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 and method for cooling at least a portion of an engine are provided. The engine is cooled using a fuel, such as a high heat sink fuel, that is subsequently used for combustion in the engine. The fuel can be used to cool one or more of the gases and/or components in the engine, thereby cooling the engine including an exhaust nozzle. For example, the fuel can be circulated through one or more heat exchanging devices that are disposed inside or outside a passage of the engine, and the fuel can absorb thermal energy from the engine or air that flows in the engine passage. In any case, the cooling of the engine can result in a reduction to the infrared signature of the engine.

Abstract: Disclosed is a methane engine for rocket propulsion. A methane supply pump (36) operated by a turbine (30) supplies a part of methane to a nozzle cooling channel (56, 156) installed on a nozzle (54, 154) of a combustor (50, 150) and supplies the other part of the methane to a combustion chamber cooling channel (53, 153) installed on a combustion chamber (52, 152) of the combustor (50, 150) so as to regulate the amount of methane supplied to a mixing head (51, 151) while maintaining the cooling properties of the combustor (50, 150), thus providing extensity of coping with changes in propulsive force and design of the combustor (50, 150). Further, a part of methane in a gas state discharged from the combustion chamber cooling channel (53, 153) is supplied to a mixing head (76) of a gas generator (94), thus providing the re-liability of the engine.

Abstract: Cavities 4 are defined within an engine nacelle 1 with an inner surface 2 and an outer surface 3. Within these cavities 4, an evaporatable liquid 5 is arranged such that a proportion of the liquid 5 is in a gaseous phase in order to facilitate heat transfer between the inner surface 2 or the outer surface 3 and a hot surface. The liquid condenses upon the inner surface 2 or the outer surface 3 and collects into a pool for subsequent further evaporation by the hot surface of an insulated gate bi-polar transistor or convertor/generator components in order to facilitate further cooling of electric components housed or arranged in the nacelle 1 or nearby.

Abstract: A bleed air cooler assembly of a gas turbine engine comprises a cooler body defining a fluid passage therein and a connector detachably affixed with the cooler body installed in an annular bypass air passage.

Abstract: Gate valve (10) for a system controlling cooling in a turbomachine, comprising a gate (20) mounted so that it can pivot about an axis (16) between a position in which it shuts off an air passage orifice (72) and a position in which it opens the orifice (72), and means (12, 14, 18) for rotating the gate (20) about the axis (16), these means comprising two superposed coaxial rotary members (14, 18) collaborating with one another via cam surfaces (40, 46) designed such that rotating the first rotary member (14) from the shut-off position causes a translational movement of the second rotary member (18) and of the gate (20) along the axis of rotation (16) followed by a rotation of this member (18) and of the gate (20) about this axis. Application of the invention to the cooling of turbojet jet pipe nozzle flaps.

Abstract: The invention concerns a turbofan provided with a pre-cooler. The invention is characterized in that to evacuate the heated cool air stream (24), at least one discharge pipe (25) is arranged in the chamber (12) and connects the pre-cooler (18) to at least one discharge orifice (26) provided in the inner fairing (10), in output of the exhaust nozzle (3) and at least more or less opposite the wing (16).

Abstract: The invention relates to an aeroengine comprising: a primary air circuit; a high-pressure compressor fed by said primary air; and a secondary air circuit. The engine further comprising at least one heat exchanger mounted in the primary air circuit upstream from said high-pressure compressor, said heat exchanger comprising a cold secondary circuit and a hot primary circuit, said hot primary circuit being fed with air from said primary air circuit and said cold secondary circuit being fed with air from the secondary air circuit. The cold circuit of the heat exchanger is fed by pipes that open out into the secondary air circuit.

Abstract: A gas turbine engine comprising a nacelle having an intake, the intake defining a generally annular chamber, an engine accessory and a heat exchanger for cooling a fluid of the accessory. The chamber is closed and the heat exchanger is disposed within the chamber operable to provide heat to prevent ice forming on the intake during engine operation.

Abstract: There is described a heat shield element comprising a wall that is provided with a hot face which can be impinged upon by a hot medium and a cold face located opposite the hot face. The heat shield element further comprises a coolant distribution system which is assigned to the cold face. In order to convectively cool the heat shield element in a particularly effective manner, a plurality of cooling ducts which extend along the hot face are provided within the wall. Said cooling ducts are fluidically connected to the distribution system such that the coolant can be distributed to the individual cooling ducts with the aid of the distribution system. The convectively cooled heat shield element can be used in a particularly advantageous fashion for the heat-resistant lining of a combustion chamber, especially a combustion chamber of a gas turbine system.

Abstract: A method and apparatus for cooling a fluid in a bypass gas turbine engine involves directing the fluid to the bypass duct of the engine to allow for heat exchange from the fluid to bypass air passing through the bypass air duct.

Abstract: An axisymmetric nozzle for a gas turbine engine has divergent flaps and seals disposed about a central longitudinal axis thereof. The divergent flaps and seals each have an inner surface defining cooling air inlet holes at an upstream portion, cooling air exit holes at a downstream portion, and cooling air channels disposed within the divergent flap and seal, and communicating at a first end with the inlet holes and at a second end with the exit holes for conducting cooling air therethrough. Some of the exit holes of the divergent flap are disposed along lateral edge regions thereof. The divergent seals are interposed between adjacent divergent flaps, and each include lateral edge regions extending laterally beyond the exit holes. The lateral edge regions of each divergent seal have an outer surface overlying at least a portion of the exit holes of an adjacent divergent flap.

Abstract: The invention provides a gas turbine installation, a cooling air supplying method, a method of modifying a gas turbine installation, which can reduce the amount of compressed air extracted for cooling of an exhaust plenum and which can increase energy efficiency. The gas turbine installation comprises an exhaust plenum comprising an exhaust casing connected to the downstream side of a turbine and made up of an outer casing and an inner casing, and an exhaust diffuser made up of an outer diffuser and an inner diffuser, which are disposed between the outer casing and the inner casing; a first cooling system for introducing, from the outer peripheral side of the outer diffuser, cooling air for cooling the exhaust plenum; and a second cooling system for introducing, from the inner peripheral side of the inner diffuser, cooling air for cooling the exhaust plenum.

Abstract: A rocket engine chamber has a back structure and a heat exchanger structure located radially inwardly of the back structure. The heat exchanger structure has a plurality of integrally formed cooling passageways with adjacent ones of the passageways having a common sidewall structure. Each passageway has a top section which forms the wall that contacts the hot gases. The top section is formed by a wall structure having an inner and outer surface formed by an arc of a circle and a substantially constant wall thickness.

Abstract: An improved engine construction is provided. The engine construction comprises a combustion chamber, a smooth wall nozzle, and a transition zone between the chamber and the smooth wall nozzle. The transition zone has a coolant system which includes a manifold formed from a non-copper material through which a coolant flows. In a high heat transfer embodiment, the transition zone includes an additional manifold formed from a copper based material.

Abstract: An improved engine construction, such as a rocket engine construction is provided. The engine construction comprises a combustion chamber, a smooth wall nozzle, and a transition zone between the chamber and the smooth wall nozzle. The transition zone has a coolant system which includes a manifold formed from a non-copper material through which a coolant flows. In a high heat transfer embodiment, the transition zone includes an additional manifold formed from a copper based material.

Abstract: A pressure vessel bypass valve forms a bypass system for enabling a part of the air compressed by a compressor to directly flow into a combustor while preventing the part of the air from flowing through a humidification tower and a recuperator. A turbine bypass valve forms another bypass system for enabling a part of a humidified air to be discharged to an exhaust duct after a heat exchange between the humidified air and a turbine exhaust gas is performed in the recuperator while preventing the part of the humidified air from flowing through a turbine. A turbine control system controls the bypass valves to be opened and closed in accordance with increase and decrease of the load.

Abstract: A system for providing an oxidizer and a fuel to a rocket engine is provided. The system includes a fuel supply system. The fuel supply system includes a fuel pump that pumps fuel to the rocket engine. The system includes a coolant supply system that supplies a coolant to the rocket engine, and a power plant that powers at least one of the fuel supply system and the coolant supply system. The power plant is powered by energy received from the coolant system.

Abstract: A method for assembling a gas turbine engine includes fabricating a heat exchanger that includes a first manifold including an inlet and an outlet, a first quantity of heat exchanger elements coupled in flow communication with the manifold inlet, a second quantity of heat exchanger elements coupled in flow communication with the manifold outlet, and a plurality of channels coupled in flow communication with the first and second quantity of heat exchanger elements to facilitate channeling compressor discharge air from the first quantity of heat exchanger elements to the second quantity of heat exchanger elements, and coupling the heat exchanger assembly to the gas turbine engine such that the heat exchanger is positioned substantially concentrically with respect to a gas turbine engine axis of rotation, and such that the heat exchanger is configured to receive compressor discharge air and channel the compressor discharge air to the combustor.

Abstract: Method and arrangement for providing a component (1) for being subjected to high thermal load during operation. The component includes a wall structure, which defines an inner space for gas flow. The component is formed by at least a first part (5) that includes an inner wall (8), an outer wall (9) and at least one cooling channel (11) between the walls. An end portion of said inner wall of the first part of the component is joined to a second part (6). The joint (18) is located at a distance from the interior of the component.

Abstract: An improved engine construction, such as a rocket engine construction is provided. The engine construction comprises a combustion chamber, a smooth wall nozzle, and a transition zone between the chamber and the smooth wall nozzle. The transition zone has a coolant system which includes a manifold formed from a non-copper material through which a coolant flows. In a high heat transfer embodiment, the transition zone includes an additional manifold formed from a copper based material.

Abstract: The invention relates to a ventilation system for a convergent divergent exhaust nozzle in a bypass turbojet comprising an afterburn chamber surrounded by an annular passage through which circulates a stream of cooling air, a convergent divergent axisymmetric nozzle arranged downstream of said afterburn chamber, each circle of flaps comprising alternately a plurality of controlled flaps, and a plurality of follower flaps, a circle of cold flaps arranged radially outside said nozzle and hinged at their upstream end to a conical shell linked to the downstream part of the casing.

Abstract: A combustion chamber, in particular for a rocket drive, includes at least one jacket madeof a composite material with a ceramic matrix. The composite material contains a fibrous structure made of carbon-containing fibers, and the fibrous structure comprises layers of fibers that form a three-dimensional matrix.

Abstract: A pulsed combustion apparatus includes a conduit having an outer wall and an inner wall. The inner wall has a number of apertures. An interior space is separated from the outer wall by the inner wall. An induction system is positioned to cyclicly admit charges to the interior space. An ignition system is positioned to ignite the charges. Flow directing surfaces are positioned to at least cyclicly direct cooling air through the apertures.

Abstract: A fuel based thermal management system includes a fuel stabilization system which permits the fuel to exceed the traditional coking temperatures. High temperature components are arranged along the fuel flow path such that even at the higher operating temperatures the fuel operates as a heat sink to transfer heat from high temperature components to the fuel. An optimal high temperature ester-based oil permits an oil-loop to exceed current oil temperature limits and achieve a high temperature to permit efficient rejection of heat to the fuel late in the fuel flow path.

Abstract: A method for assembling a gas turbine engine including a compressor and a combustor includes providing a heat exchanger assembly that includes at least one heat exchanger, and coupling the heat exchanger assembly to the gas turbine engine such that the heat exchanger is positioned substantially concentrically with respect to a gas turbine engine axis of rotation, and such that the heat exchanger is configured to channel compressor discharge air from the compressor discharge air to the combustor.

Abstract: A forced air cooling system for an APU comprising at least an oil cooler, a plenum in fluid communication with the oil cooler, and a compressor rotated by a rotating shaft of the APU such that the compressor and rotating shaft rotate at a same speed, the rotating compressor inducing a cooling air flow through the oil cooler. A method for cooling an auxiliary power unit is also provided.

Abstract: A system for cooling a portion of a rocket engine with an inert compound and transferring the thermal energy from the inert compound to the propellants. The energy absorbed by the coolant is used also to power the turbines which powers the pumps that pump the fuel, the oxidizer, and the coolant. Additionally, the coolant, which is an inert compound, is used to separate the oxidizer and the fuel before the oxidizer and the fuel enter the combustion chamber eliminating the need for a complex inert turbo pump seal package. The systems which pump or comprise the coolant physically separate the propellants before the propellants enter the rocket engine. The coolant remains substantially unconsumed in this cycle.

Abstract: A device for supplying cooling air to a hollow flap of an exhaust nozzle in a turbojet engine. The device includes a tube connecting the hollow flap to a cooling air source. The tube comprises at least one telescopic portion and at least one ball-joint connection.

Abstract: An improved engine construction, such as a rocket engine construction is provided. The engine construction comprises a combustion chamber, a smooth wall nozzle, and a transition zone between the chamber and the smooth wall nozzle. The transition zone has a coolant system which includes a manifold formed from a non-copper material through which a coolant flows. In a high heat transfer embodiment, the transition zone includes an additional manifold formed from a copper based material.

Abstract: A system and method of desalinating brackish water from an inland, underground water supply is disclosed. The system includes a source of underground water, at least one distillation and/or non-distillation desalination stage, and a source of waste heat from a power generator. The method includes desalinating brackish water drawn from an underground source using waste heat from a power generator and one or more non-distillation and/or distillation desalination stages. Potable water is recovered from both the brackish underground water using waste heat from the power generator, while dissolved salts are taken to dryness and removed as solids and from the flue gases.

Abstract: A wall structure (2) configured to be exposed to a thermal load. The wall structure having at least two layers including a first layer (5) and a second layer (6). The second layer (6) is located closer to a source of the thermal load than the first layer (5), and the layers (5, 6) are arranged so that heat is allowed to be conducted from the second layer (6) to the first layer (5). Each of the first and second layers (5, 6) are adapted to carry a significant portion of a structural load, and the second layer (6) exhibits a higher thermal conductivity and/or a lower thermal expansion than the first layer (5). The invention reduces the thermal strain in the wall structure (2).

Abstract: A combination for connecting a hot sheet and a cold sheet of a nozzle liner in spaced relationship apart includes a stud having a main body wherein the stud is attached to the hot sheet and a portion of the stud extends through an opening in the cold sheet. A support spaces the hot sheet and the cold sheet apart. A generally planar fastening collar is affixed to the portion of the stud extending through the cold sheet. A first length is defined by the portion of the stud extending through the cold sheet to a distal end of the stud. A second length is defined between a first planar surface of the collar facing the cold sheet and a second planar surface of the collar opposite the first surface. Each of the first and second lengths is less is less than a diameter of the main body of the stud.

Abstract: A convergent-divergent rocket nozzle is formed by joining two coaxially aligned conical sections at a throat plane, each diverging outward from the throat plane. Coolant channels formed in the wall of the nozzle are arranged in spirals around the nozzle axis. Preferably, the conical sections are formed from platelet laminates rolled into conical form with a single spiral seam, and at least one of the conical sections is a composite of two or more component conical sections separately formed and then joined in a nested arrangement with the seams not superimposed. A further preferred construction is one in which the convergent end of one conical section is split radially into strips that are then spread apart to serve as bonding surfaces to bond to the other conical section.

Abstract: The invention concerns a cooling system for a post combustion jet nozzle, The jet nozzle includes a primary gas duct allowing a primary flow of gas, and a secondary air duct allowing a secondary flow of air. The secondary air duct surrounds the primary gas duct and is separated from it by a protective thermal shroud. The secondary air duct has a downstream end, and dampers surrounding an output section of the primary gas duct. The system includes a protective thermal shell in the secondary air duct, at the downstream end of said duct. The protective thermal shell bears an annular diaphragm extending out in front of the dampers and being provided with support sectors and inter-sector zones equipped with slots. The inter-sector zones define spaces between the diaphragm and the protective thermal shroud.

Abstract: A variable geometry exhaust nozzle 10 for a turbine engine includes convergent and divergent flaps 12, 18 that circumscribe a gaspath 26 and define convergent and divergent nozzle sections 30, 32 with a throat 34 therebetween. A projection 56, which resides on the divergent flaps, extends radially inwardly from the nozzle. The nozzle also includes a coolant flowpath comprising an interior space 44 in the nozzle, an intake 58 for admitting coolant C into the interior space, and a coolant outlet 60 aft of the intake for discharging the coolant from the interior space. The intake resides forward of the throat. During operation, a coolant film flows along the radially inner surface of the nozzle. The projection encourages a portion of the coolant film to enter and flow through the the coolant flowpath to convectively cool the nozzle.

Abstract: The invention is a coolant system for a rocket engine. The rocket engine includes an injector, a fuel supply, an oxidizer supply, a pump for feeding fuel from the fuel supply to the injector, a pump for feeding oxidizer from the oxidizer supply to the injector, a combustor, and a nozzle, the combustor and nozzle forming a combustor and nozzle assembly. The coolant system includes a cooling jacket surrounding the combustor and nozzle assembly. A coolant fluid is supplied to the combustor jacket by a coolant pump that circulates the coolant fluid through the jacket. The coolant fluid vaporizes to cool the combustor and nozzle assembly. A turbine is operatively connected to the coolant pump and is driven by the vaporized coolant from the jacket. A heat exchanger transfers heat from the vaporized fluid to fuel or oxidizer from the supply thereof prior to the feeding of the fuel or oxidizer into the injector.

Abstract: A fluid flow actuated valve for controlling the fluid flow in a conduit comprises a support structure adapted to be mounted within the conduit and at least one flap member pivotally mounted to the support structure. The at least one flap member is pivotable between an first position for a minimum fluid flow and a second position for a maximum fluid flow. A torsion spring is attached to the at least one flap member to urge the same to pivot against a fluid pressure differential, from the first position to the second position when the fluid pressure differential is smaller than a predetermined level.

Abstract: A variable geometry exhaust nozzle 10 for a turbine engine includes convergent and divergent flaps 12, 18 that circumscribe a gaspath 26 and define convergent and divergent nozzle sections 30, 32 with a throat 34 therebetween. A projection 56, which resides on the divergent flaps, extends radially inwardly from the nozzle. The nozzle also includes a coolant flowpath comprising an interior space 44 in the nozzle, an intake 58 for admitting coolant C into the interior space, and a coolant outlet 60 aft of the intake for discharging the coolant from the interior space. The intake resides forward of the throat. During operation, a coolant film flows along the radially inner surface of the nozzle. The projection encourages a portion of the coolant film to enter and flow through the the coolant flowpath to convectively cool the nozzle.

Abstract: A combustion chamber structure (1) comprising a plurality of cooling ducts (31, 41) extending in the longitudinal direction of the combustion chamber structure (1), whereby the combustion chamber structure (1) is rotationally symmetrical in relation to an axis (4) and comprised of an outer structure (11) and a base body (16) of the combustion chamber. The base body comprises ribs (25) preferably extending in the direction of the axis (4). The ribs are arranged in the peripheral direction of the combustion chamber structure and at least one first cooling duct (31) extends in the interior of the ribs.

Abstract: A hybrid rocket motor includes a storage tank which stores an oxidizer under relatively low pressure, a turbopump preferably directly coupled to an outlet of the storage tank which pressurizes the oxidizer to a relatively high pressure, a combustion chamber including a solid fuel, and an injector between the turbopump and combustion chamber through which the oxidizer is injected into the combustion chamber. According to a preferred aspect of the invention, the turbopump is operated by an expander cycle of a heat exchanger. According to another preferred aspect of the invention, the fluid flowing through the heat exchanger is oxidizer tapped from the storage tank. A barrier is maintained between an oxidizer feed line from the turbopump and the injector until sufficient pressure is created by the turbopump to pump the oxidizer at the requisite pressure into the injector.

Abstract: A combustion chamber for a rocket engine expels a hot stream of gas and has a cooling device. The inner wall of the combustion chamber adjoins the cooling device and contains depressions formed in such a way that the stable outer layer of the stream of gas that forms in the proximity of the inner wall of the combustion chamber during operation of the combustion chamber is destabilized in the area of the depressions.

Abstract: The present invention relates to a scram jet engine for use with a hypersonic vehicle. The scram jet engine has an upper boundary wall, a lower boundary wall, and a plurality of side walls defining an inner air flowpath. The walls are formed by a plurality of tubular heat exchanger panels. Each of the heat exchanger panels comprises a plurality of structural panels, each having a plurality of cooling passages, joined together. In the scram jet engine of the present invention, fuel may be used as a coolant and supplied to the cooling passage.

Abstract: The invention concerns a cooling system for a post combustion jet nozzle (1), said jet nozzle comprising a primary gas duct (2) allowing a primary flow of gas, a secondary air duct (3) allowing a secondary flow of air, said secondary air duct surrounding the primary gas duct and separated from it by a protective thermal shroud (4), said secondary air duct having a downstream end (3a), dampers (5) surrounding an output section of the primary gas duct, the system comprising a protective thermal shell (7) in the secondary air duct, at the downstream end of said duct. According to the invention, the protective thermal shell bears an annular diaphragm (8) extending out in front of the dampers and being provided with support sectors (11) and inter-sector zones equipped with slots, said inter-sector zones defining spaces between the diaphragm and the protective thermal shroud.

Abstract: A centerbody for a gas turbine engine includes a thermal control system which minimizes the thermal stresses between the centerbody and at least one stiffener. The centerbody stiffener extends radially inward from a centerbody shell. A cavity is defined within each centerbody stiffener. The thermal control system includes a plurality of openings circumferentially disposed around the centerbody. Each opening extends through the centerbody shell into each cavity. The openings include pairs of entrance openings and exit openings which permit circumferential flow to develop within the centerbody cavity.

Abstract: A rocket propulsion unit with an outer casing and an inner casing is described whereby the inner casing is arranged with space between the outer casing, and the inner casing forms a combustion chamber and has a contour adapted to the expulsion of propellants out of the combustion chamber with a constriction for forming a combustion chamber neck. The outer casing in contrast has a contour deviating from the contour of the inner casing.

Abstract: Cooling air passages are formed in the wall of an exhaust casing connected to a turbine casing of a gas turbine. Low pressure air extracted from the low pressure stage of an air compressor of the gas turbine is supplied to the cooling air passage from the portion near the downstream end of the exhaust casing. Cooling air flows through the cooling air passage toward the upstream end of the exhaust casing and then flows into an annular cavity formed in the turbine casing near the portion corresponding to the last turbine stage. Therefore, the metal temperature of the exhaust casing near the upstream end (near the joint between the exhaust casing and the turbine casing) is lowered by the cooling air and, as cooling air of a relatively high temperature is supplied to the cavity in the turbine casing, the metal temperature of the turbine casing near the downstream end becomes higher than that provided by a conventional cooling system.

Abstract: A liner for a gas turbine engine is provided that includes a first liner section and a second liner section. The first liner section includes a first flange having a first contact surface. The second liner section includes a second flange having a second contact surface and a plurality of apertures. The first and second flanges axially overlap one another, and in a circumferential liner the second flange is disposed radially outside of the first flange. A channel is formed by the two liner sections that is open to the core gas path. In a first position, the first flange axially overlaps the second flange by a first distance and the apertures are misaligned with the first flange and disposed within the channel. Cooling air entering apertures within the second flange subsequently passes into the channel. In a second position, the first flange axially overlaps the second flange by a second distance.

Abstract: A thermal management system avoids problems associated with the recirculation of fuel through a fuel tank on an aircraft in a system that includes a fuel reservoir (70), a pump (72) for pumping fuel from the reservoir (70) to a first heat load (74) and then to a check valve (76) to a junction (78). At least one second heat load (80), (82), (84) is connected to the junction (78) and a fuel/fan air heat exchanger (104) having a fan air flow path in heat exchange relation with a fuel flow path having a fuel discharge end (106) connected to the junction (78). The fan air flow path is adapted to be connected to receive compressed air from the fan (30) of an engine (10). A bleed air/fuel heat exchanger (92) receives fuel from the second heat load and bleed air from the engine (10) and is connected to deliver fuel to the engine (10).

Abstract: Turbojet powerplant with at least one compressor (1), at least one combustion chamber (2), a high-pressure turbine (3) and a low-pressure turbine (4), characterised in that a heat exchanger (5) is arranged between the compressor (1) and the combustion chamber (2), in that at least one hot-gas line (6) branches off from an area downstream of the high-pressure turbine (3) and is connected to the heat exchanger (5), and in that at least one cold-gas line (7) connects the heat exchanger (5) with an area upstream of the low-pressure turbine (4).