Abstract: A hot fairing structure for a pre-diffuser includes a ring-strut-ring structure that comprises a multiple of hollow struts; and a multiple of diffusion passage ducts attached to the ring-strut-ring structure. A pre-diffuser for a gas turbine engine includes an exit guide vane ring having a multiple of exit guide vanes defined around an engine longitudinal axis; a ring-strut-ring structure adjacent to the exit guide vane ring to form a multiple of diffusion passages defined around the engine longitudinal axis, an inlet to each of the multiple of diffusion passages smaller than an exit from each of the multiple diffusion passage through the ring-strut-ring structure; a diffusion passage duct attached to the ring-strut-ring structure at the exit from each of the multiple diffusion passage.

Abstract: A heat exchanger has arcuate inlet and outlet manifolds and a plurality of tube banks, each tube bank coupling one of the inlet manifold outlets to an associated one of the outlet manifold inlets. Each tube bank partially nests with one or more others of the tube banks and has: a first header coupled to the associated inlet manifold outlet and the associated the outlet manifold inlet; a second header; and a plurality of tube bundles each having a first end coupled to the associated first header and a second end coupled to the associated second header. A flowpath from the each inlet manifold outlet passes sequentially through flowpath legs formed by each of the tube bundles in the associated tube bank to exit the tube bank to the associated outlet manifold inlet.

Abstract: A hot fairing structure for a pre-diffuser includes a ring-strut-ring structure that comprises a multiple of hollow struts and a multiple of inlets to a respective diffusion passage, one of the multiple of inlets formed between each one of the multiple of hollow struts located between two diffusion passages.

Abstract: A dispatchable storage combined cycle power plant comprises a topping cycle that combusts fuel to generate electricity and produce hot exhaust gases, a steam power system, a heat source other than the topping cycle, and a thermal energy storage system. Heat from the heat source, from the thermal energy storage system, or from the heat source and the thermal energy storage system is used to generate steam in the steam power system. Heat from the topping cycle may be used in series with or in parallel with the thermal energy storage system and/or the heat source to generate the steam, and additionally to super heat the steam.

Abstract: A gas turbine engine including: a tail cone; a low pressure compressor; a low pressure turbine; a low speed spool interconnecting the low pressure compressor and the low pressure turbine; and an electric generator located within the tail cone, the electric generator being operably connected to the low speed spool, wherein the electric generator includes a coolant cavity in thermal communication with one or more components of the electric generator; a structural support housing at least partially enclosing the electric generator, the structural support housing including a forward wall located on a forward end of the structural support housing, wherein the forward wall includes a first opening; a first coolant conveying tube extending through the first opening to fluidly connect to the coolant cavity; and a first electrical connector tube extending through the first opening within the first coolant conveying tube to electrically connect to the electric generator.

Abstract: An air bleed device for an aircraft engine, including a frame and a flap able to rotate about an axis in relation to the frame, the device further including a return system configured to bias the flap in a determined position about the axis and including a torsion spring, a first end of which is connected to the flap and a second end of which is connected to the frame. The second end is connected to the frame by adjusting the preload of the spring by screwing when the flap is in the determined position.

Abstract: A method is provided for maintaining a gas turbine engine installed on an aircraft, the gas turbine engine including an electric machine mounted at least partially inward of a core air flowpath of the gas turbine engine. The method includes removing a rotor mount connecting a rotor of the electric machine to a rotary component of the gas turbine engine; removing a stator mount connecting a stator of the electric machine to a stationary component of the gas turbine engine; and removing in situ the electric machine.

Abstract: An electrical power unit cooling system including at least one cooling fin configured to be thermally connected to a bus housing of the electrical power unit, the bus housing being configured to be attached to a surface of an aircraft fan housing such that the at least one cooling fin protrudes through the fan housing in order to be cooled by air flow within the fan housing.

Abstract: An engine assembly for an aircraft, including an internal combustion engine having a liquid coolant system in fluid communication with a heat exchanger, an exhaust duct in fluid communication with air passages of the heat exchanger, a fan in fluid communication with the exhaust duct for driving a cooling air flow through the air passages of the heat exchanger and into the exhaust duct, and an intermediate duct in fluid communication with an exhaust of the engine and having an outlet positioned within the exhaust duct downstream of the fan and upstream of the outlet of the exhaust duct. The outlet of the intermediate duct is spaced inwardly from a peripheral wall of the exhaust duct. The engine assembly may be configured as an auxiliary power unit. A method of discharging air and exhaust gases in an auxiliary power unit having an internal combustion engine is also discussed.

Abstract: A system is provided. The system includes a first environmental control sub-system, operating in a first mode, that receives a first medium at a first flow amount and a first pressure. The system also includes a second environmental control sub-system, operating in a second mode, that receives a second medium at a second flow amount and a second pressure. The first flow amount is greater than the second flow amount, and the second pressure is greater than the first pressure.

Abstract: A method and apparatus using a cooler to selectively cool P3 air directed into an impeller rear cavity of a gas turbine engine during engine high power levels, may include connection of the cooler to a low pressure compressor bleed-off valve apparatus that is modulated to be closed during engine high power levels to create a pressure differential over the bleed-off valve apparatus, to drive a stream of low pressure compressor air as a cooling work fluid under such pressure differential to selectively flow through the cooler.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a tie-rod and a threaded retainer. The tie-rod includes a tie-rod threaded portion and a tie-rod unthreaded portion. The threaded retainer includes a retainer threaded portion and a retainer unthreaded portion. The retainer threaded portion is mated with the tie-rod threaded portion. The retainer unthreaded portion is radially engaged with the tie-rod unthreaded portion.

Abstract: A thermal management system for a gas turbine engine and/or an aircraft is provided including a thermal transport bus having a heat exchange fluid flowing therethrough. The thermal management system also includes a plurality of heat source exchangers and at least one heat sink exchanger. The plurality of heat source exchangers and the at least one heat sink exchanger are in thermal communication with the heat exchange fluid in the thermal transport bus. The plurality of heat source exchangers are arranged along the thermal transport bus and configured to transfer heat from one or more accessory systems to the heat exchange fluid, and the at least one heat sink exchanger is located downstream of the plurality of heat source exchangers and configured to remove heat from the heat exchange fluid.

Abstract: A ducted fan gas turbine engine has a propulsive fan, a fan case surrounding the fan, a core engine, and a plurality of support structures which transmit loads from the core engine to the fan case. Each support structure has two support struts which extend from the core engine to a joint radially outwardly of the fan case. The support struts are spaced apart at the core engine but converge to meet at the joint. Each support structure further has two structural elements which extend from the joint to respective fixing positions on the fan case at opposite sides of the joint.

Abstract: A turbine engine assembly includes a core compressor configured to discharge a first airflow at a first temperature and a first pressure. The turbine engine assembly also includes a cooling system turbine configured to receive the first airflow at the first temperature and the first pressure and discharge a second airflow at a second pressure less than the first pressure. The turbine engine assembly further includes a heat exchanger configured to receive the second airflow and discharge a third airflow at a second temperature less than the first temperature. The turbine engine assembly also includes a cooling system compressor rotatably coupled to the cooling system turbine. The cooling system compressor is configured to receive the third airflow and discharge a fourth airflow at a third pressure greater than the first pressure.

Abstract: A gas turbine engine includes a compressor section, a combustor fluidly connected to the compressor section via a primary flowpath and a turbine section fluidly connected to the combustor via the primary flowpath. Also included is a cascading cooling system having a first inlet connected to a first compressor bleed, a second inlet connected to a second compressor bleed downstream of the first compressor bleed, and a third inlet connected to a third compressor bleed downstream of the second compressor bleed. The cascading cooling system includes at least one heat exchanger configured to incrementally generate cooling air for at least one of an aft compressor stage and a foremost turbine stage relative to fluid flow through the turbine section.

Abstract: A turbine injection system for a gas turbine engine includes a first end operable to receive air from a heat exchanger, a second end operable to distribute mixed cooling air to a turbine stage, an opening downstream of said first end and a mixing plenum downstream of said first end and said opening. The opening provides a direct fluid pathway into said turbine injection system.

Abstract: The present application provides a power augmentation system for use with a gas turbine engine. The power augmentation system may include a compressor discharge manifold, an inlet bleed heat system in communication with the compressor discharge manifold via an inlet bleed heat line, and an auxiliary compressor in communication with the compressor discharge manifold via the inlet bleed heat line.

Abstract: A turbine engine cooling arrangement includes a core passage for receiving a core flow for combustion, a first airflow source including a first passage adjacent the core passage for conveying a first airflow, and a second airflow source including a second passage adjacent the first passage for conveying a second airflow. A heat exchanger is thermally connected with the first passage and the second passage for transferring heat between the first airflow and the second airflow.

Abstract: A manufacturing method includes providing a substrate having one or more grooves formed therein. One or more coatings having one or more grooves formed therein are disposed on the substrate and in fluid communication with the one or more grooves in the substrate. A cover coating is disposed on a portion of an outermost surface of the one or more coatings, having one or more cooling outlets formed therein and in fluid communication with the one or more grooves in the one or more coatings. The substrate, the one or more coatings and the cover coating define therein a cooling network for cooling a component. A component having a cooling network defined therein a substrate, one or more coatings disposed on at least a portion of the substrate, and a cover coating disposed over at least a portion of an outermost coating of the one or more coatings.

Abstract: A gas turbine engine includes a heat exchanger, a bearing compartment, and a nozzle assembly in fluid communication with the bearing compartment. The heat exchanger exchanges heat with a bleed airflow to provide a conditioned airflow. The bearing compartment is in fluid communication with the heat exchanger. A first passageway communicates the conditioned airflow from the heat exchanger to the bearing compartment. A second passageway communicates the conditioned airflow from the bearing compartment to the nozzle assembly.

Abstract: Effusion cooling holes formed through a transition component provided in a combustion section of a gas turbine engine. The transition component directs a hot working gas from a combustion basket to a first row of vanes in a turbine section of the engine. The effusion cooling holes are formed through an outer wall of the transition component in a direction so that the flow of air through the effusion holes is in a direction substantially opposite to the bulk flow direction of the working gas through the transition component.

Abstract: There is disclosed a gas cooler 20 for providing high-pressure sealing gas to a bearing chamber. The cooler comprises a turbine 22; a turbine inlet 24 arranged to receive gas to drive the turbine and a turbine outlet 26 arranged to deliver gas output from the turbine; a compressor 28 arranged to be driven by the turbine 22; a compressor inlet 30 arranged to receive gas to be compressed by the compressor and a compressor outlet 32 arranged to deliver gas output from the compressor 28; and a cooler outlet 36 in fluid communication with the turbine outlet 26 and the compressor outlet 32 so as to deliver high-pressure sealing gas comprising gas merged from the turbine outlet 26 and the compressor outlet 32.

Abstract: A turbine engine includes a shaft, a fan, at least one bearing mounted on the shaft and rotationally supporting the fan, a fan drive gear system coupled to drive the fan, a bearing compartment around the at least one bearing and a source of pressurized air in communication with a region outside of the bearing compartment.

Abstract: The present application provides a passive control valve actuator cooling system to provide a flow of cooling air to a control valve actuator used with a gas turbine engine. The passive control valve actuator cooling system may include a turbine enclosure with a negative pressure therein, a radiation shield with a number of radiation shield outlets and the control valve actuator positioned therein, and a cooling air line extending from outside of the turbine enclosure to the radiation shield such that the negative pressure within the turbine enclosure pulls the flow of cooling air into and through the radiation shield so as to cool the control valve actuator.

Abstract: An aft frame for a transition duct of a gas turbine combustor includes a main body having an outer rail, an inner rail, a first side rail circumferentially separated from an opposing second side rail, a forward portion, an aft portion and an outer surface. An inlet port extends through the outer surface and an exhaust port extends through the forward portion. A serpentine cooling passage is defined within the main body beneath the outer surface. The serpentine cooling passage is in fluid communication with the inlet port and the exhaust port. A conduit may be connected to the exhaust port for routing a compressed working fluid away from aft frame.

Abstract: A diffuser assembly for a ventilation system is disclosed herein. The diffuser assembly may include a boundary defining an interior of the diffuser assembly. The diffuser assembly also may include at least one inlet to the interior of the diffuser assembly. The at least one inlet to the interior of the diffuser assembly may be in fluid communication with at least one ventilation fan. Moreover, the diffuser assembly may include an outlet from the interior of the diffuser assembly. The outlet from the interior of the diffuser assembly may be in fluid communication with an engine enclosure. In this manner, the outlet from the interior of the diffuser assembly may be configured to provide a flow of ventilation air from the at least one ventilation fan as a substantially uniform flow of ventilation air within the engine enclosure.

Abstract: One embodiment of the present invention is a unique gas turbine engine. Another embodiment of the present invention is a unique gas turbine engine bearing system. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and gas turbine engine bearing systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: Thermal isolating service tubes and assemblies thereof for gas turbine engines are provided. The thermal isolating service tube comprises an inner tubular member defining a fluid passage and at least one outer tubular member disposed about the inner tubular member. A spacing volume is defined between at least the inner tubular member and an adjacent outer tubular member. The thermal isolating service tube comprises a unitary structure and has at least one portion with a curved configuration, a non-circular cross-sectional shape, or both.

Abstract: A cooling system includes a housing having an aperture intersecting a passage that includes first and second passage portions. A plug is arranged in the aperture in an interference fit in a first position at a first temperature condition to block the passage and fluidly separate the first and second passage portions. A biasing element is arranged in the housing and is configured to move the plug from the first position to a second position at a second temperature condition to fluidly connect the first and second passage portions.

Abstract: Variable speed drive controlled discharge pumps are used for pumping chilled water from a chilled water storage tank to coils operatively associated with the air inlet of a gas turbine, wherein the gas turbine is used to generate electricity. Use of the variable speed drive controlled discharge pumps aids is protecting the temperature distribution in the chilled water storage tank so that a thermocline rises with the addition of chilled water during charging of the tank with chilled water, and lowers during discharge of the tank when using the chilled water to cool inlet air at one or more gas turbines operated at a facility.

Abstract: A system includes a turbine casing assembly that includes an outer shell and an inner shell positioned substantially concentrically within the outer shell. The inner shell includes an inner surface facing away from the outer shell and an outer surface facing toward the outer shell, and the outer surface has one or more false flanges. At least one of the one or more false flanges includes a first surface protruding from the outer surface and facing the outer shell, and a flow diverting portion extending between the first surface and the outer surface of the inner shell. The flow diverting portion includes a first portion that diverges in a first circumferential direction between the first surface and the outer surface.

Abstract: A cooling arrangement (56) having: a duct (30) configured to receive hot gases (16) from a combustor; and a flow sleeve (50) surrounding the duct and defining a cooling plenum (52) there between, wherein the flow sleeve is configured to form impingement cooling jets (70) emanating from dimples (82) in the flow sleeve effective to predominately cool the duct in an impingement cooling zone (60), and wherein the flow sleeve defines a convection cooling zone (64) effective to cool the duct solely via a cross-flow (76), the cross-flow comprising cooling fluid (72) exhausting from the impingement cooling zone. In the impingement cooling zone an undimpled portion (84) of the flow sleeve tapers away from the duct as the undimpled portion nears the convection cooling zone. The flow sleeve is configured to effect a greater velocity of the cross-flow in the convection cooling zone than in the impingement cooling zone.

Abstract: A cover (54, 54A-B) enclosing with clearance (65) an acoustic damping resonator (24) on a working gas path liner (22) of a gas turbine component (28). The cover includes a coolant inlet chamber (56, 56B) with a top wall (58, 58B) that is closer to the liner than a top wall (32) of the resonator, and extends upstream from the resonator relative to the working gas flow (48). Compressed air (26) surrounds the cover at a higher pressure than the working gas and flows (44) into and through the coolant inlet chamber, then through holes (34) in the resonator, then exits through holes (38) the liner into the working gas. The coolant inlet chamber directs the flow of compressed air over a weld (50) of the upstream wall (40) of the resonator to cool it. The cover may be formed as a box (57) or a sleeve (69).

Abstract: A cooling system for turbine starter lubricant includes one or more outflow transfer passages (54) extending from the turbine starter (10) to a secondary component (40). At least one heat exchange passage (56) affixed at a first end to an end of an outflow transfer passage (54) of the one or more outflow transfer passages (54), is located in the secondary component (40) having a lower interior temperature than the turbine starter (10). One or more return transfer passages (60) are affixed to a second end of the at least one heat exchange passage (56) and extend from the secondary component (40) to the turbine starter (10). Flowing a volume of starter lubricant (42) through the one or more outflow transfer passages (54), the at least one heat exchange passage (56), and the one or more return transfer passages (60) removes thermal energy from the volume of starter lubricant (42) and returns the volume of starter lubricant (42) to the turbine starter (10).

Abstract: The present application and the resultant patent thus provide a microchannel system for cooling a hot gas path surface of a turbine. The microchannel system may include a turbine component having an outer surface extending along a hot gas path of the turbine, a microchannel defined within the turbine component and extending about the outer surface, and a number of pockets defined within the turbine component and positioned along the microchannel. The present application and the resultant patent further provide a method of forming a microchannel system for cooling a hot gas path surface of a turbine. The method may include the steps of forming a turbine component having an outer surface extending along a hot gas path of the turbine, defining a microchannel within the turbine component and extending about the outer surface, and defining a number of pockets within the turbine component and positioned along the microchannel.

Abstract: 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.

Abstract: The invention refers to a method for producing a near-surface cooling passage in a thermally highly stressed component, which includes: a) providing a component which has a surface on a hot side in a region which is to be cooled; b) letting a channel into the surface; c) inserting a cooling tube into the channel; d) filling the channel, with the cooling tube inserted, with a temperature-resistant filling material in such a way that the inserted cooling tube is embedded into the filling material, leaving free an inlet and an outlet; and e) covering the channel, with the cooling tube embedded, with an anti-oxidation, temperature-stable cover layer. The method is inexpensive and can be used in a flexible manner in the most diverse situations in order to save cooling medium or to reduce the thermal load.

Abstract: A gas turbine engine comprises a compressor, a combustion chamber, an outer casing, an inner casing and a cooling arrangement. The outer casing surrounds the compressor and the combustion chamber and the combustion chamber has turbine nozzle guide vanes. The compressor has load carrying outlet guide vanes connected to the outer casing and the inner casing. The turbine nozzle guide vanes connect the outer casing and the inner casing. The cooling arrangement comprises a cooling air duct located between the compressor and the combustion chamber. The compressor outlet guide vanes carry at least one aerodynamic fairing. A support structure supports the cooling air duct from the inner casing at two spaced positions and the support structure forms a chamber with the inner casing. The support structure comprises at least one hollow duct and each hollow duct locates behind a respective one of the aerodynamic fairings.

Abstract: The build-up of carbon deposition on the front face of a combustor heat shield is discouraged by jetting air out from the front face of the heat shield with sufficient momentum to push approaching fuel droplets or rich fuel-air mixture way from the heat shield.

Abstract: 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. An air to air FLADE duct heat exchanger is disposed in a FLADE duct of the engine and a valving apparatus is operable for selectively switching the FLADE duct heat exchanger between the turbine cooling circuit and the air cycle system. A vapor cycle system includes a vapor cycle system condenser that may be in heat transfer cooling relationship with the air cycle system. An air cycle system heat exchanger and an engine burn fuel to air heat exchanger in the vapor cycle system condenser may be used for cooling a working fluid in a refrigeration loop of the vapor cycle system.

Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform having an outer surface and an inner surface that axially extend between a leading edge portion and a trailing edge portion. At least one augmentation feature is disposed on at least one of the leading edge portion and the trailing edge portion of the outer surface of the platform.

Abstract: A turbine exhaust cylinder for an industrial gas turbine engine with an external blower that delivers cooling air to a heat shield jacket surrounding an outer diameter cylinder to provide impingement cooling for the cylinder. Spent impingement cooling air is collected as passed through a space formed between a fairing and a strut to provide cooling to these parts. The cooling air then provides cooling for the inner diameter cylinder before being discharged into the turbine exhaust gas or out from the turbine exhaust cylinder.

Abstract: Systems, methods and apparatus for noise attenuation of a generator set are disclosed. The generator set includes an internal combustion engine enclosed within a compartment that substantially isolates the internal combustion from ambient air and a load connected to the internal combustion engine. A heat exchanger is disposed within or coupled to the compartment and is operable to cool air in the compartment without directly exchanging ambient air with compartment air.

Abstract: An aft frame of a turbine engine transition piece body is provided and includes an annular body disposed within a first annular space defined between an impingement sleeve and a compressor discharge casing and aft of a second annular space defined between the transition piece body and the impingement sleeve and including a main portion with a first surface facing the first annular space and a second surface facing the forward annular space. The main portion has an impingement hole extending therethrough from an inlet at the first surface of the annular body to an outlet at the second surface of the annular body to define a fluid path along which the first and second annular spaces communicate with one another.

Abstract: A process for management of thermal effluents of an aircraft that includes an airframe (110) and at least one propulsion system (112), whereby the at least one propulsion system (112) includes a gas turbine engine (116) that is supplied with fuel via a fuel supply circuit (122) that extends from a reservoir (124) that is arranged at the airframe (110), whereby the airframe (110) includes at least one source of thermal effluents (134), wherein the process includes at least partially dissipating—at the level of at least one propulsion system (112)—the thermal effluents that are generated at the airframe (110) by using as coolant the fuel that is used for supplying the gas turbine engine (116).

Abstract: Secondary air flow is provided for a ducted fan having a reverse flow turbine engine core driving a fan blisk. The fan blisk incorporates a set of thrust fan blades extending from an outer hub and a set of integral secondary flow blades extending intermediate an inner hub and the outer hub. A nacelle provides an outer flow duct for the thrust fan blades and a secondary flow duct carries flow from the integral secondary flow blades as cooling air for components of the reverse flow turbine engine.

Abstract: A nozzle includes a center body and a shroud circumferentially surrounding at least a portion of the center body to define an annular passage between the center body and the shroud. A plurality of apertures pass through the center body to the annular passage, and a plenum extends inside the center body and is in fluid communication with the plurality of apertures. A cooling medium is in fluid communication with the plenum. A method for cooling a nozzle includes flowing a cooling medium through a plenum across a surface of the nozzle.

Abstract: An example auxiliary power unit (APU) exhaust silencer includes cooling features to protect the outer skin and other components from heat generated by gases passing through an exhaust duct. Cooling air flow through a cooling air passage in thermal contact with the exhaust silencer carries heat away from other nearby components and the aircraft skin.

Abstract: A cooling medium supply apparatus has a sensing means and a cooling medium pump mechanism which is powered by a gas turbine engine assembly. The gas turbine engine assembly is self-contained with a fuel supply, an ignition system and a starting system. The sensing means monitors the presence of a primary supply of a cooling medium, and on detecting a loss of the primary supply, actuates the gas turbine assembly to provide a secondary supply of the cooling medium.