Actuator systems and methods of operation are disclosed. The systems include a hydraulic actuator having a primary piston having a piston head arranged within a housing defining retract and extend chambers on opposite sides of the piston head. A control element is configured to control a supply of pressure to each of the retract and extend chambers. An actuator valve is coupled to the housing and includes a secondary piston that is biased into the retract chamber in an open flow state and when the primary piston is in a fully retracted state the piston head urges the secondary piston into a closed flow state. The actuator valve defines a flow chamber where, in an open flow state, fluid can be passed through the flow chamber and in a closed flow state the fluid is prevented from passing through the flow chamber.

Methods and apparatus for cooling a surface on a flight vehicle and/or generating power include advancing the flight vehicle at a speed of at least Mach 3 to aerodynamically heat the surface. A supercritical working fluid is circulated through a fluid loop that includes compressing the supercritical working fluid through a compressor, heating the supercritical working fluid through a heat intake that is thermally coupled to the surface, expanding the supercritical working fluid in a thermal engine to generate a work output, cooling the supercritical working fluid, and recirculating the supercritical working fluid to the compressor. The work output of the thermal engine is operably coupled to the compressor, and may optionally be coupled to a generator to produce power. The supercritical working fluid absorbs heat from the surface, eliminating hot spots and permitting use of lighter and/or less expensive materials.

An aircraft comprises a machine body. The machine body encloses a turbofan gas turbine engine and a plurality of ancillary systems. The turbofan gas turbine engine comprises, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a combustor module, a turbine module, and an exhaust module. The machine body comprises a single fluid inlet aperture, with the fluid inlet aperture being configured to allow a fluid cooling flow to enter the machine body and to pass through the heat exchanger module. The heat exchanger module is configured to transfer a waste heat load from the gas turbine engine and the ancillary systems to the fluid cooling flow prior to an entry of the entire fluid cooling flow into the fan module.

A module ensuring an acoustic attenuation of a flow of a first fluid and a heat exchange between the first fluid and a second fluid. The module comprises a perforated first wall with a cutout, a second wall, a cellular structure extending from the second wall to the first wall, a recess provided in the cellular structure between a perforated bottom and a perforated top, and a heat exchanger which is fixed inside the recess between the bottom and the top and in which the second fluid circulates. Such a module ensures an attenuation of sound waves and a heat exchange without limiting the attenuation surface.

An air-breathing turbojet engine for a hypersonic vehicle is shown. The engine comprises a pump for pumping a cryogenic fuel, an inlet configured to compress inlet air by one or more shocks, a cooler to cool the compressed inlet air using the cryogenic fuel, and a turbo-compressor to compress the air further. A precooler cools the compressed inlet air using compressed cooled air from the turbo-compressor. A combustor receives compressed cooled air from the turbo-compressor and a first portion of the cryogenic fuel for combustion. A first turbine expands and is driven by combustion products, and a second turbine expands and is driven by a second portion of the cryogenic fuel. The first turbine and the second turbine drive the turbo-compressor via a shaft. An afterburner receives combustion products from the first turbine and the second portion of the cryogenic fuel from the second turbine for combustion therein.

An aircraft engine apparatus (1) includes: a rotating shaft (6); a fan (10) driven by the rotating shaft; a fan case surrounding the fan from outside in a radial direction of the rotating shaft; a nose cone (13) disposed upstream of the fan; a casing (2) that accommodates at least part of the rotating shaft and supports the fan case; a first motive force transmitter (9) that transmits motive force of the rotating shaft to the fan; and a support member (12) disposed inward of the first motive force transmitter in the radial direction, the support member coupling the nose cone to the casing such that the support member supports the nose cone in a stationary state.

An aircraft includes a machine body that encloses a turbofan gas turbine engine and a plurality of ancillary systems. The turbofan gas turbine engine includes, in axial flow sequence, a first heat exchanger module, a fan assembly, a compressor module, a combustor module, a turbine module, and an exhaust module. The aircraft includes a second heat exchanger module. The machine body comprises a single fluid inlet aperture, with the fluid inlet aperture being configured to allow a fluid cooling flow to enter the machine body and to pass through the first heat exchanger module. When a temperature of the fluid cooling flow is less than a temperature of a fluid to be cooled, the fluid to be cooled is directed to the first heat exchanger module, and when a temperature of the fluid cooling flow is greater than a temperature of the fluid to be cooled, the fluid to be cooled is directed to the second heat exchanger module and cooled using a fuel supply for the gas turbine engine.

A hypersonic aircraft includes one or more leading edge assemblies that are designed to manage thermal loads experienced at the leading edges during high speed or hypersonic operation. The leading edge assembly includes a plurality of structural layers and a plurality compliant layers alternately stacked with each other to facilitate thermal expansion and movement between the plurality of structural layers, while also providing a thermal break between the plurality of structural layers.

A method and apparatus for supplying air to a precooler. Air flow is created through a fan duct in an engine system. A first portion of the air flow is directed into a first inlet of an inlet system to feed a first half of the precooler. A second portion of the air flow is directed through the fan duct into a second inlet of the inlet system to feed a second half of the precooler.

A heat exchanger to cool an oil flow with an air flow and a fuel flow includes at least one oil flow layer to receive the oil flow, an air flow layer to receive the air flow, wherein the air flow layer is in thermal communication with the at least one oil flow layer, and a fuel flow layer to receive the fuel flow, wherein the fuel flow layer is in thermal communication with the at least one oil flow layer.