U.S. patent application number 12/766432 was filed with the patent office on 2010-11-18 for plasma sensor stall control system and turbomachinery diagnostics.
Invention is credited to Andrew Breeze-Stringfellow, David S. Clark, Curtis W. Moeckel, Aspi Wadia.
Application Number | 20100290906 12/766432 |
Document ID | / |
Family ID | 43068633 |
Filed Date | 2010-11-18 |
United States Patent
Application |
20100290906 |
Kind Code |
A1 |
Moeckel; Curtis W. ; et
al. |
November 18, 2010 |
PLASMA SENSOR STALL CONTROL SYSTEM AND TURBOMACHINERY
DIAGNOSTICS
Abstract
A system for detecting onset of a stall in a rotor is disclosed,
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 plasma sensor is capable of generating an
input signal corresponding to a flow parameter at a location near
the tip of a blade and indicative of the onset of a stall and a
correlation processor that generates a stability correlation
signal.
Inventors: |
Moeckel; Curtis W.;
(Cincinnati, OH) ; Wadia; Aspi; (Cincinnati,
OH) ; Clark; David S.; (Cincinnati, OH) ;
Breeze-Stringfellow; Andrew; (Cincinnati, OH) |
Correspondence
Address: |
GENERAL ELECTRIC COMPANY
GE AVIATION, ONE NEUMANN WAY MD F16
CINCINNATI
OH
45215
US
|
Family ID: |
43068633 |
Appl. No.: |
12/766432 |
Filed: |
April 23, 2010 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
11966242 |
Dec 28, 2007 |
|
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12766432 |
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Current U.S.
Class: |
416/31 |
Current CPC
Class: |
F01D 21/003 20130101;
F05D 2270/172 20130101; F05D 2270/101 20130101; F01D 17/06
20130101 |
Class at
Publication: |
416/31 |
International
Class: |
F01D 7/00 20060101
F01D007/00 |
Claims
1. 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 plasma sensor is capable of generating an
input signal corresponding to a flow parameter at a location near
the tip of a blade and indicative of the onset of a stall; 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
plasma sensors arranged on the static component spaced radially
outwardly and apart from tips of the row of blades.
3. A system according to claim 1 further comprising: a plurality of
plasma 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.
4. A system according to claim 1 wherein the static component is a
casing.
5. A system according to claim 4 wherein the plasma sensor is
located at a location on the static component corresponding to the
mid-chord of a blade.
6. A system according to claim 1 wherein the static component is a
shroud.
7. A system according to claim 1 wherein the rotor comprises a
plurality of fan rotors.
8. A system according to claim 1 wherein the rotor is a compressor
rotor.
9. A system according to claim 1 wherein the rotor is a booster
rotor.
10. A system for detecting onset of a stall in a compressor rotor
comprising: a plasma sensor spaced radially outwardly and apart
from tips of a circumferential row of compressor 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 plasma sensor is capable
of generating an input signal indicative of the onset of a stall in
the compressor; and a correlation processor capable of receiving
the input signal and a rotor speed signal wherein the correlation
processor generates a stability correlation signal.
11. A system according to claim 10 further comprising a plurality
of compressor rotors wherein a plurality plasma sensors are located
on the static component surrounding tips of compressor blades of at
least two compressor rotors.
12. A system according to claim 10 further comprising a plurality
of plasma sensors arranged circumferentially on the static
component around an axis of rotation of the compressor rotor and
spaced radially outwardly and apart from tips of the row of
compressor blades.
13. A system for controlling an operating tip clearance between a
rotor and a static component, 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
wherein the plasma sensor is capable of generating an input signal
corresponding to a flow parameter at a location near the tip of a
blade; a correlation processor capable of generating a correlation
signal indicative of the operating tip clearance based on the input
signal; and a control system capable of controlling the operating
tip clearance based on the correlation signal.
14. A system according to claim 13 wherein the plasma sensor is
located on the static component 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.
15. A system according to claim 13 wherein the plasma sensor is
located at a location on the static component corresponding to the
mid-chord of a blade.
16. A system according to claim 13 further comprising: a plurality
of plasma 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.
17. A system according to claim 13 wherein the static component is
a casing.
18. A system according to claim 13 wherein the static component is
a shroud.
19. A system according to claim 13 wherein the rotor is a turbine
rotor.
20. A system according to claim 13 wherein the rotor is a
compressor rotor.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application is a Continuation-in-Part (CIP) patent
application of U.S. patent application Ser. No. 11/966,242, filed
Dec. 28, 2007. The contents of that prior patent application are
incorporated herein by reference in their entirety.
BACKGROUND OF THE INVENTION
[0002] This invention relates generally to gas turbine engines,
and, more specifically, to a system for detection of a stall in a
compression system therein.
[0003] 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.
[0004] 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.
[0005] 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.
[0006] 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 flow
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 clearances 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
[0007] 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 flow parameter at a
location near the tip of a blade and indicative of the onset of a
stall and a correlation processor that generates a stability
correlation signal.
[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] FIG. 1 is a schematic cross-sectional view of a gas turbine
engine with an exemplary embodiment of the present invention.
[0011] FIG. 2 is an enlarged cross-sectional view of a portion of
the fan section of the gas turbine engine shown in FIG. 1.
[0012] FIG. 3 is an exemplary operating map of a compression system
in the gas turbine engine shown in FIG. 1.
[0013] FIG. 4a shows the formation of a region with blade tip
vortex in a fan stage.
[0014] FIG. 4b shows the spread of the blade tip vortex shown in
FIG. 4a.
[0015] FIG. 4c shows the vortex flow at blade tip region during a
stall.
[0016] FIG. 5 is a schematic cross-sectional view of the tip region
of a fan with an exemplary embodiment of a stall detection
system.
[0017] FIG. 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,
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 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 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.
[0027] 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.
[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 FIG. 2.
This tip flow breakdown is associated with tip leakage vortex
schematically shown in FIGS. 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 axially aft and tangentially from the blade
suction surface 44 to the adjacent blade pressure surface 43 as
shown in FIG. 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 FIG. 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 vortex 200 spans the leading edges 41 of adjacent
blades 40, as shown in FIG. 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 FIG. 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 FIG. 2.
[0030] 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, 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 FIG. 2, multiple sensors 502 are used around the tips of
all three fan rotor stages 12a, 12b, and 12c.
[0031] In the exemplary embodiment shown in FIG. 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 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.
[0032] During engine operation, there is an effective clearance 48
between the rotor 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, 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 correlation 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 FIGS. 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, temperature,
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
(FADEC) 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 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
sensors over rotor blade tips. The auto-correlation system in the
exemplary 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 KHz 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
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] 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 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 FIG. 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] FIG. 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 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. A single sensor has
been demonstrated to be sufficient in some applications.
[0037] FIG. 7 shows the axial location of the sensors 502, such as
the pressure sensor 202, temperature sensor 204, 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. 11, 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 aspect 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.degree. 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 turbomachinery 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 FIG. 12 comprises a leading edge plasma sensor 60 located
in the casing 50 near the tip 46 of the lead edge 41.
[0046] FIG. 12 shows an exemplary 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 FIG. 1. The exemplary embodiment shown in FIG. 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 FIGS. 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 FIG. 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 FIG. 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
FIG. 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 FIG. 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.
* * * * *