Rotor blade system for a turbine engine

Abstract

A rotor blade system includes a rotor and a plurality of blades coupled to the rotor, the plurality of blades arranged in an airfoil distribution pattern including one or more baseline blades and one or more intentionally mistuned blades. The one or more intentionally mistuned blades have a span from a base to a tip of the one or more intentionally mistuned blades. The one or more intentionally mistuned blades include a mistuning feature having a notch formed on a surface of the one or more intentionally mistuned blades and located near the tip to create a difference in a vibration frequency between the one or more intentionally mistuned blades and a baseline natural vibration frequency of the one or more baseline blades.

Description

Technical Field

[0001] This disclosure generally relates to rotor blades of turbine engines, and particularly to frequency-detuned rotor blades in turbine engines and turbine engines using such blades. Background Technology

[0002] Turbine engines (such as gas turbine engines) typically include a fan and a turbocharger. The fan includes fan blades. Attached Figure Description

[0003] The features and advantages of this disclosure will be apparent from the following description of various exemplary embodiments as shown in the accompanying drawings, wherein similar reference numerals generally indicate the same, functionally similar and / or structurally similar elements.

[0004] Figure 1 This is a schematic cross-sectional view of a turbine engine according to aspects of this disclosure.

[0005] Figure 2 This is a schematic diagram of a front view of a rotor blade system having a rotor and a plurality of blades mounted thereon, according to an embodiment of the present disclosure.

[0006] Figures 3A to 3D An example of a pattern of multiple blades according to an embodiment of the present disclosure is shown.

[0007] Figure 4 This is a schematic representation of a first intentionally detuned blade according to an embodiment of the present disclosure.

[0008] Figure 5 According to embodiments of this disclosure Figure 4 The cross-sectional view of the detuning feature shown by line 5-5.

[0009] Figure 6 This is a schematic representation of a first intentionally detuned blade according to an embodiment of the present disclosure, to illustrate various dimensions of the detuning feature and various dimensions of the first intentionally detuned blade. Detailed Implementation

[0010] The features, advantages, and embodiments of this disclosure will be set forth or apparent from consideration of the following detailed description, accompanying drawings, and claims. Furthermore, the following detailed description is exemplary and intended to provide further explanation, without limiting the scope of the claimed disclosure.

[0011] Various embodiments of this disclosure are discussed in detail below. While specific embodiments have been discussed, they are for illustrative purposes only. Those skilled in the art will recognize that other components and constructions can be used without departing from this disclosure.

[0012] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components.

[0013] The terms "upstream" and "downstream" refer to the relative directions of fluid flow within a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction from which the fluid flows.

[0014] Unless otherwise stated herein, the terms “connection,” “fixation,” “attachment,” “linkage,” etc., refer to direct connection, fixation, attachment, or linking, as well as indirect connection, fixation, attachment, or linking through one or more intermediate components or features.

[0015] Unless the context clearly indicates otherwise, the singular forms “a,” “one,” and “the” include plural references.

[0016] As used herein, the terms "axial" and "axially" refer to a direction and orientation that extends substantially parallel to the centerline of the turbine engine. Furthermore, the terms "radial" and "radially" refer to a direction and orientation that extends substantially perpendicular to the centerline of the turbine engine. Additionally, as used herein, the terms "circumferential" and "circumferentially" refer to a direction and orientation that extends in an arc around the centerline of the turbine engine.

[0017] As used in this article, "flutter" is the self-excited vibration of a blade caused by the interaction of structural dynamic forces and aerodynamic forces.

[0018] As used in this article, “flutter margin” is a measure of the pressure difference between the start of flutter and the operating line of the blade at a given flow rate.

[0019] As used in this article, “load shock” refers to the shock wave generated on the blade.

[0020] As used in this article, the "natural frequency" of a blade is the frequency at which the blade vibrates or resonates.

[0021] As used in this article, “modal shape” is the deformation that a blade will exhibit when vibrating at a specific natural frequency of the blade.

[0022] As used herein, “detuning,” “causing detuning,” and / or “detuned” refers to a change in the shape, size, and / or characteristics of a blade compared to the shape, size, and / or characteristics of another blade on the rotor. “Detuning,” “causing detuning,” and / or “detuned” includes altering or modifying the dynamic response to aerodynamic excitation relative to a baseline blade and / or relative to another intentionally detuned blade present on a rotor having a first intentionally detuned blade. The altered or modified dynamic response includes one or more of a shift or change in the natural frequency and / or a change in modal shape relative to the baseline blade or the other intentionally detuned blade.

[0023] As used herein, a "baseline" blade is a blade that includes baseline shape, baseline size, and / or baseline characteristics. A baseline blade includes baseline natural frequencies and / or baseline modal shapes. Thus, a baseline blade is not intentionally detuned and does not include intentionally detuned features.

[0024] As used herein, an "intentionally detuned blade" is a blade designed and configured to be detuned relative to a baseline blade and / or relative to another intentionally detuned blade. For example, an intentionally detuned blade is designed and configured to have a dynamic response to altered or modified aerodynamic excitations relative to its baseline blade and / or another intentionally detuned blade present on a rotor having an intentionally detuned blade. The altered or modified dynamic response includes one or more of the following: a shift or change in the natural frequency and / or modal shape relative to the baseline natural frequency and / or modal shape of the baseline blade or the natural frequency and / or modal shape of the other intentionally detuned blade.

[0025] As used herein, “intentional detuning feature” is the shape, size, and / or characteristic of an intentionally detuned blade that differs from the baseline shape, baseline size, and / or baseline characteristic of a baseline blade and / or the shape, size, and / or characteristic of another intentionally detuned blade. The intentionally detuned blade is formed and designed to offset or change the natural frequency and / or modal shape of the intentionally detuned blade relative to the baseline natural frequency and / or baseline modal shape or the natural frequency and / or modal shape of another intentionally detuned blade.

[0026] Scope limitations are combined and interchanged herein and throughout the specification and claims. Unless the context or language otherwise indicates, such scopes are identified and include all subscopes contained herein. For example, all scopes disclosed herein include endpoints, and endpoints may be combined independently of each other.

[0027] A turbine engine, such as a gas turbine engine, typically includes a fan and a turbocharger arranged in flow communication with each other, wherein the core is located downstream of the fan in the direction of flow through the turbine engine. The core of the turbine engine typically includes a compressor section, a combustion section, a turbine section, and an exhaust section in a sequential flow order. Using a multi-shaft gas turbine engine, the compressor section may include a high-pressure compressor (HPC) located downstream of a low-pressure compressor (LPC), and the turbine section may similarly include a low-pressure turbine (LPT) located downstream of a high-pressure turbine (HPT). With this configuration, the HPC is coupled to the HPT via a high-pressure shaft (HPS), and the LPC is coupled to the LPT via a low-pressure shaft (LPS). Each section of the turbine engine, including the fan, HPC, LPC, HPT, and LPT, includes a rotor and multiple blades coupled to the rotor.

[0028] Turbo engines include parameters ("AN") 2 ”), which is equal to the square of the annular middle region (A) along the rotor blade and the blade rotation speed (N). 2 The product of ) . Current engine design strives for a larger AN 2 Value. Larger AN 2 The value indicates larger rotor blades and greater power output. However, a larger AN... 2 Combining high values ​​with slender airfoil designs can introduce rotor instabilities, including flutter. Flutter can occur in stationary airfoils (e.g., wheel blades) or rotating airfoils (e.g., blades) of turbine engines, fans, turbochargers, compressors, or turbines. Flutter is a self-excited vibration of blades caused by the interaction of structural dynamic forces and aerodynamics. Flutter can lead to high cyclic fatigue (HCF) or even blade loss. The phase difference between blades when they vibrate can cause flutter. For example, if the blades are identical, the aeroelastic mode (the connected structure and aerodynamic system) is the mode of blade vibration with a constant phase angle between adjacent blades. Each aeroelastic mode has a different inter-blade phase angle. The inter-blade phase angle affects the phase between the local unstable fluid flow through the blade and the local blade motion, which in turn affects the unstable aerodynamic work acting on the blade. Opposite phase angles result in positive work being done on the blade, which leads to flutter. Flutter is usually associated with a mode shape of the blade and typically occurs at the blade's natural frequency, and can result in persistent blade vibration.

[0029] Flutter can occur at both subsonic and supersonic speeds, depending on the flow conditions at the fan inlet. Subsonic flutter typically occurs at approximately 50% to 80% of the rotor's operating corrected speed. The rotor's corrected speed is the high-equivalent speed at sea level under ambient conditions. When a turbine engine operates within the subsonic and / or supersonic flutter corrected speed ranges, shocks, also known as load shocks, are generated on the suction side of the blades and relatively close to the leading edge (LE). For example, a shock can occur at approximately 20% to 30% of the LE from the blade. The shock can cause the blades to vibrate with harmless amplitudes due to random forces available in the system. When the blades vibrate, the shock generates disturbed, unstable pressures and produces aerodynamic work, as detailed above. Positive energy can be added to the blades, causing them to vibrate with larger amplitudes, which can lead to self-excited vibrations. The shock (e.g., unstable pressures) can propagate circumferentially around the rotor and can propagate upstream and / or downstream of the blades. The various embodiments described herein and illustrated in the figures relate to mitigating the risk of flutter in engines.

[0030] Intentional detuning is a common method used to mitigate uncontrolled vibration amplification caused by random, uncontrolled detuning by introducing a more dominant mode than random detuning. In one embodiment, intentional detuning can be introduced by creating a frequency difference between a row of blades. This frequency difference is created by altering the airfoil geometry by modifying the blade tips. The process of modifying the blade tips is referred to as "groove tip modification." Using the groove tip depth as a geometry modification parameter offers several benefits, including ease of construction and minimal impact on manufacturing during final machining operations. Furthermore, since the blade construction is modified by removing material from the blade tips, it may not be necessary to change the forging dies used to manufacture the blades. Modal frequencies of models with various groove tip depths were compared. For example, in the embodiments, the modal frequency comparisons show that varying levels of modification to the airfoil tips can produce frequency differences.

[0031] Now refer to the attached diagram, Figure 1 This is a schematic cross-sectional view of a turbine engine 10 according to an embodiment of the present disclosure. Figure 1 As shown, the turbine engine 10 defines an axial direction A (extending parallel to a longitudinal centerline axis 12 provided for reference) and a radial direction R orthogonal to the axial direction A. Typically, the turbine engine 10 includes a fan section 14 and a turbocharger engine 16 disposed downstream of the fan section 14.

[0032] The depicted turbocharged engine 16 generally includes a housing 18, which is substantially tubular and defines an annular inlet 20. (As...) Figure 1As schematically shown, the housing 18 surrounds a compressor section 21 in a series flow relationship, which includes a supercharger or low-pressure (LP) compressor 22, downstream of which follows a high-pressure (HP) compressor 24; a combustion section 26; a turbine section 27, which includes a high-pressure (HP) turbine 28, downstream of which follows a low-pressure (LP) turbine 30; and an exhaust nozzle section 32. A high-pressure (HP) shaft or spool 34 drivesly connects the HP turbine 28 to the HP compressor 24 so that the HP turbine 28 and the HP compressor rotate in unison. A low-pressure (LP) shaft 36 drivesly connects the LP turbine 30 to the LP compressor 22 so that the LP turbine 30 and the LP compressor 22 rotate in unison. The compressor section 21, combustion section 26, turbine section 27, and exhaust nozzle section 32 together define a core airflow path.

[0033] for Figure 1 In the depicted embodiment, fan section 14 includes a fan 38 (e.g., a variable pitch fan) having a plurality of fan blades 40 spaced apart and coupled to disk 42. Figure 1 As depicted, fan blades 40 extend outward from disk 42 generally along a radial direction R. Each fan blade 40 is operably coupled to an actuation member 44, which is configured to uniformly and collectively change the pitch of the fan blades 40, and each fan blade 40 is rotatable relative to disk 42 about a pitch axis P. The fan blades 40, disk 42, and actuation member 44 are rotatable together via a fan shaft 45 about a longitudinal centerline axis 12, which is powered by an LP shaft 36 across a power gearbox 46. The power gearbox 46 includes multiple gears for adjusting the rotational speed of the fan shaft 45, and thus the rotational speed of the fan 38 relative to the LP shaft 36, to a more efficient fan speed.

[0034] Still referencing Figure 1 In an exemplary embodiment, the disk 42 is covered by a rotatable fan hub 48 having an aerodynamic profile to facilitate airflow through a plurality of fan blades 40. Furthermore, the fan section 14 includes an annular fan housing or nacelle 50 circumferentially surrounding at least a portion of the fan 38 and / or the turbocharged engine 16. The nacelle 50 is supported relative to the turbocharged engine 16 by a plurality of circumferentially spaced outlet guide vanes 52. Additionally, a downstream section 54 of the nacelle 50 extends above an outer portion of the turbocharged engine 16 to define a bypass airflow passage 56 therebetween.

[0035] During operation of the turbine engine 10, a volume of air 58 enters the turbine engine 10 through the inlet 60 of the nacelle 50 and / or fan section 14. As the volume of air 58 passes through the fan blades 40, a first portion 62 of the air is directed or directed into the bypass airflow passage 56, and a second portion 64 of the air is directed or directed into the upstream section of the core airflow path, or more specifically, into the annular inlet 20 of the LP compressor 22. The ratio between the first portion 62 and the second portion 64 of the air is commonly referred to as the bypass ratio. The second portion 64 of the air is then directed through the HP compressor 24 and into the combustion section 26, where the pressure of the second portion 64 increases, and the high-pressure air is mixed with fuel and burned to provide combustion gases 66.

[0036] Combustion gas 66 is directed into and expanded by the HP turbine 28. In the HP turbine 28, a portion of the thermal and / or kinetic energy from the combustion gas 66 is extracted via a series of stages consisting of HP turbine stator blades 68 connected to the housing 18 and HP turbine rotor blades 70 connected to the HP shaft or spool 34, causing the HP shaft or spool 34 to rotate, thereby supporting the operation of the HP compressor 24. Combustion gas 66 is then directed into and expanded by the LP turbine 30. Here, a second portion of the thermal and kinetic energy is extracted from the combustion gas 66 via a series of stages consisting of LP turbine stator blades 72 connected to the housing 18 and LP turbine rotor blades 74 connected to the LP shaft 36, causing the LP shaft 36 to rotate. This supports the operation of the LP compressor 22 and, via the power gearbox 46, supports the rotation of the fan 38.

[0037] Subsequently, the combustion gases 66 are directed through the exhaust nozzle section 32 of the turbocharged engine 16 to provide propulsive thrust. Simultaneously, as a first portion of air 62 is directed through the bypass airflow passage 56 before exiting from the fan nozzle exhaust section 76 of the turbocharged engine 10, the pressure of the first portion of air 62 increases significantly, also providing propulsive thrust. The HP turbine 28, LP turbine 30, and exhaust nozzle section 32 at least partially define a hot gas path 78 for directing the combustion gases 66 through the turbocharged engine 16.

[0038] Figure 1The turbine engine 10 depicted herein is merely an example. In other exemplary embodiments, the turbine engine 10 may have any other suitable configuration. For example, in other exemplary embodiments, the fan 38 may be configured in any other suitable manner (e.g., as a fixed-pitch fan), and further may be supported using any other suitable fan frame configuration. Furthermore, in other exemplary embodiments, any other suitable number or configuration of compressors, turbines, shafts, or combinations thereof may be provided. In other exemplary embodiments, aspects of this disclosure may be incorporated into any other suitable gas turbine engine, such as, for example, a turbofan engine, a propeller fan engine, a turbojet engine, and / or a turboshaft engine.

[0039] Figure 2 This is a schematic front view of a rotor blade system 100 having a rotor 102 and a plurality of blades 104 mounted thereon, according to an embodiment of the present disclosure. The rotor blade system 100 depicted herein is a fan 38 ( Figure 1 ), multiple blades 104 are fan blades 40 ( Figure 1 ), and the rotor 102 is a rotatable fan hub 48 ( Figure 1 In some examples, the rotor blade system 100 is used in the LP compressor 22. Figure 1 ) or HP compressor 24 ( Figure 1 However, the rotor blade system 100 described herein can also be used in a turbine engine 10. Figure 1 In any system. For example, the rotor blade system 100 described herein can be used in the LP turbine 30 of a turbine engine 10 ( Figure 1 ) and / or HP Turbo 28 ( Figure 1 The rotor blade system 100 can also be applied to industrial gas turbines (IGT) or power generation turbines. This rotor blade system is not limited to turbine engines, but can be used in any airfoil system (blade system) where the objective is to reduce flutter.

[0040] The rotor blade system 100 has a plurality of blades 104 mounted to a rotor 102. In some examples, the rotor 102 and the plurality of blades 104 form an integrated component, also referred to as an integral bladed disk. In some examples, the plurality of blades 104 are inserted into the rotor 102 such that the plurality of blades 104 and the rotor are separate components. In embodiments, such as Figure 2As shown, the blades 104 may be circumferentially spaced evenly around the circumference C of the rotor 102. However, in another embodiment, the blades 104 may be circumferentially spaced non-uniformly around the circumference C of the rotor 102. The plurality of blades 104 rotates as the rotor 102 rotates. The plurality of blades 104 includes a baseline blade 104A and a first intentionally detuned blade 104B. In some examples, the plurality of blades 104 also includes a second intentionally detuned blade 104C, as detailed below.

[0041] Baseline blade 104A is a blade 104 that includes baseline shape, baseline size, and / or baseline characteristics. Therefore, baseline blade 104A is a blade 104 that does not include mistuning characteristics. The first intentionally mistuned blade 104B and the second intentionally mistuned blade 104C are blades having mistuning characteristics 402 (in...). Figure 4 The blade 104 (shown in the diagram) is further detailed below. For example, the detuning feature 402 alters the natural frequency and / or modal shape of the intentionally detuned blade relative to the natural frequency and / or modal shape of the baseline blade (e.g., baseline natural frequency and / or baseline modal shape) and / or the natural frequency and / or modal shape of another intentionally detuned blade. In this way, the first intentionally detuned blade 104B and the second intentionally detuned blade 104C provide aerodynamic detuning, natural frequency detuning, and / or modal shape detuning relative to the baseline blade 104A or relative to each other to disrupt self-excited fluid structure interactions, as further detailed below. This provides the ability to intentionally detun the first intentionally detuned blade 104B and the second intentionally detuned blade 104C relative to the baseline blade 104A in order to, for example, control flutter. The detuning feature 402 (in the diagram) on the respective first intentionally detuned blade 104B or the respective second intentionally detuned blade 104C can be selected. Figure 4 (As shown in the diagram) and the location of the detuning feature 402, such that the natural frequency and / or modal shape of the first intentionally detuned blade 104B and the natural frequency or modal shape of the second intentionally detuned blade 104C can be detuned relative to the baseline natural frequency and / or baseline modal shape of the baseline blade 104A, thereby controlling and mitigating flutter. The detuning feature 402 can be selected from one or more detuning features, as further detailed below, to disrupt the self-excited fluid-structure interaction.

[0042] Baseline blade 104A extends from its base 106A to its tip 108A. Base 106A is coupled to platform 103A, and platform 103A is mounted to rotor 102. Similarly, a first intentionally detuned blade 104B extends from its base 106B to its tip 108B. Base 106B is coupled to platform 103B, and platform 103B is mounted to rotor 102. A second intentionally detuned blade 104C extends from its base 106C to its tip 108C. Base 106C is coupled to platform 103C, and platform 103C is mounted to rotor 102.

[0043] Figures 3A to 3D An example of a pattern of blade 104 according to an embodiment of the present disclosure is shown. In the embodiment, the number and pattern of the baseline blade 104A and the first intentionally detuned blade 104B can be varied to allow for flexibility in tuning the natural frequency and / or modal shape. For example, the pattern includes an airfoil distribution pattern for mitigating flutter in the first intentionally detuned blade 104B. The airfoil distribution pattern includes a circumferential distribution of the baseline blade 104A and the first intentionally detuned blade 104B on the rotor 102.

[0044] Figures 3A to 3D The embodiments described herein involve an alternating pattern of baseline blades and intentionally detuned blades. In this way, the intentionally detuned blades disrupt the self-excited fluid-structure interaction. Therefore, embodiments of this disclosure can relate to mitigating the flutter risk of turbofan engines. However, this disclosure can relate to mitigating the flutter risk of any type of turbofan engine. Embodiments of this disclosure seek to provide a way to intentionally detune blades in order to passively control flutter. For example, the embodiments described herein increase the flutter margin of intentionally detuned blades.

[0045] exist Figure 3A In one embodiment, the airfoil distribution pattern includes a first pattern P1. The first pattern P1 includes an "ABAB" pattern, where "A" is the baseline blade 104A and "B" is a first intentionally detuned blade 104B with detuning characteristics. Figure 3A The embodiment includes an alternating arrangement of a baseline blade 104A and a first intentionally detuned blade 104B. As detailed below, the first intentionally detuned blade 104B is detuned from the baseline blade 104A to disrupt the self-excited fluid structure interaction. The first intentionally detuned blade 104B may include reduced aerodynamic performance compared to the baseline blade 104A to achieve intentional detuning. The “ABAB” mode allows for a balance between the natural frequencies and / or modal shapes of the separated baseline blade 104A and the first intentionally detuned blade 104B, and the aerodynamic performance of the rotor blade system 100.

[0046] exist Figure 3BIn one embodiment, the airfoil distribution pattern includes a second pattern P2. The second pattern P2 includes an "AABAAB" pattern, in which a first intentionally detuned blade 104B is positioned between two consecutive baseline blades 104A. Compared to the "ABAB" pattern, the "AABAAB" pattern provides increased aerodynamic performance but reduces flutter mitigation. For example, the more baseline blades 104A there are, the higher the aerodynamic performance of the rotor blade system 100. However, the fewer the first intentionally detuned blades 104B, the less mechanical damping is provided to mitigate flutter.

[0047] exist Figure 3C In one embodiment, the airfoil distribution pattern includes a third pattern P3. The third pattern P3 includes an "AAABAAAB" pattern, in which a first intentionally detuned blade 104B is positioned between three consecutive baseline blades 104A. Compared to the "ABAB" and "AABAAB" patterns, the "AAABAAAB" pattern provides increased aerodynamic performance but reduces flutter mitigation.

[0048] exist Figure 3D In one embodiment, the airfoil distribution pattern includes a fourth pattern P4. The fourth pattern P4 includes an "ABC" pattern, where "B" is a first intentionally detuned blade 104B with a first detuning characteristic, and "C" is a second intentionally detuned blade 104C with a second detuning characteristic different from the first detuning characteristic (e.g., different detuning characteristics and / or different positions of the detuning characteristics on the second intentionally detuned blade 104C compared to the first intentionally detuned blade 104B). For example, compared to the first airfoil distribution pattern P1, the second airfoil distribution pattern P2, and the third airfoil distribution pattern P3, the "ABC" pattern provides increased flutter mitigation. Furthermore, the second intentionally detuned blade 104C provides additional degrees of freedom to improve the tuning of the blade pattern to a specific flutter mode. In some examples, other patterns of the airfoil distribution pattern may include the second intentionally detuned blade 104C (e.g., in an "AABC" pattern, an "AABBC" pattern, an "ABCC" pattern, etc.). The airfoil distribution pattern is selected from a group consisting of P1, P2, P3, P4, or combinations thereof. For example, the airfoil distribution pattern is selected from a group consisting of repeating patterns of AB, AAB, AAAB, ABC, AABC, AABBC, ABCC, or combinations thereof.

[0049] When developing gas turbine engines, the interactions between components can make selecting or developing a component particularly difficult during engine design and prototyping, especially when some components are at different stages of completion. For example, one or more components may be nearly complete, while one or more other components may be in the initial or preliminary stages, making only one (or a few) design parameters known. We aim to achieve design possibilities early in the design process, making the downstream selection of candidate improved designs more predictable given trade-offs. To date, this process has sometimes been more ad hoc, selecting one design or another without understanding the impact when the concept was first considered. For example, and referencing... Figure 1 The design of fan section 14 (e.g., fan 38 design, fan blade 40 design, etc.), combustion section 26 design, compressor section 21 design, turbine section 27 design, etc., may be unknown, but such components affect the aerodynamic performance of the engine and may therefore affect the design of fan blade 40.

[0050] We aim to narrow down the range of constructions or combinations of features that can produce favorable results, given constraints on design, feasibility, manufacturing, and certification requirements. For example, we seek to achieve a more favorable balance between mitigating flutter and improving aerodynamic performance, i.e., increasing efficiency in the conversion of kinetic energy in the fluid flow to mechanical energy in the turbine shaft. We also aim to make choices earlier in the design selection process to avoid wasting time and effort. In evaluating the different embodiments described herein, we inventors unexpectedly discovered a relationship between the detuning mode and the detuning characteristics of the corresponding blades. This relationship uniquely identifies a limited and readily identifiable number of advantageous embodiments (according to this disclosure) of specific architectures suitable for addressing the flutter and instability pressures experienced by fan blades.

[0051] The airfoil distribution pattern can be selected based on improved flutter mitigation (e.g., Figures 3A to 3D (Any mode described herein) without affecting the aerodynamic performance of the rotor blade system 100. For example, the first intentionally detuned blade 104B disrupts the self-excited fluid structure interaction without affecting the aerodynamic performance.

[0052] Figure 4 This is a schematic representation of a first intentionally mistuned blade 104B according to an embodiment of the present disclosure. Although reference has been made to the first intentionally mistuned blade 104B, the embodiments described herein are applicable to a second intentionally mistuned blade 104C. Figure 4As shown, the first intentionally detuned blade 104B is defined by its span in the span direction S and its chord direction C. The span of the first intentionally detuned blade 104B is defined from its base 106B to its tip 108B. The chord C of the first intentionally detuned blade 104B is defined from its leading edge 109 (LE) to its trailing edge 111 (TE). In some examples, the chord C of the first intentionally detuned blade 104B varies along its span.

[0053] exist Figure 4 In one embodiment, the first intentionally detuned blade 104B includes a detuning feature 402. As further detailed below, intentional detuning is achieved by one or more detuning features for the first intentionally detuned blade 104B relative to the baseline blade 104A (in... Figure 2 The baseline natural frequency and / or baseline mode shape are detuned (as shown in the diagram). For example, detuning feature 402 can control the first intentionally detuned blade 104B relative to the baseline blade 104A (shown in the diagram). Figure 2 The baseline natural vibrational frequency and / or baseline modal shape of the vibrational frequency and / or modal shape.

[0054] The detuning feature 402 can be located at a spanwise position S1 on the first intentionally detuned blade 104B. For example, the spanwise position S1 of the detuning feature 402 can be measured from the base 106B of the first intentionally detuned blade 104B to the edge of the detuning feature 402, which is located approximately at the midpoint 402M of the detuning feature 402. The position of the detuning feature 402 can also be measured by a depth dimension D1, which is measured from the tip 108B of the first intentionally detuned blade 104B to the edge of the detuning feature 402, which is located approximately at the midpoint 402M of the detuning feature 402. For example, the depth dimension D1 measured from the tip 108B to the edge of the detuning feature 402 can be 0% to 40% of the total wingspan S from the tip 108B to the base 106B of the first intentionally detuned blade 104B. The ratio of D1 / S is between 0% and 40%.

[0055] Therefore, the position of the mistuning feature 402 on the first intentionally mistuned blade 104B is defined by the depth dimension D1 in the radial direction. The depth dimension D1 is a percentage of the wingspan S of the first intentionally mistuned blade 104B. The position of the mistuning feature 402 on the first intentionally mistuned blade 104B is also defined by the width dimension C1 in the direction of the chord C of the first intentionally mistuned blade 104B. The width dimension C1 is a percentage of the chord C of the first intentionally mistuned blade 104B.

[0056] In one embodiment, the first end 402A of the detuning feature 402 is located at a first distance C2 from the leading edge 109 (LE) of the first intentionally detuned blade 104B, and the second end 402B of the detuning feature 402 is located at a second distance C3 from the trailing edge 111 (TE) of the first intentionally detuned blade 104B. In this embodiment, the detuning width dimension C1 can be centered within the chord C of the first intentionally detuned blade 104B, such that, for example, the first distance C2 of the first end 402A of the detuning feature 402 relative to the leading edge 109 is substantially equal to the second distance C3 of the second end 402B of the detuning feature 402 relative to the trailing edge 111.

[0057] In one embodiment, the detuning feature 402 has a thickness dimension T1. The thickness dimension T1 is perpendicular to the surface 402S of the first intentionally detuned blade 104B. In this embodiment, the thickness dimension T1 of the detuning feature 402 corresponds to the material removed from the first intentionally detuned blade 104B. Therefore, in this embodiment, the detuning feature 402 includes a notch 403 formed by removing or cutting material from the first intentionally detuned blade 104B. The detuning feature 402 is located near the tip 108B of the first intentionally detuned blade 104B. In this embodiment, the thickness dimension T1 of the detuning feature 402 may be the full thickness T(t) of the first intentionally detuned blade 104B. Figure 6 The total thickness T is 0% to 40% of the surface of the first intentionally detuned blade 104B. The total thickness T is measured from the surface of the first intentionally detuned blade 104B to the opposite surface of the first intentionally detuned blade 104B. In this embodiment, the ratio of T1 / T is 0% to 40%.

[0058] Figure 5 According to embodiments of this disclosure Figure 4 The cross-section of the detuned feature 402 (including the notch 403) is shown by line 5-5. Figure 5 The outline of the thickness dimension T1 along the width dimension C1 is shown. In one embodiment, the thickness dimension T1 along the width dimension C1 can be substantially constant. However, in another embodiment, as... Figure 5 As shown, the thickness dimension T1 varies along the width dimension C1. Therefore, the mistuning feature 402 can have a thickness dimension profile. The profile of the thickness dimension T1 and the width dimension C1 can be inclined or curved (e.g., Figure 5 (as shown), serrated, wavy, or any combination thereof, or any other desired shape.

[0059] Select the type and position of the detuning feature 402 on the first intentionally detuned blade 104B to change one or more natural frequencies and / or associated modal shapes of the first intentionally detuned blade 104B relative to the baseline natural frequency and / or baseline modal shape of the baseline blade 104A.

[0060] For example, removing mass or material from the first intentionally detuned blade 104B (e.g., cutting material to form the detuning feature 402 or notch 403) alters the natural frequency of the first intentionally detuned blade 104B relative to the baseline natural frequency of the baseline blade 104A. Changing the position of the detuning feature 402 on the first intentionally detuned blade 104B (such as the depth dimension D1 in the span direction and / or changing the width dimension C1 in the chord direction) alters the natural frequency and / or modal shape of the first intentionally detuned blade 104B relative to the baseline blade 104A. In an embodiment, the width dimension C1 is approximately 0% to 85% (e.g., 40% to 60%) of the chord C of the first intentionally detuned blade 104B. The chord C of the first intentionally detuned blade 104B is equal to the sum of the width dimension C1, the first distance C2, and the second distance C3 (i.e., C = C1 + C2 + C3). The ratio of C1 / C is between 0% and 85%.

[0061] In this embodiment, in order of frequency tuning sensitivity, the depth dimension D1 can be a first or main control parameter for adjusting the vibration frequency of the first intentionally detuned blade 104B. The second control parameter for adjusting the vibration frequency of the first intentionally detuned blade 104B can be the thickness dimension T1. The third control parameter for adjusting the vibration frequency of the first intentionally detuned blade 104B can be the width dimension C1.

[0062] Intentional detuning is typically used to mitigate uncontrolled vibration amplification caused by random, uncontrolled detuning by introducing a more dominant mode than random detuning. In one embodiment, intentional detuning can be introduced by creating a frequency difference between a first intentionally detuned blade and a baseline blade. This frequency difference is created by altering the airfoil geometry by modifying the blade tip. The process of modifying the blade tip is referred to as "groove tip modification." Using the groove tip depth as a geometry modification parameter offers several advantages, including ease of construction and minimal impact on manufacturing during final machining operations. Furthermore, since the blade construction is modified by removing material from the blade tip, it may not be necessary to change the forging die used to manufacture the blade. Modal frequencies of models with various groove tip depths were compared.

[0063] Figure 6 This is a schematic representation of a first intentionally detuned blade 104B according to an embodiment of the present disclosure, illustrating various dimensions of the detuning feature 402 and various dimensions of the first intentionally detuned blade 104B. Parametric relationships between the geometric characteristics of the detuning feature 402 are determined such that the detuning feature 402 produces a frequency difference of at least 1% from the nominal natural frequency of a blade without the detuning feature 402. In other words, the natural frequency W of the blade having the detuning feature 402 is... sqThe ratio Ra of the frequency W of the blade without detuning characteristic 402 is greater than or equal to 1.01 (i.e., W sq / W≥1.01). This ratio R can be represented by the following expression.

[0064] (1)

[0065] Among them, quality And quality (2)

[0066] Where S is the average wingspan of the intentionally detuned blade 104B, C is the average chord dimension of the intentionally detuned blade 104B, T is the average thickness dimension of the intentionally detuned blade 104B, C1 is the average chord dimension of the detuning feature 402, T1 is the average thickness dimension of the detuning feature 402, D1 is the depth dimension of the detuning feature 402, ρ is the material density of the intentionally detuned blade 104B, and k is the stiffness of the intentionally detuned blade 104B.

[0067] By replacing the mass m and mass m in expression (1) with each expression (2) sq After removing the square root and simplifying, we get expression (3).

[0068] (3)

[0069] In other words, the ratio between the first product of the average wingspan, average chord size, and average thickness size and the difference between the first product and the second product of the average chord size, average thickness size, and depth size of the detuned feature is greater than or equal to approximately 1.02.

[0070] Therefore, the ratio Ra can be adjusted by selecting appropriate values ​​for the variables in expression (3), including the average thickness dimension T1 of the detuned feature 402, the depth dimension D1 of the detuned feature 402, and the average chord dimension C1 of the detuned feature 402.

[0071] Although the first intentionally detuned blade 104B (as described above) Figure 4 ) and the first intentionally mistuned blade 105B ( Figure 6 The first intentionally detuned blade 104B or the first intentionally detuned blade 105B is described as a rotor blade, but its use is not limited to rotor blades, but can also be used in stator blades or any airfoil that may require mitigation of uncontrolled vibration amplification.

[0072] Table 1 lists the various values ​​of the variables used in the detuned blade 104B. The ratio Ra is calculated using expression (3). 2Expression (3) takes the various set values ​​of examples 1 to 8 listed in Table 1 as input. For each example 1 to 8, Ra is computed using the given set of values ​​of the input variables S, C, C1, D1, T, and T1. 2 As shown in Table 1, the ratio Ra 2 Between 1.020 and 1.081, this satisfies the inequality. .

[0073] Table 1

[0074]

[0075] Based on the values ​​provided in Table 1, various ratios were also calculated. Table 2 lists the ratios calculated using the values ​​from Examples 1 to 8. Based on the calculated ratios, the ranges of the ratios were determined. For example, the calculated ratio D1 / S ranges between 3% and 50%, the ratio T1 / T ranges between 5% and 80%, and the ratio C1 / C ranges between 5% and 95%.

[0076] Table 2

[0077]

[0078] Further aspects are provided by the following items.

[0079] A rotor blade system includes: a rotor and a plurality of blades coupled to the rotor, the plurality of blades being arranged in an airfoil distribution pattern, the airfoil distribution pattern including one or more baseline blades and one or more intentionally detuned blades. The one or more intentionally detuned blades have an wingspan from the base of the one or more intentionally detuned blades to the tip of the one or more intentionally detuned blades and a chord defined from the leading edge of the one or more intentionally detuned blades to the trailing edge of the one or more intentionally detuned blades. The one or more intentionally detuned blades include a detuning feature comprising a notch formed on a surface of the one or more intentionally detuned blades, the detuning feature being located near the tip of the one or more intentionally detuned blades to generate a frequency difference between the vibration frequency of the one or more intentionally detuned blades and the baseline natural vibration frequency of the one or more baseline blades.

[0080] According to the rotor blade system described in the preceding clause, the detuning feature is located at a depth dimension measured from the tip of the one or more intentionally detuned blades to the midpoint of the detuning feature.

[0081] The rotor blade system according to any one of the preceding clauses, wherein the detuning feature is defined by the width dimension in the direction of the chord of the one or more intentionally detuned blades.

[0082] According to any one of the preceding clauses, in the rotor blade system, the detuning feature is defined by a thickness dimension in a direction substantially perpendicular to the surface of the one or more intentionally detuned blades, the thickness dimension corresponding to material removed from the surface of the one or more intentionally detuned blades to form the notch.

[0083] In the rotor blade system according to any one of the preceding clauses, the thickness dimension of the detuning feature is substantially constant along the width dimension of the detuning feature.

[0084] According to any one of the preceding clauses, in the rotor blade system, the thickness dimension is variable along the width dimension of the detuning feature, such that the detuning feature has a thickness dimension profile that is inclined, curved, sawtooth, wavy, or any other selected shape relative to the width dimension, the thickness dimension profile being selected to generate the vibration frequency difference between the vibration frequency of the one or more intentionally detuned blades and the baseline natural vibration frequency of the one or more baseline blades.

[0085] According to any one of the preceding clauses, in the rotor blade system, the ratio of the depth dimension to the thickness dimension is selected such that the vibration frequency of the one or more intentionally detuned blades is different from the baseline natural vibration frequency of the one or more baseline blades.

[0086] According to any one of the preceding clauses, in the rotor blade system, the ratio of the depth dimension to the thickness dimension is selected such that the frequency difference between the vibration frequency of the one or more intentionally detuned blades and the baseline natural vibration frequency of the one or more baseline blades is approximately 2% to 6% of the baseline natural vibration frequency of the one or more baseline blades.

[0087] The rotor blade system according to any one of the preceding clauses, wherein the ratio of a first product of the average span, average chord, and average thickness to the difference between the first product and a second product of the average chord, average thickness, and depth of the detuned feature is greater than or equal to about 1.02.

[0088] The rotor blade system according to any one of the preceding clauses, wherein the one or more baseline blades do not include the detuning feature.

[0089] According to any one of the preceding clauses, in the rotor blade system, the airfoil distribution pattern is selected from the group consisting of: AB, AAB, AAAB, ABC, AABC, AABBC, ABCC or any combination thereof, wherein A is a baseline blade, B is a first intentionally detuned blade having a first detuned feature, and C is a second intentionally detuned blade having a second detuned feature different from the first detuned feature.

[0090] The rotor blade system according to any one of the preceding clauses, wherein a parametric relationship between the geometric characteristics of the detuning feature is determined such that the detuning feature produces a frequency difference of at least 1% from the nominal natural frequency of the blade without the detuning feature.

[0091] In any of the preceding clauses, the ratio of the natural frequency of the intentionally detuned blade having the detuning feature to the nominal natural frequency of the blade without the detuning feature is greater than or equal to 1.01.

[0092] A turbine engine includes a fan section, a compressor section, and a turbine section, wherein at least one of the fan section, the compressor section, or the turbine section includes a rotor blade system according to any one of the preceding clauses.

[0093] According to the turbine engine described in the foregoing clause, the compressor section includes a low-pressure compressor and a high-pressure compressor, wherein at least one of the low-pressure compressor or the high-pressure compressor includes the rotor blade system.

[0094] The turbine engine according to any one of the preceding clauses, wherein the turbine section comprises a high-pressure turbine and a low-pressure turbine, wherein at least one of the high-pressure turbine or the low-pressure turbine comprises the rotor blade system.

[0095] The turbine engine according to any one of the preceding clauses further includes a high-pressure shaft that drivesly connects the high-pressure turbine to the high-pressure compressor so that the high-pressure turbine and the high-pressure compressor rotate in unison.

[0096] The turbine engine according to any one of the preceding clauses further includes a low-pressure shaft that drivesly connects the low-pressure turbine to the low-pressure compressor so that the low-pressure turbine and the low-pressure compressor rotate in unison.

[0097] A turbine engine includes a fan section, a compressor section, and a turbine section. At least one of the fan section, the compressor section, and the turbine section includes a rotor blade system. The rotor blade system includes a rotor and a plurality of blades coupled to the rotor, the plurality of blades being arranged in an airfoil distribution pattern including one or more baseline blades and one or more intentionally mistuned blades. The one or more intentionally mistuned blades have a span from the base of the one or more intentionally mistuned blades to the tip of the one or more intentionally mistuned blades and a chord defined from the leading edge of the one or more intentionally mistuned blades to the trailing edge of the one or more intentionally mistuned blades. The one or more intentionally mistuned blades include a mistuning feature comprising a notch formed on the surface of the one or more intentionally mistuned blades near the tip of the one or more intentionally mistuned blades to generate a frequency difference between the vibration frequency of the one or more intentionally mistuned blades and the baseline natural vibration frequency of the one or more baseline blades.

[0098] According to the turbine engine described in the foregoing clause, the detuning feature is located at a depth dimension measured from the tip of the one or more intentionally detuned blades to the midpoint of the detuning feature.

[0099] In any of the preceding clauses, the detuning feature of the turbine engine is defined by the width dimension in the direction of the chord of the one or more intentionally detuned blades.

[0100] According to the turbine engine described in the foregoing clause, the detuning feature is defined by a thickness dimension in a direction substantially perpendicular to the surface of the one or more intentionally detuned blades, the thickness dimension corresponding to material removed from the surface of the one or more intentionally detuned blades to form the notch.

[0101] According to the turbine engine described in the foregoing clause, the thickness dimension of the detuned feature is substantially constant along the width dimension of the detuned feature.

[0102] According to the turbine engine described in the foregoing clause, the thickness dimension is variable along the width dimension of the detuned feature, such that the detuned feature has a thickness dimension profile that is inclined, curved, serrated, wavy, or any other selected shape relative to the width dimension, the thickness dimension profile being selected to generate the vibration frequency difference between the vibration frequency of the one or more intentionally detuned blades and the baseline natural vibration frequency of the one or more baseline blades.

[0103] According to the turbine engine described in the foregoing clause, the ratio of the depth dimension to the thickness dimension is selected such that the vibration frequency of the one or more intentionally detuned blades is different from the baseline natural vibration frequency of the one or more baseline blades.

[0104] According to the turbine engine described in the foregoing clause, the ratio of a first product of the average wingspan, average chord size, and average thickness size to the difference between the first product and a second product of the average chord size of the detuned feature, the average thickness size of the detuned feature, and the depth size of the detuned feature is greater than or equal to about 1.02.

[0105] According to the turbine engine described in the foregoing clause, the one or more baseline blades do not include the detuning feature.

[0106] According to the turbine engine described in the foregoing clause, the airfoil distribution pattern is selected from the group consisting of: AB, AAB, AAAB, ABC, AABC, AABBC, ABCC or any combination thereof, wherein A is the baseline blade, B is a first intentionally detuned blade having a first detuned feature, and C is a second intentionally detuned blade having a second detuned feature different from the first detuned feature.

[0107] While the foregoing description relates to preferred embodiments of the present disclosure, other variations and modifications will be apparent to those skilled in the art and can be made without departing from the present disclosure. Furthermore, features described in connection with one embodiment of the present disclosure may be used in conjunction with other embodiments, even if not explicitly stated above.

Claims

1. A rotor blade system for a turbine engine, characterized in that, The rotor blade system includes: Rotor; and Multiple blades are connected to the rotor, the multiple blades being arranged in an airfoil distribution pattern, the airfoil distribution pattern including one or more baseline blades and one or more intentionally detuned blades. The one or more intentionally mistuned blades have a span from the base to the tip of the one or more intentionally mistuned blades and a chord defined from the leading edge to the trailing edge of the one or more intentionally mistuned blades. The one or more intentionally detuned blades include a detuning feature comprising a notch formed on the surface of the one or more intentionally detuned blades, the detuning feature being located near the tip of the one or more intentionally detuned blades to generate a vibration frequency difference between the vibration frequency of the one or more intentionally detuned blades and the baseline natural vibration frequency of the one or more baseline blades.

2. The rotor blade system according to claim 1, characterized in that, The one or more baseline blades mentioned herein do not include the detuning feature.

3. The rotor blade system according to claim 1, characterized in that, The airfoil distribution pattern is selected from the group consisting of: AB, AAB, AAAB, ABC, AABC, AABBC, ABCC or any combination thereof, where A is the baseline blade, B is a first intentionally detuned blade with a first detuning feature, and C is a second intentionally detuned blade with a second detuning feature different from the first detuning feature.

4. The rotor blade system according to claim 1, characterized in that, The detuning feature is located at a depth dimension measured from the tip of the one or more intentionally detuned blades to the midpoint of the detuning feature.

5. The rotor blade system according to claim 4, characterized in that, The detuning feature is defined by the width dimension in the direction of the chord of the one or more intentionally detuned blades.

6. The rotor blade system according to claim 5, characterized in that, The detuning feature is defined by a thickness dimension in a direction substantially perpendicular to the surface of the one or more intentionally detuned blades, the thickness dimension corresponding to the material removed from the surface of the one or more intentionally detuned blades to form the notch.

7. The rotor blade system according to claim 6, characterized in that, The thickness dimension of the detuned feature is substantially constant along the width dimension of the detuned feature.

8. The rotor blade system according to claim 6, characterized in that, The thickness dimension is variable along the width dimension of the detuned feature, such that the detuned feature has a thickness dimension profile that is inclined, curved, serrated, wavy, or any other selected shape relative to the width dimension, the thickness dimension profile being selected to generate the vibration frequency difference between the vibration frequency of the one or more intentionally detuned blades and the baseline natural vibration frequency of the one or more baseline blades.

9. The rotor blade system according to claim 6, characterized in that, The ratio of the depth dimension to the thickness dimension is selected such that the vibration frequency of the one or more intentionally detuned blades is different from the baseline natural vibration frequency of the one or more baseline blades.

10. The rotor blade system according to claim 9, characterized in that, The ratio of the first product of the average wingspan, average chord length, and average thickness to the difference between the first product and the second product of the average chord length, average thickness length, and depth length of the detuned feature is greater than or equal to approximately 1.02.

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