Turbine engine airfoil

Abstract

A turbine engine includes a compressor section, a combustion section, a turbine section, and an airfoil having an outer wall defining a pressure side and a suction side, and extending between a leading edge and a trailing edge to define a mid-curve. A first thickness is defined between the pressure side and the suction side at a first location along the mid-curve.

Description

Technical Field

[0001] This disclosure relates to turbine engine airfoils, and more specifically, to cooling airfoils with customized near-tail edge thickness. Background Technology

[0002] Gas turbine engines used in commercial aircraft typically consist of a fan and a turbine. The turbine (often referred to as the core) typically comprises a compressor section, a combustor section, and a turbine section arranged in a series flow pattern. The compressor section compresses air directed to the combustor section, where the air is mixed with fuel. The mixture is then ignited to generate hot combustion gases. The combustion gases are directed to the turbine section, where energy is extracted from the combustion gases to power the compressor section and to generate work, such as propelling the aircraft in flight or powering loads (e.g., generators).

[0003] Turbine engines are typically designed to operate at high temperatures to improve engine efficiency. Cooling measures are usually provided for engine components in high-temperature environments (such as airfoils). These cooling measures reduce material wear, extend component life, and provide increased structural stability during engine operation. Attached Figure Description

[0004] The complete and feasible disclosure of this disclosure, including its best mode, is set forth in the specification with reference to the accompanying drawings, for those skilled in the art, wherein:

[0005] Figure 1 This is a schematic cross-sectional view of a turbine engine based on the various aspects described herein.

[0006] Figure 2 It is applicable to the various aspects described in this article. Figure 1 A perspective view of the airfoil assembly of a turbine engine.

[0007] Figure 3 yes Figure 2 A cross-sectional view of the airfoil assembly.

[0008] Figure 4 yes Figure 3 A cross-sectional view of the trailing edge portion of the airfoil assembly.

[0009] Figure 5 It can be used Figure 1 A cross-sectional view of another trailing edge portion of another airfoil component of the turbine engine, similar to... Figure 4 .

[0010] Figure 6 This is a graph showing the parameter space of the airfoil assembly according to the various aspects described herein. Detailed Implementation

[0011] The aspects of this disclosure generally relate to airfoils. For illustrative purposes, this disclosure will be described in the context of aircraft gas turbine engines with respect to turbine airfoils. However, it should be understood that the aspects disclosed herein are not limited thereto and may have general applicability, including in other engines and in non-aircraft applications (e.g., other mobile applications and non-mobile industrial, commercial, and residential applications).

[0012] Reference will now be made in detail to the turbine engine and airfoil, one or more examples of which are shown in the accompanying drawings. Detailed descriptions use numbers and letter labels to refer to the features in the accompanying drawings.

[0013] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components.

[0014] As used herein, the terms “front” and “rear” refer to relative positions within a gas turbine engine or vehicle, and to the normal operating posture of the gas turbine engine or vehicle. For example, for a gas turbine engine, “front” refers to the position closer to the engine inlet, while “rear” refers to the position closer to the engine nozzle or exhaust port.

[0015] As used herein, the term "upstream" refers to the direction opposite to the direction of fluid flow, while the term "downstream" refers to the direction in the same direction as the fluid flow. The terms "forward" or "front" indicate being in front of something, and "backward" or "rear" indicate being behind something. For example, in the context of fluid flow, forward / frontward refers to upstream, and backward / rearward refers to downstream.

[0016] As used herein, the term "fluid" can refer to a gas or liquid, or a multiphase system. The terms "fluid connectivity" or "fluid connection" refer to the ability of fluids to establish a connection or link between specified areas.

[0017] As used herein, the term "radial" or "radially" refers to a direction away from a common center. For example, in the overall context of a turbine engine, radial refers to the direction of a ray extending between the engine's central longitudinal axis and the engine's outer perimeter.

[0018] As used herein, unless the context clearly specifies otherwise, the singular forms “a,” “an,” and “the” include plural references. Furthermore, as used herein, the term “group” or a “group” of elements can include any number of elements, including only one.

[0019] As used in this article, the "pressure drop" across an obstacle refers to the change in the total pressure of the fluid as it passes through the obstacle. The total pressure drop across the measurement boundary is typically characterized by a pressure loss coefficient Yp. Yp is determined by the following expression 1 and is usually expressed as a percentage:

[0020] (1) Wherein, PT1 refers to the average total pressure of the fluid upstream of the obstacle, PT2 refers to the average total pressure of the fluid downstream of the obstacle, PT0 refers to the total inlet pressure at the measurement boundary, and PS1 refers to the static outlet pressure at the measurement boundary.

[0021] As used in this article, “thickness” in some examples refers to the thickness of the component wall, or in some examples refers to the thickness across the component.

[0022] As used herein, "mid-curve" refers to a line positioned between the upper surface (e.g., suction side) and the lower surface (e.g., pressure side) of the airfoil. Such a mid-curve intersects the airfoil chord at the leading and trailing edges of the airfoil.

[0023] As used herein, “cooling performance” refers to the amount of cooling provided to a component by a given amount of fluid flow. High cooling performance indicates that a low cooling airflow to the component is required to achieve the desired component metal temperature. Conversely, low cooling performance indicates that a high cooling airflow is required to achieve the same desired component metal temperature.

[0024] As used herein, “specific fuel consumption” or “SFC” refers to the engine’s specific fuel consumption, which is related to engine efficiency. SFC represents the combined effect of at least aerodynamic efficiency and cooling performance (total cooling airflow), with aerodynamic efficiency for airfoils characterized by total pressure loss. Higher (poorer) SFC occurs when total pressure loss (Yp) is higher and / or when cooling performance is lower (higher cooling airflow to maintain component metal temperature).

[0025] As used in this paper, the "thickness parameter" (denoted as "τ") represents the relationship between two thicknesses taken across the airfoil and the distance between these two thicknesses taken along the mid-curve of the airfoil, as described in this paper.

[0026] All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, rear, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, backward, etc.) are used for identification purposes only to aid the reader's understanding of this disclosure and do not impose limitations, particularly regarding the location, orientation, or purpose of aspects of this disclosure described herein. Unless otherwise stated, connective references (e.g., attachment, connection, joint, and engagement) are to be interpreted broadly and may include intermediate structural elements between sets of elements as well as relative movement between elements. Therefore, a connective reference does not necessarily imply that two elements are directly connected and have a fixed relationship with each other. Exemplary figures are for illustrative purposes only, and the dimensions, positions, order, and relative sizes reflected in the accompanying figures may vary.

[0027] In certain exemplary embodiments of this disclosure, a gas turbine engine is provided that defines a centerline and a circumferential direction about the centerline. The gas turbine engine typically includes a compressor section, a combustion section, and a turbine section arranged in series along the centerline. The gas turbine engine may define a substantially annular flow path relative to the centerline. The gas turbine engine includes at least one airfoil, such as a turbine blade positioned within the flow path. Aspects of this disclosure as described herein may be applied to airfoils in the compressor section or the turbine section.

[0028] Traditional airfoil design is typically carried out in a linear or iterative manner, beginning with aerodynamic design, where the airfoil's geometry or shape is designed to minimize the total pressure drop across the working fluid flow path, followed by cooling design, where the cooling flow rate required to achieve the desired material temperature limits is determined. In some examples, turbine airfoils often operate in environments above 1000°C, requiring cooling flow even when high-temperature alloy metallurgy is applied. Engine specific fuel consumption (SFC) and efficiency depend on both aerodynamic and cooling design. In one example, SFC is determined as follows:

[0029] (2) in It refers to the mass flow rate of fuel, and T refers to the net thrust of the engine.

[0030] The aerodynamic losses caused by the geometry or shape of airfoils can reduce engine efficiency and increase fuel consumption. Additionally, airfoil cooling designs utilize cooling air extracted from the turbine engine's airflow, which could otherwise be used for power extraction within the main flow path.

[0031] More specifically, in terms of aerodynamic and cooling design, the airfoil trailing edge design has a significant impact on engine SFC (Short-terminal pressure loss). Cooling effectiveness increases as cooling features, such as film pores, move closer to the trailing edge. However, the airfoil and trailing edge become thicker to accommodate these features, which reduces aerodynamic performance (higher total pressure loss and lower efficiency). In the conventional linear design process detailed above, the aerodynamic shape is typically defined first to minimize total pressure loss (e.g., maximize efficiency and minimize SFC) within casting constraints. Cooling features are then typically fitted within the resulting airfoil profile. Cooling airflow levels are adjusted higher until the target metal temperature criterion is met. Higher cooling airflow levels also lead to increased SFC. In this way, these competing effects result in a trade-off between aerodynamic and cooling design considerations.

[0032] Given the above, the inventors' practice involves designing airfoils by using SFC as a system-level quality factor that incorporates the net effect of competing effects to simultaneously consider aerodynamic and cooling performance. Airfoil design features (such as the distribution of trailing edge diameter and airfoil thickness upstream of the trailing edge) are specified. Cooling features are then placed at the final position allowed by the geometry definition. The net airfoil design is then evaluated for both aerodynamic and cooling performance. Aerodynamic performance is evaluated, for example, via computational fluid dynamics (CFD) modeling or cascade testing. Cooling performance is then evaluated, for example, via computational modeling with calibrated empirical models or high-fidelity CFD modeling (such as large eddy simulation (LES)). Using engine cycle information, aerodynamic and cooling performance are translated into engine SFC. Multiple iterations are performed on airfoil geometry adjustments and SFC evaluations, sometimes incorporating the implementation of advanced machine learning algorithms to obtain the optimal SFC for that particular turbine airfoil. This iterative process can continue for a considerable period until a feasible airfoil design is determined.

[0033] Figure 1 This is a schematic cross-sectional view of a gas turbine engine 10 for use in an aircraft. The engine 10 has a generally longitudinally extending axis, or engine centerline 12, extending from the front 14 to the rear 16. The engine 10 includes the following downstream sequential flow relationships: a fan section 18 comprising a fan 20; a compressor section 22 comprising a booster or low-pressure (LP) compressor 24 and a high-pressure (HP) compressor 26; a combustion section 28 comprising a combustor 30; a turbine section 32 comprising an HP turbine 34 and an LP turbine 36; and an exhaust section 38.

[0034] Fan section 18 includes a fan housing 40 surrounding fan 20. Fan 20 includes a plurality of fan blades 42 arranged radially around engine centerline 12. HP compressor 26, combustor 30 and HP turbine 34 form the core 44 of engine 10 that generates combustion gases. Core 44 is surrounded by core housing 46, which is connected to fan housing 40.

[0035] An HP shaft or spool 48, coaxially arranged around the engine centerline 12 of the engine 10, drives the HP turbine 34 to the HP compressor 26. An LP shaft or spool 50, coaxially arranged within a larger diameter annular HP spool 48 around the engine centerline 12 of the engine 10, drives the LP turbine 36 to the LP compressor 24 and the fan 20. The spools 48 and 50 are rotatable about the engine centerline 12 and are connected to a plurality of rotatable elements that collectively define a rotor 51.

[0036] Although not shown, it should be understood that engine 10 may include other components, such as, but not limited to, a gearbox. The gearbox may be located anywhere suitable within the turbine engine, such that it connects one rotating part to another. The gearbox may connect fan 20 to LP spool 50. The gearbox may allow fan 20 to operate at a speed different from the rest of engine 10.

[0037] LP compressor 24 and HP compressor 26 each include multiple compressor stages 52 and 54, respectively, in which a set of compressor blades 56 and 58 rotate relative to a corresponding set of static compressor impeller blades 60 and 62 (also referred to as nozzles) to compress or pressurize the fluid flow passing through the stage. In a single compressor stage 52 or 54, the multiple compressor blades 56 and 58 are arranged in a ring and extend radially outward from the blade platform relative to the engine centerline 12 to the blade tips, while the corresponding static compressor impeller blades 60 and 62 are positioned upstream of and adjacent to the rotating blades 56 and 58. It is worth noting that... Figure 1 The number of blades, impellers, and compressor stages shown is selected for illustrative purposes only, and other numbers are also possible.

[0038] Blades 56 and 58 for the first stage of the compressor are mounted to disk 61, which is mounted to a corresponding one of HP spool 48 and LP spool 50, with each stage having its own disk 61. Blades 56 and 58 may also be part of an integral bladed disk, rather than mounted to a disk. Impeller blades 60 and 62 for the first stage of the compressor are mounted circumferentially to the core housing 46.

[0039] HP turbine 34 and LP turbine 36 each comprise multiple turbine stages 64 and 66, respectively, in which a set of turbine blades 68 and 70 rotate relative to a corresponding set of static turbine blades 72 and 74 (also referred to as nozzles) to extract energy from the fluid flow passing through the stage. In a single turbine stage 64 and 66, the multiple turbine blades 68 and 70 are arranged in a ring and extend radially outward from the blade platform relative to the engine centerline 12 to the blade tips, while the corresponding static turbine blades 72 and 74 are positioned upstream of and adjacent to the rotating turbine blades 68 and 70. It is worth noting that... Figure 1 The number of blades, impellers, and turbine stages shown is selected for illustrative purposes only; other numbers are also possible.

[0040] Turbine blades 68 and 70 for the first stage of the turbine are mounted to disk 71, which is mounted to a corresponding one of HP spool 48 and LP spool 50, with each stage having a dedicated disk 71. Blades 72 and 74 for the first stage of the compressor can be mounted circumferentially to the core housing 46.

[0041] Complementing the rotor section, the stationary parts of the engine 10 (e.g., the static blades 60, 62, 72, 74 in the compressor section 22 and the turbine section 32) are also referred to individually or collectively as the stator 63. Therefore, the stator 63 refers to the combination of non-rotating elements throughout the engine 10.

[0042] In operation, the airflow leaving fan section 18 is split, with a portion directed to LP compressor 24. LP compressor 24 then supplies pressurized air 76 to HP compressor 26, which further pressurizes the air. The pressurized air 76 from HP compressor 26 mixes with fuel in combustor 30 and ignites, generating combustion gases. HP turbine 34 extracts some work from these gases, driving HP compressor 26. The combustion gases are discharged to LP turbine 36, which extracts additional work to drive LP compressor 24, and the exhaust is ultimately discharged from engine 10 via exhaust section 38. The drive of LP turbine 36 drives LP spool 50 to rotate fan 20 and LP compressor 24.

[0043] A portion of the pressurized air 76 is drawn from the compressor section 22 as bleed air 77. Bleed air 77 can be supplied to engine components requiring cooling. The temperature of the pressurized air 76 entering and leaving the combustor 30 increases significantly. Thus, the cooling provided by bleed air 77 is supplied to downstream turbine components (e.g., blades 68) subjected to high-temperature environments.

[0044] The remaining portion of the airflow leaving the fan section, bypass airflow 78, bypasses the LP compressor 24 and engine core 44, and exits the engine 10 through the stationary blade row, and more specifically, through the outlet guide vane assembly 80, which includes multiple airfoil guide vanes 82, located at the fan exhaust side 84. More specifically, a circumferential row of radially extending airfoil guide vanes 82 is used adjacent to the fan section 18 to exert some directional control on the bypass airflow 78.

[0045] Some of the air supplied by fan 20 can bypass engine core 44 and be used to cool parts of engine 10, especially hot parts, and / or to cool other aspects of the aircraft or power other aspects of the aircraft. In the case of a turbine engine, the hot parts of the engine are typically downstream of combustor 30, especially turbine section 32, with HP turbine 34 being the hottest part because it is directly downstream of combustion section 28. Other sources of cooling fluid may be, but are not limited to, fluid discharged from LP compressor 24 or HP compressor 26.

[0046] Figure 2 It can be used in engine 10 ( Figure 1 An exemplary airfoil assembly 90 is shown in perspective. In the example shown, the airfoil assembly 90 includes a dovetail 95 and an airfoil 100 in the form of a blade. The airfoil assembly 90 may correspond to blades 56, 58, 68, 70 or impeller blades 60, 62, 72, 74. Figure 1 One of them. In some examples, the airfoil assembly 90 includes an end wall or inner or outer band in place of the dovetail tenon 95.

[0047] The airfoil assembly 90 is formed from a material suitable for the operating environment. The materials used to form the airfoil assembly 90 can include, but are not limited to, steel, refractory metals such as titanium, or superalloys based on nickel, cobalt, or iron, as well as ceramic matrix composites. In a non-limiting example, the airfoil assembly can be formed by a variety of methods, including additive manufacturing, casting, electroforming, or direct metal laser melting. As used herein, an "additive manufacturing" part refers to a part formed by an additive manufacturing (AM) process, where the part is built layer by layer through the continuous deposition of material. AM is an appropriate name to describe the technique of building 3D objects by adding layer after layer of material, whether the material is plastic, ceramic, or metal. AM technology can utilize computers, 3D modeling software (computer-aided design or CAD), machinery, and layered materials. Once a CAD sketch is generated, an AM device can read data from the CAD file and place or add successive layers of liquid, powder, sheet, or other materials in a layer-by-layer manner to create a 3D object. It should be understood that the term "additive manufacturing" encompasses many technologies, including 3D printing, rapid prototyping (RP), direct digital manufacturing (DDM), layered manufacturing, and subsets of additive manufacturing. Non-limiting examples of additive manufacturing that can be used to form additively manufactured parts include powder bed melting, vat photopolymerization, binder jetting, material extrusion, directional energy deposition, material jetting, or sheet lamination. It is also conceivable that the process used may include printing a negative film of a part using refractory metals, ceramics, or printable plastics, and then using that film to cast the part.

[0048] The dovetail tenon 95 includes a platform 96 on which the airfoil 100 is mounted. When multiple airfoils are arranged circumferentially in a side-by-side relationship, the platform 96 helps to radially accommodate the mainstream airflow of the turbine engine and forms the radial inner wall of the annular space through which the airflow passes. The dovetail tenon 95 is configured to be mounted on the turbine rotor disk 71 of the engine 10. Figure 1 In the example shown, the dovetail joint 95 is oriented to be mounted to the turbine rotor disk 71 in the axial direction (A). Figure 1 The dovetail tenon 95 also includes at least one inlet channel 98 extending through the dovetail tenon 95 to provide fluid communication with the interior of the airfoil 100.

[0049] The airfoil 100 includes an outer wall 101 that extends between a tip 102 and a root 104 to define the wingspan direction (S). The outer wall 101 also extends between a leading edge 106 forming an upstream edge and a trailing edge 108 forming a downstream edge, as shown.

[0050] The outer wall 101 of the airfoil 100 includes a first side and a second side, the first side being shown as a concave pressure side 112 and the second side being shown as a convex suction side 114. The outer periphery of the airfoil 100 is defined by the outer wall 101.

[0051] The outer wall 101 may face the hot gas flow. The interior 116 of the airfoil 100 includes at least one cooling supply conduit 120, as shown by the dashed line. The at least one cooling supply conduit 120 is fluidly connected to the inlet passage 98 of the dovetail 95. Cooling fluid is supplied from the at least one cooling supply conduit 120. At least one cooling hole 125 is located in the outer wall 101. The at least one cooling hole 125 may be provided at any location along the outer wall 101 (including along the leading edge 106, trailing edge 108, pressure side 112, or suction side 114).

[0052] Figure 3 It shows along Figure 2 The cross-section of the airfoil 100 is taken by line III-III. The outer wall 101 (including the pressure side 112 and the suction side 114) together define the airfoil cross-sectional shape or geometric profile of the airfoil 100.

[0053] As shown, the central arc 110 extends between the leading edge 106 and the trailing edge 108. Additionally, the outer wall 101 is shown as generally solid, although this is not required. It should be understood that the outer wall 101 may include internal chambers, air chambers, gaps, in-wall cooling, etc. The central arc 110 defines an arc length 111 between the leading edge 106 and the trailing edge 108.

[0054] The outer wall 101 also defines an outer surface 103 and an inner surface 105. As shown, the inner surface 105 faces the interior 116 and is also fluidly connected to the interior 116.

[0055] At least one cooling hole 125 extends through the outer wall 101 and is fluidly connected to at least one cooling supply conduit 120. In the illustrated example, two cooling holes 125 are shown, although one or more cooling holes 125 may be provided, and the cooling holes 125 may be located on any part of the outer wall 101.

[0056] Additionally, in the illustrated example, at least one cooling supply conduit 120 forms three internal chambers within the outer wall 101. In some examples, at least one cooling supply conduit 120 is a continuous conduit, such that the three internal chambers form part of a common conduit. In some examples, at least one cooling supply conduit 120 includes a plurality of separate conduits forming chambers, as shown. Any number of chambers can be provided.

[0057] Turn Figure 4 , Figure 3Section 130 of the airfoil 100 is enlarged to show further details. Section 130 includes the near-tail edge region of the airfoil 100. In some examples, section 130 includes 0-20% of the mid-curve 110.

[0058] The trailing edge 108 is shown in a generally circular geometric profile, although this is not required. The trailing edge 108 may also include an elliptical profile, or a combination of circular and elliptical profiles, or a conical profile, or a flat or truncated profile, etc.

[0059] The central arc 110 is shown as a dashed line. Although the central arc 110 is shown as a straight line in the illustrated example, it should be understood that the central arc 110 may be curved. In some examples, the central arc 110 is locally straight or linear in the near-tailed edge region within portion 130. Additionally, although the outer wall 101 in portion 130 is shown as solid, this is not required, and the outer wall 101 may include, for example... Figure 3 The hollow region, air chamber, etc. mentioned in the text.

[0060] A first thickness T1 is defined at a first position L1 between the pressure side 112 and the suction side 114, as shown in the figure. The first position L1 is along the central arc 110. The first thickness T1 refers to the thickness measured between the outer surfaces 103 at the first position L1 in a direction perpendicular to the central arc 110. Additionally, the first position L1 is spaced from the trailing edge 108 along the central arc 110 to define a first distance (d1). In some examples, the first distance d1 is at an arc length 111 (… Figure 3 The first thickness T1 is between 0-10% of the first distance d1. As shown in the figure, the first thickness T1 is between 1-5 mm and the first distance d1 is between 4-8 mm.

[0061] The second thickness T2 is defined at the second position L2 between the pressure side 112 and the suction side 114, as shown in the figure. The second position L2 is along the mid-arc line 110. In the example shown, the second position L2 is positioned further away from the tail edge than the first position L1. The second thickness T2 refers to the thickness measured between the outer surfaces 103 at the second position L2 in a direction perpendicular to the mid-arc line 110. The second thickness T2 is in the range of 2-10 mm.

[0062] Additionally, distance (d) T The distance d is limited to the area between the first position L1 and the second position L2. T It is defined along the central arc 110. In some examples where the central arc 110 bends between the first position L1 and the second position L2, the distance d TIt can also be curved. It should be understood that in some examples, the first position L1 and the second position L2 can be offset from each other. For example, the pressure side 112 and the suction side 114 can be asymmetrical about each other with respect to the central arc 110, such that the central arc 110 is curved between the first position L1 and the second position L2.

[0063] Wall thickness (t) W The wall thickness t is defined between the outer surface 103 and the inner surface 105, as shown in the figure. In the non-limiting example shown, the wall thickness t W It is 2mm. Additionally, the air chamber distance (d) P The air chamber is defined between the inner surface 105 and the first position L1. In the non-limiting example shown, the air chamber distance d P It is 5mm. In some examples, the air chamber distance d P It is defined as the shortest distance along the mid-arc 110 between the inner surface 105 and the first position L1. In some examples, the air chamber distance d P It is the straight-line distance, or approximately the straight-line distance, for example, within 5-20% of the straight-line distance. In some examples, the air chamber distance d P It can also be a curved or non-linear distance along the middle arc 110.

[0064] An exemplary cooling hole 125 is shown in the outer wall 101. The cooling hole 125 is located in relation to... Figure 4 In the same plane as the cross-section shown. Cooling hole 125 extends between inlet 126 on inner surface 105 and outlet 127 on outer surface 103. Inlet 126 is fluidly connected to interior 116 of airfoil 100. Although cooling hole 125 is shown as a generally straight hole extending through outer wall 101, it should be understood that cooling hole 125 may have other geometries, including slits, diffusion sections, metering sections, curved sections, etc. In addition, outlet 127 of cooling hole 125 is shown on pressure side 112, and it is contemplated that cooling hole 125 may also be positioned on suction side 114.

[0065] The outlet 127 defines the leading edge 128 and the trailing edge 129, with the trailing edge 129 positioned closer to the trailing edge 108. Cooling hole width (w) H The cooling hole width w is defined at outlet 127, between leading edge 128 and trailing edge 129. H It is defined along the middle arc 110, as shown in the figure.

[0066] The cooling hole 125 also defines a third position (L3) along the central arc 110. In some examples, the third position L3 is defined by the trailing edge 129, although this is not required. In a non-limiting example, the third position L3 may also be defined by the leading edge 128, or relative to the inlet 126, or relative to the centerline defined along the cooling hole 125. Additionally, the cooling hole 125 defines a cooling hole distance (d) along the central arc 110 between the third position L3 and the first position L1. H In some examples, it can be envisioned that the airfoil 100 includes a rapid thickness expansion forward from the trailing edge 108, such as a first thickness T1 and a second thickness T2. This thickness expansion can provide cooling features, such as cooling holes 125, positioned closer to the trailing edge 108 than in conventional airfoils. In some non-limiting examples, the distance between the cooling holes and d... H Within the range of 0.01-0.1 cm, this arrangement directly provides increased cooling efficiency near the trailing edge 108.

[0067] Now for reference Figure 5 It shows that it can be used Figure 1 Another airfoil 200 of the engine 10. Therefore, similar parts will be described with similar numbers increased by 100. It should be understood that, unless otherwise stated, the description of similar parts of airfoil 100 applies to airfoil 200.

[0068] The enlarged portion 230 of the airfoil 200 is shown, similar to portion 130. Figure 3 Part 230 includes the near-tailing edge region of airfoil 200. Airfoil 200 includes an outer wall 201 defining interior 216 and having an outer surface 203, an inner surface 205, a trailing edge 208, a pressure side 212, and a suction side 214. Cooling ducts 220 extend within interior 216. A mid-curve 210 is present at the leading edge (…). Figure 4 (Not shown in the figure) extends between the outer wall 201 and the trailing edge 208. In some examples, portion 230 includes 0-20% of the middle arc 210. Additionally, a cooling hole 225 is provided in the outer wall 201, extending between an inlet 226 and an outlet 227, as shown. Inlet 226 is fluidly connected to cooling conduit 220.

[0069] The first thickness T1 is defined at the first position L1 between the pressure side 212 and the suction side 214, as shown in the figure. The first position L1 is spaced from the trailing edge 208 along the mid-arc line 210 to define a first distance d1. One difference compared to airfoil 100 is that airfoil 200 can have a larger first thickness T1. As shown in the figure, the first thickness T1 is in the range of 5-10 mm.

[0070] The second thickness T2 is defined at the second position L2 between the pressure side 212 and the suction side 214, as shown in the figure. Another difference compared to airfoil 100 is that airfoil 200 can have a larger second thickness T2. As shown in the figure, the second thickness is in the range of 10-14 mm.

[0071] Additionally, the distance d T The wall thickness t is defined along the central arc 210 between the first position L1 and the second position L2. W It is also confined between the outer surface 203 and the inner surface 205, as shown in the figure. Another difference compared to airfoil 100 is that airfoil 200 can have a greater wall thickness t. W As shown in the figure, the wall thickness is 4mm.

[0072] air chamber distance d P It is defined between the inner surface 205 and the first position L1. Another difference compared to airfoil 100 is that airfoil 200 can have a smaller air chamber distance d. P As shown in the figure, the distance d between the air chambers P It is 1mm.

[0073] The outlet 227 can define the leading edge 228 and the trailing edge 229. Cooling hole width w H It is limited to outlet 227, between leading edge 228 and trailing edge 229. Cooling hole width w H It is defined along the central arc 210, as shown in the figure. The cooling hole 225 is also defined along the central arc 210 between the trailing edge 229 and the second position L2, with a distance d between the cooling holes. H .

[0074] General reference Figure 1-5 It is understood that, among many others, airfoils 100 and 200 are merely two illustrative examples of design parameters that provide a desired balance between aerodynamic performance and cooling performance. In a non-limiting example, airfoil 100 has higher aerodynamic performance than airfoil 200, while airfoil 200 has higher cooling performance than airfoil 100, although this need not be the case.

[0075] As mentioned earlier, finding a viable airfoil solution involves first designing the airfoil's aerodynamic performance and then adjusting the design based on cooling requirements, which can lead to unfavorable system performance as assessed through SFC (Self-Fueling Function) evaluation. This is a labor-intensive and time-consuming process because it is iterative and requires finding an airfoil that presents the desired aerodynamic and cooling performance. The goal is to improve SFC, but many trade-offs need to be considered. A limited or narrow range of embodiments are desired, confined to the engine architecture and meeting design requirements.

[0076] Table 1 below shows some non-limiting examples of airfoil designs that provide feasible solutions to balancing the aerodynamic and cooling performance of a given engine architecture. A pressure loss coefficient Yp related to aerodynamic performance is also provided for each airfoil example.

[0077] Table 1

[0078]

[0079]

[0080] Unless otherwise stated, all units given in this disclosure are in centimeters (cm) unless otherwise specified as dimensionless.

[0081] The inventors discovered that for airfoil designs that produce improved results, the first thickness T1, the second thickness T2, and the distance d... T There exists a relationship between these factors. This is referred to in this paper as the thickness parameter (τ). This relationship was unexpectedly discovered during the engine design process (i.e., the design of airfoils within the engine and the evaluation of the resulting airfoils' impact on engine fuel consumption or efficiency). Furthermore, by utilizing this relationship, the inventors discovered that the number of suitable or feasible airfoil examples for placement in a turbine engine that meet design requirements can be significantly reduced, facilitating a faster downward selection of airfoil designs to consider during engine development. The discovered relationship also avoids or prevents later redesigns while providing airfoil designs that integrate high performance and cooling efficiency. The discovered relationship enables narrowing down the selection during the design process, thereby saving time, reducing material costs, and improving airfoil design for various engine architectures.

[0082] It was found that the thickness parameter τ balances the desirable aspects of aerodynamic performance and cooling performance. The thickness parameter τ is expressed as follows: (3) Where R1 is composed of the first thickness T1 and the distance d T Limited ratio:

[0083] (4) R1=d T / T1

[0084] When the first ratio R1 is between 2.5 and 6.5 (2.5≤R1≤6.5), the thickness parameter τ is between 0.275 and 0.45 (0.275≤τ≤0.45).

[0085] It should be understood that the thickness parameter τ and the distance d, as described in expressions 3 and 4, T The relationship between the first thickness T1 and the second thickness T2 can also be expressed in the following simplified form:

[0086] (5) τd T +T1=T2

[0087] In this way, the relationships shown in expressions 3-4 include dimensionless ratios, while the relationship shown in expression 5 includes dimensional terms. Regardless of the form of the expression, the thickness parameter τ itself is a dimensionless parameter.

[0088] The thickness parameter τ, the first ratio R1, and the ratio of the second thickness T2 to the first thickness T1 correspond to examples in Table 1. Some exemplary values ​​that satisfy the relationships shown in expressions 3-4 are provided in Table 2 below:

[0089] Table 2

[0090] Example: 1 2 3 4 5 6 7 8 9 τ 0.275 0.341 0.285 .290 0.450 0.275 0.375 0.275 0.325 <![CDATA[R1]]> 2.5 4.1 5.2 5.9 6.5 3.0 3.0 6.0 6.0 <![CDATA[T2 / T1]]> 1.7 2.4 2.5 2.7 3.9 1.8 2.2 2.7 2.9

[0091] Using the relationship described in expressions 3-4 (or alternatively, expression 5), the inventors were able to achieve airfoils with better cooling performance and acceptable aerodynamic performance.

[0092] Further, the first position L1 is determined based on the rate of change of the surface curvature of the outer surface 103 along the mid-arc 110. In some examples, a first position L1 is determined where the rate of change of the surface curvature exceeds a predetermined threshold. In some examples, the pressure side 112 and the suction side 114 are each treated as an unwrapped curve Rθ between the leading edge 106 and the trailing edge 108, as expressed by the following expression:

[0093] (6) Rθ=f(s)

[0094] Where s represents the surface distance along the pressure side 112 or suction side 114 measured from the leading edge 106.

[0095] The curvature (k) of the outer surface 103 is expressed as follows:

[0096] (7)

[0097] Where f(s)″ is relative to Figure 5 The second derivative of the horizontal axis 210 in the figure. The rate of change of curvature (k′) is expressed as follows:

[0098] (8)

[0099] The curvature k has a unit that is the reciprocal of length, such as cm. -1 The rate of change of surface curvature, k′, has the reciprocal unit of the square of the length, for example, cm. -2 .

[0100] Referring to expressions 6-8, the curvature k of the outer surface 103 transitions from the suction side curve 114 to the trailing edge curve 108, or from the pressure side curve 112 to the trailing edge curve 108. This transition is determined when the rate of change of surface curvature k′ exceeds a predetermined threshold. In some examples, the first position L1 is determined along the mid-arc 110, where the absolute value of the rate of change of surface curvature k′ on the outer surface 103 exceeds 960 cm. -2 (|k′|>960cm -2 In this way, the first position L1 defines the transition between the pressure side 112 and the trailing edge 108 along the middle arc 110, or at the transition between the suction side 114 and the trailing edge 108.

[0101] Further, consider the distance d between the cooling holes. H The second ratio (R2) is defined by the first thickness T1:

[0102] (9) R2=d H / T1

[0103] The second ratio R2 is between 0.3 and 3.0 (0.3≤R2≤3.0).

[0104] Further, consider the width w of the cooling hole. H (As described above, defined along the central arc 110 between the leading edge 128 and the trailing edge 129) and the first thickness T1 defines the third ratio (R3):

[0105] (10) R3=w H / T1

[0106] The third ratio R3 is between 0.3 and 3.0 (0.3≤R3≤3.0).

[0107] Turn now Figure 6 Figure 300 shows a graph where the thickness parameter τ is plotted relative to a first ratio R1, as shown. The first space 310, covered by expressions 3-4 (or alternatively, by expression 5), is shown as a first shaded area. The second space 320 is shown as a second shaded area. The second space 320 is defined by a trapezoidal region defined by embodiments having values ​​of (R1, τ) between (3.0, 0.275), (3.0, 0.375), (6.0, 0.275), and (6.0, 0.325). The second space 320 represents a subset or narrowed boundary within the first space 310.

[0108] The airfoil examples 1-9 described in Table 1-2 above are shown in graph 300. Points 301-309, from the first to the ninth, correspond to examples 1-9 respectively.

[0109] As shown in the figure, for the thickness parameter τ and the first ratio R1, the first point 301 and the fifth point 305 are located at the minimum and maximum boundaries in the first space 310, respectively. The second point 302, the third point 303, and the fourth point 304 are located in the second parameter space 320. The sixth to ninth points 306-309 are located at the boundaries of the second parameter space 320.

[0110] The pressure loss coefficient Yp (Table 1) is also indicated for points 301-309. The pressure loss coefficient Yp of the embodiments falling within space 310 can be equal to, less than, or greater than the Yp of the embodiments falling within the second space 320. As shown, embodiments 302 and 303 within the second space 320 have smaller pressure loss coefficients Yp compared to embodiment 301 within the first space 310 and outside space 320. Furthermore, embodiment 306 within the second space 320 has a larger Yp compared to embodiment 305 outside space 320 and within space 310.

[0111] The aspects disclosed herein offer several benefits. The thickness parameters and parameter space described herein provide a geometric envelope that generates rapid thickness expansion forward from the trailing edge, enabling the encapsulation or formation of cooling features closer to the trailing edge than conventional airfoils. Such a location of the cooling feature provides a lower required cooling flow, which improves engine SFC (Short-range fuel conversion). The thickness parameters and parameter space described herein further provide airfoils with higher wedge angles, providing localized near-trail-edge thickness increases and available space for cooling features, while maintaining or improving the aerodynamic performance of the airfoil. Further optimized SFC can be achieved by jointly optimizing the trailing edge shape to balance aerodynamic losses and cooling flow requirements.

[0112] For airfoils with a large amount of chargeable flow or turbine cooling flow, such as the first or second stage blades in some examples of high-pressure turbines, both cooling flow performance and aerodynamic performance have a significant impact on engine SFC or efficiency. An aspect of this disclosure provides an airfoil that exhibits improved cooling performance by increasing local trailing edge thickness or thickness expansion, and improved aerodynamic performance captured in the parameter space described herein.

[0113] This disclosure also provides for narrowing down multiple factors to a smaller design space for engine components being manufactured, the location of the engine component within the engine, the materials used, or any other design constraints. Narrowing these multiple factors down to a range of possibilities can save time, money, and resources, and result in airfoils with superior performance.

[0114] This written description uses examples to disclose this disclosure, including best practices, and also enables any person skilled in the art to practice this disclosure, including making and using any device or system and methods of making any combination. The patent scope of this disclosure is defined by the claims, but may include other examples that would occur to a person skilled in the art. Such other examples are intended to fall within the scope of the claims if they include structural elements that are not indistinguishable from the literal language of the claims, or if they include equivalent structural elements that are not substantially different from the literal language of the claims.

[0115] Further aspects of this disclosure are provided by the subject matter of the following clauses:

[0116] A turbine engine includes: a compressor section, a combustion section, and a turbine section arranged in a tandem flow configuration; and an airfoil comprising: an outer wall having an outer surface and defining an interior, the outer wall defining a pressure side and a suction side, and extending between a leading edge and a trailing edge to define a mid-curve extending between the leading edge and the trailing edge; a first thickness (T1) defined at a first position (L1) along the mid-curve in a first direction perpendicular to the mid-curve of the airfoil between the pressure side and the suction side; a second thickness (T2) defined at a second position (L2) along the mid-curve in a second direction perpendicular to the mid-curve between the pressure side and the suction side, wherein the second position L2 is positioned further away from the trailing edge than the first position L1; and a distance (d) T The distance (d) T The first thickness T1 and the distance d are defined along the middle arc line between the first position and the second position. T The first ratio (R1) is defined as R1 = d. T / T1, wherein the first thickness T1, the second thickness T2, and the first ratio R1 are related to each other. The thickness parameter (τ) is defined such that when the first ratio R1 is between 2.5 and 6.5 (2.5≤R1≤6.5), the thickness parameter τ is between 0.275 and 0.45 (0.275≤τ≤0.45).

[0117] The turbine engine according to any of the foregoing clauses, wherein the outer wall defines a cooling hole that extends between an inlet fluidly connected to the interior and an outlet on the outer surface.

[0118] The turbine engine according to any of the foregoing clauses, wherein the outlet is on the suction side.

[0119] According to any of the foregoing clauses of the turbine engine, wherein the cooling hole defines a distance (d) between the trailing edge of the cooling hole and the second position L2 along the central arc. H ), wherein the distance between the cooling holes and d H The first thickness T1 defines the second ratio (R2) as R2 = d. H / T1.

[0120] According to any of the foregoing clauses, the turbine engine wherein the cooling hole defines a cooling hole width (w) at the outlet. H ), wherein the width w of the cooling hole H The third ratio (R3) is defined by the first thickness T1 as R3 = w H / T1.

[0121] According to any of the foregoing clauses, the second ratio R2 is between 0.3 and 3.0 (0.3≤R2≤3.0).

[0122] According to any of the preceding clauses, the third ratio R3 is between 0.3 and 3.0 (0.3≤R3≤3.0).

[0123] According to any of the preceding clauses, the turbine engine, wherein the pressure side and the suction side comprise the unfolded curve (Rθ) between the leading edge and the trailing edge, Rθ = f(s).

[0124] The turbine engine according to any of the foregoing clauses further includes the curvature (k) of the outer surface, wherein

[0125] According to any of the foregoing clauses, the rate of change of curvature (k′) is defined as follows:

[0126] According to any of the preceding clauses, in a turbine engine, the first position L1 is determined along the mid-arc line, and at the first position L1, the absolute value of the rate of change of curvature k′ is greater than 960 cm. -2 (|k′|>960cm -2 ).

[0127] According to any of the preceding clauses, the turbine engine wherein the first position L1 is determined along the central arc, and at the first position L1, the rate of change of the curvature of the outer surface exceeds a predetermined threshold.

[0128] According to any of the preceding clauses, the first ratio R1 and the thickness parameter τ are in a space defined by the following: (R1 = 3.0, τ = 0.275); (R1 = 3.0, τ = 0.375); (R1 = 6.0, τ = 0.275); and (R1 = 6.0, τ = 0.325).

[0129] According to any of the preceding clauses of the turbine engine, the first position L1 is spaced apart from the trailing edge along the mid-arc line to define a first distance.

[0130] According to any of the preceding clauses, the turbine engine wherein the intermediate arc defines the arc length between the leading edge and the trailing edge, and wherein the first distance is between 0 and 10% of the arc length.

[0131] The turbine engine according to any of the foregoing clauses, wherein the pressure side and the suction side are symmetrical about the mid-curve.

[0132] An airfoil for a turbine engine, the airfoil comprising: an outer wall having an outer surface and defining an interior, the outer wall defining a pressure side and a suction side, and extending between a leading edge and a trailing edge to define a mid-curve extending between the leading edge and the trailing edge; a first thickness (T1), the first thickness (T1) being defined at a first position (L1) along the mid-curve in a first direction perpendicular to the mid-curve of the airfoil between the pressure side and the suction side; and a second thickness (T2), the second thickness (T2) being defined at a second position (L1) along the mid-curve in a first direction perpendicular to the mid-curve of the airfoil. A second position (L2) is defined between the pressure side and the suction side in a second direction perpendicular to the middle arc, wherein the second position (L2) is positioned further away from the trailing edge than the first position (L1); and a distance (d) is defined along the middle arc between the first position and the second position, wherein the first thickness (T1) and the distance (d) define a first ratio (R1) as R1 = d / T1, wherein the first thickness (T1), the second thickness (T2), and the first ratio (R1) are related to each other. The thickness parameter (τ) is defined such that when the first ratio (R1) is between 2.5 and 6.5 (2.5 ≤ R1 ≤ 6.5), the thickness parameter τ is between 0.275 and 0.45 (0.275 ≤ τ ≤ 0.45).

[0133] The airfoil according to any of the foregoing clauses further includes cooling holes in the outer wall, the cooling holes extending between an inlet fluidly connected to the interior and an outlet on the outer surface.

[0134] According to any of the foregoing clauses, the cooling hole defines a distance (d) between the trailing edge of the cooling hole and the second position L2 along the mid-arc line. H ), wherein the distance between the cooling holes and d H The first thickness T1 defines the second ratio (R2) as R2 = d. H / T1.

[0135] According to any of the foregoing clauses, the cooling hole defines the cooling hole width (w) at the outlet. H ), wherein the width w of the cooling hole H The third ratio (R3) is defined by the first thickness T1 as R3 = w H / T1.

[0136] According to any of the foregoing clauses, the second ratio R2 is between 0.3 and 3.0 (0.3≤R2≤3.0).

[0137] According to any of the foregoing clauses, the third ratio R3 is between 0.3 and 3.0 (0.3≤R3≤3.0).

[0138] According to any of the foregoing clauses, the first position (L1) is determined along the middle arc line, and at the first position (L1), the rate of change of the surface curvature of the outer surface exceeds a predetermined threshold.

[0139] According to any of the foregoing clauses, the first ratio R1 and the thickness parameter τ are in a space defined by the following: (R1 = 3.0, τ = 0.275); (R1 = 3.0, τ = 0.375); (R1 = 6.0, τ = 0.275); and (R1 = 6.0, τ = 0.325).

[0140] According to any of the foregoing clauses, the pressure side and the suction side are symmetrical about the mid-curve.

Claims

1. A turbine engine, characterized in that, include: The compressor section, combustion section, and turbine section are arranged in a series flow pattern, and Airfoil, the airfoil comprising: An outer wall having an outer surface and defining an interior, the outer wall defining a pressure side and a suction side, and extending between a leading edge and a trailing edge to define a mid-arc extending between the leading edge and the trailing edge; A first thickness T1 is defined at a first position L1 along the middle arc line between the pressure side and the suction side in a first direction perpendicular to the middle arc line of the airfoil. A second thickness T2, defined at a second position L2 along the central arc line in a second direction perpendicular to the central arc line between the pressure side and the suction side, wherein the second position L2 is positioned further away from the trailing edge than the first position L1; and Distance d T The distance d T It is defined along the central arc line between the first position L1 and the second position L2; Wherein the first thickness T1 and the distance d T The first ratio R1 is limited to ; The first thickness T1, the second thickness T2, and the first ratio R1 are related to each other. , to limit the thickness parameter τ; Among them when hour, .

2. The turbine engine according to claim 1, characterized in that, in, The outer wall defines a cooling hole that extends between an inlet fluidly connected to the interior and an outlet on the outer surface.

3. The turbine engine according to claim 2, characterized in that, in, The outlet is located on the suction side.

4. The turbine engine according to claim 2, characterized in that, in, The cooling hole defines a distance d along the mid-arc line between the rear edge of the cooling hole and the second position L2. H The distance between the cooling holes and d H The first thickness T1 limits the second ratio R2 to .

5. The turbine engine according to claim 4, characterized in that, in, The cooling hole defines the width w at the outlet. H The width w of the cooling hole H The third ratio R3 is defined by the first thickness T1. .

6. The turbine engine according to claim 5, characterized in that, in, 。 7. The turbine engine according to claim 6, characterized in that, in, 。 8. The turbine engine according to claim 1, characterized in that, in, The first position L1 is determined along the middle arc line, and at the first position L1, the rate of change of the curvature of the outer surface exceeds a predetermined threshold.

9. The turbine engine according to claim 1, characterized in that, in, The first ratio R1 and the thickness parameter τ lie within a space defined by the following: （ ）； （ ）； ( );as well as （ ）。 10. The turbine engine according to claim 1, characterized in that, in, The pressure side and the suction side are symmetrical about the middle arc line.

11. The turbine engine according to claim 1, characterized in that, in, The first position L1 is spaced apart from the tail edge along the middle arc to define a first distance.

12. The turbine engine according to claim 11, characterized in that, in, The middle arc defines the arc length between the leading edge and the trailing edge, and wherein the first distance is between 0 and 10% of the arc length.

13. An airfoil for a turbine engine, characterized in that, The airfoil includes: An outer wall having an outer surface and defining an interior, the outer wall defining a pressure side and a suction side, and extending between a leading edge and a trailing edge to define a mid-arc extending between the leading edge and the trailing edge; A first thickness T1 is defined at a first position L1 along the middle arc line between the pressure side and the suction side in a first direction perpendicular to the middle arc line of the airfoil. A second thickness T2, defined at a second position L2 along the central arc line in a second direction perpendicular to the central arc line between the pressure side and the suction side, wherein the second position L2 is positioned further away from the trailing edge than the first position L1; and The distance d is defined along the middle arc between the first position L1 and the second position L2; Wherein the first thickness T1 and the distance d define the first ratio R1 as: ; The first thickness T1, the second thickness T2, and the first ratio R1 are related to each other. , to limit the thickness parameter τ; Among them when hour, .

14. The airfoil according to claim 13, characterized in that, in, The outer wall defines a cooling hole that extends between an inlet fluidly connected to the interior and an outlet on the outer surface.

15. The airfoil according to claim 14, characterized in that, in, The cooling hole defines a distance d along the mid-arc line between the rear edge of the cooling hole and the second position L2. H The distance between the cooling holes and d H The first thickness T1 limits the second ratio R2 to .

16. The airfoil according to claim 15, characterized in that, in, The cooling hole defines the width w at the outlet. H The width w of the cooling hole H The third ratio R3 is defined by the first thickness T1. .

17. The airfoil according to claim 16, characterized in that, in, 。 18. The airfoil according to claim 17, characterized in that, in, 。 19. The airfoil according to claim 13, characterized in that, in, The first position L1 is determined along the middle arc line, and at the first position L1, the rate of change of the curvature of the outer surface exceeds a predetermined threshold.

20. The airfoil according to claim 13, characterized in that, in, The first ratio R1 and the thickness parameter τ lie within a space defined by the following: （ ）； （ ）； ( );as well as （ ）。

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