Patent Grant
9115594
The present invention generally relates to gas turbine engine compressors, and more particularly, but not exclusively, to axial compressors used in gas turbine engines.
Improving operability and performance of gas turbine engine axial flow compressors using casing treatments remains an area of interest. Some existing systems have various shortcomings relative to certain applications. Accordingly, there remains a need for further contributions in this area of technology.
One embodiment of the present invention is a unique casing treatment for a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for axial flow compressor casing treatments. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.
a and 4b depicts velocity triangles of one form of the present application.
For the purposes of promoting an understanding of the principles of the invention, reference will now be made to the embodiments illustrated in the drawings and specific language will be used to describe the same. It will nevertheless be understood that no limitation of the scope of the invention is thereby intended. Any alterations and further modifications in the described embodiments, and any further applications of the principles of the invention as described herein are contemplated as would normally occur to one skilled in the art to which the invention relates.
With reference to
One form of the gas turbine engine 50 includes a compressor 52, combustor 54, and turbine 56. The gas turbine engine 50 can take the form of an axial flow engine, but other forms are also contemplated. In just one non-limiting example, the gas turbine engine 50 can be a mixed axial/centrifugal flow engine. In some embodiments the gas turbine engine 50 can take the form of an adaptive cycle or variable cycle engine. The gas turbine engine 50 can be a turbojet or turbofan engine, among other possible engine types.
Turning now to
The compressor casing 58 of the compressor 52 includes a portion having an annular shape that forms part of a working fluid flow path through the gas turbine engine 50. In some forms the compressor casing 58 can be coupled with structure located radially inward from the casing 58 and that forms part of the working fluid flow path through the gas turbine engine 50. In some embodiments the compressor casing 58 can be segmented and can be made from a variety of materials.
In one form the compressor casing 58 includes abradable sections 70 and 72 which are designed to deteriorate when engaged with a moving portion of the gas turbine engine 50, such as a tip of the blade 62. Though the illustrative embodiment includes abradable sections 70 and 72 affixed to the casing 58, other embodiments include abradable sections at the ends of moving portions such as, but not limited to, the blades 62. Though in some embodiments the compressor casing 58 includes only one of the abradable sections 70 and 72, in other embodiments the compressor casing 58 may not have either abradable sections 70 and 72. The abradable sections 70 and 72 can be applied to the compressor casing 58 using a variety of techniques such as, but not limited to spray coating, and electroplating. In other embodiments, however, the abradable sections 70 and 72 can be mechanically coupled to the casing 58.
The main rotor 60 is operable to rotate the blades 62 at a variety of speeds to provide a compression of a working fluid for the gas turbine engine 50. The main rotor 60 and blades 62 can be made from a variety of materials and can be coupled together using a variety of techniques to form a rotating assembly. In some forms the blades 62 can be integrally formed with the main rotor 60.
The shroud 64 is disposed at the end of the blades 62 and located between the blades 62 and the airflow flow members 66. The shroud includes an axially forward portion 74 and an axially rearward portion 76. The terms “forward” and “rearward” are used herein for convenience of description and are not intended to imply the orientation of the respective portions relative to the gas turbine engine 50 and/or an aircraft with which the gas turbine engine 50 may be used. The axially forward portion 74 is depicted in the illustrative embodiment as extending forward of the blade 62 leading edge. In some embodiments, however, the axially forward portion 74 may not extend forward of the blade 62 leading edge. The axially rearward portion 76 is located at about the mid-chord position of the blade 62 but can take on different locations in other embodiments. The shroud 64 is coupled to the ends of the blades 62 and can take the form of a unitary structure in some embodiments or a segmented assembly in others. In some embodiments the shroud 64 can be formed integral with the blades 62.
The airflow member 66 is used to tangentially turn a flow of compressed working fluid traversing the passage 68, as will be discussed further hereinbelow. The airflow member 66 is disposed radially outward of the shroud 64. In the illustrative embodiment the airflow member 66 is coupled to the shroud 64 but in other embodiments can be coupled to the compressor casing 58 and thus remain stationary relative to a moving main rotor 60. The airflow member 66 can be formed integrally with the shroud in some embodiments, or can be coupled using a variety of techniques in other embodiments. Any number of airflow members 66 can be used within the passage 68. In some forms the number of airflow members 66 used in the passage 68 can be the same as the number of blades 62 of the main rotor 60. The airflow member 66 can be an airfoil shape in some embodiments.
The passage 68 conveys a compressed working fluid from a passage inlet 78 to a passage outlet 80. In some forms the passage 68 can have a relatively constant flow area between the passage inlet 78 and passage outlet 80. The passage inlet 78 is oriented rearward of the axially rearward portion 76 and forward of the trailing edge of the blade 62. In other embodiments, however, the passage inlet 78 can be located elsewhere relative to the axially rearward portion 76 and the trailing edge of the blade 62. The passage inlet 78 in the illustrated embodiment is defined by the compressor casing 58, the shroud 64, but in other embodiments can be defined by other structure of the gas turbine engine 50. The passage outlet 80 is located forward of the axially forward portion 74 of the shroud 64. The passage outlet 80 is defined by the shroud 64 and abradable section 72, but in other embodiments can be defined by other structure. To set forth just one non-limiting example, if the gas turbine engine 50 lacked an abradable section 70, the passage inlet 78 can be defined between the shroud 64 and the compressor casing 58.
a and 4b depict the blade 62, and airflow member 66, respectively, along with their respective velocity triangles to better illustrate the embodiment of the present application that includes the airflow member 66 coupled to the shroud 64. In
As will be appreciated when comparing various aspects of the velocity triangles in
As the passage inlet portion 84 is extracted from the flow of working fluid 82 and turns to flow in a direction counter to the flow of working fluid 82, a number of observations can be made. When the flow is turned the absolute tangential velocity, cU, maintains a relatively constant angular momentum, and is nearly constant for small radius change. If the flow area in the tip passage is such that the magnitude of the axial velocity is unchanged, and the assumption made that cU is changed insignificantly, then the airflow member inlet velocity triangle 96 is the mirror image of the blade extraction velocity triangle 92. As a result of the orientation of the airflow member 66 and the direction of working fluid flowing through the passage 68, the absolute tangential velocity cU is reduced across the airflow member 66. As a result of reducing absolute tangential velocity cU, the Euler equation predicts a reduction in total temperature. The Euler equation can be expressed as Δh=(U*cU2−U*cU1), where h is specific enthalpy, cU2 is the absolute axial velocity downstream of the airflow member 66, cU1 is the absolute axial velocity upstream of the airflow member 66, and U the rotational speed of the rotor. Persons of skill in the field will appreciate that the thermodynamic result of a change in specific enthalpy is a corresponding decrease in total temperature. This result reduces and could conceivably eliminate the efficiency penalty of reworking the air through the passage 68.
One aspect of the present application provides an apparatus comprising a gas turbine engine compressor having a bladed rotor enclosed by a shroud and disposed within a portion of a compressor casing section, the bladed rotor having an upstream side and a downstream side, an airflow passage formed between the compressor casing section and the shroud, the airflow passage having an inlet and an outlet, the inlet located downstream relative to the outlet, and an airflow member disposed within the airflow passage.
Another aspect of the present application provides an apparatus comprising an axial compressor having a rotor including a plurality of blades and an air extraction portion and air insertion portion located on a tip side of the plurality of blades, a compressor shroud coupled to the ends of at least some of the plurality of blades, and an airfoil member coupled to the compressor shroud and operable to reduce an absolute tangential velocity of an airflow as the airflow traverses from the extraction portion to the insertion portion.
Yet another aspect of the present application provides an axial flow compressor of a gas turbine engine comprising a gas turbine engine compressor casing, a plurality of axial compressor blades operable to rotate at a velocity to provide a compression for the gas turbine engine, a shroud coupled to the ends of the plurality of axial compressor blades, a passage located between the shroud and the gas turbine engine compressor casing, and means for altering a velocity of an airflow that has been extracted from the plurality of blades and that is flowing through the passage during operation of the axial flow gas turbine engine compressor.
Still a further aspect of the present application provides a method comprising assembling an axial flow gas turbine engine casing, locating a bladed compressor rotor having a shroud within the axial flow gas turbine engine casing, and inserting a plurality of airflow members within a passage formed between the axial flow gas turbine engine casing and the shroud.
While the invention has been illustrated and described in detail in the drawings and foregoing description, the same is to be considered as illustrative and not restrictive in character, it being understood that only the preferred embodiments have been shown and described and that all changes and modifications that come within the spirit of the inventions are desired to be protected. It should be understood that while the use of words such as preferable, preferably, preferred or more preferred utilized in the description above indicate that the feature so described may be more desirable, it nonetheless may not be necessary and embodiments lacking the same may be contemplated as within the scope of the invention, the scope being defined by the claims that follow. In reading the claims, it is intended that when words such as “a,” “an,” “at least one,” or “at least one portion” are used there is no intention to limit the claim to only one item unless specifically stated to the contrary in the claim. When the language “at least a portion” and/or “a portion” is used the item can include a portion and/or the entire item unless specifically stated to the contrary.
The present application claims the benefit of U.S. Provisional Patent Application No. 61/427,702 filed Dec. 28, 2010 which is incorporated herein by reference
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