U.S. patent application number 12/142940 was filed with the patent office on 2009-12-24 for system and method for reduction of unsteady pressures in turbomachinery.
This patent application is currently assigned to GENERAL ELECTRIC COMPANY. Invention is credited to Richard David Cedar, Kishore Ramakrishnan, Trevor Howard Wood.
Application Number | 20090317237 12/142940 |
Document ID | / |
Family ID | 41431480 |
Filed Date | 2009-12-24 |
United States Patent
Application |
20090317237 |
Kind Code |
A1 |
Wood; Trevor Howard ; et
al. |
December 24, 2009 |
SYSTEM AND METHOD FOR REDUCTION OF UNSTEADY PRESSURES IN
TURBOMACHINERY
Abstract
A turbomachinery system is provided. The system includes a first
set of blades and a second set of blades moving relative to the
first set of blades, wherein the second set of blades includes a
first subset of blades having multiple first geometric parameters.
The second set of blades also includes at least a second subset of
blades non-uniformly spaced circumferentially and axially relative
to the first subset of blades such that unsteady pressures
generated from the wakes of the first set of blades interacting
with the second set of blades is below an acceptable level.
Further, the at least second subset of blades include multiple
second geometric parameters that are different or identical to the
multiple first geometric parameters.
Inventors: |
Wood; Trevor Howard;
(Clifton Park, NY) ; Ramakrishnan; Kishore;
(Clifton Park, NY) ; Cedar; Richard David;
(Cincinnati, OH) |
Correspondence
Address: |
GENERAL ELECTRIC COMPANY;GLOBAL RESEARCH
PATENT DOCKET RM. BLDG. K1-4A59
NISKAYUNA
NY
12309
US
|
Assignee: |
GENERAL ELECTRIC COMPANY
SCHENECTADY
NY
|
Family ID: |
41431480 |
Appl. No.: |
12/142940 |
Filed: |
June 20, 2008 |
Current U.S.
Class: |
415/119 ;
29/889.21; 415/121.3 |
Current CPC
Class: |
F04D 29/321 20130101;
F01D 5/146 20130101; F04D 29/544 20130101; Y10T 29/49321 20150115;
Y02T 50/671 20130101; Y02T 50/60 20130101; F01D 25/04 20130101;
F04D 29/666 20130101; F05D 2260/96 20130101; Y02T 50/673 20130101;
F01D 5/26 20130101 |
Class at
Publication: |
415/119 ;
415/121.3; 29/889.21 |
International
Class: |
F04D 29/66 20060101
F04D029/66 |
Claims
1. A turbomachinery system, comprising: a first set of blades; and
a second set of blades moving relative to the first set of blades,
the second set of blades comprising: a first subset of blades
comprising a plurality of first geometric parameters; and at least
a second subset of blades non-uniformly spaced circumferentially
and axially relative to the first subset of blades such that
unsteady pressures generated from the wakes of the first set of
blades interacting with the second set of blades is below an
acceptable level, the at least second subset of blades comprising a
plurality of second geometric parameters that are different or
identical to the plurality of first geometric parameters.
2. The system of claim 1, wherein the first subset of blades is
stationary and the second set of blades is rotating.
3. The system of claim 1, wherein the first set of blades is
rotating and the second set of blades is stationary.
4. The system of claim 1, wherein the first set of blades and the
second set of blades are counter-rotating.
5. The system of claim 1, wherein the second set of blades further
comprises a third subset of blades spaced circumferentially and
axially relative to the first subset of blades and the second
subset of blades.
6. The system of claim 1, wherein the plurality of first geometric
parameters and the plurality of second geometric parameters
comprises a camber, a stagger, a chord, a thickness, a chordwise
distribution and a spanwise distribution of the camber, the
thickness and the stagger respectively.
7. The system of claim 1, wherein the turbomachinery system
comprises an aircraft engine, gas turbine, steam turbine, a wind
turbine, a hydro turbine, or a heating-ventillating-airconditioning
system.
8. A method for manufacturing a turbomachine comprising: providing
a first set of blades; and providing a second set of blades moving
relative to the first set of blades, the second set of blades
comprising: a first subset of blades comprising a plurality of
first geometric parameters; and at least a second subset of blades
non-uniformly spaced circumferentially and axially relative to the
first subset of blades such that unsteady pressures generated from
the wakes of the first set of blades interacting with the second
set of blades is below an acceptable level, the at least second
subset of blades comprising a plurality of second geometric
parameters that are different or identical to the plurality of
first geometric parameters.
9. The method of claim 8, wherein said providing a first set of
blades comprises rotating a first set of blades.
10. The method of claim 8, wherein said providing a second set of
blades comprises rotating the second set of blades.
11. The method of claim 8, wherein providing a second set of blades
further comprises providing a third subset of blades non-uniformly
spaced circumferentially and axially relative to the first subset
of blades and the second subset of blades.
12. The method of claim 8, wherein providing a second set of blades
further comprises providing a unique spacing circumferentially and
axially and unique geometric definition for each blade in the
second set.
Description
BACKGROUND
[0001] The invention relates generally to turbomachines and, more
particularly, to reducing unsteady pressures generated therein.
[0002] With increased public concern over aircraft-generated noise,
aircraft gas turbine engine manufacturers are faced with the
problem of developing new ways of effectively reducing noise. One
of the common noise sources includes noise generated by the
turbomachinery within the gas turbine engine. The turbomachinery
noise results from a relative motion of adjacent sets of blades,
typical of those found in compressors (including fans) and
turbines. For example, a compressor comprises multiple bladed
stages, each stage including a rotatable blade row and possibly a
stationary blade row. It has long been recognized that in
turbomachines one of the principal noise sources is the interaction
between the wakes of upstream blades and downstream blades moving
relative to the upstream set of blades. This wake interaction
results in noise at the upstream blade passing frequency and at its
harmonics, as well as broadband noise covering a wide spectrum of
frequencies.
[0003] One of the commonly used methods to reduce this wake
interaction noise is to increase the axial spacing between adjacent
sets of blades. This modification provides space for the wake to
dissipate before reaching the downstream set of blades, resulting
in less noise. Increased spacing can generally be applied to
turbomachines, however, increases in axial length of the machine
may be restricted by weight, aerodynamic performance losses, cost
and/or installation and space requirements.
[0004] Therefore, an improved means of reducing the wake
interaction effect is desirable.
BRIEF DESCRIPTION
[0005] In accordance with an embodiment of the invention, a
turbomachinery system is provided. The system includes a first set
of blades and a second set of blades moving relative to the first
set of blades. The second set of blades includes a first subset of
blades comprising multiple first geometric parameters. The second
set of blades also includes at least a second subset of blades
non-uniformly spaced circumferentially and axially relative to the
first subset of blades such that unsteady pressures generated from
the wakes of the first set of blades interacting with the second
set of blades is below an acceptable level, wherein the at least
second subset of blades comprising multiple second geometric
parameters that are different or identical to the first geometric
parameters.
[0006] In accordance with another embodiment of the invention, a
method for manufacturing a turbomachine is provided. The method
includes providing a first set of blades. The method also includes
providing a second set of blades moving relative to the first set
of blades. The second set of blades includes a first subset of
blades comprising multiple first geometric parameters. The second
set of blades also includes at least a second subset of blades
non-uniformly spaced circumferentially and axially relative to the
first subset of blades such that unsteady pressures generated from
the wakes of the first set of blades interacting with the second
set of blades is below an acceptable level. Further, the second
subset of blades comprises multiple second geometric parameters
that are different or identical to the multiple first geometric
parameters.
DRAWINGS
[0007] These and other features, aspects, and advantages of the
present invention will become better understood when the following
detailed description is read with reference to the accompanying
drawings in which like characters represent like parts throughout
the drawings, wherein:
[0008] FIG. 1 is a diagrammatic illustration of a gas turbine
engine in accordance with the invention;
[0009] FIG. 2 is a schematic top view of a two-dimensional
cross-section through an exemplary first set of blades and a second
set of blades in the turbomachinery system of FIG. 1;
[0010] FIG. 3 is a schematic graphical illustration of
nullification of an exemplary acoustic wave in accordance with an
embodiment of the invention; and
[0011] FIG. 4 is a flow chart representing steps in a method for
manufacturing a turbomachine in accordance with an embodiment of
the invention.
DETAILED DESCRIPTION
[0012] As discussed in detail below, embodiments of the invention
include a system and method for reduction of unsteady pressures in
turbomachinery. As used herein, the system and method are
applicable to various types of turbomachinery applications having
blade-wake interactions resulting in unsteady pressures. Further,
the term `unsteady pressures` as used herein refers to air unsteady
pressures and acoustics as well as blade surface unsteady pressures
that are also referred to as `aeromechanical loading`. Non-limiting
examples of such turbomachinery applications include turbojet,
turbofan, turbo propulsion engines, aircraft engines, gas turbines,
steam turbines, wind turbines, or water/hydro turbines.
[0013] FIG. 1 is a schematic illustration of an exemplary turbofan
gas turbine engine assembly 10 in accordance with the invention and
having a centerline axis 12. In the exemplary embodiment, engine
assembly 10 includes a fan assembly 13, a booster compressor 14, a
core gas turbine engine 16, and a low-pressure turbine 26 that is
coupled to fan assembly 13 and booster compressor 14. Fan assembly
13 includes a plurality of rotor fan blades 11 that extend
substantially radially outward from a fan rotor disk 15, as well as
a plurality of stator vanes 21 that are positioned downstream of
fan blades 11. Core gas turbine engine 16 includes a high-pressure
compressor 22, a combustor 24, and a high-pressure turbine 18.
Booster compressor 14 includes a plurality of rotor blades 40 that
extend substantially radially outward from a compressor rotor disk
20 coupled to a first drive shaft 31. Compressor 22 and
high-pressure turbine 18 are coupled together by a second drive
shaft 29. Engine assembly 10 also includes an intake side 28, a
core engine exhaust side 30, and a fan exhaust side 31.
[0014] During operation, air entering engine 10 through intake side
28 is compressed by fan assembly 13. The airflow exiting fan
assembly 13 is split such that a portion 35 of the airflow is
channeled into booster compressor 14 and a remaining portion 36 of
the airflow bypasses booster compressor 14 and core turbine engine
16 and exits engine 10 through fan exhaust side 31. This bypass air
36 flows past and interacts with the stators vanes 21 creating
unsteady pressures on the stator surfaces as well as in the
surrounding airflow that radiate as acoustic waves. The plurality
of rotor blades 40 compress and deliver compressed airflow 35
towards core gas turbine engine 16. Airflow 35 is further
compressed by the high-pressure compressor 22 and is delivered to
combustor 24. Airflow 35 from combustor 24 drives rotating turbines
18 and 26 and exits engine 10 through exhaust side 30.
[0015] FIG. 2 is a schematic top view of a two dimensional
cross-section through an exemplary first set of blades 52 and a
second set of blades 54 in the turbomachinery system 10 of FIG. 1.
The first set of blades 52 and the second set of blades 54 may be
located in the fan 11, booster 14, core compressor 22, or a turbine
stage 18, 26. In one embodiment, the compressor or turbine stage is
axial. In an alternative embodiment, the turbomachinery stage is
radial. In yet another embodiment, the turbomachinery stage is
mixed (radial and axial). In the illustrated embodiment, the first
set of blades 52 is rotating and the second set of blades 54 is
stationary. In an alternative embodiment, the first set of blades
52 may be stationary, while the second set of blades 54 rotates. In
yet another embodiment, the first set of blades and the second set
of blades may be counter rotating. The second set of blades 54
includes a first subset of blades 58 and at least a second subset
of blades 60. It should be noted that the second set of blades 54
may include a third subset of blades and so forth. The second
subset of blades 60 are non-uniformly spaced circumferentially, as
referenced by numeral 64 and axially, referenced by numeral 66,
relative to the first subset of blades 58 such that unsteady
pressures generated from the wakes of the first set of blades
interacting with the second set of blades is below an acceptable
level.
[0016] Various geometric parameters may be varied between the first
subset of blades 58 and the second subset of blades 60. For
example, a chord length, referenced by numeral 72, for the second
subset of blades 60 relative to the first subset of blades 58 may
be varied. In another embodiment, an inclination angle relative to
axial direction referred to as `stagger` referenced by numeral 74
and/or curvature of the blade referred to as `camber`,
respectively, may be varied relative to the first subset of blades
58. In another exemplary embodiment, a thickness of the first
subset of blades and the second subset of blades may be varied. In
yet another embodiment, a chordwise distribution of camber and/or
thickness may be varied. In another embodiment, the second set of
blades 54 may include a radial or spanwise distribution of the
foregoing parameters over different sets of blades.
[0017] As has been previously discussed, one of the principal
sources of unsteady pressures in turbomachinery is the interaction
between the wakes of the first set of blades 52 and the second set
of blades 54, moving relative to each other. As is well understood,
the wakes are defined as the region of reduced momentum behind an
airfoil evidenced by the aerodynamic drag of the blade. As
illustrated, the first set of blades 52 shed a wake 82 that is
impacted by representative second set of blades 54. However, if at
least a second subset of blades 60 are non-uniformly spaced
circumferentially and axially, the wake interaction will occur at
different and non-uniformly distributed instants of time. Further,
the first subset of blades 58 and the second subset of blades 60
may be optimally spaced such that the acoustic waves resulting from
such an interaction destructively interfere to produce less overall
noise, as described below. In another embodiment, the first subset
of blades 58 and the second subset of blades 60 may be optimally
spaced to reduce unsteady surface pressure loads on the blades 58,
60.
[0018] FIG. 3 is a schematic graphical illustration 90 of
nullification of an exemplary acoustic wave by non-uniform spacing
of the second set of blades 54 (FIG. 2). An exemplary acoustic
signal 92 representative of an acoustic wave is generated from the
interaction between the first set of blades 52 (FIG. 2) and a
subset 58 of the second set of blades 54 (FIG. 2) prior to
variation of geometric parameters. An optimal shift in the
circumferential and axial position of the second subset 60 of the
second set of blades 54 relative to 58, as described in FIG. 2,
induces acoustic radiation resulting in a signal 94 out of phase
with the signal 92. Thus, the signals 92 and 94 cancel each other
resulting in a signal 96 devoid of the acoustic energy originally
generated, a phenomenon commonly referred to as wave destructive
interference. It will be appreciated that the illustrated
embodiment is an ideal case. However, non-ideal cases may also
result in a significant or desirable reduction in noise.
[0019] FIG. 4 is a flow chart representing steps in a method 110
for manufacturing a turbomachine. The method 110 includes providing
a first set of blades in step 112. In one embodiment, a rotating
first set of blades is provided. A second set of blades is provided
in step 114 that moves relative to the first set of blades. In an
exemplary embodiment, a stationary second set of blades is
provided. The second set of blades includes a first subset of
blades having multiple first geometric parameters and at least a
second subset of blades non-uniformly spaced circumferentially and
axially relative to the first set of blades such that unsteady
pressures generated from the wakes of the first set of blades
interacting with the second set of blades is below an acceptable
level. Further, the at least second subset of blades has multiple
second geometric parameters that are different or identical to the
plurality of first geometric parameters. In one embodiment, the
second set of blades further includes a third subset of blades
non-uniformly spaced circumferentially and axially relative to the
first subset of blades and the second subset of blades. In another
embodiment, the second set of blades further includes any number of
subsets of blades non-uniformly spaced circumferentially and
axially relative to the all other subsets of blades, up to and
including the point where every blade in the second set of blades
is uniquely spaced circumferentially and axially and uniquely
defined by the multiple second geometric parameters relative to
every other blade in the second set.
[0020] The various embodiments of a system and method for reduction
of unsteady pressures in turbomachinery described above thus
provide a convenient and efficient means to reduce aerodynamic
noise and/or aeromechanical loading caused by interaction of wakes
between sets of blades moving relative to each other. The technique
provides non-uniform spacing between blades in a set of blades
resulting in a reduction in unsteady blade loading that also
results in reduced noise signals and/or a noise field that
superimposes in a way to reduce peak noise signals.
[0021] The technique can also be used to improve fuel bum by
redesigning other system or geometric parameters (e.g., reducing
the separation distance between adjacent sets of interacting
turbomachinery blades, thereby also reducing system weight) in such
a way to improve system efficiency, and employing the technique
described in this invention to maintaining desirable noise
levels.
[0022] It is to be understood that not necessarily all such objects
or advantages described above may be achieved in accordance with
any particular embodiment. Thus, for example, those skilled in the
art will recognize that the systems and techniques described herein
may be embodied or carried out in a manner that achieves or
optimizes one advantage or group of advantages as taught herein
without necessarily achieving other objects or advantages as may be
taught or suggested herein.
[0023] Furthermore, the skilled artisan will recognize the
interchangeability of various features from different embodiments.
For example, a third subset of blades described with respect to one
embodiment may include a geometric variation in stagger, camber and
thickness relative to a first subset and a second subset of blades
described with respect to another. This concept can also be
extended to the point where every blade in the set is designed
uniquely relative to all other blades in the set. Similarly, the
various features described, as well as other known equivalents for
each feature, can be mixed and matched by one of ordinary skill in
this art to construct additional systems and techniques in
accordance with principles of this disclosure.
[0024] While only certain features of the invention have been
illustrated and described herein, many modifications and changes
will occur to those skilled in the art. It is, therefore, to be
understood that the appended claims are intended to cover all such
modifications and changes as fall within the true spirit of the
invention.
* * * * *