Airfoil profile for a blade in a turbine engine

Abstract

An apparatus and method for an engine component having a working airflow split into a cooling airflow and a combustion airflow, the engine component comprising a plurality of circumferentially spaced apart airfoils rotatable about a centreline defining an axial direction, each airfoil comprising an outer wall, the outer wall bounding an interior and defining a pressure side and a suction side, the pressure side and suction side extending between a leading edge and a trailing edge to define a chordwise direction, and along a spanwise direction between a root and a tip to define a span length; wherein each airfoil has a non-ducted turning angle to stagger angle ratio defined as a percentage of the span length.

Description

[0001] Cross-references to related applications

[0002] This application claims priority to U.S. Patent Application No. 17 / 537,818, filed November 30, 2021, the entire contents of which are incorporated herein by reference. Technical Field

[0003] This disclosure generally relates to engine components with airfoil profiles, and more specifically, to blades with varying directional changes along their span length. Background Technology

[0004] A turbine engine (especially a gas or combustion turbine engine) is a rotating engine that extracts energy from a stream of combustible gas that passes through the engine and flows over multiple airfoils, including stationary impeller blades and rotating turbine blades.

[0005] Turbine blades can be subjected to flow regimes that make them prone to forming vortex structures near the hub and tip sections of the airfoil. Due to the physical clearance required to allow relative movement between the rotating blade tip and the stationary shell structure, rotating blades will form additional vortex structures due to flow leakage above the blade tip. This vortex structure limits the amount of work delivered to the compressor, and thus limits the engine's total thrust or power capability. Furthermore, controlling the upper tip leakage flow and the associated tip vortex-driven losses is crucial for unshielded blades. To control the formation and growth of these vortex structures, the local blade load distribution is uniquely adjusted to minimize the losses caused by these structures. Attached Figure Description

[0006] The specification with reference to the accompanying drawings sets forth a complete and enabling disclosure for those skilled in the art, including its best mode, wherein:

[0007] Figure 1 This is a schematic cross-sectional view of a gas turbine engine used in aircraft.

[0008] Figure 2 It comes from Figure 1 A top view of a pair of blades of a gas turbine engine.

[0009] Figure 3 Is with Figure 2 The same top view shows the unguided steering angle and the staggered angle.

[0010] Figure 4 This is a perspective view of an exemplary blade with varying unguided turning angles and staggered angles between the base and tip of the blade.

[0011] Figure 5A It is a top view of the airfoil profile of the blade at the tip.

[0012] Figure 5B It is a top view of the shape of the airfoil at the root of the tip.

[0013] Figure 6 This is a graph showing the non-guided steering angle value versus the crossover angle value.

[0014] Figure 7 It is a graph showing the ratio of the non-guided steering angle value and the staggered angle value to the span length position.

[0015] Figure 8 It is based on the method for forming a turbine disk with multiple blades disclosed herein. Detailed Implementation

[0016] The aspects of this disclosure described herein generally relate to airfoils for gas turbine engines, and more specifically, to shrouded blade airfoils. While focusing on shrouded blades, it should be understood that the disclosure herein is not limited to shrouded blades and can generally relate to airfoils in any part of an engine.

[0017] Many factors determine whether turbine blades will be shielded or not. Tensile load (stress) is one factor. A tip shroud adds weight, which can lead to an unnecessarily increased tensile load. Depending on the material used to make the blades, some materials cannot withstand the increased tensile load in some cases. For highly cooled high-pressure turbine (HPT) airfoils, a low airfoil number is preferred to keep the cooling flow to a minimum. This circumferentially drives a high blade-to-blade distance between the tips, which means a long shroud. This can lead to stress problems in the tip section. Furthermore, the high temperatures in HPTs will require shrouded blades to potentially require internal cooling within the shroud structure, increasing cost and manufacturing time. Additionally, the shroud cooling flow will result in lower cycle efficiency due to higher parasitic emissions from the compressor.

[0018] Unshrouded blades in HPTs require a balance between aerodynamics and structural optimization to ensure engine efficiency and blade life. As described in this article, the airfoil profile is defined by staggered angles and unguided steering (UGT) angles, both of which depend on the robustness of the blade design.

[0019] Managing peak Mach number and suction surface pressure diffusion between peak Mach number and the trailing edge of the airfoil requires a carefully designed airfoil profile. Local airfoil loads can vary between the root or hub region of the airfoil and the tip region radially spaced from the hub region, requiring different airfoil profile strategies at different span locations. As robustness decreases, further distributing the load to the leading edge of the airfoil, with unique considerations in the hub and tip regions, can optimize the management of Mach number and suction surface pressure diffusion. Airfoil load design needs to strike a balance between peak Mach number and suction surface pressure diffusion. High peak Mach numbers lead to impact losses and may require high levels of suction surface pressure diffusion. Abundant diffusion, high diffusivity, and especially the combination of both, result in rapid growth of the suction surface boundary layer. In extreme cases, the boundary layer may detach from the airfoil surface, leading to very high losses. Even when attached, an excessively thick suction surface boundary layer can merge with the trailing edge wake, resulting in losses higher than necessary. This paper will discuss in more detail how a lower staggered angle connected to a higher UGT angle in the hub region controls channel eddy growth, while a higher staggered angle connected to a lower UGT angle in the tip region controls leakage eddy intensity.

[0020] Hub-channel vortex growth is a function of the overall acceleration of the channel, the local airfoil load, and the channel length. For the same axial chord, a lower stagger angle is associated with a smaller true chord of the airfoil. This means a smaller flow length in the channel, resulting in a smaller surface area, which reduces drag and / or profile losses.

[0021] A higher UGT means that the local load (pressure distribution) will concentrate towards the trailing edge of the airfoil. Channel vortices begin due to velocity defects caused by the interaction between the hub boundary layer and the leading edge. Once vortices begin, their growth is enhanced by the pressure difference between the suction and pressure sides of two adjacent airfoils. Channel vortex growth can be minimized by concentrating the high pressure difference at the rear of the airfoil.

[0022] The same drive mechanism listed for the hub also applies to the tip. However, the tip has the added complexity of leakage flow, where the flow from the pressure side migrates above the tip and merges with the flow on the suction side. For the leakage region (driven by the airfoil width at a given axial position), the distribution of leakage flow from the leading edge to the trailing edge will depend on the pressure gradient or pressure variation across the tip (driven by the load) and the drag path, which is a descriptor of the type of path the flow travels. Because the leakage flow generates strong vortices as it re-enters the main flow path along the suction side, the load distribution is altered to minimize the unique growth and intensity properties of tip leakage vortices. Unlike the hub, the load distribution at the tip tends to concentrate lift towards the leading edge, where leakage drag is high due to the high airfoil thickness near the leading edge. Reducing the pressure gradient towards the thinner trailing chord section of the airfoil reduces further accumulation of leakage vortex intensity.

[0023] For illustrative purposes, this disclosure will describe turbines for use in gas turbine engines for aircraft. However, it should be understood that the aspects of this disclosure described herein are not limited thereto and can be generally applied to engines including compressors, power generation gas turbines, and in non-aircraft applications such as other mobile applications and non-mobile industrial, commercial, and residential applications.

[0024] As used in this article, “high” and “low” are descriptors about relative changes in the environment and / or relative values ​​relative to each other.

[0025] All directional references (e.g., radial, axial, proximal, distal, up, down, upward, downward, left, right, lateral, front, rear, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, backward, etc.) are for identification purposes only to aid the reader's understanding of this disclosure and should not be construed as limiting, in particular, with respect to the location, orientation, or use of aspects of this disclosure described herein. Unless otherwise stated, connection references (e.g., attachment, coupling, connection, and joining) are to be interpreted broadly and may include intermediate members between sets of elements and relative movement between elements. Thus, a connection 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. 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 “set” of elements can be any number of elements, including only one element.

[0026] As used throughout the specification and claims, approximate language is applied to modify any quantitative expression that may allow for variation without altering its underlying function. Therefore, values ​​modified by one or more terms such as “about,” “approximately,” “roughly,” and “substantially” are not limited to specified precise values. In at least some cases, approximate language may correspond to the precision of the instrument used to measure the value, or the precision of the method or machine used to construct or manufacture the component and / or system. For example, approximate language may refer to a margin of 1%, 2%, 4%, 5%, 10%, 15%, or 20% of a single value, a range of values, and / or the endpoints of a defined range of values. Scope limitations are combined and interchanged herein and throughout the specification and claims, and such scopes are identified and include all subscopes contained herein unless otherwise indicated by context or language. For example, all scopes disclosed herein include endpoints, and endpoints may be combined independently of each other.

[0027] Figure 1This is a schematic cross-sectional view of a gas turbine engine 10 for an aircraft. The engine 10 has a generally longitudinally extending axis or 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, which includes a fan 20; a compressor section 22, which includes a supercharger or low-pressure (LP) compressor 24 and a high-pressure (HP) compressor 26; a combustion section 28, which includes a combustor 30; a turbine section 32, which includes an HP turbine 34 and an LP turbine 36; and an exhaust section 38.

[0028] Fan section 18 includes a fan housing 40 surrounding fan 20. Fan 20 includes a plurality of fan blades 42 arranged radially about centerline 12. HP compressor 26, combustor 30 and HP turbine 34 form the core 44 of engine 10, which generates combustion gases. Core 44 is surrounded by core housing 46, which can be coupled to fan housing 40.

[0029] An HP shaft or spool 48, coaxially arranged around the 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 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 centerline 12 and connected to multiple rotatable elements that can collectively define the rotor 51.

[0030] 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 blades 60 and 62 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 can be arranged in a ring and can extend radially outward from the blade platform relative to the centerline 12 to the blade tips, while the corresponding static compressor blades 60 and 62 are positioned upstream of and adjacent to the rotating compressor 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.

[0031] Compressor blades 56 and 58 for the compressor stage can be mounted to (or integrated into) disc 61, which is mounted to a corresponding one of HP spool 48 and LP spool 50. Static compressor impellers 60 and 62 for the compressor stage can be mounted circumferentially to the core housing 46.

[0032] HP turbine 34 and LP turbine 36 each comprise multiple turbine stages 64 and 66, 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 may be arranged in a ring and may extend radially outward relative to a centerline 12, 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 only for illustrative purposes, and other numbers are also possible.

[0033] Turbine blades 68 and 70 for the turbine stage can be mounted on disk 71, which is mounted on a corresponding one of HP spool 48 and LP spool 50. Static turbine blades 72 and 74 for the compressor stage can be mounted circumferentially to the core housing 46.

[0034] As a complement to 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 can refer to the combination of non-rotating elements throughout the engine 10.

[0035] 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 gases are 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.

[0036] A portion of the pressurized air 76 can be drawn from the compressor section 22 as bleed air 77. Bleed air 77 can be drawn from the pressurized air 76 and supplied to engine components requiring cooling. The temperature of the pressurized air 76 entering the combustor 30 increases significantly to a temperature higher than that of the bleed air 77. Bleed air 77 can be used to reduce the temperature of core components downstream of the combustor.

[0037] The remainder of the airflow 78 bypasses the LP compressor 24 and the engine core 44, and exits the engine 10 at the fan exhaust side 84 via a stationary blade row (and more specifically, an outlet guide blade assembly 80 comprising multiple airfoil guide blades 82). More specifically, adjacent to the fan section 18, a circumferential row of radially extending airfoil guide blades 82 is used to exert some directional control on the airflow 78.

[0038] Some of the air supplied by fan 20 can bypass engine core 44 and be used to cool parts of engine 10, particularly hot parts, and / or to cool or power other aspects of the aircraft. In the context of a turbofan engine, the hot parts of the engine are typically downstream of combustor 30, particularly downstream of turbine section 32, with HP turbine 34 being the hottest part as 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.

[0039] Now for reference Figure 2 It shows the source Figure 1 The engine component of engine 10 is in the form of turbine blades 68. A plurality of circumferentially spaced airfoils 92 capable of rotating about a centerline 12 are shown as a pair of turbine blades 90. Alternatively, in a non-limiting example, the engine component may be a wheel blade, strut, service pipe, shroud, or combustion liner, or any other engine component that may require or utilize cooling passages.

[0040] Airfoil 92 may extend radially between a tip 94 and a root 96, which define the spanwise direction therebetween. In one aspect disclosed herein, tip 94 is an unshielded tip as shown. In other words, there is no tip rail at tip 94. Airfoil 92 may be mounted to platform 98 at root 96. Platform 98 facilitates radial accommodating of the turbine engine's mainstream airflow. Furthermore, airfoil 92 may include an outer wall 104 having a first side 106 and a second side 108. Outer wall 104 further extends between an upstream edge and a downstream edge to define the flow direction therebetween. It should be understood that the upstream edge may be the leading edge of airfoil 92, and the downstream edge may be the trailing edge of airfoil 92. Furthermore, as shown, the first side 106 may be the pressure side of a steering blade, and the second side 108 may be the suction side of a steering blade. It is also contemplated that airfoil 92 may be a non-steering blade, and as a non-limiting example, it may be a frame fairing. It is further envisioned that neither the first side 106 nor the second side 108 is bent to form a pressure side and / or a suction side. The outer wall 104 may define an interior 114 in which at least one cooling conduit 116 (shown in dashed lines as two exemplary cooling conduits 116) may be located.

[0041] The true chord length (L) is measured along a straight line extending between the leading edge 110 and the trailing edge 112 of the airfoil 92. The axial chord length (AW) is defined as the axial distance from the foremost point to the rearmost point of the airfoil 92. The number of blades arranged circumferentially around the centerline 12 can vary. The number of blades affects the pitch (S), which is associated with the arrangement and defined as the distance between consecutive airfoils 92. The ratio (AW / S) between the axial chord length (AW) and the pitch (S) is called the robustness parameter. Given that the axial chord length (AW) remains constant, a higher robustness number is generally associated with a higher number of blades, and vice versa.

[0042] In operation, a hot gas flow (H) (such as a burner flow) may pass along the outer side 118 of the outer wall 104 of the airfoil 92 to define a heated surface. A cooling fluid flow (C) may be provided via inlet 102 to at least one cooling duct 116. The cooling fluid flow (C) may be provided throughout the interior 114 and discharged as a cooling film through any exemplary cooling hole 120. Any surface facing the cooling fluid flow (C) may be defined as a cooled surface.

[0043] The stagnation point (SP) is located on the outer edge 118, where the hot gas flow (H) contacts the airfoil 92 at a 90-degree angle, and the velocity of the hot gas flow (H) is zero. In some cases, the stagnation point (SP) is collinear with the leading edge 110, but it should be understood that the stagnation point (SP) can vary to some extent along the leading edge 110. Furthermore, during all or part of the engine's operating conditions, the stagnation point (SP) can be temporarily or permanently different from all or part of the leading edge 110.

[0044] Turning Figure 3 A set of turbine blades 90 is shown, with some part numbers omitted for clarity. The set of turbine blades 90 includes consecutive airfoils 92, a first airfoil 92a and a second airfoil 92b, spaced apart to define a throat 122 through which a flow of hot gas (H) passes. The throat 122 has a width (W) defined as the minimum distance between a point 124 on the trailing edge 112 of the first airfoil 92a and a second side 108 (e.g., the suction side) of the second airfoil 92b. A first line 126 may extend through the point 124 tangent to the second side 108. In other words, the first line 126 is perpendicular to the width (W) measurement. A second line 128 may extend tangent to the second side 108 through the trailing edge 112 of the second airfoil 92b. An unguided steering (UGT) angle (θ) is defined as a measurement between the first line 126 and the second line 128.

[0045] Axial line 130 can be approximately parallel to center line 12 ( Figure 1The chord 132 extends within 5% of the axial line 130 and intersects the trailing edge 112 of the first airfoil 92a. The chord 132 extends between the leading edge 110 and the trailing edge 112 of the first airfoil 92a. The stagger angle (β) is defined as the measurement between the axial line 130 and the chord 132.

[0046] Although the first line 126 and the second line 128 are shown to be tangent to the second airfoil 92b, and the axial line 130 and the chord line 132 are shown to intersect the trailing edge 112 of the first airfoil 92a, it should be understood that these measurements are applicable to all airfoils 92 described herein.

[0047] Figure 4 This is a perspective view of airfoil 192, which is a variation of airfoil 92, as disclosed herein. Airfoil 192 is substantially similar to airfoil 92; therefore, similar parts will be identified by similarity numbers incremented by 100. It should be understood that, unless otherwise stated, the description of similar parts of airfoil 92 applies to airfoil 192.

[0048] Airfoil 192 includes an inflection point 240 near the leading edge 210 along the pressure side 206. The inflection point 240 defines the origin of an enlarged portion 242 extending along the pressure side 206 toward the leading edge 210. The enlarged portion 242 is defined by an outer wall 204 between the inflection point 240, the pressure side 206, the leading edge 210, a second side 208 (e.g., the suction side), and a dashed line 244 extending from the second side 208 back to the inflection point 240. The enlarged portion 242 allows for more space within the interior 114 for cooling duct 116. Figure 2 ).

[0049] Furthermore, the enlarged portion 242 can increase stagnation point (SP) stability by increasing the stagger angle (β) at the tip 194 of the airfoil 192. The inlet flow angle is the trajectory of the airflow as it enters the plane at the leading edge 210 of the airfoil 192. Under different engine operating conditions, the turbine rotor may accelerate or decelerate, and the inlet flow angle may change, which means the stagnation point (SP) will also change. If the pressure side is too flat, as shown by the dashed line 244, a small change in the inlet flow angle can lead to a large change in the stagnation point (SP) position. This can have a significant impact on cooling film coverage and durability. The enlarged portion 242 can be implemented in highly staggered tips, such as tip 194 with a high stagger angle (β), to minimize or eliminate any movement of the stagnation point (SP).

[0050] The distance between the tip 194 and the root 196 of the airfoil 192 is defined as the span length of the airfoil 192. The UGT angle (θ) and the crossover angle (β) can vary along the span length. The first lines 226r and 226t, located at the root 196 and the tip 194 respectively, can extend through the point 224 tangent to the second side 208. The second lines 228r and 228t, located at the root 196 and the tip 194 respectively, can extend tangent to the second side 208 through the trailing edge 212 of the airfoil 192. As shown, the UGT angle (θ) can change from the first UGT angle (θ1) at the root 196 to the second UGT angle (θ2) at the tip 194. The axial lines 230r and 230t, located at the root 196 and the tip 194 respectively, can be approximately parallel to the centerline 12 ( Figure 1 The chords 232r and 232t, located at the root 196 and tip 194 respectively, can extend between the leading edge 210 and trailing edge 212 of the airfoil 192. The staggered angle (β) can change from a first staggered angle (β1) at the root 196 to a second staggered angle (β2) at the tip 194.

[0051] The UGT angle (θ) can decrease along the span length moving from the root 196 toward the tip 194. As a non-limiting example, the UGT angle (θ) can be 30° at 0% of the span length or at the root 96, 23° at 50% of the span length, and 10° at 100% of the span length or at the tip 94.

[0052] On the other hand, the stagger angle (β) can increase along the span length moving from the root 196 toward the tip 194. As a non-limiting example, the stagger angle (β) is 47° at 0% of the span length or at the root 96, 53° at 50% of the span length, and 57° at 100% of the span length or at the tip 94.

[0053] Figure 5A This is a top view of the airfoil profile 250 of the airfoil member 192 at the tip 194, shown in solid lines. Dashed lines represent typical or known airfoil profiles 252 at the tip 194. It can be seen that the stagger angle (β2) increases at the tip 194 relative to the stagger angle (β3) of the typical airfoil profile 252. Furthermore, the airfoil profile 250 is elongated, such that the true chord length (L2) at the tip 194 is greater than the true chord length (L3) of the typical airfoil profile 252.

[0054] Figure 5BThis is a top view of the airfoil profile 254 of the airfoil member 192 at the root 196. The dashed lines represent a typical or known airfoil profile 256 at the root 196. It can be seen that the stagger angle (β1) decreases at the root 196 relative to the stagger angle (β4) of the typical airfoil profile 256. The inflection point 240 and further the enlarged portion 242 are more pronounced relative to the typical airfoil profile 256. As previously described herein, the enlarged portion 242 allows the elongated airfoil profile to have a stable position for stabilization.

[0055] Turning Figure 6 The graph shows the UGT angle (θ) versus the stagger angle (β) for different span lengths and blade arrangements. The graph includes lines associated with different span lengths (including 0%, 20%, 50%, 80%, and 100%). Furthermore, the graph includes lines associated with different blade arrangements, represented by different robustness values ​​“E,” “F,” and “G.” Moving from left to right on the graph indicates decreasing robustness, such that the value “E” is greater than the value “F,” and the value “F” is greater than the value “G.” The robustness value “G” can be between 0.4 and 0.6. The robustness value “G” can also be associated with a smaller number of airfoils arranged 12 circumferentially around the centerline compared to the arrangement represented by the robustness value “E.” The robustness value “E” can be between 0.6 and 0.8, and the robustness value “F” can be between 0.5 and 0.7. Compared to the robustness value "E", the UGT angle (θ) and the stagger angle (β) are typically larger at all points along the span length, exhibiting robustness of value C. Optimizing efficiency while controlling for tip loss requires consideration of robustness, stagger angle (β), and UGT angle (θ).

[0056] Turning Figure 7 The graph shows the ratio of the UGT angle (θ) and the stagger angle (β) (UGT / stagger) to the span length. Different robustness values ​​“E”, “F”, and “G” are also shown. It has been found, and can be seen in the graph, that the relationship is very close to linear regardless of the robustness value, as indicated by line 150.

[0057] All the different robustness values ​​“E”, “F”, and “G” fall within the limits shown by the upper limit line 152 and the lower limit line 154. The y-intercept between these limits can vary by a value of 0.2, or a distance of 150 from line 150. This relationship is represented by the following equation (1). The following equation (2) describes the upper limit line 152 of the curve. The following equation (3) describes the lower limit line 154 of the curve.

[0058] (1) .

[0059] (2) .

[0060] (3) .

[0061] Further, consider an upper limit line 152 and a lower limit line 154 separated from line 150 by 0.14 or... The value. Furthermore, the range can be 0.10 or 150 units from the line. The value. Even further, consider that the upper limit line 152 and the lower limit line 154 are separated from line 150 by 0.06 or... The value of . It should be understood that the narrowing of the range increases the optimization of the airfoil profile 250. When considering robustness, equation (1) can jointly adjust the stagger angle (β) and the UGT angle (θ). Optimizing the stagger angle (β) at the tip 94 improves efficiency while reducing tip losses. Since both UGT and stagger are related to robustness and have similar functional relationships, the UGT / Stagger ratio can define the optimal 3D airfoil load distribution for any selected robustness. For unshielded blades, the optimal 3D load distribution means minimizing leakage losses at the blade tip and channel vortex losses at the hub. In the examination Figure 7 When the representation is optimally constructed, several trends become apparent:

[0062] 1) In order to control hub vortex growth, the UGT tends to be high at the blade hub and low at the crossover.

[0063] 2) In order to control the leakage vortex intensity at the tip, the UGT tends to be low at the blade tip and high at the cross-cutting.

[0064] 3) Both UGT and stagger are highly dependent on robustness, where lower robustness designs often require higher levels of stagger and UGT across all spans due to increased spacing and lift requirements per airfoil.

[0065] Turning Figure 8 A flowchart illustrating a method 200 for forming an engine component is shown. The method includes forming a turbine disk 71 rotatable about a centerline 12 at block 203. A plurality of turbine blades 68, as described herein, are formed at block 205, wherein a first side 106 defines a pressure side and a second side 108 defines a suction side. At block 207, each of the plurality of blades is formed having an unguided steering angle to crossover angle ratio UGT:S relative to a span length defined using a linear equation (1) as described herein, wherein the value of SL in the following linear equation is 0%, 20%, 50%, 80%, or 100%, used to calculate the unguided steering angle to crossover angle ratio at 0%, 20%, 50%, 80%, or 100% of the span length: The method may also include mounting a plurality of turbine blades 68 to the turbine disk 71 at frame 209 with a robustness between 0.4 and 0.8.

[0066] Benefits associated with this disclosure include improved turbine efficiency and, consequently, improved overall fuel consumption rate (SFC) for a given level of robustness. The disclosure enables low-robustness designs with higher performance than conventional designs. This results in greater SFC gains due to a reduced number of blades to be cooled, by decreasing the amount of cooling flow required. With improved SFC, lower robustness can reduce initial and maintenance costs due to the fewer parts.

[0067] The cooling channels described in this article can be produced using additive manufacturing and advanced casting techniques.

[0068] It should be understood that the disclosed design is not limited to turbine engines with fan and supercharger sections, but is also applicable to turbojet engines and land-based turbine engines.

[0069] Within the scope not yet described, different features and structures of each aspect may be combined or substituted for one another as needed. A feature not shown in all examples is not to be interpreted as something that cannot be shown in this way, but rather is done for the sake of brevity. Therefore, various features of different aspects may be mixed and matched as needed to form new aspects, whether or not the new aspects are explicitly described. All combinations or permutations of the features described herein are covered by this disclosure.

[0070] This written description uses examples to illustrate aspects of the disclosure described herein, including best practices, and also enables any person skilled in the art to practice aspects of the disclosure, including making and using any apparatus or system and performing any combination of methods. The patentable scope of aspects of this disclosure is defined by the claims and 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 have 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.

[0071] Further aspects of this disclosure are provided by the subject matter of the following clauses:

[0072] An airfoil for a turbine engine includes an outer wall defining an interior, defining a pressure side and a suction side, the pressure side and the suction side extending between a leading edge and a trailing edge, and extending between a root and a tip to define a span length; wherein the airfoil has an unguided steering angle to crossover angle ratio UGT:S relative to the span length defined by the following linear equation, wherein the value of SL in the following linear equation is 0%, 20%, 50%, 80%, or 100%, used to calculate the unguided steering angle to crossover angle ratio at 0%, 20%, 50%, 80%, or 100% of the span length: UGT:S = -0.48SL + 0.66 ± 0.1.

[0073] According to any of the foregoing clauses, the y-intercept of the linear equation has a range of ±0.07.

[0074] According to any of the foregoing clauses, the y-intercept of the linear equation has a range of ±0.03.

[0075] According to any of the foregoing clauses, the airfoil further includes an inflection point along the pressure side.

[0076] According to any of the foregoing clauses, at least a portion of the leading edge includes an enlarged portion terminating at the inflection point.

[0077] According to any of the foregoing clauses, the airfoil further includes cooling channels within the interior.

[0078] According to any of the foregoing clauses, the airfoil tip is unshielded.

[0079] An airfoil as described in any of the foregoing clauses, wherein the airfoil is a blade.

[0080] An engine component for a turbine engine has a working airflow separated into a cooling airflow and a combustion airflow. The engine component includes a plurality of circumferentially spaced airfoils rotatable about a centerline defining an axial direction. Each airfoil includes an outer wall defining an interior, a pressure side, and a suction side. The pressure side and the suction side extend in a chordal direction between a leading edge and a trailing edge to define a chord, and extend in a span direction between a root and a tip to define a span length. Each airfoil has an unguided steering angle to crossover angle ratio UGT:S relative to the span length defined by the following linear equation, where the value of SL in the following linear equation is 0%, 20%, 50%, 80%, or 100%, used to calculate the unguided steering angle to crossover angle ratio at 0%, 20%, 50%, 80%, or 100% of the span length: UGT:S = -0.48SL + 0.66 ± 0.1.

[0081] Engine components according to any of the foregoing clauses, wherein the plurality of airfoils have a robustness between 0.4 and 0.8.

[0082] According to any of the preceding clauses, the engine component includes a plurality of airfoils comprising a first airfoil and a second airfoil spaced apart from each other to define a throat having a width, the width being defined as the minimum distance between the trailing edge of the first airfoil and a point along the suction side of the second airfoil.

[0083] According to any of the preceding clauses, the unguided steering angle is a measured angle between a first line tangent to the suction side of the second airfoil at the point defined by the minimum distance and a second line tangent to the trailing edge of the second airfoil along the suction side.

[0084] According to any of the preceding clauses, the engine component, wherein the interlacing is a measured angle between an axial line that intersects the trailing edge and extends parallel to the centerline (12) and the chord.

[0085] According to any of the preceding clauses, the y-intercept of the linear equation has a range of ±0.07.

[0086] According to any of the preceding clauses, the y-intercept of the linear equation has a range of ±0.03.

[0087] According to any of the preceding clauses, the engine component, wherein at least one of the plurality of airfoils includes an inflection point along the pressure side.

[0088] An engine component according to any of the foregoing clauses, wherein at least a portion of the leading edge includes an enlarged portion terminating at the inflection point.

[0089] Engine components according to any of the foregoing clauses, wherein the tip is unshielded.

[0090] A method of forming an engine component, the method comprising: forming a turbine disk rotatable about a centerline; forming a plurality of blades, each blade including a wall having a first side and a second side, each side extending radially between a root and a tip to define a span length; and forming each of the plurality of blades having an unguided steering angle to crossover angle ratio UGT:S relative to the span length using a linear equation defined as follows: where the value of SL in the linear equation is 0%, 20%, 50%, 80%, or 100%, used to calculate the unguided steering angle to crossover angle ratio at 0%, 20%, 50%, 80%, or 100% of the span length: UGT:S = -0.48SL + 0.66 ± 0.07.

[0091] The method according to any of the foregoing clauses further includes mounting the plurality of blades to the turbine disk with a robustness between 0.4 and 0.8.

Claims

1. An airfoil for a turbine engine, characterized in that, The airfoil includes: An outer wall defines the interior, defines a pressure side and a suction side, the pressure side and the suction side extending between a leading edge and a trailing edge, and extending between a root and a tip to define a span length; The airfoil has an unguided steering angle to crossover angle ratio UGT:S relative to the span length, defined by the following linear equation, where the value of SL in the following linear equation is 0%, 20%, 50%, 80%, or 100%, used to calculate the unguided steering angle to crossover angle ratio at 0%, 20%, 50%, 80%, or 100% of the span length: 。 2. The airfoil according to claim 1, characterized in that, in, The y-intercept of the linear equation has The range.

3. The airfoil according to any one of claims 1-2, characterized in that, in, The y-intercept of the linear equation has The range.

4. The airfoil according to claim 1, characterized in that, It further includes the inflection point along the pressure side.

5. The airfoil according to claim 4, characterized in that, in, At least a portion of the leading edge includes an enlarged portion that terminates at the inflection point.

6. The airfoil according to claim 5, characterized in that, It further includes cooling channels within the interior.

7. The airfoil according to claim 1, characterized in that, in, The tip is unshielded.

8. The airfoil according to claim 1, characterized in that, in, The airfoil is a blade.

9. An engine component for a turbine engine, characterized in that, The engine component has a working airflow that is separated into a cooling airflow and a combustion airflow. The engine component includes: A plurality of circumferentially spaced airfoils, the plurality of circumferentially spaced airfoils being rotatable about a centerline defining an axial direction, each airfoil comprising: The outer wall defines the interior, defines a pressure side and a suction side, the pressure side and the suction side extending in a chordal direction between the leading edge and the trailing edge to define a chord line, and extending in a span direction between the root and the tip to define a span length. Each airfoil has an unguided steering angle to crossover angle ratio UGT:S relative to the span length, defined by the following linear equation, where the value of SL in the following linear equation is 0%, 20%, 50%, 80%, or 100%, used to calculate the unguided steering angle to crossover angle ratio at 0%, 20%, 50%, 80%, or 100% of the span length: 。 10. The engine component according to claim 9, characterized in that, in, Multiple airfoil components have a robustness between 0.4 and 0.

8.

11. The engine component according to claim 9, characterized in that, in, The plurality of airfoils include consecutive first and second airfoils spaced apart from each other to define a throat having a width, the width being defined as the minimum distance between the trailing edge of the first airfoil and a point along the suction side of the second airfoil.

12. The engine component according to claim 11, characterized in that, in, The unguided steering angle is a measured angle between a first line tangent to the suction side of the second airfoil at the point defined by the minimum distance and a second line tangent to the trailing edge of the second airfoil along the suction side.

13. The engine component according to any one of claims 9-12, characterized in that, in, The intersecting angle is the measured angle between the axial line that intersects the trailing edge and extends parallel to the center line (12) and the chord.

14. The engine component according to claim 9, characterized in that, in, The y-intercept of the linear equation has a range of ±0.

07.

15. The engine component according to claim 9, characterized in that, in, The y-intercept of the linear equation has a range of ±0.

03.

16. The engine component according to claim 9, characterized in that, in, At least one of the multiple airfoils includes an inflection point along the pressure side.

17. The engine component according to claim 16, characterized in that, in, At least a portion of the leading edge includes an enlarged portion that terminates at the inflection point.

18. The engine component according to claim 9, characterized in that, in, The tip is unshielded.

19. A method for forming an engine component, characterized in that, The method includes: Forming a turbine disk capable of rotating around a central axis; Multiple blades are formed, each blade including a wall having a first side and a second side, each side extending radially between the root and the tip to define a span length; and Each of the plurality of blades is formed having an unguided steering angle to crossover angle ratio UGT:S relative to the span length, defined by the following linear equation, wherein the value of SL in the following linear equation is 0%, 20%, 50%, 80%, or 100%, used to calculate the unguided steering angle to crossover angle ratio at 0%, 20%, 50%, 80%, or 100% of the span length: 。 20. The method according to claim 19, characterized in that, This further includes mounting the plurality of blades to the turbine disk with a robustness between 0.4 and 0.8.

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