BACKGROUND OF THE INVENTION

[0001] This invention relates generally to gas turbine engines, and, more specifically, to a system for detection of a stall in a compression system therein.

[0002] In a turbofan aircraft gas turbine engine, air is pressurized in a compression system, comprising a fan module, a booster module and a compression module during operation, in large turbo fan engines, the air passing through the fan module is mostly passed into a by-pass stream and used for generating the bulk of the thrust needed for propelling an aircraft in flight. The air channeled through the booster module and compression module is mixed with fuel in a combustor and ignited, generating hot combustion gases which flow through turbine stages that extract energy therefrom for powering the fan, booster and compressor rotors. The fan, booster and compressor modules have a series of rotor stages and stator stages. The fan and booster rotors are typically driven by a low pressure turbine and the compressor rotor is driven by a high pressure turbine. The fan and booster rotors are aerodynamicaliy coupled to the compressor rotor although these normally operate at different mechanical speeds.

[0003] Operability in a wide range of operating conditions is a fundamental requirement in the design of compression systems, such as fans, boosters and compressors. Modem developments in advanced aircrafts have required the use of engines buried within the airframe, with air flowing into the engines through inlets that have unique geometries that cause severe distortions in the inlet airflow. Some of these engines may also have a fixed area exhaust nozzle, which limits the operability of these engines. Fundamental in the design of these compression systems is efficiency in compressing the air with sufficient stall margin over the entire flight envelope of operation from takeoff, cruise, and landing. However, compression efficiency and stall margin are normally inversely related with increasing efficiency typically corresponding with a decrease in stall margin. The conflicting requirements of stall margin and efficiency are particularly demanding in high performance jet engines that operate under challenging operating conditions such as severe inlet distortions, fixed area nozzles and increased auxiliary power extractions, while still requiring high a level of stability margin throughout the flight envelope.

[0004] Stalls are commonly caused by flow breakdowns at the tip of the rotor blades of compression systems such as fans, compressors and boosters, in gas turbine engine compression system rotors, there are tip clearances between rotating blade tips and a stationary casing or shroud that surrounds the blade tips. During the engine operation, air leaks from the pressure side of a blade through the tip clearance toward the suction side. These leakage flows may cause vortices to form at the tip region of the blade, A tip vortex can grow and spread when there are severe inlet distortions in the air flowing into compression system or when the engine is throttled and lead to a compressor stall and cause significant operability problems and performance losses, Conventional gas turbine engine compression systems are operated using control systems that protect against stall using worst-case assumptions. The conventional systems assume that the compressor is fully deteriorated and limit the acceleration rate, etc. of the engine based upon this. Active management systems that can detect an impending stall have been considered for some applications. The durability and temperature capability limitations of some conventional sensors may limit the use of such systems in engine operating environments.

[0005] Accordingly, it would be desirable to have the ability to measure and control dynamic processes such as flow instabilities in a compression system. It would be desirable to have a system that can measure operating parameters related to the onset of flo w instabilities and process the measured data to predict the onset of stall in a stage of a compression system, such as a fan, booster or compressor. It would be desirable to have a system wherein operational limits and tip clerances can be adjusted to take full advantage of the capabilities of a specific engine, rather than penalizing performance to account for the most deteriorated engine that may be in the field. It would also be desirable to have a system to mitigate compression system stalls based on the measurement system output, for certain flight maneuvers at critical points in the flight envelope, allowing the maneuvers to be completed without stall or surge. BRIEF DESCRIPTION OF THE INVENTION

[0006] The above-mentioned need or needs may be met by exemplary embodiments which provide a system for detecting onset of a stall in a rotor, the system comprising a plasma sensor spaced radially outwardly and apart from tips of a circumferential row of blades at a location on a static component that is between a first location and a second location wherein the first location is at a first distance of about 25% blade tip-chord length axially forward from the leading edge of a blade and the second location is at a second distance of about 25% blade tip-chord length axially aft from the trailing edge of a blade and wherein the sensor is capable of generating an input signal corresponding to a fl ow parameter at a l ocation near the tip of a blade and indicative of the onset of a stall and a correlation processor that generates a stability correlation signal.

[0007] In another embodiment, a system for detecting onset of a stall comprises a pressure sensor capable of generating a signal corresponding to the pressure at a location near the blade tip, In another embodiment, a system for detecting onset of a stall comprises a temperature sensor capable of generating a signal corresponding to the temperature at a location near the blade tip, In another embodiment, a system for detecting onset of a stall comprises a velocity sensor capable of generating a signal corresponding to the velocity at a location near the blade tip, In another embodiment, a system for detecting onset of a stall comprises an entropy sensor capable of generating a signal corresponding to the entropy at a location near the blade tip.

[0008] In another embodiment, a system for controlling an operating tip clearance between a rotor and a static component comprises a plasma sensor, a correlation processor capable of generating a correlation signal indicative of the operating tip clearance based on an input signal from the plasma sensor and a control system capable of controlling the operating tip clearance based on the correlation signal. BRIEF DESCRIPTION OF THE DRAWINGS

[0009] The subject matter which is regarded as the invention is particularly pointed out and distinctly claimed in the concluding part of the specification. The invention, however, may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which:

[0010] Figure 1 is a schematic cross-sectional view of a gas turbine engine with an exempl ary embodiment of the present invention.

[0011] Figure 2 is an enlarged cross-sectional view of a portion of the fan section of the gas turbine engine shown in Figure 1 .

[0012] Figure 3 is an exemplary operating map of a compression system in the gas turbine engine shown in Figure 1 ,

[0013] Figure 4a shows the formation of a region with blade tip vortex in a fan stage.

[0014] Figure 4b shows the spread of the blade tip vortex shown in Figure

4a.

[0015] Figure 4c shows the vortex flow at blade tip region during a stall.

[0016] Figure 5 is a schematic cross-sectional view of the tip region of a fan with an exemplary embodiment of a stall detection system.

[0017] Figure 6 is a schematic sketch of an exemplary arrangement of multiple sensors for a stall detection system.

[0018] FIG. 7 is a schematic sketch of exemplary locations of sensors in a rotor stall sensor system.

[0019] FIG. 8 is an exemplary time history of pressure from an unsteady CFD simulation of a compression system approaching stall. [0020] FIG. 9 is an exemplary time history of temperature from an unsteady CFD simulation of a compression system approaching stall.

[0021] FIG. 10 is an exemplary time history of velocity from an unsteady CFD simulation of a compression system approaching stall.

[0022] FIG. 11 is an exemplary time history of entropy from an unsteady CFD simulation of a compression system approaching stall.

[0023] FIG. 12 is a schematic cross-sectional view of the tip region of a compressor with an exemplary embodiment of a stall detection system having a plasma sensor.

DETAILED DESCRIPTION OF THE INVENTION

[0024] Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views, Figure 1 shows an exemplary turbofan gas turbine engine 10 incorporating an exemplary embodiment of the present invention. It comprises an engine centerline axis 8, fan section 12 which receives ambient air 14, a high pressure compressor (HPC) 18, a combustor 20 which mixes fuel with the air pressurized by the HPC 18 for generating combustion gases or gas flow which flows downstream through a high pressure turbine (HPT) 22, and a low pressure turbine (LPT) 24 from which the combustion gases are discharged from the engine 10. Many engines have a booster or low pressure compressor (not shown in Figure 1 ) mounted between the fan section and the H PC. A portion of the air passing through the fan section 12 is bypassed around the high pressure compressor 18 through a bypass duct 21 having an entrance or splitter 23 between the fan section 12 and the high pressure compressor 18. The HPT 22 is joined to the HPC 18 to substantially form a high pressure rotor 29. A low pressure shaft 28 joins the LPT 24 to the fan section 12 and the booster if one is used. The second or low pressure shaft 28 is rotatably disposed co-axially with and radially inwardly of the first or high pressure rotor. In the exemplar}' embodiments of the present invention shown in Figures 1 and 2, the fan section 12 has a multi-stage fan rotor, as in many gas turbine engines, illustrated by first, second, and third fan rotor stages 12a, 12b, and 12c respectively,

[0025] The fan section 12 that pressurizes the air flowing through it is axisymmetrical about the longitudinal centerline axis 8. The fan section 12 includes a plurality of inlet guide vanes (IGV) 30 and a plurality of stator vanes 31 arranged in a circumferential direction around the longitudinal centerline axis 8. The multiple fan rotor stages 12 of the fan section 12 have corresponding fan rotor blades 40a, 40b, 40c extending radially outwardly from corresponding rotor hubs 39a, 39b, 39c in the form of separate disks, or integral blisks, or annular drums in any conventional manner.

[0026] Cooperating with a fan rotor stage 12a, 12b, 12c is a corresponding stator stage comprising a plurality of circumferentially spaced apart stator vanes 31a, 31b, 31c. The arrangement of stator vanes and rotor blades is shown in Figure 2. The rotor blades 40 and stator vanes 31 define airfoils having corresponding aerodynamic profiles or contours for pressurizing the airflow successively in axial stages. Each fan rotor blade 40 comprises an airfoil 34 extending radially outward from a blade root 45 to a blade tip 46, a pressure side 43, a suction side 44, a leading edge 41 and a trailing edge 42. The airfoil 34 extends in the chordwise direction between the leadmg edge 41 and the trailing edge 42. A chord C of the airfoil 34 is the length between the leading 41 and trailing edge 42 at each radial cross section of the blade. The pressure side 43 of the airfoil 34 faces in the general direction of rotation of the fan rotors and the suction side 44 is on the other side of the airfoil. The front stage rotor blades 40 rotate within an annular casing 50 that surrounds the rotor blade tips. The aft stage rotor blades typically rotate within an annular passage formed by shroud segments 51 that are circumferentially arranged around the blade tips 46. In operation, pressure of the air is increased as die air decelerates and diffuses through the stator and rotor airfoils.

[0027] Operating map of an exemplary compression system, such as the fan section 12 in the exemplary gas turbine engine 10 is shown in Figure 3, with inlet corrected flow rate along the X- axis and the pressure ratio on the Y-axis. Operating lines 1 14, 1 16 and the stall line 1 12 are shown, along with constant speed lines 122, 124, Line 124 represents a lower speed line and line 122 represents a higher speed line. As the compression system is throttled at a constant speed, such as constant speed line 124, the inlet corrected flow rate decreases while the pressure ratio increases, and the compression system operation moves closer to the stall line 112. Each operating condition has a corresponding compressor efficiency, conventionally defined as the ratio of ideal (isentropic) compressor work input to actual work input required to achieve a given pressure ratio. The compressor efficiency of each operating condition is plotted on the operating map in the form of contours of constant efficiency, such as items 118, 120 shown in Figure 3. The performance map has a region of peak efficiency, depicted in Figure 3 as the smallest contour 120, and it is desirable to operate the compression systems in the region of peak efficiency as much as possible. Flow distortions in the inlet air flow 14 which enters the fan section 12 tend to cause flow instabilities as the air is compressed by the fan blades (and compression system blades) and the stall line 1 12 will tend to drop lower. As explained further below herein, the exemplary embodiments of the present invention provide a system for detecting the flow instabilities in the fan section 12, such as from flow distortions, and processing the information from the fan section to predict an impending stall in a fan rotor. The embodiments of the present invention shown herein enable other systems in the engine which can respond as necessary to manage the stall margin of fan rotors and other compression systems.

[0028] Stalls in fan rotors due to inlet flow distortions, and stalls in other compression systems that are throttled, are known to be caused by a breakdown of flow in the tip region 52 of rotors, such as, for example, the fan rotors 12a, 12b, 12c shown in Figure 2. This tip flow breakdown is associated with tip leakage vortex schematically shown in Figures 4a, 4b and 4c as contour plots of regions having a negative axial velocity, based from computational fluid dynamic analyses. Tip leakage vortex 200 initiates primarily at the rotor blade tip 46 near the leading edge 41. in the region of this vortex 200, there exists flow that has negative axial velocity, that is, the flow in this region is counter to the main body of flow and is highly undesirable. Unless interrupted, the tip vortex 200 propagates axiaily aft and tangentially from the blade suction surface 44 to the adjacent blade pressure surface 43 as shown in Figure 4b. Once it reaches the pressure surface 43, the flow tends to collect in a region of blockage at the tip between the blades as shown in Figure 4c and causes high loss. As the inlet flow distortions become severe, or as a compression system is throttled, the blockage becomes increasingly larger within the flow passage between the adjacent blades and eventually becomes so large as to drop the rotor pressure ratio below its design level, and causes the fan rotor to stall Near stall, the behavior of the blade passage flow field structure, specifically the blade tip clearance vortex trajectory, is perpendicular to the axial direction wherein the tip clearance voxtex 200 spans the leading edges 41 of adjacent blades 40, as shown in Figure 4c. The vortex 200 starts from the leading edge 41 on the suction surface 44 of the blade 40 and moves towards the leading edge 41 on the pressure side of the adjacent blade 40 as shown in Figure 4c.

[0029] The ability to control a dynamic process, such as a flow instability in a compression system, requires a measurement of a characteristic of the process. A continuous measurement or samples of sufficient number of discrete measurements. In order to mitigate compression system stalls for certain flight maneuvers at critical points in the flight envelope where the stability margin is small or negative, a flow parameter in the engine is first measured that can be used directly or, with some additional processing, to predict the onset of stall of a stage of a compression system, such as, for example, a multistage fan shown in Figure 2.

[0030] Figure 2 shows an exemplary embodiment of a system 500 for detecting the onset of an aerodynamic instability, such as a stall or surge, in a compression stage in a gas turbine engine 10, In the exemplary embodiment shown in Figure 2, a fan section 12 shown, comprising a three stage first rotor, 12a, 12b and 12c. The embodiments of the present invention can also be used in a single stage fan, or in other compression system in a gas turbine engine, such as a high pressure compressor 18 or a low pressure compressor or a booster, In the exemplary embodiments shown herein, a sensor 502 is used to measure a local flow property near the tip region 52 of the compression system rotor blade tips 46 during engine operation. Although a single sensor 502 can be used for the flow parameter measurements, use of at least two sensors 502 is preferred, because some sensors may become inoperable during extended periods of engine operations, In an exemplary embodiment shown in Figure 2, multiple sensors 502 are used around the tips of all three fan rotor stages 12a, 12b, and 12c.

[0031] In the exemplar} _' embodiment shown in Figure 5, the sensor 502 is located on a casing 50 that is spaced radially outwardly and apart from the fan blade tips 46. Alternatively, the sensor 502 may be located on a shroud segment 51 that is located radially outwards from the blade tips. The casing 50, or a plurality of shrouds 51, surrounds the tips of a row of blades 47. The sensors 502 are arranged cireumferentially on the casing 50 or the shrouds 51 , as shown in Figure 6. In an exemplary embodiment of using multiple sensors on a rotor stage, the sensors 502 are arranged in substantially diametrically opposite locations in the casing or shroud.

[0032] During engine operation, there is an effective clearance 48 between the rotor blade tip and the casing 50 or the shroud 51 (see Figure 5). The sensor 502 is capable of generating an input signal 504 in real time corresponding to a flow parameter, such as, for example, the dynamic pressure, temperature, velocity and/or entropy in the blade tip region 52 near the blade tip 46. A suitable high response transducer, having a response capability higher than the blade passing frequency is used. Typically these transducers have a response capability higher than 1000 Hz. in the embodiment shown herein the sensors 502 used were dynamic pressure sensors 202 made by Kulite Semiconductor Products. In an alternative embodiment, the sensor 502 is any commercially available temperature sensor 204 having suitable dynamic capabilities. In another alternative embodiment, the sensor 502 is any commercially available velocity sensor 206 having suitable dynamic capabilities. In another alternative embodiment, the sensor 502 is any commercially available entropy sensor 208 having suitable dynamic capabilities. In another embodiment (see FIG, 12 for example) the sensor 502 is a plasmas sensor 60. It is preferable to use a high frequency sampling of the flow properties measurements, such as for example, between ten and twenty times the blade passing frequency.

[0033] The flow parameter measurement from the sensor 502 generates a signal that is used as an input signal 504 by a correlation processor 510. The coiTelation processor 510 may also optionally receive as input a signal 506 corresponding to the rotational speed of the compression system rotor, such as the fan rotor 12a, 12b, 12c, as shown in Figures 1 , 2 and 5. The signal 506 indicative of the rotational speed of the rotor (referred to herein as rotational speed signal 506 or as rotor speed signal 506) may be generated by any known methods, such as using rotational speed sensors, blade proximity sensors, or other known devices and methods. The rotor speed signal, when used, can provide one method of determining the blade passing period and/or frequency. However, such a measurement of rotor speed is not necessarily a requirement. Blade passing period/ frequency can also be determined in real-time, using known methods, from the signal from an unsteady pressure, iernperature, velocity or entropy sensor near the blade tip. In the exemplary embodiments shown herein, the compression system rotor speed signal 506 is supplied by a conventional engine control system 74, that is used in gas turbine engines. Alternatively, the compression system rotor speed signal 506 may be supplied by a digital electronic control system or a Full Authority Digital Electronic Control (FAD EC) system used an aircraft engine.

[0034] The correlation processor 510 receives the input signal 504 from the sensor 502 and the rotor speed signal 506 from the control system 74 and generates a stability coiTelation signal 5 12 in real time using conventional numerical methods. Auto correlation methods available in the published literature may be used. In the exemplary embodiments shown herein, the correlation processor 510 algorithm uses the existing speed signal from the engine control for cycle synchronization. The correlation measure is computed for individual sensors over rotor blade tips. The autocorrelation system in the exemplar}' embodiments described herein sampled a signal from a sensor 502 at a frequency of 200 KHz. A relatively high value of sampling frequency in the range of about 200-400 Kl 1/ ensures that the data is sampled at a rate at least ten times the fan blade 40 passage frequency. A window of seventy two samples was used to calculate the auto-correlation showing a value of near unity along the operating line 116 and dropping towards zero when the operation approached the stall/surge line 112 (see Figure 3). For a particular compression system rotor stage, such as, for example, the fan stage 12a, 12b, 12c, when the stability margin approaches zero, the particular compression stage rotor is on the verge of stall and the correlation measure is at a minimum. In systems designed to avoid a stall or surge in a compression system, when the correlation measure drops below a selected and pre-set threshold level, a stability management system receives the stability correlation signal 512 and sends an electrical signal to the engine control system, such as for example a FADEC system, which in turn can take corrective action using the available control devices to move the engine away from surge. The methods used by the correlation processor 510 for gauging the aerodynamic stability level in the exemplary embodiment shown herein is described in the paper, "Development and Demonstration of a Stability Management System for Gas Turbine Engines ^' ", Proceedings of GT2006 ASME Turbo Expo 2006, GT2006-90324.

[0035] Figure 5 shows schematically an exemplary embodiment of the present invention using a sensor 502 located in a casing 50 near the blade tip mid- chord of a blade 40. The sensor is located in the casing 50 such that it can measure a flow property of the air in the clearance 48 between a rotor blade tip 46 and the inner surface 53 of the casing 50. In one exemplary embodiment, the sensor 502 is located in an annular groove 54 in the casing 50. in other exemplary embodiments, it is possible to have multiple annular grooves 54 in the casing 50, such as for example, to provide for tip flow modifications for stability. If multiple grooves are present, the sensor 502 is located within some of these grooves, using the same principles and examples disclosed herein. Although the sensor is shown in Figure 5 as located in a casing 50, in other embodiments, the sensor 502 may be located in a shroud 51 that is located radially outwards and apart from the blade tip 46. Also, the sensor 502 may be located in a casing 50 (or shroud 51) near the leading edge 41 tip or the trailing edge 42 tip of the blade 40.

[0036] Figure 6 shows schematically an exemplary embodiment of the present invention using a plurality of sensors 502 in a compression system, such as a fan stage, shown in Figure 2. The plurality of sensors 502 are arranged in the casing 50 for shroud 51 ) in a circumferential direction, such that pairs of sensors 502 are located substantially diametrically opposite. The correlations processor 510 receives input signals 504 from these pairs of sensors and processes signals from the pairs together, The differences in the measured data from the diametrically opposite sensors in a pair can be particularly useful in developing stability correlation signal 512 to detect the on set of a fan stall due to engine inlet flow distortions, A single sensor has been demonstrated to be sufficient in some applications.

[0037] Figure 7 shows the axial location of the sensors 502, such as the pressure sensor 202, temperature sensor204, velocity sensor 206, entropy sensor 208 or a plasma sensor 60 with respect to the compression system rotor blade leading edge 41 and trailing edge 42. In a particular application, it is possible to have any one or more types of these flow property sensors, it is not necessary to have all these sensors in particular application and a suitable combination to obtain optimum results may be used. In FIG. 7, the rotor blade tip chord 49 is shown labeled "C". The tip chord C of the airfoil 34 is the axial length between the leading 41 and trailing edge 42 at the tip of the blade. In the present invention, the sensor 502 (such as the pressure sensor 202, temperature sensor 204, velocity sensor 206, entropy sensor 208 and plasma sensor 60) is located radially outwardly and apart from tips 46 of a circumferential row of blades 47 at a location on a static component 50 (such as a casing or a shroud) that is between a first location 57 and a second location 58. See FIG. 7. The first location 57 is at a first distance 157 (labeled as "A") of about 25% blade tip-chord length 49 ("C") axially forward from the leading edge 41 of a blade 47. The second location 58 is at a second distance 158 (labeled as "B") of about 25% blade tip-chord length 49 ("C") axially aft from the trailing edge 42 of a blade 47. Thus the sensor 502 may be located at a suitable axial location in the region 159 (labeled "D" in FIG. 7). In a preferred embodiment, the sensor is located at an axial location corresponding to the mid-chord of the rotor blade tip.

[0038] FIGS. 8-11 show time history of the flow properties, pressure, temperature, velocity and entropy from an unsteady computational fluid dynamic (CFD) simulation in the rotor's relative frame of reference as the compression system approaches a stall condition. Testing experience has demonstrated that unsteady pressure measurements can be successfully used for autocorrelation calculations to predict an impending stall condition. As discussed previously herein, a lack of correlation between successive measurements in a rotor is observed when a stall is approaching. As evident in the unsteady CFD simulation shown in FIG. 8, the three local dips (items 302) below the zero non-dimensional pressure are typical of features that result in low autocorrelation when observed in the absolute frame of reference. Known autocorrelation algorithms can be used on the pressure measurements from the pressure sensor 202.

[0039] FIG. 9 shows the time history of the temperature at the location of a temperature sensor 204 in an alternative embodiment of the present invention from an unsteady computational fluid dynamic (CFD) simulation in the rotor's relative frame of reference as the compression system approaches a stall condition. In this alternative embodiment, unsteady temperature measurements using the temperature sensor 204 can be used for autocorrelation calculations to predict an impending stall condition. As discussed previously herein, a lack of correlation between successive measurements in a rotor is observed when a stall is approaching. As evident in the unsteady CFD simulation shown in FIG. 9, the three local peaks (items 304) above the zero non-dimensional temperature are typical of features that result in low autocorrelation when observed in the absolute frame of reference. Known autocorrelation algorithms can be used on the temperature measurements from the temperature sensor 204,

[0040] FIG. 10 shows the time history of the air velocity at the location of a velocity sensor 206 in an alternative embodiment of the present invention from an unsteady computational fluid dynamic (CFD) simulation in the rotor's relative frame of reference as the compression system approaches a stall condition. In this alternative embodiment, unsteady velocity measurements using the velocity sensor 206 can be used for autocorrelation calculations to predict an impending stall condition. As discussed previously herein, a lack of correlation between successive measurements in a rotor is observed when a stall is approaching. As evident in the unsteady CFD simulation shown in FIG. 10, the three local dips (items 306) below the zero non- dimensional velocity are typical of features that result in low autocorrelation when observed in the absolute frame of reference. Known autocorrelation algorithms can be used on the velocity measurements from the velocity sensor 206, [0041] FIG. 11 shows the time history of the entropy at the location of an entropy sensor 208 in an alternative embodiment of the present invention from an unsteady computational fluid dynamic (CFD) simulation in the rotor's relative frame of reference as the compression system approaches a stall condition. In this alternative embodiment, unsteady entropy measurements using the entropy sensor 208 can be used for autocorrelation calculations to predict an impending stall condition, As discussed previously herein, a lack of correlation between successive measurements in a rotor is observed when a stall is approaching. As evident in the unsteady CFD simulation shown in FIG. I I, the three local peaks (items 308) above the zero non- dimensional entropy are typical of features that result in low autocorrelation when observed in the absolute frame of reference. Known autocorrelation algorithms can be used on the entropy measurements from the entropy sensor 308.

[0042] In a preferred embodiment of the present invention of a system 500 for detecting the onset of a stall in a rotor, the sensor 502 is a known plasma sensor 60. Plasma sensors are typically composed of a pair of insulated electrodes having a small gap between them, on the order of about 0.005 inch. A plasma sensor, utilizes a high potential AC voltage between the electrodes, applied with a frequency in the range of 1-2 MHz. In one aspact of the present invention, one or more plasma sensors may be used as part of an active stall control system having high temperature capability using known auto-correlation methods. In one embodiment of the present invention, shown in FIG. 12, one or more plasma sensors are installed on the fan or booster or compressor casing 50 wall at the locations where the rotor blade 40 tip vortex may be present. Although FIG. 12 shows a plasma sensor located axially forward from the rotor blade 40 leading edge 41, other suitable locations may also be selected. The plasma sensor 60 signal processed through auto-correlation measures the level of unsteadiness generated by the tip leakage flow, thereby allowing the engine control system to calculate margin remaining before a rotor stall occurs. As impending rotor stall is detected by the system, action can be taken to avoid a stall. Examples of possible corrective actions known in the art include reverting to a variable stator vane schedule with more stall margin, altering bleed flow level to relieve compression stage loading, or reducing fuel flow, thereby slowing engine acceleration and/or unloading the compressor, Known signal processing routines and various types of known control logic can be used for processing the signals from the plasma sensor 60 and use the data to control engine operation.

[0043] The present invention of a system 500 for detecting the onset of a stall in a rotor, provides several unique advantages. Plasma sensors 60 have high temperature capabilities (about 3000°F), since they are chiefly constructed of durable metal electrodes and a dielectric. This provides for much better survivability in harsh compressor environments than more fragile conventional sensors, such as Kulite-style sensors. No cooling air is required to achieve acceptable sensor life, which reduces system complexity and improves compressor efficiency. Plasma sensors have very high response frequency capability (about 2 MHz and higher), allowing them to take measurements at the speeds required to resolve the flow features between passing blades found in high speed turbomachineiy applications. Plasma sensors are relatively small and simple to locate precisely, without creating large amounts of blockage or loss. This allows sensor locations to be optimized to maximize sensing of the vortex unsteadiness. Computational analysis indicates that suitable response can be obtained by placing the sensors as much as 25°/» blade chord upstream or downstream of the blade leading and trailing edges 41 , 42. In a preferred embodiment, the plasma sensor 60 is placed at a location corresponding to the mid-chord of the blade.

[0044] Additional turbomachinery diagnostic applications may also use plasma sensors, particularly where high temperature operation is needed, such as for example, in turbines and combustors. Active control of tip-clearances in turbines for optimum performance requires either measurement or a model of the clearances, including transients. Since the flow field in the tip region of a turbine blade is affected by the operating clearance, the output of a plasma sensor could be correlated with the operating clearance. A calibrated relationship between output of a high temperature capable plasma sensor and operating clearance can be used for controlling turbine tip clearances.

[0045] The gas turbine engine plasma sensor system 100 (see FIG. 12) includes an annular casing 50, or annular shroud segments 51, surrounding rotatable blade tips 46. A plasma generator, such as for example, a plasma sensor 60, is located on the casing 50, or the shroud segments 51, in a groove 54 or groove segments 56 spaced radially outward from the blade tips 46. The exemplary embodiment shown in Figure 12 comprises a leading edge plasma sensor 60 located in the casing 50 near the tip 46 of the lead edge 41.

[0046] Figure 12 shows an exemplar}' embodiment of a plasma sensor system 100 for evaluating the stall margin of a compression system 18. The term "compression system" as used herein includes devices used for increasing the pressure of a fluid flowing through it, and includes the high pressure compressor 18, the booster (similar to the compressor, and not shown) and the fan 12 used in gas turbine engines shown in Figure 1. The exemplary embodiment shown in Figure 12 shows a plasma sensor 60 mounted to the compressor casing 50 and includes a first electrode 62 and a second electrode 64 separated by a gap 61. The plasma sensor 60 is disposed in a radially inwardly facing surface 53 of the casing 50, In some applications, the plasma sensor 60 may be disposed in a groove 54 in a static component such as the casing or a shroud. In some gas turbine engine designs, some of the stages of the compressor 18 may have annular shroud segments 51 surrounding the blade tips. In some exemplary embodiments, multiple plasma sensors 60 may be located in the circumferential direction. In some applications, a plurality of shroud segments include an annular groove segment 56. The annular array of groove segments 56 may have a plurality of plasma sensors 60.

[0047] An AC power supply 70 is connected to the electrodes to supply a high voltage AC potential to the electrodes 62, 64, When the AC amplitude is large enough, the air ionizes in a region of largest electric potential forming a plasma 68. The plasma 68 generally begins near an edge 65 of the first electrode 62 which is exposed to the air and spreads out over an area 104 projected by the second electrode 64. The plasma 68 (ionized air) is formed in the presence of an electric field gradient. The air near the electrodes is weakly ionized, and usually there is little or no heating of the air. [0048] During engine operation, the plasma generator system 100 turns on the plasma sensor 60 to form the plasma 68 near the blade tips 46 region. An electronic controller 72 which is linked to an engine control system 74, such as for example a Full Authority Digital Electronic Control (FADEC), which controls inputs to affect the fan speeds, compressor and turbine speeds and fuel system of the engine, may be used to control the plasma sensor 60 and the auto-correlation of plasma sensor signals, or otherwise modulating it as necessary to sense and/or increase the stall margin or enhancing the efficiency of the compression system. The electronic controller 72 may also be used to control the operation of the AC power supply 70 that is connected to the electrodes to supply a high voltage AC potential to the electrodes.

[0049] In operation, when turned on, the plasma generator 100 system produces a stream of ions forming the plasma 68 enabling the sensing of the pressure distribution near the blade tip on the radially inwardly facing surface 53 of the annular casing 50. A plasma actuator, when present, provides a positive axial momentum to the fluid in the blade tip region 52 where a vortex 200 tends to form in conventional compressors as described previously and as shown in Figures 4a, 4b and 4c. The positive axial momentum applied by the plasma actuator (when present) forces the air to pass through the passage between adjacent blades, in the desired direction of positive flow, avoiding the type of flow blockage shown in Figure 4c for conventional engines. This increases the stall margin of the compressor stage and hence the compression system. Plasma sensors 60, such as for example, shown in Figure 12, may be located around the tip of some selected compressor stages where stall is likely to occur. Plasma sensors and plasma actuators may be located around tips of a plurality of the compression stages and selectively activated during engine operation using the engine control system 74 or the electronic controller 72.

[0050] Plasma sensors 60 may be placed axially at a variety of axial locations with respect to the blade leading edge 41 tip. They may be placed axially upstream from the blade leading edge 41 (see item labeled "A" in FIG. 7). They may also be placed axially downstream from the leading edge 41 , or downstream from the trailing edge 42 (see item labeled "B" in FIG. 7). Plasma sensors 60 are effective when placed in axial locations from about 25% blade tip-chord (see item labeled "C" in FIG, 7) upstream from the leading edge 41 to about 25% blade tip chord downstream from the trailing edge 42. Plasma actuators, when present, are most effective when they can act directly upon the low momentum fluid associated with the tip vortex 200 such as, for example, shown in Figure 4a. The plasma sensor 60 may be placed such that plasma 68 stream influence starts at about 10% blade tip chord, where the vortex is seen to start its growth, as shown in Figure 4a, It is more preferable to locate the plasma sensors 60 at locations from about 10% chord aft of the leading edge 41 to about 50% chord.

[0051] This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to make and use the invention. 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 have 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 languages of the claims.