RU2247845C2 - On-board aircraft engine monitoring system at limitation of rotational speed, temperature, fuel parameters and thrust - Google Patents

Abstract

FIELD: aircraft instrumentation engineering; on-board monitoring of gas-turbine engines.

SUBSTANCE: proposed system employs information received from aircraft parameter sensors and transmitted to multiplexing unit inputs and from output of this unit to on-board computer. Information received from sensors of very important parameters of engine is transmitted directly to command unit input. Measured present magnitudes are compared with specified limiting and dangerous magnitudes in on-board computer and in command unit on basis of algorithms introduced in their storage and information pertaining to dangerous parameters of aircraft engine: rotational speed of high-pressure compressor rotor, turbine outlet temperature, pressure of gas after fan doors and thrust. In case of deviation of these magnitudes from specified limits, limiting and dangerous commands are formed and are transmitted to on-board computer and to emergency information system for recording and warning the crew. Actual operating time and residual service life with commands thus obtained taken into account are determined in on-board computer and are transmitted to on-board information system for indication and recording.

EFFECT: possibility of monitoring the engine state in standard, off-standard and augmented power rating conditions.

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Description

The present invention relates to instrumentation and can be used to control an aircraft engine, mainly a gas turbine (GTE).

Known on-board aircraft engine control system containing sensors of the fuel parameters of the gas turbine engine: pressure P _fuel and fuel consumption G _fuel supplied to the engine, the outputs of which are connected to the inputs of the engine parameter conversion unit. [Aircraft engine control system. Application 2272783, UK, MKI ⁵ F 02 C 9/46; Rolls-Royce plc., No. 9224330.2, publ. 05/25/94].

The disadvantage of this system is the poor efficiency of aircraft engine control. This drawback is caused by the fact that, for the control of the gas turbine engine, firstly, a rather narrow group of controlled parameters of the aircraft engine is used - fuel parameters, and secondly, the fact that the control of the gas turbine engine does not take into account events when the controlled parameters of the aircraft engine exceed the limits set for them.

This disadvantage is deprived of the well-known aircraft engine control system. [Aviation GTE control system. Application 2262623, UK, MKI ⁵ F 02 C 9/26; Rolls-Royce rlc., No. 9126781, publ. 06/23/93].

The known system contains, in addition to the sensors of the fuel parameters of the aircraft engine, also sensors of non-fuel parameters: rotational speeds n _in and n _kW, respectively, of the rotors of the fan and compressor of high pressure, gas temperature t * _t and t * _k , respectively, behind the turbine and behind the compressor, angle α of the _ore position engine control knobs, etc., the outputs of which are connected to the inputs of the on-board computer containing the module of limit values, which allows taking into account the effect on the actual operating time of the aircraft engine of events in the output of individual controlled parameters of the aircraft engine beyond the limit values.

Limit values limit the allowable changes in the parameters of the aircraft engine in the normal mode of operation. When the aircraft engine is operating in an emergency mode, the values of individual parameters of the aircraft engine can go beyond the limits of the limiting values, without reaching the dangerous values set for these parameters.

The identification of events when the current values of the controlled parameters of the aircraft engine go beyond the limit values is important, since the consequence of such events is a significant increase in wear and tear and an increase in the actual operating time of the aircraft engine. Therefore, to assess the real technical condition of the aircraft engine during its emergency operation, it is necessary to take into account not the actually measured operating time (operating time) of the aircraft engine, but the effective operating time (effective operating time) of the aircraft engine in an emergency mode, i.e. time increased compared to the measured value. At the same time, the actual operating time of an aircraft engine operating in both standard and emergency modes should be determined as the sum of the operating time and effective operating time of the aircraft engine.

However, the known system solves the problem of determining the effective operating time very approximately, because in her calculator, the current values of the controlled parameters of the aircraft engine are compared with constant values of the limit values stored in the memory of the module of limit values. Moreover, the known system does not take into account the fact that the limiting values themselves are functions of the current values of the parameters of the aircraft engine, i.e. not constant, but “floating” values, and therefore, an accurate determination of the actual operating time of the aircraft engine during its emergency operation should be based on monitoring the parameters of the aircraft engine over floating limits. However, such control is not provided by the known system.

The closest to the proposed and adopted as a prototype onboard control system for the PS-90A aircraft gas turbine engine [V.A. Pivovarov. Diagnostics of aircraft and aircraft engines. M., MSTUGA, 1995, pp. 141-144].

The known system comprises an on-board computer with a processor, as well as a multiplexing unit, designed to normalize and multiplex the signals of the sensors of the controlled parameters of the aircraft engine, the inputs of which are connected to the outputs of the sensors of the controlled parameters of the aircraft engine, and the output is connected to the parametric input of the on-board computer with a processor, and the output of the on-board computer with the processor is designed to provide information to the on-board information system.

The known system provides control of the technical condition of the aircraft engine during its operation in the normal and emergency modes, calculates the values of the effective and actual operating time of the aircraft engine and transfers them to the on-board information system for registration and indication on call.

A disadvantage of the known system is a biased assessment of the actual operating time and technical condition of the aircraft engine when it is operating in forced mode, such as, for example, the interrupted take-off mode of a twin-engine aircraft.

The interrupted take-off mode may occur if one of the engines of a take-off aircraft equipped with two engines fails. Because termination of take-off in such a case can lead to a flight accident and is prohibited by flight regulations, the aircraft must continue to take off on one engine operating in a forced mode with increased thrust, then perform pre-landing maneuvers and then land. However, if during operation of the aircraft engine in forced mode at least one of the controlled parameters of the aircraft engine goes beyond the dangerous value set for this parameter, further flight operation of the aircraft engine is not allowed.

In order to maintain the airworthiness of the aircraft engine after it operates in forced mode with exceeding the dangerous value of the controlled parameter and to ensure the possibility of further continuation of flights using the aforementioned aircraft engine, flight regulatory documents allow us to consider the forced mode not as an emergency mode, but as one of the allowed short-term modes of operation of the aircraft engine when subject to the mandatory implementation of regulatory requirements to limit the time of the aircraft engine forcing annom mode, and also to limit the service life of an aircraft engine after its operation yield one or more actual values of monitored parameters beyond the hazardous quantities.

Therefore, in order to maintain the airworthiness of the aircraft engine, determine its effective operating time in the forced mode and residual motor resource, the technical condition of the aircraft engine during its operation in the forced mode should be monitored taking into account the parametric and time limitations provided for by this flight regulations, which is not provided by the known system and, therefore, does not allow the use of an aircraft engine that provided an interrupted take-off of the aircraft flight operation.

The objective of the invention is the provision of effective on-board control, calculation of operating time in the forced mode and the residual engine life of the aircraft engine, designed to work in the forced mode with increased thrust.

The problem is solved in that the onboard engine control system contains an onboard computer with a processor and a multiplexing unit, the inputs of which are connected to the outputs of the sensors of the aircraft engine's monitored parameters, and the output is connected to the parametric input of the onboard computer. According to the invention, what’s new is that it additionally contains a command unit, which includes a controller, limit settings module and hazardous settings module, command unit parametric inputs are connected to the outputs of some sensors of controlled parameters of the aircraft engine, the controller is connected by a bi-directional information connection to the processor the on-board computer, the limit settings module is connected to one of the controller inputs, and the hazardous settings module, on the on-board computer, is connected to the additional in addition, a timer, a module of limit algorithms, and a module of hazardous algorithms are introduced, each of which is connected to the processor inputs, and the hazardous algorithm module includes a microprocessor and hazardous algorithm cells connected to its inputs: high-speed compressor rotor speed, gas temperature behind the turbine, fuel flow and pressure, traction, - the hazardous settings module includes a microcontroller and hazardous settings micromodules connected to its inputs: the compressor rotor speed is high of pressure for turbine gas temperature, flow and pressure of the fuel rods - and avionic computer equipped with an input for receiving a signal of the forced mode aircraft engine.

When controlling an aircraft engine in the command block of the proposed system, the current values of hazardous parameters are compared with the current values of hazardous quantities, events of the exit of hazardous parameters beyond the boundaries of hazardous quantities are detected, the corresponding hazardous teams are generated and taking into account the presence of these commands and the engine’s operating time entered in the timer for forced operating mode in the on-board computer determines the operating time in the forced mode and the residual engine life of the aircraft engine when it is operating in the forced mode mode.

For a more complete disclosure of the invention, figure 1 presents a functional diagram of the claimed system, and figure 2 is a functional diagram of its two modules.

The aircraft engine control system contains a multiplexing unit 1, an on-board computer 2, which includes a processor 3, a limit algorithm module 4, a timer 5 and a dangerous algorithm module 6, a command unit 7, which includes a controller 8, a limit setting module 9, and a module 10 dangerous settings.

The on-board monitoring system interacts with the on-board information system, which includes a comprehensive indicator 11 and the flight data recorder 12, as well as with the aircraft emergency information system, which includes an alarm board 13 and a protected on-board information storage 14.

Module 6 of hazardous algorithms comprises a microprocessor 15, cell 16 of hazardous algorithms for the rotor speed of a high pressure compressor, cell 17 of hazardous gas temperature algorithms behind the turbine, cell 18 of hazardous fuel flow and pressure algorithms, and cell 19 of hazardous traction algorithms.

The dangerous settings module 10 contains a microcontroller 20, a micromodule 21 dangerous settings of the rotor speed of the high pressure compressor, a micromodule 22 dangerous settings of the gas temperature behind the turbine, a micromodule 23 dangerous settings of the fuel flow and pressure, and a micromodule 24 dangerous settings of the thrust.

In the description of the invention and in the figures the following notation:

Bx Q _m - input signal quantity of oil;

Вх n _в - input of the rotor rotor speed signal;

Bx n _kvd - signal input of the rotational speed of the rotor of the high-pressure compressor;

Вх t * _т - gas temperature signal input behind the turbine;

Bx α _ores - input signal position of the engine control handle;

Вх Р * _t / Р * _вх - the input of the ratio of the pressure signal characterizing the thrust of the aircraft engine;

Controlled parameters:

Q _m - the amount of oil;

n _in - the frequency of rotation of the rotor of the fan;

n _KVD - rotor speed of the high pressure compressor;

t * _t is the gas temperature behind the turbine;

α _ores - the angle of the engine control handle;

t * _k - gas temperature behind the compressor;

t * _in - the air temperature at the inlet to the engine;

G _fuel - fuel consumption;

p _fuel - fuel pressure;

P * _stv - gas pressure behind the fan flaps;

P * _t is the gas pressure behind the turbine;

P * _in - air pressure at the engine inlet;

P * _t / P * _in - the ratio of the above pressures characterizing the thrust of the aircraft engine;

[P * _shaft ] _min - lower limit of the dangerous gas pressure behind the fan flaps;

[G _fuel ] _min - lower limit of fuel consumption;

[P * _t / P * _in ] _min - lower limit of the dangerous pressure ratio characterizing the thrust of the aircraft engine;

τ is the current time;

[τ] is the allowed value of the aircraft engine operating time in forced mode;

n _{in max} (τ) is the limiting value of the rotor speed of the fan from (τ);

n _{KVD max} (τ) is the limit value of the rotational speed of the rotor of the high-pressure compressor from (τ);

P * _t / P * _{in min} (τ) is the dangerous value of the pressure ratio characterizing the thrust of the aircraft engine from (τ);

[P * _stem ] _min (τ) is the lower boundary of the dangerous gas pressure behind the fan flaps from (τ);

[G _fuel ] _min (τ) is the lower limit of fuel consumption from (τ);

[P * _t / P * _in ] _min (τ) is the lower boundary of the pressure ratio characterizing the thrust of the aircraft engine from (τ).

Inputs Bx Q _m, n _in Bx, Bx n _qd, Bx t * _t, ..., Bx α _ores mentioned above, the multiplexing unit 1 for receiving, multiplexing and normalize signals of current values of the controlled parameters Q _m, n _c , n _KVD , t * _t , ..., α of the aircraft engine _ores mentioned above are connected to the outputs of the respective sensors of the controlled parameters of the aircraft engine (the sensors are not shown in FIG. 1).

The output of the multiplexing unit 1 is connected to the parametric input of the on-board computer 2, which at the same time is the first input of the processor 3, which is part of this computer. Entrance Вх F.Р. on-board computer 2, designed to receive the signal “Forced mode” (F.R.), connected to the second input of the processor 3; the third input of the processor 3 is connected to the output of the timer 5. Module 4 limit algorithms and module 6 of the dangerous algorithms on-board computer 2 are connected each to the processor 3, respectively, the first and second bi-directional buses. The on-board computer 2 is connected to the command unit 7 by a third bi-directional bus connecting the processor 3 of the on-board computer 2 to the controller 8 of the command unit 7.

The output of the Output 1 of the on-board computer 2, which is simultaneously the output of the processor 3, is designed to provide information to the interacting on-board information system. The bus inputs of the microprocessor 15, which is part of module 6 of the dangerous algorithms of the calculator 2, each connected to the bus output of one of the cells 16, 17, 18 and 19 of the dangerous algorithms.

Parametric Bx n _entrances, n _qd Bx, Bx t * _t, ..., Bx * _t F / R * _Rin mentioned above, the command unit 7 designed to receive the current limit values of the parameters n and hazardous _in, n _qd, t _t , ..., P * _t / P * _{in the} aircraft engine mentioned above. These inputs are connected to the outputs of the sensors of the corresponding parameters of the aircraft engine (in Fig.1, the sensors are not shown).

The parametric inputs of the command unit 7 are simultaneously the parametric inputs of the controller 8, which is part of this unit. The remaining unidirectional inputs of controller 8 are connected one to the output of module 9 of the limit settings, the other to the output of module 10 of dangerous settings.

The output of the Output 2 of the command unit 7, simultaneously serving as the output of the controller 8, is designed to provide information to the aircraft emergency information system.

The inputs of the microcontroller 20, which is part of the module 10 of the dangerous settings of the command unit 7, each connected to the output of one of the micromodules 21, 22, 23 and 24 of the dangerous settings.

In preparing the on-board monitoring system for operation, the technical parameters of the aircraft engine, usually used to monitor its technical condition in various operating modes provided for by the flight operation regulations: standard, emergency and forced modes, are preliminarily analyzed, and a list of the minimum possible number of monitored parameters necessary and sufficient for effective control of the aircraft engine in all of the following modes:

For the monitored parameters (1) of the aircraft engine, a list of limit values is set, and a list of dangerous quantities is allocated from it. In addition, for the forced mode (F.R.), a list of operating time values for the forced mode of F.R. is also set:

Where

Δ τ ₁ , Δ τ ₂ - time settings, namely:

Δ τ ₁ - the value of the time delay information in the FR mode;

Δ τ ₂ - the value of the time for the extension of the issuance of information in the FR mode;

[τ] is the allowed value of the aircraft engine operating time in the FR mode;

[T _desig ] - the designated engine life of the aircraft engine.

From the parameters of the list (1), the limiting parameters are distinguished:

the current value of each of which can go beyond the limits of the current value of the corresponding limit value, but does not go beyond the limits of the current value of the dangerous value, and dangerous parameters are, in turn, allocated from the limit parameters:

the current value of each of which can go beyond the boundaries of the current value of the corresponding dangerous value.

When the aircraft engine is idle, time values (2) are entered into the memory of timer 5 of the on-board calculator, and mathematical expressions of limit algorithms are used in the memory of module 4 of the on-board calculator 2, designed to calculate the current values of the limit values, limiting the maximum permissible changes in the current values of the controlled parameters from above or below aircraft engine.

Expressions for limit values:

are polynomials of two types, depending on the operation mode of the aircraft engine and on the current values of the controlled parameters. Polynomials of the first kind are nonlinear binomials of the type

Where

as the parameters x and y, the parameters are selected from the above list of "Controlled parameters";

i is the serial number of a polynomial of the first kind;

z _{i max} - the current value of the limiting value in the function of the parameters x, the aircraft engine;

and _i1 is the dimensional coefficient;

f _i1 (x) and f _i2 (y) are the experimental dependences of the service functions f _i1 and f _i2 on the current values of the controlled parameters x and y of the aircraft engine in an abnormal mode of operation;

C _i1 is the additive constant (parametric setting) characterizing the abnormal mode of engine operation.

Polynomials of the second kind are linear trinomials of the type

Where

j is the serial number of a polynomial of the 2nd kind;

z _{j min} - the current value of the limit value in the function of the parameters x, the aircraft engine;

a _j1 , a _j2 - dimensional coefficients;

x and y are the current values of the controlled parameters of the aircraft engine (see above);

b _j1 , b _j2 - additive constants specifying the form of expression (7) in abnormal operation of the aircraft engine;

C _j1 - additive constant (parametric setting), characterizing the abnormal mode of operation of the aircraft engine.

Mathematical expressions of hazardous algorithms are introduced into the memory of cells 16, 17, 18, and 19 of module 6 of dangerous algorithms of the on-board calculator 2, designed to calculate the current values of hazardous quantities that limit dangerous changes from above or below the current values of the hazardous parameters (4) of the aircraft engine.

Expressions for dangerous quantities:

также представляют собой нелинейные и линейные многочлены.

also represent nonlinear and linear polynomials.

In the memory of cell 16 of module 6 of on-board calculator 2, an expression is introduced to calculate the upper boundary of the dangerous rotation speed [n _kW ] _{max of the} rotor of the high-pressure compressor

- меньшее из двух значений служебных функций

и

a

причем графики зависимостей

и

устанавливают экспериментально при стендовых испытаниях двигателя на форсированном режиме;Where

- the smaller of the two values of service functions

and

a

and dependency graphs

and

set experimentally during bench tests of the engine in forced mode;

and ₁₁ is a dimensional coefficient;

C ₁₁ - additive constant (parametric setting), characterizing the forced mode of engine operation.

In the memory of cell 18 of module 6 of on-board calculator 2, an expression is introduced to calculate the lower boundary of the dangerous pressure [P * _st ] _{min of} gas behind the fan flaps

where k ₁₁ and k ₁₂ are dimensional coefficients;

t * _in - the air temperature at the inlet to the engine;

P * _I - air pressure at the inlet to the engine,

and the values of the additive constants (parametric settings) b ₁₁ , b ₁₂ and b ₁₃ set according to the results of bench tests of the aircraft engine in forced mode.

In the memory of cell 19 of module 6 of on-board calculator 2, an expression is introduced to calculate the lower boundary of the dangerous pressure ratio [P * _t / P * _in ] _min characterizing the minimum allowable thrust of the aircraft engine

where b ₂₁ , b ₂₂ and b ₂₃ are additive constants (parametric settings);

α _ores - the angle of the engine control handle;

t * _inx - air temperature at the engine inlet;

k ₂₁ and k ₂₂ are dimensional coefficients,

moreover, the values of dimensional coefficients k ₂₁ and k ₂₂ are set according to the results of bench tests of an aircraft engine in forced mode, depending on the ratio of α _ores and t * _in , for example:

k ₂₁ = k ₂₂ = 0 at t * _in ≤ 15 ° C, α _ores ≥ 73 °;

k ₂₁ = 0.01; k ₂₂ = 0 at t * _Ix ≤ 15 ° C, 55 ° ≤ α _ores <73 °.

In addition, mathematical expressions of algorithms are introduced into the memory of the on-board calculator 2 to calculate the effective, accelerated, and actual operating time of the aircraft engine and residual engine life.

To calculate the effective T _eff operating time of the aircraft engine in an emergency mode, the sum of the form

- private effective operating time of the aircraft engine;

m is an integer equal to the number of partial sub-modes of the abnormal mode of the aircraft engine, differing in names or in the number of monitored parameters that have gone beyond the limit values;

τ _1j , τ _2j , respectively, is the start time and end time of the aircraft engine in a private sub-mode;

b _i is a constant characterizing the effect on the private effective operating time T _{eff m of the} event consisting in the output of the i-th parameter of the aircraft engine beyond the limit value;

n is the number of exit events characterizing the partial sub-mode of the aircraft engine.

For the calculation of time between the forced mode aircraft engine T _odds, the expression of type (13), in which instead of the magnitude _m T _eff is taken _odds value T _m, and

where a _i - coefficient characterizing the degree of tightening of private contingency

sub-modes of the aircraft engine when switching to forced mode.

Thus,

The actual operating time T the _{fact of the} aircraft engine is found as the sum of the operating time T, the effective operating time T _eff and the operating time in the forced mode T _{for the} aircraft engine:

and the residual motor resource [T] of the aircraft engine is defined as the difference between the assigned motor resource [T _desig ] and the actual operating time:

In the memory of module 9 of command unit 7, the numerical values of the parametric settings C _in , C _jm are entered, which are necessary for calculating the limit values in accordance with expressions (6) and (7), and in the memory of module 10, the numerical values of the parametric settings necessary for calculating dangerous quantities; in this case, setting C ₁₁ is entered into the memory of micromodule 21 of dangerous speed settings, settings b ₁₁ , b ₁₂ and b ₁₃ - into memory of micromodule 23 of dangerous pressure settings and settings b ₂₁ , b ₂₂ and b ₂₃ - into memory of micromodule 24 of dangerous thrust settings .

The values of the settings entered into the memory of modules 9 and 10 are transmitted to the controller 8 and, then, are transmitted via the third bi-directional bus to the processor 3 of the on-board computer 2. In the processor 3 of the on-board computer 2, the current values of the limit values are calculated:

and current values of hazardous quantities:

where τ is the current time, and are relayed via the third bi-directional bus to controller 8 of command unit 7.

When the aircraft engine is running, the inputs of the multiplexing unit 1 receive signals about the current values of the monitored parameters (1) of the aircraft engine, formed by the corresponding sensors; in the multiplexing unit 1, these signals are normalized, multiplexed, transmitted to the parametric input of the on-board computer 2 and, further, to the first input of the processor 3.

In processor 3, based on mathematical expressions (6), (7), which enter processor 3 via the first bi-directional bus from module 4 of limit algorithms, using the values of limit settings that enter processor 3 from module 9 of limit settings through controller 8 to the third bidirectional bus, the current values (18) of the limit values are calculated and transmitted from the processor 3 of the on-board calculator 2 to the controller 8 of the command unit 7 via the third bi-directional bus. In addition, in the processor 3 based on mathematical expressions (9), (10) and (11) entering the processor 3 via the second bidirectional bus from the hazardous algorithm module 6, using the values of the dangerous settings coming into the processor 3 from the hazardous settings module 10 through the controller 8 via the third bidirectional bus, are calculated current values (19) of dangerous quantities and are transmitted via a third bi-directional bus from processor 3 of on-board calculator 2 to controller 8 of command unit 7.

The signals about the current values of the limiting (3) and hazardous (4) parameters of the aircraft engine come from the outputs of the respective sensors directly to the inputs of the command unit 7 and then to the parametric inputs of the controller 8, in which, taking into account the information about the current values received from the processor 3 of the on-board computer 2 limit and dangerous values, the current values of the limit (3) and dangerous (4) parameters are compared with the current values of the limit (18) and dangerous (19) values in order to identify events of the output of the current values of OF DATA parameters beyond the current values of said values and the formation of marginal and dangerous commands.

When the aircraft engine is in normal mode, characterized by the absence of events when the controlled parameters of the aircraft engine exceed the limits of dangerous and dangerous values, the on-board computer 2 determines the operating time T of the aircraft engine in normal mode as the measured time of the aircraft engine in this mode and from output 1 of the on-board computer 2 value T is transmitted to an interactive airborne information system.

When the aircraft engine is operating in an emergency mode, in the event of occurrence of events when the controlled parameters of the aircraft engine exceed the limits and there are no events when the controlled parameters of the aircraft engine exceed the limits of dangerous quantities, limit commands are generated in the controller 8 of the command unit 7 and transmitted from the output of the Exit 2 of the command unit 7 to the aircraft emergency information system, as well as via the third bi-directional bus, to processor 3 of on-board computer 2. In on-board computer 2, taking into account the presence of limit commands It is calculated in accordance with expressions (13) and (12) the effective T _eff and in accordance with the expression (16) - the actual value of the operating time T _fact aircraft engine; the calculated values are transmitted from the output of Output 1 of the on-board computer 2 to the on-board information system.

When the aircraft engine is in forced mode, characterized by the presence of events when the current values of hazardous parameters go beyond the set limits, to the on-board computer 2 through its input Vkh F.R. and, further, a signal is received at the second input of processor 3 at time τ ₁ that the aircraft engine has switched to forced mode and is transmitted to controller 8 of command unit 7 via a third bi-directional bus. From timer 5, values of time settings Δ τ are transmitted to the third input of processor 3 ₁ and Δ τ ₂ are transmitted via the third bi-directional bus to controller 8 of command unit 7 and, in the presence of events when current values of hazardous parameters (4) go beyond the limits of current values of hazardous values (19), dangerous commands are generated in command block 7. Taking into account the time settings Δ τ ₁ and Δ τ _{2, the} issuance of the generated commands is delayed by the time Δ τ _{1 of the} delay in the issuance of information from the moment τ _{1 of} receiving the signal FR; after the time Δ τ _{1 of the} delay expires, dangerous commands from the output of Exit 2 of the command unit 7 are transmitted to the emergency information system of the aircraft for registration and notification of the crew, and also via the third bi-directional bus to the processor 3 of the on-board computer 2, and the transmission of dangerous commands continues over time Δ τ _{2 the} extension of information from the moment τ ₂ termination of the signal F.R., i.e. up to the point in time τ ₂ + Δ τ ₂ .

Delay for the time Δ τ ₁ and extension for the time Δ τ _{2 of the} issuance of dangerous commands are made, in connection with the special responsibility of these teams, to exclude the possibility of using false information about the forced mode of the aircraft engine. To this end, the stability of the presence of the F.R. signal is checked during the time Δ τ ₁ , and the stability of the removal of the F.R. signal during the time Δ τ ₂ to detect random outliers or malfunctions of this signal.

In avionic computer 2 using dangerous commands is calculated in accordance with expressions (14), (15) time between T _odds forced mode and in accordance with the expression (16) the actual operating time T _fact aircraft engine determined time τ _df = τ ₂ -τ ₁ operation of the aircraft engine in forced mode, is calculated in accordance with expression (17) using the value of [T _desig ] coming from the timer 5 to the third input of the processor 3, the residual motor resource [T] of the aircraft engine, taking into account the actual operating time T _fact . The calculated data is transmitted from the output of Output 1 of the on-board computer 2 to the on-board information system.

In addition, in the processor 3 of the on-board calculator 2, using the allowed time value [τ] coming from the timer 5 to the third input of the processor 3, the time τ _f.r. operation of the aircraft engine in forced mode with a permitted value of time [τ] and, in case of exceeding the last τ _f.r. > [τ], an excess command is generated. The generated command is transmitted via the third bi-directional bus to the controller 8 of the command unit 7 and from the output 2 of this unit is supplied to the aircraft emergency information system for registration and notification of the crew.

Thus, the proposed system allows to increase the efficiency of onboard control of the aircraft engine in forced mode and to use the aircraft engine after it operates in forced mode for further flight operation with limited engine life.

Claims (1)

The onboard engine control system, comprising an onboard computer with a processor, as well as a multiplexing unit, the inputs of which are connected to the outputs of the sensors of the aircraft engine's monitored parameters, and the output is connected to the parametric input of the onboard computer, characterized in that a command unit containing a controller and a limit settings module are additionally introduced and the module of dangerous settings, the parametric inputs of the command unit are connected to the outputs of the sensors of the controlled parameters of the aircraft engine, the controller is connected bi-directional information connection with the on-board computer processor, the limit settings module is connected to one of the controller inputs, and the dangerous settings module is connected to the additional one, a timer, the limit algorithm module and the dangerous algorithm module are added to the on-board computer, the output of each of which is connected to the inputs processor, and the module of dangerous algorithms includes a microprocessor and cells of dangerous algorithms connected to its inputs: compressor rotor speed high yes lignation, gas temperature behind the turbine, fuel flow and pressure, thrust, the hazardous settings module includes a microcontroller and hazardous settings micromodules connected to its inputs: rotational speed of the high-pressure compressor rotor, gas temperature behind the turbine, fuel consumption and pressure, and thrust, and the on-board computer is equipped with an input for receiving a signal about the forced mode of the aircraft engine.

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