3700results about "Continuous combustion chamber" 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 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 belongs to the field of aero-engines, relates to an engine main combustion chamber design technology, and provides a ceramic matrix composite material flame tube matching design method and system.The method comprises the steps that a flame tube wall temperature interval of a main working state point is obtained; determining the thermal-state radial dimensions of the feature points on the ceramic-based component and the metal component; converting the hot-state radial size into a cold-state radial size; carrying out two-dimensional cold-state flame tube profile design; and the cold-state radial size of the two-dimensional cold-state flame tube molded surface is converted into the hot-state size to obtain a two-dimensional hot-state flame tube molded surface, and a three-dimensional hot-state model is constructed and evaluated and iteratively optimized. According to the method, matching structure characteristic parameters of the ceramic-based component and the metal component are designed, and the molded surface of the flame tube made of the different materials is self-adaptively formed into smooth transition at main working state points, so that a backflow area is actively controlled, and the problems that the stability of a flow field of a main combustion chamber is influenced and pressure is lost due to different and relative displacement changes in the working process are solved.

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 combustor for a gas turbine (GT) system includes a combustor body including a non-additively manufactured (non-AM) part having corrosion at a first level and a replacement additively manufactured (AM) part coupled to the non-AM part. The replacement AM part has corrosion at a second level less than the first level. The replacement AM part includes at least an aft section of a tapered transition portion and an aft frame of the combustor body, and a plurality of shared parallel, sintered metal layers extending into the aft section of the tapered transition portion and the aft frame. The replacement AM part may also include a receiving element in a forwardmost end thereof configured to receive a rearwardmost remaining end of the non-AM part. The replacement AM part may also optionally include an AFS injector mount and part of a cylindrical portion of a combustion liner.

A method is provided for operating a powerplant. During this method, a hydrocarbon fuel is delivered to a turbine engine for combustion within a combustion chamber of the turbine engine during initial startup operation of the turbine engine. Ammonia is at least partially cracked into hydrogen gas and nitrogen gas. A non-hydrocarbon fuel is delivered to the turbine engine for combustion within the combustion chamber of the turbine engine during post-startup operation of the turbine engine. The non-hydrocarbon fuel is or otherwise includes the hydrogen gas or a combination of the hydrogen gas and the nitrogen gas.

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 belongs to the field of engines, and particularly relates to a dual-fuel regenerative cooling system and method. The cooling system comprises a combustion chamber, an ammonia supply device and a kerosene supply device; three layers of cooling flow channels are arranged in the wall face of the combustion chamber, the output ends of the three layers of cooling flow channels are arranged in the combustion chamber, and the three layers of cooling flow channels comprise a first cooling channel, a second cooling channel and a third cooling channel which are sequentially arranged from inside to outside. The ammonia supply device is connected with the input ends of the first cooling channel and the second cooling channel; the kerosene supply device is connected with the input ends of the third cooling channel and the second cooling channel; and ammonia or kerosene is introduced into the second cooling channel in a time-sharing manner. Through the three-layer cooling channel design and the fuel switchable function, the effects of fuel cooling, thermal protection, cracking hydrogen production and combustion organization are improved at the same time.

The application provides an integrated flame stabilizer with a tail edge cavity, which comprises a flame stabilizer, an oil injection rod, a cooling gas inlet pipe and a tail edge wing plate arranged at the rear side of the flame stabilizer, and a stable backflow ignition area is formed in the downstream area of the flame stabilizer; a plurality of cooling channels are arranged on the rear side plate of the flame stabilizer between the two tail edge wing plates, and the airflow direction at the outlet of the cooling channels is inclined to the flow direction, so that a cooling vortex is formed in the tail edge cavity, the flame stabilizer and the backflow ignition area are spaced from each other, the flame stabilizer and the downstream backflow ignition area are isolated from each other to reduce the thermal load of the components such as the flame stabilizer and the fuel injection pipe, the ablation and coking problems are avoided, meanwhile, the flame stabilizing function of the flame stabilizer is not damaged, and the influence on the backflow ignition area behind the stabilizer is small.

A multi-fuel combustion system for a turbine engine, the multi-fuel combustion system having a combustion chamber (44, 444, 544) formed by a combustor liner (34, 434, 534). The combustion chamber (44, 444, 544) defines a first combustion zone (50, 450, 550) and a second combustion zone (52, 452, 552). A first fuel system (60, 460, 560) is fluidly coupled with the first combustion zone (50, 450, 550), where a rotary fuel slinger (62, 462, 562) provides a first fuel (58, 458, 558) to the first combustion zone (50, 450, 550). A second fuel system (80, 480, 580) is fluidly coupled with the second combustion zone (52, 452, 552), where a gaseous fuel injector (82, 482, 582) provides a second fuel (74, 474, 574) to the second combustion zone (52, 452, 552).

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

The invention relates to the technical field of oil supply assemblies in afterburner components of aero-engines, in particular to a multi-channel distributed series finger-shaped oil supply header pipe and an afterburner. The multi-channel distributed series finger-shaped oil supply header pipe comprises a body, and a plurality of fuel oil channels are formed in the body; the upper end of the fuel oil channel penetrates through the top surface of the body; each fuel oil channel is correspondingly provided with a plurality of main fuel oil injection holes, the main fuel oil injection holes are communicated with the fuel oil channels, and the axes of the main fuel oil injection holes are perpendicular to the central axis of the body; the plurality of main fuel injection holes in each fuel channel are arranged at intervals from top to bottom; a plurality of bypass oil injection holes are formed in the position close to the main oil injection hole, and the bypass oil injection holes are communicated with the main oil injection hole through connecting holes; the main oil injection hole, the bypass oil injection hole and the connecting hole are all formed in the body.

The invention provides a regenerative cooling channel structure and an engine, and relates to the technical field of engine thermal protection, the regenerative cooling channel structure comprises a first channel and a second channel which are arranged between the outer wall face and the inner wall face of a combustion chamber at intervals in the radial direction, and an inlet and an outlet of the first channel are located at the inlet end and the outlet end of the combustion chamber correspondingly; an inlet and an outlet of the second channel are located at the outlet end and the inlet end of the combustion chamber respectively. An outlet of the first channel communicates with an inlet of the second channel, and a coolant flows into the first channel through the inlet of the first channel, then flows into the second channel through an outlet of the first channel and the inlet of the second channel in sequence and then flows into the combustion chamber through an outlet of the second channel. The flowing direction of the coolant in the first channel is consistent with the flowing direction of gas in the combustion chamber to form downstream cooling, the flowing direction of the coolant in the second channel is opposite to the flowing direction of the gas in the combustion chamber to form countercurrent cooling, the flowing speed and flow distribution of the coolant are improved, the heat transfer deterioration phenomenon is restrained, and the cooling effect is effectively improved.

The invention discloses an expanded air film cooling hole and a gas turbine combustion chamber wall surface, relates to the field of gas turbine combustion chambers, and aims to solve the problem of poor cooling effect of an existing air film hole. The expanding type air film cooling hole sequentially comprises a cylindrical section and an expanding section in the flowing direction of cooling air flow, the cylindrical section and the expanding section are coaxially arranged, the expanding section is gradually expanded in the flowing direction of the cooling air flow, the expanding direction is the spanwise direction, an outlet of the expanding section is in a trapezoid shape, and the four corners of the trapezoid are subjected to rounding treatment. And an outlet fillet of the expanding type air film cooling hole is formed. The expanding type air film cooling holes are formed in the wall face of the combustion chamber of the gas turbine based on the expanding type air film cooling holes, every two adjacent rows of expanding type air film cooling holes are arranged in a staggered mode, and the central axes of the expanding type air film cooling holes are parallel to each other. Generation of kidney-shaped vortexes can be weakened, the air film cooling efficiency is improved, and the cooling effect is improved. The invention is applicable to gas turbine combustion chambers.

An assembly is provided for a turbine engine. This assembly includes an engine core extending axially along an axis. The engine core includes a compressor section, a combustor, a diffuser structure, a diffuser plenum and a plurality of separators. The combustor is arranged within the diffuser plenum. The combustor includes a combustion chamber and a combustor wall between the combustion chamber and the diffuser plenum. The diffuser structure includes a plurality of diffuser passages. Each of the diffuser passages fluidly couples the compressor section to a respective one of the separators. Each of the separators includes a first outlet into the diffuser plenum and a second outlet into the combustion chamber.

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.

According to the dual-fuel afterburner without the inner duct stabilizer, an oil injection rod and an air injection rod are arranged in a supporting plate with a cavity structure, an oil injection nozzle and an air injection nozzle which are communicated are formed in the wall face of the supporting plate, and cooperative combustion of gas fuel and liquid fuel is achieved. According to the afterburner, a stable ignition source is provided through the characteristic that gas fuel is easy to burn, efficient and stable combustion is achieved in cooperation with fuel injection, and therefore a traditional inner duct stabilizer is omitted, the weight of the afterburner is remarkably reduced, and weight reduction and performance optimization of the afterburner are achieved. Evaporative flame stabilizers distributed in the circumferential direction are arranged in the outer duct area to achieve stable combustion of airflow in the outer duct, and flame stability and cross flame performance are guaranteed in cooperation with the supporting plates of the inner duct and fuel gas sprayed out of the supporting plates.

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 invention relates to an injector arrangement (1) for a gas turbine, in particular an aircraft engine, for introducing a gaseous fuel and a liquid fuel as well as air into a combustion chamber (CC), comprising an injector shaft (2) and an injector main body (3) aligned along an injector longitudinal axis (L), wherein the injector main body (3) comprises: - a first gas channel (12) arranged centrally on the injector longitudinal axis (L) with a downstream outlet opening (14) for introducing a gas flow, - at least one air channel (31, 36, 40) arranged radially around the outside of the first gas channel (12), as a second gas channel (30), and - a liquid fuel injection (24) arranged radially around the first gas channel (12) for introducing the liquid fuel into the combustion chamber (CC).An advantageous emission characteristic can be achieved by designing the first gas channel (12) exclusively for introducing the gaseous fuel into the combustion chamber (CC) (Fig. 1).

A gas turbine engine (10), comprising a compressor section (12), combustion section (14), and turbine section (16) in a serial flow arrangement, with the combustion section (14) comprising: a combustor liner (40) that at least partially defines a combustion chamber (50); and a gaseous fuel nozzle assembly (48), comprising: a rich fuel supply (102) to supply a rich mixture (106) of gaseous fuel (F) and air (70); a lean fuel supply (104) to supply a lean mixture (108) of gaseous fuel (F) and air (70) that is leaner than the rich mixture (106); a rich impingement tube (110) fluidly coupled to the rich fuel supply (102) to emit the rich mixture (106) into the combustion chamber (50); and a lean impingement tube (112) fluidly coupled to the lean fuel supply (104) to emit the lean mixture (108) into the combustion chamber (50).

The proposed solution relates to a nozzle assembly for a combustor of an engine, having at least one nozzle for injecting hydrogen into a combustion chamber of the combustor, wherein the nozzle has a nozzle main body which extends along a nozzle longitudinal axis and a nozzle head at one end of the nozzle main body and comprises a first air conduit, a fuel duct for the hydrogen to be injected, and a second air conduit. The fuel duct is arranged with a fuel exit opening for the hydrogen to be injected into the combustion chamber between the first and second air conduits with their first and second air exit openings, in relation to a radial direction running perpendicularly to the nozzle longitudinal axis. Also according to the invention, at least the second air conduit, which is situated radially further outwards, is designed and provided to provide an non-swirled air flow into the combustion chamber and the fuel duct is designed and provided to provide a non-swirled flow of hydrogen into the combustion chamber.

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.

The invention relates to a two-stage combustor and an axial staged combustion chamber. The secondary combustor comprises a fuel main pipe, a fuel branch pipe, an inner cylinder and an outer cylinder; an inner airflow channel is formed between the inner cylinder and the fuel main pipe; an inner jet hole is formed in one end, close to the outlet of the inner cylinder, of the side wall of the fuel main pipe; an outer airflow channel is formed between the outer cylinder and the inner cylinder, and the outer cylinder comprises a main cylinder section and a conical cylinder section; the conical barrel section sleeves the inlet end of the inner barrel, and at least part of the conical barrel section extends into the main barrel section; the inner diameter of the conical barrel section is gradually reduced from the first end to the second end; the axial direction of the fuel branch pipe is arranged along the radial direction of the fuel main pipe, one end of the fuel branch pipe is connected to the side wall of the fuel main pipe, and the other end penetrates through the side wall of the inner cylinder and extends into the conical cylinder section; an outer jet hole is formed in the end, stretching into the conical barrel section, of the side wall of the fuel branch pipe. The secondary combustor can effectively isolate direct coupling of secondary fuel and primary flame, and avoids unstable combustion and over-high temperature caused by combustion superposition in a coupling area.