An anti-bird strike protection net for aircraft jet engine, comprising fixed circular truncated cone, conical protection net body, connection section, and double helix electrical heating wire, wherein the fixed circular truncated cone is secured on the cowl inlet's outside of the aircraft jet engine via which the conical protection net body is connected to the jet engine with the connection section: hinge, catches, and gas springs; The conical protection net body comprises the base of the conical protection net body, the conical support frame, the net which covers the outside of the conical support frame; the double helix electrical heating wire spirally surrounding the surface of the net is used as deicing device.

This invention is an aftermarket add-on device and can be an original integration part as jet engine manufacturing. It's simple and easy to be secured on the cowl inlet without changing the structure of the aircraft jet engine.

Splitter apparatus for gas turbine engines are disclosed. An example splitter apparatus may include a splitter including an annular outer wall substantially defining a convex leading edge; an annular splitter support positioned radially within the outer and including a forward end disposed substantially against a splitter inner; and an annular first bulkhead spanning between the outer wall and the splitter support. The outer wall, the splitter support, and the first bulkhead may define a generally annular splitter plenum. The forward end of the splitter support may include spaced apart, radially oriented metering slots. The outer wall may include an inner portion disposed radially inward from the splitter inner surface extending aft and including spaced-apart exit slots. The splitter plenum, the metering slots, and the exit slots may conduct airflow from the plenum, through the metering slots against the splitter inner surface, and through the exit slots.

Splitter apparatus for gas turbine engines are disclosed. An example splitter apparatus may include a splitter including an annular outer wall substantially defining a convex leading edge; an annular splitter support positioned radially within the outer and including a forward end disposed substantially against a splitter inner; and an annular first bulkhead spanning between the outer wall and the splitter support. The outer wall, the splitter support, and the first bulkhead may define a generally annular splitter plenum. The forward end of the splitter support may include spaced apart, radially oriented metering slots. The outer wall may include an inner portion disposed radially inward from the splitter inner surface extending aft and including spaced-apart exit slots. The splitter plenum, the metering slots, and the exit slots may conduct airflow from the plenum, through the metering slots against the splitter inner surface, and through the exit slots.

A bleed air system for an aircraft includes a turbocompressor including a turbine portion coupled to drive a compressor portion. The compressor portion includes a compressor inlet and a compressor air discharge. The turbine portion includes a turbine inlet and a turbine discharge. A passage delivers air to the compressor inlet from one or more locations of an engine. A passage delivers air to the turbine inlet from at least one or more locations of the engine. A passage receives air from the compressor discharge selectively delivered to one or more locations of the engine. An outlet passage receives air from the turbine discharge communicating airflow to an aircraft system. A controller directs inlet and discharge airflow of the compressor portion and the turbine portion to control operation of the turbocompressor. A gas turbine engine and a method are also disclosed.

An aircraft includes a fuselage and an engine housed in the fuselage. A ram-air engine inlet is formed on an exterior of the fuselage. The ram-air engine inlet is defined, at least in part, by a cowling door. The cowling door is moveable between a closed position and an open position. An intake duct fluidly couples the ram-air engine inlet to the engine. A filter is disposed across the intake duct. A bypass duct is formed in the ram-air engine inlet aft of the intake duct. The bypass duct is operable to be selectively opened and closed.

A method of operating an aircraft engine having a bladed rotor coupled thereto during an icing condition is provided. The method comprises controlling the aircraft engine to alternatingly achieve an ice accretion rotational speed of the bladed rotor at which ice accretion on blades of the bladed rotor occurs, and a reduced rotational speed of the bladed rotor to promote ice shedding from the blades of the bladed rotor. The reduced rotational speed is lower than the ice accretion rotational speed.

A method of operating an aircraft engine having a bladed rotor coupled thereto during an icing condition is provided. The method comprises controlling the aircraft engine to alternatingly achieve an ice accretion rotational speed of the bladed rotor at which ice accretion on blades of the bladed rotor occurs, and a reduced rotational speed of the bladed rotor to promote ice shedding from the blades of the bladed rotor. The reduced rotational speed is lower than the ice accretion rotational speed.

A gas turbine engine comprises a fan rotor, a lower speed compressor rotor and a higher speed compressor rotor, and a lower speed turbine rotor and a higher speed turbine rotor. The lower speed turbine rotor rotates the lower speed compressor rotor, and rotates a gear reduction to, in turn, rotate the fan rotor. The higher speed turbine rotor rotates the higher speed compressor rotor. An electrical generator is driven to rotate with one of the lower speed turbine rotor and the fan rotor.

A gas turbine engine comprises a fan rotor, a lower speed compressor rotor and a higher speed compressor rotor, and a lower speed turbine rotor and a higher speed turbine rotor. The lower speed turbine rotor rotates the lower speed compressor rotor, and rotates a gear reduction to, in turn, rotate the fan rotor. The higher speed turbine rotor rotates the higher speed compressor rotor. An electrical generator is driven to rotate with one of the lower speed turbine rotor and the fan rotor.

An engine bleed control system for a gas turbine engine of an aircraft is provided. The engine bleed control system includes an engine bleed tap coupled to a fan-air source or a compressor source of a lower pressure compressor section before a highest pressure compressor section of the gas turbine engine and a turbo-compressor in fluid communication with the engine bleed tap. The engine bleed control system also includes a controller operable to selectively drive the turbo-compressor to boost a bleed air pressure as pressure augmented bleed air and control delivery of the pressure augmented bleed air to an aircraft use.