Turbine engine with shock wave attenuation

Abstract

Aspects of the present disclosure generally relate to a turbine engine and a method of operating a turbine engine having an engine core including a compressor, a combustor, and a turbine in an axial flow arrangement such that a working airflow passes from the compressor through the engine core to the turbine to define a flow direction through the engine core. The method includes generating a shock wave in a working gas flow propagating in a flow direction, and at least partially attenuating the shock wave.

Description

[0001] This application is a divisional application of the patent application with application number 202110779855.9 and invention title "Turbine Engine with Shock Wave Attenuation" filed on July 9, 2021.

[0002] Research sponsored by the federal government

[0003] This invention was made with government support under contract. The government may have certain rights in this invention. Technical Field

[0004] The present disclosure generally relates to turbine engines, and more particularly, to turbine engines configured to attenuate shock waves within the engine. Background Art

[0005] A turbine engine (particularly a gas or combustion turbine engine) is a rotary engine that extracts energy from a process air or working air that flows through the engine via a plurality of compressor stages, through a combustor, and then through a plurality of turbine stages. The compressor stages and turbine stages include pairs of rotating blades and stationary vanes arranged axially.

[0006] During engine operation, the rotation of the blades causes flow disturbances, such as shock waves, within the working air flow as the working air flow advances axially through the engine. The shock waves can propagate through the engine and encounter components located therein, thereby generating pressure disturbances that negatively affect the axial flow of the working air, which results in a reduction in the operating efficiency of the engine. Summary of the Invention

[0007] Aspects and advantages of the present disclosure will be set forth in part in the following description, or may be apparent from the description, or may be learned by practice of the present disclosure.

[0008] In one aspect, the present disclosure relates to a method of operating a turbine engine having an engine core that includes a compressor, a combustor, and a turbine arranged in an axial flow configuration such that a working air flow passes from the compressor through the engine core to the turbine to define a flow direction through the engine core. The method includes generating a shock wave in the working air flow propagating in the flow direction and directing the shock wave onto at least one attenuation structure of at least one component within the turbine engine to at least partially attenuate the shock wave.

[0009] In another aspect, the present disclosure relates to a turbomachine. The turbomachine includes: an engine core having a compressor, a combustor, and a turbine arranged in an axial flow configuration; a flow path extending from the compressor through the engine core to the turbine to define a flow direction of a working gas flow through the engine core; a rotatable set of airfoils in one of the compressor or the turbine such that rotation of the set of airfoils forms a working gas flow having shock waves; and at least one component in the flow path, the at least one component including an outer surface having at least one attenuation structure configured to at least partially attenuate the shock waves flowing thereon.

[0010] These and other features, aspects, and advantages of the present disclosure will be better understood with reference to the following description and the appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate aspects of the present disclosure and, together with the description, serve to explain the principles of the present disclosure. BRIEF DESCRIPTION OF THE DRAWINGS

[0011] A complete and enabling disclosure of the present invention, including the best mode thereof, for the ordinary skill in the art, is set forth in the specification, which makes reference to the accompanying drawings, in which:

[0012] Figure 1 is a schematic cross-sectional view of a gas turbine engine of an aircraft including an exemplary component in the form of a turbine center frame in accordance with various aspects described herein.

[0013] Figure 2 is Figure 1 a schematic cross-sectional view of the turbine center frame of

[0014] Figure 3 is a schematic perspective view of an airfoil usable in a turbine center frame having an attenuation structure in accordance with various aspects described herein Figure 1 for

[0015] Figure 4 is Figure 3 a cross-sectional view along line IV-IV of a portion of the airfoil of

[0016] Figure 5 is Figure 3 a cross-sectional view along line V-V of another portion of the airfoil of

[0017] Figure 6 is a schematic perspective view of a pair of airfoils usable in a turbine center frame having another attenuation structure in accordance with various aspects described herein Figure 1 for

[0018] Figure 7 is a schematic perspective view of a pair of airfoils usable in a turbine center frame having another attenuation structure in accordance with various aspects described hereinFigure 1 Schematic perspective view of another airfoil in the turbine center frame.

[0019] Figure 8 is Figure 6 Cross-sectional view of a part of the airfoil along line VIII-VIII.

[0020] Figure 9 is a schematic perspective view of another airfoil in the turbine center frame that can be used with another damping structure according to various aspects described herein. Figure 1 Schematic perspective view of another airfoil in the turbine center frame.

[0021] Figure 10 is Figure 8 Cross-sectional view of a part of the airfoil along line X-X.

[0022] Figure 11 is a schematic perspective view of another airfoil in the turbine center frame that can be used with another damping structure according to various aspects described herein. Figure 1 Schematic perspective view of another airfoil in the turbine center frame.

[0023] Figure 12 is a schematic perspective view of another airfoil in the turbine center frame that can be used with another damping structure according to various aspects described herein. Figure 1 Schematic perspective view of another airfoil in the turbine center frame.

[0024] Figure 13 is a schematic perspective view of another airfoil in the turbine center frame that can be used with another damping structure according to various aspects described herein. Figure 1 Schematic perspective view of another airfoil in the turbine center frame.

[0025] Figure 14 is a schematic perspective view of another airfoil in the turbine center frame that can be used with another damping structure according to various aspects described herein. Figure 1 Schematic perspective view of another airfoil in the turbine center frame. Detailed Description

[0026] Recent trends in engine technology, including turbomachinery technology, have resulted in increased turbine blade loading requirements, including local Mach numbers and the resulting compressible flow phenomena, such as shock waves and expansion waves in the working gas flow. Such flow characteristics can propagate downstream along the engine flow path, creating pressure losses and reduced efficiency in downstream components. Additionally, some shock waves reflect from downstream components and create subsequent aerodynamic effects, which also negatively impact system performance. Further, some engine architectures have closely coupled high-pressure turbines (HPTs) and low-pressure turbines (LPTs), which results in increased aerodynamic or aeroelastic interactions between the HPT and LPT. Aspects of the present disclosure relate to mitigating or attenuating such flow phenomena, including shock waves, by using surface features on engine components that may be exposed to shock waves or reflected shock waves in the flow path.

[0027] For illustrative purposes, an exemplary component will be described in the form of a turbine engine having a turbine center frame positioned between a high-pressure turbine and a low-pressure turbine. In a non-limiting example, such a turbine engine can be in the form of a gas turbine engine, a turboprop engine, a turboshaft engine, or a turbofan engine. However, it will be understood that the aspects of the present disclosure described herein are not limited thereto and can have general applicability to other turbine engine components or within other engine systems. For example, the present disclosure can be applicable to systems in other engines or vehicles and can be used to provide benefits in industrial, commercial, and residential applications.

[0028] As used herein, the term "upstream" refers to the direction opposite to the fluid flow direction, while the term "downstream" refers to the direction the same as the fluid flow direction. The terms "front" or "forward" refer to in front of something, and "rear" or "backward" refer to behind something. For example, when used in the context of fluid flow, front / forward can denote upstream, while rear / backward can denote downstream.

[0029] Further, as used herein, the term "shock wave" or "shock" will broadly refer to a flow perturbation in the form of a rapid or sharp pressure wave traveling or propagating through a medium. Such shock waves used herein can travel faster or slower than the speed of sound through the medium. In one example, a shock wave can be generated by a body moving through the medium at a high speed relative to the speed of sound in the medium (e.g., in a non-limiting example, greater than 80% of the speed of sound, or between 70% - 150% of the speed of sound). In another example, a shock wave can be generated by a supersonic expansion of a fluid flowing over or past a stationary body. In yet another example, a shock wave can refer to a pressure wave within a subsonic fluid flow.

[0030] Furthermore, as used herein, the terms "radial" or "radially" refer to a direction away from a common center. For example, in the overall context of a turbine engine, radial refers to the direction along a ray extending between the central longitudinal axis of the engine and the outer engine circumference. Additionally, as used herein, the term "group" or "a group of" elements can be any number of elements, including only one element.

[0031] 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 assist the reader in understanding the disclosure and should not be construed as limitations on the embodiments, particularly as to the position, orientation, or use of aspects of the disclosure described herein. Unless otherwise stated, connection references (e.g., attached, coupled, fixed, connected, joined, etc.) should be construed broadly and may include intermediate members between assemblies of elements and relative movement between elements. Thus, a connection reference does not necessarily infer that two elements are directly connected and have a fixed relationship to each other. Unless the context clearly dictates otherwise, the singular forms "a," "an," and "the" include plural references.

[0032] As used throughout the specification and claims, approximating language is applied to modify any quantitative representation that can vary without resulting in a change in the basic function associated therewith. Accordingly, values modified by terms such as "about," "approximately," and "substantially" are not limited to the specified exact values. In at least some instances, the approximating language can 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 components and / or systems. For example, the approximating language can refer to within a range of 10%.

[0033] Exemplary drawings are for illustrative purposes only, and the dimensions, positions, sequences, and relative sizes reflected in the attached drawings may vary.

[0034] Figure 1 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 that extends from a front 14 to a rear 16. The engine 10 includes, in a downstream serial flow relationship: a fan section 18 that includes a fan 20; a compressor section 22 that includes a booster or low-pressure (LP) compressor 24 and a high-pressure (HP) compressor 26; a combustion section 28 that includes a burner 30; a turbine section 32 that includes an HP turbine 34 and an LP turbine 36; and an exhaust section 38.

[0035] The fan section 18 includes a fan casing 40 that surrounds the fan 20. The fan 20 includes a plurality of fan blades 42 radially disposed about the centerline 12. The LP compressor 24, HP compressor 26, combustor 30, HP turbine 34, and LP turbine 36 together form the core 44 of the engine 10, which generates combustion gases. The core 44 is surrounded by a core casing 46, which may be coupled to the fan casing 40.

[0036] An HP shaft or spool 48 coaxially disposed about the centerline 12 of the engine 10 drivingly connects the HP turbine 34 to the HP compressor 26. An LP shaft or spool 50 coaxially disposed about the centerline 12 of the engine 10 within the larger diameter annular HP spool 48 drivingly connects the LP turbine 36 to the LP compressor 24 and the fan 20. The spools 48, 50 are rotatable about the engine centerline and are coupled to a plurality of rotatable elements that together may define a rotor 51.

[0037] The LP compressor 24 and HP compressor 26 each include a plurality of compressor stages 52, 54, where a set of compressor blades 56, 58 rotate relative to a corresponding set of stationary compressor vanes 60, 62 to compress or pressurize the fluid flow passing through the stage. In a single compressor stage 52, 54, the plurality of compressor blades 56, 58 may be arranged in a ring and may extend radially outward from a blade platform to a blade tip relative to the centerline 12, while the corresponding stationary compressor vanes 60, 62 are positioned upstream and adjacent to the rotating blades 56, 58. It should be noted that Figure 1 the number of blades, vanes, and compressor stages shown is chosen for illustrative purposes only, and other numbers are possible.

[0038] The blades 56, 58 of the compressor stage may be mounted to (or integrated with) a disk 61, which is mounted to the corresponding one of the HP and LP spools 48, 50. The vanes 60, 62 of the compressor stage may be mounted to the core casing 46 in a circumferential arrangement.

[0039] The HP turbine 34 and LP turbine 36 each include a plurality of turbine stages 64, 66, where a set of turbine blades 68, 70 rotate relative to a corresponding set of stationary turbine vanes 72, 74 (also referred to as nozzles) to extract energy from the fluid flow passing through the stage. In a single turbine stage 64, 66, the plurality of turbine blades 68, 70 may be arranged in a ring and may extend radially outward relative to the centerline 12, while the corresponding stationary turbine vanes 72, 74 are positioned upstream and adjacent to the rotating blades 68, 70. It should be noted that Figure 1 the number of blades, vanes, and turbine stages shown is chosen for illustrative purposes only, and other numbers are possible.

[0040] The blades 68, 70 of the turbine stage can be mounted to a disk 71 which is mounted to the corresponding one of the HP and LP spools 48, 50. The vanes 72, 74 of the compressor stage can be mounted to the core housing 46 in a circumferential arrangement.

[0041] As a complement to the rotor portion, the stationary portions of the engine 10 (such as the stationary vanes 60, 62, 72, 74 in the compressor and turbine sections 22, 32) are also referred to individually or collectively as the stator 63. Thus, the stator 63 can refer to the combination of non-rotating elements throughout the engine 10.

[0042] In operation, the airflow leaving the fan section 18 is split such that a portion of the airflow is directed along a flow path 75 that extends through the core 44 and defines a flow direction 76. More specifically, the airflow moves along the flow path 75 into the LP compressor 24 which then supplies pressurized air to the HP compressor 26 that further pressurizes the air along the flow path 75. The pressurized air from the HP compressor 26 is mixed with fuel and ignited in the combustor 30 to produce combustion gases to form a working airflow. The HP turbine 34 that drives the HP compressor 26 extracts some work from these gases. The combustion gases are discharged into the LP turbine 36 which extracts additional work along the flow path 75 to drive the LP compressor 24, and the exhaust gases are finally discharged from the engine 10 via the exhaust section 38. The driving of the LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP compressor 24.

[0043] A portion of the pressurized airflow can be extracted as bleed air 77 from the compressor section 22. The bleed air 77 can be extracted from the pressurized airflow and supplied to engine components that require cooling. The temperature of the pressurized airflow entering the combustor 30 increases significantly to be higher than the bleed air temperature. The bleed air 77 can be used to reduce the temperature of the core components downstream of the combustor.

[0044] The remaining portion of the airflow bypasses the LP compressor 24 and the engine core 44 and leaves the engine assembly 10 through a stationary vane row (and more specifically, an exit guide vane assembly 80 including a plurality of airfoil guide vanes 82 at the fan exhaust side 84). More specifically, a circumferential row of radially extending airfoil guide vanes 82 is used near the fan section 18 to exert some directional control on the airflow.

[0045] In addition, during operation, shock waves 90 are generated in the working gas flow within the engine core 44. In one example, the rotation of a set of airfoils (such as the HP turbine blades 68) can cause the HP turbine blades 68 to move through the working gas flow at a sufficient speed to generate shock waves 90. In another non-limiting example, shock waves 90 can be generated by supersonic expansion between adjacent HP turbine vanes 72. In yet another example, the rotation of the upstream HP turbine blades 68 can at least partially form the working gas flow, and shock waves 90 can be formed or generated within the working gas flow by rotating the HP turbine blades 68, or by supersonic expansion of the working gas flow between adjacent HP turbine vanes 72, etc., or any combination thereof. Regardless of how it is formed, shock waves 90 can propagate or travel along the flow path 75 and encounter other components within the engine core 44, such as another HP turbine blade 68, HP turbine vane 72, LP turbine blade 70, LP turbine vane 74, core housing 46, etc. Shock waves 90 represent flow disturbances within the working gas flow, such as a rapid increase in pressure propagated through the combustion gases in the engine core 44.

[0046] A portion of the shock wave 90 can also be reflected from engine components to define a reflected portion 91 that moves upstream (i.e., opposite to the flow direction 76) through the engine core 44, as shown. The reflected portion 91 can propagate upstream and also encounter other components within the engine core 44, including the HP turbine blade 68, HP turbine vane 72, LP turbine blade 70, LP turbine vane 74, core housing 46, etc. Shock waves 90 and the reflected portion 91 represent flow disturbances that can interfere with the normal or standard axial flow of the working gas flow through the engine core 44.

[0047] For visual clarity, shock waves 90 and the reflected portion 91 are shown by arrows herein. It should be understood that such arrows represent the direction of travel of shock waves 90 and the reflected portion 91, as is commonly understood for wave phenomena. It should also be understood that shock waves 90 and the reflected portion 91 can encounter multiple components within the engine core 44, including reflecting from multiple components to form additional reflected portions. For the purpose of visual clarity, a single shock wave 90 and a single reflected portion 91 will be shown, and it should be understood that multiple shock waves 90 and reflected portions 91 can exist within the working gas flow.

[0048] In addition, for illustrative purposes, an exemplary component that can be encountered by shock waves 90 or the reflected portion 91 will be described in the form of a turbine center frame 92 disposed between the HP turbine 34 and the LP turbine 36. It should be understood that the present disclosure is not limited thereto, and aspects of the present disclosure can also be applied to other components within the engine 10, including portions of the HP turbine 34, LP turbine 36, core housing 46, or inter-turbine ducts or end walls, etc.

[0049] The turbine center frame (also known as the turbine intermediate frame) is typically used in the transition duct between the high-pressure and low-pressure turbines. Such a frame can provide diffusion or smoothing of the turbulent airflow leaving the high-pressure turbine before entering the low-pressure turbine. The airflow entering the turbine center frame may have unsteady, three-dimensional, and non-uniform turbulent motions. These flow characteristics typically result in a high level of mixing and poor performance. The turbine center frame typically includes airfoils to divert, turn, or accelerate the airflow for the purposes of diffusion or smoothing as described above.

[0050] Figure 2 The turbine center frame 92 is shown in more detail. The turbine center frame 92 is part of the transition duct 88 disposed between the HP turbine 34 and the LP turbine 36. The turbine center frame 92 can extend from an inlet 93 at the rear end of the HP turbine 34 to an outlet 94 at the front end of the LP turbine 36.

[0051] The turbine center frame 92 includes a set of airfoils 95. The set of airfoils 95 can be used to change the fluid flow discharged from the HP turbine 34, for example, by diffusing, smoothing, accelerating, or reducing turbulence, so as to provide a smoother airflow entering the LP turbine 36 and extract more work from the airflow, as described above.

[0052] The set of airfoils 95 is illustrated within the turbine center frame 92. In the example shown, an exemplary airfoil 100 of the set of airfoils 95 can extend between a radial inner wall 97 and a radial outer wall 98 within the turbine center frame 92. The span height 106 and chord length 107 of the airfoil 100 are shown for reference. The set of airfoils 95 can include any suitable number of stationary or rotating airfoils.

[0053] A shock wave 90 and a reflected portion 91 are also shown within the engine core 44. For illustrative purposes, the shock wave 90 will be described as originating from a portion of the HP turbine 34, moving through the turbine center frame 92, encountering the set of airfoils 95, and reflecting from a portion of the LP turbine 36 to form the reflected portion 91. It should be understood that the shock wave 90 and the reflected portion 91 can be generated by any suitable engine components within the core 44.

[0054] Steering Figure 3, an exemplary airfoil 100 that can be used for a turbine center frame 92 is shown. It should be understood that aspects of the present disclosure can be used for any airfoil within a turbine engine 10, including any rotating or non-rotating airfoil at any location within the turbine engine 10. The airfoil 100 has an outer wall 101 that extends from a leading edge 102 to a trailing edge 103 and defines a pressure side 104 and a suction side 105. At least one damping structure 110 can be provided with the airfoil 100 and is illustrated as being located on the pressure side 104. In the example shown, the damping structure 110 is in the form of a plurality of planar structures along the outer wall, illustrated as a hexagonal or "honeycomb" structure 112. In one example, the honeycomb structure 112 can be formed of the same material as the outer wall 101. Alternatively, the honeycomb structure 112 can include a different material compared to the outer wall 101. It should be understood that the damping structure 110 can be located at any position on the airfoil 100, the radial inner wall 97, or the radial outer wall 98 ( Figure 2 ). In another non-limiting example where the airfoil 100 is in the form of a blade extending from a platform, at least one damping structure can be located on the blade, the platform, or a combination thereof. In yet another non-limiting example where the airfoil 100 is in the form of a vane extending between an inner band and an outer band, at least one damping structure can be provided on the vane, the inner band, the outer band, or a combination thereof.

[0055] A first group 114 of honeycomb structures 112 is shown near the trailing edge 103. The first group 114 of honeycomb structures 112 includes a first panel 116 and a second panel 118. The first panel 116 can protrude further into the flow path 75 compared to the second panel 118, thereby forming a "raised" first panel 116.

[0056] In addition, a second group 120 of honeycomb structures 112 is shown near the trailing edge 103. The boundaries between adjacent honeycomb structures 112 can form raised ridges 122 that protrude or project into the flow path 75. The ridges 122 can be continuous or discontinuous on the outer wall 101; multiple ridges 122 can also be provided.

[0057] It should be understood that Figure 3 the arrangement of the honeycomb structures 112 shown is provided for illustrative purposes and is not intended to limit the present disclosure. Multiple honeycomb structures 112 can be used on any part of the airfoil 100, including covering the entire outer wall 101 or a portion thereof. In one non-limiting example, the pressure side 104 can be covered by multiple first panels 116 and second panels 118, while the suction side 105 can be covered by multiple raised ridges 122.

[0058] Figure 4The first panel 116 and the second panel 118 are shown in cross-section. As shown, the first panel 116 defines a first panel height 124 and the second panel 118 defines a second panel height 126. In the example shown, the first panel height 124 is greater than the second panel height 126 such that the first panel 116 projects further into the flow path 75 as described above, although this need not be the case. In non-limiting examples, the first panel height 124 or the second panel height 126 can be between 1 mm and 7 mm, or between 5 mm and 50 mm, or between 1 cm and 10 cm. Additionally or alternatively, the first panel height 124 or the second panel height 126 can be designed relative to the dimensions of the airfoil 100. For example, in non-limiting examples, either or both of the first panel height 124 or the second panel height 126 can be between 0% and 1% of the span height 106 of the airfoil 100, or between 0% and 2% of the span height 106, or between 0% and 10% of the chord length 107 of the airfoil 100. It is further contemplated that the first panel height 124 can be equal to the second panel height 126.

[0059] Figure 5 The ridge 122 between adjacent honeycomb structures 112 is shown in cross-sectional view. The honeycomb structures 112 are shown as layers disposed on the outer wall 101 and having a layer height 128. The ridge 122 can define a ridge height 130 that is greater than the layer height 128, as measured from the outer wall 101. In non-limiting examples, the layer height 128 can be between 1 mm and 5 mm, or between 5 mm and 50 mm, or between 1 cm and 10 cm. In other non-limiting examples, the ridge height 130 can be between 2 mm and 10 mm, or between 7 mm and 60 mm, or between 1 cm and 10 cm. In non-limiting examples, it is also contemplated that the ridge height 130 can be sized, formed, designed, etc. relative to the dimensions of the airfoil 100, including between 0% and 2% of the span height 106 of the airfoil 100, or between 0% and 2% of the chord length 107 of the airfoil 100.

[0060] During operation, the shock wave 90 or the reflected portion 91 ( Figure 3) Any one or both of them may propagate along the flow path 75 and flow along or on the outer wall 101. The shock wave may encounter at least one attenuation structure 110 protruding into the flow path 75, including any one or both of the ridges 122, the raised first panel 116 or the second panel 118 of the first set 114 or the second set 120 of honeycomb structures 112. The attenuation structure 110 may be configured to at least partially attenuate the shock wave 90 or the reflected portion 91 flowing thereon. "Attenuation" (also known as "dissipation", "weakening" or "dispersion") may include absorbing the energy carried by the shock wave 90 or the reflected portion 91, reducing the amplitude of the shock wave 90 or the reflected portion 91, reducing or preventing the reflection of the shock wave 90 or the reflected portion 91, scattering the shock wave 90 or the reflected portion 91, etc., or a combination thereof.

[0061] When the shock wave 90 or the reflected portion 91 encounters the outer wall 101, the portion of the attenuation structure 110 protruding into the flow path 75, such as the ridge 122, the raised first panel 118 or the second panel 120, can be used as an elastic member that can vibrate, resonate, etc. to absorb the energy carried by the shock wave 90 or the reflected portion 91. This energy transfer reduces the amplitude, weakens, etc. of the shock wave 90 or the reflected portion 91, thereby at least partially attenuating the shock wave 90 or the reflected portion 91. Compared with a smooth or flat outer wall that provides a bounce of the incident surface wave or flow interruption, this energy transfer can also reduce or prevent the reflection of the shock wave 90 or the reflected portion 91 from the outer wall 101. In this way, the attenuation structure 110 can be in the form of a protruding surface irregularity on an engine component within the flow path 75 and is configured to attenuate the shock wave 90 or the reflected portion 91.

[0062] Now referring to Figure 6 , the first airfoil 200A and the second airfoil 200B that can be used in the turbine center frame 92 are shown. The first airfoil 200A and the second airfoil 200B are similar to the airfoil 100. Therefore, similar components will be identified with similar numbers increased by 100 and appended with A or B. It should be understood that the description of the similar components of the airfoil 100 applies to the first airfoil 200A and the second airfoil 200B unless otherwise specified.

[0063] The first airfoil 200A is circumferentially spaced from the second airfoil 200B. The first airfoil 200A and the second airfoil 200B have respective first outer walls 201A and second outer walls 201B. The first outer wall 201A extends from a first leading edge 202A to a first trailing edge 203A and defines a first pressure side 204A and a first suction side 205A. The second outer wall 201B extends from a second leading edge 202B to a second trailing edge 203B and defines a second pressure side 204B and a second suction side 205B. The first and second airfoils 200A, 200B can be in the form of stationary vanes that extend radially from a radial inner wall 97 of a turbine center frame 92 to a radial outer wall 98 ( Figure 2 ; not shown in Figure 6 for visual clarity).

[0064] In addition, at least one damping structure 210 can be provided on at least one of the first airfoil 200A or the second airfoil 200B. In the example shown, the at least one damping structure 210 is in the form of a first set 231 of scaling structures 234 on the first outer wall 201A and a second set 232 of scaling structures 234 on the second outer wall 201B, each scaling structure having a curved outer surface 233 that projects into the flow path 75. In the example shown, each of the first set 231 and the second set 232 of scaling structures 234 can project into the flow path 75 from the respective first pressure side 204A and second pressure side 204B, although this need not be the case.

[0065] The first set 231 is shown over a portion of the first pressure side 204A. The second set 232 is shown covering the entire second pressure side 204B. It will be understood that any number of scaling structures 234 can be used in any pattern or arrangement in the first set 231 and the second set 232 and on any part of the first outer wall 201A or the second outer wall 201B, including the first pressure side 204A, the second pressure side 204B, the first suction side 205A, or the second suction side 205B. Further, the scaling structures 234 are shown in Figure 6 an example where the outer surface 233 has a generally U-shaped geometric profile. It should be understood that the outer surface 233 of the scaling structure 234 can have any suitable geometry or dimensions, including hemispherical, parabolic, pitted, symmetric, asymmetric, etc., or combinations thereof.

[0066] During operation, the shock wave 90 or the reflected portion 91 can propagate between the first airfoil 200A and the second airfoil 200B along the flow path 75. The shock wave 90 can move or flow along the first outer wall 201A or the second outer wall 201B and encounter an attenuation structure 210, namely any one or both of the first set 231 and the second set 232 of scaling structures 234 that project into the flow path 75. As the shock wave 90 or the reflected portion 91 moves along the scaling structure 234, the curved outer surface 233 of each scaling structure 234 can reflect portions of the shock wave 90 or the reflected portion 91 along multiple different directions, thereby scattering or dispersing the shock wave 90 or the reflected portion 91. The first set 231 and the second set 232 of scaling structures 234 can thus be configured to at least partially attenuate the shock wave 90 or the reflected portion 91 as described above.

[0067] Now referring Figure 7 , another airfoil 300 that can be used in the turbine center frame 92 is shown. The airfoil 300 is similar to the airfoils 100, 200A, 200B. Accordingly, like components will be identified with like numerals further incremented by 100, and it should be understood that the description of like components of the airfoils 100, 200A, 200B applies to the airfoil 300 unless otherwise stated.

[0068] The airfoil 300 has an outer wall 301 that extends from a leading edge 302 to a trailing edge 303 and defines a pressure side 304 and a suction side 305. At least one attenuation structure 310 can be provided on the outer wall 301 and is illustrated as being located on the pressure side 304. One difference is that the attenuation structure 310 is in the form of a resonant cavity 335 that extends into the outer wall 301. In the example shown, a plurality of resonant cavities 335 are provided along the pressure side 304. It is contemplated that any number of resonant cavities 335 can be provided along any portion of the outer wall 301 and in any suitable arrangement or pattern.

[0069] Figure 8 Two resonant cavities 335 are shown in cross-section. Each resonant cavity includes a neck 336 that defines a neck volume 337, and the neck 336 has an opening 338 on the outer wall 301 that is fluidly coupled to the flow path 75 ( Figure 7 ). A chamber 340 is located within the outer wall 301 and defines a chamber volume 341 that is fluidly coupled to the neck 336. Although the chamber 340 and the neck 336 are illustrated as having a substantially rectangular geometric profile, any suitable geometric profile can be utilized, including circular, curved, asymmetric, or irregular geometric profiles.

[0070] The relative dimensions of the chamber volume 341 and the neck volume 337 can be selected or designed to attenuate a specific frequency or frequency range of the shock wave 90 passing through the flow path 75. In other words, the resonator cavity 335 can define a Helmholtz resonator that can be tuned to attenuate the sound waves flowing over the airfoil 300, including the shock wave 90. The resonator cavities 335 can all be the same, or they can have different chamber volumes 341 and neck volumes 337 at different portions of the outer wall 301. In yet another example (not shown), the first airfoil can have a first set of resonator cavities all tuned to a first frequency, while the second airfoil can have a second set of resonator cavities all tuned to a second frequency.

[0071] In operation, the shock wave 90 or the reflected portion 91 can flow over the outer wall 301 of the airfoil 300. The air within the resonator cavity 335 can resonate in response to the presence of the shock wave 90 or the reflected portion 91, thereby removing energy from the shock wave 90 or the reflected portion 91 and at least partially dissipating the shock wave 90 or the reflected portion 91. It is also contemplated that the energy absorbed from the shock wave 90 or the reflected portion 91 can be released back into the flow path 75 via the resonance of the air within the resonator cavity 335, as is known in the art. In this manner, the attenuation structure 310 can be in the form of a resonant mechanism that provides attenuation of the shock wave 90 or the reflected portion 91.

[0072] Now referring Figure 9 , another airfoil 400 that can be used in the turbine center frame 92 is shown. The airfoil 400 is similar to the airfoils 100, 200A, 200B, 300. Accordingly, like components will be denoted with like numerals further incremented by 100, it being understood that the description of like components of the airfoils 100, 200A, 200B, 300 applies to the airfoil 400 unless otherwise specified.

[0073] The airfoil 400 has an outer wall 401 that extends from a leading edge 402 to a trailing edge 403 and defines a pressure side 404 and a suction side 405. At least one attenuation structure 410 can be provided on the outer wall 401 and is illustrated as being located on the pressure side 404. One difference is that the attenuation structure 410 is in the form of at least one cooling hole 445 having an outlet 446 in the outer wall 401. Cooling air 447 from within the airfoil 400 is illustrated as leaving the outlet 446 in a direction that is at least partially opposite to the flow direction 76. In other words, the cooling holes 445 of the attenuation structure 410 can direct the cooling air 447 at least partially in an upstream direction.

[0074] Figure 10Shows a cross-sectional view of one of the cooling holes 445. The cooling hole 445 extends along a passage 448 between an inlet 449 and an outlet 446. The inlet 449 may be fluidly coupled to a source of cooling air within the airfoil 400, including bleed air 77( Figure 1 ). As shown, the outlet 446 may define an outlet centerline 450 that is at least partially opposed to the flow direction 76. Although the passage 448 is illustrated as having a linear geometry, it should be understood that any suitable geometric profile may be utilized, including straight or curved segments, metering passages, diffuser segments, etc., or combinations thereof.

[0075] During operation, the shock wave 90 or the reflected portion 91 may propagate along the flow direction 76 and impinge on or propagate along the outer wall 401. The cooling air 447 may flow out of the cooling hole 445 at least partially opposite to the flow direction 76, facing the shock wave 90 or the reflected portion 91. Thus, the cooling air 447 may remove energy or otherwise disrupt the shock wave 90 or the reflected portion 91, thereby dissipating or attenuating the shock wave 90 or the reflected portion 91. In this manner, the attenuation structure 410 may utilize flow interruption to attenuate the shock wave 90 or the reflected portion 91.

[0076] Now referring to Figure 11 , another airfoil 500 that may be used in the turbine center frame 92 is shown. The airfoil 500 is similar to the airfoils 100, 200A, 200B, 300, 400. Accordingly, like components will be identified with like numerals further incremented by 100, it being understood that the description of like components of the airfoils 100, 200A, 200B, 300, 400 applies to the airfoil 500 unless otherwise stated.

[0077] The airfoil 500 has an outer wall 501 that extends from a leading edge 502 to a trailing edge 503 and defines a pressure side 504 and a suction side 505. At least one attenuation structure 510 may be provided on the outer wall 501 and is illustrated as being located on the pressure side 504. One difference is that the attenuation structure 510 is in the form of a set of ramp structures 560 that project into the flow path 75. The set of ramp structures 560 is illustrated on the pressure side 504, and it is also contemplated that the ramp structures may be located on the suction side 505.

[0078] Each of the set of ramp structures 560 is shown as extending entirely on the pressure side 504 of the airfoil 500, although this need not be the case. The set of ramp structures 560 may have a generally triangular or "sawtooth" geometric profile, where a surface 562 extends from the outer wall 501 and projects into the flow path 75 to form a ramp end 564 as shown. The surface 562 and the ramp end 564 may be oriented relative to the flow direction 76 to face a reflected portion 91 of a shock wave 90 traveling against the flow direction 76 as shown. In this way, the attenuation structure 510 may be arranged to attenuate a reflected portion of a shock wave generated along the flow path during operation of the engine 10. It is contemplated that the surface 562 and the ramp end 564 may be formed with multiple orientations to attenuate or weaken a shock wave 90 or a reflected portion 91 moving in multiple directions while flowing along the outer wall 501.

[0079] Now referring Figure 12 , another airfoil 600 that may be used in the turbine center frame 92 is shown. The airfoil 600 is similar to airfoils 100, 200A, 200B, 300, 400, 500. Accordingly, like components will be identified with like numerals further incremented by 100, and it should be understood that descriptions of like components of airfoils 100, 200A, 200B, 300, 400, 500 apply to the airfoil 600 unless otherwise noted.

[0080] The airfoil 600 has an outer wall 601 that extends from a leading edge 602 to a trailing edge 603 and defines a pressure side 604 and a suction side 605. At least one attenuation structure 610 may be provided on the outer wall 601 and is illustrated as being located on the pressure side 604. One difference is that the attenuation structure 610 is in the form of a set of curved structures 670 that project into the flow path 75. The set of curved structures 670 is shown on the pressure side 604, and it is also contemplated that the curved structures may be located on the suction side 605 as well.

[0081] Each of the set of curved structures 670 is shown as extending entirely on the pressure side 604 of the airfoil 600, although this need not be the case. As shown, the set of curved structures 670 is shown as having a semi-circular outer surface 672 that extends from the outer wall 501 and projects into the flow path 75. The outer surface 672 may define a maximum width 674 that extends away from the airfoil outer wall 601. In non-limiting examples, the maximum width 674 may be between 1 mm and 20 mm, or between 1 cm and 5 cm. In non-limiting examples, it is also contemplated that the maximum width 674 may be selected, sized, etc. relative to the dimensions of the airfoil 600, such as between 0% and 1% of a span height similar to the span height 106, or between Figure 2) between 0% and 1% of the chord length. This arrangement can provide attenuation of the shock wave 90 or the reflected portion 91 as the shock wave 90 or the reflected portion 91 moves over the bend of the attenuation structure 610.

[0082] Now referring to Figure 13 , another airfoil 700 that can be used in the turbine center frame 92 is shown. The airfoil 700 is similar to the airfoils 100, 200A, 200B, 300, 400, 500, 600. Accordingly, like components will be identified with like numerals further increased by 100, and it should be understood that the description of like components of the airfoils 100, 200A, 200B, 300, 400, 500, 600 applies to the airfoil 700 unless otherwise stated.

[0083] The airfoil 700 has an outer wall 701 that extends from a leading edge 702 to a trailing edge 703 and defines a pressure side 704 and a suction side 705. As shown, the outer wall 701 extends in the spanwise direction S, and the circumferential direction C for the turbine center frame 92 ( Figure 2 ) is also shown.

[0084] At least one attenuation structure 710 may be provided on the outer wall 701. One difference is that the attenuation structure 710 is in the form of an arcuate wall portion 775. In the example shown, both the pressure side 704 and the suction side 705 form arcuate portions 775 in the circumferential direction C as shown. In other words, both the pressure side 704 and the suction side 705 also extend at least partially in the circumferential direction C to form an arcuate wall portion 775 as they move in the spanwise direction S. The arcuate wall portion 775 may be convex or concave.

[0085] During operation, the shock wave 90 or the reflected portion 91 may encounter the arcuate wall portion 775, which may serve to disrupt, redirect, or introduce turbulence into the shock wave 90 or the reflected portion 91, thereby attenuating the shock wave 90 or the reflected portion 91. In one example, the arcuate wall portion 775 may create a local pressure gradient within the working airflow such that the shock wave 90 or the reflected portion 91 is disrupted as it flows over the arcuate wall portion 775. In another example, the arcuate wall portion 775 may focus or otherwise redirect the incident shock wave 90 or the reflected portion 91 towards a benign region of the flow path 75, such as a platform, end wall, or other region where the incident shock wave may have reduced adverse effects. In this way, the arcuate wall portion 775 can serve to disperse the incident shock wave or reflect the incident shock wave towards a predetermined location within the engine 10.

[0086] Now referring to Figure 14, shows another airfoil 800 that can be used in the turbine center frame 92. The airfoil 800 is similar to the airfoils 100, 200A, 200B, 300, 400, 500, 600, 700. Accordingly, like components will be identified with like numerals further incremented by 100, and it should be understood that the description of like components of the airfoils 100, 200A, 200B, 300, 400, 500, 600, 700 applies to the airfoil 800 unless otherwise stated.

[0087] The airfoil 800 has an outer wall 801 that extends from a leading edge 802 to a trailing edge 803 and defines a pressure side 804 and a suction side 805. As shown, the outer wall 801 extends in the spanwise direction S, and the circumferential direction C for the turbine center frame 92 ( Figure 2 ) is also shown.

[0088] At least one attenuation structure 810 may be provided on the outer wall 801. One difference is that the attenuation structure 810 is in the form of a convex bow portion 880 located on the pressure side 804. In the example shown, the convex bow portion 880 extends fully between the leading edge 802 and the trailing edge 803. The convex bow portion 880 projects outwardly into the flow path 75 in the circumferential direction C.

[0089] During operation, the shock wave 90 or the reflected portion 91 will encounter the convex bow portion 880, which can serve to disrupt, redirect, or introduce turbulence into the shock wave 90 or the reflected portion 91. For example, the convex bow portion 880 can create a local pressure gradient within the working airflow such that the shock wave 90 or the reflected portion 91 is disrupted as it flows over the convex bow portion 880. In another example, the convex bow portion 880 can redirect the shock wave 90 or the reflected portion 91 in multiple directions in order to scatter the shock wave 90 or the reflected portion 91. In this way, the attenuation structure 810 that projects into the flow path 75 can at least partially attenuate, dissipate, disperse, or scatter the shock wave 90 or the reflected portion 91.

[0090] Reference Figure 1 - 14 , aspects of the present disclosure provide a method of operating a turbine engine 10 that has an engine core 44 having a compressor, a combustor, and a turbine arranged in an axial flow configuration, such as an LP compressor 24, an HP compressor 26, a combustor 30, an HP turbine 34, and an LP turbine 36 ( Figure 1), whereby working air flow passes through engine core 44 from compressors 24, 26 to turbines 34, 36 to define a flow direction 76 through engine core 44. The method includes generating a shock wave 90 in the working air flow propagating in flow direction 76. In one example, a shock wave 90 can be formed by rotating a set of airfoils in one of the compressors or turbines (such as LP compressor blades 56, HP compressor blades 58, HP turbine blades 68 or LP turbine blades 70( Figure 1 ))). Additionally or alternatively, shock wave 90 can be formed by supersonic expansion of the working air flow through adjacent blades (including HP turbine blades 72 or LP turbine blades 74( Figure 1 ))). The method further includes directing shock wave 90 onto at least one attenuation structure 110, 210, 310, 410, 510, 610, 710, 810 on at least one component (such as airfoils 100, 200A, 200B, 300, 400, 500, 600, 700, 800) within engine 10 to at least partially attenuate shock wave 90. The method can further include directing shock wave 90 between two circumferentially spaced airfoils (such as airfoils 200A, 200B( Figure 6 ))). The method can further include directing a reflected portion 91 of shock wave 90 onto a second attenuation structure 110, 210, 310, 410, 510, 610, 710, 810 on a second component (such as airfoils 100, 200A, 200B, 300, 400, 500, 600, 700, 800) within engine 10 to attenuate reflected portion 91. The method can include flowing a cooling air flow (such as cooling air 447) from the interior of at least one component (such as airfoil 400) into flow path 75 in a direction opposite to flow direction 76( Figure 9 - 10 ).

[0091] Some specific operating examples will be described below. It should be understood that such examples are intended to illustrate aspects of the present disclosure and do not limit the present disclosure in any way.

[0092] In one example, a shock wave can be generated by the high-speed rotation of an HP turbine blade. The shock wave can propagate along a flow path, pass through a turbine center frame, and encounter a downstream engine component in the form of an LP turbine vane that has a plurality of attenuation structures in the form of scaled structures on a pressure side and a suction side. The shock wave can form a reflected portion that travels upstream from a leading edge of the LP turbine vane and a transmitted portion that travels on the LP turbine vane and downstream. Due to the attenuation structures, the amplitude of the transmitted portion of the shock wave can be reduced by 50%. The reflected portion can travel upstream, return through the turbine center frame, and encounter a second component in the form of an airfoil in the turbine center frame that has a second set of attenuation structures in the form of a ramp structure that has a ramp surface facing the reflected portion. The reflected portion of the shock wave can travel on the ramp structure, which serves to attenuate the reflected portion of the shock wave. In this way, attenuation structures can be utilized on engine components to attenuate shock waves emanating from both upstream and downstream directions.

[0093] In another example, based on the type of shock wave most likely to be encountered at that location, the airfoil in the turbine center frame can have multiple types of attenuation structures. More specifically, the pressure side of the airfoil and the platform at the airfoil root can include a plurality of resonant cavities that are tuned to attenuate specific harmonics present from shock waves traveling downstream, while the suction side of the airfoil and the adjacent platform portion can include a plurality of raised panels to attenuate reflected shock waves from other downstream components.

[0094] Aspects of the present disclosure can be used to mitigate the effects of shock waves on downstream engine components, including by using local surface features or protrusions. Shock waves can be generated by the rotation of upstream blades or by supersonic expansion through upstream nozzles. Regardless of how the shock wave is formed, the attenuation structures described herein can disperse such shock waves by reducing primary reflections, reducing harmonic reflections, removing energy from the wavefront, using surface or counterflow cooling air to disrupt the wavefront, etc.

[0095] Multiple benefits can be realized from the present disclosure. One advantage is that the use of attenuation structures provides a reduction in the pressure gradient within the working air flow to minimize losses and mixing losses between upstream and downstream waves. Another benefit is the dispersion or "smearing" of the shock wave gradient and a reduction in the intensity of the shock wave traveling in the flow path, which provides more work extracted from the working air flow and a longer component operating life. Another benefit is a reduction in the amplitude of unsteady blade loading, thereby minimizing unsteady losses on the airfoil surface due to shock pressure gradients. Still another benefit is the reduction of shock-wake interactions via diffusion of the shock wave through the attenuation structures such that the reflected shock wave traveling upstream mixes with the shock wave traveling downstream with lower losses.

[0096] Within the scope not yet described, the different features and structures of the various aspects can be used in combination as needed or substituted for each other. The fact that a feature is not described in all examples does not mean that it cannot be so described, but rather this is done for the sake of brevity of description. Thus, the various features of the different aspects can 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.

[0097] This written description uses examples to describe the aspects of the disclosure described herein, including the best mode, and also enables any person skilled in the art to practice the aspects of the disclosure, including making and using any device or system and performing any incorporated method. The patentable scope of the aspects of the disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. If these other examples have structural elements that are indistinguishable from the literal language of the claims, or if they include equivalent structural elements that do not differ materially from the literal language of the claims, then these other examples are intended to fall within the scope of the claims.

[0098] Further aspects of the invention are provided by the subject matter of the following clauses:

[0099] 1. A method of operating a turbine engine having an engine core, the engine core including a compressor, a combustor, and a turbine arranged in an axial flow configuration such that a working gas flow passes from the compressor through the engine core to the turbine to define a flow direction through the engine core, the method comprising: generating a shock wave in the working gas flow propagating in the flow direction; and directing the shock wave onto at least one attenuation structure on at least one component within the turbine engine to at least partially attenuate the shock wave.

[0100] 2. The method according to any of the preceding clauses, wherein the at least one component includes at least one of an airfoil, an inner band, or an outer band.

[0101] 3. The method according to any of the preceding clauses, wherein the at least one attenuation structure includes at least one of a scaled structure, a raised panel, a raised ridge, a resonant cavity, or an arcuate surface portion.

[0102] 4. The method according to any of the preceding clauses, wherein the at least one component includes an airfoil having an outer surface that defines a pressure side and a suction side and extends between a leading edge and a trailing edge.

[0103] 5. The method according to any of the preceding clauses, wherein the at least one attenuation structure is located on the pressure side of the airfoil.

[0104] 6. The method according to any of the preceding clauses, wherein the at least one component further comprises a turbine center frame having the airfoil.

[0105] 7. The method according to any of the preceding clauses, further comprising flowing a cooling air stream from the interior of the at least one component into the working air stream in a direction opposite to the flow direction.

[0106] 8. The method according to any of the preceding clauses, further comprising directing a reflected portion of the shock wave onto a second attenuation structure on a second component to attenuate the reflected portion.

[0107] 9. The method according to any of the preceding clauses, wherein the second attenuation structure comprises one of a scaled structure, a raised panel, a raised ridge, a resonant cavity, or an arcuate surface portion.

[0108] 10. The method according to any of the preceding clauses, wherein the at least one attenuation structure comprises a resonant cavity having an inlet on an outer surface of the at least one component and in fluid communication with the working air stream.

[0109] 11. The method according to any of the preceding clauses, wherein the at least one component comprises two circumferentially spaced airfoils, and wherein the directing further comprises directing the shock wave between the two circumferentially spaced airfoils.

[0110] 12. The method according to any of the preceding clauses, wherein the shock wave comprises a pressure wave within a subsonic flow.

[0111] 13. A turbine engine, comprising: an engine core including a compressor, a combustor, and a turbine arranged in an axial flow configuration; a flow path extending from the compressor through the engine core to the turbine to define a flow direction of a working air stream through the engine core; a rotatable set of airfoils in one of the compressor or the turbine such that rotation of the set of airfoils generates a shock wave in the working air stream; and at least one component in the flow path, the at least one component including an outer surface having at least one attenuation structure configured to at least partially attenuate the shock wave flowing thereover.

[0112] 14. The turbine engine according to any of the preceding clauses, wherein the at least one attenuation structure comprises at least one of a scaled structure, a raised panel, a raised ridge, a resonant cavity, or an arcuate surface portion.

[0113] 15. A turbine engine according to any of the preceding clauses, wherein the at least one damping structure projects from the outer surface into the flow path.

[0114] 16. A turbine engine according to any of the preceding clauses, further comprising a second component including a second damping structure configured to damp reflected shock waves flowing thereon.

[0115] 17. A turbine engine according to any of the preceding clauses, wherein the at least one component includes an airfoil having the outer surface that defines a pressure side and a suction side and extends between a leading edge and a trailing edge.

[0116] 18. A turbine engine according to any of the preceding clauses, wherein the at least one damping structure includes an arcuate portion in the outer surface.

[0117] 19. A turbine engine according to any of the preceding clauses, wherein the airfoil includes a set of cooling holes having corresponding outlets that are at least partially aligned in a direction opposite to the flow direction.

[0118] 20. A turbine engine according to any of the preceding clauses, wherein the at least one damping structure includes a resonance cavity having an inlet on the outer surface of the at least one component and fluidly coupled to the flow path.

Claims

1. A turbine engine, characterized in that, Comprising: An engine core, said engine core including a compressor, a combustor, and a turbine arranged in axial flow; A flow path, said flow path extending from said compressor through said engine core to said turbine to define a flow direction of the working air flow through said engine core; A component, said component including an outer surface having at least one attenuation structure, said at least one attenuation structure protruding into said flow path; And A rotatable set of airfoils, said rotatable set of airfoils being in one of said compressor or said turbine, whereby rotation of said set of airfoils forms said working air flow having shock waves, said shock waves propagating to said outer surface of said component and being at least partially attenuated by said at least one attenuation structure; Wherein said at least one attenuation structure includes at least one of a set of raised ridges in said outer surface or a set of raised panels in said outer surface, and at least one of said set of raised ridges or said set of raised panels is arranged to form a set of honeycomb structures.

2. The turbine engine according to claim 1, characterized in that, Wherein said at least one attenuation structure includes at least one of a scaling structure, a ramp structure, a bending structure, or a bow-shaped wall portion.

3. The turbine engine according to claim 1, characterized in that, Wherein said at least one attenuation structure includes a set of scaling structures, said set of scaling structures covering at least a portion of said outer surface and protruding into said flow path.

4. A method of operating a turbine engine, characterized in that, Said turbofan engine has an engine core, said engine core including a compressor, a combustor, and a turbine arranged in axial flow, whereby the working air flow passes from said compressor through said engine core to said turbine to define a flow path through said engine core, the method comprising: Generating a shock wave in said working air flow, said shock wave propagating to an outer surface of a component within said turbofan engine, said outer surface including at least one attenuation structure, said at least one attenuation structure protruding into said flow path; and Directing said shock wave onto said at least one attenuation structure to at least partially attenuate said shock wave; Wherein said at least one attenuation structure includes at least one of a set of raised ridges in said outer surface or a set of raised panels in said outer surface, and said at least one of said set of raised ridges or said set of raised panels is arranged to form a set of honeycomb structures.

5. The method according to claim 4, characterized in that, Wherein said component includes one of an airfoil, an inner band, or an outer band.

6. The method according to claim 4, characterized in that, Wherein said at least one attenuation structure includes at least one of a scaling structure, a ramp structure, a bending structure, or a bow-shaped wall portion.

7. A turbine engine, characterized in that, Comprising: An engine core, said engine core including a compressor, a combustor, and a turbine arranged in axial flow; A flow path, said flow path extending from said compressor through said engine core to said turbine to define a flow direction of the working air flow through said engine core; A first airfoil, said first airfoil including an outer wall and at least one attenuation structure, said at least one attenuation structure including a resonant cavity extending into said outer wall; And A rotatable set of airfoils, the rotatable set of airfoils being in one of the compressor or the turbine, wherein rotation of the set of airfoils forms a working air flow having shock waves that propagate to the outer surface of the first airfoil and are at least partially attenuated by the at least one attenuation structure.

8. The turbine engine according to claim 7, characterized in that, The resonator cavity includes a neck and a chamber; wherein the neck includes an opening at the outer wall; wherein the chamber is fluidly coupled to the neck; the neck volume of the neck is less than the chamber volume of the chamber; and wherein the attenuation structure includes one or more additional resonator cavities having additional necks and additional chambers, the additional necks having additional openings at the outer wall, the additional chambers being fluidly coupled to the additional necks.

9. The turbine engine according to claim 8, characterized in that, wherein the one or more additional resonator cavities include a plurality of additional resonator cavities having different additional neck volumes from each other.

10. The turbine engine according to claim 9, characterized in that, wherein the resonator cavity defines a Helmholtz resonator.

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