12711 results about "Combustor" patented technology

The invention relates to the technical field of control and regulation systems, in particular to a hearth auxiliary regulation control system which comprises a data processing module, an analysis module, a control module, a regulation and control module, an optimization module and a safety early warning module. The data processing module collects temperature data of a hearth combustor area, a water cooling wall and a tail flue in real time through a distributed temperature sensor network, and a dynamic combustion field three-dimensional model is generated through Kalman filtering denoising, redundancy cross validation and a temperature compensation algorithm in combination with combustor flame position signals and an air door opening state. And the optimization module updates the strategy weight based on the combustion efficiency improvement rate, and reinforces learning to extend the adaptive boundary of the strategy library. And the safety early warning module calculates the matching coefficient of the pipe wall temperature and the allowable stress in real time, triggers a grading intervention strategy to reduce the overtemperature risk, forms full-process closed-loop control, and improves the thermal efficiency and the operation safety of the boiler.

The invention discloses a rotational flow micro-mixing nozzle with built-in plasma combustion supporting and a micro-mixing combustor. The rotational flow micro-mixing nozzle comprises a tangential rotational flow part, a fuel pipe and a mixing pipe which are coaxially arranged, and plasma combustion supporting exciters arranged on the fuel pipe and the mixing pipe. The plasma combustion-supporting exciter is integrated in the nozzle, the structural design of the micro-mixing combustor can be simplified, and the size and complexity of the micro-mixing combustor are reduced; the plasma combustion-supporting exciter is installed at the top of the mixed gas channel of the rotational flow micro-mixing nozzle, stable adjustment of the combustion process under the variable working conditions can be achieved by controlling opening and closing and the discharge intensity of the plasma combustion-supporting exciter when the load changes, and meanwhile the response speed under the different working conditions is increased; air or an oxidizing agent can be fully mixed with fuel at a specific rotational flow degree through rotational flow air flow generated by the tangential rotational flow component, the combustion range can be widened, and the stable combustion effect under the extreme working condition is enhanced.

The invention discloses an ammonia-coal co-combustion burner which comprises a plasma pulverized coal pre-combustion level module, an ammonia two-level combustion module, a pulverized coal main combustion level module and an air classification secondary air module. Plasma is generated by discharging between electrodes in the plasma pulverized coal pre-combustion stage module, pulverized coal carried by primary air is combusted through a plasma fire core area to generate first pulverized coal flame, and the first pulverized coal flame ignites ammonia fuel on the upper stream of the ammonia two-stage combustion module to generate stable ammonia-rich flame. Combustion products and unburnt reactants are mixed with downstream air in the ammonia two-stage combustion module to generate lean-ammonia flame for heating pulverized coal in the pulverized coal main combustion stage module to generate second pulverized coal flame, and the pulverized coal and ammonia gas are promoted to be fully combusted in the air classification secondary air module through secondary air, so that stable combustion of ammonia-coal mixed fuel is realized; the problems that ammonia coal combustion is unstable, NOx emission is too high and ammonia escapes are effectively solved.

A method of operating a gas turbine engine comprising an engine core comprising a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system arranged to supply oil to the gearbox; and a heat exchange system comprising an air-oil heat exchanger through which the oil flows, a fuel-oil heat exchanger through which the oil and the fuel flow; and a valve arranged to allow a proportion of the oil sent via at least one of the heat exchangers to be varied, is disclosed, the method comprising controlling the valve such that, under cruise conditions, an oil flow ratio of:rate⁢ of⁢ oil⁢ flow⁢ into⁢ air-oil⁢ heat⁢ exchangerrate⁢ of⁢ oil⁢ flow⁢ into⁢ fuel-oil⁢ heat⁢ exchangeris in the range from 0 to 0.59.

A method of operating a gas turbine engine including: an engine core with a turbine, compressor, a fuel combustor, and a core shaft connecting the turbine to the compressor; a fan; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system arranged to supply oil to the gearbox; and a heat exchange system including: an air-oil heat exchanger through which the oil flows; a fuel-oil heat exchanger through which the oil and the fuel flow; a temperature sensor downstream of the fuel-oil heat exchanger; and a valve arranged to allow oil or air flow rate through at least one heat exchanger to be varied, the method including: determining if the fuel temperature has increased above a set threshold at cruise conditions; and in response, controlling the valve so as to change the flow rate through the heat exchanger.

The invention relates to the technical field of steam turbine testing, in particular to a testing system for mismatch correction of a wide-load-interval thermal power generating unit speed regulation model, which comprises a thermodynamic system component, a speed regulation system device, a dynamic signal acquisition device and a control device, a dynamic signal processing module analyzes the relevance between main steam pressure sudden change and control valve opening oscillation in real time, a dynamic response database including a time sequence is constructed, and a control valve opening feed-forward compensation instruction and a combustor fuel distribution gradient adjustment signal are generated according to correction parameters. Power over-modulation wave crests caused by sudden load change are suppressed, and primary frequency modulation action delay is reduced; and the simulation verification module compares measured data with model output through an error feedback mechanism, dynamically adjusts high-pressure cylinder flow characteristic curve calibration parameters, optimizes a control strategy in combination with a network-related safety margin evaluation result, and realizes improvement of dynamic response precision in a full peak regulation interval and enhancement of power grid frequency stability.

The invention relates to the technical field of mechanical stress and thermal deformation monitoring, in particular to a combustion fine adjustment method based on full-hearth expansion monitoring, which is characterized in that expansion sensor arrays are arranged on each layer of a main burner area of a boiler hearth, and vibration monitoring units are arranged in a layered manner from a main burner to an SOFA air area synchronously; normal deformation, three-dimensional expansion displacement data and vibration spectrum data of the four furnace walls are collected; generating a standardized monitoring data set through wavelet noise reduction and temperature compensation processing, fusing boiler operation parameters, and inputting the boiler operation parameters into the reinforcement learning model to generate a combustor adjustment instruction; water-cooled wall expansion rate change data is collected in real time to optimize model weight and update a control strategy, and a closed-loop feedback mechanism is formed. According to the method, through multi-source sensing data fusion and dynamic modeling, sub-meter analysis of thermal load distribution and quick response of the combustion state are achieved, the boiler operation efficiency is improved, and the service life of equipment is prolonged.

A method of operating a gas turbine engine having a combustor having a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber. The fuel spray nozzles have a first subset of fuel spray nozzles and a second subset of fuel spray nozzles. The combustor is operable in a condition in which the first subset of fuel spray nozzles are supplied with more fuel than the second subset of fuel spray nozzles. A ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. The method includes providing a fuel to the plurality of fuel spray nozzles having a calorific value of at least 43.5 MJ / kg. A gas turbine engine can be for an aircraft.

A fuel cavity, an air cavity and a combustion cavity are sequentially and coaxially arranged at intervals in a sealed mode from bottom to top, a plurality of nozzle assemblies are fixedly arranged in the air cavity, and the nozzle assemblies are symmetrically arranged relative to the center of the air cavity to form an annular multi-layer array structure; the nozzle assembly comprises an air flow guide component, a tangential rotational flow component, a fuel pipe and a mixing pipe which are coaxially arranged. Through a combined adjusting mechanism of the air flow guide component and the tangential rotational flow component, different rotational flow degrees of a single nozzle can be realized, so that the nozzle adapts to fuels with different properties; according to different working conditions, different air rotational flow degrees can be set for the nozzle assemblies on different layers in a staggered mode, or different air rotational flow degrees can be set for the multiple nozzle assemblies on the same layer in a staggered mode, so that staged coupling combustion of the multiple nozzle assemblies in a large space is achieved, and the purposes of enhancing combustion stability and improving temperature uniformity are achieved.

A method of operating a gas turbine engine includes an engine core with a turbine, compressor, fuel combustor, and core shaft; a fan upstream of the engine core; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system supplying oil to the gearbox; and a heat exchange system including: air-oil and fuel-oil heat exchangers, and wherein the oil loop system branches such that part of the oil flows along each branch and the air-oil and fuel-oil heat exchangers are parallel on different branches; and a modulation valve allows the oil sent via each branch to be varied, the method including controlling the heat exchange system wherein, under cruise conditions, a heat transfer ratio of:rate⁢ of⁢ heat⁢ transfer⁢ from⁢ oil⁢ to⁢ airrate⁢ of⁢ heat⁢ transfer⁢ from⁢ oil⁢ to⁢ fuelis in the range from 0 to 0.67.

An aircraft gas turbine engine includes a combustor, with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include first and second subsets. The first subset of nozzles is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A lean cruise nvPM emissions index ratio isEIcruise⁡(lean),SAFEIcruise⁡(lean),FF,where EIcruise(lean),SAF isEImaxTO,SAF+EIclimb,SAF2and EIcruise(lean),FF isEImaxTO,FF+EIclimb,FF2.EImaxTO,SAF is nvPM emissions index operating at 100% available thrust and EIclimb,SAF is the nvPM emissions index operating at 85% available thrust if fuel includes sustainable aviation fuel. EImaxTO,FF is the nvPM emissions index operating at 100% available thrust and EIclimb,FF is the nvPM emissions index at 85% available thrust if fuel is a fossil-based hydrocarbon. The lean cruise nvPM emissions index ratio is less than 1.

A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. A thrust nvPM emissions index ratio is defined as:EImaxTO / FmaxTOEIidle / Fidlewhere: EIidle is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine if operating at around 7% available thrust for given operating conditions; EImaxTO is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions; FmaxTO is the thrust of the gas turbine engine at around 100% available thrust in kN for the given operating conditions; and Fidle is the thrust of the gas turbine engine at around 7% available thrust in kN for the given operating conditions. The thrust nvPM emissions index ratio is greater than 0.001. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.

Systems and methods for providing lubrication to a gas turbine engine are described. An example gas turbine engine includes a plurality of bearings disposed along an axial length of a rotatable shaft to provide rotational support between a static structure and at least one of a compressor and a turbine of the gas turbine engine. The gas turbine engine further includes a fuel system operable to deliver a mixture of fuel and lubricant sequentially to at least one of the plurality of bearings and the combustor so that during operation of the gas turbine engine, at least the lubricant within the mixture is in a liquid phase prior to being received by the at least one of the plurality of bearings, and in a vapor phase once the lubricant is delivered to the at least one of the plurality of bearings.

A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:6. A MTO nvPM emissions index ratio is defined as:EImaxTO,SAFEImaxTO,FFwhere: EImaxTO,SAF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EImaxTO,FF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1. The gas turbine engine is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles. Also disclosed are methods of operating the gas turbine engine.

Disclosed is a water heater including a case provided with an air supply hole and an air discharge hole, a burner provided in an interior of the case, and a humidified air supply part that provides humidified air to the burner together with a fuel. The humidified air supply part includes a first accommodation tank that accommodates water to generate vapor, and a humidified air supply tank that provides a space for supplying the vapor supplied from the first accommodation tank to air introduced into an interior of the case through the air supply hole.

A gas turbine engine for an aircraft. The gas turbine engine comprising: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. An MTO nvPM emissions index ratio-modified fuel flow is defined as:EImaxTO,SAFEImaxTO,FF×Wf,maxTOwhere: EImaxTO,SAF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); EImaxTO,FF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and Wf,maxTO is the mass flow rate of fuel provided to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at around 100% available thrust for the given operating conditions. The MTO nvPM emissions index ratio-modified fuel flow of the gas turbine engine in kg / s is less than 2. The gas turbine engine is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.

A gas turbine engine includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The nozzles include a first and second subset. Each of the nozzles of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to nozzles in the second subset from 1:3 to 1:6. A fuel-flow nvPM emissions index ratio isEIidle×Wf,idleRImaxTO×Wf,maxTO.EIidle and EImaxTO are the system loss corrected nvPM emissions index in mg / kg operating at around 7% and 100% available thrust for the given operating conditions, respectively. Wf,idle and Wf,maxTO are the rate of fuel flow to the fuel spray nozzles in kg / s at around 7% and 100% available thrust for the given operating conditions, respectively. The fuel-flow nvPM emissions index ratio of the gas turbine engine is less than 0.3.

The invention discloses a plasma-enhanced direct-injection type liquid ammonia atomizing nozzle and a liquid ammonia combustor, and belongs to the field of liquid fuel combustion. The plasma-enhanced direct-injection type liquid ammonia atomizing nozzle comprises a liquid ammonia channel, a compressed air channel, an atomizing channel, a secondary air channel and a heating assembly, wherein the outlet end of the compressed air channel and the outlet end of the liquid ammonia channel are both communicated with the atomizing channel; the heating assembly comprises a high-voltage electrode, a ground electrode and a central air channel, the ground electrode is arranged on the outer side of the high-voltage electrode, and an outlet of the central air channel faces a discharge area formed by the high-voltage electrode and the ground electrode; a discharge area formed by the high-voltage electrode and the ground electrode is close to the liquid ammonia channel; a ground electrode hole is formed in the position, close to the liquid ammonia channel, of the ground electrode. Self-evaporation of preheated liquid ammonia is achieved through small plasma flames, efficient atomization of large-flow liquid ammonia is achieved, and consumption of atomization media is reduced while stable combustion is achieved.

A turbine engine for an aircraft including a combustor, a turbine, and a steam system. The combustor generates combustion gases. The turbine receives the combustion gases and is drivingly coupled to a shaft. The steam system includes a boiler, a superheater, and a steam turbine. The boiler receives water and is fluidly connected to the combustor to receive a first portion of the combustion gases from the combustor and to boil the water to generate steam. The superheater is located downstream of the combustor to receive a second portion of the combustion gases from the combustor and to heat the steam. The steam turbine is fluidly coupled to the superheater to receive the steam from the superheater and to rotate the steam turbine. The steam turbine is drivingly coupled to the shaft.

A gas turbine engine includes: a rich burn, quick quench, lean burn combustor having a number of fuel spray nozzles in the range 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. An nvPM emissions index ratio is defined as:EIidle×Wf,idleEImaxTO×Wf,maxTOwhere: EIidle and EImaxTO are respectively the nvPM emissions index in mg / kg of gas turbine engine if operating at around 7% or 100% available thrust for given operating conditions; Wf,idle and Wf,maxTO are respectively the rate of fuel flow to fuel spray nozzles in kg / s at around 7% or 100% available thrust for given operating conditions. The fuel-flow nvPM emissions index ratio is less than 0.08. The gas turbine engine is configured to provide fuel including sustainable aviation fuel to fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.

The invention relates to a SOEC (Solid Oxide Electrolyte Cell) hydrogen production system and an operation method thereof. The SOEC hydrogen production system comprises an SOEC galvanic pile module, a deionized water pump, a deionized water preheater, a heat exchange type steam generator, an ejector, a feed gas preheater, a hydrogen preheater, a fan, an air preheater, a combustor, a stop valve, a heat tracing band, a power supply, a control system, a pipeline, a temperature sensor and the like. The system is provided with the combustor, the heat exchange type steam generator and the heat exchanger, heat generated by combustion of natural gas is used for providing heat for high-temperature raw material gas and high-temperature air required by the SOEC hydrogen production system in a heat exchange mode, meanwhile, heat of high-temperature gas at an outlet of an SOEC galvanic pile module is recycled, and the power consumption of the SOEC hydrogen production system is effectively reduced; in addition, the high-temperature raw material gas containing hydrogen is heated in a heat exchange mode instead of an electric heating mode, so that the safety of the hydrogen production system is effectively improved; the system is low in power consumption, stable, reliable and high in operability.

A turbine engine for an aircraft. The turbine engine includes a combustor, a turbine, a boiler, a steam turbine, and a reheat boiler. The combustor generates combustion gases and the turbine is positioned downstream of the combustor to receive the combustion gases and to rotate the turbine. The boiler is positioned downstream of the combustor to receive the combustion gases and to boil water to generate steam. The steam turbine is fluidly coupled to the boiler to receive the steam from the boiler and to rotate the steam turbine. Each of the turbine and the steam turbine is drivingly coupled to a core shaft to rotate the core shaft. The reheat boiler is fluidly coupled to the steam turbine to receive the steam from the steam turbine and to reheat the steam. The combustor is fluidly coupled to the reheat boiler to receive the reheated steam.

A gas turbine engine for an aircraft. The gas turbine engine comprising: a rich burn, quick quench, lean burn (RQL) combustor having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. An MTO nvPM emissions index ratio is defined as:EImaxTO,SAFEImaxTO,FFwhere: EImaxTO,SAF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EImaxTO,FF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the fuel spray nozzles is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1. The gas turbine engine is configured to provide fuel comprising a SAF to the fuel spray nozzles. Also disclosed is a method of operating a gas turbine engine.

A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:6. A MTO nvPM emissions index ratio is defined as:EImaxTO,SAFEImaxTO,FFwhere: EImaxTO,SAF is the nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EImaxTO,FF is the nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1. The gas turbine engine is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles. Also disclosed are methods of operating the gas turbine engine.

A gas turbine engine includes: a rich burn, quick quench, lean burn combustor having a number of fuel spray nozzles in range of 14-22 or a number of fuel spray nozzles per unit engine core size in range 2 to 6. An MTO nvPM emissions index ratio is:EImaxTO,SAFEImaxTO,FF×Wf,maxTOwhere: EImaxTO,SAF and EImaxTO,FF are respectively the nvPM emissions index in mg / kg when operating around 100% available thrust for given operating conditions if fuel provided to fuel spray nozzles includes either sustainable aviation fuel (SAF) or fossil-based hydrocarbon fuel; Wf,maxTO is mass flow rate of fuel provided to fuel spray nozzles in kg / s when gas turbine engine is operating at around 100% available thrust for given operating conditions. MTO nvPM emissions index is less than 2. The gas turbine engine is configured to provide fuel including SAF to fuel spray nozzles. Also disclosed is method of operating gas turbine engine.

The invention relates to the technical field of pulverized coal fired boiler combustors, and particularly discloses a double-high-concentration-area oil-saving ignition pulverized coal combustor. The double-high-concentration-area oil-saving ignition pulverized coal burner comprises a flow guide device and an ignition device arranged in a combustion cylinder. The flow guide device comprises a main cylinder and a turbulent flow structure arranged on the periphery of the main cylinder, the main cylinder is connected to the inner wall of the combustion cylinder through a supporting piece, and the turbulent flow structure is of a structure protruding in the radial direction and used for disturbing the flow direction of mixed gas. Therefore, the problem that in the prior art, when the structural design of an oil-saving ignition pulverized coal burner is in a non-ignition working condition, the pulverized coal concentration distribution mode of'thick inside and thin outside 'is opposite to the design of'thick outside and thin inside' of an original burner of a boiler, and the air powder flow field characteristics of an outlet of the burner and the interior of a hearth are prone to being remarkably influenced is solved.

The utility model discloses a stable combustion device for high-pressure rotating sliding arc plasma ignition. The stable combustion device comprises a cathode assembly, a locknut, an insulating supporting ring, a connecting plate, a rear side air inlet, an anode air duct, a rotational flow air chamber and a front side air inlet. The cathode assembly comprises a binding post, a guide rod, a central electric pole, a conical cathode seat and a central cathode inner core which are sequentially connected into a whole in a threaded mode, the cathode assembly is fixed to the center of the anode air duct through an insulating supporting ring, and a multi-groove flow guide ring, a rotational flow air chamber and a rotational flow hole are formed in the front side of the anode air duct; after medium gas is introduced, under the excitation of an electric field, a plasma arc generated between the electrodes slides to the central cathode inner core along the conical cathode seat under the pushing of airflow, and a cycle process of breakdown-stretching-arc quenching-re-breakdown is periodically carried out, so that sliding arc plasma is generated; meanwhile, the plasma arc has the advantages of being high in energy density, improving the chemical reaction rate and the like, and ignition and stable combustion functions are provided for the industrial combustor. The structure is simple, starting is rapid, and the electrode service life is long.

A gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3 that is defined byVGVG⁢B.VG is a gutter volume of the gutter and VGB is a gearbox volume. A gas turbine engine includes the gearbox assembly, a fan, a combustor that generates combustion gases, a turbine, a nacelle that surrounds the fan, and a steam system. The combustor and the turbine define a core air flowpath. Core air flows through the core air flowpath. The steam system that extracts water from the combustion gases, vaporizes the water to generate steam, and injects the steam into the core air flowpath to add mass flow to the core air. A bypass ratio of the gas turbine engine is in a range of 18:1 to 100:1.

Systems and methods of the present disclosure includes a mixer configured to mix air and hydrocarbons and a burner configured to receive the mixed air and hydrocarbons and to generate syngas. The system also includes a cooling system configured to receive the syngas and to cool the syngas to form cooled syngas. The system further includes a collector configured to collect carbon from the cooled syngas as soot. Moreover, the system includes a flare stack configured to receive the cooled syngas and to burn off at least part of the cooled syngas.

The invention discloses a mobile modular ammonia-hydrogen fuel cell system and a power generation method. The mobile modular ammonia-hydrogen fuel cell system comprises an ammonia storage module, an ammonia filling module, a hydrogen supply module and a fuel cell module, the ammonia storage module is connected with the ammonia filling module, the ammonia filling module is connected with the hydrogen supply module, and the fuel cell module is communicated with the hydrogen supply module; the hydrogen supply module comprises a plurality of reactors and combustors which are separately arranged; the fuel cell module comprises a fuel cell and a lithium battery which are connected in series; wherein two reactors are flue gas heating reactors, and the other reactor is an electric heating reactor; the combustor is connected with the flue gas heating reactor; the two flue gas heating reactors are simultaneously communicated with the electric heating reactor; the lithium battery is communicated with the electric heating sheet, and a purification unit is arranged between the electric heating reactor and the fuel cell. The mobile modular ammonia-hydrogen fuel cell system can realize flexible carrying of equipment and rapid supply of fuel, can realize rapid power supply under emergency conditions, and has the advantages of being green, pollution-free, wide in adaptive condition and the like.