7207results about "Jet propulsion plants" patented technology

The invention belongs to the field of data processing, and particularly relates to a dynamic cooperative control system and method for a gas turbine and a micro-grid, and the method comprises the steps: constructing a micro-grid real-time monitoring module, continuously collecting distributed energy real-time output, controllable load demands, bus voltage frequency and equipment state parameters, and carrying out the filtering and noise reduction through a preprocessing unit, thereby guaranteeing the data precision; calculating a real-time power difference value based on the preprocessed data, calling an adaptive neural fuzzy inference system, taking the power difference value, the bus voltage deviation and the frequency deviation as input, and judging whether the power difference value, the bus voltage deviation and the frequency deviation exceed a preset threshold value by means of a fuzzy rule base and a neural network model; if the threshold values are not exceeded, the current states of the gas turbine and the energy storage system are maintained; if any one exceeds the threshold value, a dynamic cooperative control instruction is triggered, precise cooperative control of the gas turbine and the micro-grid is achieved, and the operation stability, the operation efficiency and the reliability of the micro-grid are improved.

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.

The invention relates to the technical field of energy, in particular to a dynamic self-adaptive control method and system for a gas-steam combined cycle unit, and the method comprises the steps: constructing a multi-modal model library comprising a plurality of operation modes; operating parameters and environment parameters of the unit are collected in real time, and multi-source data fusion and system state evaluation are carried out; carrying out multi-time scale prediction on the load demand, fuel characteristics and environmental parameters of the unit based on a deep learning model; dynamically selecting model precision according to a prediction result, and implementing a multi-mode collaborative optimization decision to generate a control instruction; and executing the control instruction and monitoring an execution effect, and carrying out adaptive adjustment on model parameters and strategies according to feedback. According to the method, the unit can maintain efficient and stable operation under the scenes of power grid peak regulation, rapid load fluctuation, complex environment change and the like, meanwhile, pollutant emission is reduced, and the service life of key equipment is prolonged.

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 heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.

A gas turbine engine is provided. The gas turbine engine includes: a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches; wherein the gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (FnTotal) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: FnTotal×EGT / (AHPCExit2×1000).

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.

A thermal management system defines a thermal management system flowpath to provide a flow of a fluid to an electric machine and a power electronics assembly electrically connected to the electric machine. The thermal management system includes a first heat exchanger thermally connected to the thermal management system flowpath and to the electric machine, and a second heat exchanger thermally connected to the thermal management system flowpath downstream of the first heat exchanger. The second heat exchanger is thermally connected to the power electronics assembly.

A gas turbine engine includes a turbomachine comprising compressor, combustion, and turbine sections. The gas turbine engine defines a maximum exhaust gas temperature, a maximum drive turbine shaft torque, and a corrected specific power. The gas turbine engine includes a controller configured to autonomously regulate performance of the gas turbine engine in response to at least one of: a thrust demand, an energy efficiency target, or a flight profile condition.

An integrated hydrogen-electric engine includes, an air compressor system; a hydrogen fuel source; a fuel cell; an elongated shaft connected to the air compressor system or a propulsor; and a motor assembly disposed in electrical communication with the fuel cell, wherein the air compressor system includes a plurality of electrically driven compressors configured to run in series or parallel.

An air / oil heat exchanger arranged in a vein. The exchanger includes internal parietal corridors and / or external parietal corridors and central corridors, the radial height of the internal and / or external parietal corridors increasing from the inlet of the exchanger to its outlet.

A power turbine blade for a gas turbine engine includes at least one of a third-stage blade, a fourth-stage blade, or a fifth-stage blade. The power turbine blade includes a platform and an airfoil that extends in a spanwise direction from the platform to a tip shroud. The airfoil has an external airfoil surface formed in conformance with multiple cross-section airfoil profiles described by a set of Cartesian coordinates.

A gas turbine engine includes: a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath; a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio during operation of the gas turbine engine in a cruise operating mode; and a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan; wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8.

A system for supplying power to a portable battery pack including a battery enclosed by a wearable and replaceable pouch or skin using at least one solar panel is disclosed, wherein the pouch or skin can be provided in different colors and / or patterns. Further, the pouch or skin can be MOLLE-compatible. The battery comprises a battery element housed between a battery cover and a back plate, wherein the battery element, battery cover, and back plate have a slight curvature or contour. Further, the battery comprises flexible leads.

A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.

The present application relates to the field of metal material manufacturing, and in particular to a GH4151 alloy bar and a preparation method therefor and a use thereof. The method comprises the following steps: (1) carrying out homogenization heat treatment on a GH4151 alloy cast ingot to obtain a heat treated bar; (2) forging the heat treated bar to obtain a forged bar; (3) carrying out hot rolling treatment on the forged bar to obtain a hot rolled bar; (4) carrying out solution treatment on the hot rolled bar to obtain a solid solution bar; (5) carrying out cold drawing treatment on the solid solution bar to obtain a cold drawn bar; and (6) carrying out aging treatment on the cold drawn bar to obtain a GH4151 alloy bar. According to the GH4151 alloy bar prepared in the present application, grains can be effectively refined, the surface dimensional precision is improved, and the tensile strength of the prepared GH4151 alloy bar is improved, so that the GH4151 alloy bar has excellent mechanical properties.

Methods and apparatus for anti-ice heat supply from waste heat recovery systems are disclosed. An apparatus for an aircraft, the apparatus comprising a fuel heat exchange system, an anti-ice heat exchanger, a waste heat recovery heat exchanger, and a conduit coupled to the fuel heat exchange system, the anti-ice heat exchanger, and the waste heat recovery heat exchanger, the conduit including a first portion and a second portion distinct from the first portion, wherein the first portion of the conduit carries a first portion of a thermal transfer fluid from the waste heat recovery heat exchanger to the anti-ice heat exchanger in which the thermal transfer fluid supplies anti-ice heat to a portion of the aircraft, wherein the second portion of the conduit carries a second portion of the thermal transfer fluid from the waste heat recovery heat exchanger to the fuel heat exchange system.

A triple-flow turbine engine for an aircraft, this turbine engine having a longitudinal axis and including upstream suspension elements which are located in a first plane perpendicular to the axis and are connected or fixed to the gas generator of the turbine engine; downstream suspension elements which are located in a second plane perpendicular to the axis and are connected or fixed to the gas generator; and thrust-absorbing rods which include first ends which are connected or fixed to the gas generator and opposing second ends which are located in a third plane perpendicular to the axis, wherein the first, second and third planes are located at a cold compartment of the gas generator.

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.

A lubrication system for a turbine engine. The turbine engine includes a propulsor and one or more rotating components. The lubrication system includes a sump, a primary lubrication system, an auxiliary lubrication system, and a strain energy storage system. The sump stores lubricant therein. The primary lubrication system supplies the lubricant from the sump to the one or more rotating components during normal operation of the turbine engine. The auxiliary lubrication system includes an auxiliary pump including an auxiliary pump shaft. The strain energy storage system includes a spring drivingly coupled to the auxiliary pump shaft. The spring stores strain energy during normal operation of the turbine engine and releases the strain energy when the propulsor is windmilling to rotate the auxiliary pump shaft to power the auxiliary pump such that the auxiliary pump pumps the lubricant from the sump to the one or more rotating components.

The application provides an aero-engine acceleration limiting method, and relates to the technical field of aero-engines, and comprises the following steps: receiving a flight control instruction, calculating a first fuel supply amount according to the flight control instruction, and executing single / dual-outer-envelop working mode conversion; detecting a position state of an actual mode selection valve and a rotating speed feedback value in real time; judging whether the current state of the engine meets the mode switching condition of single / dual-outer-envelop according to the position state of the mode selection valve; calculating a second fuel supply amount, a third fuel supply amount, and a fourth fuel supply amount according to the position state of the mode selection valve, combining the first fuel supply amount, a fifth fuel supply amount, taking the minimum value among the five fuel supply amounts, comparing the minimum value with a sixth fuel supply amount to take the maximum value, taking the minimum value with a seventh fuel supply amount, taking the maximum value of the comparison result with an eighth fuel supply amount as a control fuel supply amount output; and judging whether the engine is in an acceleration / deceleration state according to a rotating speed difference. The application solves the safety problem possibly caused by the acceleration / deceleration process of the dual-outer-envelop engine.

The invention belongs to the field of aircrafts, relates to an aircraft power system design technology, and provides an aero-engine of a contra-rotating blade configuration and a control method, and the aero-engine comprises a gas turbine engine, contra-rotating blades and a main transmission shaft. The gas turbine engine is located in the main flow channel, the contra-rotating blades are arranged in the additional duct, an inlet adjusting piece is arranged at the inlet position of the additional duct, and an outlet adjusting piece is arranged at the outlet position of the additional duct. One end of the main transmission shaft is selectively connected or disconnected with the gas turbine engine through a clutch, and the other end of the main transmission shaft is connected with the contra-rotating blades through a connecting structure. Through the innovative power architecture design, the core problem that the thermodynamic cycle efficiency and the thrust characteristic of a traditional aero-engine under the complex flight working condition are difficult to dynamically adapt is solved, the bypass ratio is adjusted through the external contra-rotating blades, multi-working-condition self-adaptive optimization is achieved, and the reliability of the aero-engine is improved. Therefore, the aircraft power system can efficiently and reliably operate in a wide-area complex environment.

An assembly for an aircraft propulsion system includes an open propulsor rotor and an open guide vane structure. The open propulsor rotor is rotatable about a rotational axis. The open guide vane structure is axially next to and downstream of the open propulsor rotor. The open guide vane structure includes a plurality of open guide vanes arranged circumferentially about the rotational axis. The open guide vanes include a first open guide vane. The first open guide vane includes a camber line, a leading edge, a trailing edge, a leading edge section and a trailing edge section. The first guide vane extends longitudinally along the camber line from the leading edge to the trailing edge. The leading edge section is configured to pivot about a pivot axis. The trailing edge section is configured to move relative to the leading edge section to change a camber of the first open guide vane.

The invention belongs to the technical field of overall aerodynamic design of helicopters, and particularly relates to an infrared suppressor fairing of a helicopter and an air heat optimization design method of the infrared suppressor fairing. The fairing comprises a fairing body, a shutter air inlet, a square protective net air inlet, a second-stage exhaust pipe and a first-stage mixed exhaust pipe. The front of the fairing body is regarded as a half ellipse which is laterally cut along a long axis; a shutter air inlet is formed in the front portion of the fairing body, and the plane of a shutter is crescent A square protective net air inlet is formed in front of the top of the fairing body; an oval second-stage exhaust pipe is arranged behind the top of the fairing body so that high-temperature exhaust can be further cooled by means of downwash airflow of the rotor wings. A first-stage mixed exhaust pipe is sleeved in the second-stage exhaust pipe; and the primary mixing exhaust pipe is connected with an outlet of the transmitter. According to the method, the problem that the pneumatic performance and the thermal performance are difficult to consider in a traditional design method is solved, comprehensive optimization of the pneumatic performance and the thermal performance of the fairing is achieved, and a pneumatic design scheme of the infrared suppressor fairing is formed.

A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows:(TOUTADTExit)2*EGTAHPCExit*10-11;wherein CSP is greater than 0.0001194×EGT2−0.103×EGT+22.14 and less than 0.0003294×EGT2−0.306×EGT+77.91; and wherein EGT is greater than 525 degrees Celsius and less than 1250 degrees Celsius.