RU2796563C1 - Method for operation of an aircraft gas turbine according to its technical condition - Google Patents

Abstract

FIELD: operation; diagnostics of aircraft gas turbine engines.

SUBSTANCE: method for operating an aircraft gas turbine engine according to its technical condition includes determining the accumulated damage of each main part of the engine, taking into account engine operating modes and then determining the residual engine life by comparing the integral length of a crack that develops in one or more dangerous places of the highest stress concentration of the main part during engine operation from a defect not detected by non-destructive testing methods, with a maximum allowable value, in the event that the calculated value of the integral crack length approaches the maximum allowable value, the presence and actual length of the crack are controlled by non-destructive testing devices and operation is stopped if the control reveals that the actual crack length has approached its maximum permissible value, the ranges of flight conditions for implementation of engine loading cycles are predetermined, the ranges of flight conditions for the implementation of loading cycles are divided into zones, for each zone and each main part, dependence of the stress intensity factor on the length of the crack is calculated and, when determining the crack growth for the flight and the integral crack lengths the registered parameters that determine the operation of the engine are taken into account, including parameters that characterize the flight conditions for the implementation of each loading cycle.

EFFECT: ensuring a more complete use of the potential capabilities of engine parts in terms of resource due to the use of an improved mechanism for calculating the length of the crack.

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Description

The invention relates to the field of operation and diagnostics of aircraft gas turbine engines, taking into account specific operating conditions.

The prototype of the claimed invention is a known method for operating an aircraft gas turbine engine according to its technical condition (RU 2439527, IPC G01M 15/14, published on 01/10/2012), according to which the parameters that determine the operation of the engine are recorded, the accumulated damage to each main engine part is determined, taking into account the modes engine operation and the subsequent determination of the residual life by comparing the accumulated damage with the maximum allowable value. As a parameter characterizing the damageability of an engine part, a well-known patent uses the length of a crack that develops in the zone of the highest stress concentration during engine operation from a defect not detected by non-destructive testing methods. When determining the damage rate, the increase in the crack length per flight and the integral crack length accumulated in flights are calculated, taking into account the actual engine loading cycles during operation and the predetermined calculated dependence of the stress intensity factor (SIF) on the crack length. When determining the residual resource in the event that the calculated value of the crack length approaches the maximum allowable value, the presence and actual length of the crack are controlled by non-destructive testing devices. Operation is terminated if during the control it is revealed that the actual length of the crack has approached its maximum allowable value.

The disadvantage of the known method is the low accuracy of determining the residual engine life due to the bias of the crack length calculation mechanism, which does not take into account the influence of flight conditions on the crack growth from engine loading cycles. In a known method, to determine the crack growth of each main part of each loading cycle, the calculated dependence of the stress intensity factor on the length of the crack is used, previously determined under maximum loading conditions of the entire operating range. However, as practice shows, about 80% of the operation of engines of highly maneuverable aircraft (LA) is carried out at subsonic speeds and altitudes up to 10 kilometers, at which the load of the main engine parts is much lower than the maximum [Gogaev G.P., Nemtsev D.V. "Improving the methodology for monitoring the development of a resource by low-cycle fatigue of the main parts of the gas turbine engines of highly maneuverable aircraft", Collection "XLII of the International Youth Scientific Conference "Gagarin Readings" MAI, Moscow, 2018, pp. 124-125]. Also, operating experience and mathematical models of engines and aircraft indicate that each engine operation mode has different implementation limits for flight conditions.

Thus, the use of this mechanism for calculating the length of the crack leads to incomplete use of the potential capabilities of engine parts in terms of resource and, as a result, to an increase in the cost of the engine life cycle due to the replacement of engine parts that have not exhausted their resource during repair.

The technical result achieved by using the claimed method is a more complete use of the potential of engine parts in terms of resource due to the use of an improved crack length calculation mechanism. The achievement of the maximum allowable values of the length of the crack in parts when using the proposed method occurs after a longer period of operation compared to the prototype. Thus, the use of the claimed method reduces the life cycle cost of the engine.

The specified technical result is achieved by the fact that in the claimed method of operating an aircraft gas turbine engine according to its technical condition, including the registration of parameters that determine the operation of the engine, the determination of the accumulated damage of each main engine part, taking into account the operating modes of the engine, and the subsequent determination of the residual engine life by comparing the integral length of the crack , developing in one or several dangerous places of the highest stress concentration of the main part during engine operation from a defect not detected by non-destructive testing methods, with a maximum allowable value, while the integral crack length is determined taking into account the increase in the crack length for each flight, real engine loading cycles in the process work and a predetermined calculated dependence of the stress intensity factor on the length of the crack, when determining the residual life of the engine in the event that the calculated value of the integral length of the crack approaches the maximum allowable value, the presence and actual length of the crack are controlled by non-destructive testing devices and the operation of the main part is stopped if during the control it is revealed, that the actual length of the crack has approached its maximum allowable value, the ranges of flight conditions for the implementation of engine loading cycles are preliminarily determined, the ranges of flight conditions for the implementation of loading cycles are divided into zones, for each zone and each main part, the calculated dependence of the stress intensity factor on the length of the crack is determined and at in determining the crack growth for the flight and the integral length of the crack, the registered parameters that determine the operation of the engine, including the parameters characterizing the flight conditions for the implementation of each loading cycle, are taken into account.

Dividing the range of flight conditions for the implementation of loading cycles of an aircraft gas turbine engine into zones, determining for each zone the calculated dependence of the stress intensity on the crack length for each main part, as well as determining the crack growth for the flight and the integral length of the crack, taking into account the registered parameters characterizing the flight conditions for the implementation of each of the loading cycle, makes it possible to use, when determining the integral length of the crack, different values of the increase in the length of the crack according to the belonging of the recorded parameters corresponding in time to the peak of the selected loading cycle, to the selected zone of engine flight conditions. If the peak of the selected cycle corresponds to the boundary of the zones, then to determine the crack growth for dangerous places of the highest stress concentration of the main part, the crack growth belonging to one of the adjacent zones, in which its value will be the smallest, is used, as a result, the calculation of the accumulated damage will be as close as possible to flight conditions. Thus, the claimed method takes into account the flight conditions under which the selected loading cycle was implemented.

Significant features may have development and addition.

The ranges of flight conditions for the implementation of engine loading cycles are determined based on the data of mathematical models of the engine and the aircraft or on the basis of the results of the analysis of serial operation.

When dividing the entire range of flight conditions for the implementation of loading cycles into zones, the number and size of zones are selected individually for each main part in order to achieve the required accuracy in calculating the integral crack length. As a criterion for determining the number and size of zones, a relative change in the damage value between zones can be used in such a way that for neighboring zones this change should not exceed the selected tolerance. The minimum tolerance value corresponds to such a value of the relative change in the damage between zones, at which a further decrease in the size and a corresponding increase in the number of zones does not lead to a significant refinement of the accumulated damage and, at the same time, requires a significant increase in the required calculations.

The parameters at the engine inlet are used as parameters characterizing the flight conditions for the implementation of loading cycles - full temperature and full pressure, due to the fact that these parameters determine the conditions of thermomechanical loading of the main parts.

The claimed invention is explained further by a more detailed description of its implementation with reference to the figures, where:

in fig. 1 - graph of the ranges of flight conditions for the implementation of engine loading cycles in the coordinates of the total temperature and total pressure at the engine inlet at various engine operating modes;

in fig. 2 - graph of the ranges of flight conditions for the implementation of engine loading cycles with division into zones;

in fig. 3 is a plot of the dependence of the stress intensity factor on the length of the crack for different zones of the ranges of flight conditions;

in fig. 4 - graph of the function of changing the engine speed over time;

in fig. 5 is a graph of the dependence of the integral length of the crack on the duration of operation for various methods of operating an aircraft engine.

The method of operating an aircraft gas turbine engine is based on the fact that in the process of operation the actual operating time of the engine and the integral crack length in each of the dangerous places of its main parts are compared with their maximum allowable values determined by the results of tests of the fatigue properties of the material of the parts and calculations of the stress-strain state of dangerous detail areas.

The claimed method is carried out as follows.

1) At the stage of designing and fine-tuning the engine, in order to simplify the consideration of the variety of modes of its operation, loading is schematized. The rotor speed is taken as the main characteristic parameter of the engine operation, which determines the loading mode. To schematize the loading of the engine, the entire range of changes in its operation in terms of rotational speed is divided into a number of designated modes and the ranges of rotational speed values corresponding to them are determined. The number of assigned engine operating modes and, accordingly, typical loading cycles can vary and are determined taking into account the technical requirements for the engine, its control system, the effect of changing the rotor speed on the development of cyclic durability of the main engine parts (such as compressor and turbine disks, combustion chamber, shafts, as well as other parts, the destruction of which can lead to dangerous consequences [Aviation Rules - Part 33. Airworthiness Standards for Aircraft Engines]), as well as the purpose of the aircraft in which this engine is used.

On the example of selection in the range of changes in engine operation by the rotation frequency of the assigned modes: MG - low gas, CR - cruising mode, MAX U - maximum training mode, MAX B - maximum combat mode - the following typical loading cycles are distinguished:

N1U - corresponds to the change in rotational speed n ₀ - n _{MAX Y} - n ₀ ;

N2Y - corresponds to a change in the rotational speed n _MG - n _{MAX Y} - n _MG ;

N3U - corresponds to a change in rotational speed n _KR - n _{MAX Y} - n _KR ,

N1B - corresponds to a change in rotational speed n ₀ - n _{MAX B} - n ₀ ;

N2B - corresponds to a change in the rotational speed n _MG - n _{MAX B} - n _MG ;

N3B - corresponds to a change in rotational speed n _KR - n _{MAX B} - n _KR ,

where n ₀ - rotational speed equal to zero (engine off);

n _MG - range of rotational speeds in idle mode;

n _CR - speed range in cruising mode;

n _{MAX Y} - range of rotational speeds at the maximum training mode,

n _{MAX B} - range of rotational speeds at the maximum combat mode.

2) Further, at the stage of designing and debugging the engine, the ranges of flight conditions for the implementation of loading cycles are determined, for example, based on the data of mathematical models of the engine and aircraft or based on the results of serial operation. The most reliable way to determine the ranges of flight conditions for the implementation of loading cycles is the analysis of serial operation. It is recommended to use ExCon v. 1.0 [Computer program of the Russian Federation No. 2018665849 dated 12/11/2018] or its subsequent versions, allowing in an automated mode to analyze the implementation of flight conditions in operation. As parameters characterizing the flight conditions, the values of speed (Mach number M or indicated speed V _pr ), flight altitude H, total temperature at the inlet to the engine T* _inlet , total pressure at the inlet to the engine P* _inlet , and others can be used. parameters that allow determining the conditions of thermomechanical loading of the main parts. In the considered example of the implementation of the claimed method, the total temperature T* _in and the total pressure P* _in at the engine inlet are used as parameters characterizing the flight conditions. For various combinations of parameters M and H, the parameters at the engine inlet T* _in and P* _in can be similar, in addition, T* _in and P* _in are classical perturbing effects of the theory of control systems for aircraft power plants, which, together with a given engine operation mode unambiguously determine the conditions of thermomechanical loading of components and parts. Thus, it becomes possible to group different conditions according to the values of the parameters M and H having similar values of the parameters T* _in and P* _in at the motor inlet. Parameters T* _in and P* _in are measured directly on the engine or calculated from the parameters measured directly on the engine.

An example of the ranges of flight conditions for the implementation of loading cycles B (N1, N2, N3) and Y (N1, N2, N3) is shown in Fig. 1.

3) The obtained ranges of flight conditions for the implementation of loading cycles are divided into zones. The number and size of zones are selected individually for each type of loading cycle for each main part. Increasing the number of zones makes it possible to increase the accuracy of crack growth control, but increases the required amount of calculations.

As an example, in FIG. 2 shows the division into zones for N1Y cycles.

4) For selected dangerous places of the highest stress concentration of each main part, determine the size of the initial defect L _0i undetected by non-destructive testing methods. For each main part, one or more dangerous places, denoted by i, can be selected. Dangerous places are determined based on the stress-strain state of the main part, the possibility of studying these places by non-destructive testing methods, the possible presence of metallurgical and technological defects in these places. Depending on the above conditions, each major part may have one or more hazardous locations.

5) For the materials used for the main parts, standard tests of crack resistance properties at operating temperature (according to OST 1 92127-90) are carried out to determine the characteristics C, m, ΔK _c characterizing the rate of crack propagation in the material under cyclic loading, where C and m are the coefficients of the equation Paris, ΔK _c is the critical value of the stress intensity factor at which the crack growth rate becomes unacceptably high.

6) Carry out the necessary calculations to determine the values of the stress intensity factors ΔK corresponding to arbitrarily given lengths L of cracks in each place of each main part for each type of cycle under different flight conditions in the specified zones according to paragraph 3. The determination of the values of the stress intensity factors from the crack length is carried out by any of methods known in fracture mechanics (for example, V.Z. Parton, E.M. Morozov Mechanics of elastic-plastic fracture, M .: Nauka, 1974).

Based on the obtained values, the dependences of the stress intensity factors on an arbitrary crack length ΔK _ijk =f(L) are constructed for each dangerous place of each main part i of each type of cycle j of each zone of implementation of cycle k.

An example of building dependencies for different zones is shown in Fig. 3.

7) According to the obtained graphs of the dependence of the stress intensity factor on the length of the crack ΔK _ijk =f (L) for each dangerous place of each main part, find the critical crack length L _ci corresponding to the critical value of the stress intensity factor ΔK _c determined when testing samples for crack resistance. Since each dangerous place corresponds to several specific dependences of the stress intensity factor on the length of the crack in cases for different zones, the final value ΔK _c is taken to be the smallest of the obtained values in different zones (in reserve).

8) For each dangerous zone OD, the maximum allowable crack length L _dopi is determined by the formula L _dopi =L _ci /Z, where Z is the safety factor.

9) When operating the engine, to control the achievement of the limiting values of L _dopi , algorithms for processing the recorded flight information are used, which make it possible to identify typical loading cycles. These algorithms are based on the function of changing the engine speed over time.

The loading cycles are determined in the following sequence:

a) For the cyclogram of the change in the engine speed for one flight, all extrema of the time function n=f(τ) are determined (Fig. 4);

b) In accordance with the methods of schematization of random processes (GOST 25.101-83), all loading cycles of the function n=f(τ) are distinguished.

During the flight, the parameters T* _in and P* _in are recorded with the required frequency. After selecting typical loading cycles, the values of the recorded parameters T* _in and _P * in corresponding in time to the peaks of the selected cycles are determined (Fig. 4).

According to the belonging of the selected parameters T* _in and P* _in to the selected zone of engine flight conditions, the appropriate dependence of the stress intensity factor on the crack length ΔK _ijk =f(L) is selected, which is subsequently used to calculate the increment in the crack length ΔL _ni according to the Paris formula ΔL _ni =CΔK _i (L _Σ(n-1)i ) ^m , where: ΔK _i (L _Σ(n-1)i ) - stress intensity determined for each dangerous place of each main part according to the corresponding schedule (Fig. 3) from integral crack length L _Σ(n-1)i .

The integral length of the crack before the first flight is taken equal to the length of the initial defect L _0i undetected by non-destructive testing methods. After the first loading cycle, the crack growth will be ΔL _1i =CΔK _i (L _0i ) ^m , and the crack length L _Σ1i =L _0i +ΔL _1i . Further, the crack growth from subsequent loading cycles will be determined taking into account the integral length of the crack after previous loading cycles.

10) At the end of the flight, the obtained integral length of the crack L _Σi for each dangerous place is compared with the maximum allowable value of the length of the crack L _dopi . If the condition L _Σi ≥ L _addi is met for any of the dangerous places of the main part, the operation of the main part is stopped and the dangerous places of the main part are inspected for possible development of a crack in it from the alleged defect.

In case of detection of a developed defect, the part is removed from further operation and replaced with a new one. In the absence of a developed defect, it is assumed that the defect has not developed, its length is again taken equal to the size of the initial defect L _0i , not detected by the applied non-destructive testing methods, and the operation of the main part continues as described above.

Thus, a new approach to the implementation of the method of operating an aircraft engine according to the invention makes it possible to take into account the influence of flight conditions in determining the increment and integral crack length, which will increase the service life without removing parts and reduce the cost of the engine life cycle. In FIG. Figure 5 shows the integral increments of the crack length for one dangerous place in the main part, taking into account the influence of flight conditions and taking into account operation in the zone of maximum loading conditions.

Claims (4)

1. A method of operating an aircraft gas turbine engine according to its technical condition, including registering the parameters that determine engine operation, determining the accumulated damage of each main engine part, taking into account the engine operating modes, and then determining the residual engine life by comparing the integral length of a crack developing in one or more dangerous places of the highest stress concentration of the main part during engine operation from a defect not detected by non-destructive testing methods, with a maximum allowable value, while the integral length of the crack is determined taking into account the increase in the length of the crack for each flight, real engine loading cycles during operation and a predetermined calculated dependence stress intensity factor on the length of the crack, when determining the residual life of the engine in the event that the calculated value of the integral length of the crack approaches the maximum allowable value, the presence and actual length of the crack are controlled by non-destructive testing devices and the operation of the main part is stopped if the control reveals that the actual length of the crack has approached its maximum permissible value, characterized in that the ranges of flight conditions for the implementation of engine loading cycles are preliminarily determined, the ranges of flight conditions for the implementation of loading cycles are divided into zones, for each zone and each main part, the calculated dependence of the stress intensity factor on the length of the crack is determined and when determining the increase cracks per flight and the integral length of the crack take into account the registered parameters that determine the operation of the engine, including parameters that characterize the flight conditions for the implementation of each loading cycle. 2. The method according to p. 1, characterized in that the ranges of flight conditions for the implementation of engine loading cycles are determined based on the data of mathematical models of the engine and the aircraft or according to the results of the analysis of serial operation. 3. The method according to claim 1, characterized in that when dividing the entire range of flight conditions for the implementation of loading cycles into zones, the number and size of zones are selected individually for each main part. 4. The method according to claim 1, characterized in that the parameters characterizing the flight conditions for the implementation of loading cycles are used as parameters at the engine inlet - full temperature and full pressure.

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