Patent Application
20200232330
The subject matter disclosed herein generally relates to gas turbine engines and, more particularly, to a method and apparatus for improving the flow capacity by introducing the additional geometry design parameters of recesses on blades or vanes of gas turbine engines during transonic flow.
High performance with supersonic airfoils of a gas turbine engine is often challenging due to the presence of a strong shock inside the passage of two adjoining airfoils, which is referred to as a passage shock. Passage shocks may form proximate a throat between two adjacent blades or stator vanes as air flowing through a gas turbine engine between the two adjacent blades or stator vanes increases in velocity and turns transonic. The throat may be the aerodynamic throat or geometric throat between two adjacent blades or stator vanes. The shock waves may tend to impede air flowing through the throat between the two adjacent blades or stator vanes. The passage shock may incur high losses and limits increases in flow capacity.
According to one embodiment, a component system of a gas turbine engine is provided. The component system including: a first component having an outer surface; a second component having an outer surface, the second component and the first component being in a facing spaced relationship defining an air passageway therebetween; and a first recess located in at least one of the outer surface of the first component proximate the air passage and the outer surface of the first component proximate the air passage, wherein the first recess is located proximate a throat within the air passageway stretching between the first component and the second component.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first component and the second component are blades.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first component and the second component are stator vanes.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first recess is located in the outer surface of the first component proximate the air passage, and wherein the component system further comprises: a second recess located in the outer surface of the second component proximate the air passage, wherein the second recess is located proximate the throat within the air passageway stretching between the first component and the second component.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first component further includes a suction side proximate the air passage and a pressure side opposite the suction side, wherein the second component further comprises a pressure side proximate the air passage and a suction side opposite the pressure side, wherein the first recess is located in the outer surface of the first component proximate the air passage on the suction side of the first component, and wherein the second recess is located in the outer surface of the second component proximate the air passage on the pressure side of the second component.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first component further includes a suction side proximate the air passage and a pressure side opposite the suction side, and wherein the first recess is located in the outer surface of the first component proximate the air passage on the suction side of the first component.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the second component further comprises a pressure side proximate the air passage and a suction side opposite the pressure side, wherein the first recess is located in the outer surface of the second component proximate the air passage on the pressure side of the second component.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first recess is located in the outer surface of the first component proximate the air passage, and wherein the first recess stretches at least partially across a span of the first component.
In addition to one or more of the features described above, or as an alternative, further embodiments may include: a second recess located in the outer surface of the second component proximate the air passage, wherein the second recess is located proximate the throat within the air passageway stretching between the first component and the second component.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the second recess stretches at least partially across a span of the second component.
According to another embodiment, a gas turbine engine is provided. The gas turbine engine including: a fan section; a compressor section; a turbine section; and a component system located within at least one of the fan section, the compressor section, and the turbine section, the component system comprising: a first component having an outer surface; a second component having an outer surface, the second component and the first component being in a facing spaced relationship defining an air passageway therebetween; and a first recess located in at least one of the outer surface of the first component proximate the air passage and the outer surface of the first component proximate the air passage, wherein the first recess is located proximate a throat within the air passageway stretching between the first component and the second component.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first component and the second component are blades.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first component and the second component are stator vanes.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first recess is located in the outer surface of the first component proximate the air passage, and wherein the component system further comprises: a second recess located in the outer surface of the second component proximate the air passage, wherein the second recess is located proximate the throat within the air passageway stretching between the first component and the second component.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first component further includes a suction side proximate the air passage and a pressure side opposite the suction side, wherein the second component further comprises a pressure side proximate the air passage and a suction side opposite the pressure side, wherein the first recess is located in the outer surface of the first component proximate the air passage on the suction side of the first component, and wherein the second recess is located in the outer surface of the second component proximate the air passage on the pressure side of the second component.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first component further includes a suction side proximate the air passage and a pressure side opposite the suction side, and wherein the first recess is located in the outer surface of the first component proximate the air passage on the suction side of the first component.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the second component further comprises a pressure side proximate the air passage and a suction side opposite the pressure side, wherein the first recess is located in the outer surface of the second component proximate the air passage on the pressure side of the second component.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first recess is located in the outer surface of the first component proximate the air passage, and wherein the first recess stretches at least partially across a span of the first component.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the component system further comprises: a second recess located in the outer surface of the second component proximate the air passage, wherein the second recess is located proximate the throat within the air passageway stretching between the first component and the second component.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the second recess stretches at least partially across a span of the second component.
The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, that the following description and drawings are intended to be illustrative and explanatory in nature and non-limiting.
The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike:
The detailed description explains embodiments of the present disclosure, together with advantages and features, by way of example with reference to the drawings.
A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the Figures.
High performance with supersonic airfoils of a gas turbine engine is often challenging due to the presence of a strong shock inside the passage of two adjoining airfoils, which is referred to as a passage shock. Passage shocks may form proximate a throat between two adjacent blades or stator vanes as air flowing through a gas turbine engine between the two adjacent blades or stator vanes increases in velocity and turns transonic. The throat may be the aerodynamic throat or geometric throat between two adjacent blades or stator vanes. The shock waves may tend to impede air flowing through the throat between the two adjacent blades or stator vanes. The passage shock may incur high losses and limits increases in flow capacity.
Inlet conditions to the airfoil such as Mach number, air angle and airfoil geometry determines and control the strength and location of this passage shock. This passage shock will also vary along the airfoil surface depending on the operating flight condition. Specific goals for an airfoil design such as reducing losses, increasing operating range or increasing flow capacity are typical objectives that can be in conflict with one another. Targeting the airfoil design to meeting one of these goals will end up having an optimized shape that may be unique for that condition at the expense of the others. The embodiments disclosed herein seek to address these issues by manipulating the shape along the surface of the airfoil at different locations in effect optimizing for each specific goal.
Embodiments disclosed herein include apparatuses and methods to reduce, delay, and/or eliminate the shock waves forming between two adjacent blades of a gas turbine engine. Advantageously, by placing a recess (i.e., indentation or groove) in the blade proximate the throat where the shock wave is predicted to form, the shockwave may be reduced, delayed, and/or eliminated because the recess expands the through as discussed further herein.
The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. An engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The engine static structure 36 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 m/sec).
Referring now to
Each of the blades 100 includes a leading edge 106 and a trailing edge 108 opposite the leading edge 106 of the blade 100. The blade 100 spans from the leading edge 106 to the trailing edge 108 along the chord X1 of the blade 100. The blade 100 may have an airfoil shape, as shown in
The blades 100 are in a facing spaced relationship forming an air passage 102 therebetween. The two adjacent blades 100 are separated by a distance D1 defining the size of the air passage 102. The distance D1 (i.e., size of the air passage) varies in size along the chord X1 from the leading edge 106 of the blade 100 to the trailing edge 108 of the blade 100. The distance D1 (i.e., size of the air passage) also varies in size along the span Y1 from the root 122 of the blade 100 to the tip 124 of the blade 100 (see
The air passage 102 includes a throat 101. The throat 101 is located where the blades 100 are closest together along the chord X1 and the span Y1 (i.e., where the distance D1 is at a minimum). Passage shock waves begin to form proximate the throat 101 between the two adjacent blades 100 as air 10 flowing through a gas turbine engine 20 between the two adjacent blades 100 increases in velocity and turns transonic. The shock waves may tend to impede air 10 flowing through the throat 101 between the two adjacent blades 100.
One or both of the blades 100 may include a recess 150 in the outer surface 110 of blade 100 proximate the throat 101 and proximate the air passageway 102. There may be one or more recesses 150 along the chord X1 of the blade 100. The recess 150 is located proximate the throat 101 within the air passageway 102 stretching between the first blade 101a and the second blade 101b. The recess 150 is recessed inward into the blade 100 relative to the outer surface 110 of the blade 100. The recess 150 may extend a width W1 over the outer surface 110 of the blade 100. The width W1 of the recess 150 may be measured along the chord X1 of the blade 100. The width W1 of the recess 150 may vary of the span Y1 (see
In the example illustrated in
Referring now to
Technical effects of embodiments of the present disclosure include locally recessing a surface of at least one of two adjacent blades proximate a throat between the two adjacent blades to reduce and/or eliminate shockwave formation proximate the throat.
The term “about” is intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application. For example, “about” can include a range of ±8% or 5%, or 2% of a given value.
The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof.
While the present disclosure has been described with reference to an exemplary embodiment or embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.
This invention was made with Government support under Contract No. N00014-09-D-0821-M801 awarded by the United States Navy. The Government has certain rights in the invention.
