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.

A deicing apparatus for aircraft comprises a passive vortex tube adapted to be mounted at a location at or adjacent a component of an aircraft and adapted to heat the component. A deicing method is also disclosed.

A deicing apparatus for aircraft comprises a passive vortex tube adapted to be mounted at a location at or adjacent a component of an aircraft and adapted to heat the component. A deicing method is also disclosed.

An aircraft may include at least one secondary power system requiring secondary power, at least two main engines, and at least three non-propulsive utility power (NPUP) generation systems. The NPUP generation systems may each be configured to provide full-time secondary power during operation of the aircraft. The NPUP generation systems may be configured to provide at least a portion of the secondary power required by the secondary power system.

An aircraft may include at least one secondary power system requiring secondary power, at least two main engines, and at least three non-propulsive utility power (NPUP) generation systems. The NPUP generation systems may each be configured to provide full-time secondary power during operation of the aircraft. The NPUP generation systems may be configured to provide at least a portion of the secondary power required by the secondary power system.

An engine bleed air system of an aircraft, comprising a heat exchanger which comprises a rectangular core with a rectangular outlet section, a cylindrical outlet duct and a transition area between the rectangular outlet section of the rectangular core and the cylindrical outlet duct, and at least a downstream system, further comprising a flow deflector located at least in the cylindrical outlet duct, such that the outlet flow characteristics homogeneity are improved and a particular flow deflection and distribution of the outlet flow is achieved, avoiding the damage of the at least one downstream system. The invention also provides a method for homogenizing the temperature of the outlet flow of a heat exchanger outlet in an engine bleed air system of an aircraft.

An engine bleed air system of an aircraft, comprising a heat exchanger which comprises a rectangular core with a rectangular outlet section, a cylindrical outlet duct and a transition area between the rectangular outlet section of the rectangular core and the cylindrical outlet duct, and at least a downstream system, further comprising a flow deflector located at least in the cylindrical outlet duct, such that the outlet flow characteristics homogeneity are improved and a particular flow deflection and distribution of the outlet flow is achieved, avoiding the damage of the at least one downstream system. The invention also provides a method for homogenizing the temperature of the outlet flow of a heat exchanger outlet in an engine bleed air system of an aircraft.

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.

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.

An engine bleed control system for a gas turbine engine of an aircraft is provided. The engine bleed control system includes a multi-tap bleed array including engine bleed taps coupled to a compressor source of a lower pressure compressor section before a highest pressure compressor section of the gas turbine engine. A highest stage of the engine bleed taps has a maximum bleed temperature below an auto-ignition point of a fuel-air mixture of the aircraft at idle engine power at a maximum aircraft altitude and a pressure suitable for pressurizing the aircraft at the maximum aircraft altitude. The engine bleed control system also includes a plurality of valves operable to extract bleed air from each of the engine bleed taps. A controller is operable to selectively open and close each of the valves based on a bleed air demand and control delivery of the bleed air to an aircraft use.