1032results about "Turbine/propulsion engine cooling" patented technology

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 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.

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.

The invention relates to the technical field of aviation aircraft power systems, and discloses an aero-engine liquid hydrogen fuel conveying system and a control method thereof, which adopt multi-branch high-precision metering to realize wide-range high-precision flow regulation requirements, determine branch through-flow capability by a root order method, and realize high-precision flow regulation of the aero-engine liquid hydrogen fuel conveying system. The number of hydrogen supply branches, the available regulation ratio of the regulating valve and the total flow regulation ratio demand are matched, and the problem of high-precision fuel conveying in an extremely wide flow regulation range is solved. The circulating coupling heat exchanger is adopted to reasonably utilize waste heat of the system to raise the temperature of hydrogen fuel to the level suitable for combustion, reasonable optimization of heat / energy utilization of the system is achieved, fuel consumption is reduced, and circulating heat efficiency is improved. The hydrogen temporary storage tank is adopted to solve the problem that the supply flow of an upstream liquid hydrogen pump and a heat exchanger is not matched with the flow required by a downstream regulating valve due to hysteresis in the heat exchange process in the fuel flow regulation transient process.

Fuel thermal management systems and related methods are disclosed. An example system includes a flowline to carry a fluid, a waste heat recovery heat exchanger coupled to the flowline, the waste heat recovery heat exchanger to heat the fluid to a first temperature, and a feed tank coupled to the flowline, the feed tank including a first inlet and a second inlet, the first inlet to receive a first portion of the fluid from the waste heat recovery heat exchanger at the first temperature, the second inlet to receive a second portion of the fluid at a second temperature less than the first temperature, wherein the first portion of the fluid and the second portion of the fluid mixes in the feed tank to form a third portion having a third temperature between the first temperature and the second temperature.

The invention relates to an aero-engine, a thermal management method, a thermal management system and a computer readable medium. The thermal management method comprises the steps that it is judged that the aero-engine is in a first working condition, a second working condition, a third working condition or a fourth working condition; if in the first working condition, the three-way valve adjusts fuel oil to only enter the combustion chamber, the air-liquid heat exchanger and the air-fuel heat exchanger do not work, and the hybrid liquid cooling circulation system passes through the battery, the controller, the motor and the synovial fluid heat exchanger, then is cooled through the fuel liquid heat exchanger and then returns to the circulating pump; if in the second working condition, the fuel liquid heat exchanger, the air-liquid heat exchanger and the air-fuel heat exchanger all work; if in a third working condition, the fuel liquid heat exchanger and the air-fuel heat exchanger do not work; and if in a fourth working condition, the hybrid liquid cooling circulation system can realize internal heat balance, the hybrid cooling liquid directly returns to the circulation pump after absorbing heat through the controller, the motor and the synovial fluid heat exchanger, and then the battery is heated, so that the whole circulation is completed.

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 to exchange heat between the cryogenic fuel and at least one of (1) the hydraulic fluid or (2) a working fluid thermally coupled to the hydraulic system.

An assembly is provided for a turbine engine. This assembly includes an engine structure, and the engine structure includes a plenum, a reverse flow combustor, a nozzle structure and a cooling circuit. The reverse flow combustor is disposed in the plenum and includes a combustion chamber. The nozzle structure is arranged at an outlet from the combustion chamber. The nozzle structure includes a first platform, a second platform and a plurality of vanes arranged circumferentially about an axis. The vanes extend across the flowpath from the first platform to the second platform. The cooling circuit extends in the first platform and between a circuit inlet into the cooling circuit and a circuit outlet from the cooling circuit. The circuit inlet is fluidly coupled with and disposed along the plenum. The circuit outlet is fluidly coupled with and disposed along the flowpath.

Fuel thermal management systems and related methods are disclosed. An example system includes a flowline to carry a fluid, a waste heat recovery heat exchanger coupled to the flowline, the waste heat recovery heat exchanger to heat the fluid to a first temperature, and a feed tank coupled to the flowline, the feed tank including a first inlet and a second inlet, the first inlet to receive a first portion of the fluid from the waste heat recovery heat exchanger at the first temperature, the second inlet to receive a second portion of the fluid at a second temperature less than the first temperature, wherein the first portion of the fluid and the second portion of the fluid mixes in the feed tank to form a third portion having a third temperature between the first temperature and the second temperature.

An assembly is provided for an aircraft powerplant. This assembly includes a vane structure and a heat exchanger. The vane structure extends longitudinally to a leading edge of the vane structure. The vane structure includes a translating body, a stationary body and an inlet passage. The translating body is configured to translate longitudinally between a first position and a second position. The translating body is configured to form the leading edge of the vane structure and close an inlet into the inlet passage when in the first position. The translating body is configured to open the inlet into the inlet passage when in the second position. The heat exchanger is disposed within the stationary body. The inlet passage projects into the stationary body from the inlet into the inlet passage to the heat exchanger.

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*1⁢0-1⁢1;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 fuel system for a gas turbine engine including a compressor section, a combustion section, and a turbine section includes a fuel tank for storing a fuel, a plurality of heat exchangers downstream of the fuel tank, and a valve downstream of the plurality of heat exchangers. The valve includes a fuel inlet in fluid communication with the plurality of heat exchangers, a first fuel outlet in fluid communication with a first fluid pathway, and a second fuel outlet in fluid communication with a second fluid pathway. The fuel system includes a fuel cell including an anode inlet and an anode outlet. The anode inlet is in fluid communication with the first fluid pathway. The fuel system also includes a combustion chamber of the combustion section in fluid communication with the second fluid pathway and the anode outlet of the fuel cell and a controller communicatively coupled to the valve.

The invention discloses a variable-pore phase-change sweating cooling bionic structure based on fractal characteristics. The variable-pore phase-change sweating cooling bionic structure comprises a variable-pore porous sweating layer and a regulation and control unit, wherein the variable-pore porous sweating layer is composed of fractal flow equalizing plates which are sequentially arranged in a stacked mode in the heat flow direction; the fractal flow equalizing plate is of an N-stage bifurcated tree-shaped structure and is used for equally dividing one path of cooling working medium into 2N paths of cooling working medium, and N is an integer ranging from 4 to 6; the variable-pore porous sweating layer is divided into an upper-layer high-heat-flow area and a lower-layer low-heat-flow area in the thickness direction, the upper-layer high-heat-flow area is used for promoting phase change and efficient heat exchange of a cooling working medium, and the lower-layer low-heat-flow area is used for promoting liquid-phase permeation of the cooling working medium; and the regulation and control unit predicts the phase change leading edge position S (t) in real time based on a neural network, and dynamically regulates the cooling working medium flow Q (t) input into the inlet of the fractal flow equalizing plate according to the S (t). The invention further discloses a design method of the structure. By introducing the bionic variable porosity characteristic, coupling of flowing heat exchange is achieved, the cooling efficiency of thermal protection is improved, and consumption of a heat exchange working medium is reduced.

The invention discloses a hydrogen fuel SOFC actual combined cycle power generation system based on gas turbine pressurization, and belongs to the technical field of energy power equipment. In the starting stage, air is compressed by a gas compressor and then enters a combustion chamber to be combusted with hydrogen, exhausted gas is introduced into a turbine to do work, electric energy is provided for temperature rise before SOFC reaction through a motor, and turbine exhausted gas sequentially heats hydrogen and air at an inlet of the combustion chamber. Under the stable operation working condition, air sequentially passes through a gas compressor, a flow divider and a heat exchanger to be introduced into an SOFC cathode, hydrogen is heated through the heat exchanger and enters an SOFC anode to be subjected to an oxidation reaction, cathode and anode tail gas and supplemented gas enter a combustion chamber to be combusted, exhaust gas enters a turbine to do work and then is exhausted, and turbine exhaust gas sequentially heats hydrogen of SOFC anode inlet gas and air of cathode inlet gas; wherein the SOFC anode tail gas heats hydrogen supplemented by the combustion chamber through the heat exchanger, and the SOFC cathode tail gas heats air supplemented by the combustion chamber through the heat exchanger.

A double-flow turbomachine has an exhaust casing through which a hot primary flow passes and surrounded by a cold secondary flow, and a cone carried by the exhaust casing. The exhaust casing has a hollow radial arm passing through the primary flow to convey air into the cone in order to form a cooling flow for components located in the cone. The cone has an outer wall and an inner wall which runs alongside the outer wall and is radially spaced therefrom to delimit a thermal insulation space through which the cooling flow passes from upstream to downstream before it is discharged through an outlet opening at the end of the cone. It has an inner cover surrounding a component to channel the cooling flow so that it runs along the component to cool it before passing through the thermal insulation space to be discharged there.

A gas turbine package ventilation system is disclosed. The gas turbine package ventilation system comprises an enclosure (20, 20', 20'') surrounding a gas turbine (10), the gas turbine (10) including, in the direction of flow: a compressor (11), a combustor (12) and a compressor driving turbine (13), wherein a ventilation air distribution system is adapted to separate two divergent ventilation flows, a first ventilation flow, flowing downstream from a central portion of the gas turbine (10) towards an exhaust gas outlet (16) of the gas turbine (10), and a second ventilation flow, flowing upstream from a central portion of the gas turbine (10) towards a fresh air inlet (14) of the gas turbine (10).

This invention discloses a method for converting a gas turbine regenerative cycle to a humidified cycle. The compressor draws in air, pressurizes it, and discharges it into an aftercooler, where the air exchanges heat with cold water, thus cooling the air. A venting path is installed on the air pipeline between the aftercooler and the humidifier. The amount of air vented is adjusted according to different humidified cycle operating conditions to maintain a relatively constant compressor outlet pressure. Air and hot water come into counter-current contact in the humidifier. The humid air from the humidifier exchanges heat with the gas turbine exhaust in the regenerator before entering the combustion chamber. Natural gas mixes and combusts with the humid air before entering the turbine to perform work. The turbine drives the compressor and generator, which generate electricity. The turbine exhaust passes through the regenerator and economizer, where it exchanges heat with the humid air and water respectively to recover waste heat before being discharged into the atmosphere. This method achieves flow matching between the two systems without modifying the core components of the gas turbine (compressor and turbine), greatly reducing the difficulty of converting existing gas turbine units to a humidified cycle and maintaining a high system energy utilization rate.

A gas turbine engine assembly includes a core engine (20) that includes a core flow path where a core airflow (25) is compressed in a compressor section (24), communicated to a combustor section (56), mixed with fuel and ignited to generate an exhaust gas flow (74, 75) that is expanded through a turbine section (28). The turbine section (28) is coupled to drive the compressor section (24) through an engine drive shaft (88). A tap (64) is at a location up stream of the combustor section (56) for drawing a bleed airflow (66, 65). A bleed air heat exchanger (72) places the bleed airflow (66, 65) in thermal communication with an auxiliary flow (68) for heating the bleed airflow (66, 65). An exhaust heat exchanger (76) is configured to transfer thermal energy from the exhaust gas flow (75) into the bleed airflow (66, 65).

The present application relates to the technical field of cooling water system control, especially to a self-shutdown control method for a closed cycle cooling water system of a gas turbine combined cycle unit. The closed cycle cooling water system comprises multiple closed water pumps and multiple closed coolers. The present application realizes full automatic control of three closed water pumps, optimizes the standby logic of the closed water pumps, sets up a primary standby function and a secondary standby function, and integrates all the three closed water pumps into the shutdown control process. Through the closed cooler shutdown process and the closed water pump shutdown process, the most suitable closed cooler and closed water pump are automatically selected for shutdown, and the shutdown of the closed cycle water system is realized in an orderly and efficient manner according to the established rules. Without the need for human judgment and operation by the operation personnel, the full automatic shutdown of the closed cycle cooling water system is realized without disturbing the stable operation state of the unit, and the labor intensity of the operation personnel is reduced, and the safety and the automation level of the unit are improved.

This disclosure relates to a gas turbine and a thermal power generation system. The gas turbine includes a main compressor assembly, an auxiliary compressor, a cooling assembly, and a turbine assembly. The main compressor assembly includes a primary compression unit, a final compression unit, and multiple intermediate compression units located between the primary and final compression units. Each intermediate compression unit has an exhaust port communicating with its internal cavity. The intake port of the auxiliary compressor is connected to the exhaust port of one of the intermediate compression units via the cooling assembly. The exhaust port of the auxiliary compressor is connected to the internal chamber of the turbine assembly. The cooling assembly is disposed between the intermediate compression unit connected to the intake port of the auxiliary compressor and the turbine assembly, and is used to cool the gas before it enters the internal chamber of the turbine assembly.

This stator blade comprises a blade body having a cross-sectional shape that forms a blade profile. The blade body includes: a trailing end space and an adjacent space that are located between a leading edge and a trailing edge; and a plurality of trailing end cooling passages that penetrate from the trailing end space to the trailing edge. The adjacent space is located further on the leading edge side than the trailing end space. The edge of the trailing end space on the leading edge side is open and communicates with the trailing end space. A plurality of divider ribs, which are aligned in the blade height direction and divide the trailing end space in the blade height direction, and a plurality of pins are arranged in the trailing end space. The plurality of pins all both joined to a positive pressure-side inner wall surface and a negative pressure-side inner wall surface that define the trailing end space. Among the plurality of divider ribs, at least one divider rib forms an extending divider rib which extends to the interior of the adjacent space and which is joined to an inner wall surface demarcating the adjacent space.

Disclosed is a turbine airfoil including an airfoil body including an outer wall configured to define an internal space, in which a cooling hole is formed in the outer wall and allows the internal space S and an external space of the airfoil body to communicate with each other. The inner surface of the outer wall, which defines the cooling hole, includes a first inner surface, and a second inner surface configured to define an outer hole region of the cooling hole H that communicates with the external space, in which the outer hole region is provided as a plurality of outer hole regions spaced apart from one another, and in which the outer wall includes a partition section provided between the two adjacent outer hole regions.

A turbine engine (20') includes a duct (16') defining an annular passage, at least two heat exchangers (72') arranged within the annular passage and spaced circumferentially apart, a passage (76) between the at least two heat exchangers (72'), and a forward flow control device (82) operable for controlling airflow through the passages (76).

A rotary combustion engine includes a cylindrical housing with a plurality of internal cavities. Each of the internal cavities is a cylindrical segment of the housing, and a plurality of stationary blades are attached to its inner wall. The stationary blades extend inward to ride along the surface of a central shaft, contacting the shaft with a seal. Each cavity contains a rotor blade attached to the shaft and configured to drive the shaft. Each rotor blade extends outwardly to an inner curve of the housing with a sealing means. Each of the plurality of rotor blades has a reciprocating circular motion inside each of the cavities along the shaft between each of the stationary blades, thereby generating rotational power, delivering compressed rotational force, and provides cooling for the cavity.

The utility model relates to a gas turbine intake air treatment technical field, concretely relates to a gas turbine intake air cold and hot double -effect integrated heat exchange device, the utility model discloses according to the special requirement of gas turbine to intake air, also can install the intake air cooling heat exchanger before the intake air filter element, heating and cooling share a finned tube heat exchanger, indispensable temperature control system is designed, and the reasonable control of intake air temperature / humidity is realized, in most areas, not only can effectively heat the intake air of gas turbine, prevent the wet block and ice block of filter element, but also can make the intake air temperature of gas turbine reduce 3~10 DEG C in high temperature season, improve the output of gas turbine 3~10%, reduce the fuel consumption of gas turbine 1.2~4%, reduce NO X Emission 30% or more, the engineering cost and energy consumption are lower.

The oil-electricity hybrid aero-engine comprises a tail nozzle and an outer duct, and further comprises a motor and an electronic cooling loop, the system comprises a thermoacoustic engine and a first thermoacoustic heat pump, the thermoacoustic engine is arranged at the tail nozzle and used for absorbing heat of tail gas at the tail nozzle, and the first thermoacoustic heat pump is arranged at the tail nozzle and used for absorbing heat of the tail gas at the outer duct. Sound power is output; the first thermo-acoustic heat pump comprises a first cold-end heat exchanger, a first heat regenerator and a first hot-end heat exchanger, the first cold-end heat exchanger is used for absorbing heat to form a low-temperature area and is used for cooling the electronic cooling loop, part of heat is transmitted to the first hot-end heat exchanger through the first heat regenerator by utilizing acoustic power, and the first hot-end heat exchanger releases heat to the outer duct. The system utilizes a thermo-acoustic energy conversion technology to enhance the cooling of electronic parts, and can also recover tail gas waste heat to convert the tail gas waste heat into heat with a higher temperature level.

An aircraft propulsion system (20) includes a bottoming cycle (60) where a working fluid is heated and expanded through a bottom turbine to generate shaft power. A first heat exchanger (44) provides thermal communication between a fuel cell exhaust flow and the working fluid of the bottoming cycle (60). The working fluid is cooled by a cryogenic fuel flow (56) in a fuel / working fluid heat exchanger (54).