Abstract: Systems and methods for operating an engine in a multi-engine rotorcraft are described herein. A first parameter indicative of torque of a first engine is obtained. A decrease of the first parameter is detected. In response to detecting the decrease of the first parameter, an autorotation of the rotorcraft is accommodated, A second parameter indicative of torque of a second engine of the rotorcraft is assessed while accommodating the autorotation. If the second parameter has not decreased, a shaft shear of the first engine is identified and accommodating of the autorotation is ended. If the second parameter has decreased, the accommodating is maintained.

Abstract: An aspect includes a variable cycle system of a gas turbine engine. The variable cycle system includes an actuation system, an electric component, and a controller. The actuation system is configured to adjust a variable cycle of turbomachinery of the gas turbine engine. The electric component is operable to provide a shaft power supply or a load corresponding respectively to an adjustment of the turbomachinery. The controller is operable to adjust an output of either or both of the actuation system and the electric component for separate control of thrust and cycle responses.

Abstract: Gearboxes for aircraft gas turbine engines, in particular to arrangements for journal bearings such gearboxes, and to related methods of operating such gearboxes and gas turbine engines. Example embodiments include a gearbox for an aircraft gas turbine engine, the gearbox including: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin.

Abstract: Systems, program products, and methods for adjusting operating limit (OL) thresholds for compressors of gas turbine systems based on mass flow loss are disclosed herein. The systems may include at least one computing device in communication with the gas turbine system, sensor(s) measuring operational characteristic(s) of the gas turbine system, and a pressure sensor measuring an ambient fluid pressure surrounding the gas turbine system. The computing device(s) may be configured to adjust operational parameters of the gas turbine system by performing processes including determining a mass flow loss between an estimated, first mass flow rate and a calculated, second mass flow rate for the compressor of the gas turbine system, and adjusting an OL threshold for the compressor of the gas turbine system based on the mass flow loss. The OL threshold for the compressor may be below a predetermined surge threshold for the compressor.

Abstract: A method for operating a gas turbine engine comprises providing fuel flow and compressed airflow to a combustor with a fuel-to-air ratio, the compressed airflow being from a compressed air source; detecting at least one parameter indicative of the fuel-to-air ratio being below a predetermined value; and bleeding compressed air from the compressed air source when the at least one parameter indicative of the fuel-to-air ratio is below the predetermined value to increase the fuel-to-air ratio to at least the predetermined value.

Abstract: The present disclosure provides systems and methods for power production. In particular, the systems and methods utilize the addition of heat to an expanded turbine exhaust stream in order to increase the available quantity of heat for recuperation and use therein for heating a compressed carbon dioxide stream for recycle back to a combustor of the power production system and method.

Abstract: A method and control system for controlling compressor output for a gas turbine engine is disclosed. The power output of a gas turbine engine can vary and be below desired output levels due to operating conditions such as ambient temperature and elevation. These operating conditions can lead to lower output of the gas compressor of the turbine engine and lower operating temperatures within or proximate to a turbine of the gas turbine engine and lead to less power output. Additional fuel can be added to increase power to the gas producer shaft and increase turbine temperature of the gas turbine engine. A power transfer device can be used to remove or add power to the gas producer shaft to balance the gas producer mechanical limits and turbine thermal limits at maximum levels and lead to higher power output.

Abstract: A gas turbine engine comprises a compressor having a plurality of blades mounted to a hollow, annular compressor drum. The compressor comprises an electric storage device mounted within the hollow compressor drum.

Abstract: A method, including: operating an industrial gas turbine engine having a plurality of combustor cans arranged in an annular array, each can having burner stages and a pilot burner arrangement having a premix pilot burner and a diffusion pilot burner; operating in asymmetric combustion, wherein at least one can is a warm can where respective burners stages are off and remaining cans operate as hot cans where respective burner stages are on; and while maintaining a constant rate of fuel flow to the pilot burner arrangement of the warm can, changing fuel fractions within the pilot burner arrangement of the warm can.

Abstract: A method of reducing applied heat within an inlet duct of a gas turbine generating electricity includes applying heat to the inlet duct of the gas turbine to attain an initial temperature set point and to produce conditions sufficient for preventing formation of ice within the inlet duct, measuring a position of an inlet guide vane (IGV) of the gas turbine, an inlet duct temperature, and an inlet duct relative humidity to determine a thermodynamic state in the inlet duct, evaluating the thermodynamic state to determine if the conditions are sufficient for preventing formation of ice within the inlet duct, and in response to determining that sufficient conditions exist within the inlet duct for preventing formation of ice, adjusting the applied heat to maintain the measured inlet duct temperature.

Abstract: An EV driving sound control system may include a sound output device of generating a driving sound of an EV vehicle; a torque measurement sensor of measuring a torque of a motor in the EV vehicle; an indoor noise measurement sensor of detecting a noise of the EV vehicle; and a signal processing controller for controlling the sound output device to reach a target tone by receiving signals from the torque measurement sensor and the indoor noise measurement sensor, wherein the signal processing controller extracts and outputs only the acceleration feeling sound component from an engine sound upon acceleration, and masks high-frequency noise upon deceleration by the sound output device, by use of the traveling information related to the EV vehicle inputted from the CAN communication in real time and the signals inputted from the indoor noise measurement sensor and the torque measurement sensor.

Abstract: An oil supply system for a gas turbine engine has a lubricant pump delivering lubricant to an outlet line. The outlet line is split into at least a hot line and into a cool line, with the hot line directed primarily to locations associated with an engine that are not intended to receive cooler lubricant, and the cool line directed through one or more heat exchangers at which lubricant is cooled. The cool line then is routed to a fan drive gear system of an associated gas turbine engine. A method and apparatus are disclosed. The heat exchangers include at least an air/oil cooler wherein air is pulled across the air/oil cooler to cool oil. The air/oil cooler is provided with an ejector tapping compressed air from a compressor section to increase airflow across the air/oil cooler.

Abstract: A tank assembly (100) for a gas turbine engine is provided comprising a tank (102) and a plurality of restraints (112, 114). The restraints (112, 114) include a fixing part (124) for securing the tank (102) to a support structure (104). A first restraint (112) has a first rigidity in the direction of a length of the tank (102) and at least one second restraint (114) has a second rigidity in the direction of the length of the tank (102). Upon thermal expansion or contraction of the tank (102) relative to the support structure (104), the or each second restraint (114) flexes to a greater extent than the first restraint (112).

Abstract: Flow multiplier systems for aircraft are described herein. A flow multiplier system includes a turbo-compressor having a compressor, a turbine, and a drive shaft coupled between the compressor and the turbine. A compressor outlet of the compressor is fluidly coupled to an ejector in a gas turbine engine. The system also includes a supply line fluidly coupling a compressed air tank and a turbine inlet and a valve coupled to the supply line. The system includes a controller configured to, based on an input signal requesting to increase output power of the gas turbine engine, send a command signal to open the valve to enable a flow of pressurized air from the compressed air tank to the turbine inlet. The turbine drives the compressor to create high pressure air at the compressor outlet, which is provided into the gas turbine engine to increase the output power.

Abstract: A hybrid-electric propulsion system includes a turbomachine and an electrical system, the electrical system including an electric machine coupled to the turbomachine. A method for operating the propulsion system includes receiving, by one or more computing devices, a command to accelerate the turbomachine to provide a desired thrust output; receiving, by the one or more computing devices, data indicative of a temperature parameter approaching or exceeding an upper threshold; and providing, by the one or more computing devices, electrical power to the electric machine to add power to the turbomachine to provide, or assist with providing, the desired thrust output in response to receiving the command to accelerate the turbomachine and receiving the data indicative of the temperature parameter approaching or exceeding the upper threshold.

Abstract: Herein provided are methods and systems for controlling an engine having a variable geometry mechanism. A power level difference between a requested engine power level and a current engine power level is determined at a computing device. The power level difference is compared to a predetermined power threshold at the computing device. When the power level difference exceeds the predetermined power threshold, a position control signal for changing a position of the variable geometry mechanism is generated and output at the computing device, the position control signal generated based on a requisite bias level, the requisite bias level being based on the power level difference.

Abstract: A gas turbine engine includes a lubrication system, fuel system and thermoelectric heat exchanger adapted for selective operation in response to operational states of the gas turbine engine.

Abstract: An aircraft comprising a main engine to provide a supply of bleed air, an APU housing within an APU compartment and having an APU bleed valve, an APU bleed air duct connecting main engine with the APU, and a bleed air heating system for the APU compartment comprising an auxiliary pipeline connecting the APU bleed air duct with the APU compartment, a temperature sensor, an auxiliary pipeline valve to control the discharge of bleed air into the APU compartment, and a temperature controller configured to establish a heating operation mode, when the sensed temperature falls below a minimum temperature threshold value, and a standby operation mode, when the sensed temperature surpasses a maximum temperature threshold value. The temperature controller operates the main engine and the valves to establish these operation modes.

Abstract: An aircraft turboprop engine includes at least a low-pressure body and a high-pressure body. The low-pressure body drives a propeller by means of a gearbox. The turboprop engine also includes means for supplying air to an air-conditioning circuit of an aircraft cabin, wherein the supply means has at least one compressor borne by the gearbox and of which the rotor is coupled to the low-pressure body by means of the gearbox.

Abstract: A gas turbine engine includes a first shaft coupled to a first turbine and a first compressor, a second shaft coupled to a second turbine and a second compressor, and a third shaft coupled to a third turbine and a fan assembly. The turbine engine includes a heat rejection heat exchanger configured to reject heat from a closed loop system with air passed from the fan assembly, and a combustor positioned to receive compressed air from the second compressor as a core stream. The closed-loop system includes the first, second, and third turbines and the first compressor and receives energy input from the combustor.

Abstract: A system for controlling a plurality of electromechanical effectors operably connected to an engine to control engine parameters. The system also includes a plurality of sensors operably connected to measure a state or parameter of each effector, a power supply configured to supply power to the plurality of effectors, and a controller operably connected to the plurality of sensors, the plurality of effectors, and the power supply. The controller executes a method for an adaptive model-based control for controlling each effector, The method includes receiving a request indicative of a desired state for each effector, receiving a weighting associated each request, obtaining information about a current state of each effector, and updating an adaptive model based control (MBC) based upon the information. The method also includes generating a control command for an effector based upon the adaptive MBC and commanding the effector based upon the control command.

Abstract: A two-shaft gas turbine facility includes a two-shaft gas turbine, an induction motor connected to a compressor of the two-shaft gas turbine, a secondary battery, a first frequency converter that controls power transmission and reception between a power system and the induction motor, and a second frequency converter that controls charging and discharging of power between the secondary battery and the power system. A first control unit of a control device causes power transmission and reception to be performed between the induction motor and the power system by the first frequency converter if a required output change rate is higher than a maximum output change rate. A second control unit causes the secondary battery to be charged and discharged by the second frequency converter if power to be transmitted to and received from the induction motor has reached maximum allowable power.

Abstract: An engine system and method for operating an internal combustion engine in dynamically varying conditions. An exemplary system comprises an internal combustion engine configured to receive both a primary fuel and a secondary fuel into one or more chambers in which a combustion process occurs, a fuel injection system, an air intake manifold and a fuel manifold; an electronic system which controls timing and metering of the primary fuel and/or the secondary fuel in the combustion process; and a plurality of sensors positioned to measure one or more variables associated with combustion of the primary fuel in the presence of the secondary fuel. The electronic system is configured to apply a control signal to adjust an engine setting to reduce NOx emissions based in part on the magnitude of the variable.

Abstract: A fuel and bleed controller is provided. The fuel and bleed controller includes a processor and a memory. The memory stores program instructions thereon. The program instructions are executable by the processor to cause the fuel and bleed controller to send status requests to systems of the aircraft. The systems comprise a bleed system and bleed user controllers. The program instructions are further executable by the processor to cause the fuel and bleed controller to receive status responses from the systems of the aircraft and determine fuel requirements based on the status responses in advance of operational needs by the systems of the aircraft. The program instructions are further executable by the processor to cause the fuel and bleed controller to control engines of the aircraft based on the fuel requirements.

Abstract: A method for determining an emission behaviour of a gas turbine engine. In order to provide a reliable operation of the gas turbine engine the method includes: parameterising the emission behaviour of the gas turbine engine for at least one selected first state variable of the gas turbine engine by using a model, which reflects a state behaviour of the gas turbine engine, and determining the emission behaviour of the gas turbine engine by using the parameterisation.

Abstract: A temperature control device includes a temperature difference calculation unit that calculates a temperature difference between an inlet and an outlet of a boost compressor on the basis of a pressure ratio of the inlet and the outlet of the boost compressor and an IGV opening degree of the boost compressor, the boost compressor outputting cooling air obtained by cooling compressed air from a compressor to a cooling target; a temperature information calculation unit that calculates temperature information for feedback control for at least one of the inlet and the outlet of the boost compressor on the basis of the temperature difference between the inlet and the outlet of the boost compressor; and a control unit that performs feedback control by using the temperature information for feedback control such that at least one of an inlet temperature and an outlet temperature of the boost compressor approaches a setting value.

Abstract: A turbine engine includes a fan connected to a fan shaft, a combustion chamber, and an electric motor/generator in communication with the fan shaft. A controller is configured to direct power into the electric motor/generator during engine accelerations from steady state such that air flow to the combustion chamber is increased. The controller is further configured to direct power out of the electric motor/generator during engine decelerations from steady state such that air flow to the combustion chamber is decreased.

Abstract: An apparatus and method of controlling a multi-stage compressor of a gas turbine engine having a front-block of stages with variable stator vanes (VSVs), a rear block of stages downstream of the front-block, and a bleed air off-take from at least one of the stages. The method includes sensing the pressure of the bleed air from the bleed air off-take and adjusting the position of the VSV for the front-block.

Abstract: A gas turbine engine has a turbine driving a gear reduction. The gear reduction drives a fan drive shaft assembly to, in turn, drive a fan rotor at a reduced speed relative to the turbine. A generator generates current upon rotation of the fan drive assembly. The generator supplies electric current to an electric motor driven pump. The electric motor driven pump delivers lubricant to components within the gear reduction.

Abstract: In one embodiment, a system is provided. The system includes a turbine control system, comprising a processor. The processor is configured to receive an input for transitioning between a normal load path (NLP) of a turbine system and a cold load path (CLP) of the turbine system. The processor is additionally configured to determine a carbon monoxide (CO) setpoint based on the input. The processor is further configured to apply a temperature control based on the CO setpoint, wherein the normal load path comprises higher emissions temperatures as compared to the cold load path.

Abstract: A method for controlling a gas turbine when the gas turbine is started may include measuring the number of rotations of the gas turbine, determining a target acceleration rate based on the number of rotations of the gas turbine, measuring a current acceleration rate, determining a flow rate of fuel based on a difference between the current acceleration rate and the target acceleration rate, and supplying fuel to the gas turbine at the determined flow rate. The flow rate of the fuel may be determined between a maximum flow rate of the fuel that has been previously stored and a minimum flow rate of the fuel. A temperature of the exhaust gas after controlling the flow rate of the fuel may be monitored.

Abstract: A method and system for use with actuation system to control a plurality of vanes disposed within a turbine engine, wherein a vane position sensor provides a vane position signal corresponding to a vane position of the plurality of vanes, the actuation system includes a plurality of motors engaged in response to the vane position signal, and a differential gearbox having a plurality of inputs operatively coupled to the plurality of motors and an output operatively coupled to the plurality of vanes, wherein an output speed of the output is a sum of a plurality of input speeds of the plurality of inputs.

Abstract: A method includes: obtaining a measurement; determining whether the measurement is in a first range, a second range, or a third range; controlling operation of a valve in a full-freeze mode when the data measurement falls within the first range; controlling the operation of the valve of the turbine system in a semi-freeze mode when the data measurement falls within the second range; and controlling the operation of the valve of the turbine system in a full-operation mode when the data measurement falls within the third range.

Abstract: The embodiments relate to a method for the computer-aided control and/or regulation of a technical system, particularly a power generation installation. The actions to be performed in the course of regulation or control are ascertained using a numerical optimization method (e.g., a particle swarm optimization). In this case, the numerical optimization method uses a predetermined simulation model that is used to predict states of the technical system and, on the basis thereof, to ascertain a measure of quality that reflects an optimization criterion for the operation of the technical system.

Abstract: A computer-implemented method for controlling and/or regulating a technical system, in which actions to be carried out on the technical system are first of all determined using an action selection rule which was determined through the learning of a data-driven model and, in particular, a neural network. On the basis of these actions a numerical optimization searches for actions which are better than the original actions according to an optimization criterion. If such actions are found, the technical system is regulated or controlled on the basis of these new actions, such that the corresponding actions are applied to the technical system in succession. The method is suitable, in particular, for regulating or controlling a gas turbine, wherein the actions are preferably optimized with respect to the criterion of low pollutant emission or low combustion chamber humming.

Abstract: Embodiments are directed to an engine of an aircraft, a starter generator configured to receive power to cause the engine to start and to provide electrical power to at least one load, and a battery system comprising: a battery, and at least one of a controller and a regulator configured to limit an in-rush current to the battery following the starting of the engine.

Abstract: A lubricant supply system for a gas turbine engine has a lubricant lube pump delivering lubricant to an outlet line. The outlet line is split into at least a hot line and into a cool line, with the hot line directed primarily to locations associated with an engine that are not intended to receive cooler lubricant, and the cool line directed through one or more heat exchangers at which lubricant is cooled. The cool line then is routed to a fan drive gear for an associated gas turbine engine. A method and apparatus are disclosed.

Abstract: Various embodiments include a system having: at least one computing device configured to tune a set of gas turbines (GTs) by performing actions including: commanding each GT in the set of GTs to a base load level, based upon a measured ambient condition for each GT; commanding each GT in the set of GTs to adjust a respective power output to match a nominal power output value, and subsequently measuring an actual emissions value for each GT; adjusting an operating condition of each GT in the set of GTs based upon a difference between the respective measured actual emissions value and a nominal emissions value at the ambient condition; and setting a target operating condition for each GT to match the adjusted operating condition after the adjusting of the operating condition.

Abstract: Various embodiments include a system having: at least one computing device configured to tune a set of gas turbines (GTs) by performing actions including: commanding each GT in the set of GTs to a base load level, based upon a measured ambient condition for each GT; commanding each GT in the set of GTs to adjust a respective power output to match a nominal power output value, and subsequently measuring an actual emissions value for each GT; adjusting an operating condition of each GT in the set of GTs based upon a difference between the respective measured actual emissions value and a nominal emissions value at the ambient condition; and building an independent emissions model for each GT based upon the measured actual emissions value for each GT and the adjusted operating condition of each GT.

Abstract: A control system for a gas turbine engine including a power turbine is disclosed. The control system may include a control module to receive engine operating goals and an estimated current engine state, wherein the estimated current engine state is produced by a model-based estimation module using a bandwidth signal produced by an adaptation logic module. The control module is operative to determine fuel flow, inlet guide vane schedules and stability bleed schedules based at least in part on the received engine operating goals and the estimated current engine state, and to send signals to a gas generator of the gas turbine engine in order to control the gas generator according to the determined fuel flow, inlet guide vane schedules and stability bleed schedules.

Abstract: A system method of reducing engine induced aircraft cabin resonance in an aircraft includes sensing a parameter representative of current aircraft flight conditions, and sensing a parameter representative of current engine operating conditions of at least one of a first turbofan gas turbine engines or a second turbofan gas turbine engine. In a control system, the parameter representative of current aircraft flight conditions and the parameter representative of current engine operating conditions are processed to supply a variable inlet guide vane (VIGV) offset value. The VIGV offset value is applied to a VIGV reference command associated with one of the first or second turbofan gas turbine engine, to thereby cause the VIGVs of one of the first or second turbofan gas turbine engine to move to a more closed position.

Abstract: A method of controlling the angular orientation of at least one variable fluidfoil. The method includes the steps of receiving detected fluidfoil angular orientation data for the at least one fluidfoil; receiving a fluidfoil angular orientation demand; generating a modified angular orientation demand by modifying the fluidfoil angular orientation demand before it is used to control adjustment of the at least one fluidfoil, the modification being performed using the detected fluidfoil angular orientation data; and outputting the modified demand and controlling the at least one fluidfoil angular orientation in accordance with the modified demand.

Abstract: A combustion device (10) includes at least one combustion chamber (11) with a plurality of burners (B1, . . . , Bn) operating in parallel which produce in each case a flame (F1, . . . , Fn) which reaches into the combustion chamber (11), wherein each of the burners (B1, . . . , Bn), via a fuel distribution system (18), is supplied with a fuel from a fuel supply (16), which fuel distribution system (18) includes control elements (V1, . . . , Vm) for manual or controlled regulating of the fuel supply and/or fuel composition of individual burners (B1, . . . , Bn) and/or groups of burners (B1, . . . , B3; Bn?2, . . . , Bn). In a method of using the combustion device, a quick optimization or homogenization is achieved by a function (F) of the flame temperatures of the burners (B1, . . . , Bn) being provided in dependence upon the positions of the control elements (V1, . . .

Abstract: An example method of initiating a maintenance action on a component includes monitoring an electrical current required to maintain a steady state position. The method then initiates a maintenance action on the component based on the monitored current.

Abstract: A control system for a gas turbine engine including a power turbine is disclosed. The control system may also include an outer loop control module to determine a torque request. The outer loop control module may include a feedback control component operative to provide regulation of the power turbine, a feed-forward component operative to anticipate a load on the power turbine, and a hybrid control component operative to prevent output of a torque request that cannot currently be delivered by the power turbine. The control system may also include an inner loop control module to receive the torque request from the outer loop control module, to determine fuel flow and inlet guide vane schedules based at least in part on the received torque request, and to send signals to control a gas generator of the gas turbine engine according to the determined fuel flow and inlet guide vane schedules.

Abstract: The present invention provides a method of testing gas supply of a gas appliance, which includes the steps of: monitoring a flow rate of a gas flow through a gas regulator of a pipeline to have a detected gas flow rate; comparing the detected gas flow rate with the ideal range of gas flow rate; and providing an alarm when the detected gas flow rate exceeds the ideal range of gas flow rate. The present invention further provides a compensating method when an abnormal condition is detected. The compensating method will change the gas flow rate or the air flow rate to get a proper air fuel ratio of the mixed gas.

Abstract: A method for adjusting the operation of a turbomachine integrated with an exhaust gas recirculation (EGR) system is provided. The method may utilize the composition of an inlet fluid entering the turbomachine. The method may also utilize a variety of turbomachine operating data.

Abstract: Systems, devices, and methods for controlling a fuel supply for a turbine or other engine using direct and/or indirect indications of power output and optionally one or more secondary control parameters.

Abstract: In a fuel supply device of a gas turbine engine of the present invention, a fuel divider 66 includes a fuel entrance E1 into which the fuel supplied from a collecting fuel passage 63 is introduced; a pilot port 76 connected to the pilot fuel passage 64; a main port 77 connected to the main fuel passage 65; a pilot port needle valve element 101 which adjusts an opening degree of the pilot port 76; a main port needle valve element 102 which opens and closes the main port 77; and a drive element 100 which is actuated according to the fuel pressure at the fuel inlet E1 to drive the pilot port needle valve element 101 and the main port needle valve element 102.

Abstract: A method for determining a cooling flow parameter of a cooling medium supplied through a gas turbine is disclosed. The method may generally include receiving a signal associated with a first value of a combustion product parameter at a location within a combustion zone of the gas turbine, receiving a signal associated with a second value of a combustion product parameter at a location downstream of the combustion zone, comparing the first and second values of the combustion product parameter and determining a cooling flow parameter of the cooling medium based on the comparison of the first and second values.