Abstract: A heat exchanger is provided for warming fuel prior to introduction of the fuel to an engine, the heat exchanger including: a fuel inlet and a fuel outlet; a heat exchange matrix having heat transfer components past which the fuel flows between said inlet and said outlet and which are arranged to be heated; and swirl inducing means arranged between said inlet and said matrix arranged to cause fuel from said inlet to swirl prior to entering said matrix. By causing the fuel to swirl prior to entering the heat exchange matrix, entrained ice in the fuel can be caused to concentrate at an outer region of the heat exchanger thereby allowing fuel to continue to flow through the heat exchanger to the engine even when ice is present in the fuel.

Abstract: The gas turbine engine comprises at least one electrical generator and an electrical heater associated with a gas path of the engine. The heater is powered by the generator to selectively add heat to the gas turbine cycle.

Abstract: Setting in place of one or several coolers in the wall of a secondary flow of a bypass turbomachine. The wall extends from an intermediate casing toward a leading edge of a separator nose between a primary flow and the secondary flow. The wall includes a series of support arms attached to an intermediate casing, distributed over the perimeter of the wall and directed upstream. A series of surface air-oil heat exchangers forming wall segments are arranged end-to-end on the support arms, so as to form an annular wall. A shroud having a leading edge is arranged and fixed in the area of the upstream edges of the heat exchangers, so as to complete the wall. The support arms include hydraulic connectors connected to one another on each arm, adapted to cooperate with corresponding connectors in the area of the heat exchangers and in the area of the intermediate casing.

Abstract: An oil cooler is provided for a wall of an air flow passage of a gas turbine engine. The oil cooler has a heat exchanger with an oil conduit for carrying a flow of heated oil proximal to the wall of the air flow passage. The heat exchanger is arranged to transfer heat from the heated oil to the air flowing through the passage. The oil cooler further has a system for altering the flow of air across the heat exchanger such that the heat transfer effectiveness of the heat exchanger can be adjusted.

Abstract: A wall-cooling device for cooling a wall subjected to a heat source and to an outer fluid. The device includes a cooling part with a fastening portion and a cooling portion. The cooling part is made of a thermally-conductive shape-memory alloy having a transition temperature T. The fastener portion is for fastening the device rigidly on a face of the wall. The cooling portion is shaped so that it takes up a first position that is substantially parallel to the wall when the temperature of the wall is lower than the temperature T and so that it takes up a second position in which a plane of the second portion defines, relative to a plane tangential to the wall, a dihedral angle lying in the range 45° to 90° .

Abstract: One embodiment of a heat exchange system is disclosed having a heat exchange compressor and multiple evaporators capable of operating at different heat transfer requirements. The heat exchange compressor may be a single-stage or multi-stage compressor. In one form the evaporators return working fluid in separate streams to the heat exchange compressor. The heat exchange compressor may be integrated with a gas turbine engine and includes a number of inlets that correspond to a number of separate evaporators. Each inlet can be configured to receive working fluid at different locations within a pressure and velocity flow field created in the compressor. The heat exchange compressor may be driven by a shaft of the gas turbine engine and may be positioned at a variety of locations.

Abstract: A heat exchange system for use in fluid operated equipment to provide air and working fluid heat exchanges to cool the working fluid in airstreams on a stream side of a wall. An actuator is mounted to be substantially located on a side of the wall opposite the stream side thereof having a positionable motion effector. A heat exchanger core having a plurality of passageway structures therein to enable providing the working fluid to, and removal therefrom. The heat exchanger core is mounted on the motion effector so as to be extendable and retractable thereby through the opening for selected distances into that region to be occupied by the airstreams.

Abstract: The present invention provides a regenerative superheater system for an ejector ramjet engine. The invention includes a superheater in thermal communication with the combustion chamber of the ramjet engine. The superheater transfers thermal energy from combustion chamber to an ejectant which is then redirected upstream to the ramjet ejector. In one embodiment of the invention the temperature of the ejectant is modulated by a variable geometry cooler that controls the amount of thermal energy removed from the superheater system by ambient air. In an alternate embodiment of the invention, the temperature of the ejectant is modulated by a variable geometry superheater that controls the amount of thermal energy added to the superheater system through combustion gas.

Abstract: A system and method are provided for controlling the temperature of engine bleed air from a turbofan gas turbine engine. The system includes a fan air valve and a fan air valve controller. The fan air valve is adapted to receive a flow of fan air from a turbofan gas turbine engine intake fan. The fan air valve is coupled to receive valve position commands and is configured, in response to the valve position commands, to move to a valve position to thereby control the engine bleed air temperature. The fan air valve controller is configured to implement a linear quadratic regulator (LQR) control. The fan air valve controller is adapted to receive a plurality of sensor signals, each sensor signal representative of one or more system parameters, and is configured, in response to the sensor signals, to supply the valve position commands to the fan air valve.

Abstract: An aircraft adaptive power thermal management system for cooling one or more aircraft components includes an air cycle system, a vapor cycle system, and a fuel recirculation loop operably disposed therebetween. An air cycle system heat exchanger is between the air cycle system and the fuel recirculation loop, a vapor cycle system heat exchanger is between the vapor cycle system and the fuel recirculation loop, and one or more aircraft fuel tanks are in the fuel recirculation loop. An intercooler including a duct heat exchanger in an aircraft gas turbine engine FLADE duct may be in the air cycle system. The system is operable for providing on-demand cooling for one or more of the aircraft components by increasing heat sink capacity of the fuel tanks.

Abstract: An aircraft propulsion unit including a turbojet and a heat exchanger located above the turbojet and drawing a cooling air stream and a hot air stream in the turbojet. Wherein the cooling air intake and hot air intake surfaces in the housing are directed forward of the turbojet and have normal lines inclined relative to the axis of the turbojet. The embodiment also concerns an aircraft equipped with at least one such propulsion unit.

Abstract: A system for managing heat fluxes of an aircraft is provided, including a fuel tank supplying a turbomachine with fuel via a fuel supply circuit; a cell cooling circuit connected to a thermal effluents source, which integrates a first heat exchanger ensuring heat transfer between a coolant circulating in the cell cooling circuit and air flow channeled into a cooling air channel extending from an air intake to an exhaust; and a second heat exchanger ensuring a heat transfer between the coolant circulating in the cell cooling circuit and the fuel. The air intake includes a blocker to open/close it. The system includes a fuel transfer circuit, which connects the fuel tank to the second heat exchanger and provides fuel to the second heat exchanger, including a valve selectively changing a fuel supply direction from the second heat exchanger between the directions of the tank and the fuel supply circuit.

Abstract: A method of assembling a heat exchanger in a gas turbine engine assembly is provided. The method includes providing a structural body including a wall having an elongated slot, providing a heat exchanger operably associated with a bracket, and providing a slip joint for coupling the bracket to the structural body. The slip joint includes a standoff, an elongated member, and a biasing member disposed about the elongated member. The standoff is inserted through the wall so that the biasing member imparts a predetermined biasing force between the bracket and the structural body. The slip joint allows relative sliding movement between the bracket and the structural body when a sliding force due to thermal expansion or contraction is greater than the biasing force.

Abstract: A turbogenerator having a gas turbine engine powering generator and a cooling system and an annular heat exchanger powered by a fan disposed across an outer fan duct of the engine. Fan variable inlet and outlet guide vanes may be used to vary power between the fan and the generator which are drivenly connected to a low pressure turbine. Inner and outer portions separated by a rotating shroud of the fan are disposed in annular inner and outer fan ducts respectively. A directed energy weapon may be powered by the generator and cooled by the cooling system. A refrigeration apparatus may be operably disposed between the annular heat exchanger and the directed energy weapon for cooling the directed energy weapon and conditioning power electronics for the weapon. The refrigeration apparatus may include a vapor cycle cooling system and a cold storage containing a phase change material.

Abstract: A thermal management system for a gas turbine engine includes a first heat exchanger that provides a first conditioned fluid, a second heat exchanger in series flow communication with the first heat exchanger and that exchanges heat with the first conditioned fluid to provide a second conditioned fluid, and a third heat exchanger in parallel flow communication with the second heat exchanger and that selectively exchanges heat with the first conditioned fluid to provide a third conditioned fluid. The second conditioned fluid and the third conditioned fluid combine to provide a mixed conditioned fluid. A first portion of the mixed conditioned fluid including a first temperature is communicated to a first engine system. A second portion of the mixed conditioned fluid including a second temperature is communicated to a second engine system. The second temperature includes a greater temperature than the first temperature.

Abstract: A device for producing electrical power has a thermoelectric device coupled to an aircraft bleed system for generating electrical power using temperature differentials between ram air and bleed air. A system control is coupled to the ram air and the bleed air to control a temperature of the bleed air and to control a temperature gradient of the thermoelectric device.

Abstract: The present invention relates to a cooling device (2) for cooling combustion gases from a rocket motor (1) in an antitank weapon. The cooling device (2) comprises a coolant reservoir (3, 21) containing a coolant (4). The coolant reservoir (3, 21) is arranged such that the coolant (4) is transferred from the coolant reservoir (3, 21) to the combustion gases in the gas outlets (6) of the rocket motor (1) in response to a pressure increase in the rocket motor (1). The coolant reservoir (3, 21) constitutes an integral part of the rocket motor (1) and is connected to the combustion chamber (7) by at least two gas inlets (8) for pressurization of the coolant (4). The coolant reservoir is furthermore connected to the rocket motor nozzle (9) by at least two coolant outlets (10) for transfer of coolant (4) from the coolant reservoir (3, 21) to the rocket motor nozzle (9).

Abstract: A lubricant supply system for a gas turbine engine has a lubricant lube pump delivering lubricant to an outlet line. The outlet line is split into at least a hot line and into a cool line, with the hot line directed primarily to locations associated with an engine that are not intended to receive cooler lubricant, and the cool line directed through one or more heat exchangers at which lubricant is cooled. The cool line then is routed to a fan drive gear for an associated gas turbine engine. A method and apparatus are disclosed.

Abstract: A method of assembling a gas turbine engine includes providing at least one heat exchanger assembly including a heat exchanger and a lobed mixer extending downstream from the heat exchanger, wherein the mixer includes a plurality of lobes that each define a first chute and at least one second chute that extends between each pair of adjacent spaced-apart lobes, and coupling the at least one heat exchanger assembly within a bypass duct of the engine such that the at least one heat exchanger assembly is coupled to at least one of an outer engine casing and an inner engine casing of the turbine engine.

Abstract: Air bleed device for cooling components in a turbine engine, including an annular conduit (18) having a radially internal portion swept by an airflow (22) and which comprises at least one air inlet orifice (20) formed in an upstream radial wall (38) of the conduit (18), a flap valve (40) for controlling the airflow entering through the orifice (20) and of which the flapper (41) includes a plate (42) applied against the wall (38) and capable of being moved by sliding on said wall (38) by a maneuvering member (50) mobile in translation parallel to the wall (38) between a position in which the plate (42) closes off the orifice (20) and a position in which the plate (42) opens said orifice.

Abstract: A precooler for cooling compressor bleed air for an environmental control system includes a heat exchanger in fluid communication with a source of cooling air and operable for cooling the bleed air. A variable bypass valve between a bleed air source and environmental control system is operable for bypassing at least a portion of the compressor bleed air around the heat exchanger. The cooling air may be a portion of fan air modulated by a variable fan air valve. The bleed air source may be selectable between the low pressure bleed air source and a high pressure bleed air source. One method includes flowing the compressor bleed air from a single low pressure source only and increasing thrust sufficiently to meet a minimum level of pressure of the bleed air during one engine out aircraft operating condition during approach or loitering.

Abstract: A method for assembling a turbine engine includes assembling a heat exchanger assembly that includes at least a radially inner plate, a radially outer plate, and a heat exchanger coupled between the radially inner and outer plates, forming the heat exchanger assembly such that the heat exchanger assembly has a substantially arcuate shape, and coupling the heat exchanger to a fan casing such that the heat exchanger is positioned upstream or downstream from the fan assembly.

Abstract: In a heating apparatus for heating the air sucked into a gas turbine by a heat exchanger, the temperature fluctuation of the heated air is suppressed even in the period, for which a steam source to be fed to the heat exchanger is changed. For suppression, a heat exchanger is fed with both a self-can steam, the feed rate of which is controlled by a self-can steam control valve, and the auxiliary-steam, the feed rate of which is controlled by an auxiliary-steam control valve. At starting time, the quantity of the auxiliary-steam is reduced at a constant rate, and that of the self-can steam is increased while a feedback control and a feedforward control are being made. At stopping time, the quantity of the self-can steam is reduced at a constant rate, and that of the auxiliary-steam is increased while the feedback control and the feedforward control are being made.

Abstract: Closed-cycle rocket engine assemblies including a combustor assembly, a combustor jacket, a turbine, a first pump and a mixing chamber are disclosed. The combustor jacket facilitates the transfer of heat from the combustor assembly into a fluid and the turbine is driven by a heated fluid from the combustor jacket. The mixing chamber may include a first inlet to receive a fluid from the turbine, a second inlet to receive a fluid from a first reactant reservoir, and an outlet to deliver a fluid to the first pump. Additionally, the first pump may be coupled to and powered by the turbine and the first pump may be configured to deliver at least a portion of the fluid from the mixing chamber into the combustion chamber of the combustor assembly. Related methods of operating such rocket engine assemblies are also disclosed.

Abstract: A method for starting up an aircraft turbine engine using an electric starter is provided. The aircraft turbine engine is provided with at least one hydraulic fluid circuit for lubricating the aircraft turbine engine, the aircraft turbine engine lubricating circuit is provided with a temperature-regulating section that is able to remove the heat generated by the aircraft turbine engine when the aircraft turbine engine is running, the electric starter is provided with at least one hydraulic fluid circuit for lubricating parts of the electric starter, and the lubricating circuit of the electric starter is provided with the temperature-regulating section that is able to remove the heat emitted by the electric starter when the lubricating circuit of the electric starter is operating.

Abstract: A gas turbine engine system is disclosed which includes a core passage and a bypass passage which can be configured as a fan bypass duct or a third stream bypass duct. The core passage and bypass passage are routed to flow through a nozzle before exiting overboard an aircraft. The nozzle includes moveable members capable of changing a configuration of the nozzle. In one form the moveable members are capable of changing throat area for portions of the nozzle that receive working fluid from the core passage and the bypass passage. The bypass passage can include a branch. In one form the branch can include a heat exchanger. The bypass passage can also provide cooling to one or more portions of the nozzle, such as cooling to a deck of the nozzle.

Abstract: A vehicle has a body and a source of a propellant. An engine is carried by the body. The engine reacts the propellant to produce thrust. The engine has a heat exchanger transferring heat from the reaction to at least a component of the propellant and generating electricity thermoelectrically.

Abstract: A system, methods, and user-friendly program product to optimize energy recovery for a process or cluster of processes under all possible combinations of given process changes and stream-specific minimum temperature approach values without enumeration, are provided. The systems, methods, and program product can include steps/operations to identify a set of common-structure heat exchanger network designs which allow for construction of a physically exchanger network easily retrofittable to accommodate time-dependent new operating modes, disturbances, and uncertainty schemes.

Abstract: An aircraft propulsion arrangement including a gas turbine aircraft engine having a compressor, an oil system configured to route engine oil through a heat exchanger mounted to an external surface of an aircraft component so as to define part of an aerodynamic surface to a flow of ambient air, and a duct arrangement fluidly connecting the compressor to the heat exchanger. A method of operating the gas turbine aircraft engine, the method involving the steps of: (a) flowing engine oil through the heat exchanger and thus into heat-exchange relationship with the ambient air; and (b) directing a bleed flow of compressor gas drawn from the compressor along the duct arrangement and into heat-exchange relationship with the engine oil in the heat exchanger, wherein the directing step (b) is performed selectively.

Abstract: A heat engine for use in conjunction with a power generating plant, including a turbine section having a number of turbines, a heat exchanger section having a number of modules through which the expanded working fluid of the power generating plant and other sources of heat are circulated, a laminar flow inducing section, and a tower section for providing a pressure differential across the turbines of the turbine section. In use, the heat engine provides the dual function of: heating air to generate an updraft such that air forces its way into the turbine sections to drive the turbines and generate additional electricity; and using incoming colder air to condense the expanded working fluid and cool other sources of heat.

Abstract: A cooling system in an aircraft gas turbine engine includes a heat exchanger positioned within an annular nacelle space surrounding a bypass duct of the engine. The heat exchanger has a radial coolant passage extending between the bypass duct and ambient air surrounding the nacelle, and a flow passage extending substantially normal to the radial passage for direction of a fluid to be cooled therethrough. A configuration of this nature may assist in defining a no-flow length of the heat exchanger in a third direction normal to the other two mentioned directions, which may allow for improved performance within a given radial envelope.

Abstract: A pre-cooler having an annular cross-sectional shape about the axis of a pod and arranged inside the rear part of an inner shroud in external contact with a cold flow exiting a fan duct channel and in close contact with a cooling air stream drawn from the cold flow.

Abstract: Gas turbine engine systems involving integrated fluid conduits are provided. In this regard, a representative engine casing for a gas turbine engine includes: a casing wall; and a fluid conduit integrated into the casing wall.

Abstract: An engine system includes a thrust chamber that has a cooling channel. The cooling channel is adapted to provide sustained cracking conditions for a fluid at steady-state operating conditions. A turbine has an input in fluid communication with an output of the cooling channel. A pump is mechanically coupled with the turbine and is in fluid communication with the cooling channel.

Abstract: Disclosed is a turbo pump in which a pump impeller is connected to one end of a rotary shaft and a turbine is connected to the other end of the rotary shaft. The turbo pump is designed such that an equivalent region, between a turbine efficiency curve obtained on the basis of a conditional expression where the number of rotations of the rotary shaft is maintained constant regardless of a pump flow rate and a turbine efficiency curve of an actual machine, becomes an operation region.

Abstract: A cryogenic-propellant rocket engine includes: at least a first tank for a first liquid propellant; a second tank for a second liquid propellant; a third tank for an inert fluid; an axisymmetrical nozzle including a combustion chamber, a device for injecting first and second liquid propellants into the combustion chamber, a nozzle throat, and a divergent section; and a heater device including at least one duct for conveying the inert fluid and arranged outside the nozzle in immediate proximity thereof, but without making contact therewith, to recover energy of thermal radiation emitted when the rocket engine is in operation and to heat the inert fluid.

Abstract: A cooling system of a gas turbine engine, includes a heat exchanger having a common wall shared by a first air passage for directing a portion of a compressor air flow to be used as cooling air, and a second air passage for directing a portion of a bypass air flow, the portion of the compressor air flow being thereby cooled by the portion of the bypass air flow through the common wall.

Abstract: A turbine engine component of a turbofan engine fitted with a bypass air valve includes at least one turbine engine component having a surface with at least one aperture, said turbine engine component located from between a bypass fan duct and a turbine exhaust nozzle of the turbofan engine; a bypass air valve includes a liner concentrically disposed about the turbine engine component and parallel to a centerline of the turbofan engine, the liner has a surface including at least one aperture and at least one impermeable region, and means for actuating the liner about the turbine engine component; and a flow transfer location comprising an area proximate to a turbine exhaust stream flow.

Abstract: A propulsion system for an aircraft includes a heat exchange unit through which a hot air stream and a cooling air stream flow. The heat exchange unit includes a hot air inlet pipe situated on a rear face of the heat exchange unit, to carry the hot air stream into the heat exchange unit, and a cooling air discharge pipe situated on the rear face of the heat exchange unit to expel the cooling air stream from the heat exchange unit, the hot air inlet pipe and the cooling air discharge pipe being concentric with each other. The hot air stream flows from the rear face forward to a front of the heat exchange unit, then from the front of the heat exchange unit, to a rear of the heat exchange unit, so as to be expelled via the rear face of the heat exchange unit.

Abstract: A method for using a heat exchange system in operating equipment in which a working fluid is utilized in providing selected operations thereof, including for use in lubricating systems for aircraft turbofan engine equipment, the heat exchange system for providing air and working fluid heat exchanges to cool the working fluid at selectively variable rates in the operating equipment developed airstreams. A heat exchanger core is provided in a controlled air flow duct system opening at its entrance to those airstreams and having its outlet end opening downstream in those airstreams.

Abstract: A bypass turbine engine includes a precooler having an annular cross-sectional shape about the axis of a pod and arranged inside the rear part of an inner shroud in contact with a cold fluid exiting a fan duct.

Abstract: A heat exchange system for use in fluid operated equipment to provide air and working fluid heat exchanges to cool the working fluid in airstreams on a stream side of a wall. An actuator is mounted to be substantially located on a side of the wall opposite the stream side thereof having a positionable motion effector. A heat exchanger core having a plurality of passageway structures therein to enable providing the working fluid to, and removal therefrom. The heat exchanger core is mounted on the motion effector so as to be extendable and retractable thereby through the opening for selected distances into that region to be occupied by the airstreams.

Abstract: A system, methods, and user-friendly program product to optimize energy recovery for a process or cluster of processes under all possible combinations of given process changes and stream-specific minimum temperature approach values without enumeration, are provided. The systems, methods, and program product can include steps/operations to synthesize multiple heat exchanger network designs according to different process variations schemes, analyze the impact of process variations on each heat exchanger network, and identify which design provides a more optimal heat exchanger network design and which process variation scheme represents a more optimal flexibility index.

Abstract: Systems, methods, and program product to synthesize a grass-roots heat exchanger network for non-thermodynamically constrained problems for a plurality of hot process streams to be cooled and a plurality of cold process streams to be heated according to a plurality of utilities targets, are provided. An exemplary system includes a heat exchange network synthesizing computer and heat exchange network synthesizing program product configured to analyze the waste heat recovery problem without decomposition to produce a result having an optimal number of network heat exchanger units through application of advanced process-to-process matching schemes for non-thermodynamically constrained problems.

Abstract: In a by-pass gas turbine engine having a lobed type by-pass air and turbine exhaust mixer unit, the reheat gutters are arranged in the mixer in the turbine exhaust duct. The gutters extend radially and are substantially Y-shaped, each of the two legs of the Y extending radially along the sides of a mixer lobe.

Abstract: A gas turbine reheat system is provided with channel-shaped gutters, with a wall portion spaced from each gutter to form a space therebetween. The space is supplied with cooling air conveniently from the engine by-pass duct.

Abstract: An external jet nozzle mixer includes identically formed lobes. The external mixer works with the internal mixer further to mix the engine internal bypass flow with the internal jet engine core flow to level the disparate flow velocities, to reduce the peak velocities from the jet engine core and increase the lower bypass velocities of the engine internal bypass flow, and thereby reduce noise. The internal lobe contours act as lifting flutes, causing mixing of the primary hot and cold flows to mix before exiting the nozzle. The external lobe contours act as venturi chutes, accelerating the cooler ambient secondary air flow. The lobes thus act collectively as an injector to force the cooler ambient secondary flow into the previously mixed primary flow as it exits the nozzle. Also obtained is an increased thrust efficiency and, consequently, decreased fuel consumption and engine emissions.

Abstract: An aircraft system includes a heat source and a passage near the heat source for carrying fluid having a cooling capacity to cool the heat source. The passage includes a catalyst that endothermically cracks the fluid to increase the cooling capacity.

Abstract: A propellant tank for storing a liquid propellant A and supplying vapor produced by evaporation of part of the liquid propellant A to an external location comprises a tank body for storing the liquid propellant A, a mesh member arranged inside the tank body to cover a liquid surface of the liquid propellant A to divide an interior of the tank body into a liquid propellant storing area LA and a gas storing area GA by utilizing surface tension of the liquid propellant, and a heater arranged to a gas storing area GA side of the tank body to keep the gas storing area GA at higher temperature than temperature in the liquid propellant storing area LA. The tank body has a propellant inlet open into the liquid propellant storing area LA and a gas outlet open into the gas storing area GA.

Abstract: A propulsion unit comprises a turbojet engine and a thermal exchanger, situated above the turbojet engine. The turbojet engine comprises a regulation device to regulate an air flow-rate in a cooling air lead-in conduit conveying a stream of cooling air into a thermal exchanger, The regulation device regulates the air flow-rate upstream with respect to the cooling air lead-in conduit.