2102results about "Turbine/propulsion engine cooling" patented technology

A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.

A gas turbine engine is provided. The gas turbine engine includes: a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches; wherein the gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (FnTotal) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: FnTotal×EGT / (AHPCExit2×1000).

A gas turbine engine is provided, having: a bifurcation; and a turbomachine further including a cooled cooling air (CCA) system, the CCA system having: a cold side bleed assembly defining an inlet positioned to be in fluid communication with an airflow over the bifurcation; a CCA heat exchanger in thermal communication with the cold side bleed assembly downstream of the inlet of the cold side bleed assembly; and a hot side bleed assembly defining an inlet in fluid communication with a working gas flowpath through a compressor section, at a compressor discharge cavity, or both, the hot side bleed assembly in thermal communication with the CCA heat exchanger to cool an airflow through the hot side bleed assembly, the hot side bleed assembly further in thermal communication with a hot component of the turbomachine to cool the hot component of the turbomachine.

A thermal management system defines a thermal management system flowpath to provide a flow of a fluid to an electric machine and a power electronics assembly electrically connected to the electric machine. The thermal management system includes a first heat exchanger thermally connected to the thermal management system flowpath and to the electric machine, and a second heat exchanger thermally connected to the thermal management system flowpath downstream of the first heat exchanger. The second heat exchanger is thermally connected to the power electronics assembly.

Heat exchangers for cryogenic fuel distribution systems and related methods are disclosed herein. An example aircraft disclosed herein includes a cryogenic fuel system including a cryogenic fuel, a hydraulic system including a hydraulic fluid, a gas turbine engine fluidly coupled to the cryogenic fuel system, and a heat exchanger fluidly coupled to the cryogenic fuel system, the heat exchanger is used to exchange heat between the cryogenic fuel and at least one of (1) hydraulic fluid or (2) working fluid thermally coupled to the hydraulic system.

A gas turbine engine includes a turbomachine comprising compressor, combustion, and turbine sections. The gas turbine engine defines a maximum exhaust gas temperature, a maximum drive turbine shaft torque, and a corrected specific power. The gas turbine engine includes a controller configured to autonomously regulate performance of the gas turbine engine in response to at least one of: a thrust demand, an energy efficiency target, or a flight profile condition.

A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows:(TOUT / √{square root over (ADTExit)})2*EGT / AHPCExit*10−11;wherein CSP is greater than 0.0001194×EGT2−0.103×EGT+22.14 and less than 0.0003294×EGT2−0.306×EGT+77.91; and wherein EGT is greater than 525 degrees Celsius and less than 1250 degrees Celsius.

The present invention relates to a system, apparatus, and method for the dual-purpose functionality of capturing carbon dioxide (CO2) from atmospheric and flue gas sources and concurrently generating electrical energy. More specifically, this patent application focuses exclusively on the non-ionized hydrogen gas turbine variant of the apparatuses and methods for capturing carbon dioxide and electrical energy generating system invention. The invention embodies an integrated system designed for the efficient capture of CO2 and simultaneous production of clean electricity. To simultaneously generate electric power at maximum efficiency and capture carbon dioxide, the present invention comprises different integrated processes, integrated systems, and techniques. The present system comprises; a non-ionized hydrogen gas turbine system unit, a carbon dioxide capturing system unit, a hybrid solar Hydrogen-Oxygen gas generator system system unit and waste heat recovery system unit. Moreover, the system comprises waste-heat-based carbon dioxide absorption and waste-heat-based regeneration sub-units to power carbon dioxide capturing parts by waste heat. All the systems of units are developed, integrated, hybrid, and co-functionally working to simultaneously enhance both CO2 capture efficiency and electricity generation. Through these integrated processes, systems, and techniques, the invention aims to generate and maximize clean electrical power output while effectively capture CO2 and reducing CO2 emission impacts on the environment with a dual solution approach. The invention further includes various alternative embodiments, and invention versions for versatile application in CO2 capture and energy production fields.

An air / oil heat exchanger arranged in a vein. The exchanger includes internal parietal corridors and / or external parietal corridors and central corridors, the radial height of the internal and / or external parietal corridors increasing from the inlet of the exchanger to its outlet.

A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.

Methods and apparatus for anti-ice heat supply from waste heat recovery systems are disclosed. An apparatus for an aircraft, the apparatus comprising a fuel heat exchange system, an anti-ice heat exchanger, a waste heat recovery heat exchanger, and a conduit coupled to the fuel heat exchange system, the anti-ice heat exchanger, and the waste heat recovery heat exchanger, the conduit including a first portion and a second portion distinct from the first portion, wherein the first portion of the conduit carries a first portion of a thermal transfer fluid from the waste heat recovery heat exchanger to the anti-ice heat exchanger in which the thermal transfer fluid supplies anti-ice heat to a portion of the aircraft, wherein the second portion of the conduit carries a second portion of the thermal transfer fluid from the waste heat recovery heat exchanger to the fuel heat exchange system.

A micro gas turbine engine may include a single-piece rotor having a compressor, a turbine, and a shaft manufactured from an additive manufacturing process. The compressor, turbine, and shaft are a single uniform piece and not separately joined together. The micro gas turbine engine may further include a ceramic cover having built in fuel injectors and stator blades for the compressor. The micro gas turbine engine may further include a inner combustor lining having built in stator blades for the turbine.

A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.

A method of operating an aircraft gas turbine engine including: a core including a turbine, a compressor, a combustor, and a core shaft connecting the turbine to the compressor; a fan upstream of the core; a fan shaft; at least one bearing supporting the fan shaft; at least one auxiliary system using some of the fuel; an oil loop system to supply oil to the bearing; and a heat exchange system including: a primary fuel-oil heat exchanger through which oil in the oil loop system and substantially all the fuel flow to transfer heat between the oil and the fuel; and a secondary fuel-oil heat exchanger through which oil in the oil loop system and a portion of the fuel flow, to transfer heat between the oil and the fuel; and a fuel pump; wherein, under cruise conditions, fuel temperature on leaving the second heat exchanger is 120° C. to 200° C.

A gas turbine engine assembly includes a core engine that includes a core flow path where a core airflow is compressed in a compressor section, communicated to a combustor section, mixed with fuel and ignited to generate an exhaust gas flow that is expanded through a turbine section. The turbine section is coupled to drive the compressor section through an engine drive shaft. A tap is at a location up stream of the combustor section for drawing a bleed airflow. A bleed air heat exchanger places the bleed airflow in thermal communication with an auxiliary flow for heating the bleed airflow. An exhaust heat exchanger is configured to transfer thermal energy from the exhaust gas flow into the bleed airflow.

A tail cone ventilation system including a tail cone case defining a tail cone interior and a tail cone exterior, the tail cone having a forward portion and an aft portion separated axially along a tail cone axis; a distribution manifold located within the tail cone interior proximate the forward portion, wherein the distribution manifold comprises nozzles arranged radially around the axis, the nozzles configured to direct a cooling air over at least one electronic component within the tail cone interior and along an inner surface of the tail cone case; an air inlet fluidly coupled with the distribution manifold through ducting, the air inlet located externally from the tail cone interior; and a tail cone discharge located proximate the tail cone aft portion, the tail cone discharge being fluidly coupled with the distribution manifold.

A fuel injector cooling system, comprising: a blower including a blower inlet and a blower outlet; a valve including a valve inlet and a valve outlet; a duct defining a flow passage between the blower outlet and the valve inlet; a fuel injector including a flange and a valve housing extending radially outward from the flange, wherein the valve housing defines an outer surface of the fuel injector; and a cooling jacket defining an inner surface and a cooling air inlet in fluid communication with the flow passage, wherein the cooling jacket defines a cooling flow passage between the inner surface of the cooling jacket and the outer surface of the valve housing, wherein the cooling flow passage is in fluid communication with the flow passage.

A stator for an electric machine includes a stator core defining a slot extending in an axial direction from a first end to a second end, a wire disposed in the slot, a coolant pipe disposed in the slot and extending in the axial direction from the first end to the second end, the coolant pipe extending from a pipe inlet to a pipe outlet, and a thermally conductive potting material disposed in the slot between the coolant pipe, the wire, and the stator core.

A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows:(TOUTADTExit)2*EGTAHPCExit*10-11;wherein CSP is greater than 0.0001194×EGT2−0.103×EGT+22.14 and less than 0.0003294×EGT2−0.306×EGT+77.91; and wherein EGT is greater than 525 degrees Celsius and less than 1250 degrees Celsius.

A hybrid-electric aircraft propulsion system includes a gas turbine engine and an electrical assembly. The gas turbine engine includes a first rotational assembly. The first rotational assembly is rotatable about a rotational axis of the gas turbine engine. The first rotational assembly includes a first shaft, a bladed first compressor rotor, and a bladed first turbine rotor. The first shaft interconnects the bladed first compressor rotor and the bladed first turbine rotor. The electrical assembly includes a first motor-generator, a first motor control unit, a motor-generator (MG) cooling system, and a motor control unit (MCU) cooling system. The first motor-generator is coupled to the first shaft. The first motor control unit is electrically connected to the first motor-generator. The MG cooling system is connected in fluid communication with the first motor-generator. The MCU cooling system is connected in fluid communication with the first motor control unit. The MCU cooling system is independent of the MG cooling system.

Disclosed are a gas turbine generator set and mobile power generation equipment. The gas turbine generator set includes a gas turbine, a generator, and an equipment cabin. The gas turbine is arranged inside the equipment cabin. The generator is connected to the gas turbine for generating electric power. The generator may be arranged outside the equipment cabin, so that a heat dissipation load of the equipment cabin is reduced. As such, difficulty in designing an efficient heat dissipation and cooling scheme for the gas turbine inside the equipment cabin is reduced. The solution also facilitates convenient maintenance of the generator without opening the equipment cabin or entering the equipment cabin.

The invention relates to the technical field of aircrafts, in particular to a turbine engine pre-cooling system and an aircraft. The turbine engine precooling system comprises a propulsion channel, a precooler, an air turbine, a radiator, a circulating pump and an engine. The precooler is located on the side, close to the air inlet, of the engine. The precooler is provided with a first heat exchange medium channel and an air channel; the radiator is arranged on the outer side of the propelling channel and provided with a second heat exchange medium channel and a fuel channel, and the second heat exchange medium channel communicates with the first heat exchange medium channel; the air turbine is arranged in a hollow channel in the center of the precooler, and the air inlet direction of the air turbine is parallel to the extending direction of the air channel. The circulating pump is arranged on the radiator and connected with the air turbine so as to drive the heat exchange medium to circularly flow between the first heat exchange medium channel and the second heat exchange medium channel; and a valve is arranged at an exhaust port of the air turbine, so that the heat exchange efficiency and the working condition adaptability are improved.

A combustor for a gas turbine system includes a combustor body, a head end fuel nozzle assembly, and an axial fuel stage (AFS) injector. An air flow passage defined at least partially by the combustor body is configured to deliver a first portion of an air supply to the head end fuel nozzle assembly, and the AFS injector is configured to receive a second portion of the air supply. A valve is operatively positioned between the air supply and the AFS injector. With the AFS injector on (i.e., fueled), the valve is open, and a third, additional portion of the air supply flows to the AFS injector. With the AFS injector off (i.e., unfueled), the valve is closed to block the third portion flowing to the AFS injector and to deliver it to the head end fuel nozzle assembly, which reduces the firing temperature for highly reactive fuels.

A hybrid-electric propulsion system includes a gas turbine engine comprising a high pressure system and a low pressure system, an electric machine coupled to one of the high pressure system or the low pressure system, and a thermal management system defining one or more thermal management system flowpaths to provide one or more heat exchange fluids to the electric machine. A controller collects one or more signals from one or more sensing nodes connected to at least one of the thermal management system flowpaths or the electric machine, analyzes the one or more signals to detect a thermal anomaly corresponding to at least one of the electric machine or the thermal management system, and responsive to detecting the thermal anomaly, performs at least one of adjusting a flow of at least one heat exchange fluid of the one or more heat exchange fluids or derating the electric machine.

The invention relates to the technical field of engine thermal management, and particularly provides an active cooling method for coupling rotational flow cooling and regenerative cooling, which comprises the following steps: arranging a rectangular auxiliary inlet channel on the side wall of a rectangular main inlet channel of a regenerative cooling channel, and constructing a rotational flow regenerative cooling channel; when the rectangular main inlet channel performs regenerative cooling, coolant induced rotational flow is injected into the rectangular auxiliary inlet channel, and the rotational flow strength is regulated and controlled by adjusting the ratio of the flow of the coolant induced rotational flow to the total flow; and a cooling process simulation model is constructed, and the cooling effect of the rotational flow regeneration cooling channel is analyzed based on simulation calculation data. By means of the innovative rotational flow regeneration cooling channel design, controllable rotational flow is introduced into a traditional regeneration cooling channel, and the problems that a scramjet engine combustion chamber is insufficient in heat transfer in the extremely high-temperature environment and the fuel chemical heat sink utilization rate is low are solved. And fluid mixing is enhanced through rotational flow, a thermal boundary layer is destroyed, the fuel pyrolysis reaction is optimized, and efficient heat exchange is achieved.

A hydraulic system for a hydrogen fueled gas turbine engine. The hydraulic system comprises a closed-loop hydraulic fluid conduit (232), a hydrogen-hydraulic fluid heat exchanger (230) configured to exchange heat between cryogenic hydrogen fuel and hydraulic fluid within the hydraulic fluid conduit (232), and a second fuel-hydraulic heat exchanger (518) configured to exchange heat between a second, non-hydrogen fuel and hydraulic fluid in the hydraulic fluid conduit (232).

A tail cone ventilation system including a tail cone case defining a tail cone interior and a tail cone exterior, the tail cone having a forward portion and an aft portion separated axially along a tail cone axis; a distribution manifold located within the tail cone interior proximate the forward portion, wherein the distribution manifold comprises nozzles arranged radially around the axis, the nozzles configured to direct a cooling air over at least one electronic component within the tail cone interior and along an inner surface of the tail cone case; an air inlet fluidly coupled with the distribution manifold through ducting, the air inlet located externally from the tail cone interior; and a tail cone discharge located proximate the tail cone aft portion, the tail cone discharge being fluidly coupled with the distribution manifold.

Disclosed herein is an integrated in-line emergency lube pump assembly for a planetary gearbox, including a main lube pump, a coaxial lube pump coaxially aligned with and coupled to the planetary gearbox, and an oil tank having a main section and an auxiliary section. There is no physical boundary between the main section and the auxiliary section of the oil tank. When the detected pressure in the oil supply line is equal to or exceeds a predetermined pressure threshold, the main lube pump operates to causes the main pump to pump oil from the main section of the oil tank to the planetary gearbox. Alternatively, when the detected pressure in the oil supply line is less than the predetermined pressure threshold the main lube pump ceases operation to cause the coaxial lube pump to draw oil from the auxiliary section of the oil tank to the planetary gearbox.

A system for providing mobile electric power includes a first transport including a gas turbine and a generator and a second transport including an air inlet filter housing and an exhaust stack. The two transports are connectable to each other at an intake connection and an exhaust connection between facing sides of the two transports. The first transport includes first and second inlet ports, and first, second, and third outlet ports on the facing side of the first transport. The intake connection provides filtered air from the air inlet filter housing to the first inlet port as combustion air for the gas turbine, and the second inlet port as cooling air to cool an enclosure of the gas turbine. The exhaust connection exhausts air to the exhaust stack from the first, second and third outlet ports.

An aircraft engine assembly includes a gas turbine engine having an intake channel configured to receive an incoming flow of air and form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction, and an electric machine coupled with the low pressure shaft and located at the aft end of the gas turbine engine proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated.