Devices and methods of rocket propulsion are disclosed. In one aspect, a staged combustion liquid rocket engine with preburner and turbopump unit (TPU) integrated into the structure of the combustion chamber is described. An initial propellant mixture is combusted in a preburner combustion chamber formed as an annulus around a main combustion chamber, the combustion products from the preburner driving the turbine of the TPU and subsequently injected into the main combustion chamber for secondary combustion along with additional propellants, generating thrust through a supersonic nozzle. The preburner inner cylindrical wall is shared with the outer cylindrical wall of the engine's main combustion chamber and the turbine is axially aligned with the main combustion chamber. Liquid propellants supplied to the engine are utilized for regenerative cooling of the combustion chamber and preburner, where the liquid propellants are gasified in cooling manifolds before injection into the preburner and main combustion chamber.

A shaft bearing assembly having a shaft supported in a bearing housing is described. The shaft bearing assembly includes a pressure reduction device for generating an underpressure inside the bearing housing. The pressure reduction device is operated by employing an incompressible fluid. Further, a method of reducing a pressure inside a bearing housing supporting a shaft is described. The method includes using an incompressible fluid for operating a pressure reduction device for generating an underpressure inside the bearing housing.

The invention provides various designs of an apparatus and method for attemperating a gas stream temperature. The apparatus of the present invention provides a body through which a gas stream passes that permits, as desired, a second gas, such as gas outside of the gas duct or such as ambient air, to be added to the main gas stream to attemperate the temperature of the main gas stream. The body or device may be referred to as a variable eductor having a plurality of openings through which a second gas may pass into the main gas stream. The openings may be opened or closed, and the variable eductor provides control over which openings and the degree to which each opening is opened. In some designs the variable eductor is inserted between two portions of a gas duct. The variable eductor has widespread application, such as downstream of a gas turbine to attemperate the exhaust gas temperature during startup.

Disclosed herein are methods, control systems and ejector systems related to introducing combustible vent gas to a unit using the combustible vent gas as a fuel, and/or to a gas compressor. Certain embodiments have a combustible gas provided to the unit through a first fuel line having a control valve to control the flow of the combustible gas. A portion of the combustible gas is selectively flowed to a second fuel line having a control valve to control the flow of combustible gas in the second fuel line. The flow of combustible gas in the second fuel serves as a primary flow to an ejector. The secondary flow to the ejector is vent gas. The output from the ejector is recombined with the flow in the first fuel line. Various control methodologies are possible using the control valves, vent gas flow and/or pressure of fuel to the unit.

The disclosure provides devices for cooling a turbine casing for a turbomachine and methods of using the same. The devices extend around an axis (X) and include air-distribution means configured to take in air and convey it to the casing. The air-distribution means include a first ramp and a second ramp extending circumferentially about the axis (X), each ramp having air ejection orifices intended to be directed towards the casing in order to cool it The devices further include adjustment means capable of adjusting the flow rate of air ejected at the level of the first ramp with respect to the flow rate of air ejected at the level of the second ramp.

A jet pump for a system for supplying fluid to a turbomachine. The jet pump includes an active fluid inlet pipe including a tube delimiting the inlet pipe, and a passive fluid inlet pipe that is fluidically separated from the active fluid inlet pipe by the tube. The active fluid inlet pipe includes at least one twisted blade that is positioned within the tube and is configured to make the active fluid rotate with respect to the axis of the tube.

Power and cooling systems including a drive system, a power generation unit, and a cooled fluid generation unit. A primary working fluid that is expanded within a turbine of the drive system and compressed within compressors in a closed-loop cycle. The power generation unit includes a generator and a heat source configured to heat the primary working fluid prior to injection into the turbine. T cooled fluid generation unit includes an ejector downstream of the compressors and a separator arranged downstream of the ejector and configured to separate liquid and gaseous portions of the primary working fluid. The gaseous portion is directed to the compressors and the liquid portion is directed to an evaporator heat exchanger to generate cooled fluid.

A fuel pumping system of an aircraft, having a first boost pump; a main pump; a main pump feed conduit fluidly coupled to the main pump and configured to receive fuel after it has passed through the first boost pump; a main pump discharge conduit fluidly coupled to the main pump; an ejector fluidly coupled to the main pump feed conduit, between the first boost pump and the main pump; a bypass conduit fluidly coupled to the main pump discharge conduit and the ejector such that the bypass conduit is configured to provide fluid to the ejector drawn from the output of the first boost pump into the main pump; and a backpressure regulating valve fluidly coupled to the bypass conduit, and configured to transition to an open state when pressure in the main pump discharge conduit is above a threshold.

The invention relates to an assembly for an aircraft turbomachine, comprising a central element (1) for the ejection of gas, and a connecting flange (9) interposed between, upstream, a gas outlet (22a) made of metal for a turbomachine, and, at the downstream end, the central element (1). The connecting flange comprises an annular portion (9a) and flexible tabs (11) having axially: a first end (111a) where the tab is connected to the said annular portion, and a second free end (111b), projecting radially inwardly with respect to the first end and to which said tab is attached with the central element (1).

A gas turbine engine includes: a main passage through which compressed air extracted from a compressor is guided to a bearing; and an oil supply device that includes an oil tank storing oil and supplies the oil to the main passage through which the compressed air flows to generate oil mist. The oil supply device includes: a pressure reducer disposed at the main passage; a bypass passage that bypasses the pressure reducer and connects an upstream passage and a downstream passage; a holding chamber disposed at the bypass passage and having a predetermined volume; a switching mechanism that permits or blocks communication between the holding chamber and the oil tank, communication between the holding chamber and the upstream passage, and communication between the holding chamber and the downstream passage; and a controller that controls the switching mechanism.