Circumferentially varying fan casing treatments for reducing fan noise effects

A rotary component for a gas turbine engine includes a plurality of rotor blades operably coupled to a rotating shaft extending along the central axis and an outer casing arranged exterior to the plurality of rotor blades in a radial direction of the gas turbine engine. The outer casing defines a gap between a blade tip of each of the plurality of rotor blades and the outer casing. The outer casing includes a plurality of features formed into an interior surface thereof. Each of the plurality of features includes one or more design parameters that are perturbed about a mean design parameter for stall performance so as to provide a circumferential variation in wake strengths associated with the plurality of rotor blades, thereby reducing operational noise of the gas turbine engine.

FIELD

The present disclosure relates generally to gas turbine engines, and more particularly to circumferentially varying fan casing treatments for reducing fan noise effects.

BACKGROUND

A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere. Turbofan gas turbine engines typically include a fan assembly that channels air to the core gas turbine engine, such as an inlet to the compressor section, and to a bypass duct. Gas turbine engines, such as turbofans, generally include fan casings surrounding a fan assembly including the fan blades. The compressor section typically includes one or more compressors with corresponding compressor casings.

As is known, an axial compressor for a gas turbine engine may include a number of stages arranged along an axis of the compressor. Each stage may include a rotor disk and a number of compressor blades, also referred to herein as rotor blades, arranged about a circumference of the rotor disk. In addition, each stage may further include a number of stator blades, disposed adjacent the rotor blades, and arranged about a circumference of the compressor casing. During operation of a gas turbine engine using a multi-stage axial compressor, a set of turbine rotor blades are turned at high speeds by a turbine so that air is continuously induced into the compressor. The air is accelerated by the rotating compressor blades and swept rearwards onto the adjacent rows of stator blades. Each rotor blade/stator blade stage increases the pressure of the air.

As such, the art is continuously seeking new and improved methods of reducing noise associated with various gas turbine engine components.

DETAILED DESCRIPTION

The term “at least one of” in the context of, e.g., “at least one of A, B, and C” refers to only A, only B, only C, or any combination of A, B, and C.

The term “turbomachine” refers to a machine including one or more compressors, a heat generating section (e.g., a combustion section), and one or more turbines that together generate a torque output.

The term “gas turbine engine” refers to an engine having a turbomachine as all or a portion of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, etc., as well as hybrid-electric versions of one or more of these engines.

The term “combustion section” refers to any heat addition system for a turbomachine. For example, the term combustion section may refer to a section including one or more of a deflagrative combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other appropriate heat addition assembly. In certain example embodiments, the combustion section may include an annular combustor, a can combustor, a cannular combustor, a trapped vortex combustor (TVC), or other appropriate combustion system, or combinations thereof.

The terms “low” and “high”, or their respective comparative degrees (e.g., -er, where applicable), when used with a compressor, a turbine, a shaft, or spool components, etc. each refer to relative speeds within an engine unless otherwise specified. For example, a “low turbine” or “low speed turbine” defines a component configured to operate at a rotational speed, such as a maximum allowable rotational speed, lower than a “high turbine” or “high speed turbine” of the engine.

As may be used herein, the terms “first”, “second”, “third” and so on may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

The term “adjacent” as used herein with reference to two walls and/or surfaces refers to the two walls and/or surfaces contacting one another, or the two walls and/or surfaces being separated only by one or more nonstructural layers and the two walls and/or surfaces and the one or more nonstructural layers being in a serial contact relationship (i.e., a first wall/surface contacting the one or more nonstructural layers, and the one or more nonstructural layers contacting the a second wall/surface).

Approximating language, as used herein throughout the specification and claims, is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about”, “approximately”, and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value, or the precision of the methods or machines for constructing or manufacturing the components and/or systems. For example, the approximating language may refer to being within a 1, 2, 4, 10, 15, or 20 percent margin. These approximating margins may apply to a single value, either or both endpoints defining numerical ranges, and/or the margin for ranges between endpoints.

During operation of a gas turbine engine, air passing through the various inlets may include distortion, such as pressure gradients, velocity gradients, and/or swirl or angular variations associated with the incoming flow. Furthermore, the fan assembly of the gas turbine engine may generate fan-outlet-guide-vane (OGV) interaction noise, inlet distortion/fan interaction noise, and/or fan self-noise. Such distortion may be circumferentially varying around the inlet of the fan blades or the compressor blades. Moreover, such distortion may propagate through each subsequent stage of the compressor. Such distortion may affect the stall point of the compressor and/or the fan assembly. Stall on the compressor and/or fan blades may generally reduce the efficiency of the engine. For example, fan or compressor stalls may reduce the pressure ratio and reduce the airflow, thereby adversely affecting the efficiency and/or operability range of the engine.

In particular, as the fan starts to stall, the flow in the tip region weakens, aerodynamic blockage develops and flow diverts to adjacent tip passages and the stall blockage cell rotates. This instability eventually grows into full stall. When the inlet is subjected to high crosswinds, the flow around the inlet lip can separate, which can trigger this stalling mechanism earlier than in undistorted conditions.

As such, the present disclosure is directed to tailored fan casing treatments for reducing fan noise (such as fan self-noise, fan-outlet-guide vane interaction noise, and/or fan inlet distortion-related noise). In particular, the fan casing treatments of the present disclosure are configured to reduce the noise source, with acoustic liners optionally employed in the fan duct to attenuate the noise before the noise propagates outside the engine. More specifically, it is generally known that fan operation and stall limits can be extended through the use of casing treatments. However, the present disclosure is directed to uniquely designed circumferentially varying patterns of fan casing treatments through which fan noise can be reduced as described herein. As such, the novel concept of the present disclosure can be extended to address fan inlet distortion related noise as will also be described.

In an embodiment, for example, the fan casing treatments of the present disclosure are configured to account for the azimuthal sectors that suffer greater levels of distortion to mitigate the operability and aeromechanics risks before stall dynamics begin to develop. For example, in an embodiment, the fan casing treatments may include feature(s) having design parameters (such as a slot design) arranged in a wave pattern, wherein the circumferential wave number is equal to a multiple of outlet guide vanes. In such embodiments, the design parameters may be perturbed about a mean design parameter for stall performance (e.g., to include zero mean casing treatment for specific cases where stall performance does not require mean casing treatments) so as to provide a circumferential variation in wake strengths associated with the fan blades. The design parameters may include, for example, blade overlap, passage width, slot depth, length and relative blade position, and the like, and combinations thereof. Moreover, the wave pattern of the design parameters may be clocked circumferentially relative to the OGVs to ensure the low noise wakes impinge on the leading edges of the outlet guide vanes and the high noise wakes (where “low noise wakes” and “high noise wakes” generally denote wakes with lower/higher energy content that would lead to quieter/louder fan-outlet-guide vane interaction noise, respectively, relative to the mean state) pass between the outlet guide vanes for at least one of a set of noise-sensitive operating condition(s).

In another embodiment, the feature(s) may include a circumferential variation in at least one of the design parameters described herein, which can be designed to equilibrate N/rev blade unsteady loading in response to the inflow distortion for at least one of a set of noise-sensitive operating condition(s). In such embodiments, the design parameters can take into account an aggressive short inlet windward side crosswind distortion and/or a high takeoff angle-of-attack distortion.

In still further embodiments, the feature(s) may include acoustic treatment concepts integrated at outer diameter surfaces of the feature(s) for fan self-noise reduction. The acoustic treatments may include, for example, a bulk absorber feature, a perforated material resistance layer, a wire mesh resistance layer, a single-degree-of-freedom acoustic liner, a double-degree-of-freedom acoustic liner, a multiple-degree-of-freedom acoustic liner, or combinations thereof.

Referring now to the drawings,FIG.1illustrates a cross-sectional view of an embodiment of a gas turbine engine10that may be utilized within an aircraft in accordance with aspects of the present disclosure. More particularly, for the embodiment ofFIG.1, the gas turbine engine10is a high-bypass turbofan jet engine, with the gas turbine engine10being shown having a longitudinal or axial centerline axis12extending therethrough along an axial direction A for reference purposes. The gas turbine engine10further defines a radial direction R extending perpendicular from the centerline12. Further, a circumferential direction C (shown in/out of the page inFIG.1) extends locally perpendicular (i.e., circumferentially about the centerline12) to both the centerline12and the radial direction R. Although an exemplary turbofan embodiment is shown, it is anticipated that the present disclosure can be equally applicable to turbomachinery in general, such as a turboshaft, turbojet, or a turboprop configuration, including marine and industrial turbine engines and auxiliary power units.

In general, the gas turbine engine10includes a core gas turbine engine (indicated generally by reference character14) and a fan section16positioned upstream thereof. The core engine14generally includes a substantially tubular outer casing18that defines an annular inlet20. In addition, the outer casing18may further enclose and support a low pressure (LP) compressor22for increasing the pressure of the air that enters the core engine14to a first pressure level. A multi-stage, axial-flow high pressure (HP) compressor24may then receive the pressurized air from the LP compressor22and further increase the pressure of such air. The pressurized air exiting the HP compressor24may then flow to a combustor26within which fuel is injected into the flow of pressurized air, with the resulting mixture being combusted within the combustor26. The high energy combustion products60are directed from the combustor26along the hot gas path of the gas turbine engine10to a high pressure (HP) turbine28for driving the HP compressor24via a high pressure (HP) shaft or spool30, and then to a low pressure (LP) turbine32for driving the LP compressor22and fan section16via a low pressure (LP) drive shaft or spool34that is generally coaxial with HP shaft30. After driving each of turbines28and32, the combustion products60may be expelled from the core engine14via an exhaust nozzle36to provide propulsive jet thrust.

Additionally, as shown inFIGS.1and2, the fan section16of the gas turbine engine10generally includes a rotatable, axial-flow fan rotor38configured to be surrounded by an annular fan casing40. In particular embodiments, the LP shaft34may be connected directly to the fan rotor38or a rotor disk39, such as in a direct-drive configuration. In alternative configurations, the LP shaft34may be connected to the fan rotor38via a speed reduction device37such as a reduction gear gearbox in an indirect-drive or geared-drive configuration. Such speed reduction devices may be included between any suitable shafts/spools within the gas turbine engine10as desired or required. Additionally, the fan rotor38and/or rotor disk39may be enclosed or formed as part of a fan hub41.

It should be appreciated by those of ordinary skill in the art that the fan casing40may be configured to be at least partially supported relative to the core engine14by a plurality of substantially radially-extending, circumferentially-spaced outlet guide vanes42. As such, the fan casing40may enclose the fan rotor38and its corresponding fan rotor blades (fan blades44). Moreover, a downstream section46of the fan casing40may extend over an outer portion of the core engine14so as to define a secondary, or by-pass, airflow conduit48that provides additional propulsive jet thrust.

During operation of the gas turbine engine10, it should be appreciated that an initial airflow (indicated by arrow50) may enter the gas turbine engine10through an associated inlet52of the fan casing40. The airflow50then passes through the fan blades44and splits into a first compressed airflow (indicated by arrow54) that moves through the by-pass conduit48and a second compressed airflow (indicated by arrow56) which enters the LP compressor22. The LP compressor22may include a plurality of rotor blades (LP rotor blades45) enclosed by the outer casing18. The pressure of the second compressed airflow56is then increased and enters the HP compressor24(as indicated by arrow58). Additionally, the HP compressor24may include a plurality of rotor blades (HP rotor blades47) enclosed by the outer casing18. After mixing with fuel and being combusted within the combustor26, the combustion products60exit the combustor26and flow through the HP turbine28. Thereafter, the combustion products60flow through the LP turbine32and exit the exhaust nozzle36to provide thrust for the gas turbine engine10.

Referring now toFIGS.3A,3B, and4, various views of an embodiment of a rotary component61of the gas turbine engine10ofFIG.1are illustrated. Specifically, as shown, the rotary component61may be configured as a portion of the fan section16, but it should be appreciated that the rotary component61may be configured as the LP compressor22, the HP compressor24, the HP turbine28, the LP turbine32, and/or any other rotary component61of the gas turbine engine10. As indicated,FIG.4is taken along section line4-4ofFIG.3A. As shown, the rotary component61includes a one or more sets of circumferentially spaced rotor blades62, such as fan blades44, which extend radially outward towards an outer casing64from a hub66. As such, the rotor blades62may be coupled to a rotating shaft (such as fan rotor38as shown inFIG.1). Further, the outer casing64may be arranged exterior to the rotor blades62in the radial direction R. It should be appreciated that the outer casing64may be a part of the fan casing40(FIG.1) or a standalone component coupled thereto. Each of the rotor blades62may be circumscribed by the outer casing64, such that a tip clearance gap68is defined between the outer casing64and a rotor blade tip63of each rotor blade62.

During operation, the aerodynamic loading and efficiency of the rotary component61is generally affected by the tip leakage flow, as indicated by directional arrows74(FIG.3A), proximate the rotor blade tips63, as well as aerodynamic loading and drag effects as the airflow passes over the rotor blades62through the passages between adjacent rotor blades62. This results in wakes from the rotor blades62that evolve downstream and appear upon arriving at the OGVs42(FIG.2), as illustrated inFIG.4, as a plurality of airflow variations71, such as fan or rotor wakes. In particular, as shown inFIG.4, these fan or rotor wakes are illustrated as an undistorted wakes, wherein the inflow to the fan section16(FIG.1) is substantially uniform about the circumferential direction.

Additionally, the initial airflow50traveling through the inlet52through the annular inlet20may include flow non-uniformities75both radially and circumferentially (as shown inFIGS.3A-5). Such flow non-uniformities75may be present in the initial airflow50passing through the inlet52, and/or the fan section16may form such circumferential flow non-uniformities75. These flow non-uniformities75upon interaction with the rotor blades62result in circumferentially distorted wakes, which result in perturbed wake shapes to what is illustrated as the airflow variations71ofFIG.4. Such distorted wakes are deeper and/or wider when associated with overloaded fan blade responses to the distorted inflow, and vice versa for underloaded responses. To reduce fan-OGV interaction noise, one must address both the wakes associated with a mean undistorted fan inflow response as well as the potentially distorted wakes resulting from a distorted inflow fan response, which is the subject of this disclosure as described below.

Referring now toFIG.5, a cross-section of the fan section16is illustrated along the radial direction R and circumferential direction C. Further, as illustrated, the initial airflow50may be directed across the fan blades44(not shown) and into the core engine14(FIG.1). Further, as shown, the initial airflow50may contain flow non-uniformities (illustrated as triangles75inFIG.5) such as but not limited to pressure, temperature, velocity, and/or swirl or angular variations. As shown, the flow non-uniformities75may vary along the circumferential direction C of the fan casing40. As such, the initial airflow50may define one or more high distortion locations88and one or more low distortion locations90. Further, the high distortion locations88, which are generally caused by crosswind and angle of attack effects, may affect the stall margins and/or efficiency of the fan section16.

Further, such high distortion level locations88may be introduced by one or more stages70(FIGS.3A-3B) of the rotary component(s)61(FIGS.3A-3B) and affect stages70of the rotary component(s)61downstream thereof. For example, high distortion level locations88may be introduced by the fan section16and affect the LP compressor22(FIG.1) and HP compressor24(FIG.1) downstream of the fan section16. Further, high distortion level locations88flowing through annular inlet20(FIG.1) may affect the LP compressor22and HP compressor24. Further, high distortion locations88may be introduced by the LP compressor22(such as by the upstream stages70of the LP compressor22) and affect downstream stages70of the LP compressor22or the HP compressor24. In addition, high distortion locations88may be introduced by the HP compressor24(such as by the upstream stages70of the HP compressor24) and affect downstream stages70of the HP compressor24.

Referring specifically toFIGS.3A and3B, novel features78for the outer casing64that address the circumferential flow non-uniformities75described herein to reduce fan noise are illustrated. In particular, as shown in each ofFIGS.3A and3B, the outer casing64may define an interior surface76. Further, as shown, the tip clearance gap68may be defined between the rotor blades62and the interior surface76. Further, the interior surface76may include a plurality of the features78. For instance, as shown inFIG.3A, the feature(s)78may include one or more slots80. In particular embodiments, as shown, the slots80may extend generally along the axial direction A. However, in further embodiments, the slot(s)80may extend along any suitable direction. Such feature(s)78may be formed in the outer casing64after manufacturing the outer casing64(e.g., the features78may be machined in the interior surface76of the outer casing64). However, in other embodiments, the feature(s)78may be formed integrally with the outer casing64(e.g., the features78may be formed in the outer casing64during an additive manufacturing process or casting process). Furthermore, as shown inFIG.3B, the feature(s)78may include one or more circumferential grooves120.

Thus, as shown inFIGS.3A and3B, one or more of the features78may be positioned radially outward from one or more of the rotor blades62and into the volume of the enshrouding nacelle or casing structure delimited by interior surface76and outer casing64. Further, the feature(s)78may be positioned axially between a leading edge84and a trailing edge86of the rotor blade(s)62. For instance, each of the features78may be positioned between the leading edges84and trailing edges86of the rotor blades62of a stage70of the rotary component61. As such, the feature(s)78may be positioned within one or more of the blade passages positioned between the rotor blade tips63of a stage70and the outer casing64. In further embodiments, one or more of the features78of may be positioned at least partially radially outward from one or more of the rotor blades62on the interior surface76of the outer casing64. For instance, the feature(s)78may be positioned at least partially axially forward of the leading edge84or at least partially axially rearward of trailing edge86of the rotor blade(s)62, such as the rotor blades62of a stage70. As such, one or more of the features78may be positioned partially within one or more of the blade passages.

Referring to now generally toFIGS.6-16, multiple embodiments of rotary components61including features78on the interior surface76of the outer casing64are illustrated in accordance with aspects of the present disclosure. Particularly,FIG.6illustrates a partial, front view of an embodiment of the rotary component61with the features78on the interior surface76of the outer casing64in accordance with aspects of the present disclosure;FIGS.7A-7Fillustrate various plots of example slot configurations for the outer casing64of the rotary component61and their corresponding wave patterns in accordance with aspects of the present disclosure; andFIG.8illustrates a partial, front view of another embodiment of the rotary component61with the features78on the interior surface76of the outer casing64in accordance with aspects of the present disclosure.

More specifically, as shown inFIGS.6and8and mentioned, the fan section16includes the outer casing64, such as fan casing40, arranged exterior to the plurality of fan blades44in the radial direction. The fan casing40defines a tip clearance gap68(FIGS.11and12) between rotor blade tips63of the fan blades44and the fan casing40. Moreover, as shown inFIGS.6and8and mentioned, the fan casing40includes a plurality of features78formed into an interior surface76thereof. For example, in an embodiment, as shown inFIGS.6and8, one or more of the features78may be a slot80, such as an axial slot. In another embodiment, as shown inFIGS.3B and16, one or more of the features78may be a circumferential groove120.

Furthermore, as shown in the embodiment ofFIGS.6and8, each of the plurality of features78includes one or more design parameters. For example, as shown inFIG.8, the design parameter(s) of the features78may be perturbed about a mean design parameter (e.g., such as a mean depth81) for stall performance so as to provide a circumferential variation in wake strengths associated with the plurality of fan blades44, thereby reducing fan-outlet-guide-vane interaction noise of the gas turbine engine10. As used herein, the term “design parameter” generally refers to a predetermined characteristic of one or more of the features78relating to design of the feature(s)78that can be ascertained or set at the design stage of the outer casing40. More specifically, for example, in the context ofFIGS.6and8, the term “design parameter” refers to one of, or a combination of two or more of: blade overlap, a radial height, an axial dimension, a circumferential dimension, an orientation, relative blade position, slot type, and separation distance between two adjacent features of the plurality of features, and combinations thereof. Moreover, as used herein, the term “separation distance between adjacent features” refers to a distance from one feature to the next closest feature (with no features positioned therebetween).

In further embodiments, as shown inFIGS.6and8, the design parameter(s) of the plurality of features78may define a wave pattern72with a plurality of wave cycles73or periodic repetitions. In particular, as shown inFIGS.6and8, a single wave cycle73of the wave pattern72extends from a first crest77to a second crest79with a trough85in between. Thus, in an embodiment, each of the plurality of wave cycles73may be circumferentially equal to a multiple of downstream guide vanes, such as outlet guide vanes42(FIG.2) (such as 1×, 2×, 3×, etc. a multiple of downstream guide vanes). Moreover, as shown inFIGS.6and8, in an embodiment, the fan blades44rotate in a clockwise direction94and generate wakes96. As such, in an embodiment, the wave pattern72of the design parameter(s) of the plurality of features78may be clocked with the outlet guide vanes42to ensure low noise wakes (such as wakes96) impinge on leading edges92of the outlet guide vanes42and high noise wakes pass between the outlet guide vanes42(as depicted inFIG.9Band described further below) for one or more noise-sensitive operating conditions, such as acoustic certification conditions that may include sidelines, cutback, and/or approach.

Furthermore, in such embodiments, the number of wave cycles73or periodic repetitions of the wave pattern72of the features78may be selected as a function of a number of the downstream outlet guide vanes42and one or more design considerations such that blade tip loading is varied and wake shed from the fan blades44is clocked to the downstream outlet guide vanes42. In particular embodiments, for example, the number of wave cycles73of the wave pattern72may be selected using the relationship below:
NCT=k*NVWhere NVis the number of outlet guide vanes42downstream of the fan blades44, k is any integer 1, 2, 3 etc. depending on other design considerations or preferences, andNCTis the number of waves or cycles of features, which can be the number of physical features or the periodicity of the wave or repeating pattern (relating to the circumferential non-uniformity in their design parameters).

In an embodiment, for example, as shown inFIG.6, the features78may include a single slot80per downstream guide vane. Thus, as shown at83, the full array of slots80may be clocked relative to the array of downstream outlet guide vanes42in that the full array of slots80is shifted circumferentially to favorably align the wakes96to account for how the wakes96evolve, swirling circumferentially as the wakes evolve downstream (FIG.6). As used herein, the term “clocking” generally describes the process of circumferentially arranging the features78with respect to the downstream guide vanes in a manner such that low noise wakes (such as the tip vicinity of the wakes96) impinge on the leading edges92of the outlet guide vanes42and high noise wakes pass between the outlet guide vanes42. Such clocking may be determined using design methods and/or computational models well known to one skilled in the art of turbomachinery aerodynamics. For example, for a fan exit flow swirl angle α2and fan trailing edge circumferential location θ, the circumferential location where a fan tip wake impinges on an OGV tip leading edge is given approximately by the relationship

θ+L⁢tan⁢α2Rtip.
In such embodiments, this relationship can be used to determine the circumferential location or clocking offset relative to the OGVs of the non-uniform casing treatment parameters such that the low noise wakes impinge on the OGV leading edges and the high noise wakes pass between the OGVs.

Furthermore, as shown inFIGS.7A-7F, various plots of example slot configurations for the slots80(FIGS.6and8) in accordance with aspects of the present disclosure are provided. In particular,FIG.7Aillustrates a single slot square wave pattern according to the present disclosure, wherein the width, w, of the slots can be designed relative to the spacing between vanes, S=2πRTIP/NV, to affect what can be thought of as a duty cycle w/S.FIG.7Billustrates a double slot wave pattern according to the present disclosure. The concept naturally extends to multiple slots per wave such asFIGS.7C and7Dillustrate for two examples of generally symmetric wave patterns according to the present disclosure.FIGS.7E and7Fillustrate two examples of generally asymmetric wave patterns according to the present disclosure. Here, symmetrical generally refers to symmetry of the wave patterns about the mid-wavelength location as pictured inFIGS.7C-7D. It should be understood thatFIGS.7A-7Fare provided for illustrative purposes only and are not meant to be limiting.

The wake interaction effects between the fan blades44and the outlet guide vanes42as well as clocking of the wave pattern72(FIGS.6and8) of the design parameter(s) of the plurality of features78(FIGS.6and8) with the outlet guide vanes42can be better understood with respect toFIGS.9A-9C. In particular, as shown inFIG.9A, at time (t)=t0, the instant when the tip portion of the fan wakes impinge on the leading edge92of the outlet guide vane42is identified (as shown at v′). As shown inFIG.9B, the wake (gust) strengths are varied circumferentially about the mean solution. Thus, quiet portions (as shown by v′) can be aligned (e.g., t=t0−1/(vane passing frequency, VPF) to reduce the outlet guide vane interaction noise and loud portions (denoted as v′+) can be aligned (e.g., t=t0−0.5/VPF) between the outlet guide vane42to compensate for an average mean flow behavior as desired. Accordingly, as shown inFIG.9C, the design parameter(s) can be selected to achieve wake modifications, clocked circumferentially to align with optimal timing for suitably chosen operating conditions of high aircraft noise sensitivity.

Referring now toFIGS.10-12, various views of another embodiment of a rotary component, such a fan section of a gas turbine engine, according to aspects of the present disclosure is illustrated. In particular, as shown inFIG.10, the gas turbine engine10may be part of an aircraft100, such as being attached to a wing103of the aircraft. Moreover, as shown inFIGS.11and12, the fan section16of the gas turbine engine10may include a plurality of fan blades44operably coupled to a rotating shaft extending along the central axis and an outer casing64, such as fan casing40, arranged exterior to the plurality of fan blades44in the radial direction. Moreover, as shown and mentioned, the outer casing64includes a plurality of features78formed into the interior surface76thereof. Furthermore, as shown in the illustrated embodiment, certain of the plurality of features78include one or more design parameters that are locally varied at one or more high distortion locations88where distortion leads to overloading, thereby reducing the overloading and inlet distortion/fan interaction noise of the gas turbine engine10. As mentioned, the design parameter(s) may include, for example, blade overlap, a radial height, an axial dimension, a circumferential dimension, an orientation, relative blade position, slot type, separation distance between two adjacent features of the plurality of features, or combinations.

In such embodiments, the outer casing64may include a location of a first distortion level101on the interior surface76of the outer casing64and a location of a second distortion level102on the interior surface76of the outer casing64. For instance, the interior surface76of the outer casing64may define a first distortion level101at the high distortion locations88. Or, more particularly, the inflow distortion comes from the inlet geometry subjected to off-design conditions (e.g., crosswind, and high angle-of-attack) and the fan blades44, including the tip leakage flows, respond to that distortion. Thus, in an embodiment, the locally modified slot geometry is generally designed to equilibrate the fan blade loading as it rotates through the distorted flow to favorably impact that response, and hence the fan performance, operability, and noise.

Similarly, the interior surface76of the outer casing64may define a second distortion level102at other low distortion locations90. Or, more particularly, the airflow passing through the tip clearance gap68between the interior surface76of the outer casing64and the rotor blade tips63at the other low distortion locations90may define the second distortion level102at or upstream of the stage70(FIGS.3A-3B) of the rotary component61. It should be appreciated that the first distortion level101may be defined at a high distortion location88whereas the second distortion level102may be defined at a low distortion location90. As such, the second distortion level102may be less than the first distortion level101.

More particularly, as shown inFIGS.11and12, the high distortion location(s)88may be a windward side98of the outer casing64or a keel99of the outer casing64. Such locations may be selected due to crosswind distortion (as indicated by arrow105inFIG.10) at the windward side98and/or high angle-of-attack distortion (as indicated by arrow107inFIG.10) at the keel99. Thus, at such high distortion locations88, the design parameters, such as the radial depths, are locally varied in that the radial depths may increase up to a maximum depth at the high distortion location(s)88. Furthermore, as shown, the plurality of features at low distortion locations90outside of the high distortion location(s)88may include uniform design parameters (such as features78having equal radial depths, including zero depth and greater). In particular, as shown, the low distortion locations90outside of the high distortion location(s) may include a crown or top of the outer casing64, a leeward side of the outer casing64(i.e., the side opposite the windward side98), and/or an entirety of the circumference of the outer casing64.

Referring now toFIGS.13A-16, still further embodiments of the features78that may be formed into the outer casing64of the fan section16(FIG.1) to reduce fan self-noise according to aspects of the present disclosure are illustrated. In particular,FIG.13Aillustrates a front view of one of the features78according to the present disclosure andFIG.13Billustrates a side view of one of the features78according to the present disclosure, particularly illustrating a flow path108of air through the fan section16.

Moreover, as shown in the illustrated embodiment ofFIGS.13A and13B, the features78may include an acoustic liner feature104integrated therein to reduce self-noise of the fan section16. As used herein, an acoustic liner feature generally refers to a feature applied to the features78to intentionally attenuate or suppress noise. More specifically, as generally shown inFIGS.13A-15B, the acoustic liner feature104may be defined by a separate volume (e.g., portions114separated from portions116via a resistance layer106) in the features78such that the air communication to each volume is different in the porous surface to generally allow for only acoustic communication. Thus, in certain embodiments, the resistance layer106is represented as a dashed line at the porous interfaces in each ofFIGS.13A-15Band can be a perforated face-sheet resistance layer or a wire mesh resistance layer (as well as any other example provided herein), with the separate volume being an acoustic volume.

In particular embodiments, for example, as shown inFIGS.13A,13B, and14A, the acoustic liner feature104of the plurality of features78may be integrated at an outer diameter110of the features78. Furthermore, in an embodiment, the acoustic liner feature104may include a bulk absorber material, a perforated face-sheet resistance layer, a wire mesh resistance layer, a single-degree-of-freedom acoustic liner, a double-degree-of-freedom acoustic liner, a multiple-degree-of-freedom acoustic liner, or combinations thereof. As used herein, an acoustic liner generally refers to a sandwich panel constructed of a porous top layer, a honeycomb structure, and an impervious bottom layer. Thus, as used herein, single-, double-, and multi-degree-of-freedom acoustic liners generally refer to acoustic liners having a perforated sheet over a liner core of resonant chambers, with the double- and multi-degree-of-freedom acoustic liners having an additional porous mesh or perforated septa embedded within the resonant chambers of the liner core. In particular,FIG.13Aillustrates the acoustic liner feature104having a perforated/wire mesh resistance layer106.

In still further embodiments, as shown inFIG.14A, portions114of the features78containing the acoustic liner feature104may extend at an angle112with respect portions116of the features78without the acoustic liner feature104. In particular, as shown, the angle112may be an obtuse angle, chosen for example so as to enable tighter packaging of the combined feature within a radially constrained nacelle volume. In other embodiments, the angle112may be any suitable angle ranging from 0 degrees to 180 degrees.

Moreover, in an embodiment, as shown inFIGS.14A and15A, portions114of the features78containing the acoustic liner feature104may be uniform (e.g., portions114of the features78containing the acoustic liner feature104may have equal radial depths). In addition, as shown inFIG.14A, the portions116of the features78without the acoustic liner feature104may also be uniform (e.g., may have the same radial depths). In contrast, as shown inFIGS.15A and15B, the portions114of the features78without the acoustic liner feature104may be non-uniform (e.g., may have different radial depths). Furthermore, as shown, the portions116of the features78with the acoustic liner feature104may be non-uniform (e.g., may have different radial depths).

In yet another embodiment, as shown inFIGS.14B and14C, every other one of the features78may include the acoustic liner feature104to create an alternating configuration. Furthermore, as shown particularly inFIG.14B, the plurality of features78may extend generally perpendicular within the outer casing64with respect to the interior surface76. In contrast, as shown inFIG.14C, the plurality of features78may be slanted (e.g., and may extend at an angle118) within the outer casing64with respect to the interior surface76thereof such that the plurality of features78extend in a direction towards a rotational direction of the fan section (with the rotational direction of the fan being in the clockwise direction inFIG.14C).

Referring now toFIG.16, still another embodiment of one of the features78that may be formed into the outer casing64of the fan section16(FIG.1) to reduce fan self-noise according to aspects of the present disclosure is illustrated. In particular, as shown, the feature78may be a characteristic of a circumferential groove120, e.g., a circumferentially varying wave cycle shape with mode order kV, where k is any integer 1, 2, 3, etc. In such embodiments, as shown, the wave cycle shape of the circumferential groove120may be aligned with a spacing122of the outlet guide vanes42and clocked so as to align the quiet portions of the wakes with the downstream outlet guide vane leading edges to reduce fan-outer-guide-vane interaction noise.

A rotary component for a gas turbine engine, the rotary component comprising: a plurality of rotor blades operably coupled to a rotating shaft extending along the central axis; and an outer casing arranged exterior to the plurality of rotor blades in a radial direction of the gas turbine engine, the outer casing defining a gap between a blade tip of each of the plurality of rotor blades and the outer casing, wherein the outer casing comprises a plurality of features formed into an interior surface thereof, each of the plurality of features comprising one or more design parameters that are perturbed about a mean design parameter for stall performance so as to provide a circumferential variation in wake strengths associated with the plurality of rotor blades, thereby reducing operational noise of the gas turbine engine.

The rotary component of any preceding clause, wherein the one or more design parameters of the plurality of features are related to a periodicity of downstream guide vanes of the gas turbine engine.

The rotary component of any preceding clause, wherein the one or more design parameters of the plurality of features together define a wave pattern with a plurality of circumferential wave cycles, wherein a number of the plurality of circumferential wave cycles is equal to a multiple of the downstream guide vanes.

The rotary component of any preceding clause, wherein the wave pattern of the one or more design parameters of the plurality of features is clocked with respect to the downstream guide vanes to ensure wakes associated with quieter interaction noise impinge on leading edges of the downstream guide vanes and wakes associated with louder interaction noise pass between the downstream guide vanes for one or more noise-sensitive operating conditions.

The rotary component of any preceding clause, wherein the plurality of wave cycles of the wave pattern is selected as a function of a number of the downstream guide vanes and one or more design considerations such that blade tip loading is varied and wake shed from the plurality of rotor blades after passing the plurality of features are clocked relative to the downstream guide vanes to minimize the operational noise of the gas turbine engine.

The rotary component of any preceding clause, wherein the one or more design parameters comprising at least one of blade overlap, a radial dimension, an axial dimension, a circumferential dimension, an orientation, relative blade position, slot type, or a separation distance between two adjacent features of the plurality of features.

The rotary component of any preceding clause, wherein the plurality of features each comprise a slot or a characteristic of a circumferential groove.

The rotary component of any preceding clause, wherein the rotary component is a fan section of the gas turbine engine, and wherein the operational noise comprises fan-outlet-guide-vane interaction noise of the fan section of the gas turbine engine.

A rotary component for a gas turbine engine, the rotary component comprising: a plurality of rotor blades operably coupled to a rotating shaft extending along the central axis; and an outer casing arranged exterior to the plurality of rotor blades in a radial direction of the gas turbine engine, the outer casing defining a gap between a blade tip of each of the plurality of rotor blades and the outer casing, wherein the outer casing comprises a plurality of features formed into an interior surface thereof, and wherein each of the plurality of features comprise one or more design parameters that define a wave pattern with a plurality of wave cycles around the outer casing, wherein the plurality of wave cycles of the wave pattern is clocked to an arrangement of downstream guide vanes of the rotary component such that wakes associated with quieter interaction noise impinge on leading edges of the downstream guide vanes and wakes associated with louder interaction noise pass between the downstream guide vanes.

The rotary component of any preceding clause, wherein the plurality of features comprise a plurality of slots, and wherein a duty cycle of the wave pattern is a function of a width, shape, orientation of the plurality of slots, and a spacing between the downstream guide vanes.

The rotary component of any preceding clause, wherein the one or more design parameters comprising at least one of a radial dimension, an axial dimension, a circumferential dimension, an orientation, slot type, or separation distance between two adjacent features of the plurality of features.

The rotary component of any preceding clause, wherein the plurality of features is clocked to an array of downstream guide vanes.

The rotary component of any preceding clause, wherein the wave pattern is symmetrical.

The rotary component of any preceding clause wherein the wave pattern is asymmetrical.

A rotary component for a gas turbine engine, the rotary component comprising: a plurality of rotor blades operably coupled to a rotating shaft extending along the central axis; and an outer casing arranged exterior to the plurality of rotor blades in a radial direction of the gas turbine engine, the outer casing defining a gap between a blade tip of each of the plurality of rotor blades and the outer casing, wherein the outer casing comprises a plurality of features formed into an interior surface thereof, and wherein certain of the plurality of features comprise one or more design parameters that are locally varied at one or more distortion locations where distortion leads to overloading, thereby reducing the overloading and operational noise of the gas turbine engine.

The rotary component of any preceding clause, wherein the one or more design parameters of the plurality of features are related to a periodicity of downstream guide vanes of the gas turbine engine.

The rotary component of any preceding clause, wherein the one or more distortion locations comprise at least one of a windward side of the outer casing and/or a keel section of the outer casing.

The rotary component of any preceding clause, wherein the plurality of features at locations outside of the one or more distortion locations comprise uniform design parameters, and wherein the locations outside of the one or more distortion locations comprise at least one of a crown of the outer casing or a leeward side of the outer casing.

The rotary component of any preceding clause, wherein the rotary component is a fan section of the gas turbine engine.

The rotary component of any preceding clause, wherein the operational noise comprises at least one of inlet distortion noise and fan outlet-guide-vane (OGV) interaction noise of the fan section of the gas turbine engine.

A rotary component, comprising: a plurality of rotor blades operably coupled to a rotating shaft extending along a central axis of the rotary component; and an outer casing arranged exterior to the plurality of rotor blades in a radial direction of the rotary component, the outer casing defining a gap between a blade tip of each of the plurality of rotor blades and the outer casing, wherein the outer casing comprises a plurality of features formed into an interior surface thereof, and wherein at least one feature of the plurality of features is a slot having an acoustic liner feature integrated therein to reduce operational noise of the rotary component.

The rotary component of any preceding claim, wherein the at least one feature comprising the acoustic liner feature is related to a periodicity of one or more downstream guide vanes.

The rotary component of any preceding claim, wherein each of the plurality of features comprises one or more design parameters that together define a wave pattern with a plurality of circumferential wave cycles, wherein a number of the plurality of circumferential wave cycles is equal to a multiple of the downstream guide vanes.

The rotary component of any preceding claim, wherein the at least one feature comprising the acoustic liner feature is independent of a periodicity of one or more downstream guide vanes.

The rotary component of any preceding claim, wherein the acoustic liner feature of the at least one feature is integrated at an outer diameter thereof.

The rotary component of any preceding claim, wherein a portion of the at least one feature comprising the acoustic liner feature extends at an angle with respect to another portion of the at least one feature without the acoustic liner feature.

The rotary component of any preceding claim, wherein the one or more design parameters of a portion of the at least one feature comprising the acoustic liner feature are uniform and the one or more design parameters of another portion of the at least one feature without the acoustic liner feature is non-uniform.

The rotary component of any preceding claim, wherein the one or more design parameters of a portion of the at least one feature comprising the acoustic liner feature and another portion of the at least one feature without the acoustic liner feature are non-uniform.

The rotary component of any preceding claim, wherein the one or more design parameters comprise at least one of a blade overlap, radial height, an axial dimension, a circumferential dimension, an orientation, relative blade position, slot type, or separation distance from an adjacent feature.

The rotary component of any preceding claim, wherein every other of the plurality of features comprises the acoustic liner feature to create an alternating configuration of features with and without the acoustic liner feature.

The rotary component of any preceding claim, wherein the plurality of features extend generally perpendicular within the outer casing with respect to the interior surface thereof.

The rotary component of any preceding claim, wherein the acoustic liner feature comprises at least one of a bulk acoustic liner feature, a perforated material, a wire mesh resistance layer, a single-degree-of-freedom acoustic liner, a double-degree-of-freedom acoustic liner, a multiple-degree-of-freedom acoustic liner, or combinations thereof.

The rotary component of any preceding claim, wherein the at least one feature has a maximum depth on a windward side or a keel of the outer casing.

The rotary component of any preceding claim, wherein the rotary component is a fan section of a gas turbine engine.

The rotary component of any preceding claim, wherein the plurality of features extend at an angle within the outer casing with respect to the interior surface thereof such that the plurality of features extend in a direction towards a rotational direction of the fan section.

The rotary component of any preceding claim, wherein the operational noise comprises fan self-noise of the fan section.

A fan section, comprising: a plurality of fan blades operably coupled to a rotating shaft extending along a central axis of the fan section; and an outer casing arranged exterior to the plurality of fan blades in a radial direction of the fan section, the outer casing defining a gap between a blade tip of each of the plurality of fan blades and the outer casing, wherein the outer casing comprises at least one feature formed into an interior surface thereof, and wherein the at least one feature comprises an acoustic liner feature integrated therein to reduce fan self-noise of the fan section.

A fan section, comprising: a plurality of fan blades operably coupled to a rotating shaft extending along a central axis of the fan section; and an outer casing arranged exterior to the plurality of fan blades in a radial direction of the fan section, the outer casing defining a gap between a blade tip of each of the plurality of fan blades and the outer casing, wherein the outer casing comprises a plurality of features formed into an interior surface thereof, the plurality of features comprising one of a slot or a circumferential groove, wherein the plurality of features comprises one or more design parameters that together define a wave pattern with a plurality of circumferential wave cycles, wherein a number of the plurality of circumferential wave cycles is equal to a multiple of outlet guide vanes of the fan section to reduce fan-outlet-guide-vane (OGV) interaction noise of the fan section, and wherein at least one of the plurality of features comprises an acoustic liner feature integrated therein to reduce fan self-noise of the fan section.

The fan section of any preceding claim, wherein the design parameters are locally varied at one or more high distortion locations where distortion causes overloading, thereby reducing the overloading and operational noise of the fan section.

The fan section of any preceding claim, wherein the one or more distortion locations comprise at least one of a windward side of the outer casing and/or a keel section of the outer casing.