Abstract: A thermal management system for a gas turbine engine and/or an aircraft is provided including a thermal transport bus having a heat exchange fluid flowing therethrough. The thermal management system also includes a plurality of heat source exchangers and at least one heat sink exchanger. The plurality of heat source exchangers and the at least one heat sink exchanger are in thermal communication with the heat exchange fluid in the thermal transport bus. The plurality of heat source exchangers are arranged along the thermal transport bus and configured to transfer heat from one or more accessory systems to the heat exchange fluid, and the at least one heat sink exchanger is located downstream of the plurality of heat source exchangers and configured to remove heat from the heat exchange fluid.

Abstract: An assembly for a combustion chamber of a turbomachine. The assembly comprises an annular shell extending along a longitudinal axis, the shell having an inner wall intended to be turned towards a furnace of the combustion chamber and an outer wall opposite the inner wall and a projecting element extending radially from an area of the outer wall. The shell has at least one deflector projecting from the outer wall and located downstream of the projecting element with respect to a direction of gas flow along the outer wall, oriented along the longitudinal axis. The deflector extends circumferentially. Perforations are provided axially between the projecting element and the or each deflector.

Abstract: An exhaust nozzle for use with a gas turbine engine includes an outer shroud and a nozzle-plug assembly coupled to the outer shroud. The nozzle-plug assembly includes an inner plug and at least one support vane that is coupled to the outer shroud to support the inner plug in an exhaust nozzle flow path.

Abstract: A system includes an engine case. A heat exchanger is included inside the engine case. The heat exchanger includes an air passage and a fuel passage. The air passage and fuel passage are in fluid isolation from one another, but are in thermal communication with one another for exchange of heat.

Abstract: A variable exhaust nozzle for a gas turbine engine includes an outer shroud and an inner plug. The outer shroud is arranged circumferentially about an axis. The inner plug extends along the axis and is at least partially located within the outer shroud. At least one of the outer shroud and the inner plug are movable selectively to cause a variable area region of the variable exhaust nozzle to change.

Abstract: A thermal management system is provided for incorporation into at least one of a gas turbine engine or an aircraft. The thermal management system includes: a thermal transport bus having a heat exchange fluid flowing therethrough, the thermal transport bus further including: a first flow loop comprising at least one first heat source exchanger, at least one first heat sink exchanger, and a first pump to move the heat exchange fluid through the first flow loop; and a second flow loop comprising at least one second heat source exchanger, at least one second heat sink exchanger, and a second pump to move the heat exchange fluid through the second flow loop; wherein the first flow loop is isolated from the second flow loop, and wherein the first heat source exchanger and the second heat source exchanger are configured with a common heat source.

Abstract: Combustor assemblies and methods for assembling combustor assemblies are provided. For example, a combustor assembly comprises an annular inner liner and an annular outer linear, each extending generally along an axial direction. The outer liner includes an outer flange extending forward from its upstream end. The combustor assembly also comprises a combustor dome extending between an inner liner upstream end and the outer liner upstream end and including an inner flange extending forward from a radially outermost end of the combustor dome. The inner liner, outer liner, and combustor dome define a combustion chamber therebetween, and the combustor dome and a portion of the outer liner together define an annular cavity of the combustion chamber. The inner and outer flanges define an airflow opening therebetween, and a chute member is positioned within the airflow opening to define an air chute for providing a flow of air to the annular cavity.

Abstract: A system for regulating a thermal parameter associated with a heat exchange assembly of a turbomachine includes at least one device for measuring or estimating the thermal parameter, a controlled valve acting on the flow rate of a fluid in the assembly, and a regulator including a comparator determining the sign of an error signal relating to the difference between the thermal parameter and a setpoint. The regulator further includes a switch configured to deliver, based on the sign, to a valve control system, a maximum control current switching the valve in a closed state or a minimum control current switching the valve in an open state, and a phase-shifter configured to apply on the error signal a phase advance determined from an estimate of a time constant of the control system.

Abstract: A fluid mixing apparatus includes mixing conduits that extend through a fluid plenum. The fluid plenum, which surrounds a first wall defining a main passage fluidly coupled to a low-pressure fluid source, is surrounded itself by a second wall defining a high-pressure plenum fluidly coupled to a high-pressure fluid source. An insulated tube disposed at the inlet of the first wall delivers a third fluid. The mixing conduits fluidly couple the high-pressure plenum to the main passage, where the high-pressure fluid is mixed with low-pressure fluid and the third fluid. Optionally, the fluid plenum may house a fourth fluid that is injected through injection holes in the mixing conduits. The fluid mixing apparatus may be used to mix one or more fuels with high- and low-pressure air in a gas turbine combustor. Alternately, the fluid mixing apparatus may mix a fluid with high- and low-pressure water streams.

Abstract: An exhaust nozzle of a gas turbine engine includes an outer nozzle wall, a centerbody arranged in a flow channel, and two struts connecting the centerbody to the wall. A first strut is connected to the wall by a first connection allowing movement of the strut relative to the nozzle wall in the axial direction. Another strut is connected to the wall by a second connection allowing movement of the strut relative to the wall in the radial and axial directions. The second connection is formed by a sliding element and a receiving slot, wherein the sliding element includes an interaction zone interacting with an actuator for axial movement. The interaction zone has a radial length such that the interaction between the actuator and the interaction zone is maintained when the sliding element is moved in the radial direction by radial thermal expansion of the strut and/or the centerbody.

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan including an array of fan blades rotatable about an engine axis, a compressor including a first compressor section and a second compressor section, the second compressor section including a second compressor section inlet with a compressor inlet annulus area, a fan duct including a fan duct annulus area outboard of the second compressor section inlet, and a turbine having a first turbine section driving the first compressor section, a second turbine section driving the fan through an epicyclic gearbox, the second turbine section including blades and vanes, and wherein the second turbine section defines a maximum gas path radius and the fan blades define a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.6.

Abstract: A nacelle for an aircraft having a thrust reverser including a translating cowl and a blocking panel. The blocking panel is connected by a pivot connection and is connected to the translating cowl by a sliding connection including an engagement member slidingly engaged in a track, the track having a first portion parallel to the longitudinal axis of the outer casing and a second portion non-parallel to the longitudinal axis. Motion of the translating cowl along the longitudinal axis slides the engagement member within the track, and is performed without moving the blocking panel when the engagement member is within the first portion of the track. The motion of the translating cowl causes the blocking panel to pivot about the pivot connection when the engagement member is within the second portion of the track.

Abstract: Method for optimizing the limitation of dust emissions from a gas turbine or combustion plant comprising a line for supplying liquid fuel oil, a line for generating fuel oil atomizing air, and a central controller, wherein: a first definition step, starting from a nominal temperature of the fuel oil and a nominal pressure ratio of the atomizing air of the fuel oil, and by controlling the injection of the soot inhibitor, of a nominal operating point corresponding to the maximum permissible level of emitted dust; a second step of controlling a first parameter, taken from the group of the fuel oil temperature and the pressure ratio of the fuel oil atomizing air, in order to reach another operating point; and a third step of controlling the soot inhibitor injection to achieve the maximum permissible level of emitted dust.

Abstract: In a cooling system of an aircraft turbojet engine surrounded by a nacelle, a heat-transfer fluid is circulated that is less flammable than a lubricant (H) for the turbojet engine and liquid at temperatures between ?70° C. and +175° C., the heat-transfer fluid (C) being intended to circulate in a closed-circuit circulation duct including a first surface heat-exchanger between the heat-transfer fluid (C) and air (F), at the level of at least one of the outer fairing and the inner fairing of the nacelle, and a second surface heat-exchanger between the heat-transfer fluid (C) and the lubricant (H).

Abstract: A gas turbine engine including a fan, a core including at least one rotatable shaft, and a gearbox mechanically coupling at least one rotatable shaft of the core to the fan is provided. The gas turbine engine also includes a coupling system for mounting the gearbox within the gas turbine engine. The coupling system includes a flexible coupling connected to at least one of a fan frame or a core frame, as well as a torque frame connected to the flexible coupling and the gearbox. Moreover, a deflection limiter is provided, loosely attaching the flexible coupling to the gearbox to provide a predetermined axial range of motion, radial range of motion, and circumferential range of motion between the gearbox and the frame to which the flexible coupling is attached.

Abstract: A turbomachine includes a combustion chamber with a radially extending downstream flange at its downstream end. A distributor is disposed downstream of the combustion chamber and includes a platform from which at least one vane extends radially. The platform has an upstream edge extending radially and delimiting, with the downstream flange disposed opposite, a gap opening into the combustion chamber at its radially inner end and closed at its radially outer end by sealing means fixed to the distributor. The downstream flange of the combustion chamber has at least one rectilinear cooling orifice passing through the flange and opening into the gap opposite the platform of the distributor.

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan including a circumferential array of fan blades, a low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area, and a fan duct annulus area outboard of the low pressure compressor section inlet, and a fan drive turbine section. The fan drive turbine section includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is equal to or greater than 0.35, and is less than 0.55.

Abstract: A case for a gas turbine engine includes a case wall and a boss that extends from the case wall. The boss includes a perimeter step.

Abstract: An igniting device for igniting a mixture, in particular for an engine, comprises an energy converting device and a fluid flow injecting device. The energy converting device is configured for converting fluid flow energy of at least one fluid flow into heat, thereby igniting the mixture. The energy converting device comprises an ignition chamber for the at least one fluid flow. The fluid injecting device is configured for injecting a plurality of fluid flows into the ignition chamber. The injection takes place such that a first fluid flow is injected into the ignition chamber with a higher fluid flow velocity than a second fluid flow.

Abstract: A disclosed turbofan engine includes a gas generator section for generating a gas stream flow and a propulsor section for generating propulsive thrust as a mass flow rate of air through a bypass flow path. The propulsor section includes a fan driven by a power turbine through a speed reduction device at a second rotational speed lower than a first rotational speed of the power turbine. An Engine Unit Thrust Parameter (“EUTP”) defined as net engine thrust divided by a product of the mass flow rate of air through the bypass flow path, a tip diameter of the fan and the first rotational speed of the power turbine is between 0.05 and 0.13 during operation of the turbofan engine.