247 results about "Fuel spray" 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.

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.

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 a low-temperature liquefied natural gas spray cooling performance test system and a test method, belongs to the technical field of low-temperature fuel spray cooling, and solves the problem that a test system and a test method for taking liquefied natural gas fuel as a spray cooling material are lacked in the prior art. And the performance of the liquefied natural gas fuel as the spray cooling material cannot be accurately analyzed. The test system comprises a double-layer structure spraying chamber, a nozzle, a heat source and a data acquisition system, lNG mist sprayed by the nozzle flows in the inner space of the double-layer structure mist spraying chamber; liquid nitrogen flows in the interlayer space of the double-layer structure spraying chamber and is used for low-temperature cooling of LNG spraying. According to the test system, a double-layer heat preservation structure is adopted, and the LNG partially gasified due to storage and transportation can be cooled by using low-temperature liquid nitrogen, so that the phase change heat absorption capacity of the LNG is exerted in spray cooling. And meanwhile, the influence of the external environment on the test result can be well isolated, and the external interference is reduced.

The invention relates to the technical field of coating preparation, in particular to a spraying process of a compact yttrium oxide coating, which comprises the steps of pretreatment, powder activation, matrix surface modification, meltallizing spraying, cooling curing, de-shielding treatment and post-cleaning. Yttrium oxide powder is activated through low-temperature plasma, sand blasting parameters and a high velocity oxy-fuel spraying process are optimized, the porosity of the coating is remarkably reduced to 1.5% or below, and the hardness is improved to 500 HV0.3. Internal stress is reduced in a sectional cooling mode, and balance of coating performance and economical efficiency is achieved in combination with a parameter optimization algorithm based on particle deposition dynamics. The preparation method can meet the requirement of the high-end semiconductor manufacturing field on a high-performance protective coating, and solves the problem of high porosity of a traditional process.

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.

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.

A gas turbine engine has a first subset of fuel spray nozzles and a second subset of fuel spray nozzles. A combustor is operable with the first subset of fuel spray nozzles supplied with fuel at a greater fuel flow rate than each the second subset of fuel spray nozzles. A a ratio of the first subset of fuel spray nozzles to the second subset of fuel spray nozzles is 1:2 to 1:5. A first idle-MTO nvPM emissions index ratio is:EIidleEImaxTOwhere: EIidle is the system loss corrected nvPM emissions index in mg / kg operating at around 7% available thrust; and EImaxTO is the system loss corrected nvPM emissions index in mg / kg operating at around 100% available thrust; and the first idle-MTO nvPM emissions index ratio of the gas turbine engine is less than 60. A sustainable aviation fuel (SAF) can be provided to the fuel spray nozzles.

A gas turbine engine has a first subset of fuel spray nozzles and a second subset of fuel spray nozzles. A combustor is operable with the first subset of fuel spray nozzles supplied with fuel at a greater flow rate than each of the second subset of fuel spray nozzles. A ratio of the first subset of fuel spray nozzles to the second subset of fuel spray nozzles is 1:2 to 1:5. A first idle-MTO nvPM emissions index ratio is:EIidleEImaxTO,where: Elidle is the nvPM emissions index in mg / kg operating at around 7% available thrust; ElmaxTO is the nvPM emissions index in mg / kg of the gas turbine engine operating at around 100% available thrust; and the first idle-MTO nvPM emissions index ratio of the gas turbine engine is less than 60. A sustainable aviation fuel (SAF) can be provided to the fuel spray nozzles.

Disclosed are an apparatus for a simulation test of diesel engine fuel spray in an alpine environment and a test method, the apparatus comprising a fuel storage and pretreatment device, a fuel injection system, an optical visualization system, an environment chamber, and a spray pretreatment system. In alpine environmental conditions, the apparatus can achieve high-pressure injection of liquid, separate injected atomized mist into fine droplets and impinge same on a wall surface by means of a multi-droplet generation apparatus and a wall surface apparatus, and observe and record the whole process by means of an optical visualization path. According to the present invention, the problem of difficulty in distinguishing droplets due to the coalescence of spray into a sheet is avoided, the costs of a traditional spray and droplet test solution can be significantly reduced, and the problem of a high level of dependence on a control program in a constant volume vessel solution is solved. According to the present invention, simulation of spray test in an alpine environment is possible, and the test apparatus has a simple structure. Moreover, it is possible to conduct regular research on free spray, multi-droplet wall impingement, spray wall impingement, and development characteristics thereof under conditions such as different altitudes and different temperatures.

The invention discloses a (NiCr-Cr3C2) / NiCr composite coating and a preparation method thereof, and belongs to the technical field of composite material preparation.The preparation method of the (NiCr-Cr3C2) / NiCr composite coating comprises the following steps that an HVOF technology (high-speed oxygen fuel spraying technology) is adopted, NiCr powder and NiCr-Cr3C2 powder are sequentially sprayed on the surface of a base material, and a NiCr transition layer and a NiCr-Cr3C2 working layer are formed respectively; and then annealing treatment is carried out, and the (NiCr-Cr3C2) / NiCr composite coating is obtained. Through the annealing treatment, especially the annealing at 700 DEG C, a nanoscale secondary carbide and polycrystalline ring net structure can be formed, so that the coating obtains the optimal comprehensive performance, namely the highest coating section microhardness (1242HV0.2), the optimal toughness (7.9 MPa.m < 1 / 2 >) and the lowest wear rate (0.37 * 10 <-5 > mm < 3 > / N.m).

A gas turbine engine for aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The fuel spray nozzles include a first subset and a second subset of nozzles. Each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of the first subset to the second subset is 1:2 to 1:5. A MTO nvPM emissions index ratio isEImaxTO,SAFEImaxTO,FF.EImaxTO,SAF is nvPM emissions index in mg / kg of the engine when operating at around 100% available thrust if fuel provided to the fuel spray nozzles includes sustainable aviation fuel. EImaxTO,FF IS nvPM emissions index in mg / kg of the engine when operating at around 100% available thrust if fuel provided to the fuel spray nozzles is fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio is less than 1.

A fuel spray nozzle includes a fuel circuit having in series a gallery, circumferentially spaced passages arranged in a row around the nozzle, and an annular spin chamber. Each passage has an inlet for receiving a respective portion of the fuel from the gallery and a metering orifice for discharging its portion of the fuel. The passages are configured such that, when the flow of liquid fuel to the inlet port is shut off, a respective differential static pressure develops across stagnant liquid fuel remaining between the inlet and the metering orifice of each passage, and the passages are further configured such that one or more selected passages develop a different differential static pressure to the remaining passages causing a flow of purging air to enter the gallery from the combustor through the selected passages and exit through the remaining passages, thereby purging the gallery and the passages of fuel.

The invention discloses a combustion chamber fuel injection and oil-gas mixing system suitable for a low-temperature environment, and belongs to the field of combustion chambers of gas turbines, the combustion chamber fuel injection and oil-gas mixing system comprises a flame tube, a centrifugal nozzle and a two-stage cyclone which are coaxially arranged, after fuel is atomized through the centrifugal nozzle and enters, the fuel is further atomized through rotational flow of an inner ring of the two-stage cyclone and enters the flame tube, and the fuel is injected into the combustion chamber through the centrifugal nozzle. External high-pressure air is swirled through the outer ring of the two-stage swirler to form swirling air, and internal fuel spray is crushed, atomized and mixed; a venturi tube is arranged at the end, close to the flame tube, of the inner ring of the two-stage cyclone, vortex generators are evenly arranged on the inner wall of the expansion section of the venturi tube in the circumferential direction, and oil mist with the small and even particle size is formed after the atomized mist is acted. Through multi-stage rotational flow, oil mist with fine and uniform granularity is formed, the atomization effect is optimized, and the fuel oil atomization quality and combustion stability in a low-temperature environment are improved.

The invention relates to the technical field of hypersonic flame spraying control, in particular to a hypersonic flame spraying closed-loop control system, which comprises a hypersonic flame particle parameter acquisition module for acquiring a particle rate sequence and a flame flow temperature sequence of each time period in a hypersonic flame spraying process; the hypersonic flame particle parameter analysis module is used for obtaining flame flow fluctuation quantity of each extreme point in the particle rate sequence corresponding to each time period; the flame velocity temperature influence quantity of the supersonic flame in each time period is obtained; obtaining a filtering parameter closed-loop regulating variable of high velocity oxy-fuel spraying control; and the hypersonic flame spraying control module is used for adjusting filtering parameters in IMC control of each time period according to the difference between the filtering parameter closed-loop adjusting quantity of each time period and the filtering parameter closed-loop adjusting quantity of the adjacent previous time period. The control quality and efficiency of high velocity oxy-fuel spraying are improved, and the spraying process is more accurate and stable.

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 fuel-flow nvPM emissions index ratio is defined as:E⁢Iidle×Wf,idleE⁢ImaxTO×Wf,maxTOwhere: EIidle is the 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 nvPM emissions index in mg / kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions; Wf,idle is the rate of fuel flow to the fuel spray nozzles in kg / s at around 7% available thrust for the given operating conditions; and Wf,maxTO is the rate of fuel flow to the fuel spray nozzles in kg / s at around 100% available thrust for the given operating conditions. The fuel-flow nvPM emissions index ratio of the gas turbine engine is less than 0.3. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles. Also disclosed are methods of operating a gas turbine engine.

This invention discloses a high-disturbance spray boom structure based on plasma-excited coupled aerodynamically assisted atomization, relating to the field of afterburner technology for aero-engines. It includes: a shell with an air inlet; a fuel pipe with a fuel delivery channel; the fuel pipe extending into the shell and forming an air channel with it; a first insulating platform with a collision atomization zone and an atomization development zone; the first insulating platform is connected to the output end of the fuel pipe, and its neck has several fuel-gas atomization channels; a second insulating platform with a plasma discharge region; the second insulating platform is snapped into the first insulating platform; a high-voltage electrode located between the first and second insulating platforms; and a grounding electrode located between the second insulating platform and the shell. This invention employs the coupling of plasma excitation with aerodynamically assisted atomization and cross-collision to perform three consecutive actions on the fuel spray, thereby improving the fuel atomization performance of the high-disturbance spray boom.

The invention discloses an iron-based amorphous alloy suitable for a marine environment and a method for preparing a coating based on the iron-based amorphous alloy. The chemical formula of the iron-based amorphous alloy is Fe < 35 > Ni < 20-x > Cr < 20 > Mo < 5 > B < 4 > C < 4 > P < 12 > Ce < x >, wherein the value of x ranges from 0.5 to 1.5. On the basis of the iron-based amorphous alloy, a coating formed by adopting a high velocity oxy-fuel spraying mode not only has good seawater corrosion resistance, but also has good microbial corrosion resistance; after the coating is soaked in a culture medium solution inoculated with pseudomonas aeruginosa for 14 days, the self-corrosion potential is-314 mV, the self-corrosion current density is 5.9 [mu] A cm <-2 >, the passivation interval exceeds 1200 mV, and the passivation current density is 5.1 [mu] A cm <-2 >.

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 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. A fuel-flow nvPM emissions index ratio is defined as:EIidle×Wf,idleEImax⁢TO×Wf,maxTOwhere: 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; Wf,idle is the rate of fuel flow to the fuel spray nozzles in kg / s at around 7% available thrust for the given operating conditions; and Wf,maxTO is the rate of fuel flow to the fuel spray nozzles in kg / s at around 100% available thrust for the given operating conditions. The fuel-flow nvPM emissions index ratio of the gas turbine engine is less than 0.08. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.

The embodiment of the invention provides a preparation method of an anti-erosion gradient coating. The preparation method comprises the steps that a to-be-repaired area on the surface of a workpiece is subjected to cleaning and sand blasting treatment; the first powder, the second powder and the third powder are sequentially sprayed to the to-be-repaired area through the high velocity oxy-fuel spraying technology, and the gradient coating is prepared; the first powder, the second powder and the third powder all comprise hard alloy powder and base material powder made of the same material as the workpiece, the content of the hard alloy powder in the first powder is smaller than that of the hard alloy powder in the second powder, and the content of the hard alloy powder in the second powder is smaller than that of the hard alloy powder in the third powder; and after the workpiece is sprayed and naturally cooled, the surface of the sprayed workpiece is subjected to protection treatment. According to the method, the anti-erosion coating (gradient coating) can be prepared on the basis of hypersonic flame spraying. And under the action of the gradient coating, a compact coating with smaller tensile stress can be obtained, and the properties of erosion resistance, water erosion resistance and the like of the workpiece are improved.

The invention relates to the technical field of fuel injection, and provides a turbulent burner. The turbulent burner comprises an outer wall ring, a central bluff body, a plurality of rotational flow blades, a turbulent flow bluff body and an anti-vibration pile; a central channel is formed in the middle of the outer wall ring; the central bluff body is located in the central channel, a first fuel channel and a second fuel channel are formed in the central bluff body, and a first fuel spray hole communicated with the first fuel channel and a second fuel spray hole communicated with the second fuel channel are formed in the downstream of the central bluff body; the rotational flow blades are connected with the central bluff body and the outer wall ring; the turbulent flow bluff bodies are circumferentially distributed along the downstream of the central bluff body and are positioned on the upstream of the first fuel spray holes; a rotational flow channel is formed among the outer wall ring, the central bluff body, the turbulent flow bluff body and the rotational flow blades, air flows through the rotational flow channel to form air rotational flow, and the first fuel spraying holes and the second fuel spraying holes are used for spraying fuel and mixed with the air rotational flow for combustion in a transverse jet flow mode. The anti-vibration pile is connected with the outer wall ring, located at an outlet of the turbulent burner and arranged in the circumferential direction of the outer wall ring.

A gas turbine engine has a first subset of fuel spray nozzles and a second subset of fuel spray nozzles. A combustor is operable with the first subset of fuel spray nozzles supplied with fuel at a greater flow rate than each of the second subset of fuel spray nozzles. A ratio of the first subset of fuel spray nozzles to the second subset of fuel spray nozzles is 1:2 to 1:5. A first idle-MTO nvPM emissions index ratio is:EIidleEImax⁢TO,where: EIidle is the nvPM emissions index in mg / kg operating at around 7% available thrust; EImaxTO is the nvPM emissions index in mg / kg of the gas turbine engine operating at around 100% available thrust; and the first idle-MTO nvPM emissions index ratio of the gas turbine engine is less than 60. A sustainable aviation fuel (SAF) can be provided to the fuel spray nozzles.

A gas turbine engine has a first subset of fuel spray nozzles and a second subset of fuel spray nozzles. A combustor is operable with the first subset of fuel spray nozzles supplied with fuel at a greater fuel flow rate than each the second subset of fuel spray nozzles. A a ratio of the first subset of fuel spray nozzles to the second subset of fuel spray nozzles is 1:2 to 1:5. A first idle-MTO nvPM emissions index ratio is:EIidleEImaxTOwhere: EIidle is the system loss corrected nvPM emissions index in mg / kg operating at around 7% available thrust; and EImaxTO is the system loss corrected nvPM emissions index in mg / kg operating at around 100% available thrust; and the first idle-MTO nvPM emissions index ratio of the gas turbine engine is less than 60. A sustainable aviation fuel (SAF) can be provided to the fuel spray nozzles.