U.S. patent application number 11/682015 was filed with the patent office on 2008-11-06 for flutter sensing and control system for a gas turbine engine.
Invention is credited to Robert J. Morris, William E. Rosenkrans.
Application Number | 20080273961 11/682015 |
Document ID | / |
Family ID | 39363931 |
Filed Date | 2008-11-06 |
United States Patent
Application |
20080273961 |
Kind Code |
A1 |
Rosenkrans; William E. ; et
al. |
November 6, 2008 |
FLUTTER SENSING AND CONTROL SYSTEM FOR A GAS TURBINE ENGINE
Abstract
A gas turbine engine system includes a nacelle, a fan casing
within the nacelle, a variable area fan nozzle, a sensor and a
controller. The sensor detects an airfoil flutter condition. The
controller communicates with the sensor and is operable to move the
variable area fan nozzle to influence a discharge airflow area in
response to the detection of the airfoil flutter condition.
Inventors: |
Rosenkrans; William E.;
(Columbia, CT) ; Morris; Robert J.; (Portland,
CT) |
Correspondence
Address: |
CARLSON, GASKEY & OLDS/PRATT & WHITNEY
400 WEST MAPLE ROAD, SUITE 350
BIRMINGHAM
MI
48009
US
|
Family ID: |
39363931 |
Appl. No.: |
11/682015 |
Filed: |
March 5, 2007 |
Current U.S.
Class: |
415/13 ; 415/118;
415/151 |
Current CPC
Class: |
F01D 17/08 20130101;
F02C 9/20 20130101; F01D 17/14 20130101; F05D 2270/10 20130101 |
Class at
Publication: |
415/13 ; 415/118;
415/151 |
International
Class: |
F02C 9/22 20060101
F02C009/22 |
Claims
1. A gas turbine engine system, comprising: a nacelle; a fan casing
within the nacelle; a variable area fan nozzle; a sensor which
detects an airfoil flutter condition; and a controller that
communicates with said sensor, wherein said controller is operable
to move said variable area fan nozzle to influence a discharge
airflow area associated with said variable area fan nozzle in
response to the detection of said airfoil flutter condition.
2. The system as recited in claim 1, wherein said sensor is mounted
to said fan casing adjacent to a blade tip area of a fan
section.
3. The system as recited in claim 1, wherein said airfoil flutter
condition comprises self-induced oscillations simultaneously
experienced by adjacent airfoils of the gas turbine engine
system.
4. The system as recited in claim 3, wherein said adjacent airfoils
are circumferentially disposed about a fan section of the gas
turbine engine system.
5. The system as recited in claim 1, wherein said controller
influences said discharge airflow area by moving said variable area
fan nozzle between a first position having a first discharge
airflow area and a second position having a second discharge
airflow area greater than said first discharge airflow area in
response to detection of said airfoil flutter condition.
6. The system as recited in claim 1, wherein said discharge airflow
area extends between said variable area fan nozzle and a core
engine casing.
7. A gas turbine engine, comprising: a nacelle; a fan casing within
the nacelle; a variable area fan nozzle moveable to influence a
discharge airflow area associated with said variable area fan
nozzle; a fan section within said fan casing; a compressor section
and a turbine section positioned downstream of said fan section; a
combustor section positioned between said compressor section and
said turbine section; a sensor which detects an airfoil flutter
condition; and a controller that communicates with said sensor,
wherein said controller is operable to move said variable area fan
nozzle in response to the detection of said airfoil flutter
condition.
8. The gas turbine engine as recited in claim 7, wherein said
variable area fan nozzle includes at least one flap assembly, said
at least one flap assembly moveable between a first position having
a first discharge airflow area and a second position having a
second discharge airflow area greater than said first discharge
airflow area.
9. The gas turbine engine as recited in claim 8, comprising an
actuator assembly in communication with said controller and
operable to move said at least one flap assembly between said first
position and said second position.
10. The gas turbine engine as recited in claim 7, wherein said
sensor is mounted to said fan casing adjacent to a blade tip area
of said fan section.
11. The gas turbine engine as recited in claim 10, wherein said fan
section includes a plurality of airfoils circumferentially disposed
about said fan section.
12. The gas turbine engine as recited in claim 7, wherein the gas
turbine engine is a geared turbofan engine.
13. The gas turbine engine system as recited in 7, wherein said
airfoil flutter condition comprises self-induced oscillations
simultaneously experienced by adjacent airfoils of said fan
section.
14. The gas turbine engine system as recited in 7, wherein said
discharge airflow area extends between said variable area fan
nozzle and a core engine casing.
15. A method for controlling a gas turbine engine, comprising the
steps of: (a) sensing a flutter condition; and (b) influencing a
discharge airflow area associated with a variable area fan nozzle
in response to sensing the flutter condition.
16. The method as recited in claim 15, wherein the flutter
condition comprises self-induced oscillations simultaneously
experienced by adjacent airfoils of the gas turbine engine.
17. The method as recited in claim 15, wherein said step (a)
comprises: mounting a sensor to a fan section of the gas turbine
engine for sensing the flutter condition.
18. The method as recited in claim 15, wherein the variable area
fan nozzle is moveable between a first position having a first
discharge airflow area and a second position having a second
discharge airflow area greater than the first discharge airflow
area, wherein said step (b) comprises: influencing the discharge
airflow area by moving the variable area fan nozzle from the first
position to the second position in response to sensing the flutter
condition.
19. The method as recited in claim 18, further comprising the step
of: (c) returning the variable area fan nozzle to the first
position in response to sensing that the flutter condition has
ceased.
Description
BACKGROUND OF THE INVENTION
[0001] This invention generally relates to a gas turbine engine,
and more particularly to a flutter sensing system for a gas turbine
engine.
[0002] Gas turbine engines typically include a compressor section,
a combustor section and a turbine section. Air is pressurized in
the compressor section and is mixed with fuel and burned in the
combustor section to add energy to expand the air and accelerate
the airflow into the turbine section. The hot combustion gases that
exit the combustor section flow downstream through the turbine
section, which extracts kinetic energy from the expanding gases and
converts the energy into shaft horsepower to drive the compressor
section.
[0003] In a turbofan gas turbine engine, for example, a fan section
is included upstream of the compressor section. Combustion gases
are discharged from the gas turbine engine through a core exhaust
nozzle and fan air is discharged through an annular fan exhaust
nozzle defined at least partially by a nacelle surrounding the core
engine. A majority of propulsion thrust is provided by the
pressurized fan air which is discharged through the fan exhaust
nozzle, while the remaining thrust is provided from combustion
gases discharged through the core exhaust nozzle.
[0004] A fan section, the compressor section and the turbine
section may include multiple airfoils disposed circumferentially
about an engine longitudinal centerline axis. At certain aircraft
operating conditions, these airfoils may be subjected to flutter,
or self-induced oscillations. The flutter conditions are caused by
the interaction between adjacent airfoils. During flutter,
aerodynamic forces couple with each airfoil's elastic and inertial
forces, which may increase the kinetic energy of each airfoil and
produce negative damping. The negative damping is enhanced where
adjacent airfoils vibrate in unison. Disadvantageously, the airfoil
oscillations caused by flutter may become so severe that fracture
or failure of the airfoils is possible.
[0005] Methods are known for mitigating the negative effects of
flutter. For example, many gas turbine engine systems include high
pressure compressors having variable vane rows (i.e., vanes that
are rotatable about a perpendicular axis relative to a longitudinal
centerline axis of the gas turbine engine). The variable vane rows
have been used effectively to schedule the engine around flutter
conditions by controlling the angle of incidence of the airfoils
relative to a direction of flowing airflow. Also, bleed or valve
systems are known which bleed airflow downstream from the airfoils
to throttle airflow and mitigate flutter. Additionally, airfoil
designs are known which tailor a leading edge of each airfoil to
obtain improved local airfoil incidence and adjacent airfoils
having different natural frequencies. Finally, having inconsistent
airfoil spacing in a forward stage varies the intermittent air
pulses communicated to a following airfoil stage, thus reducing
natural frequency excitation. Disadvantageously, all of these
methods result in system compromises, small to moderate performance
losses and may be expensive to incorporate into existing gas
turbine engine systems.
[0006] Accordingly, it is desirable to provide a gas turbine engine
having a closed-loop flutter sensing system which achieves reduced
flutter operation and minimizes performance losses of the gas
turbine engine.
SUMMARY OF THE INVENTION
[0007] A gas turbine engine system includes a nacelle, a fan casing
within the nacelle, a variable area fan nozzle, a sensor and a
controller. The sensor detects a flutter condition. The controller
communicates with the sensor and is operable to move the variable
area fan nozzle in response to a detection of the flutter
condition.
[0008] A method of controlling a gas turbine engine includes
sensing a flutter condition and increasing a discharge airflow area
associated with a variable area fan nozzle in response to sensing
the flutter condition.
[0009] The various features and advantages of this invention will
become apparent to those skilled in the art from the following
detailed description. The drawings that accompany the detailed
description are briefly described below.
BRIEF DESCRIPTION OF THE DRAWINGS
[0010] FIG. 1 illustrates a general partial cut-away view of a gas
turbine engine;
[0011] FIG. 2 is a perspective view of a section of a variable area
fan nozzle (VAFN);
[0012] FIG. 3 is a schematic view of an example gas turbine engine
having a variable area fan nozzle (VAFN); and
[0013] FIG. 4 illustrates a partial cut-away view of a fan section
of the gas turbine engine.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
[0014] FIG. 1 illustrates a gas turbine engine 10 which suspends
from a pylon 11 and may include (in serial flow communication) a
fan section 12, a low pressure compressor 14, a high pressure
compressor 16, a combustor 18, a high pressure turbine 20 and a low
pressure turbine 22. During operation, air is pulled into the gas
turbine engine 10 by the fan section 12, is pressurized by the
compressors 14, 16, and is mixed with fuel and burned in the
combustor 18. Hot combustion gases generated within the combustor
18 flow through the high and low pressure turbines 20, 22, which
extract energy from the hot combustion gases.
[0015] In a two spool design, the high pressure turbine 20 utilizes
the extracted energy from the hot combustion gases to power the
high pressure compressor 16 through a high speed shaft 19, and a
low pressure turbine 22 utilizes the energy extracted from the hot
combustion gases to power the low pressure compressor 14 and the
fan section 12 though a low speed shaft 21. However, the invention
is not limited to the two spool gas turbine architecture described
and may be used with other architectures such as a single spool
axial design, a three spool axial design and other architectures.
That is, the present invention is applicable to any gas turbine
engine, and to any application.
[0016] The example gas turbine engine 10 is in the form of a high
bypass ratio turbofan engine mounted within a nacelle 26, in which
a significant amount of the air pressurized by the fan section 12
bypasses the core engine for the generation of propulsion thrust.
The nacelle 26 partially surrounds a fan casing 28 and an engine
casing 31. The example illustrated in FIG. 1 depicts a high bypass
flow arrangement in which approximately 80% of the airflow entering
the fan section 12 may bypass the core engine via a fan bypass
passage 30 which extends between the nacelle 26 and the core engine
casing 31 for receiving and communicating a discharge airflow F1.
The high bypass flow arrangement provides a significant amount of
thrust for powering an aircraft.
[0017] In one example, the bypass ratio (i.e., the ratio between
the amount of airflow communicated through the fan bypass passage
30 relative to the amount of airflow communicated through the core
engine itself) is greater than ten, and the fan section 12 diameter
is substantially larger than the diameter of the low pressure
compressor 14. The low pressure turbine 22 has a pressure ratio
that is greater than five, in one example. The engine 10 may
include a gear train 23 which reduces the speed of the rotating fan
section 12. The gear train 23 can be any known gear system, such as
a planetary gear system with orbiting planet gears, a planetary
system with non-orbiting planet gears, or other type of gear
system. In the disclosed example, the gear train 23 has a constant
gear ratio. It should be understood, however, that the above
parameters are only exemplary of a contemplated geared turbofan
engine. That is, the invention is applicable to a traditional
turbofan engine as well as other engine architectures.
[0018] The discharge airflow F1 is communicated within the fan
bypass passage 30 and is discharged from the engine 10 through a
variable area fan nozzle (VAFN) 40 defined radially between the
nacelle 26 and the core engine casing 31. Core exhaust gases C are
discharged from the core engine through a core exhaust nozzle 32
defined between the core engine casing 31 and a center plug 34
defined coaxially therein around a longitudinal centerline axis A
of the gas turbine engine 10.
[0019] In one example, the VAFN 40 concentrically surrounds the
core engine casing 31 near an aftmost segment 29 of the nacelle 26.
However, the VAFN 40 may be positioned at other locations of the
engine 10. A discharge airflow area 36 is associated with the VAFN
40 and extends between the VAFN 40 and the core engine casing 31
for axially discharging the fan discharge airflow F1.
[0020] FIG. 2 illustrates the components of the VAFN 40. This
structure is exemplary only, and, as other embodiments would
similarly vary the discharge airflow area 36, will only be briefly
discussed herein. The VAFN 40 generally includes a synchronizing
ring 41, a static ring 43 and at least one flap assembly 45. Other
VAFN actuation mechanisms may be used. The flap assembly 45 is
pivotally mounted to the static ring 43 at multiple hinges 47 and
linked to the synchronizing ring 41 through a linkage 49. An
actuator assembly 51 selectively rotates the synchronizing ring 41
relative to the static ring 43 to adjust the flap assembly 45
through the linkage 49. The radial movement of the synchronizing
ring 41 is converted to tangential movement of the flap assembly 45
to vary the discharge airflow area 36 of the VAFN 40, as is further
discussed below.
[0021] FIG. 3 illustrates a flutter sensing system 50 of the gas
turbine engine 10. The discharge airflow area 36 may be influenced
during certain flight conditions, such as flutter conditions, by
opening or closing the VAFN 40. Flutter conditions represent
self-induced oscillations. Flutter conditions are caused by
unsteady aerodynamic conditions such as the interaction between
adjacent airfoils. During flutter, aerodynamic forces couple with
each airfoil's elastic and inertial forces, which may increase the
kinetic energy of each airfoil and produce negative damping. The
negative damping is enhanced where adjacent airfoils begin to
vibrate together.
[0022] In one example, the VAFN 40 is moveable between a first
position X and a second position X' (represented by phantom lines).
A discharge airflow area 37 of the second position X' is greater
than the discharge airflow area 36 of the first position X.
[0023] The VAFN 40 is selectively moved to the second position X'
to control the air pressure of the discharge airflow F1 within the
fan bypass passage 30. For example, closing the VAFN 40 (i.e.,
moving the VAFN to the first position X) reduces the discharge
airflow area which restricts the fan airflow F1 and produces a
pressure build up (i.e., an increase in air pressure) within the
fan bypass passage 30. Opening the VAFN 40 to the second position
X' increases the discharge airflow area, allowing additional fan
airflow, which reduces the pressure build up (i.e., a decrease in
air pressure) within the fan bypass passage 30. That is, opening
the VAFN 40 creates additional thrust power for the gas turbine
engine 10.
[0024] The flap assemblies 45 (See FIG. 2) of the VAFN 40 are moved
from the first position X to the second position X' in response to
detecting a flutter condition of the gas turbine engine 10, in one
example. In another example, the VAFN 40 is moved in response to
detecting a cross-wind condition. However, it should be understood
that the VAFN 40 may additionally be actuated in response to other
operability conditions such as take-off or ground operations.
[0025] The flutter sensing system 50 is a closed-loop system and
includes a sensor 52 and a controller 54. The sensor 52 actively
and selectively detects the flutter condition and communicates with
the controller 54 to move the VAFN 40 between the first condition X
and the second position X' or any intermediate position via the
actuator assemblies 51. Of course, this view is highly schematic.
In one example, the sensor 52 is a time of arrival type sensor. A
time of arrival sensor times the passage (or arrival time) of an
airfoil as the airfoil passes a fixed, case-mounted sensor as the
airfoil rotates about the engine longitudinal centerline axis A. In
the example shown in FIG. 3, the arrival time of the fan section 12
airfoils 60 are timed by the sensor 52. Of course, other airfoils
may similarly be timed. The controller 54 is programmed to
differentiate between which airfoil arrival times correlate to a
flutter condition and which airfoil arrival times correlate to
non-flutter conditions.
[0026] It should be understood that the sensor 52 and the
controller 54 are programmable to detect flutter conditions or
other conditions. A person of ordinary skill in the art having the
benefit of the teachings herein would be able to select an
appropriate sensor 52 and program the controller 54 with the
appropriate logic to communicate with the sensor 52 and the
actuator assembly 51 to move the VAFN 40 between the first position
X and the second position X' or any intermediate position in
response to a flutter condition or any other condition.
[0027] The VAFN 40 is returned to the first position X from the
second position X', which is otherwise indicated when the flutter
conditions subside. In one example, the sensor 52 communicates a
signal to the controller 54 where the flutter conditions are no
longer detected by the sensor 52. Therefore, the efficiency of the
gas turbine engine 10 is improved during both flutter and
non-flutter conditions. Also, airfoil damage due to continued
operation in a flutter condition is reduced.
[0028] FIG. 4 illustrates an example mounting location for the
sensor 52 of the flutter sensing system 50. In one example, the
sensor 52 is mounted to the fan casing 28 which surrounds the fan
section 12. In another example, the sensor 52 is mounted directly
adjacent to a blade tip area T of the fan section 12. The blade tip
area T of the fan section 12 is the area of the fan casing 28 which
is directly adjacent to the tips 62 of each airfoil 60 (only one
shown in FIG. 4) of the fan section 12 as the airfoils 60 are
rotated about the engine centerline axis A. In yet another example,
multiple sensors 52 are circumferentially disposed about the core
engine casing 31 adjacent to the blade tip area T of each airfoil
60. The sensor 52 may also be mounted adjacent to the blade tip
area of the airfoils of the compressor sections 14, 16 or the
turbine sections 20, 22.
[0029] The foregoing description shall be interpreted as
illustrative and not in any limiting sense. A worker of ordinary
skill in the art would recognize that certain modifications would
come within the scope of this invention. For that reason, the
following claims should be studied to determine the true scope and
content of this invention.
* * * * *