A stator temperature control system for a gas turbine engine is provided. The stator temperature control system includes a casing circumferentially surrounding a stator assembly, the casing having a top portion and a bottom portion; an air source having an inlet and an outlet; and a supply line in fluid communication with the outlet of the air source and the bottom portion of the casing, wherein the bottom portion of the casing receives a flow of air from the air source via the supply line to increase a temperature of the bottom portion of the casing.

A heat exchanger assembly that in a preferred embodiment comprises: an inlet duct lower wall interfacing with a bypass duct; an outlet duct lower wall interfacing with a bypass duct; a heat exchanger coupled between the inlet duct lower wall and the outlet duct lower wall wherein the heat exchanger is at a compound angle with respect to an inlet duct air flow direction; and a fairing coupled to the top of the heat exchanger wherein the fairing forms the inlet duct upper wall and the outlet duct upper wall.

A cooling system configured to reduce a temperature within a gas turbine engine in a shutdown mode of operation includes a first gas turbine engine including a compressor having a bleed port. In a first operating mode of the gas turbine engine, the compressor bleed port is configured to channel a high pressure flow of air from the compressor. During a shutdown mode of operation, the compressor bleed port is configured to channel an external flow of cooling air into the compressor. The cooling system also includes a source of cooling air and a conduit coupled in flow communication between the compressor bleed port and the source of cooling air. The source of cooling air configured to deliver a flow of cooling air into the compressor through the compressor bleed port.

A geothermal turbine includes: a rotor; a casing which houses the rotor; a plurality of rotor blades disposed around the rotor; a plurality of stationary vanes supported on the casing; at least one seal portion disposed in a gap between the rotor and the casing at an upstream side of a first-stage rotor blade of the plurality of rotor blades so as to seal leakage steam which flows out inward in a radial direction of the rotor from between a first-stage stationary vane of the plurality of stationary vanes and the first-stage rotor blade; and a steam passage configured to extract a part of steam after passing the first-stage stationary vane and discharge the part of steam to the gap through an interior of the first-stage stationary vane.

A system is provided, including a heat management system. The heat management system includes a thermal delivery system configured to providing heating, cooling, or a combination thereof, to a first zone of a turbomachinery, and a controller operatively coupled to the thermal delivery system and configured to control the heating, the cooling, or the combination thereof, of the first zone, to minimize or to eliminate positional changes, structural changes, or a combination thereof, in one or more components of the turbomachinery due to thermal energy.

A gas-circulation system for conditioning a disk of an aircraft may comprise a first takeoff port configured to extract a combusted gas and a second takeoff port configured to extract an uncombusted gas. A first valve may comprise an inlet in fluid communication with the first and second takeoff ports and an outlet of the first valve in fluid communication with the disk.

An airfoil has a leading end, a trailing end, and first and second sides that join the leading end and the trailing end. A rib extends from the first side to the second side and partitions an internal core cavity into a forward cavity and an aft cavity. A baffle is disposed in the forward cavity and has a showerhead array of impingement orifices adjacent the leading end of the airfoil wall. A cooling passage network is embedded in the airfoil wall between inner and outer portions of the airfoil wall. The cooling passage network has an inlet orifice through the inner portion of the airfoil wall, an array of pedestals, and at least one outlet orifice through the outer portion. The inlet orifice opens to the forward cavity at a location aft of the showerhead array of impingement orifices.

Apparatus for selective delivery of pressurised gas comprises a containment wall which at least partially bounds of a volume of pressurised gas during use of the apparatus, and a valve directly coupled to an output port of the containment wall such that the containment wall and the valve retain the pressurised gas when the valve is closed. The apparatus obviates the need for a duct coupling the output port to the input of the valve, and hence reduces the number of interfaces between the output port and the input of the valve from two, as in arrangements of the prior art, to one. The apparatus presents fewer failures modes compared to the case where a duct couples the output port to the valve.

A hybrid propulsion system is provided. The system may comprise a gas turbine engine and a secondary engine, an inlet, an exhaust, a pressurized tank, and an expansion valve. The inlet may be in fluid communication with the ambient environment. The gas turbine engine may have a core passage including a compressor, a combustion chamber, and a turbine. The core passage may be in selective fluid communication with the inlet. The exhaust may be in fluid communication with the ambient environment and the core passage. The pressurized tank may be located upstream of the core passage. The pressurized tank may contain a cooling fluid. The expansion valve may be in fluid communication with the pressurized tank and the core passage. The pressurized tank may provide cooling fluid to the core passage to cool the gas turbine engine during operation of the secondary engine.

An air turbine starter device includes a gear assembly, a rotor arranged in a cavity of a housing and operably connected the gear assembly, a first manifold having a cavity with a first manifold port operative to direct compressed air to the rotor, and a second manifold having a cavity with a second manifold port operative to direct compressed air to the rotor. The first manifold is larger than the second manifold, the second manifold is fluidly connected in parallel with the first manifold, and the first manifold port and the second manifold port are operative to drive the rotor in a common direction for starting a gas turbine connected to the gear assembly. Air turbine starter systems and methods of starting gas turbine engines are also described.