Methods and devices for cooling systems (100, 700) are provided that are in fluid communication with bleed air from a jet engine compressor. The cooling systems include: a first precooler (210) receiving bleed air from the jet engine compressor; a heat exchanger (730) downstream from the first precooler (210); a cooling system compressor (220) downstream from the first precooler (210), wherein the heat exchanger (730) and the cooling system compressor (220) are in separate flow paths from the first precooler (210); a cooling system precooler (230) downstream from the cooling system compressor (220); a cooling system turbine (240) with variable guide vanesVGTand downstream from the cooling system precooler (230); and a discharge conduit (245) downstream from the cooling system turbine (240) and the heat exchanger (730). A bypass line (290) can also be included that bypasses the cooling system turbine (240).

A thermal management system for transferring heat from a heat load includes a composite structural member that supports a heat load source and a heat transfer member in thermal contact with the composite structural member, and in thermal contact with a heat sink. The system further includes at least one thermally-conductive first fastener that is in thermal contact with the heat transfer member, couples the heat load source to the composite structural member, and conducts heat from the heat load source into the heat transfer member. The heat transfer member conducts heat from the thermally-conductive first fastener to the heat sink.

A single-piece gas turbine engine bleed duct for a gas turbine engine including: a main airflow conduit configured to transmit a bleed flow to a location outside the gas turbine engine; a pressure regulating valve for regulating airflow through the main airflow conduit; a first inlet duct for directing airflow to the main airflow conduit and toward the pressure regulating valve, the first inlet duct including a non-return valve; and a second inlet duct for directing airflow to the main airflow conduit and toward the pressure regulating valve, the second inlet duct including a control valve; wherein each of the first and second inlet ducts are directly connectable to an engine casing of a gas turbine engine.

A thermal management system for transferring heat from a heat load includes a first composite structural member that supports a heat load source, a second composite structural member, and a heat transfer member positioned between the first composite structural member and the second composite structural member and in thermal contact with at least the first composite structural member, and in thermal contact with a heat sink. The system further includes at least one thermally-conductive first fastener that is in thermal contact with the heat transfer member, couples the heat load source to at least the first composite structural member, and conducts heat from the heat load source into the heat transfer member. The heat transfer member conducts heat from the thermally-conductive first fastener to the heat sink.

A propulsion engine for an aeronautical vehicle defines a radial direction and a cooling air flowpath. The propulsion engine includes a power source; and a fan including a fan blade rotatable by the power source and extending generally along the radial direction, the fan blade defining an inlet, an outlet, and a cooling air passage extending between the inlet and the outlet and in airflow communication with the cooling air flowpath, the inlet being positioned inward from the outlet along the radial direction to provide a cooling airflow through the cooling air flowpath.

A propeller system for a tail section of an aircraft includes a propeller hub located at the tail section of the aircraft, a plurality of propeller blades mounted to and extending outwardly from the propeller hub, a propeller shaft coupled to the propeller hub and operable to rotate the propeller hub about an axis of rotation, and a propeller gearbox connected to the propeller shaft. The propeller gearbox is fluidly cooled by an airflow within the tail section. A spinner assembly surrounds the propeller hub. The spinner assembly includes at least one outlet opening formed therein downstream from the propeller hub relative to the airflow. The spinner assembly is rotatable to draw the airflow into at least one cooling flow inlet formed in the tail section and across the propeller gearbox to cool the propeller gearbox and out the at least one outlet opening.

Heat transfer devices and methods for enclosing a heat source and facilitating convective heat transfer from the heat source. A heat transfer device includes an outer wall having an outer surface exposed to an environment of the heat transfer device and defining an outer shape of the heat transfer device, and an inner wall defining a flow passage through the heat transfer device. The outer wall and the inner wall collectively define an internal volume that is configured to house the heat source. The flow passage includes an inlet configured to receive a fluid from the environment, and an outlet configured to exhaust the fluid from the flow passage that includes a core region extending between the inlet and the outlet and configured to deliver the fluid from the inlet to the outlet and allow heat to exchange between the fluid within the core region and the internal volume.

The invention concerns a helicopter ventilation architecture, said helicopter comprising at least two avionics bays (112a, 112b) comprising electronic equipment (116a, 116b) to be ventilated, said architecture comprising, for each avionics bay, an air inlet (120a, 120b) allowing outside air to enter the avionics bay in order to ventilate said avionics bay, and an air outlet (124a, 124b), allowing the air ventilating the avionics bay to exit the avionics bay, characterised in that the ventilation architecture further comprises a mixing chamber (134), connected to the air outlets, configured to receive the air originating from all the avionics bays, at least one air duct (138a, 138b), connected to the mixing chamber and to an outlet (130a, 130b) for discharging the air to the outside, and at least two fans (128a, 128b), arranged and distributed in the air duct or ducts.

An aircraft refrigeration system comprises an ambient air line, through which ambient air flows and which is connected to conduct ambient air to an aircraft, an ambient air compressor, arranged in the ambient air line, for compressing the ambient air in the ambient air line, a cabin exhaust air turbine, connected to a cabin exhaust air line and coupled to the ambient air compressor in the ambient air line and configured to expand cabin exhaust air flowing through the cabin exhaust air line and to drive the ambient air compressor, a transmission, configured to couple the cabin exhaust air turbine to the ambient air compressor and to set a speed of the ambient air compressor, and a motor, coupled to the transmission and configured to drive the ambient air compressor.

A laminar flow control surface having a foraminous section (also referred to below as a foraminous portion, 251) of a skin (250) of a vehicle to permit low temperature fluid to flow through the foraminous section to a heat exchanger (236), to reduce drag of the vehicle and to dissipate heat from the heat exchanger. The foraminous section (251) and the heat exchanger (236) synergistically reduce drag and transfer heat from the heat exchanger. The vehicle may be an aircraft with a laminar flow control system including a foraminous portion on a leading edge of the aircraft. While the aircraft is in flight, a portion of air impinging near the foraminous portion may flow laminarly about the leading edge, and another portion of the impingingair may flow through the foraminous portion to a heat exchanger to transfer heat from the heat exchanger to the air.