RU2489592C1 - Method of controlling fuel feed to gas turbine engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: additionally, in case signal "Fire in nacelle" is generated at takeoff by aircraft anti-fire system turbo compressor current rpm is fixed to be used as preset rpm magnitude for preset time interval. Said interval elapsed, fuel feed is terminated to start the engine. This allows normal operation of the engine at preset thrust required for normal takeoff even in the case of fire in nacelle.

EFFECT: higher reliability and safety.

1 dwg

Description

The invention relates to the field of aircraft engine manufacturing and can be used in electronic systems (ACS) for automatic control of gas turbine engines (GTE).

A known method of controlling a gas turbine engine implemented in an electronic hydromechanical self-propelled guns of a supervisory type. Keba I.V. “Flight operation of helicopter gas turbine engines”, M .: “Transport”, 1976, p.123-125.

The method consists in the fact that in order to improve control accuracy, the control action of the hydromechanical controller is adjusted in a limited range by an electronic corrector.

The disadvantage of this method is its low efficiency.

Closest to this invention in technical essence is a method of controlling fuel consumption in a gas turbine engine, which consists in measuring the position of the engine control lever (ORE), the speed of the turbocharger and free turbine (ST), the pressure and temperature of the air at the engine inlet, temperature of gases behind the gas generator turbine, form a predetermined value of the turbocompressor speed as a function of the position of the throttle pressure and air temperature at the engine inlet, set the limits for this engine the temperature of the gases behind the gas generator turbine and the rotational speed CT, compare the set value of the turbocharger rotation speed and the measured one, compare the gas temperature limit for the engine behind the gas generator turbine and the measured one, compare the temperature limit of the CT speed for the engine and the measured one, compare the obtained discrepancies to a minimum with the signal of the pick-up machine (AL), the selected value is fed to the proportional-integral (PI) controller, where the correcting effect on the fuel consumption meter supplied to the combustion chamber (CS) of the engine.

The disadvantage of this method is the following.

For new generation engines, for example, the PD-14 engine developed by Aviadvigatel OJSC, Perm, which is part of the power plant (SU) of the MS-21 aircraft developed by Irkut OJSC, Moscow, the following requirement is imposed: the engine is in the process of take-off the aircraft must provide take-off thrust even in the event of a fire in the engine nacelle.

When using the known method in self-propelled guns PD-14, this requirement cannot be fulfilled due to the following reasons.

When a fire occurs in the engine nacelle, the “external” sensors and their communication lines (sensors and communication lines located outside the engine casing: pressure gauges of the throttle pressure and air temperature at the engine inlet) are the first to fail. In contrast to them, the sensors located in the “hot” part of the engine work practically “to the end”: thermocouples, speed sensors located in the cooled engine mounts.

The electronic engine controller (for the PD-14 engine is the RED-14 unit developed by STAR OJSC, Perm) and the executive part of the self-propelled guns, which provide fuel metering into the combustion chamber and control engine mechanization (for the PD-14 engine, is the DG unit -14 developments of STAR OJSC), have special protection that allows working in conditions of elevated ambient temperature.

Despite this, when the known control method is implemented in self-propelled guns, the loss of information about the position of the throttle throttle pressure and air temperature at the engine inlet caused by a fire in the engine nacelle will not allow maintaining the take-off thrust of the engine. This reduces the reliability of the SU and the safety of the aircraft.

The aim of the invention is to increase the reliability of the SU and the safety of the aircraft.

This goal is achieved by the fact that in the method of controlling fuel consumption in a gas turbine engine, which consists in measuring the position of the throttle throttle speed of the turbocompressor and CT, the pressure and temperature of the air at the engine inlet, the temperature of the gases behind the gas generator turbine, form a predetermined value of the speed of the turbocompressor as the function of the throttle position of the pressure and air temperature at the engine inlet, the gas temperature limits for the engine behind the gas generator turbine and the speed of rotation CT are set for this engine, comparing the target value of the turbocharger speed is measured and measured, the temperature limit of the gas for the engine behind the gas generator turbine is measured and measured, the limit of the speed of the gas for the engine and the temperature measured for the engine are compared, the resulting discrepancies are selected to a minimum with the AP signal, the selected value is fed to PI regulator. where a control action is formed on the fuel consumption meter supplied to the engine КС, additionally, when the “Fire in engine nacelle” signal generated by the aircraft’s fire alarm system is received during take-off, the current value of the turbocharger speed is recorded and used as the set value of the turbocharger speed in for a predetermined time, after which the fuel supply to the compressor station is stopped and the engine is turned off.

The figure shows a diagram of a device that implements the inventive method.

The device contains a series-connected block 1 of sensors (DB), an electronic controller 2 of the engine (RED), an electrohydrogen converter 3 (EGP), a fuel dispenser 4, a stop valve 5 (KO), a controlled input KO 5 connected to the output of the RED 2.

RED 2 is an on-board digital computer (BCM) containing read-only memory (ROM), which contains software (software) that implements engine control algorithms. Additionally, the digital computer is equipped with input / output devices (I / O) of physical signals (from database 1 and in EGP 3), random access memory (RAM), which is necessary for the processor to process information coming from the air-blast computer, reprogrammable memory (RPZU), necessary for storing information relating to the individual characteristics of the engine (operational adjustments, operating hours, remaining life). A computer, ROM, software, air-blast, RAM, processor, RPG are not shown in the figure.

The device operates as follows.

In RED 2 using DB 1 measure the position of the throttle speed of the turbocharger and CT, the pressure and temperature of the air at the engine inlet, the temperature of the gases behind the gas generator turbine.

According to the dependencies stored in ROM RED 2 in advance:

- form a predetermined value of the speed of the turbocharger as a function of the position of the throttle pressure and air temperature at the engine inlet (an example of such a dependence is given, for example, in the book Besekersky VA, Popov EP "Theory of automatic control", M., "Science", 1975, p. 34-35),

- set the temperature limits for a given engine for the temperature of the gases behind the gas generator turbine and the rotational speed of the CT (for the PD-14 engine these values are 1370K for the temperature of the gases and 8000 rpm for the rotational speed of the CT).

Then, in RED 2, the set value of the turbocharger speed is measured and measured with the help of DB 1, the temperature limit of the gases for the engine behind the gas generator turbine and the value measured with the help of DB 1 is compared, the limit value of the rotational speed CT for this engine is measured and measured with the help of DB 1 .

The resulting discrepancies are selected to a minimum with the AP signal (not shown in the figure), working, for example, according to the program

$$\text{GT}={\text{f (α}}_{\text{RUD}}{\text{, T}}^{\ast }{}_{\text{VX}}{\text{,R}}^{\ast }{}_{\text{VX}}{\text{,R}}_{\text{to}}\text{, nk)}\text{(one)}$$

Where

GT - the maximum allowable fuel consumption for a given engine operation mode,

α _ORE - position ORE

T * _BX - air temperature at the engine inlet,

P * _BX , - air pressure at the engine inlet,

P _to - the air pressure behind the engine compressor,

nк - engine compressor speed.

The selected value is fed to the PI controller (not shown in the figure), where a control action is formed on the fuel consumption meter, supplied with the help of air-blast red 2 through EGP 3 to meter 4, with which the fuel consumption is controlled in the engine CS. The signal from RED 2 to KO 5 is absent and KO 5 is in the open position.

Additionally, when the signal “Fire in the engine nacelle” generated by the aircraft’s fire alarm system (not shown) is received in RED 2 during the take-off of the aircraft, the current value of the turbocharger speed is recorded and used as the set value of the turbocharger speed for a predetermined time, after which, according to the commands of RED 2 using EGP 3, dispenser 4 and KO 5 stop the flow of fuel to the compressor station and turn off the engine.

For the PD-14 engine and MS-21 aircraft, the take-off mode is determined by the simultaneous fulfillment of the following conditions:

- the installation angle of the ore is greater than 80 °;

- the parking brake is off;

- compressor rotor speeds are greater than 99%.

A predetermined time during which even with the signal “Fire in the engine nacelle”, the self-propelled gun retains the take-off mode of the engine - 5 minutes (set in the statement of work on the self-propelled guns of the PD-14 engine).

T.O. by improving the quality of fuel consumption control in the engine's CS, the aircraft takes off normally even in the event of a fire in the engine nacelle. After the aircraft takes off, the engine, in the engine nacelle of which a fire occurred, is turned off, the fire is localized and eliminated (using the aircraft’s fire protection system). After that, the aircraft, even with the engine turned off, can perform a safe landing on one engine at the airport of departure (if it is a twin-engine aircraft of the Tu-204 or MS-21 type) or continue the flight (if it is a four-engine aircraft of the IL-96-400 type).

This increases the reliability of the engine as an element of the aircraft SU and the safety of the aircraft itself.

Claims (1)

A method of controlling fuel consumption in a gas turbine engine, which consists in measuring the position of the engine control lever (ORE), the speed of the turbocharger and a free turbine (ST), the pressure and air temperature at the engine inlet, the temperature of the gases behind the gas generator turbine, form a predetermined frequency value the rotation of the turbocharger as a function of the position of the throttle, pressure and air temperature at the engine inlet, set the gas temperature behind the gas turbine and the rotation speed C , compare the set value of the turbocharger speed and the measured one, compare the limit value for the engine gas temperature behind the gas generator turbine and the measured one, compare the limit value for the given engine speed ST and the measured one, the received discrepancies are selected to a minimum with the signal of the pick-up circuit (AP), selected the value is fed to the proportional-integral (PI) controller, where a control action is formed on the fuel flow meter dispenser supplied to the chamber combustion engine (CS), characterized in that in addition to the signal “Fire in the engine nacelle” generated by the aircraft’s fire alarm system during take-off, the current value of the turbocharger speed is recorded and used as the set value of the turbocharger speed for a predetermined time after which the fuel supply to the compressor station is stopped and the engine is turned off.

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