Patent Application
20170276070
The present disclosure generally relates to a gas turbine propulsion system, and more particularly relates to an axi-centrifugal compressor in which a proportioned distribution of the pressure rise across the axial and centrifugal compressor sections is achieved.
Gas turbine propulsion systems for aircraft must deliver high performance in a compact, lightweight configuration. This is particularly important in smaller jet propulsion systems typically used in regional and business aviation applications as well as in other turbofan, turboshaft, turboprop and rotorcraft applications. A well-known way to improve engine efficiency is to increase the overall pressure rise of the compressor. However, this is typically done by increasing the number of compressor stages, which increases the volume, weight, and cost of the engine; all of which reduce the value and competitive position of the engine. Larger commercial transport engines typically utilize “all-axial” high pressure compressors. Relative to the larger commercial transport engines, these relatively-smaller propulsion engines more frequently utilize “axi-centrifugal” compressors. Compressors of this type typically include one or more axial compressor stages followed by a centrifugal stage. The centrifugal compressor provides a second mechanism to increase overall pressure rise but often results in an increase of the diameter of the impeller in the centrifugal stage, resulting in increased weight and engine size.
Moreover, axial and centrifugal compressors have different optimal operating regimes, different physical characteristics, and different sensitivities to geometric and/or operational variation. As a result, conventional gas turbine propulsion systems pursue a very high level of aerodynamic loading in the centrifugal compressor section, relative to the axial compressor section, in an attempt to benefit from the centrifugal stabilizing effect and performance robustness. Unfortunately, this may require a relatively large and heavy centrifugal compressor section to achieve the overall pressure rise across the compressor.
Thus, there is a need for an axi-centrifugal compressor configuration, which strategically distributes the pressure rise between the axial and centrifugal compressor sections so that the compressor may deliver the overall pressure rise with fewer stages, lighter weight, and lower cost than previously available without compromising other critical performance characteristics of the engine such as compressor stability and durability. A compressor section of this type in a gas turbine propulsion system will meet the demands for a new level of performance for a smaller propulsion engine to robustly provide a high level of overall pressure rise with the highest efficiency and stability, while concurrently decreasing parts count, length, weight, and cost.
This summary is provided to describe select concepts in a simplified form that are further described in the Detailed Description. This summary is not intended to identify key or essential features of the claimed subject matter, nor is it intended to be used as an aid in determining the scope of the claimed subject matter.
The present disclosure provides an axi-centrifugal compressor configuration having a proportioned pressure ratio distribution across the compressor. In accordance with the present disclosure, a relatively high level of pressure rise is provided on the axial portion of the compressor, which leads to a reduction in the diameter and weight of the centrifugal stage compared to conventional compressor designs. A high level of pressure rise is also provided on each individual axial stage to minimize the stage count, which leads to a shorter length, lighter weight, lower cost, and high performance compared to conventional compressor designs. Aerodynamic over-loading of the axial stages that would otherwise result in boundary layer separations, low efficiency, and poor compressor operability is minimized.
An axi-centrifugal compressor is provided in a gas turbine propulsion system typically used in regional and business aviation applications as well as in other turbofan, turboshaft, turboprop and rotorcraft applications (e.g., less than 15 klbf Sea Level Take Off Thrust). The compressor is rotatably supported on a shaft assembly in a housing and operable to affect a pressure ratio along a flow path between a compressor inlet and a compressor exit. An axial compressor section having one or more axial stages is operable to affect a first pressure ratio (PRax) along the flow path between the compressor inlet and a first section exit. A centrifugal compressor section is operable to affect a second pressure ratio (PRc) along the flow path between a second section inlet and the compressor exit. The first section exit associated with the axial compressor section is in fluid communication with the second section inlet associated with the centrifugal section along the flow path. The compressor has a tuning factor defined as PRax/PRc in the range between 2.8 and 4.5 and a loading factor defined as (PRax)1/n/PRc in the range between 0.6 and 0.8, where n is the number of stages in the axial compressor section.
A method is also provided for compressing a fluid along a flow path in a gas turbine propulsion system typically used in regional and business aviation applications as well as in other turbofan, turboshaft, turboprop and rotorcraft applications (e.g., less than 15 klbf Sea Level Take Off Thrust). The fluid is drawn in at a first inlet pressure along the flow path through a first inlet. The fluid is compressed along the flow path in an axial compressor section having one or more axial stages downstream from the first inlet to a first exit in the axial compressor section at a first outlet pressure. A first pressure ratio (PRax) is affected across the axial compressor section. The fluid is communicated from the first exit into a second inlet along the flow path at a first outlet pressure. The fluid is compressed along the flow path in a centrifugal compressor section from the second inlet to a second exit. A second pressure ratio is affected across the centrifugal compressor section. The pressure rise through the compressor is distributed between the axial and centrifugal compressor sections such that a tuning factor defined as PRax/PRc is in the range between 2.8 and 4.5 and a loading factor defined as (PRax)1/n/PRc is in the range between 0.6 and 0.8, where n is the number of stages in the axial compressor section.
Furthermore, other desirable features and characteristics of the apparatus and method will become apparent from the subsequent detailed description and the appended claims, taken in conjunction with the accompanying drawings and the preceding background.
The present disclosure will hereinafter be described in conjunction with the following drawing figures, wherein like numerals denote like elements, and wherein:
The following detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. As used herein, the word “exemplary” means “serving as an example, instance, or illustration.” Thus, any embodiment described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other embodiments. All of the embodiments described herein are exemplary embodiments provided to enable persons skilled in the art to make or use the invention and not to limit the scope of the invention which is defined by the claims. Furthermore, there is no intention to be bound by any expressed or implied theory presented in the preceding technical field, background, brief summary, or the following detailed description.
Broadly, exemplary embodiments discussed herein include an axi-centrifugal compressor configuration having an proportioned pressure ratio distribution across the compressor. A relatively high pressure rise is provided on the axial portion of the compressor and across each stage in the axial compressor section. In addition, the aerodynamic over-loading in each axial stage is minimized. In this way, a proportioned pressure rise can be achieved in a small, light-weight and cost-effective manner.
Reference now is made to the drawings in which
The turbine section 140 rotates to drive equipment in the engine 100 via rotors or spools concentrically disposed about an axis of rotation 170 within the shaft assembly 160. Specifically, the turbine section 140 may include one or more rotors 142, 144 driven by the expanding exhaust fluids to rotate to the shaft assembly 160 and drive the compressor section 120 including the axial compressor section 122 and the centrifugal compressor section 124. While
The axial compressor section 122 progressively compresses fluids flowing generally axially (i.e., parallel to axis 170) along the flow path 180. The axial compressor section 122 may include one or more axial compressor stages 122.1, 122.2, 122.3. For example, as shown in
The centrifugal compressor section 124 compresses the fluid and directs the flow radially outward (i.e., in a direction which increases in a radial direction away from the axis 170) through an impeller assembly 210 driven on the shaft assembly 160. The rotor assemblies 200, 202, 204 and the impeller assembly 210 shown in
As indicated above, the axial compressor section 122 includes a first stage 122.1 immediately downstream of the compressor inlet 182, a second stage 122.2 downstream of the first stage 122.1, and a third stage 122.3 downstream of the second stage 122.2. Each of the axial compressor stages 122.1, 122.2, 122.3 contributes to a pressure rise from the compressor inlet 182 to the axial compressor exit 186. The performance of the axial compressor section 122 can be characterized according to a first pressure rise (TPRA) and a first pressure ratio (PRax) across the axial compressor section 122, as well as the pressure ratio per axial stage 122.1, 122.2, 122.3 (PR/stage ax) as provided below:
TPR
A
=PE
1
−PI
1
PR
ax
=PE
1
/PI
1
PR/stage ax=(PE1/PI1)1/n
wherein:
Likewise, the performance of the centrifugal compressor section 124 can be characterized according to a second pressure rise (TPRc) and a second pressure ratio (PRc) across the centrifugal compressor section 124 as provided below:
TPR
c
=PE
2
−PI
2
PR
c
=PE
2
/PI
2
wherein:
As noted above, operation of the compressor section 120, and particularly the contribution of the axial compressor section 122 and the centrifugal compressor section 124 are proportioned, while the aerodynamic over-loading of the axial stages 122.1, 122.2, 122.3 is minimized. Specifically, the compressor 120 has a tuning factor which satisfies the following condition:
and a loading factor which satisfies the following condition:
While a tuning factor in the range between 2.8 and 4.5 provides advantages described herein, additional advantages may be gained for a tuning factor in the range between 3.5 and 4.0. Likewise, while a loading factor in the range between 0.6 and 0.8 provide advantages described herein, additional advantages may be gained for a loading factor in the range between 0.65 and 0.75.
To satisfy these conditions, a proportioned distribution of pressure rise across the axial stage compressor 122 and the centrifugal compressor 124, as well as across axial stage 122.1, 122.2, 122.3 (i.e., stage matching) must be achieved. Moreover, the span-wise gradient of pressure rise across each axial stage should be configured to maximize the overall pressure rise attained. While a basic axial compressor section 122 has been illustrated and described herein, one skilled in the art will understand that additional compressor elements may be included to ensure that the axi-centrifugal compressor 120 satisfies the tuning and loading factors over the range of expected operating conditions. For example, the use of stability enhancing devices such as rotor tip casing treatments; variable stagger inlet guide vanes and stators; boundary layer separation control devices such as fluidic actuators (suction and blowing), vortex generators and plasma actuators placed on the end walls and/or airfoils to control destabilizing boundary layer separations; and airflow bleed (in or out) may be used to modify the stage matching at different operating conditions.
A compressor section 120 satisfying both of these conditions effectively distributes the pressure rise among the axial compressor stage 122 and the centrifugal compressor section 124 so that the compressor section 120 may achieve the desired overall pressure rise with fewer stages, lighter weight and lower cost as compared to conventional compressor sections. In particular, a compressor section 120 with a tuning factor within a range between 2.8 and 4.5 and a loading factor within a range between 0.6 and 0.8 achieves a relatively high pressure rise across the axial compressor section by providing a high level of pressure rise on each individual axial stage, while avoiding detrimental aerodynamic over-loading of any axial stage. The result is a compact and efficient compressor section in which the axial compressor section is shorter, lighter, lower cost and higher performance, and the centrifugal compressor section is smaller in diameter and lower in weight.
Unless otherwise explicitly indicated, the term “pressure” as used herein is intended to mean a total pressure at a given location, for example the total pressure at an inlet or exit of either the axial compressor section or the centrifugal compressor section. “Total pressure” refers to the sum of static pressure and dynamic pressure as expressed by Bernoulli's principle. Any contribution attributable to gravitational head may also be included in the “total pressure.” Accordingly, the terms “pressure rise” and “pressure ratio” are also considered in terms of total pressures, unless explicitly indicated otherwise.
In this document, relational terms such as first and second, and the like may be used solely to distinguish one entity or action from another entity or action without necessarily requiring or implying any actual such relationship or order between such entities or actions. Numerical ordinals such as “first,” “second,” “third,” etc. simply denote different singles of a plurality and do not imply any order or sequence unless specifically defined by the context in which it is used. The sequence of the text in any of the claims does not imply that process steps must be performed in a temporal or logical order according to such sequence unless it is specifically defined by the context in which it is used. The process steps may be interchanged in any order without departing from the scope of the invention, provided an interchange in order does not contradict the claim language and is not logically nonsensical.
Furthermore, depending on the context, words such as “connect” or “coupled to” used in describing a relationship between different elements do not imply that a direct physical connection must be made between these elements. For example, two elements may be connected to each other physically, electronically, logically, or in any other manner, through one or more additional elements.
While at least one exemplary embodiment has been presented in the foregoing detailed description of the invention, it should be appreciated that a vast number of variations exist. It should also be appreciated that the exemplary embodiment or exemplary embodiments are only examples, and are not intended to limit the scope, applicability, or configuration of the invention in any way. Rather, the foregoing detailed description will provide those skilled in the art with a convenient road map for implementing an exemplary embodiment of the invention. It being understood that various changes may be made in the function and arrangement of elements described in an exemplary embodiment without departing from the scope of the invention as set forth in the appended claims.
