RU2249119C2 - Aircraft engine monitoring method - Google Patents

Abstract

FIELD: aircraft industry.

SUBSTANCE: invention relates to aircraft instruments and it can be used for in-flight monitoring of aircraft engine, mainly, gas-turbine engine, using on-board computer and command unit. According method, limit and critical parameters are revealed from controllable parameters of aircraft engine and limit and critical values of said parameters are found, parameter settings are introduced into memory of on-board computer through command unit for calculating limit and critical values and also time of tolerable operation of aircraft engine at afterburning and algorithms of calculations of current values of limit and critical values as function of current values of parameters of aircraft engine and algorithms of calculation of effective, forced and actual running times and residual service life of aircraft engine, and current values of limit and critical values are calculated in on-board computer and results are delivered into command unit, current values of controllable parameters are supplied to input of on-board computer, and current values of limit and critical parameters of aircraft engine are supplied to input of command unit, and current values of said parameters are compared in command unit with current values of limit and critical values, and if parameter values comes out of limits of limit or critical values, corresponding commands are formed and transmitted into aircraft emergency system and into on-board computer, and said computer calculates actual values of running time and residual service life of aircraft engine using said commands and measured time of operation of engine at afterburning and transmits them into interacting on -board information systems, measured time of operation of aircraft engine at afterburning is compared with tolerable time and if tolerable time value is exceeded, corresponding command is formed which is transmitted through command unit into aircraft emergency system.

EFFECT: possibility of checking condition of aircraft engine at standard, emergency and afterburning operation.

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Description

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

There is a known method for controlling an aircraft gas turbine engine and an onboard engine control system that implements this method, based on the control of the fuel parameters of a gas turbine engine: pressure P _fuel and flow rate G _fuel fuel supplied to the engine. [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 method is the lack of effectiveness of the aircraft engine control. This drawback is caused by the fact that, for the control of gas turbine engines, firstly, they use a fairly narrow group of controlled parameters of the aircraft engine - fuel parameters, and secondly, because they do not take into account the events when the controlled parameters of the aircraft engine exceed the maximum limits established for them. This disadvantage is deprived of the well-known method of controlling an aircraft gas turbine engine and the onboard control system of an aircraft engine [Aircraft gas turbine engine control system. Application 2262623, UK, MKI ⁵ F 02 C 9/26; Rolls-Royce rlc., No. 9126781, publ. 06/23/93].

In the known method, carried out using an on-board computer, for controlling a gas turbine engine, firstly, in addition to the fuel parameters, non-fuel parameters of the aircraft engine are also used: rotational speeds n _in and n _kW , respectively, of the rotors of the fan and compressor of high pressure, gas temperature t * t * _t and _to, respectively, behind the turbine and the compressor, the angle α _ores motor position control handle et al., and secondly, to take into account the effect on the actual operating time of an aircraft engine events, consisting in the separate control of the outlet Rui aircraft engine parameters beyond their assigned 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 method solves the problem of determining the effective operating time very approximately, because based on a comparison of the current values of the controlled parameters of the aircraft engine with constant values of limit values. This does not take into account the fact that the limiting values are themselves functions of the current values of the parameters of the aircraft engine, i.e. not constant, but “floating” values. In this regard, 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, which is not provided in a known manner.

The closest to the proposed and adopted as a prototype method for controlling an aircraft engine, implemented in the on-board monitoring system of an aircraft gas turbine engine PS-90A [V.A. Pivovarov, is free from this drawback. Diagnostics of aircraft and aircraft engines. M., MSTUGA, 1995, p. 141-144].

When implementing this method, an on-board computer is used to determine the controlled parameters of the aircraft engine, establish the limit values and determine their functional dependencies, extract the limit parameters of the aircraft engine from the controlled parameters of the aircraft engine, and introduce mathematical expressions of algorithms to calculate the limit values of the effective and actual operating time of the aircraft engine.

As a result of the implementation of this method, the determination of the actual operating time of the aircraft engine, obtained by summing the operating time and effective operating time of the aircraft engine, is made quite accurately.

The disadvantage of this method 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 state 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 normative documents, which is not provided in a known manner , and, therefore, does not allow the use of an aircraft engine, providing an interrupted take-off of the aircraft, for further 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.

To solve the problem in the method of controlling an aircraft engine using an on-board computer, in which the controlled parameters of the aircraft engine are determined, limit values are determined and their functional dependencies are determined, the maximum parameters of the aircraft engine are extracted from the controlled parameters of the aircraft engine, mathematical expressions of algorithms are introduced into the on-board computer to calculate the maximum values of the effective and the actual operating time of the aircraft engine, new, according to the invention, is that d they additionally use a command unit connected by bidirectional information communication with the on-board computer, using the command unit enter the numerical values of the parametric settings necessary for calculating the limit values, translate these values into the on-board computer, calculate the current values of the limit values in the on-board computer and transfer them to the command unit , receive, normalize, multiply and transmit them to the command unit, receive, normalize, multiply and transmit to the on-board computer signals about the current values of the controlled parameters of the aircraft engine, receive and transmit signals to the command block about the current values of the limit parameters of the aircraft engine, in the command block compare the current values of the limit parameters with the current values of their corresponding limit values, in case the current value of the limit parameter exceeds the current value of the corresponding form a limit command and transfer it to the aircraft emergency information system and to the on-board computer , in the on-board computer using the adopted limit commands, the values of the effective and actual operating time of the aircraft engine are calculated and transmitted to the on-board information systems, the dangerous values of the controlled parameters of the aircraft engine are determined and the functional dependencies of the dangerous values are determined, the dangerous parameters are extracted from the limit parameters, the current value of each of which go beyond the boundaries of the corresponding dangerous value, enter into the on-board computer the numerical values of the time limits to, determining the output of information during the operation of the aircraft engine in forced mode, as well as the numerical values of the allowed operating time of the aircraft engine in forced mode and the designated engine life of the aircraft engine, additionally introduce mathematical expressions of algorithms for calculating dangerous values, operating time of the aircraft engine in forced mode and the residual engine life of the aircraft engine, in command the unit additionally enter the numerical values of the dangerous parametric settings necessary for calculating the dangerous values, and these values are transferred to the on-board computer, the current values of dangerous quantities are calculated in the on-board computer and transferred to the command unit, a signal is sent to the on-board computer that the aircraft engine is in forced mode and transmitted to the command unit, in the command unit the current values of dangerous parameters are compared with the current dangerous values, in case the current value of the dangerous parameter goes beyond the boundaries of the current value of the corresponding dangerous value, a dangerous command is generated, after a time information delivery holders after giving a signal about the transition of the aircraft engine to the forced mode transmit the generated dangerous commands to the aircraft emergency information system and to the on-board computer until the signal about the transition of the aircraft engine to the forced mode stops and, in addition, during the extension of the issuance of information after its termination , in the on-board computer, the operating time of the aircraft engine in forced mode is calculated taking into account the dangerous commands received and the operating time of the aircraft engine in forced mode e, specify the value of the actual operating time of the aircraft engine, taking into account the operating time of the aircraft engine in forced mode, determine the value of the residual engine life of the aircraft engine, transfer the mentioned values to the on-board information systems, compare the duration of the signal about the transition of the aircraft engine to the forced mode with the permitted time value and, if the latter is exceeded, form an excess command and transmit it through the command unit to the aircraft emergency information system.

For a more complete disclosure of the invention, figure 1 presents a functional diagram of an onboard control system for an aircraft engine implementing the claimed method, and figure 2 is a functional diagram of two modules of this system.

The onboard engine engine control system comprises a multiplexing unit 1, an onboard computer 2, which includes a processor 3, a limit algorithm module 4, a timer 5, and a hazardous algorithm module 6 containing a microprocessor 7, a cell 8 of hazardous high-pressure compressor rotor speed algorithms, cell 9 dangerous algorithms for gas temperature behind the turbine, cell 10 dangerous algorithms for fuel flow and pressure, cell 11 dangerous algorithms for gas pressure behind the fan flaps and cell 12 dangerous algorithms for traction, command unit 13, which includes a controller 14, a module 15 of the limit settings and a module 16 of dangerous settings, containing a microcontroller 17, a micromodule 18 of the dangerous settings of the rotor speed of the high pressure compressor, a micromodule of 19 dangerous settings of the gas temperature behind the turbine, a micromodule of 20 dangerous settings of the flow and fuel pressure, micromodule 21 dangerous gas pressure settings behind the fan flaps and micromodule 22 dangerous draft settings.

In the description of the invention and in the drawings, 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 * _inx - air temperature at the engine inlet;

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 - the 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 a dangerous value of the pressure ratio characterizing the thrust of the aircraft engine from (τ);

[P * _shaft ] _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).

1 multiplexing unit inputs the I Q _m, n _in Bx, Bx n _qd ,, _t t * Bx, ..., Bx α _ores mentioned above, intended for receiving and multiplexing signals normalizing the current values of monitored 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 calculator 2 is connected to the block 13 by a third bi-directional bus connecting the processor 3 of the calculator 2 with the controller 14 of the block 13.

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 systems. The bus inputs of microprocessor 7, which is part of module 6 of dangerous algorithms of calculator 2, are each connected to the bus output of one of the cells 8, 9, 10, 11 and 12 of dangerous algorithms.

Inputs _to the I n, n _qd Bx, Bx t * _t, ..., Bx * _t F / R * _Rin mentioned above, the command block 13 are designed to receive the current limit values of the parameters n and hazardous _in, n _qd, t * _t , ..., P * _t / P * _in 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 mentioned inputs of the command block 13 are simultaneously parametric inputs of the controller 14, which is part of this block. The remaining inputs of the controller 14 are connected one to the output of the module 15 limit at the rates, the other to the output of the module 16 of the dangerous rates.

The output of the Output 2 of the command unit 13, simultaneously serving as the output of the controller 14, is intended for the issuance of information in the emergency information system of the aircraft.

The inputs of the microcontroller 17, which is part of the module 16 of the dangerous settings of the command unit 13, each connected to the output of one of the micromodules 18, 19, 20, 21 and 22 of the dangerous settings.

The following is an example implementation of the inventive method of controlling an aircraft engine.

Example.

Consider an example of the implementation of the claimed method using an onboard engine control system, newly developed for the PS90-A engine of the Tu-214 aircraft.

In this method, the technical parameters of the aircraft engine, usually used to control its technical condition in various operating modes stipulated by the flight operation regulations: regular, emergency and forced modes, are preliminarily analyzed, and a list of the minimum possible number of controlled parameters necessary and sufficient for effective control of the aircraft engine is formed:

For the monitored parameters (1) of the aircraft engine, a list of limit values is established, and a list of dangerous quantities is isolated from it. In addition, for the forced mode (F.R.), a list of time values is also established: the time value Δ τ _{1 of the} delay, the time value Δ τ _{2 of the} extension of the output of information in the F.R. mode, the allowed value [τ] of the aircraft engine operation time on the forced the mode assigned motor resource [T _desig ] of the aircraft engine and enter the set time values into the memory of the timer 5 of the on-board computer 2.

From the parameters of the list (1), limit 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 a dangerous value, and dangerous parameters, in turn, are allocated from the limit parameters, for example:

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, mathematical expressions of limit algorithms are introduced into the memory of module 4 of the on-board calculator 2, designed to calculate the current values of the limit values, limiting the maximum or lower maximum permissible changes in the current values of the controlled parameters of the 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;

a _i1 - 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 rate) 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;

and _j1 , and _j2 are 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 (6) in abnormal operation of the aircraft engine;

C _j1 is the additive constant (parametric rate) characterizing the abnormal mode of operation of the aircraft engine.

Mathematical expressions of dangerous algorithms are introduced into the memory of cells 8, 9, 10, 11, and 12 of module 6 of the on-board calculator 2, designed to calculate the current values of dangerous quantities, limiting dangerous changes from the top or bottom of the current values of the controlled parameters of the aircraft engine.

Expressions for Dangerous Values

also represent nonlinear and linear polynomials.

In the memory of cell 8 of module 6 of calculator 2, an expression is _entered 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

, а

, причем графики зависимостей

,

и

устанавливают экспериментально при стендовых испытаниях двигателя на форсированном режиме,Where

- the smaller of the two values of service functions

, 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 rate), characterizing the forced mode of engine operation.

In the memory of cell 9 of module 6 of calculator 2, an expression is introduced to calculate the upper boundary of the dangerous temperature [t * _t ] _max gas behind the turbine

- меньшее из двух значений служебных функций

и

, а

, причем графики зависимостей

,

устанавливают экспериментально при стендовых испытаниях двигателя на форсированном режиме,Where

- the smaller of the two values of service functions

and

, a

, and dependency graphs

,

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

and ₂₁ is a dimensional coefficient;

C ₂₁ - additive constant (parametric setting), characterizing the forced mode of engine operation.

An expression for calculating the lower boundary of the dangerous pressure [P * _st ] _{min of} gas behind the fan flaps is introduced into the memory of cell 11 of module 6 of the on-board computer 2

where k ₁₁ and k ₁₂ are dimensional coefficients;

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

P * _in - the air pressure at the engine inlet, and the values of additive constants (parametric settings) b ₁₁ , b ₁₂ and b ₁₃ are set according to the results of bench tests of the aircraft engine in forced mode.

An expression for calculating the lower boundary of the dangerous pressure ratio, [P * _t / P * _in ] _min , characterizing the minimum allowable thrust of the aircraft engine, is introduced into the memory of cell 12 of module 6 of the on-board calculator 2

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

α _ores - the angle of the engine control handle;

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

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 * _bx ≤ 15 ° С, α _ores > 73 °;

k ₂₁ = 0.01; k ₂₂ = 0 at t * _bx ≤ 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 the residual engine life.

To calculate the effective T _eff operating time of the aircraft engine in an emergency mode, use 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 partial 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.

To calculate the operating time in the forced mode of the aircraft engine T _for use an expression of the type (13), in which instead of the value of T _{eff m} take the value of T _{for m} , and

where a _i is a coefficient characterizing the degree of tightening of the private emergency sub-mode of the aircraft engine during the transition to the forced mode.

Thus,

The actual operating time T _fact found as the sum of the aircraft engine operating time T, the effective operating time T _Effi forced _odds aircraft engine operating time T:

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 15 of command unit 13, numerical values of the parametric settings C _in , C _{jm are entered} , which are necessary for calculating the limit values in accordance with expressions (5) and (6), and in the memory of module 16, the numerical values of the parametric settings necessary for calculating dangerous quantities; at the same time, at heading C _11, 18 dangerous speed settings are entered into the memory of the micromodule; at heading C ₂₁ , into the memory of the micromodule 19, dangerous temperature settings, settings b ₁₁ , ₁₂ and b ₁₃ are entered into the memory of the micromodule 21 dangerous pressure settings and b ₂₁ , b ₂₂ and b ₂₃ - in the memory of the micromodule 22 dangerous draft settings.

The values of the settings entered into the memory of modules 15 and 16 are transmitted to the controller 14 and then transmitted through 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 relay them on the third bi-directional bus to the controller 14 of the command unit 13.

When the aircraft engine is operating, the signals about the current values of the monitored parameters (1) of the aircraft engine, generated by the corresponding sensors, are fed to the inputs of the multiplexing unit 1, normalize, multiplex and transmit the signals of the sensors to the parametric input of the on-board computer 2 and then to the first input of the processor 3.

In the processor 3, based on mathematical expressions (5), (6), which are transmitted to the processor 3 via the first bi-directional bus from the limit algorithm module 4, using the limit settings that are transmitted to the processor 3 from the limit limit module 15 through the controller 14 through the third bidirectional bus, calculate the current values (18) of the limit values and transfer them from the processor 3 of the on-board calculator 2 to the controller 14 of the command unit 13 on the third bi-directional bus, In addition, in the processor 3 based on mathematical expressions ( 8), (9), (10), (11), which are transferred to 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 16 through the controller 14 according to the third bidirectional the bus, calculate the current values (19) of dangerous quantities and on the third bi-directional bus transfer them from the processor 3 of the on-board calculator 2 to the controller 14 of the command unit 13.

The signals about the current values of the limiting (2) and hazardous (3) parameters of the aircraft engine are fed directly to the inputs of the command unit 13, and then to the parametric inputs of the controller 14, in which, taking into account the information received from the processor 3 of the on-board computer 2 about the current values of the limiting and dangerous values, compare the current values of the limit (2) and dangerous (3) parameters, respectively, with the current values of the limit (18) and dangerous (19) values in order to identify events when the current values of the controlled parameters go beyond the boundaries of the tech values of the mentioned values and the formation of the respective teams.

When the aircraft engine is operating in the 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 operating time T of the aircraft engine in the normal mode is determined as the measured time of the aircraft engine in this mode and the operating time is transmitted to the interacting on-board information systems for registration.

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 14 of the command unit 13 and transmit them from the output of the Exit 2 of the command unit 13 to the aircraft emergency information system, and also 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 is calculated in accordance with expressions (13) and (12) the effective T _eff and in accordance with the expression (16) - the value of the actual T _fact developments aircraft engine and transmitting the computed values of the output O 1 onboard computer 2 to interact onboard information system for registration.

When the aircraft engine is in forced mode, characterized by the presence of events when dangerous parameters go beyond the established boundaries, fed into the on-board computer 2 through its input Vh F.R. and then, to the second input of processor 3, at a time τ ₁ , a signal is received about the aircraft engine switching to forced mode, and this signal is transmitted to the controller 14 of command unit 13 via a third bidirectional bus. From timer 5, temporal rates Δ τ are supplied to the third input of processor 3 ₁ and Δ τ ₂ , transfer them via the third bi-directional bus to the controller 14 of the command unit 13 and, if there are events when the current values of the hazardous parameters (3) go beyond the boundaries of the current values of the dangerous quantities (19), dangerous commands are generated in the command block 13, I delay t the issuance of the generated commands for a time Δ τ ₁ delay the issuance of information from the moment τ ₁ receiving signal F.R. after the time Δ τ _{1, the} delay sends dangerous commands from the output of Exit 2 of the command unit 13 to the aircraft emergency information system 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 for time Δ τ ₂ information renewal from the moment τ ₂ termination signal F.R., i.e. up to the point in time τ ₂ + Δ τ ₂ .

The delay for the time Δ τ ₁ and the extension for the time Δ τ _{2 of the} issuance of dangerous commands is performed to exclude the possibility of using false information about the forced mode of the aircraft engine. To this end, during the time Δ τ ₁ check the stability of the presence of the signal F.R., and during the time Δ τ ₂ - stability of the removal of the signal F.R. to detect random outliers or malfunctions of this signal.

In the on-board computer 2, using dangerous commands, the forced _{Tfor is} calculated in accordance with expressions (14), (15) and, in accordance with expression (16), the actual T _{fact of the} aircraft engine operating time is determined by the time τ _F.R. = τ ₂ -τ ₁ operation of the aircraft engine in forced mode, calculated in accordance with expression (17) using the value of [T _desig ], which is transferred 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 The fact and transmit the calculated data from the output of the Output 1 of the on-board computer 2 in the interacting on-board information systems.

In addition, in the processor 3 of the on-board calculator 2, using the allowed time value [τ], which is transmitted from the timer 5 to the third input of the processor 3, the time τ _{fp is compared.} 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], the excess command is generated, it is sent via the third bi-directional bus to the controller 14 of the command unit 13, and from the output 2 of this unit, the generated command is transmitted to the aircraft emergency information system for registration and notification of the crew.

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

Claims (1)

A method of controlling an aircraft engine using an onboard computer, in which the controlled parameters of the aircraft engine are determined, limit values are determined and their functional dependencies are determined, the maximum parameters of the aircraft engine are extracted from the controlled parameters of the aircraft engine, mathematical expressions of algorithms are entered into the onboard computer to calculate the maximum values of the effective and actual operating time of the aircraft engine, characterized in that they additionally use a command unit connected by two direct communication with the on-board computer, using the command unit enter the numerical values of the parametric settings necessary for calculating the limit values, translate these values into the on-board computer, calculate the current values of the limit values in the on-board computer and transfer them to the command unit, receive, normalize, multiply and transmit signals to the on-board computer about the current values of the controlled parameters of the aircraft engine, receive and transmit signals to the command unit about the current values values of the aircraft engine, in the command unit, compare the current values of the limit parameters with the current values of the corresponding limit values, in the case of the current value of the limit parameter exceeding the current value of the corresponding limit value, form the limit command and transmit it to the emergency information system of the aircraft and to the onboard calculator, in the on-board calculator using the adopted limit commands calculate the values of the effective and actual operating time of the aircraft the engine and transfer them to the on-board information systems, establish the dangerous values of the controlled parameters of the aircraft engine and determine the functional dependences of the dangerous values, extract the dangerous parameters from the limit parameters, the current value of each of which can go beyond the boundaries of the corresponding dangerous value, enter the numerical values of time settings that determine the issuance of information when the aircraft engine is in forced mode, as well as the numerical values of the allowed time nor the operation of the aircraft engine in forced mode and the designated engine life of the aircraft engine, they additionally introduce mathematical expressions of algorithms for calculating dangerous quantities, the operating time of the aircraft engine in forced mode and the residual engine life of the aircraft engine, additionally enter the numerical values of the dangerous parametric settings necessary for calculating dangerous values and transmit these values in the on-board computer, in the on-board computer calculate the current values of the dangerous quantities and transmit them in the command unit, a signal is sent to the on-board computer about the aircraft engine switching to forced mode and transmitted to the command unit, in the command unit, the current values of the hazardous parameters are compared with the current values of the hazardous values, if the current value of the hazardous parameter exceeds the current value of the corresponding dangerous the values form a dangerous command, after the delay time for issuing information after giving a signal about the transition of the aircraft engine to the forced mode, the generated dangerous to the emergency alarm system of the aircraft and to the on-board computer until the moment the signal about the transition of the aircraft engine to the forced mode ceases and, in addition, during the extension of the issuance of information after its termination in the on-board computer, the operating time of the aircraft engine in the forced mode is calculated taking into account the dangerous commands received and operating time of the aircraft engine in forced mode, specify the value of the actual operating time of the aircraft engine, taking into account the operating time of the aircraft engine in forced mode, determine the residual engine resource of the aircraft engine, the mentioned values are transmitted to the on-board information systems, the signal duration of the aircraft engine switching to the forced mode is compared with the allowed time value, and if the latter is exceeded, an excess command is generated and transmitted through the command unit to the aircraft emergency information system.

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