RU2556477C1 - Vibration diagnostic method of gas-turbine engines in operation as per information of onboard devices - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: method is applied for monitoring of a vibration state of aircraft GTE both on ground processing devices and in real time in onboard engine control systems.

EFFECT: obtaining high reliability parameters of GTE vibration diagnostic results immediately under operating conditions by means of all-around consideration of factors determining vibration state of GTEs at their operation on the ground and under flight conditions.

6 cl, 1 dwg

Description

The invention relates to the field of monitoring the technical condition of aircraft gas turbine engines (GTE) equipped with standard measuring equipment, the signals from which during operation are also recorded by a standard on-board recording device (BUR) installed on board the corresponding aircraft (Aircraft). The method is used to monitor the vibrational state of aircraft gas turbine engines both on ground processing devices (NLD) and in real time in onboard engine control systems (BSCD).

A known method of vibration diagnostics of a gas turbine engine - ed. St. SU 1816986, class G01M 15/00, 2004. The method includes measuring and recording the values of the vibration signal and the rotational speed of the motor rotor in transient conditions, comparing the measured and reference values of the vibration signal for the characteristic speeds and determining the technical condition of the engine by deviating the measured value of the vibration signal from the reference, the measurement and registration of the values of the vibration signal and speed is carried out when adjusting the fuel equipment to the minimum and maximum excess fuel.

This method has a significant error and does not allow to reliably determine the actual vibrational state of the engine, has very narrow functional capabilities and cannot be used to determine the vibrational state of a gas turbine engine in flight conditions.

A known method for the diagnosis of gas turbine engines - patent RU 2297613, cl. G01M 15/14, 2007. The method includes measuring the vibration of a running engine, spectral analysis of vibration, and comparing the obtained data with the same values measured in the initial state of the engine. In this case, a spectral analysis of the envelope of the vibration signal emitted at characteristic frequencies is carried out, the amplitudes of the components of the obtained spectrum are measured in the range from zero to the rotational speed of the rotor having the highest rotation speed, the obtained values are compared with the same values in the initial state, about the place of the main sources of vibration change judged by the frequencies of the components having the largest deviations of the measured values from the original, and the localization of the defect is carried out by the vibration spectra and a wide frequency range by measuring and comparing the values of the modulation components of the vibration in the ranges of only those carriers whose frequencies are multiples of the fundamental frequency sources.

This method causes great difficulty in determining the reliable relationship between the causes of increased vibration and changes in certain spectral components. In addition, this method involves the installation of additional vibration measuring equipment on a gas turbine engine and requires the presence of experienced spectral analysis specialists in the operating organization who are capable of issuing competent decisions about the gas turbine engine vibration state.

A known method for diagnosing the technical condition of parts, assemblies and drive units of a gas turbine engine - patent RU 2379645, class. G01M 15/14, 2008. The method includes measuring and digitally processing vibration signals from the body structures of the gas turbine engine and drive units to obtain information about the technical condition of the diagnosed parts, components and drive units of the gas turbine engine. In this case, the measurement of vibration signals from the casing structures of the gas turbine engine and drive units is carried out remotely and non-contact by means of a laser vibration transducer at the informative points on the surface of the casing of the gas turbine engine and drive units close to the diagnosed parts, components and drive units of the gas turbine engine within the measurement zones determined by a radius of mainly equal to a quarter the length of the bending wave in the body structures of the gas turbine engine and drive units, and the digital processing of vibration signals is carried out with the calculation that the depth of modulation in the envelope of the discrete spectrum components in the high frequency vibration range of the oscillation hull structures CCD drive units and to obtain information about the mechanical condition diagnosed parts, assemblies of the drive units TBG.

This method is not suitable for use in operational conditions, because requires installation and fine-tuning of special vibration-measuring laser equipment. In addition, the use of such equipment to implement the method in flight conditions is practically difficult, and it also requires a lot of labor and the presence of narrow specialists to process information from laser sensors and interpret its results.

The closest analogue is the method of vibration diagnostics GTE - patent RU 2499240, class. G07M 15/00, 2013. The method includes obtaining a reference vibration characteristic for ground engine tests, obtaining a flight vibration characteristic, comparing the reference and flight vibration characteristics and determining the technical condition of the engine by deviating the flight vibration characteristic from the reference one, while obtaining the standard vibration characteristic is carried out by forming a basic vibration characteristic that is carried out by measuring and recording the values of the vibration signal at the operating frequencies of the rotor rotation during ground tests of the engine, as well as the formation of operational vibration characteristics, for which a series of flights is carried out, on each of the flights of the series, according to the values of the vibration signal at the operating rotational speeds of the rotor, a local operational vibration characteristic is formed, a threshold for deviation of local operational vibration characteristics from the base is set, each local vibration characteristic of a given series of experimental flights is obtained compared with the base and local vibration characteristics, the values of which do not go beyond the established threshold when compared to the basic characteristic is formed reference vibration characteristics.

This method forms a model of the standard vibrational state of a gas turbine engine on a test bench in the factory, which does not allow taking into account individual design features for installing a gas turbine engine on board an aircraft, since often the cause of increased vibration is a violation in the engine installation technology. In addition, the control of the vibrational state of a gas turbine engine is not entirely correct, because performed using various measuring equipment, which includes ground-based equipment and registration channels in the drill in flight aircraft.

The objective of the present invention is to achieve high reliability indicators of the results of vibration diagnostics of a gas turbine engine directly under operating conditions by comprehensively taking into account factors that determine the vibration state of a gas turbine engine when they work on the ground and in flight conditions.

The specified result is achieved by the fact that the method of vibration diagnostics of gas turbine engines in operation according to the information of the onboard registration devices includes recording flight information of the aircraft and parameters characterizing the vibration state of the gas turbine engines by the onboard registration device; reading registered information and parameters from the aircraft’s onboard registration device; identification of information and parameters read from the on-board recording device by entering passport and service data and formular data of an operating gas turbine engine; the formation of the first reference vibration characteristics of a gas turbine engine based on the results of engine operation on ground tests or tests on board an aircraft, on ground stages of aircraft movement from the moment the engine starts until the start of the take-off, and also from the moment the aircraft lands up until the engine shuts off; the formation of the second reference vibration characteristics of the gas turbine engine in flight at each moment of flight, taking into account the influence of linear and rotational forces and accelerations acting on the aircraft in flight, as well as the angular position of the aircraft in space; determination of the boundaries of the description of the first reference vibration characteristics when the engine is operating in ground conditions and threshold values that separate the vibration signal of the serviceable and malfunctioning state of a gas turbine engine; determination of the boundaries of the description of the second reference vibration characteristics during engine operation in flight conditions and threshold values that separate the vibration signal of the serviceable and faulty state of the gas turbine engine; monitoring the current state of the first vibration characteristic of a gas turbine engine by comparing it with the corresponding reference vibration characteristic; monitoring the current state of the second vibration characteristics of the gas turbine engine by comparing it with the corresponding reference vibration characteristics taking into account the state of the first vibration characteristics of the gas turbine engine; determination of the technical state of the gas turbine engine by the totality of deviations of the first and second vibration characteristics from the corresponding reference vibration characteristics at each current time of the gas turbine engine operation.

In this case, the formation of the first reference vibration characteristic of the gas turbine engine is carried out according to the adaptive principle, which allows the adjustment of the first reference vibration characteristic with each new operation mode of the gas turbine engine during ground tests on board the aircraft, provided that the gas turbine engine is in good condition and based on the selected component of the vibration signal from the high rotor speed pressure allocated by the component of the vibration signal from the rotor speed st pressure vibration component extracted from the gas temperature, the selected component of the vibration from the engine operation time.

The formation of the second reference vibration characteristic of the gas turbine engine is carried out according to the adaptive principle, which allows for the adjustment of the second reference vibration characteristic with each new flight mode of the aircraft, provided that the gas turbine engine is in good condition and based on the selected component of the vibration signal from the rotational speed of the high pressure rotor, extracted component of the vibration signal from the rotor speed low pressure, the selected component of the vibration signal from the temperature gases, the selected component of the vibration signal from the air pressure at the engine inlet, the selected component of the vibration signal from the engine time, the selected component of the vibration signal caused by the vertical overload of the aircraft, the selected component of the vibration signal caused by the longitudinal overload of the aircraft, the extracted component of the vibration caused by the lateral effect aircraft overload, the selected component of the vibration signal caused by the influence of angular velocity axes around the vertical axis of the aircraft, the selected component of the vibration signal caused by the influence of angular velocity around the transverse axis of the aircraft, the selected component of the vibration signal, introduced depending on the angular position relative to the transverse axis of the aircraft (pitch angle), the highlighted component of the vibration signal introduced in depending on the angular position relative to the longitudinal construction axis of the aircraft (roll angle), the Method has the ability ustomize - self-monitoring system of vibration and its self - under each separate turbine engine, operated on a particular Sun and in specific conditions. By vibration state, one should understand the magnitude of the vibration signal of the vibration detector at time t, i.e. value of V (t). According to the basic principle of diagnosis at the current time t, the vibrational state of a gas turbine engine can be described as:

where V _VIB0 (t ₀ ) is the initial (operational) vibration state of the gas turbine engine, t ₀ is the initial moment of operation of the gas turbine engine (normally t ₀ is the time interval of the initial engine operation on the aircraft), ΔV _VIB (t) is a diagnostic random process that describes change in vibration state at an arbitrary time t of engine operation; η (t) is a random process of vibrational state change not observed with the help of measuring instruments, independent of the past or current state.

Therefore, to control the current vibrational state of the gas turbine engine, it is proposed to implement the method in the form of a structural diagram (Fig. 1). GTE 2 is installed on board the aircraft 1, which is mandatory equipped with a drill 3, recording both the parameters of the power plant and the parameters of the movement of aircraft 1 and its position in space. GTE 2 is equipped with standard measuring equipment, most of the signals of which are fed to the control unit 3. For the description of the initial (reference) vibration state of GTE 2, not only the vibration sensor and rotor speed sensor are used, but also a number of other sensors installed on the GTD 2, as well as on aircraft 1.

In fact, modern gas turbine engines in operation are two- and three-shaft. The source of vibration perceived by the sensor may be an imbalance of rotating masses on any of the shafts, as well as the production or defects of shaft and pillow bearings. Therefore, the root-mean-square value of the amplitude of the vibration signal obtained at the output of the standard vibration measurement equipment on board the aircraft will include components that are multiples of the frequency of rotation of each of the GTE shafts. In relation to a twin-shaft gas turbine engine, we denote them as n ₁ is the rotational speed of the low pressure rotor (RND), and n ₂ is the rotational speed of the high pressure rotor (RVD).

Therefore, of course, when controlling the vibrational state, it is necessary to use measurements from the RVD speed sensor 5 and the RND speed sensor 6.

The magnitude of the vibration signal (vibration velocity or acceleration) depends on the heat state of the structural elements of the engine, because during his work heat exchange processes occur continuously in them and not always these processes lead to a uniform change in the geometric dimensions of structural elements. The latter phenomenon, in turn, again leads to an imbalance of the rotating masses of the engine and, as a consequence, to an increase in the amplitude of certain components of the vibration signal, measured by a standard sensor mounted on the motor housing. The heat condition of the engine design is estimated by the value of the gas temperature measured by thermocouples behind (t ₄ ^* ) or before (t ₃ ^* ) the RND turbine. Therefore, the signals from thermocouples 7 (we denote them as t _Г ^* ) installed on a gas turbine engine should be used to control the vibration state of the latter.

In flight, a certain load on the fan part (inlet stages) of the gas turbine engine compressor 2, especially on aircraft 1, which reach transonic and supersonic flight speeds, is exerted by the pressure of the inlet air flow. From this load, the level of the registered BUR 3 vibration signal changes, i.e. the air pressure at the engine inlet in flight affects its vibrational state. The pressure of the inhibited air flow at the inlet to each engine P _BX is usually measured by a corresponding sensor and the BUR 3 is recorded. Therefore, measurements from a standard sensor 8 P _{BX of the} engine are also used to monitor the vibration state of the engine.

The operating time t _{PAB_DB} GTE means the period of time from the start of engine start on the ground until it _turns off at the end of the flight or ground testing. For an engine with a normal vibration state, there should not be any significant dependence of the level of the vibration signal in the same modes on the operating time of the engine. And, on the contrary, for engines with a deteriorated vibration state, this dependence is quite, but not always necessary, can take place. Therefore, when monitoring the vibrational state of the engine in order to assess possible trend phenomena during one flight (operation cycle), the dependence of the vibration signal on the engine operating time should be taken into account. In general terms, the composition of the parameters of a gas turbine engine, on which its vibrational state depends, can be written as a vector

In flight on a gas turbine engine, the rotor part of which is essentially a massive gyroscope, linear and rotational speeds and accelerations, caused by the dynamics of the aircraft, operate. They, in turn, act on rotating structural elements of the engine (turbine rotors), causing gyroscopic moments from Coriolis forces on them, and, in addition, affect the signal level from vibration sensors installed on the gas turbine engine. Therefore, it is necessary to take into account the parameters of the aircraft motion when monitoring the vibrational state of the engine, as a result of which measurements are used from the following sensors that are part of the aircraft equipment: vertical overload sensor 9; sensor 10 longitudinal overload; lateral overload sensor 11; aircraft pitch angle sensor 12; sensor 13 angle of heel of the aircraft; yaw angle sensor 14; sensor 15 angular velocity around the transverse construction axis of the aircraft; angular velocity sensor 16 around the longitudinal construction axis of the aircraft; sensor 17 angular velocity around the vertical construction axis of the aircraft.

- вектор перегрузок, действующих на ВС, $$\overrightarrow{\theta }$$

- вектор углового положения ВС в пространстве, $$\overrightarrow{\omega }$$

- вектор угловых скоростей ВС в связанной системе координат.We denote these parameters as three vectors: $$\overrightarrow{g}$$

is the vector of overloads acting on the aircraft, $$\overrightarrow{\theta }$$

is the vector of the angular position of the aircraft in space, $$\overrightarrow{\omega }$$

is the vector of angular velocities of the aircraft in a coupled coordinate system.

Then the vibrational state of a gas turbine engine in flight conditions can be described by a functional relationship:

According to the principle of control incorporated in the method, the following actions are performed.

The information recorded by the BUR 3 is transferred using the device 20 reader from the aircraft.

Then, a device 21 for inputting passport and service data identifying the flight performed is used to select the read information to determine whether the information belongs to a specific aircraft, date and flight number. The device 21 input passport and service data is a keyboard, display and processor with solid state memory. The data entered from device 21, in addition to the date and number of the flight performed (flight number), includes the names of take-off and landing aerodromes, a short, usually encoded description of the flight mission. Using the same device, the formal data of the engines operated on a particular aircraft 1 are also entered, first of all, the serial number and installation location, as well as the formal data of the aircraft 1 itself to uniquely identify the ownership of the data being read from the control panel 3. The device 21 records the calibration data of the sensors and measurement systems, information from which is recorded on BUR 3, including sensors installed on GTE 2.

The registered BUR 3 information, passport and service data are received in block 22 for generating characteristics of the reference vibration state (BFCE) of the controlled gas turbine engine 2. This block at the initial stage of operation of the gas turbine engine 2 (after it is installed on board aircraft 1) determines the initial (reference) vibration characteristics of the gas turbine engine 2:

- one - for the work of a gas turbine engine on the ground, i.e. during its work on ground tests, as well as on the ground stages of flight from the moment the GTE is launched until the start of the aircraft take-off at the start, and from the moment the aircraft lands up to shutdown the GTE;

- the second - for the operation of a gas turbine engine in flight conditions, i.e. taking into account the influence of linear and rotational forces and accelerations acting on the aircraft in flight, as well as the angular position of the aircraft in space.

To form a standard vibration characteristic of a gas turbine engine during its operation on the ground at the beginning of operation, the following operations are performed in order with the information from the control panel accumulated at this stage:

;- the selection of the component of the vibration signal from the rotational speed of the high pressure rotor, i.e. values $$\Delta {\text{V}}_{{\text{VIB\_n}}_{\text{2}}}$$

;

;- selection of the component of the vibration signal from the rotational speed of the low pressure rotor, i.e. values $$\Delta {\text{V}}_{{\text{VIB\_n}}_{\text{one}}}$$

;

;- the selection of the component of the vibration signal from the temperature of the gases, i.e. values $$\Delta {\text{V}}_{{\text{VIB\_t}}_{\text{G}}^{\text{*}}}$$

;

;- the selection of the vibration component from the time the engine is running, i.e. values $$\Delta {\text{V}}_{{\text{VIB\_t}}_{\text{SLAV\_DV}}}$$

;

- the combination of the selected components in the reference vibration characteristics of the gas turbine engine when it is working on the ground, recorded as a functional relationship:

- determination of the boundaries of the description of the reference vibration characteristics during engine operation in terrestrial conditions and threshold (tolerance) values that separate the vibration signal of the healthy and malfunctioning state of the gas turbine engine.

To form the reference vibration characteristics of the gas turbine engine during its operation in flight at the beginning of operation, the following operations are performed in order with the information from the BUR accumulated during the first flights at this stage:

;- the selection of the component of the vibration signal from the rotational speed of the high pressure rotor, i.e. values $$\Delta {\text{V}}_{{\text{VIB\_n}}_{\text{2}}}$$

;

;- selection of the component of the vibration signal from the rotational speed of the low pressure rotor, i.e. values $$\Delta {\text{V}}_{{\text{VIB\_n}}_{\text{one}}}$$

;

;- the selection of the component of the vibration signal from the temperature of the gases, i.e. values $$\Delta {\text{V}}_{{\text{VIB\_t}}_{\text{G}}^{\text{*}}}$$

;

;- the selection of the component of the vibration signal from the air pressure at the engine inlet, i.e. values $$\Delta {\text{V}}_{{\text{VIB\_R}}_{\text{VX}}}$$

;

;- the selection of the component of the vibration signal from the time of engine operation, i.e. values $$\Delta {\text{V}}_{{\text{VIB\_t}}_{\text{SLAV\_DV}}}$$

;

;- the selection of the component of the vibration signal caused by the effect of vertical overload of the aircraft, i.e. values $$\Delta {\text{V}}_{{\text{VIB\_n}}_Y}$$

;

;- the selection of the component of the vibration signal caused by the impact of the longitudinal overload of the aircraft, i.e. values $$\Delta {\text{V}}_{{\text{VIB\_n}}_{\text{X}}}$$

;

;- the allocation of the component of the vibration signal caused by the impact of lateral overload of the aircraft, i.e. values $$\Delta {\text{V}}_{{\text{VIB\_n}}_{\text{Z}}}$$

;

;- the selection of the component of the vibration signal caused by the influence of the angular velocity around the vertical construction axis of the aircraft, i.e. values $$\Delta {\text{V}}_{\text{VIB\_}{\omega }_{\text{Y}}}$$

;

;- the selection of the component of the vibration signal caused by the influence of the angular velocity around the transverse construction axis of the aircraft, i.e. values $$\Delta {\text{V}}_{\text{VIB\_}{\omega }_{\text{Z}}}$$

;

;- the selection of the component of the vibration signal, introduced depending on the angular position relative to the transverse construction axis of the aircraft (pitch angle), i.e. values $$\Delta {\text{V}}_{\text{VIB\_}\vartheta }$$

;

;- the selection of the component of the vibration signal, introduced depending on the angular position relative to the longitudinal construction axis of the aircraft (roll angle), i.e. values $$\Delta {\text{V}}_{\text{VIB\_}\vartheta }$$

;

;- the selection of the component of the vibration signal, introduced depending on the angular position relative to the vertical construction axis of the aircraft (yaw angle), i.e. values $$\Delta {\text{V}}_{\text{VIB\_}\psi }$$

;

- the combination of the selected components in the standard vibration characteristics of the gas turbine engine during its operation in flight, written in the form of a functional relationship:

- determination of the boundaries of the description of the reference vibration characteristics during engine operation in flight conditions and threshold (tolerance) values that separate the vibration signal of the healthy and malfunctioning state of the gas turbine engine.

In the proposed method, the description of the standard vibrational state of the gas turbine engine is formed not on the test bench in the factory, but in the operating conditions of the gas turbine engine on board the aircraft. This approach allows you to take into account individual design features for installing a gas turbine engine on board the aircraft, as often the cause of increased vibration is a violation in the engine installation technology. In addition, the control of the vibration state of a gas turbine engine is correct if it is performed using the same measuring equipment, which also includes the registration channels in the BUR.

In addition, the definition of reference vibration characteristics is performed according to the adaptive principle. This principle is based on the fact that at the initial stage of operation of a gas turbine engine (during the first testing on the ground, during the first flight) it is impossible to cover all possible modes of operation of a gas turbine engine, especially dynamic ones, and the full range of flight modes. Therefore, the reference vibration characteristics during the operation of the gas turbine engine on the ground and in flight are not formed for all modes. During the further operation of the gas turbine engine, more and more new modes are gradually being covered. Therefore, the reference vibration characteristics of the gas turbine engine must be adjusted taking into account the measurement results of the sensors obtained in these new modes encountered. The main condition for such an adjustment is that the gas turbine engine is still in a good vibrational state. Therefore, the adaptation process, that is, the adjustment of the reference vibration characteristics, does not stop after a finite number of flights, but continues until there are new modes of operation of the gas turbine engine and new flight modes, provided that the engine is in good vibration condition.

After the formation of the reference vibration characteristics during the operation of the gas turbine engine on the ground and in flight, the method provides for the transition to monitoring the current vibration state of the engine by comparing it with its reference vibration state. The value of the vibration signal measured on the ground or in flight is compared with the reference value of the vibration signal obtained in the same gas turbine operation mode and in the same flight conditions. To ensure effective control of the vibrational state of the gas turbine engine on board the aircraft in flight, the on-board equipment includes a discriminator 18. It receives real-time flight signals from all five sensors installed on the gas turbine engine and nine motion sensors and the position of the aircraft in space from onboard equipment. Using the analog-to-digital converter and processor included in the block discriminator, a vibration signal is generated at each moment of flight time that strictly belongs to the current mode of operation of the gas turbine engine and the flight mode. Further, this unit compares the selected vibration signal with its reference value, according to the results of which the light, text or speech signaling device 19 of the aircraft crew can be turned on.

The values of the reference levels of the vibration state before the flight is transmitted to the block discriminator 18 using the ground-based device 23 for transmitting reference data (PED). This device includes interface channels for interfacing a BFC with an on-board discriminator unit. It should be noted that there is no need to transfer reference data before each aircraft flight. Data is transmitted to the BFCE only in cases where the GTE reference vibration characteristics are corrected on the ground or in flight.

When controlling the vibrational state of a gas turbine engine using the ground component of the proposed method, a block 24 of trend and predictive control (Trpr) is used, which is a specialized computer that performs two types of trend analysis: short-term and long-term. Short-term trend and predictive control is performed on the basis of a time series:

- разность между текущим и эталонным значением вибросигнала в момент времени t_i в текущем полете при выходе ГТД на j-й контролируемый режим.Where $$\Delta {\text{V}}_{\text{j}\text{VIB}}$$

- the difference between the current and the reference value of the vibration signal at time t _i in the current flight when the gas turbine engine enters the j-th controlled mode.

Long-term trend and predictive control is carried out on the basis of a time series:

- среднее изменение вибрации двигателя на j-м режиме в ном полете, T_НАРi -наработка двигателя на j-м режиме в i-м полете, T_{НАР ТЕК} - наработка двигателя на j-м режиме в текущем полете.Where $$\Delta {\text{V}}_{\text{j}\text{VIB}}$$

- the average change in engine vibration in the j-th mode in the flight, T _NARi is the engine _operating time in the j-th mode in the i-th flight, T _{NAR TEK} is the engine operating time in the j-th mode in the current flight.

Moreover, as an assessment of the danger of changes in the vibrational state of the engine as a whole, the mode with the most significant trend and the worst forecast is selected.

In addition, an integral element in the work of the ground component of the method is decision block 25, which is an integrated device, the main element of which is a specialized computer with high performance and a large amount of non-volatile memory. The composition of the decision-making unit also includes: a device (display) for displaying the results of the BUR information processing; a device for documenting the vibration state estimates of each individual gas turbine engine in an aircraft fleet located in the operating organization.

In cases where block 25 is not capable of automatically making a decision on the vibration state of the gas turbine engine, the interactive dialogue system is connected with the operator (decision maker). At the same time, access to the expert-reference system is opened, in which there is a database and knowledge of how to operate a specific sample of a gas turbine engine and the whole gas turbine engine park in the operating organization.

Claims (5)

1. The method of vibration diagnostics of gas turbine engines in operation according to information from onboard recording devices, including:
- registration of flight information of the aircraft and parameters characterizing the vibration state of gas turbine engines with an onboard registration device;
- reading registered information and parameters from the aircraft’s onboard registration device;
- identification of information and parameters read from the on-board recording device by entering passport and service data and formular data of an operating gas turbine engine;
- the formation of the first reference vibration characteristics of a gas turbine engine based on the results of engine operation on the ground or on board the aircraft, on the ground stages of the aircraft movement from the moment the engine starts to the start of the take-off, as well as from the moment the aircraft lands until the engine shuts off;
- the formation of the second reference vibration characteristics of the gas turbine engine in flight at each moment of flight, taking into account the influence of linear and rotational forces and accelerations acting on the aircraft in flight, as well as the angular position of the aircraft in space;
- determination of the boundaries of the description of the first reference vibration characteristics when the engine is operating in terrestrial conditions and threshold values that separate the vibration signal of a working and faulty state of a gas turbine engine;
- determination of the boundaries of the description of the second reference vibration characteristics during engine operation in flight conditions and threshold values that separate the vibration signal of the healthy and faulty state of the gas turbine engine;
- monitoring the current state of the first vibration characteristics of the gas turbine engine by comparing it with the corresponding reference vibration characteristics;
- monitoring the current state of the second vibration characteristic of the gas turbine engine by comparing it with the corresponding reference vibration characteristic taking into account the state of the first vibration characteristic of the gas turbine engine;
- determination of the technical condition of the gas turbine engine from the totality of deviations of the first and second vibration characteristics from the corresponding reference vibration characteristics at each current time of the gas turbine engine operation. 2. The method according to p. 1, characterized in that the formation of the first reference vibration characteristics of the gas turbine engine is carried out according to the adaptive principle, which allows for the adjustment of the first reference vibration characteristics with each new operation mode of the gas turbine engine during ground tests on board the aircraft, provided that the gas turbine engine is in good condition. 3. The method according to p. 1 or 2, characterized in that the formation of the first reference vibration characteristics of the gas turbine engine is carried out on the basis of the selected component of the vibration signal from the rotational speed of the high pressure rotor, the selected component of the vibration signal from the rotational speed of the low pressure rotor, the selected component of the vibratory signal from the gas temperature, the selected component of vibration from the time of engine operation. 4. The method according to p. 1, characterized in that the formation of the second reference vibration characteristics of the gas turbine engine is carried out according to the adaptive principle, allowing the adjustment of the second reference vibration characteristics with each new flight mode of the aircraft, provided that the gas turbine engine is in good condition. 5. The method according to p. 1 or 4, characterized in that the formation of the second reference vibration characteristics of the gas turbine engine is based on the selected component of the vibration signal from the rotational speed of the high pressure rotor, the selected component of the vibration signal from the rotational speed of the low pressure rotor, the selected component of the vibratory signal from the gas temperature, the selected component of the vibration signal from the air pressure at the engine inlet, the selected component of the vibration signal from the time of engine operation, the selected composition of the vibrating signal caused by the effect of vertical overload of the aircraft, the distinguished component of the vibro-signal caused by the effect of longitudinal overload of the aircraft, the distinguished component of the vibro-signal caused by the effect of lateral overload of the aircraft, the distinguished component of the vibro-signal caused by the influence of angular velocity around the vertical construction axis of the aircraft, the distinguished component of the vibro-signal caused by the influence of angular velocity around the transverse structural the axis of the aircraft, the selected component of the vibration signal introduced depending on the angular position relative to the transverse construction axis of the aircraft, the selected component of the vibration signal, introduced depending on the angular position relative to the longitudinal construction axis of the aircraft, the selected component of the vibration signal, introduced depending on the angular position relative to vertical construction axis of an aircraft.

Images