7107 results about "Combustor" patented technology

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.

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.

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.

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 machine learning multi-objective optimization control method based on coal quality characteristics and combustion state feedback, and belongs to the technical field of industrial process automatic control. According to the method, collaborative optimization control over the combustion efficiency, NOx emission and combustion stability is achieved under the multi-working-condition operation condition of the combustor. According to the method, coal quality characteristic parameters and combustion state data are collected, working condition clustering is carried out after data fusion, and a historical combustion operation data set coupling coal quality characteristics and combustion states is constructed; a combustion state evaluation model and a combustion control model are established by adopting machine learning, on the basis of identifying working conditions in real time and evaluating combustion performance, a control parameter adjustment amount is generated in combination with trend extrapolation and a multi-objective cost function, and a combustion state feedback rolling correction model and a control strategy are utilized to realize combustion control. The self-adaptive closed-loop collaborative optimization control of the combustion process under multiple coal types and multiple working conditions is realized, the boiler efficiency is improved, the NOx emission is reduced, and the operation robustness is enhanced.

The invention relates to the technical field of boiler control, and discloses a gas-fired boiler starting control method and system. The method comprises the following steps: acquiring initial operation parameters such as hearth pressure, water inlet temperature, flue gas oxygen content and fuel valve opening of the gas-fired boiler, establishing a thermodynamic state model in a starting stage according to the initial operation parameters, and dividing dynamic response thresholds of different load intervals; and calculating a theoretical fuel supply rate of the combustor based on the model, generating and dynamically updating a fuel supply correction coefficient in combination with a threshold value, synchronously marking key nodes according to abnormal distribution of hearth temperature, and adjusting a spatial air distribution strategy of the combustor. And after the correction coefficient is subjected to multi-stage buffer smoothing, a fuel control instruction is output to an executing mechanism. And switching to steady state control after the boiler reaches the preset load. The method adapts to the characteristics of a power plant unit, so that the starting process is more stable, the damage to a heated surface and pollutant emission are reduced, and rapid grid connection is assisted.

A combustor includes a combustion liner that defines a combustion chamber and an outer sleeve spaced apart from the combustion liner, such that an annulus is defined between the combustion liner and the outer sleeve. The combustor further includes a fuel injection assembly having a boss coupled to at least one of the combustion liner and the outer sleeve. The boss includes an inner flange coupled to the combustion liner, an outer flange, and a body extending between the outer flange and the inner flange. The boss defines a cooling circuit having a post-impingement plenum defined in the inner flange between an inner wall of the inner flange and an impingement plate in cooling proximity to the inner wall; and a plurality of film cooling channels that are defined in the inner flange and fluidly couple the post-impingement plenum to the combustion chamber.

A gas turbine engine for an aircraft includes a fan and a turbo-engine having a combustor that combusts compressed air and fuel to generate combustion gases and a power turbine coupled to a shaft for driving the fan. The gas turbine engine includes a direct contact condenser for condensing steam in a first flow into liquid water droplets and to generate a second flow including the liquid water droplets, a water tank for collecting the liquid water droplets of the second flow, and a water pump for introducing the liquid water droplets from the water tank into the direct contact condenser to act as condensation nuclei for the steam in the first flow. The first flow is an exhaust-vapor flow extracted from the combustion gases and the second flow is an exhaust-condensate flow from the direct contact condenser.

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.

The invention discloses a condensation-free efficient heat exchange control method and system for a full-premixing water heater, and the method comprises the steps: collecting the multi-source information, such as air inlet temperature, humidity, gas flow, flue gas oxygen concentration and wall temperature, in real time, calculating a predicted dew point based on a dynamic dew point prediction model, and setting a dynamic safety boundary; the condensation risk is judged by comparing the wall temperature with the safety boundary, and active intervention is carried out in a staged cooperation mode that fuel gas and air volume are continuously adjusted firstly and then a partial combustion zone sectioning valve is rapidly turned off; the system comprises a main controller, a mesh-shaped infrared radiation burner, a sectional valve assembly, a gas proportional valve, a frequency conversion fan and related sensors. According to the invention, the system runs close to the condensation boundary through intelligent control, and the advantage of high operation heat efficiency is realized on the premise of ensuring high reliability.

A mobile heater adapted to heat air from the environment and to supply the heated air into a cabin like a recreational vehicle or a tent, the heater comprises a housing, a supply air inlet arranged at a side of the housing, a supply air outlet arranged at another side of the housing, a supply air channel inside the housing extending the between the supply air inlet and the supply outlet, means for generating a supply air flow inside the supply air channel from the supply air inlet to the supply air outlet, and flexible supply air hose that can be removably coupled to the supply air outlet and by means of which the heated air can be led inside the cabin, the heater further comprises a burner chamber inside the housing, wherein the supply air channel passes through the burner chamber, a burner arranged inside the burner chamber, a heat exchanger arranged inside the burner chamber and being designed as a tube section that forms at least a section of the supply air channel, a control unit and a user interface , wherein the burner extends below and along a vertical projection of the heat exchanger, and the heat exchanger has a plurality of firstlamellas and second lamellas, wherein the first lamellas extend from the outer surface of the tube section inside the burner chamber and parallel to a vertical plane, and the second lamellas extend from the inner surface of the tube section inside the supply air channel and along the direction of the supply air flow.

The invention discloses a combustion chamber and turbine guider integrated composite flow channel outlet flow field uniformity evaluation method, and belongs to the technical field of gas turbine pneumatic design. The method comprises the steps that a geometric model of a combustion chamber and turbine guider integrated composite flow channel is built, a combustion chamber function section and a guider pneumatic section are fused through integral pneumatic modeling to form a continuous flow channel, a front section adopts a diffusion structure to form a low-speed combustion area, and a rear section adopts a gradually-shrinking channel to achieve gas accelerated expansion; performing three-dimensional cold-state flow field numerical simulation by adopting a structured grid division and delay separation vortex simulation technology; flow parameters including total pressure, Mach number, airflow angle and vorticity are extracted from the cross section of a flow channel outlet, conversion from a Cartesian coordinate system to a polar coordinate system is carried out, and a uniformly distributed aerodynamic parameter field is reconstructed by using an interpolation algorithm; the directional coupling degree is calculated based on the spatial correlation of the pressure gradient vector and the vorticity vector, and the vortex-pressure coupling entropy production index representing the overall aerodynamic loss of the flow channel is established in combination with the pressure pulsation intensity and the vorticity concentration ratio. According to the method, a three-dimensional evaluation system is constructed by integrating circumferential and radial non-uniformity and entropy production parameters, and accurate quantitative evaluation of flow characteristics and uniformity of the combustion chamber and turbine integrated structure is realized.

The invention relates to the technical field of flue gas recirculation, in particular to a bell-type annealing furnace oxygen-rich combustion flue gas recycling system and method, and the system comprises a circulating flue gas heating cylinder, an oxygen-rich combustor and a system control valve group; oxygen of the oxygen-rich combustor is mixed with fuel, after the fuel enters the circulating flue gas heating cylinder, the oxygen is mixed with part of the fuel to form root flame, and the unburned fuel and oxygen are completely combusted in the circulating flue gas heating cylinder to form flameless flame; the number of the circulating flue gas heating cylinders is one or more, each circulating flue gas heating cylinder is provided with an oxygen-rich combustor, heat generated by combustion is used for heating flue gas generated by combustion of oxygen and fuel in the circulating flue gas heating cylinders, the heated flue gas is circulated to the bell-type furnace through the high-temperature and high-pressure circulating flue gas fan, and redundant flue gas is recycled in a centralized mode. According to the system, heat is fully utilized through flue gas circulation, heating uniformity and quality are improved, combustion flue gas is high-purity carbon dioxide, and zero emission of the flue gas can be achieved through trapping and recycling.

A gas turbine engine for an aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include a first and second subset. The combustor is operable so each 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 is 1:3 to 1:6. A thrust nvPM emissions index ratio is EImaxTO / FmaxTO / EIidle / Fidle. EIidle is system loss corrected nvPM emissions index in mg / kg of the engine operating at around 7% available thrust. EImaxTO is system loss corrected nvPM emissions index in mg / kg of the engine at around 100% available thrust. FmaxTO is thrust of the engine at around 100% available thrust. Fidle is thrust at around 7% available thrust. The thrust nvPM emissions index ratio is between 0.0009 and 0.02.

A gas turbine engine for an aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include a first and second subset. The combustor is operable so each 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 the second subset is 1:3 to 1:6. A fuel-flow nvPM emissions index ratio isEIidle × Wf,idleEImaxTO × Wf,maxTO.EIidle is the system loss corrected nvPM emissions index in mg / kg of the engine at around 7% available thrust. EImaxTO is the index at around 100% available thrust. Wf,idle is the rate of fuel to the nozzles in kg / s at around 7% available thrust. Wf,maxTO is the rate of fuel to the nozzles in kg / s at around 100% available thrust. The index ratio is between 0.357 and 8.

The utility model discloses a kitchen range. The kitchen range comprises a burner and a pot rack, a plurality of first fire outlet holes are formed in the top of the combustor; the pot rack comprises a pot rack main body and a plurality of supporting legs, the pot rack main body is arranged around the periphery of the burner, the supporting legs are arranged at the bottom of the pot rack, and the supporting legs are arranged around the peripheral side of the pot rack main body at intervals; the inner circumferential side face of the pot rack body comprises a flow guide air inlet face, the flow guide air inlet face extends upwards from the bottom of the pot rack body, the flow guide air inlet face and the burner are arranged in a spaced mode, and a first air inlet channel communicated with the first fire outlet holes is formed between the flow guide air inlet face and the outer circumferential side face of the burner. An air inlet channel does not need to be formed in the pot rack body, so that the sealing integrity of the pot rack body can be guaranteed, and the heat preservation performance of the pot rack body can be improved.

The invention provides a CNN-PINN-based gas turbine combustion chamber performance prediction method, belongs to the technical field of combustion chamber combustion, and aims to solve the problems of high resource consumption and high grid dependence of the conventional CFD calculation at present, and the method comprises the steps: S1, collecting historical operation data of a combustion chamber, generating simulation data through numerical simulation, and carrying out the calculation of the simulation data; performing comparative analysis on the simulation data and the collected experimental operation data to prepare a data set; s2, preprocessing the data set; s3, establishing a CNN-PINN neural network prediction model according to the data characteristics and the target; s4, establishing a loss function containing a data error term and a physical information error term; s5, training and evaluating the established model by using the preprocessed data; and S6, performing performance prediction on target data by using the trained model, and performing reverse normalization on a prediction result to obtain prediction data corresponding to multiple targets.

The utility model discloses a flame-catalysis integrated combustor, which solves the problem that catalytic combustion and flame combustion are not integrated in the prior art, and has the beneficial effects of realizing integrated combustion and improving combustion efficiency, and the specific scheme is as follows: the flame-catalysis integrated combustor comprises a shell, one side of the shell is provided with an air inlet, and the other side of the shell is provided with an air outlet; a fuel-air mixer, a flame tube and a catalytic component are sequentially arranged in the shell from one side of the air inlet to the other side of the shell, the fuel-air mixer comprises a plurality of fuel distribution chambers which are sequentially arranged in a nested mode, and a gas circulation channel is formed between every two adjacent fuel distribution chambers; the fuel distribution chambers are provided with fuel spray holes facing the gas circulation channel, the outlet end of the fuel-air mixer is inserted into the flame tube, the flame tube comprises a shell, the shell and the shell are arranged at an interval, and the catalytic component is fixed to the end, away from the fuel-air mixer, of the shell of the flame tube.

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. A first idle-MTO nvPM emissions index ratio is defined as:EIidle / EImaxTO where: 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; and 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. The first idle-MTO nvPM emissions index ratio of the gas turbine engine is less than 0.8. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles. A method of operating the gas turbine engine is also disclosed.

The utility model discloses a hydrogen ammonia coupling low nitrogen burner relates to burner technical field, including first tank and second tank, be equipped with connecting subassembly between the first tank and the second tank, first tank far away from the one end of second tank is fixedly connected with the gas outlet pipe, the gas outlet pipe is communicated with first tank, and the gas outlet pipe is connected with the gas outlet pipe. A combustion assembly is arranged at one end of the gas outlet pipe away from the first tank; through the arrangement of the ammonia gas inlet pipe and the pipe connector, hydrogen and ammonia gas in a specific proportion can be injected into the cavity formed by the first tank body and the second tank body during working, the hydrogen and the ammonia gas are mixed through the premixing assembly and then sprayed out of the nozzle, the combustion temperature can be effectively reduced through the addition of the ammonia gas, NOx generation is reduced, meanwhile, the combustion speed of the ammonia gas is low, and the combustion efficiency is improved. And after the ammonia gas is mixed, the overall combustion speed can be reduced, the combustion stability is improved, so that the combustor has good combustion stability, and meanwhile, the emission of NOx is greatly reduced.

The invention discloses a combustor and gas equipment, and relates to the technical field of gas equipment. Wherein the burner comprises a fire grate and an injection device, and the fire grate is provided with an air inlet; the injection device is installed at an air inlet of the fire grate and provided with an injection channel and a fuel gas channel arranged outside the injection channel, the injection channel is provided with an air inlet end, an air outlet end and a fuel gas introduction port, the air inlet end of the injection channel is used for inputting air, the fuel gas introduction port is communicated with the fuel gas channel, the air outlet end of the injection channel is communicated with the air inlet, and the fuel gas introduction port is communicated with the fuel gas channel. The gas-air mixer is used for conveying gas-air mixture into the fire grate. According to the technical scheme, fuel gas and air can be mixed more sufficiently, sufficient combustion can be achieved, and emission of nitric oxide is reduced.