associated with the shutdown/cooldown cycle.

Abstract: A mechanical connection system is provided that includes a ceramic skin panel, at least one bushing, and at least one bolt. The ceramic skin panel includes a first surface, a second surface, and at least one aperture extending between the first surface and the second surface. The aperture has an axial centerline, a chamfered portion contiguous with the first surface, and an axial portion contiguous with the second surface. The chamfered portion includes a tapered surface disposed at a first angle such that a line extending along the tapered surface intersects the axial centerline at a plane of the second surface. The bushing has a housing and a collar. The housing is configured to mate with the chamfered portion of the aperture disposed within the ceramic skin panel. The collar is configured for engagement with the housing.

Abstract: A duct for a gas turbine engine, the duct has a wall defining a flow passage for an airflow, the wall defining a baseline surface, the duct comprising an off-take port for drawing a portion of the airflow out of the duct, the off-take port including a projected portion projecting away from the baseline surface into the flow passage, a conduit opening defined within the projected portion for receiving the portion of the airflow, and an off-take conduit communicating with the conduit opening for directing the portion of the airflow away from the duct, the projected portion having a hump upstream of the conduit opening relative to a direction of the airflow and a scoop downstream of the opening relative to the direction of the airflow.

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 thrust reverser for a turbojet engine nacelle includes thrust reversal cascade vanes arranged about an annular flow duct having a stationary front casing, at least one movable cover moving backwards into a thrust reversal open position, and tilting flaps connected by rear pivots to the movable cover, which, in an open position, tilt and thereby at least partially close the annular flow duct. The cascade vanes move backwards with the movable cover, in that, in the closed position, the tilting flaps and the cascade vanes are arranged outside the stationary casing, and each tilting flap is connected to the stationary casing by slides which cause the tilting flaps to tilt when the movable cover moves backwards.

Abstract: A gas turbine engine includes a main compressor section, a combustor, and a main turbine section. A fuel pump delivers fuel to the combustor. A tap taps air compressed by the main compressor section, and is connected for delivering the tapped air through a first heat exchanger and to a boost compressor. Air downstream of the boost compressor is connected to cool a component. Driving compressed air is connected to be delivered to a power turbine. The power turbine is connected to drive both the boost compressor and the fuel pump.

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.