2266results 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 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.

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.

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.

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 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, comprising a compressor section, combustion section, and turbine section in serial flow arrangement, with the combustion section comprising: a combustor liner at least partially defining a combustion chamber; a wall coupled to the combustor liner; a first fuel supply to supply a first fuel; a gaseous fuel supply to supply a gaseous hydrogen fuel; and a fuel nozzle assembly coupled to the wall and fluidly coupled to the first fuel supply and the gaseous hydrogen fuel supply, the fuel nozzle assembly comprising: a main mixer; and a fuel nozzle comprising an outer wall defining a pilot channel and an outer wall fuel orifice to emit at least one of the first fuel or the gaseous fuel radially outward into the main mixer; and a secondary mixer disposed in the pilot channel.

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.

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 system for a hydrogen fuelled aircraft propulsion system comprises a fuel tank configured to store hydrogen fuel, a main fuel conduit configured to provide fuel to a combustor of a gas turbine engine, a fuel heater comprising a catalytic combustor configured to catalytically combust a portion of the hydrogen fuel prior to delivery to the combustor and a heat exchanger configured to exchange heat between exhaust gases from the fuel heater and hydrogen fuel in the fuel conduit.

The application relates to a fixing mechanism of a combustion chamber tile and a combustion chamber heat shield device, wherein the circumferential side edge of the tile is provided with a stepped installation slot, the stepped surface of which constitutes an anchoring surface; the shell is provided with a shell ventilation hole and a containing through hole; the fixing mechanism comprises a pressing disc and a connecting bolt, the pressing disc comprises a middle connecting part, a limiting side plate and a pressing plate, the limiting side plate is located outside the middle connecting part, the lower end of the limiting side plate is connected with the lower end of the middle connecting part, and a cooling cavity is arranged between the two, the pressing plate is connected with the upper end of the limiting side plate and extends outward, the lower end of the middle connecting part is provided with a cooling groove and an abutting surface, a disc-in-plate ventilation hole and a connecting through hole are arranged in the middle connecting part and are connected with the cooling groove and the cooling cavity, one end of the connecting bolt is provided with a pressing head, and the other end is a fixed end; during installation, the limiting side plate of the pressing disc is located in the lower slot part, the connecting bolt passes through the connecting through hole and the containing through hole, the middle connecting part is pressed on the shell by the pressing head, the cooling groove is aligned with the shell ventilation hole, and the pressing plate is pressed on the anchoring surface.

There is provided a flameholder 370 for a reheat assembly 300, 300A of a gas turbine engine 10. The flameholder comprises an internal flow passageway 376 extending from an inlet aperture 374 and defining a flow direction F for flow through the flameholder 370. The internal flow passageway 376 is defined by an internal surface 34 of the flameholder 370. A grid of recesses 31 is formed in the internal surface 34 of the flameholder 370.

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.

A cooling flow arrangement for a ceramic matrix composite (CMC) component of a turbine engine, wherein the CMC component includes a hot side configured for exposure to a hot gas path of the turbine engine and an opposing cold side. At least one cavity is provided in the CMC component, the at least one cavity having an entrance disposed on the cold side that is configured for receiving a cooling flow. A seal is disposed adjacent the at least one cavity and the cold side of the CMC component and is configured to provide a seal between the CMC component and an upstream turbine component. At least one hole extends through the seal and is positioned to provide the cooling flow to the at least one cavity of the CMC component.

A gas turbine engine, comprising: a compressor section, a combustion section, and a turbine section in a serial flow arrangement, with the combustion section comprising: a combustor liner that at least partially defines a combustion chamber; and a fuel nozzle assembly, comprising: a liquid fuel supply to supply a liquid fuel; a hydrogen fuel supply to supply a gaseous hydrogen fuel; and a fuel nozzle body defining a fuel nozzle assembly centerline, the fuel nozzle body fluidly coupled with the liquid fuel supply and the hydrogen fuel supply to provide the liquid fuel and the gaseous hydrogen fuel to the combustion chamber.

A fuel nozzle for a turbine engine. The fuel nozzle has a premixer body, vortex generator, air injection orifice, and fuel injection orifice. The premixer body defines a primary flow path. The air injection orifice is provided in the premixer body and located downstream of the vortex generator. The fuel injection orifice is provided in the premixer body and opens into the primary flow path.

The invention discloses a micro-mixing nozzle integrating heat pipe cooling and heat regeneration and a combustor, the micro-mixing nozzle comprises a heat pipe assembly I, a micro-mixing sleeve, an air mixing module and a fuel pipe, the micro-mixing sleeve, the air mixing module and the fuel pipe are coaxially arranged, and the heat pipe assembly I comprises a heat conduction ring, perforated fins, a heat pipe and heat pipe radiating fins; the overall combustor comprises a fuel cavity, an air cavity and a cooling cavity which are sequentially, coaxially and fixedly arranged at intervals from bottom to top, and a plurality of micro-mixing nozzles which are symmetrically arranged relative to the center of the overall geometric central axis are installed in the air cavity and the cooling cavity. According to the micro-mixing nozzle integrating cooling and backheating of the heat pipe and the combustor, the inlet area of the combustion chamber can be rapidly cooled, the service life of the combustor is prolonged, meanwhile, heat of the inlet area of the combustion chamber can be effectively utilized to preheat air, and the combustion efficiency is improved. The problem that the service life of an existing combustor is short due to thermal damage of a high-temperature backflow area to a micro-mixing nozzle can be solved.

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.

A fuel injector (10) includes a mount (14), a stem (16), a nozzle (18), and a heat shield (12). The stem (16) extends from the mount (14). The nozzle (18) is at a distal end of the stem (16) opposite the mount (14) and configured to discharge an oxidant-fuel mixture along a nozzle axis (C). The heat shield (12) extends from the mount (14) towards the nozzle (18). The heat shield (12) includes a stem segment (12A) surrounding the stem (16) and at least a portion of the nozzle (18) and a compliant segment (12B) extending from the stem segment (12A) to the nozzle (18), a lip (26) extending from the nozzle (18), or a cap (28) surrounding the nozzle (18).

A gas turbine engine including a fuel injector having an outer valve body with an inner passage, a wall at least partially enclosing the inner passage, the wall having an inner surface, an outer surface, and one or more channels passing through the wall between the outer surface of the wall and the inner surface of the wall, and a retractable valve body positioned within the inner passage of the outer valve body. The retractable valve body has one or more slots in fluid communication with the fluid spin chamber. The retractable valve body is positioned within the inner passage to move along a reciprocating axis at one or more valve positions including an advanced valve position, a retracted valve position, an open valve position, and a closed valve position.