A turbine assembly with a hollow aerofoil having a main cavity with an impingement tube, insertable inside the main cavity for impingement cooling of an inner surface of the main cavity, and a platform at a radial end of the hollow aerofoil, and a cooling chamber for cooling the platform arranged relative to the hollow aerofoil on an opposed site of the platform. The cooling chamber is limited at a first radial end by a wall segment of the platform and at an opposed radial second end from a cover plate. The impingement tube extends in span wise direction through the cooling chamber from the platform to the cover plate and restricts a sub-cavity of the main cavity. The wall segment includes an entry aperture for a cooling medium to enter from the cooling chamber of the platform into the sub-cavity of the hollow aerofoil.

A stator vane assembly for a gas turbine engine in which both higher pressure cooling air and lower pressure cooling air are both supplied to the stator vane assembly to cool both an airfoil and inner and outer diameter endwall cavities of the stator vane assembly, in which the spent higher pressure cooling air is then discharged into a combustor of the gas turbine engine. The higher pressure cooling air flows through a closed loop cooling circuit formed within the stator vane assembly while the lower pressure cooling air is discharged through exit holes into the hot gas stream of the turbine.

A gas turbine engine includes a fan located at a forward portion of the gas turbine engine, a compressor section and a turbine section arranged in serial flow order. The compressor section and the turbine section together define a core airflow path. A rotary member is rotatable with at least a portion of the compressor section and with at least a portion of the turbine section. An outlet guide vane assembly includes multiple outlet guide vanes located in an exhaust airflow path downstream of the turbine section. The multiple outlet guide vanes being spaced-apart circumferentially from each other over an angular range of about 360 degrees, and each multiple outlet guide vane defining a radial extent. At least one of the multiple outlet guide vanes includes a cold fluid passageway extending at least partially radially therethrough through which a fluid coolant flows and another of the multiple guide vanes includes a heated fluid passageway extending at least partially radially therethrough through which the fluid coolant flows and receives heat from exhaust airflow from the core airflow path.

A gas turbine has at least one rotor and inner housing part to form an annular chamber therebetween. The annular chamber is fluidically connected to a compressor portion at one end and expansion turbine portion at the other, and is supplied with cooling fluid. First and second swirl supply lines supply the annular chamber with cooling fluid. The cooling fluid is supplied to the surface of the rotor with a tangential flow component, and a first seal element in the annular chamber acts as a flow resistor. A discharge line in the rotor between the first seal element and expansion turbine portion receives and discharges cooling fluid from the second swirl supply line. No bypass lines are provided from the first swirl supply line such that the cooling fluid is conducted around the second swirl supply line in order to be returned to a location of the annular chamber.

An aircraft having an engine, a dihydrogen tank, devices to be heated, a first air intake for taking in air at a low pressure or at an intermediate pressure, a second air intake for taking in air at a high pressure, a first heat exchanger, a first pipe which passes through the first heat exchanger and feeds the devices to be heated. Upstream of the first heat exchanger, the first pipe is divided into two sub-pipes connected respectively to the first air intake and the second air intake, and a fuel pipe that is connected between the tank and the combustion chamber and passes through the first heat exchanger. The use of heat exchangers on the dihydrogen pipe allows a regulation of the temperature of the devices to be heated and of the engine and to increase the temperature of the dihydrogen before its combustion.

An inlet air chilling system for use with a gas turbine engine is disclosed. The inlet air chilling system may include an inlet air filter house, an air chilling/heating coil positioned within the inlet air filter house, and an energy recovery/heating coil positioned downstream of the air chilling/heating coil within the inlet air filter house and in communication with the air chilling/heating coil.

An embodiment of an independent cooling circuit for selectively delivering cooling fluid to a component of a gas turbine system includes: a plurality of independent circuits of cooling channels embedded within an exterior wall of the component, wherein the plurality of circuits of cooling channels are interwoven together; an impingement plate; and a plurality of feed tubes connecting the impingement plate to the exterior wall of the component and fluidly coupling each of the plurality of circuits of cooling channels to at least one supply of cooling fluid, wherein, in each of the plurality of circuits of cooling channels, the cooling fluid flows through the plurality of feed tubes into the circuit of cooling channels only in response to a formation of a breach in the exterior wall of the component that exposes at least one of the cooling channels of the circuit of cooling channels.

A cooling device for cooling platforms of a guide vane ring of a gas turbine is arranged downstream inside a main flow channel of a combustion chamber. Cooling air passages are arranged in a wall of the platforms or of an intermediate piece that is connected therewith to guide cooling air for film cooling the surfaces of the platforms. At least in certain areas, the wall is configured with at least two layers having—as viewed from the main flow channel—an outer wall and a spaced apart inner wall forming a hollow space, wherein the hollow space can be impinged by cooling air through at least one cooling air blow-in opening inside the outer wall, and at least one cooling air blow-out opening is arranged inside the inner wall extending in the downstream direction to the surfaces of the platforms.

A method for cooling a component of a gas turbine engine includes flowing a cooling airflow through a cooling passage of a turbine rotor blade, wherein the cooling passage includes an inlet and an outlet formed on a blade tip of the turbine rotor blade. The method further includes receiving at least a portion of the cooling airflow exiting the outlet of the cooling passage with an aperture defined in a casing of the gas turbine engine, wherein the casing is spaced from the blade tip along the radial direction. In addition, the method includes providing the cooling airflow received with the aperture defined in the casing to the component of the gas turbine engine through a coolant duct assembly of the gas turbine engine.

A turbine includes an inner casing to which at least a stator vane of a turbine section is mountable, and an outer casing arranged around the inner casing in such a way that an outer cooling channel is formed between the inner casing and the outer casing. The outer cooling channel includes a fluid inlet through which a cooling fluid is injectable from an outer volume of the turbine into the outer cooling channel. The cooling channel includes a fluid outlet such that the cooling fluid is exhausted into an inner volume of the turbine. The fluid inlet is located with respect to the fluid outlet such that the cooling fluid inside the outer cooling channel includes a flow direction which has a component that is orientated in opposite direction with respect to a main flow direction of a working fluid of the turbine.