KR20110130292A - Method of detecting failure of aeroplane engine and apparatus for detecting failure of aeroplane engine performing the same - Google Patents

Abstract

PURPOSE: A failure detect method of an airplane engine and a detect device using the same are provided to accurately detect the failure of an airplane engine based on the boundary value by varying a failure detect boundary value according to the mission profile of an airplane. CONSTITUTION: A failure detect method of an airplane engine comprises the following steps. The mission profile of the airplane is discriminated. The failure detect boundary value is determined according to the mission profile(110,120). The failure of the airplane engine is determined based on the failure detect boundary value(130). The mission profile comprises one among gliding, take-off, cruise, descending and landing. The failure detect boundary value is based on a predetermined signal generated in the airplane engine.

Description

FIELD OF DETECTING FAILURE OF AEROPLANE ENGINE AND APPARATUS FOR DETECTING FAILURE OF AEROPLANE ENGINE PERFORMING THE SAME}

The present invention relates to a failure detection method of an aircraft engine and an aircraft engine failure detection device for performing the method, and more particularly, to a failure detection method of an aircraft engine using a signal generated from the aircraft engine and an aircraft engine performing the method. It relates to a failure detection device.

Engines used in aircraft may be divided into reciprocating engines and gas turbine engines. Gas turbine engines are commonly used in terms of jet engines.

 Unlike reciprocating engines, jet engines are composed of a series of intake-compression-combustion-exhaust air, which are relatively low in vibration and highly efficient.

On the other hand, when flying at low speed and low altitude, the reciprocating engine has higher efficiency than the gas turbine engine and is easier to maintain and manage. Therefore, the reciprocating engine is widely used for ultralight aircraft and small helicopter engines.

In detecting a failure of an aircraft engine, a jet engine performs a failure detection method based on a flight data recorder (FDR) or a health & usage monitoring system (HUMS).

The FDR, also known as the black box, is installed in a heat-resistant shock-resistant capsule attached to the tail of the aircraft and is a recording device for investigating the cause of the aircraft's accidents. Five data of altitude, air velocity, nose orientation, vertical acceleration, and time Carved with diamond needle on stainless steel tape.

As the development of the aircraft is advanced, it is difficult to explain the cause of the accident with only this FDR.In addition to the five data, 19 types including the attitude of the aircraft, the control of the wing of the aircraft, the thrust of each engine, and the transmission of VHFHH communication are available. A digital flight data recorder (DFDR) is used to record and record data from the data.

HUMS, another failure detection device used in jet engines, is an aircraft condition monitoring device that can retrieve and diagnose all status information of an aircraft in real time. HUMS enables the real-time diagnosis and analysis of all aircraft's internal and external environmental and operational status during flight to ensure safe flight and efficient gas management.

HUMS performs early diagnosis of damage to the aircraft, part condition diagnosis and system defect recognition, operation and flight time identification with automatic data processing, engine condition and engine oil contamination condition diagnosis.

In small aircrafts, FADEC (Full Authority Digital Engine Control) is used to detect engine failures for reciprocating engine failures. In addition, HUMS is used to detect structural failures of the aircraft, but unlike jet engines, reciprocating engines are limited in terms of failure diagnosis due to economic constraints, and the performance of computing devices to be mounted is also limited.

It is a first object of the present invention to provide a method for detecting a failure of an aircraft engine, which can detect a failure that may occur in an aircraft engine using a simple method.

In addition, a second object of the present invention is to provide an aircraft engine failure detecting apparatus for performing a failure detection method of an aircraft engine capable of detecting a failure that may occur in an aircraft engine by a simple method.

In the fault diagnosis method of an aircraft engine according to an aspect of the present invention for achieving the first object of the present invention described above, determining a mission shape of the aircraft and determining a failure detection threshold value according to the mission shape, the failure And determining whether the aircraft engine has failed based on a detection threshold value. The mission shape of the aircraft may be one of slide, takeoff, cruise, descent, and landing. The fault detection threshold is a value based on a predetermined signal generated by the aircraft engine and may vary according to the mission shape of the aircraft. The failure detection threshold may be a value generated based on at least one of a vibration signal of the aircraft engine, RPM information of a crankshaft of the aircraft engine, and torque information of a crankshaft included in the aircraft engine. The determining of the failure of the aircraft engine based on the failure detection threshold may determine whether the crankshaft included in the aircraft engine has failed. The determining of the failure of the aircraft engine based on the failure detection threshold may include converting a signal provided by the sensor into a frequency domain using a sensor installed in the aircraft engine and when the aircraft engine is in a normal state. The failure of the aircraft engine may be determined by comparing a value obtained by converting a signal provided from the sensor installed in the engine of the aircraft into a frequency domain. The determining of the failure of the aircraft engine based on the failure detection threshold may include determining whether the aircraft engine has at least one of a mass imbalance, an axis alignment error, a bearing engagement error, an axis friction, an axis imbalance, and oil rotation. You can judge. The aircraft engine may be a reciprocating engine having a natural vibration frequency due to the four stroke provided in the small aircraft.

In addition, the aircraft engine failure diagnosis apparatus according to an aspect of the present invention for achieving a second object of the present invention is a sensor unit for detecting a signal generated by the aircraft engine, the signal generated in the sensor unit in the frequency domain A signal processor for converting the signal converted into the frequency domain provided by the signal processor to convert the normal signal generated when the engine of the aircraft generated from the aircraft engine to the frequency domain according to the mission shape of the aircraft It may include a failure detection unit for determining whether or not the aircraft engine failure compared to one signal. The failure detection unit may determine whether the crankshaft included in the aircraft engine. The failure detector may determine whether at least one of a mass imbalance, an axis alignment error, a bearing engagement error, an axis friction, an axis imbalance, and oil rotation occurs in the aircraft engine. The mission shape of the aircraft may be one of slide, takeoff, cruise, descent, and landing. The aircraft engine may be a reciprocating engine having a natural vibration frequency due to the four stroke provided in the small aircraft.

According to the above-described failure detection method of an aircraft engine and an aircraft engine failure detection device that performs the above method, a failure that may occur in an engine of an aircraft based on the threshold value by changing the failure detection threshold value according to the mission shape of the aircraft. Is detected.

Therefore, the accuracy of aircraft engine failure detection can be improved, and the failure of the aircraft engine can be effectively detected using a simple method.

In addition, by providing a method for easily detecting a failure that may occur in the engine of the aircraft, it can be applied not only to a large aircraft but also to a small aircraft having a limitation on the performance of the computing device to be mounted for failure detection.

1 is a flowchart illustrating a method for detecting a failure of an aircraft reciprocating engine according to an embodiment of the present invention.
Figure 2 is a conceptual diagram showing the mission shape of the aircraft according to an embodiment of the present invention.
3 is a conceptual diagram illustrating a state in which a sensor for measuring a vibration signal of an engine installed in an engine of an aircraft according to an embodiment of the present invention is installed in an engine.
4 is a graph illustrating a method of determining whether a failure occurs by converting a vibration signal of an engine into a frequency domain according to an embodiment of the present invention.
5 is a graph showing a result of simulation of a failure detection method of an aircraft engine according to an embodiment of the present invention using the above-described method.
6 is a graph showing a result of simulation of a failure detection method of an aircraft engine according to an embodiment of the present invention using the above-described method.
7 is a block diagram illustrating a failure detection apparatus for performing a failure detection method of an aircraft engine according to an exemplary embodiment of the present invention.

As the invention allows for various changes and numerous embodiments, particular embodiments will be illustrated in the drawings and described in detail in the written description. However, this is not intended to limit the present invention to specific embodiments, it should be understood to include all modifications, equivalents, and substitutes included in the spirit and scope of the present invention. Like reference numerals are used for like elements in describing each drawing.

The terms first, second, etc. may be used to describe various components, but the components should not be limited by the terms. The terms are used only for the purpose of distinguishing one component from another. For example, without departing from the scope of the present invention, the first component may be referred to as the second component, and similarly, the second component may also be referred to as the first component. And / or < / RTI > includes any combination of a plurality of related listed items or any of a plurality of related listed items.

It is to be understood that when an element is referred to as being "connected" or "connected" to another element, it may be directly connected or connected to the other element, . On the other hand, when a component is referred to as being "directly connected" or "directly connected" to another component, it should be understood that there is no other component in between.

The terminology used herein is for the purpose of describing particular example embodiments only and is not intended to be limiting of the present invention. Singular expressions include plural expressions unless the context clearly indicates otherwise. In this application, the terms "comprise" or "have" are intended to indicate that there is a feature, number, step, operation, component, part, or combination thereof described in the specification, and one or more other features. It is to be understood that the present invention does not exclude the possibility of the presence or the addition of numbers, steps, operations, components, components, or a combination thereof.

Hereinafter, with reference to the accompanying drawings, it will be described in detail a preferred embodiment of the present invention. Hereinafter, the same reference numerals are used for the same components in the drawings, and duplicate descriptions of the same components are omitted.

1 is a flowchart illustrating a method of detecting a failure of an aircraft reciprocating engine according to an exemplary embodiment of the present invention.

Referring to FIG. 1, first, information related to the current altitude, speed, and pressure of an aircraft may be provided, and the current mission shape of the aircraft may be determined based on the information (step 110).

An aircraft may have a variety of mission shapes in service.

2 is a conceptual diagram showing a mission shape of the aircraft.

Referring to FIG. 2, the mission shape refers to the operational state of the current aircraft, and the mission shape of the aircraft is, for example, the slide 210, the takeoff 220, the cruise 230, the descent 240, and the landing 250. It can be one. This is a simplified representation of the aircraft's condition from the departure of a particular airport to the destination airport.

As the mission shapes of the aircraft, such as the slide 210, takeoff 220, cruise 230, descent 240, landing 250 changes during operation, various information such as altitude information, speed information, pressure information, etc. Can be changed to a value.

The mission shape of the aircraft can be determined based on various pieces of information that can be provided by the aircraft. For example, if the altitude information or pressure information of the aircraft changes drastically, it may be determined that the takeoff 220 or descending 240 state, the altitude and pressure is constantly changing and the speed is maintained within a certain range In this case, it may be determined that the current state of the aircraft is the cruise 230 state.

The altitude information, speed information, and pressure information may be used as an example for determining the mission shape of the aircraft, and may determine the mission shape of the aircraft using other information capable of identifying the mission shape of the other aircraft.

In addition, it is also possible to determine the mission shape of the aircraft by further subdividing the mission shape of the aircraft 210, takeoff 220, cruise 230, descent 240, landing 250.

When the mission shape of the aircraft is determined through step 110, a threshold value according to a signal generated by the engine according to the mission shape of the aircraft may be formed (step 120).

For example, RPM (revolution per minute) value and torque value of the crankshaft, which is a component included in the engine, may have various values according to the mission shape of the aircraft. Therefore, if the RPM value and torque value of the crank shift can vary according to the mission shape of the aircraft, the vibration signal of the engine including the crankshaft also changes the mission shape of the aircraft when at least one of the RPM value and the torque value of the crankshaft is changed. It may vary.

The predetermined signal generated by the engine in the steady state may vary the boundary value of the normal range according to the mission shape of the aircraft. If a predetermined signal generated by the engine is beyond the boundary of the normal range according to the mission shape, it may be determined that the engine has failed.

For example, the boundary value of the normal range according to the mission of the aircraft is a rotational RPM value of the crankshaft included in the engine, and may vary according to the mission shape. It is also possible to set the boundary value of the normal range according to the mission shape of the.

That is, the signal generated by the engine according to the mission shape of the aircraft may be a constant signal that may occur in some components of the engine, and when the type of the constant signal is different, the boundary value of the normal range according to the mission shape of the aircraft is different from each other. Can have

In addition to measuring, signals from some components of the engine can be measured, as well as signals from the engine itself.

The failure of the engine may be determined using the set threshold value (step 130).

3 is a conceptual diagram illustrating a state in which a sensor for measuring a vibration signal of an engine installed in an engine of an aircraft according to an embodiment of the present invention is installed in an engine.

As a sensor capable of measuring vibration signals of the engine, a non-contact sensor, a speedometer sensor, an accelerometer sensor, or the like may be used.

Referring to FIG. 3, it is illustrated that the accelerometer sensor 310, which is easy to install functionally and accurately measured in pepper wave number, is installed in a direction perpendicular to the cowl of the engine 300 in a manner recommended by ISO, but has a function of measuring vibration. In the case of a sensor having the present invention can be used in a range that does not depart from the essence of the present invention, and also in the installation of the sensor is not limited to the installation position, the installation method, the number of sensors and the like.

Determination of the failure of the aircraft engine may be made through frequency analysis of the vibration signal detected using the vibration sensor 310 installed in the engine 300. By analyzing the magnitude of the acceleration at a specific frequency obtained through frequency analysis and the magnitude of the acceleration at a specific frequency when the crankshaft is normal, it is possible to determine whether the failure.

For example, assuming that the vibration signal of an aircraft engine has a peak value of acceleration magnitude at 100 Hz during cruising in a steady state and an abnormally high peak value at 50 Hz when a certain component of the engine has failed, the vibration signal of the aircraft engine is analyzed and analyzed. The failure which occurred in the component part of an engine can be detected.

In the method for detecting a fault signal, as described above, a threshold value can be set based on a vibration signal generated from the engine to detect a fault. However, the present invention is not limited thereto. Signals and the like may be analyzed to detect boundary values and failures for determining whether an aircraft engine has failed.

In addition, in order to individually detect failures of various components of the aircraft, if a separate component fails, different signals may be separately provided to have a normal range boundary value according to the mission shape that is most effective in detecting failures according to each component. Can be.

For example, if a bearing of an engine fails, frequency analysis using a rotation sensor is performed, and if the crankshaft of the engine fails, a vibration sensor is used to diagnose the failure. You can use a sensor that can accurately judge.

As a result of the determination in step 130, when a failure occurs in the aircraft engine, the failure detection may be notified.

It is possible to notify the pilot or base station that a failure has occurred in the engine of the aircraft through a specific signal or record that the failure has occurred in the black box included in the aircraft.

4 is a graph illustrating a method of determining whether a failure occurs by converting a vibration signal of an engine into a frequency domain according to an embodiment of the present invention.

Referring to FIG. 4, the acceleration magnitude according to the frequency-converted value based on the vibration signal of the engine was used as a value used for the measurement of the failure signal, but as described above, another value other than the acceleration magnitude based on a signal other than the vibration signal is also used in the aircraft. Can be used to detect engine failures.

Figure 4a is a graph showing the amplitude of the vibration signal (Acceleration Amplitude) according to the frequency when the aircraft is in a steady state according to an embodiment of the present invention.

Referring to FIG. 4A, the acceleration magnitude increases and decreases periodically with peak values at 100 Hz (1x) as the first rotation frequency, 200 Hz (2x) as the harmonic frequency, 300 Hz (3x), and 400 Hz (4x). Can be.

4B is a graph illustrating conversion of a vibration signal of an engine into a frequency domain when an axis imbalance occurs in an aircraft engine according to an exemplary embodiment of the present invention.

Axial imbalance is a fault signal that occurs due to an unbalance of the weight of the rotor with respect to the center of rotation.

Referring to FIG. 4B, when the axis imbalance signal is generated, it can be seen that the acceleration magnitude peak value of 100 Hz, which is the primary rotational frequency, is increased compared to the normal state, and the increase amount of the acceleration magnitude may be proportional to the degree of axis imbalance. Therefore, it is possible to detect an axial imbalance occurring in the engine based on the signal characteristics in this frequency domain.

4C is a graph illustrating conversion of a vibration signal of an engine to a frequency domain when axial friction occurs according to an embodiment of the present invention.

Axial friction is a failure that can be caused by micro clearance contact between the rotor and the fixture.

Referring to FIG. 4C, when an axial friction occurs, a peak value of an acceleration magnitude may be detected in a frequency region corresponding to 0.5 times the primary rotation frequency. Therefore, it is possible to detect the axial friction generated in the engine based on the signal characteristics in this frequency domain.

4D is a graph illustrating conversion of a failure signal that can be detected when an oil rotation occurs in a frequency domain according to an embodiment of the present invention.

Oil rotation is a failure that occurs when oil between the bearing and the shaft is attracted to the shaft to rotate at shaft speed.

Referring to FIG. 4D, it can be seen that the peak value of the acceleration magnitude signal is detected at a frequency of 2 / 5x when oil rotation occurs. Thus, oil rotation generated in the engine can be detected based on the signal characteristics in this frequency domain.

5 is a graph showing a result of simulation of a failure detection method of an aircraft engine according to an embodiment of the present invention using the above-described method.

For the simulation, the aircraft engine was simulated using AMEsim (Advanced Modeling Environment for Simulation of engineering system).

As the fault signal, the axis imbalance and the axial friction signal were applied with time difference and the mission status of the aircraft was taken off and the threshold value was set.

Figure 5a is a graph showing the acceleration value according to the time that occurs during takeoff of the aircraft in the steady state.

5B is a graph showing the magnitude of acceleration according to the frequency appearing when an axis imbalance signal is applied as a failure signal.

Referring to FIG. 5B, it can be seen that when an axis imbalance occurs, an acceleration magnitude value at an initial rotation frequency (50 Hz) has an abnormal peak value beyond a normal range. That is, when the axis imbalance occurs during takeoff, it can be seen that the peak value of abnormal magnitude is detected at the first rotational frequency (50 Hz), and it can be seen that the axis imbalance occurred in the engine based on this.

5C is a graph showing acceleration magnitudes according to frequencies appearing when an axial friction signal is applied as a failure signal.

Referring to FIG. 5C, it can be seen that when an axial friction occurs, an acceleration magnitude value at 25 Hz corresponding to 0.5 times the primary rotation frequency (50 Hz) has an abnormal peak value beyond the normal range. Therefore, when axial friction occurs, it can be seen that an abnormal peak value occurs at a frequency of 0.5 times the primary rotation frequency, and it can be seen that axial friction has occurred in the engine based on this.

6 is a graph showing a result of simulation of a failure detection method of an aircraft engine according to an embodiment of the present invention using the above-described method.

AMEsim (Advanced Modeling Environment for Simulation of engineering system) was used to simulate aircraft engine failures.

As the fault signal, the axis imbalance signal, the axis friction signal, and the oil rotation signal were applied with time difference, and the mission status of the aircraft was placed in the cruise state, and the threshold value for fault detection was set.

Figure 6a is a graph showing the magnitude of the acceleration value according to the time that occurs when the aircraft cruises in the steady state.

FIG. 6B is a graph showing an acceleration magnitude according to a frequency displayed when an axial friction signal is applied as a failure signal.

Referring to FIG. 6B, when the axial friction occurs, an acceleration magnitude value at 25 Hz corresponding to 0.5 times the primary rotation frequency (50 Hz) has an abnormal peak value beyond the normal range. Therefore, when axial friction occurs, it can be seen that an abnormal peak value occurs at a frequency of 0.5 times the primary rotation frequency, and it can be seen that axial friction has occurred in the engine based on this.

6C is a graph showing acceleration magnitudes according to frequencies appearing when an oil rotation signal is applied as a failure signal.

Referring to Figure 6c it can be seen that when the oil rotation occurs, the acceleration magnitude value at about 20Hz corresponding to 0.4 times the primary rotation frequency (50Hz) has an abnormal peak value beyond the normal range. Therefore, when oil rotation occurs, it can be seen that an abnormal peak value occurs at a frequency of 0.4 times the primary rotation frequency, and it can be seen that oil rotation has occurred in the engine based on this.

6D is a graph showing acceleration magnitudes according to frequencies appearing when an axis imbalance signal is applied as a failure signal.

Referring to FIG. 6D, it can be seen that when an axis imbalance occurs, an acceleration magnitude value at an initial rotation frequency (50 Hz) has an abnormal peak value beyond a normal range. That is, it can be seen that the peak value of abnormal magnitude is detected at the first rotational frequency (50 Hz), and it can be seen that the shaft imbalance occurred in the engine based on this.

 These failure detection signals are examples of failure signals that may occur in the engine of the aircraft. In addition to the axis imbalance signal, the axis friction signal, the oil rotation signal, other failure signals such as the mass imbalance signal, the axis alignment error signal, the bearing defect signal, etc. Can be detected using.

As described above, various signals that may occur in the engine, such as rotation signals and torque signals, other than vibration signals, may be used as signals that may be used for setting threshold values and detecting failures. It can have a value.

7 is a block diagram illustrating a failure detection apparatus for performing a failure detection method of an aircraft engine according to an exemplary embodiment of the present invention.

Referring to FIG. 7, a failure detection apparatus 700 of an aircraft engine may include a sensor unit 710, a signal processor 720, and a failure detector 730.

The sensor unit 710 attaches a sensor that can detect a signal that can be used for failure detection, such as a vibration signal or a rotation signal, to a portion capable of detecting a signal that may occur in an engine such as an engine or a crankshaft. Can be implemented.

The signal processor 720 may convert the failure detection signal generated by the sensor unit 710 into the frequency domain.

The failure detector 730 may determine whether a failure has occurred in the aircraft engine based on a failure detection signal converted into a frequency range provided from the signal processor 720 by changing a threshold value according to the mission type of the aircraft.

According to an embodiment of the present invention, a method for detecting a failure of an aircraft engine and an aircraft engine failure detecting apparatus for performing the method may be provided and used in a reciprocating engine having a four-stroke period of a small aircraft. It is possible to detect the failure of the reciprocating engine by using the vibration generated.

In the above, the present invention has been described in detail with reference to preferred embodiments, but the present invention is not limited to the above embodiments, and various modifications and changes by those skilled in the art within the spirit and scope of the present invention. This is possible.

Claims (13)

In the failure diagnosis method of the aircraft engine,
Determining a mission shape of the aircraft and determining a failure detection threshold according to the mission shape; And
And determining whether the aircraft engine has failed based on the failure detection threshold. According to claim 1, The mission shape of the aircraft,
Method for diagnosing a failure of an aircraft engine, characterized in that one of the slide, takeoff, cruise, descent, landing. The method of claim 1, wherein the failure detection threshold is:
A method for diagnosing a failure of an aircraft engine, characterized in that the value is based on a predetermined signal generated by the aircraft engine and changes according to the mission shape of the aircraft. The method of claim 1, wherein the failure detection threshold is:
And a value generated based on at least one of a vibration signal of the aircraft engine, RPM information of a crankshaft of the aircraft engine, and torque information of a crankshaft included in the aircraft engine. The method of claim 1, wherein the determining of the aircraft engine failure based on the failure detection threshold comprises:
Failure diagnosis method of the aircraft engine, characterized in that determining whether the crankshaft included in the aircraft engine failure. The method of claim 1, wherein the determining of the failure of the aircraft engine based on the failure detection threshold comprises:
A value obtained by converting a signal provided by the sensor into a frequency domain using a sensor installed in the aircraft engine and a value obtained by converting a signal provided by the sensor installed in the engine of the aircraft into a frequency domain when the aircraft engine is in a normal state Fault determination method of the aircraft engine, characterized in that for determining the failure of the aircraft engine. The method of claim 1, wherein the determining of the failure of the aircraft engine based on the failure detection threshold comprises:
And determining whether at least one of a mass imbalance, an axis alignment error, a bearing engagement error, an axis friction, an axis imbalance, and oil rotation has occurred in the aircraft engine. The method of claim 1, wherein the aircraft engine,
The fault diagnosis method of the aircraft engine, characterized in that the reciprocating engine having a natural vibration frequency due to the four strokes provided in the small aircraft. In the fault diagnosis apparatus of the aircraft engine,
A sensor unit detecting a signal generated from the aircraft engine;
A signal processing unit converting the signal generated by the sensor unit into a frequency domain; And
The signal converted into the frequency domain provided by the signal processing unit is compared with the signal converted into the frequency domain normal signal generated when the engine of the aircraft generated in the aircraft engine in the normal state according to the mission shape of the aircraft Failure diagnosis apparatus for an aircraft engine including a failure detection unit for determining whether or not the aircraft engine failure. The method of claim 8, wherein the failure detection unit,
Failure diagnosis apparatus for an aircraft engine, characterized in that determining whether the crankshaft included in the aircraft engine failure. The method of claim 8, wherein the failure detection unit,
And determining whether at least one of a mass imbalance, an axis alignment error, a bearing engagement error, an axis friction, an axis imbalance, and oil rotation has occurred in the aircraft engine. The method of claim 8, wherein the mission shape of the aircraft,
Device for diagnosing a failure of an aircraft engine, characterized in that the slide, takeoff, cruise, descent, landing. The method of claim 8, wherein the aircraft engine,
The fault diagnosis apparatus of the aircraft engine, characterized in that the reciprocating engine having a natural vibration frequency due to the four strokes provided in the small aircraft.

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