|Combination engines for aircraft - An engine combination for generating forces with a gas turbine engine generating force, and an internal combustion engine provided in the combination as an intermittent combustion engine generating force having an air intake, there being an air transfer duct connected from a compressor in the gas turbine engine to the air intake to transfer compressed air thereto.
|Stator assembly including bleed ports for turbine engine compressor - A stator assembly for a turbine engine is disclosed. The stator assembly has a stator support ring, a plurality of stator vanes protruding radially inwardly from the stator support ring, and at least one bleed port disposed on the stator support ring between at least two adjacent stator vanes of the plurality of stator vanes. The at least one bleed port defines an oblong profile having a first end terminating near a leading edge of a first stator vane of the at least two adjacent stator vanes, and a second end, substantially opposed from the first end, terminating near a trailing edge of a second stator vane of the at least two adjacent stator vanes. The profile may be concave toward a suction sidewall of the second stator vane of the at least two adjacent stator vanes.
|Air-bleed gas turbine - An air-bleed gas turbine includes a compressor; a combustor; a turbine; and an air bleed portion by which compressed air, which is supplied from the compressor to a circular space that surrounds the combustor, is partially supplied to a device outside the air-bleed gas turbine. The air bleed portion includes an air bleed passage that is communicated with an air bleed port which is connected to the device outside the air-bleed gas turbine, and a communication passage that is communicated with the circular space. The air bleed passage is communicated with the circular space through the communication passage. The flow passage area of the communication passage is increased as the distance from the air bleed port increases in the circumferential direction of the circular space.
|Dual Flow Turbine Engine Equipped with a Precooler - The invention concerns a precooler (
|TIP TURBINE ENGINE WITH ASPIRATED COMPRESSOR - A tip turbine engine includes an axial compressor having a plurality of airfoils compressing core airflow. The airfoils include bleed air openings on their suction side surfaces. The bleed air openings prevent separation of the compressed airflow, which permits each airfoil stage to perform increased compression without separation of the airflow. As a result, the number of stages can be reduced, thereby shortening the overall length of the turbine engine.
|Jet engine with compressor air circulation and method for operating the jet engine - In a jet engine with compressor air circulation provided for stabilizing the flow conditions, the compressed hot compressor air tapped from the flow path of the compressor (
|APPARATUS AND METHOD FOR SUPPRESSING DYNAMIC PRESSURE INSTABILITY IN BLEED DUCT - An apparatus and method for suppressing dynamic instability in a bleed duct of a gas turbine engine includes a fan bypass duct configured to permit a flow of air through the gas turbine engine. The fan bypass duct defines a fan duct surface, and the bleed duct has an inlet in fluid communication with the fan bypass duct and a flow control valve having an opened position and a closed position. A flow diverter is positioned on the fan duct surface proximate the inlet of the bleed duct and diverts the flow of air from the inlet when the flow control valve is closed, while permitting a portion of the flow of air to enter the inlet when the flow control valve is opened.
|ANTI-ICING APPARATUS AND METHOD FOR AERO-ENGINE NOSE CONE - An aero-engine nose cone anti-icing system (
|Re-pressurisation device - A gas turbine engine re-pressurisation device (
|Aeroengine bleed valve - A bleed valve bleeds hot gases from a compressor into a bypass duct of a gas turbine engine. The valve comprises a diffuser having opposing walls and a divider located between the walls and which define at least two passages through which a bleed fluid flows and into a fluid flow through the bypass duct. The passages are angled towards each other to force the two gas flows together to form a high aspect ratio plume. This plume has a relatively large surface area that that improves mixing with the bypass flow to cool it and prevented otherwise hot bleed flows from impinging on heat sensitive components adjacent the bypass duct.
|Bleed Air Supply System and Method to Supply Bleed Air to an Aircraft - The invention relates to a bleed air supply system and a safety device for closing a valve in at least one of the several bleed air supply lines depending on an alarm signal that indicates an opening in at least one of the several bleed air supply lines. To stop the supply of bleed air irrespective of the type of the bleed air source, the safety device comprises a filter device for generating the alarm signal depending on the ground connection signal and on an opening signal from an opening signal device that indicates any opening of the bleed air supply line that is connected to the high-pressure ground connection.
|UTILIZING INLET BLEED HEAT TO IMPROVE MIXING AND ENGINE TURNDOWN - An improved inlet bleed heat system for a gas turbine engine is disclosed. The inlet bleed heat system provides improved mixing in an inlet region while permitting the engine to be operated at lower power settings. The inlet bleed heat system comprises a supply conduit, a plurality of feed tubes extending from the supply conduit, and a guide tube for receiving opposing ends of the feed tubes. The plurality of feed tubes each have a plurality of injection orifices and the feed tubes are oriented such that the injection orifices generally face into a flow of oncoming air with the feed tubes being positioned forward of a plurality of sound attenuating baffles.
|TURBINE BLADE TIP CLEARANCE APPARATUS AND METHOD - A method for adjusting a clearance between a blade tip of a turbine engine and a blade track spaced radially outward of the blade tip is disclosed herein. The method includes the step of operably coupling an elongate member to a blade track in a turbine engine. The method also includes the step of directing a fluid stream having a temperature in proximity to the elongate member. The temperature of the fluid stream can change over time. The method also includes the step of transferring heat between the fluid stream and elongate member to a change a size of the elongate member and move the blade track radially relative to a centerline axis of the turbine engine. An exemplary apparatus for carrying out the method is also disclosed.
|TURBINE SHROUD THERMAL DISTORTION CONTROL - A shroud suitable for use in a gas turbine engine exhibits substantially uniform thermal growth.
|Gas turbine unit and its cooling method - It is the object of the present invention to feed cooling air suitable for cooling the high-temperature part of the gas turbine.
|Heat exchanger, propulsion unit and aircraft provided therewith - A heat exchanger provided with a heat exchanger housing for circulating a cooling air flow and a hot air flow therein including a hot air input channel placed on the housing rear side for supplying the hot air flow to the heat exchanger housing and a cooling air discharge channel arranged on the rear side of said heat exchanger housing for exhausting the cooling air flow therefrom, wherein the channels are concentrically shaped. The disclosed embodiments also relate to a propulsion unit provided with the inventive heat exchanger, wherein the cooling air discharge channel arranged in the rear part of the housing crosses the strut in such a way that the cooling air flow is exhausted behind a jet engine. An aircraft provided with at least one inventive propulsion unit is also disclosed.
|Gas-turbine compressor with bleed-air tapping - A gas-turbine compressor has a casing
|TURBOPROP ENGINE WITH AN APPARATUS FOR THE GENERATION OF A COOLING AIRFLOW - A turboprop engine (
|FAST ACTING AIRCRAFT ENGINE BLEED AIR CHECK VALVE - An engine air supply system includes an air source, which is a gas turbine engine in one example. A component, such as an environmental control system component, is interconnected to the air source by a conduit that includes a supply flow direction. A check valve is arranged in the conduit. The check valve includes a valve portion moveable in the supply flow direction from a closed position to an open position. The closed position is arranged at an acute angle relative to the supply flow direction. The check valve includes first and second valve portions that are hemi-elliptical in shape. The first and second valve portions are non-planar and arranged approximately 90 degrees apart when in the closed position. The first and second valve portions move toward one another from the closed position to the open position. Each of the first and second valve portions moves approximately 30-40 degrees from the open position to the closed position in response to a stall condition, which significantly reduces the response time and forces generated by the first and second valve portions when closing.
|System for diffusing bleed air flow - In one aspect, the present disclosure is directed an apparatus configured to diffuse a flow of bleed air. The apparatus having an inlet collar configured to receive the flow of bleed air in a direction substantially along a longitudinal axis of the apparatus. The apparatus further having an end wall longitudinally spaced apart from the inlet collar and configured to block the flow of bleed air in a direction substantially along the longitudinal axis. The apparatus also having a first diffuser wall spaced concentrically relative to a second diffuser wall, each of the first and second diffuser walls positioned between the inlet collar and the end wall and including a plurality of perforations configured to permit the flow of bleed air to exit the apparatus at an angle relative to the longitudinal axis.
|COOLING AIR BLEED DEVICE IN A TURBINE ENGINE - Air bleed device for cooling components in a turbine engine, including an annular conduit (
|COOLING AIR BLEED DEVICE IN A TURBOMACHINE - Air bleed device for cooling components in a turbine engine, including an annular conduit (
|TURBINE WITH FLUIDICALLY-CONTROLLED VALVE AND SWIRLER WITH A BLEED HOLE - A fluidically-controlled valve is provided. The fluidically-controlled valve includes a main flow channel with a main flow entrance, a flow exit and a constricted channel section located between the main flow entrance and the flow exit. The fluidically-controlled valve also includes a control flow channel including a jet forming control entrance, a first branch channel, a second branch channel, a common channel section, and a convex channel wall. The common channel section follows the control entrance, the first branch channel emerges from the common channel section and leads to the main flow entrance, the second branch channel emerges from the common channel section and leads to the constricted channel section, and the convex channel wall extends from the common channel section into the first branch channel. The fluidically-controlled valve can be used in bypasses present in turbines or in swirlers of gas turbine combustors.
|Bleed valve arrangement - Problems can arise with regard to noise created by bleed valve arrangements (
|DIFFUSER WITH ENHANCED SURGE MARGIN - A diffuser for a centrifugal impeller assembly of a gas turbine engine includes a diffuser case having a plurality of vanes extending therein defining a plurality of circumferentially distributed angled passages in communication with an inlet space. Each vane includes a bleed port defined in a suction surface thereof, in proximity of the leading edge. The diffuser case includes a passive fluid communication defined at least partially through each one of the vanes between each bleed port and the inlet space upstream of the leading edge, such that air bled through the bleed ports is recirculated upstream of the leading edges to the inlet space to increase a surge margin of the diffuser.
|Bleed assembly - A noise reduction device
|Noise reduction device - A noise reduction device
|FLUSH INLET SCOOP DESIGN FOR AIRCRAFT BLEED AIR SYSTEM - A bleed air duct that preferably includes an inlet section configured to include a flush scoop and a louver. The louver is located and configured such that in the desired operating flow range of the duct, the fluid entering the flush scoop is disturbed and as a result creates a low pressure region downstream of the louver. The low pressure region substantially eliminates the generation of any pressure pulses and acoustic resonance also known as Helmholtz resonance.
|DEVICE FOR MOVING A PLURALITY OF HATCHES IN A GAS TURBINE ENGINE - A device is provided for moving a plurality of hatches, spaced at intervals from one another in a circumferential direction of a gas turbine engine, between a first position and a second position, wherein each of the hatches is arranged to keep an opening in a wall in a closed position, when in the first position, and to keep the opening in an open position, when in the second position. At least one of the hatches is provided with two connecting members which are mutually separated in the circumferential direction of the device, and the device includes a linkage for the movement, the linkage being connected to the hatch at the two mutually separated connecting members.
|HEAT EXCHANGER TO COOL TURBINE AIR COOLING FLOW - 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
|METHOD AND SYSTEM FOR OPERATING A COMBINED CYCLE POWER PLANT - An embodiment of the present invention may provide a variable speed booster, which receives air from a compressor of a gas turbine through an intercooler, to supply air at a relatively constant pressure to an air processing unit. An embodiment of the present invention may provide an external power source to energize the variable speed booster; which provides air at a relatively constant pressure to an air processing unit.
|TURBINE COOLING AIR FROM A CENTRIFUGAL COMPRESSOR - A gas turbine engine turbine cooling system includes an impeller and a diffuser directly downstream of the impeller, a bleed for bleeding clean cooling air from downstream of the diffuser, and one or more channels in fluid communication with the bleed. Each of the channels having a generally radially extending section followed by a generally axially aftwardly extending section terminating at an annular cooling air plenum connected to accelerators. The radially and axially aftwardly extending sections may be connected by a bend section of the cooling air channel and the axially aftwardly extending section may be angled radially inwardly going from the bend section to the cooling air plenum. Each of the cooling channels includes an inner wall formed by a forward end wall extending radially outwardly from an inner combustor casing, an annular cover covering a radially inner portion of the forward end wall, and the inner combustor casing.
|Gas Turbine - Disclosed is a gas turbine enabling turn down operation, of which the system and the operation can be simplified, and also the production cost and the maintenance cost can be reduced. Specifically disclosed is a gas turbine comprising: a compressor section (
|Variable Geometry Ejector - An ejector for a turbine engine is described herein. The ejector may include a variable geometry motive nozzle and a variable geometry mixing tube positioned downstream of the variable geometry motive nozzle.
|INLET BLEED HEAT SYSTEM THERMAL LAG MITIGATION USING COMPRESSOR INTERSTAGE BLEEDS - A system to mitigate thermal lag of start-up operations of a gas turbine is provided and includes a housing which is disposed upstream from and which is fluidly coupled to a compressor of the gas turbine, a first system coupled to an aft portion of the compressor and to the housing which, when activated, receives a first air supply from the compressor and delivers the first air supply to the housing, and a second system, coupled to the first system, which removes a second air supply from an intermediate portion of the compressor and delivers the second air supply to the first system prior to the activation of the first system.
|OPTIMIZATION OF LOW-BTU FUEL-FIRED COMBINED-CYCLE POWER PLANT BY PERFORMANCE HEATING - Disclosed is an optimized approach of using bleed-off of compressed air flow from a gas turbine compressor in a combined-cycle power plant and performance heating to augment plant performance. In one embodiment, a diverted portion of a by-product off gas and the bleed-off of compressed air flow are fired heated to produce a high temperature flue gas that is used to performance heat a pressurized mixture of fuel prior to being supplied to the gas turbine combustor.
|FLOW DISCHARGE DEVICE - A flow discharge device, such as a compressor bleed outlet discharging into a bypass duct of a gas turbine engine, comprises an outlet panel
|DIRECT TRANSFER AXIAL TANGENTIAL ONBOARD INJECTOR SYSTEM (TOBI) WITH SELF-SUPPORTING SEAL PLATE - An apparatus for cooling turbine blades in a turbine engine including a direct transfer axial tangential onboard injector (TOBI) for a turbine rotor and a self-supporting seal plate disposed on a rotating disk for the turbine engine. The TOBI includes a plurality of openings emanating a flow of cooling air. The self-supporting seal plate comprises a plurality of shaped cooling holes in fluid communication with the flow of cooling air emanating from the TOBI. The rotating disk includes a plurality of turbine blade slots formed therein. The plurality of cooling holes are in fluid communication with the plurality of turbine blade slots for directing the flow of cooling air to provide cooling to the plurality of turbine blades. The plurality of openings, the plurality of cooling holes and the plurality of turbine blade slots are in axial alignment and optimized to minimize radial and hoop stresses.
|Compressor for a gas turbine - An axial compressor providing a thermal adjustment of a housing of a compressor of a stationary gas turbine to the rotor is provided. A partial flow is decoupled from the compressor air flow for cooling gas turbine components. The contact of the partial flow decoupled from the compressor with the interior side of the housing is substantially limited, or even avoided, by a separating element in a collection chamber annularly encompassing the flow path in order to prevent the premature thermal heating of the gas turbine or of the housing during cold starting.
|AIRCRAFT PROPULSION ASSEMBLY COMPRISING HOT AIR BLEED SYSTEMS - An aircraft propulsion unit includes a nacelle (
|APU BLEED VALVE WITH INTEGRAL ANTI-SURGE PORT - A compressor bleed air valve in an auxiliary power unit (“APU”) modulates bleed air flow both for accessory pneumatic systems powered by the bleed air and for reducing or removing surge in the compressor.
|AIR FILTRATION SYSTEM FOR GAS TURBINE ENGINE PNEUMATIC SYSTEM - A gas turbine engine pneumatic system for controlling an air bleed valve includes a variable orifice device for controllably discharging air from the pneumatic system in order to reduce a maximum air pressure applied to the air bleed valve, to a predetermined level. A filter device is provided to filter the air discharged from the variable orifice.
|CONTINUOUS SLOT IN SHROUD - A compressor shroud has a continuous slot defined in a gaspath side surface thereof for communicating through the body of the shroud. Spaced-apart structural bridges span a rear portion of the slot in the shroud to maintain structural rigidity of the shroud. The method of forming the continuous slot is also described.
|Systems and Methods for Gas Turbine Combustors - A gas turbine system comprising, a diffuser operative to diffuse an airstream output from a compressor, a fuel nozzle operative to receive fuel and emit the fuel in a combustor, and at least one bleed duct operative to direct bleed air from down stream of the combustor to the fuel nozzle.
|AIR MANIFOLD IN A TURBOMACHINE - A turbomachine including an intermediate casing including, fastened to the end thereof, an outer casing of a high pressure compressor, and an air bleed mechanism bleeding air downstream from the stream through the compressor and including an outlet connected to an air reinjection mechanism reinjecting air upstream from the compressor via an annular manifold surrounding the inner wall of the intermediate casing and situated radially between the inner wall and an outer wall defining a secondary flow stream of the turbomachine.
|BLEED AIR TRANSFER TUBE - A fluid-conveying device including an inner tubular member with a circumferential end portion, a non-elastomeric ring received in a depression of the end portion, and an outer tubular member. The ring has a peripheral surface with a rounded contour defined along a longitudinal direction configured to remain out of the depression to engage the outer tubular member and facilitate sealing thereof.
|METHOD AND APPARATUS FOR CONTROLLING FLUID FLOW - A method and apparatus for controlling fluid flow is provided. The valve assembly includes a valve body including a longitudinal axis and a radially inner internal flow passage extending axially between an inlet opening and an outlet opening, a disk selectively positionable to modulate a flow of working fluid through the flow passage, and a radially outer external surface. The valve assembly further includes a first actuator positioned radially outboard of the external surface, a second actuator positioned radially outboard of the external surface, and a first linkage including a first actuator end configured to couple to the first actuator, a second actuator end configured to couple to the second actuator, and a rod end configured to couple to the disk, wherein an actuating force applied from each actuator to each respective actuator end is combined through the first linkage to the rod end.
|BLEED ASSEMBLY FOR A GAS TURBINE ENGINE - A bleed assembly for a gas turbine engine is provided. The assembly includes: a duct having an inlet and an outlet; a bleed valve that controls the flow of bleed fluid into the inlet; and a dome-shaped diffuser screen which covers the outlet. The diffuser screen has a plurality of through-holes for passage of the bleed fluid. Each through-hole has one or more nearest-neighbour through-holes at a nearest-neighbour spacing. At the periphery of the diffuser screen, the average nearest-neighbour spacing of the through-holes at a given radial distance from the centre of the diffuser screen increases with increasing radial distance.
|GAS TURBINE ENGINE SURGE MARGIN BLEED POWER RECUPERATION - The present invention provides in one embodiment of the present invention a surge margin power recuperation system. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for recuperating power from surge margin bleed air. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.
|Ejector-OBB Scheme for a Gas Turbine - A cooling circuit in a gas turbine includes an over board bleed (OBB) circuit bleeding air from a last stage of the compressor, an extracted air circuit extracting air from an upstream stage of the compressor, and an ejector receiving input from the OBB circuit and the extracted air circuit. Ejector outlet flow from the ejector is communicated to the turbine.
|BLED DIFFUSER FED SECONDARY COMBUSTION SYSTEM FOR GAS TURBINES - The present subject matter provides a system for modifying static pressure recoveries and emissions formation within a gas turbine. The system includes a bled diffuser positioned downstream from a compressor section of the gas turbine and a bleed duct extending from the bled diffuser. The bleed duct may be configured to direct bleed air from the pressurized airflow exiting the compressor section to a secondary combustion system located downstream from the main combustion system in a combustor. The bleed air flowing into the secondary combustion system may be mixed with fuel to form an air/fuel mixture.
|AIR CONTAMINATION DETECTION IN AN AIRCRAFT AIR SYSTEM - An aircraft air system includes a gas turbine engine, a bleed air duct directing compressed air bled from a compressor to an inner compartment within the aircraft, and an air contamination detector located downstream of the gas turbine engine compressor. The air contamination detector includes a visual indicator which detects the presence of a fluid contaminant within the bleed air.
|BEARING COMPARTMENT PRESSURIZATION AND SHAFT VENTILATION SYSTEM - An assembly and method for providing buffer air and/or ventilation air within a gas turbine engine, the assembly includes an accessory gearbox and a centrifugal compressor. The accessory gearbox is connected to a driven shaft of the gas turbine engine. The centrifugal compressor is driven by the accessory gearbox during operation of the gas turbine engine to receive and compress a bleed air to produce the buffer air and/or the ventilation air. The buffer air and/or the ventilation air is communicated within the gas turbine engine.
|FLOW MIXING VENT SYSTEM - A vent system is disclosed having a first flow stream flowing over a first surface in a flow path, a conduit that channels a second flow stream and an aero-chimney that is in flow communication with the conduit and located near the first surface wherein the aero-chimney has a body having an aerodynamic shape having a leading edge portion and a trailing edge portion such that the first flow stream flows around the aero-chimney near the first surface.
|GAS TURBINE ENGINE SYSTEM WITH BLEED AIR POWERED AUXILIARY ENGINE - One embodiment of the present invention is a unique gas turbine engine with a bleed air powered auxiliary engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for bleed air powered auxiliary engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.
|AIR BLEED HAVING AN INERTIAL FILTER IN THE TANDEM ROTOR OF A COMPRESSOR - The invention relates to a secondary air system for a compressor of centrifugal or mixed type including a rotor presenting an axis of rotation, said compressor being adapted to compress an oxidizer gas. The secondary air system includes an oxidizer gas bleed system arranged in the rotor.
|Inner Bleed Structure of 2-Shaft Gas Turbine and a Method to Determine the Stagger Angle of Last Stage Stator of Compressor for 2-Shaft Gas Turbine - An inner bleed structure of the 2-shaft gas turbine includes a slit for leading part of compressed air to a cavity is formed between a wall surface of a rotor wheel of the compressor equipped with a last stage rotor of the compressor which is connected to a first rotating shaft and end of an inner casing, and a bleed hole for leading part of compressed air after flowing down the last stage of the compressor to a cavity formed in the inner side of the inner casing at the downstream side of the last stage of the compressor.
|BLEED VALVE - A bleed valve having an inlet port and an outlet port. The bleed valve comprises mutually perpendicular radial, longitudinal and tangential axes arranged such that the outlet port is substantially in a plane defined by the longitudinal and tangential axes. The bleed valve also comprises at least one vane spanning the outlet port and having a first side and a second side, the first side being closer to the inlet port than the second side. The at least one vane defining a vector from its first to its second side that has non-zero radial, longitudinal and tangential components.
|CENTRIFUGAL COMPRESSOR WITH BLEED FLOW SPLITTER FOR A GAS TURBINE ENGINE - An impeller includes a plurality of vanes formed around a hub, each of the plurality of vanes defines an offset between a leading edge and a trailing edge.
|METHODS AND APPARATUS FOR ASSEMBLING A LOW NOISE EJECTOR MOTIVE NOZZLE - A method of assembling an ejector is provided, wherein the method includes providing a motive nozzle tip having a centerline axis and including a nozzle tip edge having at least one protrusion extending through a plane substantially normal to the centerline axis. The method also includes coupling the motive nozzle tip to the ejector.
|GAS TURBINE ENGINE AND HEAT EXCHANGE SYSTEM - One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique gas turbine engine heat exchange system. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and heat exchange systems for gas turbine engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.
|TURBOJET INCLUDING AN AUTOMATICALLY VARIABLE FLOW RATE BLEED CIRCUIT FOR COOLING AIR - Bleeding cooling air to cool a subassembly, e.g. such as a turbine, with automatic adjustment of the air flow section as a function of the speed of the engine. According to the invention, a shutter element is fastened to co-operate with a bleed hole, with the material that constitutes either the shutter element or the wall in which the hole is formed being of a type in which it is possible to create eddy currents, and a magnet is mounted to move past said arrangement.
|COMPRESSOR TIP CLEARANCE CONTROL AND GAS TURBINE ENGINE - One embodiment of the present invention is a unique compressor. Another embodiment of the present invention is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for blade tip clearance control for compressors and gas turbine engine compressors. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.
|GAS TURBINE ENGINE AND SYSTEM FOR MODULATING SECONDARY AIR FLOW - One embodiment of the present invention is a unique gas turbine system having a system for modulating secondary air flow. Another embodiment is a compressor vane stage with a plenum employed in providing secondary air flow. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for secondary air flow systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.
|PITOT TUBE WITH INCREASED PARTICLE SEPARATION EFFICIENCY - A pneumatically actuated valve includes a valve body, a pneumatic actuator, and a pitot tube. The valve body includes a valve disc positioned in a valve housing that defines a flow passage with a flow passage diameter. The pneumatic actuator rotates the valve disc in the valve housing. The pitot tube is connected to the pneumatic actuator and has a tip extending into the flow passage. The tip is spaced apart from the valve housing by at least 1/20 of the flow passage diameter. The tip is substantially free of downstream obstructions for a distance equal to at least ½ of the flow passage diameter when the valve is in an open position.
|AIRCRAFT GAS TURBINE - An aircraft gas turbine is constituted by accommodating a compressor (
|FLADE DUCT TURBINE COOLING AND POWER AND THERMAL MANAGEMENT - An aircraft compound cooling system includes a power thermal management system for cooling one or more aircraft components, an air cycle system, a vapor cycle system, and a turbine cooling circuit for cooling bleed air and cooling turbine components in a high pressure turbine in the engine.
|METHOD AND SYSTEMS FOR BLEED AIR SUPPLY - A method and system for an integrated ejector valve assembly is provided. The integrated ejector valve assembly includes a first valve assembly configured to control a flow of relatively lower pressure fluid from a first inlet port, a second valve assembly configured to control a flow of relatively higher pressure fluid from a second inlet port, a first actuation chamber configured to close the first valve assembly, a second actuation chamber configured to close the second valve assembly, and a third actuation chamber configured to open the second valve assembly.
|FLOW DISCHARGE DEVICE - A bleed flow discharge device (
|GAS TURBINE INLET HEATING SYSTEM - A gas turbine inlet heating system is disclosed. In one embodiment, the system includes: a compressor having: an inlet bellmouth adjacent to a set of inlet guide vanes (IGVs); and an outlet fluidly connected to the inlet bellmouth; and a conduit coupled to an outlet of the compressor, the conduit including a control valve, the conduit for diverting a first portion of compressed air from the outlet of the compressor to the inlet bellmouth.
|PRESSURIZED AUXILIARY POWER UNIT LUBRICATION SYSTEM - An exemplary lubrication system includes a flow path that carries fluid from a fluid supply through at least a portion of an auxiliary power unit to an outlet separate from the fluid supply. The fluid supply is pressurized to move the fluid along the flow path.
|GAS TURBINE ENGINE COOLING SYSTEMS HAVING HUB-BLEED IMPELLERS AND METHODS FOR THE PRODUCTION THEREOF - Embodiments of a gas turbine engine cooling system for deployment within a gas turbine engine are provided, as are embodiment of a method for producing a gas turbine engine cooling system. In one embodiment, the gas turbine engine cooling system includes an impeller having a hub, a plurality of hub bleed air passages, and a central bleed air conduit. The plurality of hub bleed air passages each have an inlet formed in an outer circumferential surface of the hub and an outlet formed in an inner circumferential surface of the hub. The central bleed air conduit is fluidly coupled to the outlets of the plurality of hub bleed air passages and is configured to conduct bleed air discharged by the plurality of hub bleed air passages to a section of the gas turbine engine downstream of the impeller to provide cooling air thereto.
|COMPRESSOR BLEED COOLING FLUID FEED SYSTEM - A compressor bleed cooling fluid feed system for a turbine engine for directing cooling fluids from a compressor to a turbine airfoil cooling system to supply cooling fluids to one or more airfoils of a rotor assembly is disclosed. The compressor bleed cooling fluid feed system may enable cooling fluids to be exhausted from a compressor exhaust plenum through a downstream compressor bleed collection chamber and into the turbine airfoil cooling system. As such, the suction created in the compressor exhaust plenum mitigates boundary layer growth along the inner surface while providing flow of cooling fluids to the turbine airfoils.
|COOLNG STRUCTURE FOR RECOVERY-TYPE AIR-COOLED GAS TURBINE COMBUSTOR - In a cooling structure for a recovery-type air-cooled gas turbine combustor having a recovery-type air-cooling structure that bleeds, upstream of the combustor, and pressurizes compressed air supplied from a compressor, that uses the bled and pressurized air to cool a wall, and that recovers and reuses the bled and pressurized air as combustion air for burning fuel in the combustor together with a main flow of the compressed air, wall cooling in which cooling air is supplied to cooling air passages formed in the wall of the combustor to perform cooling involves a downstream wall region, closer to a turbine, that is cooled using the bled and pressurized air as the cooling air and an upstream wall region, closer to a burner, that is cooled using, as the cooling air, bled compressed air bled from a main flow of the compressed air through a housing inner space.
|METHOD FOR OPTIMIZING THE OPERABILITY OF AN AIRCRAFT PROPULSIVE UNIT, AND SELF-CONTAINED POWER UNIT FOR IMPLEMENTING SAME - A method for optimizing operability of an aircraft propulsive unit, and a self-contained power unit implementing the method. The method removes mechanical bleed constraints in engines during transient flight phases of an aircraft to optimize operability of the engine assembly during the phases. To this end, a supply of power is provided, particularly during the phases, by an additional indirectly propulsive engine power source. The method for optimizing operability of the propulsive unit of an aircraft including main engines as main drive sources includes, using a main engine power unit GPP as a power source, providing all the non-propulsive power and, during the transient engine phases, at most partially providing additional power to the body of the main engines.
|ENVIRONMENTAL CONTROL SYSTEM FOR AIRCRAFT UTILIZING TURBO-COMPRESSOR - An environmental control system includes a higher pressure tap to be associated with a higher compression location in a main compressor section associated with a gas turbine engine. A lower pressure tap is associated with a lower pressure location, which is at a lower pressure than the higher pressure location. The lower pressure tap communicates to a first passage leading to a downstream outlet and a second passage leading into a compressor section of a turbocompressor. The higher pressure tap leads into the turbine section of the turbocompressor such that air in the higher pressure tap drives the turbine section to in turn drive the compressor section of the turbocompressor. A combined outlet of the compressor section of the turbocompressor and the turbine section intermix and pass downstream to be delivered to an aircraft use.
|GAS TURBINE ENGINE BEARING CHAMBER SEALS - A gas turbine engine sealing air supply system comprising a bearing chamber seal to prevent lubricant fluid loss from a fluid chamber, the sealing effected by ingress of sealing air. An air supply duct is coupled to the bearing chamber seal to provide the sealing air. A first duct is coupled between a starter air system of the gas turbine engine and the air supply duct to supply sealing air during gas turbine engine starting.
|METHOD OF USING EXTERNAL FLUID FOR COOLING HIGH TEMPERATURE COMPONENTS OF GAS TURBINE FOR A PROCESS POWER PLANT - An external fluid in a closed loop is used to cool hot gas path components of gas turbine. After cooling the turbine components, the heated external fluid is dumped either in the compressor discharge casing or in the one of the turbine's stages. Where the external fluid is nitrogen to be dumped in the turbine compressor's discharge casing, the nitrogen is compressed using diluent nitrogen compressors. Alternatively, where the external fluid is nitrogen to be dumped in one of the stages of the turbine, the nitrogen is not compressed at all. The turbine blade heat exchangers in the turbine stages through which the nitrogen passes can be connected in parallel or in series for cooling the hot gas path components in the turbine stages. The nitrogen can optionally be mixed with air or steam or not mixed at all.
|GAS TURBINE ENGINE IN-BOARD COOLED COOLING AIR SYSTEM - A system for supplying turbine cooling air flow includes a turbofan engine, a heat exchanger, and a door. The turbofan engine includes an engine case that has an inner volume within which at least a gas turbine engine is mounted, and a bypass flow passage that is defined by an outer fan duct and an inner fan duct and that is configured to direct fan air flow therethrough. The heat exchanger is disposed within the turbofan engine, is coupled to receive fluid and cooling air from the bypass flow passage, and is configured to transfer heat between fluid and the cooling air. The door is movably mounted in the turbofan engine and is movable between a closed position, in which the cooling air will not flow through the heat exchanger, and an open position, in which the cooling air may flow through the heat exchanger.
|HEAT EXCHANGER - A gas turbine engine has a fan, a compressor section, a combustor, and a turbine section. The fan delivers a portion of air into the compressor, and into a duct, as bypass air. A bleed air system bleeds a quantity of air from the compressor into a chamber at least at low power conditions of the engine. The bleed air system has an opening which may be selectively closed to block bleed air, or opened to allow bleed flow from the compressor into the chamber. A heat exchanger is positioned such that a first surface of the heat exchanger is contacted by bypass air in the duct, and a second surface of the heat exchanger is contacted by bleed air when the bleed air system directs air from the compressor into the chamber.
|CUSTOMER BLEED AIR PRESSURE LOSS REDUCTION - A bleed air supply system for a gas turbine engine comprising a duct having an inlet end and extending to an outlet end. The inlet end of the duct is provided with a central insert. In another feature, there may be a plurality of ducts, and inlet ends of the plurality of ducts being spaced by at least 90°. In another feature, a compressor may have a diffuser with a shroud ending upstream of the downstream end of an inner shroud, having an outer shroud ending at a location upstream of a downstream end of an inner shroud at locations circumferentially aligned with an inlet end of the duct.
|Outer Fuel Nozzle Inlet Flow Conditioner Interface to End Cap - An interface between an inlet flow conditioner and a compressor discharge air passage in a gas turbine includes a cap back plate including a curved exterior surface, and an inlet flow conditioner (IFC) cooperable with the cap back plate. An end of the IFC includes a curved exterior surface that is continued from the curved exterior surface of the cap back plate. The interface provides for more uniform and higher pressure air as well as more air supply, resulting in more uniform combustion.
|MID-SECTION OF A CAN-ANNULAR GAS TURBINE ENGINE WITH A COOLING SYSTEM FOR THE TRANSITION - A cooling system is provided for a transition (
|CONSTANT SPEED PUMP SYSTEM FOR ENGINE ECS LOSS ELIMINATION - A gas turbine engine has an impeller pump for delivering air to an environmental control system and a speed control pump connected to the impeller pump for driving the impeller pump at a constant speed.
|PUMP SYSTEM FOR HPC EPS PARASITIC LOSS ELIMINATION - An engine includes a duct containing a flow of cool air and a pump system for providing air to an environmental control system. The pump system has an impeller having an inlet for receiving cool air from the duct and an outlet for discharging air to the environmental control system.
|AIRCRAFT GAS TURBINE HAVING A BOOSTER BLEED DUCT IN A STATOR VANE ROOT ELEMENT OF A BYPASS DUCT - The present invention relates to an aircraft gas turbine having a core engine and a bypass duct surrounding the latter, where the core engine includes in its inflow zone a booster, in the area of which at least on booster bleed duct is provided for supplying an airflow from the booster into the bypass duct, and a stator vane row arranged n the bypass duct downstream of a fan, characterized in that a vane root of the stator vane is designed as a structural element mounting the core engine, and that the booster bleed duct is provided inside the vane root and issues downstream of the stator vane into the bypass duct.
|AIR COOLER SYSTEM FOR GAS TURBINE ENGINES - A buffer air cooler system for gas turbine engines disposed in a bypass duct of the engine, includes a housing for containing the buffer air cooler therein and an inlet portion attached to the housing. In one embodiment, the inlet portion has a double-skin configuration in at least one region of a top, bottom and sides of the inlet portion.
|FUEL NOZZLE CAP - Certain embodiments include a first individual sector configured to fit together with a plurality of individual sectors to form a combustor cap assembly of a turbine combustor, wherein the first individual sector is configured to fixedly attach to a first fuel nozzle of a plurality of fuel nozzles, the first individual sector comprises a first substantially enclosed cavity configured to surround the first fuel nozzle, and the first substantially enclosed cavity is configured to receive a cooling air flow.
|Machined Aerodynamic Intercompressor Bleed Ports - An air bleed system for a jet engine is provided. The air bleed system is located in the low pressure compressor chamber and comprises a movable bleed valve and a stationary, annular bleed case having a forward section and an aft section and ligaments connecting the two sections and providing structural support to the bleed case. The ligaments are machined from the bleed case into an aerodynamic shape that maximizes the flow of air through the bleed ports.
|AERODYNAMIC INTERCOMPRESSOR BLEED PORTS - An air bleed system for a gas turbine engine includes an annular bleed case with a manifold therein. The annular bleed case has a forward section and an aft section and ligaments connecting the two sections. The forward section, the aft section and the ligaments define bleed ports. The manifold is disposed radially outward of the ligaments and communicates with the series of bleed ports.
|BAFFLE ASSEMBLY FOR BLEED AIR SYSTEM OF GAS TURBINE ENGINE - A baffle assembly for a housing of a combustor system of a gas turbine engine is provided. The baffle assembly includes a baffle retainer and a baffle. The baffle retainer includes a retainer tube and a collar. The retainer tube is received within a bleed port of the housing. The collar is defined at an upper end of the retainer tube and is configured to rest against the housing. The baffle includes a baffle tube rigidly attached to an inner surface of the retainer tube.
|FLOW DISCHARGE DEVICE - A bleed flow discharge device (
|GAS TURBINE - A gas turbine provided with an air bleeder tube (
|GAS TURBINE ENGINE BUFFER SYSTEM - A gas turbine engine includes a compressor section, a combustor in fluid communication with the compressor section, a turbine section in fluid communication with the combustor, and a buffer system. The buffer system can include a first circuit that supplies a first buffer supply air and a second circuit that supplies a second buffer supply air. The first circuit can include a first bleed air supply a second bleed air supply and at least one of an ejector and a valve. The second circuit can include a third bleed air supply, a fourth bleed air supply and at least one of an ejector and a valve.
|TURBOMACHINE WITH BLEED VALVES LOCATED AT THE INTERMEDIATE CASE - An assembly including an intermediate case of a bypass turbojet engine and of an inter-jet case extending upstream of the intermediate case to separate a primary air jet of the turbojet engine from its bypass air jet, the inter-jet case including, passing through it, a closable duct for diverting part of the primary flow to the bypass flow thereby forming a blow-off valve for the LP compressor, the intermediate case including arms passing across the bypass flow and the inter-jet case in its internal cavity including a first chamber situated upstream of the arms and a second chamber situated level with the arms, the duct being open or closed off by an annular component capable of axial movement set in motion by an arm that can rotate about a fixed pivot under action of a control cylinder, the cylinder being positioned in the second chamber.
|BLEED TUBE ATTACHMENT - An anti-vortex assembly for a gas turbine engine includes a support ring including a seat surrounding an opening and a bleed tube including a base with a bearing face sealing against the seat. A retaining ring is mounted within the support ring and engages the base of the bleed air tube to hold the bleed tube within the support ring.
|BLEED AIR SLOT - A disclosed gas turbine engine includes a case for a case for a compressor section including a bleed air slot. The bleed air slot includes an inlet having a first area radially inward of an outlet having a second area with the second area being greater than the first area. The bleed air slot further includes a center portion disposed along a radial line extending from an axis of the engine and an elongated portion extending from the opening at an angle relative to a line normal to the radial line.
|AIR COOLING SHAFT AT BEARING INTERFACE - A gas turbine engine including a compressor rotor and a turbine rotor connected by a compressor shaft portion connected to the compressor rotor and a turbine shaft portion connected to the turbine rotor. The compressor shaft portion and the turbine shaft portion are connected axially together by a shaft coupling, between the compressor rotor and the turbine rotor, and at least a bearing rotatably coupled to the compressor shaft portion adjacent the shaft coupling. The compressor shaft and/or the turbine shaft are provided with openings permitting cooling air to enter air passages in the area of the shaft coupling and surrounding the end of the turbine shaft portion, in order to dissipate heat originating at the turbine rotor and thus reducing the thermal stresses at the bearing.
|SYSTEM FOR CONTROLLING A COOLING FLOW FROM A COMPRESSOR SECTION OF A GAS TURBINE - A system for controlling a flow rate of a compressed cooling medium between a compressor section and a turbine section of a gas turbine includes a flow path that is defined between the compressor section and the turbine section of the gas turbine and a thermally actuated variable flow valve disposed within the flow path. The variable flow valve defines an opening that changes in size based on a temperature of the compressed cooling medium flowing therethrough.
|SYSTEM AND METHOD FOR EXPANDING A COMPRESSOR BLEED FLOW - A system includes a bleed system configured to direct a bleed flow from a high pressure region to a low pressure region. The bleed system includes a valve configured to control the bleed flow through the bleed system and a staged bleed conduit configured to incrementally depressurize the bleed flow. The staged bleed conduit includes an inlet coupled to the valve, a first stage configured to depressurize the bleed flow that is coupled to the inlet, a second stage configured to depressurize the bleed flow that is coupled to the first stage, and an outlet coupled to the second stage. The outlet is configured to direct the bleed flow to the low pressure region. The inlet, the first stage, the second stage, and the outlet are disposed along parallel axes.
|LOW PRESSURE COMPRESSOR BLEED EXIT FOR AN AIRCRAFT PRESSURIZATION SYSTEM - An aircraft pressurization system, includes an auxiliary compressor for further compressing compressed air received from a low pressure compressor section of a gas turbine engine while the compressed air is below a predetermined pressure level; a bleed passage for fluidically connecting the auxiliary compressor to the low pressure compressor section; and an environmental control system coupled to an output of the auxiliary compressor for conditioning the compressed air to a predetermined level.
|GAS TURBINE ENGINE BUFFER SYSTEM - A gas turbine engine includes a compressor section, a combustor in fluid communication with the compressor section, a turbine section in fluid communication with the combustor, a fan section configured to be driven by the turbine section via a geared architecture, and a buffer system that communicates buffer air to a portion of the gas turbine engine. The buffer system includes a first circuit configured to selectively mix a first bleed air supply having a first pressure and a second bleed air supply having a second pressure that is greater than the first pressure to provide a first buffer supply air having an intermediate pressure compared to the first pressure and the second pressure.
|GAS TURBINE ENGINE BUFFER SYSTEM - A gas turbine engine includes a fan, a compressor section, and a turbine section configured to drive the compressor section and the fan. A buffer system is configured to communicate a buffer supply air to a portion of the gas turbine engine. The buffer system includes a first bleed air supply having a first pressure, a second bleed air supply having a second pressure that is greater than the first pressure, and an ejector that selectively augments the first bleed air supply to prepare the buffer supply air for communication to the portion of the gas turbine engine. A method and a buffer system are also disclosed.
|Nacelle Anti-Ice Valve Utilized as Compressor Stability Bleed Valve During Starting - A gas turbine engine has a compressor section received within an inner housing. An is an outer housing is spaced radially outwardly of the inner core housing. A nacelle has an anti-icing system which taps compressed air from the compressor section through an anti-ice valve and to the nacelle. The anti-ice valve is opened at startup of the gas turbine engine to assist compressor stability.
|CONTAMINATION FREE REVERSE FLOW FITTING - A precooler for an aircraft engine system includes a precooler core and a precooler inlet to direct a compressor bleed flow into the precooler core to cool the compressor bleed flow. The precooler further includes a precooler outlet to direct the compressor bleed flow from the precooler to a selected component of the aircraft engine system and a precooler bleed port through which a portion of the compressor bleed flow is diverted to a secondary component of the aircraft engine system. The precooler bleed port is oriented such that flow entering the precooler bleed port must substantially reverse direction from a direction of the compressor bleed flow through the precooler.
|Aircraft having an Engine, a Fuel Tank, and a Fuel Cell - An aircraft includes a propulsion unit with a compressor and a bleed air device for providing bleed air from the propulsion unit, a fuel tank for a fuel, at least one fuel cell, a reactor for reforming fuel from the fuel tank to a hydrogen-containing fuel gas, and at least one feed unit with a bleed air inlet, a fuel inlet, an oxidizing agent outlet, and a fuel outlet. The feed unit can selectively and not simultaneously feed an oxidizing agent by means of the oxidizing agent outlet or fuel by means of the fuel outlet into the reactor. The feed unit is connected to the bleed air device and designed to provide the oxidizing agent based on bleed air.
|GAS TURBINE ENGINE WITH BLEED DUCT FOR MINIMUM REDUCTION OF BLEED FLOW AND MINIMUM REJECTION OF HAIL DURING HAIL INGESTION EVENTS - A gas turbine engine includes a bleed structure which includes a forward wall and a rear structural wall to define a deposit space downstream of the bleed structure for a hail event of a predetermined duration.
|SPLIT RING VALVE - A valve includes a body with bleed ports and a ring. The ring surrounds the body and includes two adjacent ring segments. The ring is movable between an open position and a closed position, the latter of which prevents flow through the ports.
|BLEED DUCT ASSEMBLY FOR A GAS TURBINE ENGINE - A bypass duct has a support unit comprising a pair of aerofoils arranged as an “A” frame. A bleed duct assembly is provided on the radially inner wall of the bypass duct annulus and the aerofoils project from the surface and extend across the annulus between the inner wall and an outer wall of the annulus. The aerofoils lean at an acute angle to the surface with the first flank facing toward the inner wall and adjoining a bleed duct opening. The bleed duct having a bleed duct passage and a submerged scoop.
|AIRCRAFT ENVIRONMENTAL CONTROL SYSTEM SELECTIVELY POWERED BY THREE BLEED PORTS - A bleed air supply system configured to supply bleed air from an engine to an environmental control system (ECS) pack is provided including a high pressure port configured to bleed air from a high spool of the engine. An intermediate pressure port is configured to bleed air from the high spool of the engine. The bleed air at the intermediate pressure port has a pressure less than the bleed air at the high pressure port. A precooler is fluidly coupled to the ECS pack, the high pressure port and the intermediate pressure port. Bleed air from the high pressure port and the intermediate pressure port is conditioned in the precooler before being provided to the ECS pack. A low pressure port is configured to bleed air from a low spool of the engine. The bleed air from the low pressure port bypasses the precooler and is supplied directly to the ECS pack.
|VORTEX FLUID FLOW DEVICE - The invention concerns a vortex fluid flow device comprising a vortex chamber, multiple forward flow inlets to the chamber and a reverse flow inlet to the chamber. The vortex fluid flow device is arranged such that in use there is a greater pressure drop across the device when there is reverse fluid flow than when there is forward fluid flow.
|TURBINE ENGINE INCORPORATING THERMOELECTRIC GENERATORS - A front-fan turbojet engine including at least one fluid circuit and an air/fluid heat exchanger by which the fluid is cooled by air external to the turbojet engine and a splitter for splitting a flow downstream of the fan between a primary flow and a secondary flow. The heat exchanger is associated with a thermoelectric generator including a first and a second thermal exchange surface, of which the first surface is in thermal contact with the airflow and the second surface is in thermal contact with the fluid to be cooled in the exchanger.
|COMBUSTOR COOLING STRUCTURE - The invention refers to a transition piece of a combustor of a gas turbine including an impingement cooling zone, a sequential disposed liner having at least one cooling arrangement and a closing plate with respect to the sequential disposed liner. The sequential disposed liner has a cooling channel structure. The cooling channel structure forms a closed loop cooling scheme or a quasi-closed loop cooling scheme. The cooling channel structure is operatively connected to a cooling medium to cool at least one part of the sequential disposed liner.
|METHOD AND SYSTEM FOR CONTROLLING COMPRESSOR FORWARD LEAKAGE - A system for controlling compressor forward leakage to the compressor flow path in a gas turbine is provided. The system includes a first compressor rotor wheel coupled to an upstream face of a compressor spacer wheel. A second compressor rotor wheel is coupled to a downstream face of the compressor spacer wheel. The compressor spacer wheel includes at least one intake recess defined in the downstream face, at least one discharge recess defined in the upstream face, and at least one axial passage coupling the at least one intake recess in fluid communication with the at least one discharge recess.
|GAS TURBINE ENGINE BUFFER SYSTEM - A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a compressor section, a combustor in fluid communication with the compressor section and a turbine section in fluid communication with the combustor. A buffer system includes a first circuit that supplies a first buffer supply air and a second circuit that supplies a second buffer supply air. The first circuit includes a first bleed air supply and a second bleed air supply. The second circuit includes a third bleed air supply and a fourth bleed air supply and at least one of the first circuit and the second circuit includes a variable area ejector.
|FLUID SYSTEM - A gas turbine engine is disclosed including a bifurcation fairing located in a bypass duct of the gas turbine engine and traversing the radial extent of the bypass duct. The bifurcation fairing has a delivery conduit inlet leading to a delivery conduit extending inside the bifurcation fairing, the delivery conduit being arranged in use for delivery of bypass air to one or more components of the gas turbine engine. A diverter conduit has a diverter conduit inlet from the delivery conduit upstream of the delivery conduit reaching the one or more components of the gas turbine engine. The diverter conduit has an outlet to a location other than the one or more components of the gas turbine engine.
|OFF-LINE WASH SYSTEMS AND METHODS FOR A GAS TURBINE ENGINE - The present application and the resultant patent provide a wash system for a gas turbine engine. The wash system may include a water source containing a volume of water therein, and a surface filming agent source containing a volume of a surface filming agent therein. The wash system also may include a mixing chamber in fluid communication with the water source and the surface filming agent source, wherein the mixing chamber is configured to mix the water and the surface filming agent therein to produce a film-forming mixture. The wash system further may include an aerosolizing device in fluid communication with the mixing chamber and configured to form an aerosol spray of the film-forming mixture and a propellant. The wash system still further may include a number of supply lines in fluid communication with the aerosolizing device and configured to direct the aerosol spray into the gas turbine engine.
|HIGH PRESSURE TURBINE COOLING - A gas turbine engine includes a compressor section having a plurality of compressor stages, a combustor fluidly connected to the compressor section, a turbine section fluidly connected to the combustor section, the turbine section having at least one stage, a compressor bleed structure disposed in one of the plurality of compressor stages and operable to remove air from the compressor stage, a heat exchanger having an input connected to the compressor bleed, and an output connected to an active cooling system of at least one turbine stage, and wherein the compressor stage in which the compressor bleed structure is disposed includes airflow at a pressure above a minimum pressure threshold, and wherein the airflow has a temperature above a maximum temperature threshold.
|GAS TURBINE ENGINE ANTI-SURGE VALVE DISCHARGE TUBE - An anti-surge mechanism has a duct for tapping compressed air and delivering that air into an exhaust flow. A selectively open valve allows the tapped air to flow into the exhaust flow through a plurality of holes. The holes are sized to tune a frequency of a sound created by the tapped air to a frequency range outside of normal human hearing. A gas turbine engine is also disclosed.
|BOLTED DUCT JOINTS - A first duct segment may include a first flange including a first planar mating face and a first recess. A second duct segment may include a second flange including a second planar mating face and a second recess. The first duct segment and the second duct segment may be installed in a gas turbine engine. A first E-seal may be inserted in the first recess, and a second E-seal may be inserted in the second recess. A gap may be measured between the first flange and the second flange. A flat shim may be selected based on the size of the gap. The flat shim may be inserted between the first flange and the second flange. A plurality of bolts may be inserted through the first flange, the flat shim, and the second flange to seal the first duct segment to the second-duct segment.
|COMBUSTOR CAP ASSEMBLY - A combustor cap assembly includes an impingement plate coupled to an annular shroud and a cap plate which is coupled to the impingement plate to form an impingement air plenum therebetween. The combustor cap assembly further includes a flow conditioning plate coupled to a forward end portion of the shroud. The flow conditioning plate includes an inner band portion, an outer band portion and an annular portion. The annular portion defines a plurality of flow conditioning passages. The inner band portion at least partially defines a cooling air plenum within the combustor cap assembly. The inner band portion defines an exhaust channel which is in fluid communication with the impingement air plenum and an exhaust outlet. The flow conditioning plate further defines a cooling air passage which provides for cooling air flow into the cooling air plenum.
|DEVICE FOR THE EXTRACTION OF BLEED AIR AND AIRCRAFT ENGINE WITH AT LEAST ONE DEVICE FOR THE EXTRACTION OF BLEED AIR - A device for the extraction of bleed air from flowing air at or in an aircraft engine is provided. The device includes an adjusting means for the specific adjustment of an inlet cross section of an opening for the bleed air in or at a wall during operation of the aircraft engine. The opening for the bleed air is arranged in or at a deformable base and the adjusting means acts on the deformable base for the purpose of modifying the inlet cross section of the opening relative to the flowing air, wherein the deformable base is part of a metallic housing in the aircraft engine.
|DEVICE FOR THE EXTRACTION OF BLEED AIR AND AIRCRAFT ENGINE WITH AT LEAST ONE DEVICE FOR THE EXTRACTION OF BLEED AIR - A device for the extraction of bleed air from flowing air at or in an aircraft engine is provided. The device includes a means for specific adjustment of an inlet cross section of an opening at or in the area of a wall of the aircraft engine and a flow guide means for a boundary layer flow.
|MOUNTING ARRANGEMENT FOR AN ELECTRICAL HARNESS - A mounting arrangement for an electrical harness comprises an elongate tray, the tray having a base and raised sides to define a channel. The electrical harness is mounted within the channel and cooling air is directed into the channel to cool the electrical harness.
|INLET BLEED HEAT MANIFOLD INCLUDING ACOUSTICALLY TREATED FEED PIPE - An inlet bleed heat (IBH) system manifold for a compressor inlet housing is provided. The manifold includes: a plurality of feed pipes for delivering a compressor discharge air, each feed pipe extending across a duct of the compressor inlet housing. Each feed pipe includes: an elongated inner feed pipe for delivering the compressor discharge air, the inner feed pipe including a plurality of orifices along at least a portion of a length of the inner feed pipe, each orifice extending through a wall of the inner feed pipe allowing the compressor discharge air to exit the inner feed pipe; and a noise attenuating material disposed about the inner feed pipe and the plurality of orifices, the noise attenuating material configured to attenuate noise created by the compressor discharge air exiting the plurality of orifices.
|HEAT EXCHANGER - Heat exchanger includes tubular element within outer casing and inner sleeve within tubular element. Outer casing and outer surface of tubular element define a first annular passage through which first fluid flow path extends. Inner sleeve and inner surface of tubular element define second annular passage through which second fluid flow path extends. First annular passage is sealed against outer surface of tubular element and second annular passage is sealed within inner surface of tubular element. Second heat exchanger having tubular element within outer casing. Tubular element has outer fins and inner fins. Outer casing and outer surface of tubular element define a first annular passage through which first fluid flow path extends. Inner surface of tubular element defines a second passage through which second fluid flow path extends. Outer fins are integral with outer surface of tubular element and inner fins are integral with inner surface of tubular element.
|Hybrid Power Generation System
|HIGH PRESSURE COMPRESSOR DIFFUSER FOR AN INDUSTRIAL GAS TURBINE ENGINE