RU2249712C2 - Onboard monitoring system of engine at limited rotational speed, temperature and thrust - Google Patents

Abstract

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

SUBSTANCE: proposed monitoring system makes it possible to check aircraft engine working in standard, off-standard and augmented power rating modes. Monitoring is based on use of information received from sensors of aircraft engine parameters and transmitted to inputs of multiplexing unit and from output of this unit to onboard computer. Information received from sensors pertaining to very important parameters of aircraft engine is fed directly to command unit input. Information received from sensors of very important parameters of aircraft engine, viz.: rotational speed of high-pressure compressor rotor, turbine exhaust gas temperature and thrust are compared in onboard computer and in command unit on basis of computational algorithms introduced in their storage and received from sensors with limiting and dangerous magnitudes; in case these magnitudes are beyond limiting and dangerous magnitudes, limiting and dangerous commands are formed and are transmitted to onboard computer and to emergency information system of aircraft for recording and for warning the crew. Magnitudes of actual operating hours of aircraft engine and remaining service life are determined in onboard computer with the received commands taken into account and are transmitted to onboard information system for indication and recording.

EFFECT: enhanced efficiency of onboard monitoring of aircraft engine working at augmented power rating and increased thrust.

2 dwg

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 sensors for the fuel parameters of the aircraft engine, also sensors for 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 $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ and t $\genfrac{}{}{0ex}{}{\text{*}}{\text{to}}$ respectively, behind the turbine and behind the compressor, the angle α of the _{ores of the} position of the engine control handle, etc., the outputs of which are connected to the inputs of the on-board computer containing a module of limit values, allowing to take into account the effect on the actual operating time of the aircraft engine of events consisting in the output of individual controlled parameters of the aircraft engine beyond the limit quantities.

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, the command unit is additionally introduced into it, which includes the controller, the limit settings module and the dangerous settings module, the command unit parametric inputs are connected to the outputs of the sensors part of the 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 limit algorithm module, and a dangerous algorithm module 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-pressure compressor rotor speed, gas temperature behind the turbine, traction, - the hazardous settings module includes a microcontroller and micromodules of dangerous settings connected to its inputs: high-pressure compressor rotor speed, temperature ha behind the turbine, thrust, - and the on-board computer is equipped with an input for receiving a signal about the forced mode of the aircraft engine.

When controlling the aircraft engine in the command unit of the proposed system, the current values of the hazardous parameters are compared with the current values of the hazardous quantities, events of the exit of the hazardous parameters beyond the boundaries of the hazardous quantities are detected, the corresponding dangerous 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 algorithms for gas temperature behind a turbine, and cell 18 of hazardous traction algorithms.

The dangerous settings module 10 contains a microcontroller 19, a micromodule 20 dangerous settings of the rotor speed of the high-pressure compressor, a micromodule 21 dangerous settings of the gas temperature behind the turbine and a micromodule 22 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;

Bx t $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ - gas temperature signal input behind the turbine;

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

Bx R $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ /R $\genfrac{}{}{0ex}{}{\text{*}}{\text{in}}$ - 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 $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ - gas temperature behind the turbine;

α _ores - the angle of the engine control handle;

t $\genfrac{}{}{0ex}{}{\text{*}}{\text{to}}$ - gas temperature behind the compressor;

t $\genfrac{}{}{0ex}{}{\text{*}}{\text{in}}$ - air temperature at the engine inlet;

G _fuel - fuel consumption;

p _fuel - fuel pressure;

R $\genfrac{}{}{0ex}{}{\text{*}}{\text{st}}$ - gas pressure behind the fan flaps;

R $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ - gas pressure behind the turbine;

R $\genfrac{}{}{0ex}{}{\text{*}}{\text{in}}$ - air pressure at the engine inlet;

R $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ /R $\genfrac{}{}{0ex}{}{\text{*}}{\text{in}}$ - the ratio of the above pressures, characterizing the thrust of the aircraft engine;

[R $\genfrac{}{}{0ex}{}{\text{*}}{\text{st}}$ ] _min is the lower limit of the dangerous gas pressure behind the fan flaps;

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

[R $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ /R $\genfrac{}{}{0ex}{}{\text{*}}{\text{in}}$ ] _min is 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);

R $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ /R $\genfrac{}{}{0ex}{}{\text{*}}{\text{min min}}$ (τ) is the dangerous value of the pressure ratio characterizing the thrust of the aircraft engine from (τ);

[R $\genfrac{}{}{0ex}{}{\text{*}}{\text{st}}$ ] _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);

[R $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ /R $\genfrac{}{}{0ex}{}{\text{*}}{\text{in}}$ ] _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 $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ , ..., Bx α _ores mentioned above of the multiplexing unit 1, intended for receiving, normalizing and multiplexing signals about the current values of the monitored parameters Q _m , n _in , n _qdd , t $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ , ..., α _{ores of the} aircraft engine mentioned above are connected to the outputs of the respective sensors of the controlled parameters of the aircraft engine (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 on-board computer 2, are each connected to the bus output of one of the cells 16, 17, 18 of the dangerous algorithms.

Parametric Bx n _entrances, n _qd Bx, Bx t $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ , ..., Bx P $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ /R $\genfrac{}{}{0ex}{}{\text{*}}{\text{in}}$ mentioned above, command unit 7 are designed to receive the current values of the limit and hazardous parameters n _in , n _KVD , t $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ ,... , R $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ /R $\genfrac{}{}{0ex}{}{\text{*}}{\text{in}}$ 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 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 19, which is part of the module 10 of the dangerous settings of the command unit 7, are each connected to the output of one of the micromodules 20, 21, 22 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, the time values (2) are entered into the memory of timer 5 of the on-board calculator 2, and mathematical expressions of limit algorithms are used in the memory of module 4 of the limit algorithms of the on-board calculator 2, designed to calculate the current values of the limiting quantities that limit the maximum permissible changes of the current ones from above or below 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 parameters x and y, 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 limit value in the function of the parameters x, y of the aircraft engine;

a _i1 - dimensional coefficient;

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

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

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, y of 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.

In the memory of cells 16, 17, 18 of dangerous algorithms and microcontroller 19 of module 6 of dangerous algorithms of the on-board computer 2, mathematical expressions of dangerous algorithms are introduced to calculate the current values of dangerous quantities, limiting dangerous changes from above or below the current values of dangerous parameters (4) of the aircraft engine.

Expressions for dangerous quantities:

also represent nonlinear and linear polynomials.

In the memory of the cell 16 of the dangerous algorithms of the module 6 of the dangerous algorithms of the on-board computer 2, an expression is introduced to calculate the upper boundary of the dangerous speed [n _kvd ] _{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;

a 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 the dangerous algorithms of module 6 of the dangerous algorithms of the on-board computer 2, an expression is introduced to calculate the lower boundary of the dangerous pressure ratio [P $\genfrac{}{}{0ex}{}{\text{*}}{\text{t}}$ /R $\genfrac{}{}{0ex}{}{\text{*}}{\text{in}}$ ] _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 $\genfrac{}{}{0ex}{}{\text{*}}{\text{in}}$ - 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:

≤ 15° C, α _руд≥ 73° ;k ₂₁ = k ₂₂ = 0 for

≤ 15 ° C, α _ores ≥ 73 °;

≤ 15° C, 55° ≤ α _руд<73° .k ₂₁ = 0.01; k ₂₂ = 0 for

≤ 15 ° C, 55 ° ≤ α _ores <73 °.

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

Where

- 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, the start time and the end time of the aircraft engine in the private j-th 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.

To calculate the operating time in the forced mode of the T _for aircraft engine, an expression of the type (12) is used, in which instead of the value of T _{eff m} is taken

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 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 the limit settings of the command unit 7, the numerical values of the parametric settings C _in , C _jm are entered, which are necessary to calculate the limit values in accordance with expressions (6) and (7), and in the memory of the module 10 of dangerous settings, the numerical values of the parametric settings, necessary for calculating hazardous quantities; in this case, the setpoint C ₁₁ is entered into the memory of the micromodule 20 of dangerous settings, and the setpoint b ₂₁ is entered into the memory of the micromodule 22 of dangerous setpoints of thrust.

The values of the settings entered in the memory of module 9 of the limit settings and module 10 of dangerous settings are transmitted to the controller 8 and then 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 relayed on the third bi-directional bus to the controller 8 of the 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 then 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 (17) of the limit values are calculated, and the third bidirectional bus is transferred from the processor 3 of the on-board calculator 2 to the controller 8 of the command unit 7. In addition, in the processor 3 on the basis of mathematical expressions of those entering the processor 3 via the second bi-directional bus from the hazardous algorithm module 6, using the values of the dangerous settings coming to the processor 3 from the module 10 of the dangerous settings via the controller 8 via the third bi-directional bus, the current values (18) of the dangerous quantities are calculated and the third the bi-directional bus is transferred from the processor 3 of the on-board calculator 2 to the controller 8 of the 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 current information received from the processor 3 of the on-board computer 2 values of limit and dangerous values, the current values of limit (3) and dangerous (4) parameters are compared with the current values of limit (17) and dangerous (18) values in order to identify events of the output of current values of associated parameters beyond the current values of said quantities and limit the formation of dangerous and commands.

When the aircraft engine is operating 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 (12) and (11) the effective T _eff and in accordance with the expression (15) - 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 then, to the second input of the processor 3 at a time instant τ ₁ , a signal is received about the aircraft engine switching to the forced mode and transmitted via the third bi-directional bus to the controller 8 of the command unit 7. From the timer 5, the values of the time settings Δ τ ₁ and Δ are transmitted to the third input of the processor 3 τ ₂ , are transmitted via the third bi-directional bus to the controller 8 of the command unit 7 and, in the presence of events when the current values of the hazardous parameters (4) go beyond the boundaries of the current values of the dangerous values (18), dangerous commands are generated in the 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 .

The delay for the time Δ τ ₁ and the 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. For this purpose, 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 (13), (14) time between T _odds forced mode and in accordance with the expression (15) the actual operating time T _fact aircraft engine determined time τ _df = τ ₂ -τ ₁ operation of the aircraft engine in forced mode, calculated in accordance with expression (16) using the value of [T _desig ] coming from timer 5 to the third input of 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 _{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], 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 pressure, gas temperature behind the turbine, thrust, the hazardous settings module includes a microcontroller and hazardous settings micromodules connected to its inputs: high-pressure compressor rotor speed, gas temperature behind the turbine, thrust, and the on-board computer is equipped with an input for receiving a forced mode signal aircraft engine.

Images