US20100284785A1

US20100284785A1 - Fan Stall Detection System

Info

Publication number: US20100284785A1
Application number: US11/966,242
Authority: US
Inventors: Aspi Rustom Wadia; Seyed Gholamali Saddoughi; Clark Leonard Applegate
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2007-12-28
Filing date: 2007-12-28
Publication date: 2010-11-11
Also published as: CA2710009A1; GB2467715B; DE112008003400T8; GB2467715A; WO2009086358A1; DE112008003400T5; GB201010130D0; JP2011508155A

Abstract

A system for detecting onset of a stall in a rotor is disclosed, the system comprising a sensor located on a static component spaced radially outwardly and apart from tips of a row of blades arranged circumferentially on the rotor wherein the sensor is capable of generating an input signal corresponding to a flow parameter at a location near the tip of a blade, a control system capable of generating a rotor speed signal, and a correlation processor capable of receiving the input signal and the rotor speed signal wherein the correlation processor generates a stability correlation signal.

Description

BACKGROUND OF THE INVENTION

This invention relates generally to gas turbine engines, and, more specifically, to a system for detection of a stall in a compression system therein, such as a fan.
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 aerodynamically coupled to the compressor rotor although these normally operate at different mechanical speeds.
Operability in a wide range of operating conditions is a fundamental requirement in the design of compression systems, such as fans, boosters and compressors. Modern 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.
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.
Accordingly, it would be desirable to have the ability to measure and control dynamic processes such as flow instabilities in a fan. It would be desirable to have a system that can measure an engine parameter related to the onset of flow instabilities, such as the dynamic pressure near the blade tips, and process the measured data to predict the onset of stall in a stage of a compression system, such as a multistage fan. 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

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 sensor located on a static component spaced radially outwardly and apart from tips of a row of blades arranged circumferentially on the rotor wherein the sensor is capable of generating an input signal corresponding to a flow parameter at a location near the tip of a blade, a control system capable of generating a rotor speed signal, and a correlation processor capable of receiving the input signal and the rotor speed signal wherein the correlation processor generates a stability correlation signal.
In another embodiment, a system for detecting onset of a stall in a multi-stage fan rotor comprises a pressure sensor located on a casing surrounding tips of a row of fan blades wherein the pressure sensor is capable of generating an input signal corresponding to the dynamic pressure at a location near the fan blade tip.

BRIEF DESCRIPTION OF THE DRAWINGS

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:

FIG. 1 is a schematic cross-sectional view of a gas turbine engine with an exemplary embodiment of the present invention.

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

FIG. 3 is an exemplary operating map of a compression system in the gas turbine engine shown in FIG. 1.

FIG. 4 a shows the formation of a region with blade tip vortex in a fan stage.

FIG. 4 b shows the spread of the blade tip vortex shown in FIG. 4 a.

FIG. 4 c shows the vortex flow at blade tip region during a stall.

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

FIG. 6 is a schematic sketch of an exemplary arrangement of multiple sensors for a stall detection system.

DETAILED DESCRIPTION OF THE INVENTION

Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views, FIG. 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 FIG. 1) mounted between the fan section and the HPC. 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 exemplary embodiments of the present invention shown in FIGS. 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 12 a, 12 b, and 12 c respectively.
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 40 a, 40 b, 40 c extending radially outwardly from corresponding rotor hubs 39 a, 39 b, 39 c in the form of separate disks, or integral blisks, or annular drums in any conventional manner.
Cooperating with a fan rotor stage 12 a, 12 b, 12 c is a corresponding stator stage comprising a plurality of circumferentially spaced apart stator vanes 31 a, 31 b, 31 c. The arrangement of stator vanes and rotor blades is shown in FIG. 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 leading 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 the air decelerates and diffuses through the stator and rotor airfoils.
Operating map of an exemplary compression system, such as the fan section 12 in the exemplary gas turbine engine 10 is shown in FIG. 3, with inlet corrected flow rate along the X-axis and the pressure ratio on the Y-axis. Operating lines 114, 116 and the stall line 112 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 FIG. 3. The performance map has a region of peak efficiency, depicted in FIG. 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 112 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.
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 the fan rotors 12 a, 12 b, 12 c shown in FIG. 2. This tip flow breakdown is associated with tip leakage vortex schematically shown in FIGS. 4 a, 4 b and 4 c 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 axially aft and tangentially from the blade suction surface 44 to the adjacent blade pressure surface 43 as shown in FIG. 4 b. 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 FIG. 4 c 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 vortex 200 spans the leading edges 41 of adjacent blades 40, as shown in FIG. 4 c. 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 FIG. 4 c.
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 using a continuous measurement method or using samples of sufficient number of discrete measurements. In order to mitigate fan 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 multistage fan shown in FIG. 2.
FIG. 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 FIG. 2, a fan section 12 shown, comprising a three stage first rotor, 12 a, 12 b and 12 c. 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 pressure sensor 502 is used to measure the local dynamic pressure near the tip region 52 of the fan 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 FIG. 2, multiple pressure sensors 502 are used around the tips of all three fan rotor stages 12 a, 12 b, and 12 c.
In the exemplary embodiment shown in FIG. 5, the pressure sensor 502 is located on a casing 50 that is spaced radially outwardly and apart from the fan blade tips 46. Alternatively, the pressure 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 pressure sensors 502 are arranged circumferentially on the casing 50 or the shrouds 51, as shown in FIG. 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.
During engine operation, there is an effective clearance 48 between the fan blade tip and the casing 50 or the shroud 51 (see FIG. 5). The sensor 502 is capable of generating an input signal 504 in real time corresponding to a flow parameter, such as the dynamic pressure 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 exemplary embodiments shown herein the sensors 502 used were made by Kulite Semiconductor Products. It is preferable to use a high frequency sampling of the dynamic pressure measurement, such as for example, approximately ten times the blade passing frequency.
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 correlation processor 510 also receives as input a fan rotor speed signal 506 corresponding to the rotational speed of the fan rotor 12 a, 12 b, 12 c, as shown in FIGS. 1, 2 and 5. In the exemplary embodiments shown herein, the fan rotor speed signal 506 is supplied by a conventional engine control system 74, that is used in gas turbine engines. Alternatively, the fan rotor speed signal 506 may be supplied by a digital electronic control system or a Full Authority Digital Electronic Control (FADEC) system used an aircraft engine.
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 correlation signal 512 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 pressure transducers over rotor blade tips. The auto-correlation system in the exemplary embodiments described herein sampled a signal from a pressure sensor 502 at a frequency of 200 KHz. This relatively high value of sampling frequency 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 FIG. 3). For a particular fan stage 12 a, 12 b, 12 c when the stability margin approaches zero, the particular fan stage 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.
FIG. 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 the dynamic pressure of the air in the clearance 48 between a fan 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 pressure sensor 502 is located within some of these grooves, using the same principles and examples disclosed herein. Although the sensor is shown in FIG. 5 as located in a casing 50, in other embodiments, the pressure sensor 502 may be located in a shroud 51 that is located radially outwards and apart from the blade tip 46. Also, the pressure 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.
FIG. 6 shows schematically an exemplary embodiment of the present invention using a plurality of sensors 502 in a fan stage, such as item 40 a in FIG. 2. The plurality of sensors 502 are arranged in the casing 50 (or 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.
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.

Claims

1. A system for detecting onset of a stall in a rotor, the system comprising:

a sensor located on a static component spaced radially outwardly and apart from tips of a row of blades arranged circumferentially on the rotor wherein the sensor is capable of generating an input signal corresponding to a flow parameter at a location near the tip of a blade;

a control system capable of generating a rotor speed signal; and

a correlation processor capable of receiving the input signal and the rotor speed signal wherein the correlation processor generates a stability correlation signal.

2. A system according to claim 1 further comprising:

a plurality of sensors arranged on the static component spaced radially outwardly and apart from tips of the row of blades.

3. A system according to claim 2 wherein the sensor is a pressure sensor.

4. A system according to claim 2 wherein the sensor is a pressure sensor capable of generating a pressure signal corresponding to the dynamic pressure at a location near the blade tip.

5. A system according to claim 1 further comprising:

a plurality of sensors arranged circumferentially on the static component around an axis of rotation of the rotor and spaced radially outwardly and apart from tips of the row of blades.

6. A system according to claim 2 wherein the static component is a casing.

7. A system according to claim 2 wherein the static component is a shroud.

8. A system according to claim 1 wherein the rotor comprises a plurality of fan rotors.

9. A system according to claim 1 wherein the sensor is located at a location on the static structure corresponding to the mid-chord of a blade.

10. A system according to claim 1 wherein the sensor is located at a location on the static structure corresponding to the lead edge of a blade.

11. A system for detecting onset of a stall in a fan rotor comprising:

a pressure sensor located on a static component surrounding tips of a row of fan blades wherein the pressure sensor is capable of generating an input signal corresponding to the dynamic pressure at a location near the blade tip;

a control system capable of generating a fan rotor speed signal; and

a correlation processor capable of receiving the input signal and the fan speed signal wherein the correlation processor generates a stability correlation signal.

12. A system according to claim 11 further comprising a plurality of fan rotors wherein a plurality pressure sensors are located on the static component surrounding tips of a row of fan blades of at least two fan rotors.

13. A system according to claim 11 further comprising a plurality of sensors arranged circumferentially on the static component around an axis of rotation of the rotor and spaced radially outwardly and apart from tips of the row of fan blades.

14. A system according to claim 11 wherein a fan blade tip operates at a supersonic speed during the generation of the pressure signal.

15. A system according to claim 11 wherein the correlation processor receives the input signal from a plurality of pressure sensors and the rotor speed signal to generate a correlation signal.

16. A system according to claim 11 wherein the correlation processor generates a correlation signal based on the input signal from a plurality of pressure sensors and the rotor speed signal.

17. A system according to claim 11 wherein the correlation processor generates a correlation signal based on the input signal from pressure sensors located on the static component surrounding tips of a row of fan blades of at least two fan rotors.

18. A system according to claim 11 wherein the static component is a casing.

19. A system according to claim 11 wherein the static component is a shroud.

20. A system according to claim 11 wherein the sensor is located at a location on the static structure corresponding to the mid-chord of a fan blade.

21. A system according to claim 11 wherein the sensor is located at a location on the static structure corresponding to the leading edge of a fan blade.