JP2003184570A - Estimating method for deteriorated gas turbine engine - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide an estimating method for a deteriorated gas turbine engine by which damage accumulated in a component of the gas turbine engine is accurately determined by utilizing parameter monitored and obtained during the operation of the engine. <P>SOLUTION: In the estimating method for the deteriorated gas turbine engine for estimating the deterioration of the component of the gas turbine engine, an accumulated distortion energy applied to the component is calculated in view of stress σ and distortion ε applied to the component in a certain cycle; that is, by using a formula: Σ(1/2 × maximum stress × distortion)×(cycle). Based on this accumulated energy, the deterioration of the component is estimated. <P>COPYRIGHT: (C)2003,JPO

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a method of estimating a deteriorated gas turbine engine for determining the maintenance timing of a gas turbine engine for an aircraft.

[0002]

2. Description of the Related Art A gas turbine engine for an aircraft includes a compressor 10, a combustor 11 and a turbine 1 as shown in FIG.
2. Combustion chamber 1 which is composed of jet nozzle 13 and compresses air entering from front air intake 14 with compressor 10.
1, burns the fuel injected from the fuel injection valve 15 with the compressed air, injects the high-temperature and high-pressure gas from the turbine nozzle 16, and rotates the turbine 12 that drives the compressor 10. At a high speed, the jet nozzle 13 in the rear portion is jetted to generate a propulsive force.

Since a gas turbine engine for an aircraft is repeatedly subjected to various loads during operation, in order to maintain its safety, maintenance such as repair or replacement is performed before deterioration of components of the gas turbine engine. I need.

[0004] As parameters showing the deterioration tendency of the engine, performance values such as thrust, compressor, fuel consumption, etc. are used as indexes, and these are detected by sensors, and the values are constantly monitored individually. It is acquired as navigation data. The acquired data becomes the navigation history of the gas turbine engine in a certain cycle, and the abnormality or malfunction of the gas turbine can be detected from the data.

However, from the viewpoint of safety, it is necessary to replace or maintain the components of the gas turbine at an appropriate timing before any abnormality or malfunction occurs.

[0006] Conventionally, as a method of determining the maintenance timing of a gas turbine engine for an aircraft which has been operated for a long time, the cycle and the usage time are used for judgment.

[0007]

However, in an actual engine, there are variations in the damage accumulated in the component parts due to individual differences in the engine, and it is not uniformly determined only by the cycle and time.

For example, thermal stress due to temperature, stress due to rotational centrifugal force, and gas pressure act on the blades that form the turbine 12, and these act repeatedly on the blades, which accumulates as damage on the turbine and It is estimated that if the accumulated strain energy reaches a certain value, it will break or break, but the temperature and stress acting on the turbine differ depending on the operation and the engine model. Are the same,
The damage varies.

Further, since the accumulated damage is subjected to a composite effect of temperature and stress, it cannot be managed or judged by evaluating individual parameters. That is, these parameters are data during navigation such as takeoff,
Even if these individual parameters are evaluated, it is difficult to evaluate the damage accumulated in the component parts.

Therefore, an object of the present invention is to solve the above problems and to accurately determine the damage accumulated in the component parts of the gas turbine engine by using the parameters monitored / acquired during the engine operation. Is to provide an estimation method of.

[0011]

In order to achieve the above object, the invention of claim 1 is a method for estimating a deteriorated gas turbine engine for estimating deterioration of a component of a gas turbine engine, and is configured in a certain cycle. This is a method for estimating a deteriorated gas turbine engine in which accumulated strain energy applied to a component is obtained from stress and strain applied to the component, and deterioration of the component is estimated based on the accumulated energy.

According to a second aspect of the present invention, the accumulated strain energy is obtained by Σ (1/2 × maximum stress × strain) × (cycle) from the maximum stress and strain applied to the component parts and the cycle thereof. Is a method of estimating a deteriorated gas turbine engine of.

According to the third aspect of the present invention, the accumulated strain energy with respect to the cycle is sequentially plotted on a graph to obtain an accumulated strain energy line, and from the line, the replacement time of the component parts is predicted. This is a method for estimating a turbine engine.

[0014]

BEST MODE FOR CARRYING OUT THE INVENTION Preferred embodiments of the present invention will be described in detail below with reference to the accompanying drawings.

First, the basic structure of the gas turbine engine is as described in FIG.

Table 1 shows the performance image of cycle, inlet temperature, rotation speed (turbine), turbine temperature, and fuel flow rate, which is output and obtained for each takeoff of a certain gas turbine engine (model IH ○ △). It shows the data. This take-off performance image data is recorded for each engine mounted on the aircraft, and is a simulation of the acquired data.

[0017]

[Table 1]

Table 1 shows only a part of the takeoff performance image data. From this data, a turbine blade is taken as an example.
A method of obtaining the stored strain energy will be described.

FIG. 2 shows the stress / strain hysteresis of the turbine blade, and shows the hysteresis of the maximum strain when the maximum stress is applied.

In FIG. 2, ΔW _TP is the tensile-side inelastic strain energy, ΔW _Te is the tensile-side elastic strain energy, and when the total tensile-side strain energy (strain energy per unit volume) is ΔW _TT , ΔW _TT is ΔW _TT = ΔW _TP + ΔW _Te ≈ 1/2 × σ
× ε.

FIG. 3 shows the accumulated strain energy accumulated in the turbine blade. Strain energy is accumulated as creep fatigue, and this creep has
Although a factor of time is also involved, it is assumed that strain energy is accumulated by cycles (number of times) in the figure.

From FIGS. 2 and 3, the accumulated strain energy is as follows: accumulated strain energy = Σ (1/2 × maximum stress × strain) × (cycle)

Here, the maximum stress is obtained by the following formula: maximum stress ≈ centrifugal force ∝ maximum rotation speed ² = 30 × maximum rotation speed ² / (10,010) ² . However, the above formula is the rotation speed 10,010 which is the data obtained by the experiment at the altitude 0 and the stationary state.
It can be applied in the case of an engine that produces a stress of 30 Ksi at rpm.

FIG. 4 shows the relationship between temperature / stress vs. strain.

Thus, strain is expressed as a function of turbine blade temperature by:

Distortion ∝ Turbine temperature = 2E-1
9e ^{(0.0465 x turbine temperature)} Turbine temperature is detected by a temperature sensor embedded in the turbine blade, and since the responsiveness is poor due to heat conduction, the gas temperature may not always be accurately indicated.
The temperature may be calculated based on the fuel flow rate value.

From the empirical formula, the turbine temperature in this case is Turbine temperature = 0.047 x fuel flow rate + 477.73 Ask in.

As described above, the accumulated strain energy can be obtained by obtaining the maximum stress from the maximum rotation speed and the strain from the turbine temperature.

FIG. 1 shows the accumulated strain energy accumulated in the turbine blade for each cycle based on the take-off performance data shown in Table 1, and the accumulated strain energy for each cycle is calculated for each gas turbine engine. The energy is plotted and shown in the graph.

In FIG. 1, a stored strain energy line a is a refurbished engine, one with severe energy storage, a stored strain energy line b is a refurbished engine, other than line a, and a stored strain energy line c is a new engine. , Line d indicates a line that becomes a breakage suspect in 2500 cycles, and the accumulated strain energy is 250,000.
0 is the limit of breakage suspect of the turbine blade.

Conventionally, as indicated by the line d, it is assumed that the breakage suspect is reached in 2500 cycles, and the maintenance is performed before reaching 2500 cycles uniformly.

However, as shown by the accumulated distortion energy line a, in the case of the refurbishment engine in which the energy accumulation is severe, 2
At 2000 cycles reaching 500 cycles, the breakage suspect limit is reached, and conversely, in the new engine, the cycle reaching the breakage suspect limit is 3500 cycles as shown by the accumulated strain energy line c.

Therefore, in the repair engine of the stored strain energy line a, 180 before reaching 2000 cycles.
With 0 cycles, in the engine for rehabilitation of accumulated strain energy line b, 2500 cycles with line c
With the new engine of the above, maintenance such as repair of turbine blades may be performed in 3000 cycles.

As described above, the cycle of reaching the breakage suspect differs depending on the individual engine.
As in the present invention, the accumulated strain energy is calculated for each takeoff,
This is plotted on a graph for each cycle as a stored strain energy line, and the cycle at which the limit of the breakage suspect is reached can be predicted from this line, whereby the maintenance timing of the engine can be accurately estimated.

In the above-mentioned embodiment, the repair of the turbine blade which is particularly subjected to the heat load has been described. However, in addition to the turbine blade, components such as the compressor blade, the bearing, the injection nozzle of the combustion chamber, etc., which need to be repaired. Of course, it can also be applied to.

Although the cycle can be obtained as direct data of the aircraft, the cycle number shown in FIG. 1 is an example for a specific model (IH ○ Δ), and if the model is different, it will be used as data. Of course, the numbers of the obtained cycles are also different.

[0037]

In summary, according to the present invention, it is possible to accurately determine the maintenance timing of a component by obtaining the accumulated strain energy applied to the component and estimating the deterioration from the accumulated energy.

[Brief description of drawings]

FIG. 1 is a diagram showing a relationship of accumulated strain energy with respect to a cycle obtained by a method of the present invention.

FIG. 2 is a diagram showing stress / strain hysteresis for explaining accumulated strain energy in the present invention.

FIG. 3 is a diagram showing accumulated strain energy accumulated in a turbine blade in the present invention.

FIG. 4 is a diagram showing a relationship between temperature / stress vs. strain in the present invention.

FIG. 5 is a schematic cross-sectional view of a gas turbine engine.

[Explanation of symbols]

12 turbine σ stress ε strain

Claims (3)

[Claims] 1. A method of estimating a deteriorated gas turbine engine for estimating deterioration of a component of a gas turbine engine, wherein accumulated strain energy applied to the component is obtained from stress and strain applied to the component in a certain cycle, and A method for estimating a deteriorated gas turbine engine, which comprises estimating deterioration of a component based on accumulated energy. 2. The accumulated strain energy is calculated by Σ (1/2 ×) from the maximum stress and strain applied to the component and its cycle.
The method for estimating a deteriorated gas turbine engine according to claim 1, wherein the maximum stress x strain) x (cycle) is calculated. 3. A method for estimating a deteriorated gas turbine engine according to claim 2, wherein accumulated strain energy with respect to cycle is sequentially plotted on a graph to obtain a accumulated strain energy line, and the replacement time of the component parts is predicted from the line. .