The invention relates to a de-icing device designed to supply de-icing air to a turbomachine separation nozzle extending along a longitudinal axis, the turbomachine comprising:—the separation nozzle which is designed to be positioned downstream from a turbomachine fan and comprises an internal casing and an external casing which form a separation between a primary flow vein for a primary stream and a secondary flow vein for a secondary stream, said streams issuing from the fan, the internal casing and the external casing defining an inter-vein space;—turbomachine guide vanes designed to be secured by screws to the internal casing such that the screws extend into the inter-vein space, the de-icing device being intended to be positioned in the inter-vein space and comprising—an air inlet;—an air outlet;—a plurality of channels extending from the air inlet toward the air outlet; the channels being arranged in relation to one another such that they are designed to extend from the air inlet toward the air outlets, passing between the screws for securing the guide vanes.

An aircraft fuel system can include a fuel line configured to transport fuel therein, an exposed aircraft structure in direct or indirect thermal communication with the fuel in the fuel line to receive heat from the fuel to provide a deicing or anti-icing heat to the exposed aircraft structure. The exposed aircraft structure can include at least one internal fuel channel in fluid communication with the fuel line for direct thermal communication with the fuel. The system can include a fuel/fluid heat exchanger in fluid communication with the fuel line to transfer heat from the fuel to a fluid to provide indirect thermal communication between the fuel and the exposed aircraft structure.

An aircraft fuel system can include a fuel line configured to transport fuel therein, an exposed aircraft structure in direct or indirect thermal communication with the fuel in the fuel line to receive heat from the fuel to provide a deicing or anti-icing heat to the exposed aircraft structure. The exposed aircraft structure can include at least one internal fuel channel in fluid communication with the fuel line for direct thermal communication with the fuel. The system can include a fuel/fluid heat exchanger in fluid communication with the fuel line to transfer heat from the fuel to a fluid to provide indirect thermal communication between the fuel and the exposed aircraft structure.

The invention relates to a nozzle (1) for separating a primary flow from a secondary flow in a double-flow turbomachine. The separation nozzle comprises an outer annular wall (12), an inner annular wall (13), a radial annular wall (14) and an inner annular baffle (16). and defines a first cavity (17) between the outer annular wall (12) and the inner annular baffle (16), and a second cavity (18) between the inner annular wall (13), the radial annular wall (14) and the inner annular baffle (16).

The invention relates to a nozzle (1) for separating a primary flow from a secondary flow in a double-flow turbomachine. The separation nozzle comprises an outer annular wall (12), an inner annular wall (13), a radial annular wall (14) and an inner annular baffle (16). and defines a first cavity (17) between the outer annular wall (12) and the inner annular baffle (16), and a second cavity (18) between the inner annular wall (13), the radial annular wall (14) and the inner annular baffle (16).

A system for lubricating components of a gas turbine engine includes a supply pump to pump lubricant from a reservoir through a supply conduit to lubricated component. An outer body deicing circuit is fluidly coupled in series with the supply conduit. Moreover, a scavenge conduit extends from the lubricated component to the reservoir and a scavenge pump pumps the lubricant from the lubricated component through the scavenge conduit. In addition, a valve is fluidly coupled in series with the supply conduit and the scavenge conduit and a bypass conduit extends from the valve to a location on the supply conduit downstream of the deicing circuit. When the lubricant is supplied to the valve, the valve directs the lubricant flowing through the scavenge conduit into the reservoir. When the supply of lubricant to the valve is halted, the valve directs the lubricant flowing through the scavenge conduit into the bypass conduit.

A fluid valve is provided including an inner shell and an outer shell. The inner shell includes a sidewall having a first opening and an interior surface defining an inner chamber. The outer shell includes a sidewall having a second opening and an exterior surface defining an outer chamber. The inner shell is positioned within the outer shell and the inner shell is movable relative to the outer chamber between a first position and a second position by a change in fluid conditions of a fluid supplied to the fluid valve. The first opening and the second opening overlap to define a passageway extending from the interior surface of the inner shell to the exterior surface of the outer shell. Relative movement of the inner shell from the first position toward the second position reduces a cross-sectional area of the passageway.

A gas turbine engine de-icing system includes a heat exchanger. A coolant loop is in fluid communication with the heat exchanger and is configured to circulate a heat transfer fluid. An engine oil loop is in fluid communication with the heat exchanger and is configured to transfer heat to the heat transfer fluid. A gas turbine engine inlet structure has at least one fan inlet guide vane. A spray bar is disposed at least partially in at least one fan inlet guide vane. The spray bar is in fluid communication with the coolant loop. The spray bar is configured to spray the heat transfer fluid onto an inner surface of the at least one fan inlet guide vane to de-ice the fan inlet guide vane.

A gas turbine engine de-icing system includes a heat exchanger. A coolant loop is in fluid communication with the heat exchanger and is configured to circulate a heat transfer fluid. An engine oil loop is in fluid communication with the heat exchanger and is configured to transfer heat to the heat transfer fluid. A gas turbine engine inlet structure has at least one fan inlet guide vane. A spray bar is disposed at least partially in at least one fan inlet guide vane. The spray bar is in fluid communication with the coolant loop. The spray bar is configured to spray the heat transfer fluid onto an inner surface of the at least one fan inlet guide vane to de-ice the fan inlet guide vane.

An air temperature sensor with a housing having a skin with a first and second portion, a temperature sensor having at least a portion extending through the housing, a set of fluid passageways defined within an interior of the housing, and a tube to receive bleed air from an aircraft engine located within the interior and to allow hot bleed air into the set of fluid passageways.