SYSTEM AND METHOD FOR DETECTING AIRFOIL CLASH WITHIN A COMPRESSOR

Abstract

A system for detecting airfoil clashing within a compressor comprises a compressor having a row of rotor blades and an adjacent row of stator vanes, an acoustic emission detection system that generates an acoustic emissions signal, a rotor blade monitoring system that generates a blade pass signal, and a processing unit that is in electronic communication with the acoustic emission detection system and the rotor blade monitoring system. The processing unit is configured to fuse the acoustic emissions signal with the blade pass signal to provide a combined signal that is indicative of an airfoil clashing event between the rotor blades and the adjacent row of stator vanes.

Description

FIELD OF THE INVENTION

The present invention generally involves an axial compressor. More specifically, the invention relates to a system and a method for detecting airfoil clash between adjacent compressor rotor blades and stator vanes disposed within the axial compressor.

BACKGROUND OF THE INVENTION

A gas turbine typically includes an axial flow compressor, one or more combustors that are disposed downstream from the compressor, a turbine that is disposed downstream from the one or more combustors and a shaft that extends axially through the gas turbine. The compressor includes an outer casing that circumferentially surrounds at least a portion of the shaft. The compressor further includes alternating rows of compressor rotor blades and stator vanes that are disposed within the outer casing. The compressor rotor blades are coupled to the shaft and extend radially outward towards the outer casing. The stator vanes are arranged annularly around the shaft and extend radially inward from the outer casing towards the shaft. A stage within the compressor generally comprises of one row of the compressor rotor blades and an axially adjacent row of the stator vanes.

In operation, a working fluid such as ambient air enters an inlet of the compressor and is progressively compressed as it flows through each stage within the compressor. The compressed working fluid is routed from the compressor to each combustor where it is mixed with a fuel and burned to produce hot combustion gases. The combustion gases are routed into the turbine. Thermal and kinetic energy are transferred from the combustion gases to multiple rows of turbine blades that are coupled to the shaft, thereby causing the shaft to rotate and produce mechanical work.

For safe and reliable operation of the gas turbine it is essential to maintain absolute and relative position within defined tolerances of the compressor rotor blades with respect to the stator vanes. Tolerances within the compressor, particularly axial tolerances between the adjacent rows of compressor rotor blades and the stator vanes are generally optimized to accommodate for some axial movement of the shaft and axial deflection of the compressor stator vanes and/or the compressor rotor blades during normal operation of the gas turbine such as when the compressor is running at a constant speed. However, during abnormal conditions such as during a sudden acceleration or a surge event the shaft may surge or move axially and the compressor rotor blades and/or the stator vanes may temporarily experience axial deflections that are outside of design tolerances. As a result, a tip portion of the stator vanes may rub or clash against a root portion of the rotating compressor rotor blades, thereby potentially resulting in damage such as cracks at the root portion of the compressor rotor blades and/or at the tip portion of the stator vanes. This event is commonly known in the industry as compressor airfoil clashing. Once formed, the cracks may propagate, thereby potentially resulting in extended shut down periods to repair or replace the damaged components.

Various systems and methods exist to monitor the condition or health and to detect damage to the compressor rotor blades and the stator vanes. However, existing systems and methods are generally ineffective at detecting the actual occurrence of the airfoil clashing event which may allow an operator to modify a maintenance schedule of the gas turbine. In addition, by alerting an operator or a monitoring station in real-time, appropriate measures may be taken to reduce potential damage to the rotor blades and/or the stator vanes. Therefore, it would be desirable to have a system and method for detecting the airfoil clashing event in real-time.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention are set forth below in the following description, or may be obvious from the description, or may be learned through practice of the invention.

One embodiment of the present invention is a system for detecting airfoil clashing within a compressor. The system comprises a compressor having a row of rotor blades and an adjacent row of stator vanes, an acoustic emission detection system that generates an acoustic emissions signal, a rotor blade monitoring system that generates a blade pass signal, and a processing unit that is in electronic communication with the acoustic emission detection system and the rotor blade monitoring system. The processing unit is configured to fuse the acoustic emissions signal with the blade pass signal to provide a combined signal that is indicative of an airfoil clashing event between the rotor blades and the adjacent row of stator vanes.

Another embodiment of the present invention is a gas turbine. The gas turbine includes a compressor having an outer casing that surrounds a row of rotor blades coupled to a shaft and an adjacent row of stator vanes arranged in an annular array around the shaft. A combustor is disposed downstream from the compressor and a turbine disposed downstream from the combustor. The gas turbine further includes a system for detecting airfoil clashing between the row of rotor blades and the row of stator vanes. The system comprises an acoustic emission detection system that generates an acoustic emissions signal, a rotor blade monitoring system that generates a blade pass signal, and a processing unit that is in electronic communication with the acoustic emission detection system and the rotor blade monitoring system. The processing unit fuses the acoustic emissions signal with the blade pass signal to provide a combined signal that is indicative of an airfoil clashing event between the rotor blades and the adjacent row of stator vanes.

The present invention may also include a method for detecting airfoil clashing between a row of rotor blades and an adjacent row of stator vanes within a compressor. The method includes sensing acoustic emissions from the stator vanes, generating an acoustic emissions signal from the sensed acoustic emissions, sensing time of arrival information of the rotor blades, generating a blade pass signal from the sensed time of arrival information, communicating the acoustic emission signal and the blade pass signal to a processing unit. The method further includes analyzing the acoustic emission signal and the blade pass signal simultaneously to detect an airfoil clashing event.

Those of ordinary skill in the art will better appreciate the features and aspects of such embodiments, and others, upon review of the specification.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including the best mode thereof to one skilled in the art, is set forth more particularly in the remainder of the specification, including reference to the accompanying figures, in which:

FIG. 1 provides an example of a gas turbine as may incorporate various embodiments of the present invention;

FIG. 2 provides a diagrammatic illustration of a system for detecting airfoil clashing according to at least one embodiment of the present invention;

FIG. 3 provides a diagrammatic illustration of the system for detecting airfoil clashing as shown in FIG. 2, according to at least one embodiment of the present invention; and

FIG. 4 provides a flow chart for detecting an airfoil clashing event between a row of rotor blades and an adjacent row of stator vanes within a compressor, according to one embodiment of the present invention.

DETAILED DESCRIPTION OF THE INVENTION

Reference will now be made in detail to present embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the invention. As used herein, the term “radially” refers to the relative direction that is substantially perpendicular to an axial centerline of a particular component, and the term “axially” refers to the relative direction that is substantially parallel to an axial centerline of a particular component.

Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that modifications and variations can be made in the present invention without departing from the scope or spirit thereof. For instance, features illustrated or described as part of one embodiment may be used on another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents.

Although exemplary embodiments of the present invention will be described generally in the context of an axial flow compressor used in an industrial gas turbine for purposes of illustration, one of ordinary skill in the art will readily appreciate that embodiments of the present invention may be applied to any device having a row of rotating blades that is positioned adjacent to a row of stationary or stator vanes and is not limited to an axial-flow compressor unless specifically recited in the claims. For example, the present invention may be incorporated into a compressor of a jet engine, a high speed ship engine, a small scale power station, or the like. In addition, the present invention may be incorporated into a compressor used in varied applications, such as large volume air separation plants, blast furnace applications, propane dehydrogenation, or the like.

Referring now to the drawings, wherein like numerals refer to like components, FIG. 1 illustrates an example of a gas turbine 10 as may incorporate various embodiments of the present invention. As shown, the gas turbine 10 generally includes an axial flow compressor 12, a combustion section 14 disposed downstream from the compressor 12 and a turbine 16 disposed downstream from the combustion section 14. The compressor 12 generally includes multiple rows 18 of rotor blades 20 arranged circumferentially around a shaft 22 that extends at least partially through the gas turbine 10. The compressor 12 further includes multiple rows 24 of stator vanes 26 arranged circumferentially around the shaft 22. The stator vanes may be fixed to an outer casing 28 that extends circumferentially around the rows 18 of the rotor blades 20. The compressor 12 may also include one or more rows of actuatable inlet guide vanes 30 disposed substantially adjacent to an inlet 32 to the compressor 12. The combustion section 14 includes at least one combustor 34. The shaft 22 may extend axially between the compressor and the turbine 16.

In normal operation, air 36 is drawn into the inlet 32 of the compressor 12 and is progressively compressed to provide a compressed air 38 to the combustion section 18. The compressed air 38 is mixed with fuel in the combustor 34 to form a combustible mixture. The combustible mixture is burned in the combustor 34, thereby generating a hot gas 40 that flows from the combustor 34 across a row of turbine nozzles 42 and into the turbine section 16. The hot gas 38 rapidly expands as it flows across alternating stages of turbine blades 44 connected to the shaft 22 and the turbine nozzles 42. Thermal and/or kinetic energy is transferred from the hot gas 40 to each stage of the turbine blades 44, thereby causing the shaft 22 to rotate and produce mechanical work. The shaft 22 may be coupled to a load such as a generator (not shown) so as to produce electricity. In addition or in the alternative, the shaft 22 may drive the compressor section 12 of the gas turbine.

During certain operational events, such as during a sudden loss of load on the gas turbine, rotational speed of the shaft 22 may suddenly increase or surge, thereby causing a momentary axial surge or movement of the shaft 22 and/or static and/or dynamic deflection of the rotor blades 20 and/or the stator vanes 26. In certain instances the axial surge or movement of the shaft 22 and/or static and/or dynamic deflection of the rotor blades 20 and/or the stator vanes 26 exceeds axial clearance design tolerances. As a result, a root portion (not shown) of the rotor blades 20 may strike or rub a tip portion (not shown) of the stator vanes 26. This event is commonly known in the art as an airfoil clashing event. An airfoil clashing event may result in damage such as cracking to either or both of the rotor blades 20 or the stator vanes 26.

FIG. 2 provides a diagrammatic illustration of an exemplary system 100 for detecting an airfoil clashing event between the rotor blades 20 and the stator vanes 26 within the compressor 12, herein referred to as “system 100”, according to various embodiments of the present invention. As shown in FIG. 2, the system 100 includes the compressor 12, at least one row 18 of the rotor blades 20, an adjacent row 24 of the stator vanes 26, an acoustic emissions detection system 200 for monitoring acoustic emission (AE) waves propagating through the stator vanes 26, a rotor blade monitoring system 300 for monitoring time of arrival (TOA) information corresponding to the rotation of the rotor blades 20 that is representative of at least one of static or dynamic deflections of the rotor blades 20, and a controller 400 that is in electronic communication with the acoustic emission detection system 200 and the rotor blade monitoring system 300.

As shown in FIG. 2, the acoustic emission detection system 200 includes a plurality of sensing devices 202 that are dispersed on an outer surface 204 of the outer casing 28. The acoustic emission detection system 200 may also include circuitry (not shown) commonly available for sensing acoustic emissions such as pressure transients, acoustic frequencies, magnitude of acoustic energy release, and/or shock waves produced during a compressor airfoil clashing event and/or during crack initiation and/or propagation. The sensing devices 202 may include a magnetostrictive material sensing device, a piezoelectric sensing device and/or a capacitive sensing device that converts stress waves to electrical signals 206 respectively. The sensing devices 202, for example, may include an optical sensing device, an acoustic emission sensing device, a radio frequency wireless sensing device or the like. The sensing devices 202 may be arranged circumferentially around the outer surface 204 of the outer casing 28.

It is noted that the present acoustic emission detection system 200 may include an optimal number of the sensing devices 202 based upon the size of the compressor 12 and precision expected in monitoring the stator vanes 26. An optimal location or position of the plurality sensing devices 202 on the outer surface 204 of the outer casing 28 may be determined using triangulation techniques. As used herein, the term “optimal location” refers to locations for distribution of the plurality of sensing devices 202 on the outer surface 204 of the outer casing 28 such that acoustic emission (AE) waves generated by each of the stator vanes 26 is captured by the sensing devices 202.

When the compressor 12 is operating under stress or under uncharacteristic operating conditions such as during an airfoil clashing event, the one or more of the stator vanes 26 generate acoustic emission (AE) waves. The AE waves travel or propagate through the stator vanes 26 and the outer casing 28 to reach the outer surface 202 of the outer casing 28. When these AE waves reach the outer surface 202 of the outer casing 28, the sensing devices 202 measure the AE waves to generate acoustic emission or AE signals 206. The AE signals 206 may then be communicated to the controller 400. The AE signals 206 may be a time-series signals in voltage.

In certain embodiments, the AE signals 206 may be preprocessed by one or more intermediate devices 208 before being communicated to the controller 400. The intermediate devices 208, for example, may include an amplifier, an interface unit, a data acquisition system, a processing unit and the like. The one or more intermediate devices 208 may increase the strength and quality of the AE signals 206 and/or may execute computer algorithms to combine or fuse multiple AE signals 206 to provide a sensor level AE signal 210 to the controller 400. The AE signals 206, 210 may include information that is indicative of amplitude of the AE waves, magnitude of acoustic energy release, duration of AE waves, as well as any other acoustic emissions information that may be indicative of an airfoil clashing event.

The rotor blade monitoring system 300, as shown in FIG. 2, includes at least one blade pass sensor 302 for sensing time of arrival or TOA of the rotor blades 20 with respect to a reference mark (not shown) on the shaft 22 (FIG. 1), the rotor blades 20 and/or a keyphasor. A keyphasor may comprise, for example, a proximity switch used to identify the beginning and completion of each revolution of the rotor blades 20. The operation of a keyphasor is well known to those skilled in the art and is discussed in greater detail in THE KEYPHASOR—A Necessity for Machinery Diagnosis, Bentley Nev., November, 1977.

In one configuration of the rotor blade monitoring system 300, the blade pass sensor 302 may comprise two types of sensors. For example, the blade pass sensors 302 may include a leading edge sensor 304 and trailing edge sensor 306. The leading edge sensors 304 sense TOA of a leading edge 308 of the rotor blades 20 whereas the trailing edge sensors 306 sense TOA of a trailing edge 310 of the rotor blades 28. A discrepancy or lag in TOA of the leading edge 308 and the TOA of trailing edge 310 may indicate blade twist and/or static and dynamic deflection of the rotor blades 20 which may be associated with an airfoil clashing event.

The blade pass sensors 302 including the leading edge sensors 304 and the trailing edge sensors 306 may be mounted on the outer surface 204 of the outer casing 28 in a position such that arrival of each rotor blade 20 can be sensed efficiently. In one embodiment, the blade pass sensors 302 may sense TOA of at least one rotor blade 20 by taking a keyphasor as a reference for determination of completion of each revolution of the rotor blades 20 and then communicating a corresponding blade pass signal 312 to the controller 400.

The number of the blade pass sensors 302 and the number of each leading edge and trailing edge sensors 304, 306 depends on a variety of factors including, for example, the number of rows 18 of the rotor blades 20 and the number of rows 24 of the stator vanes 26 in the compressor 12. The leading edge and trailing edge sensors 304, 306 may be of similar or different types. For example, the blade pass sensors 302, the leading edge sensors 304 and the trailing edge sensors 306 may be capacitive, eddy current and/or magnetic sensors.

The blade pass signals 312 may be preprocessed by one or more intermediate devices 314 before being communicated to the controller 400. The intermediate devices 314, for example, may include an amplifier, an interface unit, a data acquisition system, a sensor level fuser and the like. The one or more intermediate devices 314 may increase the strength and quality of the blade pass signals 312 and/or may combine or fuse multiple blade pass signals 312 to provide a sensor level blade pass signal 316 to the controller 400. The blade pass signals 312, 316 that are communicated to the controller 400 may include information that is indicative of rotor blade twist, static deflection of the rotor blades 20, dynamic deflection of the rotor blades 20 or any other information that may be indicative of a compressor airfoil clashing event.

FIG. 3 provides a diagrammatic illustration of an exemplary controller 400 as shown in FIG. 2, as may be incorporated in the present invention. The controller 400 may include various components such as microprocessors 402, coprocessors, and/or memory/media elements 404 that store data, store software instructions, and/or execute software instructions. The various memory/media elements 404 may be one or more varieties of computer readable media, such as, but not limited to, any combination of volatile memory (e.g., RAM, DRAM, SRAM, etc.), non-volatile memory (e.g., flash drives, hard drives, magnetic tapes, CD-ROM, DVD-ROM, etc.), and/or other memory devices (e.g., diskettes, magnetic based storage media, optical storage media, etc.). Any possible variations of data storage and processor configurations will be appreciated by one of ordinary skill in the art.

During a surge event such as during a sudden loss of load on the gas turbine, an instantaneous increase of the rotational speed of the shaft 22 will occur. In particular instances, the surge event results in axial movement of the shaft 22 and/or static and/or dynamic deflection of the rotor blades 20 which exceeds an allowable deflection threshold value which is predetermined by design tolerances. When the static and/or dynamic deflection of the rotor blades 20 exceeds the deflection threshold, the rotor blades 20 will strike or rub against the stator vanes 26, thereby resulting in a compressor airfoil clashing event.

The airfoil clashing event will result in a simultaneous anomaly in the blade pass signals 312 and/or the sensor level blade pass signal 316, and in the AE signals 206 and/or the sensor level AE signals 210 that are communicated to the controller 400. The blade pass signals 312 and/or the sensor level blade pass signal 316 will contain information indicative of a momentary static and/or dynamic deflection of the rotor blades 20 which exceeds the allowable deflection threshold value. For example, the airfoil clashing event will cause an instantaneous change in the TOA of the rotor blades 20 as detected by the blade pass sensors 302, 304 and 306. Simultaneously, the airfoil clashing event will cause a sudden release of acoustic energy as the metallic rotor blades 20 strike the metallic stator vanes 26.

The AE signals 206 and/or the sensor level AE signals 210 provided by the AE sensors 202 will contain information indicative of at least one of a sudden change in amplitude of the AE waves, a sudden release of acoustic energy from the stator vanes, a change in the duration of the AE waves being emitted from the stator vanes, as well as any other AE wave or other acoustic emissions information that may be indicative of an anomalous occurrence indicative of airfoil clashing at the stator vanes.

In particular embodiments, the controller 400 executes an algorithm which combines or fuses the AE signals 206 and/or the sensor level AE signals 210 with the blade pass signals 312 and/or the sensor level blade pass signals 316 to detect if the compressor airfoil clashing event has occurred. The controller 400 then generates an output signal 406 such as an alarm signal 408 if the fusion of the AE signals 206 and/or the sensor level AE signals 210 with the blade pass signals 312 and/or the sensor level blade pass signals 316 indicate simultaneous anomalies in both the AE signals 206, 210 and the blade pass signals 312, 316. Fusion of these signals can be accomplished by using various techniques. For example, fusion of the AE signals 206, 210 and the blade pass signals 312, 316 can be accomplished using Fuzzy, Bayesian methods, Demster methods, Demster-shafer methods, neural network methods, tree logic methods, voting logic methods or rule based fusion methodologies.

The controller 400 may use each, some or any combination of the blade pass signals 312, 316 the AE signals 206, 210 or the output signal 406 to determine the particular location or rows 18, 24 (FIG. 2) within the compressor 12 (FIG. 2) where the airfoil clashing event occurred. For example, impacted rotor blades 20 may be detected by analyzing information indicative of TOA that is present in the blade pass signals 312, 316 communicated to the controller 400. The location of the affected stator vanes 26 may be determined by triangulation methods. Severity of the airfoil clashing event may be determined based upon signal strength of at least one of the AE signals 206, 210 or the blade pass signals 312, 316.

In particular embodiments, as shown in FIG. 2, the system 100 further includes one or more operation parameter sensors 500 disposed on and/or around the gas turbine 10 (FIG. 1). The operation parameter sensors 500 provide real-time or near real-time measurements of various operating parameters of the compressor 12 or other associated equipment such as the combustors 34 (FIG. 1) or the turbine 16 (FIG. 1) which operate in conjunction with the compressor 12.

Commonly measured operating parameters of the compressor 12 may include, for example, compressor discharge temperature, compressor pressure ratio, inlet guide vane angle, bearing temperatures, bearing vibrations, rotor vibrations, etc. Commonly measured operating parameters of associated equipment may include, for example, gas turbine load, fuel stroke reference, shaft speed, turbine exhaust temperatures, etc. Each operation parameter sensor 500 transmits a parameter signal 502 reflective of the operating parameter to the controller 400 for further processing.

The controller 400 may process/monitor the parameter signals 502 to detect anomalies in the parameter signals 502 such as a sudden increase in the shaft speed, malfunctioning inlet guide vanes or an increase in compressor pressure ratio that may be indicative of an upcoming or occurring surge event. The controller 400 may communicate and/or utilize this information to confirm the occurrence of the airfoil clashing event and/or to determine a root cause of the airfoil clashing event. The controller 400 may also use the output signal 406 to trip the gas turbine 10 and/or move the gas turbine compressor 12 out of the surge condition to avoid or reduce damage to the rotor blades 20 and/or the stator vanes. The output signal 406 may drive an alarm circuit 408, actuate a safety circuit, or trigger a combination of the two to ensure prompt operator action to address the situation.

The parameter signal(s) 502, the blade pass signals 312, 316 the AE signals 206, 210 and/or the output signal 406 may be communicated to a visualization device 410 such as an LCD or CRT monitor for real-time visualization and monitoring of the system. In addition or in the alternative, any or all of the AE signals 206, 210, the blade pass signals 312, 316, the parameter signal(s) 502 and/or the output signal 406 may be transmitted to one or more off-site data storage and/or monitoring facilities 600 via a wired or wireless communication network.

The system 100 illustrated in FIGS. 2 and 3 and as described herein, may be used to provide a method 700 for detecting an airfoil clashing event between the row 18 of the rotor blades 20 and a row 24 of the stator vanes 26 within the compressor 12. FIG. 4 provides a flow diagram of the method 700 for detecting an airfoil clashing event between the row 18 of the rotor blades 20 and a row 24 of the stator vanes 26 within the compressor. Specifically, as provided in step 702, the AE sensors 202 are used to sense the acoustic emission waves from the stator vanes 26. At step 704, AE signals 206 are generated from the sensed acoustic emissions. The AE signals 206 may be preprocessed by one or more of the intermediate devices 208. For example, multiple AE signals 206 may be amplified, filtered, fused or combined into a single AE signal 210.

At step 706, the method 700 further includes using the blade pass sensors 302 to sense the time of arrival (TOA) information of the rotor blades 20. At step 708 the method 700 includes generating the blade pass signal 312 from the sensed TOA information. The blade pass signal 312 may be preprocessed by one or more of the intermediate devices 314. For example, the blade pass signals 312 from the blade pass sensors 302 may be amplified, filtered, fused or combined into a single blade pass signal 316.

At step 710, the AE signal 206, 210 and the blade pass signal 312, 316 are then communicated to the processing unit 400. At step 712, the processing unit 400 executes one or more algorithms to analyze the AE signal 206, 210 and the blade pass signal 312, 316 simultaneously to detect an airfoil clashing event. For example, the algorithm may identify a static and/or dynamic deflection that exceeds a predetermined design threshold that occurs simultaneously or nearly simultaneously as an anomalous AE signal 206, 210 that may correspond to a sudden release of acoustic energy, a spike in frequency, etcetera.

In particular embodiments, the AE signals 206, 210 include information that is indicative of at least one of frequency amplitude, magnitude of acoustic energy release or duration of acoustic energy release. The method 700 may further include fusing or combining the AE signals 206, 210 with the blade pass signals 312, 316 to provide the output signal 406 that is indicative to the airfoil clashing event. In further embodiments, the method 700 includes sensing one or more of the operating parameters of the gas turbine 10 to generate the parameter signal 502, and communicating the parameter signal 502 to the processing unit 400 to predict an impending airfoil clashing event and/or to determine the root cause of the airfoil clashing event.

The method 700 may further include communicating at least one of the output signal 406, the AE signal 206, 210, the blade pass signal 312, 316, the operating parameter signal 502 or the alarm signal 408 to the off-site data storage and/or monitoring facilities 600 via a wired or wireless communication network. In further embodiments, the method 700 may include communicating at least one of the parameter signal 502, the blade pass signals 312, 316 the AE signals 206, 210 and/or the output signal 406 to a visualization device 410 such as an LCD or CRT monitor for real-time visualization and monitoring of the system 100. The method 700 also may include using the output signal 406 and/or the alarm signal 408 to cause the gas turbine 10 to trip, reduce in speed and/or shut down based on the detection and/or prediction of the clashing event. This step may be accomplished using controls logic or other controls based algorithms.

This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.

Claims

1. A system for detecting airfoil clashing within a compressor, comprising:

a. a compressor having a row of rotor blades and an adjacent row of stator vanes;

b. an acoustic emission detection system that generates an acoustic emissions signal;

c. a rotor blade monitoring system that generates a blade pass signal; and

d. a processing unit in electronic communication with the acoustic emission detection system and the rotor blade monitoring system, wherein the processing unit fuses the acoustic emissions signal with the blade pass signal to provide a combined signal that is indicative of an airfoil clashing event between the rotor blades and the adjacent row of stator vanes.

2. The system as in claim 1, wherein the airfoil clashing event corresponds to an acoustic emissions signal that exceed a predetermined threshold.

3. The system as in claim 2, wherein the acoustic emission signal includes information that is indicative of at least one of amplitude, magnitude of acoustic energy release or duration of acoustic energy release.

4. The system as in claim 1, wherein the blade pass signal includes information that is indicative of time of arrival of the rotor blades that falls outside of a predetermined threshold.

5. The system as in claim 1, wherein the blade monitoring system identifies rotor blade position of damaged rotor blades within the row of rotor blades.

6. The system as in claim 1, wherein the processing unit generates an alarm signal that corresponds to the airfoil clashing event.

7. The system as in claim 1, wherein the compressor is a component of a gas turbine, the system further comprising one or more operating sensors that detect one or more operating parameters of the gas turbine.

8. The system as in claim 7, wherein the processing unit determines root cause for the airfoil clashing event based on the detected one or more operating parameters.

9. A gas turbine, comprising:

a. a compressor having an outer casing that surrounds a row of rotor blades coupled to a shaft and an adjacent row of stator vanes arranged in an annular array around the shaft;

b. a combustor disposed downstream from the compressor;

c. a turbine disposed downstream from the combustor; and

d. a system for detecting airfoil clashing between the row of rotor blades and the row of stator vanes, the system comprising: i. an acoustic emission detection system that generates an acoustic emissions signal; ii. a rotor blade monitoring system that generates a blade pass signal; and iii. a processing unit in electronic communication with the acoustic emission detection system and the rotor blade monitoring system, wherein the processing unit fuses the acoustic emissions signal with the blade pass signal to provide a combined signal that is indicative of an airfoil clashing event between the rotor blades and the adjacent row of stator vanes.

10. The gas turbine as in claim 9, wherein the airfoil clashing event corresponds to an acoustic emissions signal that exceed a predetermined threshold.

11. The gas turbine as in claim 10, wherein the acoustic emission signal includes information that is indicative of at least one of amplitude, magnitude of acoustic energy release or duration of acoustic energy release.

12. The gas turbine as in claim 9, wherein the blade pass signal includes information that is indicative of time of arrival of the rotor blades that falls outside of a predetermined threshold.

13. The gas turbine as in claim 9, wherein the blade monitoring system identifies rotor blade position of damaged rotor blades within the row of rotor blades.

14. The gas turbine as in claim 9, wherein the processing unit generates an alarm signal that corresponds to the airfoil clashing event.

15. The gas turbine as in claim 9, further comprising one or more operating sensors that detect one or more operating parameters of the gas turbine.

16. The gas turbine as in claim 15, wherein the processing unit determines root cause for the airfoil clashing event based on the detected one or more operating parameters.

17. A method for detecting airfoil clashing between a row of rotor blades and an adjacent row of stator vanes within a compressor, comprising:

a. sensing acoustic emissions from the stator vanes;

b. generating an acoustic emissions signal from the sensed acoustic emissions;

c. sensing time of arrival information of the rotor blades;

d. generating a blade pass signal from the sensed time of arrival information;

e. communicating the acoustic emission signal and the blade pass signal to a processing unit; and

f. analyzing the acoustic emission signal and the blade pass signal simultaneously to detect an airfoil clashing event.

18. The method as in claim 17, wherein the acoustic emissions signal includes information that is indicative of at least one of frequency amplitude, magnitude of acoustic energy release or duration of acoustic energy release.

19. The method as in claim 17, further comprising fusing the acoustic emissions signal with the blade pass signal to provide at least one of an output signal or an alarm signal indicative to an actual or predicted airfoil clashing event.

20. The method as in claim 19, further comprising tripping the gas turbine based on at least one of the output signal or the alarm signal.