US20090317237A1

US20090317237A1 - System and method for reduction of unsteady pressures in turbomachinery

Info

Publication number: US20090317237A1
Application number: US12/142,940
Authority: US
Inventors: Trevor Howard Wood; Kishore Ramakrishnan; Richard David Cedar
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2008-06-20
Filing date: 2008-06-20
Publication date: 2009-12-24

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.

Description

BACKGROUND

The invention relates generally to turbomachines and, more particularly, to reducing unsteady pressures generated therein.
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.
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.
Therefore, an improved means of reducing the wake interaction effect is desirable.

BRIEF DESCRIPTION

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.
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

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:

FIG. 1 is a diagrammatic illustration of a gas turbine engine in accordance with the invention;

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;

FIG. 3 is a schematic graphical illustration of nullification of an exemplary acoustic wave in accordance with an embodiment of the invention; and

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

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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.

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:

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.