Class / Patent application number | Description | Number of patent applications / Date published |
060266000 | Including heat exchange means | 36 |
20080271433 | LOW PROFILE BLEED AIR COOLER - 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. | 11-06-2008 |
20080314019 | Engine cooling | 12-25-2008 |
20090025366 | PROPULSION UNIT FOR AIRCRAFT AND AIRCRAFT COMPRISING AT LEAST ONE SUCH PROPULSION UNIT - 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. | 01-29-2009 |
20090205314 | Combustion Chamber Wall - The invention relates to a combustion chamber wall in a combustion chamber, comprising a support structure with a number of heat shield which are attached to the support structure with the aid of a fastening means. Said fastening means encompasses a spring device which is fixedly connected to the support structure and is composed at least of a spring that is mounted in a jacket. The spring is secured against falling out on the side facing away from the combustion chamber by means of a first removable retainer. The interior of the spring device is provided with an axial hole for accommodating the fastening means. The spring is secured against falling out on the side facing the combustion chamber by means of a second removable retainer. The spring device can be dismounted on the side facing the combustion chamber. | 08-20-2009 |
20090301057 | TURBOREACTOR FOR AIRCRAFT - The invention concerns a turbojet for aircraft including engine located in nacelle, and thermal exchanger intended to cool a fluid participating in the engine propulsive system, characterized in that said thermal exchanger is located on engine external wall, an interstitial space within which air can circulate being arranged between the engine external wall and a lower wall of said thermal exchanger. The invention also concerns an aircraft provided with at least one such turbojet. | 12-10-2009 |
20100000200 | IMPINGEMENT COOLING DEVICE - An impingement cooling sleeve includes a sleeve body having an inner surface to face a transition duct and an outer surface facing opposite the inner surface. At least one cooling hole is formed within the sleeve body and is used to direct cooling air toward the transition duct. At least one conduit member is attached to the sleeve body and is associated with the at least one cooling hole. The conduit member has a first opening to define an air inlet and a second opening to define an air outlet. In one example, the first opening is spaced apart from the outer surface of the sleeve body by a distance. In one example, the first opening comprises an annular end face surface that defines a plane that is obliquely orientated relative to the outer surface of the sleeve body. | 01-07-2010 |
20100043396 | GAS TURBINE ENGINE FAN BLEED HEAT EXCHANGER SYSTEM - A heat exchanger system for a gas turbine engine includes: (a) a fan having at least two stages of rotating fan blades surrounded by a fan casing, the fan operable to produce a flow of pressurized air at a fan exit; (b) at least one heat exchanger having a first flowpath in fluid communication with the fan at a location upstream of the fan exit; and (c) a fluid system coupled to a second flowpath of the at least one heat exchanger. The first and second flowpaths are thermally coupled to each other. | 02-25-2010 |
20100101209 | SYSTEM AND METHOD FOR CHANGING THE EFFICIENCY OF A COMBUSTION TURBINE - An embodiment of the present invention takes the form of an application and process that incorporates an external heat source to increase the temperature of the airstream entering a compressor section of a combustion turbine. An embodiment of the present invention may perform an anti-icing operation that may not require an Inlet Bleed Heat system (IBH) to operate. An embodiment of the present invention may perform an anti-icing operation that may allow for the IGV angle to remain nearly constant. An embodiment of the present invention may increase the output and efficiency of a combustion turbine operating at partload by delaying IBH operation and delaying the closing IGVs. | 04-29-2010 |
20100146934 | Hot Gas Chamber - A hot gas chamber, such as a combustion chamber for a power unit for the discharge of a hot gas flow, particularly for a rocket propulsion unit, has a combustion chamber wall, which has an internal surface and cooling ducts adjoining the latter. The internal surface of the combustion chamber wall facing the combustion chamber is further developed at least in areas for avoiding a film condensation. | 06-17-2010 |
20100162685 | A METHOD FOR COOLING AIR AND DEVICES - A method of constructing self-powered air-conditioner comprises a convergent divergent nozzle where powered fan pushes air into said nozzle. While the pushed air accelerates toward the nozzle throat it becomes colder as air internal energy transformed into kinetic energy. An axial turbine installed within the nozzle throat extracts energy from the air in the nozzle and drives an electrical generator that provides electricity to the fan electric motor. Alternatively the turbine and fan are installed on common shaft, which could be the electric generator shaft. The cold air within the nozzle throat cools the nozzle throat skin, which serves as air-conditioner core. The cold nozzle skin is wrapped with coiled pipes in which liquid flows, becomes colder and this cold liquid flows away to heat exchanger where air is flowing through it and becomes colder. This cold air is then flows into spaces needed to be air-conditioned. | 07-01-2010 |
20100180575 | ENGINE WALL STRUCTURE AND A METHOD OF PRODUCING AN ENGINE WALL STRUCTURE - An engine wall structure includes an inner wall to which hot gas is admitted during engine operation, an outer wall, which is colder than the inner wall during engine operation, and at least two webs that connect the inner wall with the outer wall and delimit a cooling duct between the walls. The webs are mainly formed by a first material and the inner wall is mainly formed by a second material of other composition and other heat conductivity than the first material. | 07-22-2010 |
20100236217 | HEAT TRANSFER SYSTEM AND METHOD FOR TURBINE ENGINE USING HEAT PIPES - A heat transfer system is provided for a turbine engine of the type including an annular casing with an array of thermally conductive, generally radially-extending strut members disposed therein. The heat transfer system includes at least one arcuate heat pipe disposed in contact with an outer surface of the casing within fore-and-aft limits of the axial extent of the strut members. The heat pipe is thermally coupled to a heat source, such that heat from the heat source can be transferred through the heat pipe and the casing to the strut members. | 09-23-2010 |
20100242437 | FUEL-COOLED FLEXIBLE HEAT EXCHANGER WITH THERMOELECTRIC DEVICE COMPRESSION - An apparatus includes a thermoelectric (TE) device, a gas flow conduit proximate to one side of the thermoelectric device, a plurality of flexible tubes proximate to a second side of the thermoelectric device, and a spring to control contact force between the flexible tubes and the thermoelectric device. The spring comprises a coil spring at least partially circumscribing the gas flow conduit. The thermoelectric device converts a temperature differential between the flexible tubes and the gas flow conduit into electrical energy. | 09-30-2010 |
20100275577 | ROCKET ENGINE INJECTORHEAD WITH FLASHBACK BARRIER - Propellants flow through specialized mechanical hardware that is designed for effective and safe ignition and sustained combustion of the propellants. By integrating a micro-fluidic porous media element between a propellant feed source and the combustion chamber, an effective and reliable propellant injector head may be implemented that is capable of withstanding transient combustion and detonation waves that commonly occur during an ignition event. The micro-fluidic porous media element is of specified porosity or porosity gradient selected to be appropriate for a given propellant. Additionally the propellant injector head design integrates a spark ignition mechanism that withstands extremely hot running conditions without noticeable spark mechanism degradation. | 11-04-2010 |
20100326049 | COOLING SYSTEMS FOR ROTORCRAFT ENGINES - A cooling system is provided for an engine of a rotorcraft configured to generate a downwash into an atmosphere during operation. The system includes an inlet coupled to the engine and configured to receive a liquid from the engine; an outlet coupled to the engine and configured to return the liquid to the engine; and a body defining a conduit having a first end coupled to the inlet and a second end coupled to the outlet such that the liquid flows from the inlet to the outlet through the conduit and transfers heat to the atmosphere via the downwash from the rotorcraft. | 12-30-2010 |
20110214410 | PROPELLANT TANK AND VAPOR JET EMITTING DEVICE INCLUDING SAME - 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. | 09-08-2011 |
20110265448 | JET NOZZLE MIXER - 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. | 11-03-2011 |
20120240551 | HEAT ENGINE - 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. | 09-27-2012 |
20160069265 | AIR GUIDING DEVICE AND AIRCRAFT ENGINE WITH AIR GUIDING DEVICE - An air guiding device in an aircraft engine, comprising at least one connection device of a core engine shroud with an external wall of a bypass duct of the aircraft engine is provided. At least one first air inlet opening for inflowing air is connected to a connection device or is arranged inside the connection device. | 03-10-2016 |
20190145317 | GAS TURBINE ENGINE HAVING AN AIR-OIL HEAT EXCHANGER | 05-16-2019 |
060267000 | For a liquid | 16 |
20080245054 | Cooling System for an Aircraft, Aircraft Comprising the Cooling System and Cooling Method - A cooling system for an aircraft includes a cooling circuit for transporting a coolant, with the cooling circuit running to a component in the aircraft that is heated up during flying in order to take up heat from the heated-up component via the coolant. The cooling circuit runs to a jet engine for propulsion of the aircraft in order to release heat to a flow of gas in the jet engine via the coolant. | 10-09-2008 |
20080264035 | COOLANT FLOW SWIRLER FOR A ROCKET ENGINE - A thrust chamber assembly cooling system includes a twisted ribbon/wire of any one of many variety of cross-sectional shapes located within a nozzle assembly cooling passage to direct the coolant to flow in a swirling manner which induces mixing and breaks up the boundary layer in the coolant passage to enhance the convective heat transfer. | 10-30-2008 |
20080276596 | Engine Wall Structure and a Method of Producing an Engine Wall Structure - An engine wall structure includes an inner wall to which hot gas is admitted during engine operation, an outer wall, which is colder than the inner wall during engine operation, and at least two webs that connect the inner wall with the outer wall and delimit a cooling duct between the walls. The webs are mainly formed by a first material and the inner wall is mainly formed by a second material of other composition and other heat conductivity than the first material. | 11-13-2008 |
20090288390 | SIMPLIFIED THRUST CHAMBER RECIRCULATING COOLING SYSTEM - In some implementations a propulsion system includes a thrust chamber having a gap between an inner shell and an outer shell, the inner shell and the outer shell being attached together to form the thrust chamber. In some implementations, the rocket engine also includes a recirculating cooling system operably coupled to the gap in at least two locations and operable to recirculate a convective coolant through the gap. | 11-26-2009 |
20090308051 | HEAT EXCHANGER TUBE AND AIR-TO-AIR INTERCOOLER - An improved heat exchanger tube which can be used in an air-to-air intercooler or other air-to-air heat exchangers. The cooling tube has a first tube with an inner and outer surface and a second tube with an inner and outer surface. The second tube is located inside the first tube. There are one or more walls extending from the inner surface of the first tube to the outer surface of the second tube. A plurality of fins are located on the outer surface of the first tube. The fins can take the form of individual circular fins or one or more helical fins. | 12-17-2009 |
20100107603 | Systems and methods for thermal management in a gas turbine powerplant - A thermal management system for a gas turbine powerplant with an engine oil line and an engine fuel line incorporates a heat transfer control module that includes a reversible heat pump with a heat pump compressor for circulating working fluid in forward and reverse directions through a working fluid line of the heat pump. The heat control module also includes a first heat exchanger having a heat exchange path for the working fluid between the compressor and a heat pump expansion valve and another heat exchange path for the engine oil. A second heat exchanger has a heat exchange path for the working fluid between the compressor and the expansion valve and another heat exchange path for the engine fuel. The heat pump can be operated in forward or reverse directions depending on whether heat is to be transferred from the engine oil or the fuel to the heat pump working fluid. In another embodiment an engine oil reservoir located between the first heat exchanger and the engine collects the oil before it is introduced to the engine and thus acts as a heat capacitor for the system. | 05-06-2010 |
20100205933 | Regeneratively cooled porous media jacket - The fluid and heat transfer theory for regenerative cooling of a rocket combustion chamber with a porous media coolant jacket is presented. This model is used to design a regeneratively cooled rocket or other high temperature engine cooling jacket. Cooling jackets comprising impermeable inner and outer walls, and porous media channels are disclosed. Also disclosed are porous media coolant jackets with additional structures designed to transfer heat directly from the inner wall to the outer wall, and structures designed to direct movement of the coolant fluid from the inner wall to the outer wall. Methods of making such jackets are also disclosed. | 08-19-2010 |
20100263350 | APPARATUS AND METHOD FOR COOLING A TURBINE USING HEAT PIPES - The turbine section of the turbine engine is provided with a flow of cooling air which is taken from a compressor section of the turbine engine. The air received from the compressor section is itself cooled before the air is delivered to the turbine. Heat is removed from the flow of air by a plurality of heat pipes which conduct heat away from the flow of air to lower the temperature of the air before it is provided to the turbine. | 10-21-2010 |
20100275578 | COOLING EXCHANGER DUCT - 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. | 11-04-2010 |
20100300066 | TURBOJET ENGINE FOR AIRCRAFT - A turbojet engine for an aircraft that includes an engine provided in a nacelle and at least one heat exchanger for cooling down a hot fluid collected in the propulsion system of the turbojet engine before re-injecting the aforementioned partially-cooled hot flow into the aforementioned propulsion system, wherein at least one heat exchanger is a radial heat exchanger extending in the lower portion of the turbojet engine at a lower branching of the turbojet engine. | 12-02-2010 |
20100300067 | COMPONENT CONFIGURED FOR BEING SUBJECTED TO HIGH THERMAL LOAD DURING OPERATION - A component configured for being subjected to a high thermal load during operation includes a wall structure with a tubular shape, wherein the wall structure includes a plurality of cooling channels for handling a coolant flow. The wall structure is divided in a plurality of sectors in a circumferential direction of the wall structure. Each sector includes at least two cooling channels and the wall structure is configured to prevent coolant flow communication between the cooling channels in adjacent sectors. | 12-02-2010 |
20110192137 | Method for Manufacturing a Regeneratively Cooled Nozzle Extension of a Rocket Combustion Chamber and Nozzle Extension - A method for manufacturing a regeneratively cooled nozzle extension of a rocket combustion chamber, the nozzle extension including a first wall and a second wall arranged coaxially to each other and between which a number of cooling channels is configured that are laterally delimited by cooling channel webs. The first wall and the second wall are connected to each other by a positive fit by cooling channel webs of the first wall engaging with corresponding recesses of the second wall for forming the positive fit. The positive fit is produced by a forming process in the region of the cooling channels of the second wall having the recesses. | 08-11-2011 |
20120090292 | COMBUSTION CHAMBER COMPRISING A CONDENSATION-PROOF BARRIER ON A REGENERATIVE CIRCUIT - The invention concerns a combustion chamber ( | 04-19-2012 |
20130160427 | COOLING DEVICE FOR COOLING COMBUSTION GASES FROM RECOILLESS ANTI-TANK WEAPONS - The present invention relates to a cooling device ( | 06-27-2013 |
20130232950 | Exit Manifold Flow Guide - An exit manifold is disclosed which includes a manifold body that includes a plurality of inlets. The manifold body provides communication between the inlets and a discharge port. At least one of the inlets directs flow in a first direction and at least one of the inlets directs flow in a second direction. The first and second directions are opposite and the material flowing in these opposite directions collides in front of the discharge outlet. The collision of these two oppositely-directed flows creates a high pressure stagnation region that may block or impede flow from one or more inlets that may be in alignment with the discharge port. | 09-12-2013 |
20160177874 | DEVICE FOR PRESSURIZING PROPELLANT TANKS OF A ROCKET ENGINE | 06-23-2016 |