U.S. patent application number 11/966242 was filed with the patent office on 2010-11-11 for fan stall detection system.
Invention is credited to Clark Leonard Applegate, Seyed Gholamali Saddoughi, Aspi Rustom Wadia.
Application Number | 20100284785 11/966242 |
Document ID | / |
Family ID | 40383778 |
Filed Date | 2010-11-11 |
United States Patent
Application |
20100284785 |
Kind Code |
A1 |
Wadia; Aspi Rustom ; et
al. |
November 11, 2010 |
Fan Stall Detection System
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.
Inventors: |
Wadia; Aspi Rustom;
(Loveland, OH) ; Saddoughi; Seyed Gholamali;
(Clifton Park, NY) ; Applegate; Clark Leonard;
(West Chester, OH) |
Correspondence
Address: |
GENERAL ELECTRIC COMPANY
GE AVIATION, ONE NEUMANN WAY MD F16
CINCINNATI
OH
45215
US
|
Family ID: |
40383778 |
Appl. No.: |
11/966242 |
Filed: |
December 28, 2007 |
Current U.S.
Class: |
415/118 |
Current CPC
Class: |
F05D 2270/101 20130101;
F01D 21/003 20130101; F04D 27/001 20130101 |
Class at
Publication: |
415/118 |
International
Class: |
F04D 29/00 20060101
F04D029/00 |
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.
Description
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, such as a fan.
[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 aerodynamically 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. 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.
[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.
[0005] 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
[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 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.
[0007] 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
[0008] 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:
[0009] FIG. 1 is a schematic cross-sectional view of a gas turbine
engine with an exemplary embodiment of the present invention.
[0010] FIG. 2 is an enlarged cross-sectional view of a portion of
the fan section of the gas turbine engine shown in FIG. 1.
[0011] FIG. 3 is an exemplary operating map of a compression system
in the gas turbine engine shown in FIG. 1.
[0012] FIG. 4a shows the formation of a region with blade tip
vortex in a fan stage.
[0013] FIG. 4b shows the spread of the blade tip vortex shown in
FIG. 4a.
[0014] FIG. 4c shows the vortex flow at blade tip region during a
stall.
[0015] FIG. 5 is a schematic cross-sectional view of the tip region
of a fan with an exemplary embodiment of a stall detection
system.
[0016] FIG. 6 is a schematic sketch of an exemplary arrangement of
multiple sensors for a stall detection system.
DETAILED DESCRIPTION OF THE INVENTION
[0017] 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.
[0018] 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.
[0019] 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.
[0020] 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.
[0021] 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 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.
[0022] 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.
[0023] 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 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 12a, 12b, and 12c.
[0024] 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.
[0025] 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.
[0026] 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 12a, 12b, 12c, 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.
[0027] 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 12a, 12b, 12c 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.
[0028] 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.
[0029] FIG. 6 shows schematically an exemplary embodiment of the
present invention using a plurality of sensors 502 in a fan stage,
such as item 40a 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.
[0030] 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.
* * * * *