Patent Description:
The engines of an aircraft propelled by gas turbine engines produce varying amounts of audible noise during all stages of operation, including during takeoff and landing. For example, a gas turbine engine typically operates at or near maximum thrust as the aircraft departs from an airport, generating large amounts of engine noise, and at a lower thrust as the aircraft approaches an airport. Some aircraft engine noise can be partially suppressed at the engine nacelle inlet and the exhaust nozzle and center body by noise absorbing structures. These structures can absorb acoustic energy by canceling reflected acoustic waves or by converting acoustic energy into heat. The structures typically consist of a porous skin and three or more non-perforated walls to form one or more chambers. The porous skin and the non-perforated walls of such chambers combine to form a plurality of Helmholtz resonators that resonate in response to certain sound frequencies or certain bands of frequencies and cancel sound waves reflected between the porous face skin and non-perforated walls or subsequently convert the sound energy into heat (via elastic or mechanical hysteresis caused by the resonant response of air within the resonator cavities and of the liner components), and thereby effectively absorb or dissipate at least a portion of generated engine noise.

In general, relatively thin acoustic panels may be utilized to attenuate noise with relatively short wavelengths and high frequencies, whereas relatively thick acoustic panels may be utilized to attenuate noise with relatively long wavelengths and low frequencies. However, as noise wavelengths become longer as a byproduct of new engine designs while space allocation for noise attenuation structures decreases, traditional acoustic panel configurations may not attenuate noise to acceptable levels, which are often mandated by government regulations. To achieve further reductions in the noise levels of gas turbine engines used on modern aircraft, especially during aircraft takeoffs and approaches, it is desirable to dissipate some of the long-wavelength and low-frequency noise generated by the combustor and the exhaust system of a gas turbine engine, particularly where the combustor and exhaust noise has one or more frequencies less than about <NUM>,<NUM>.

<CIT> discloses an acoustic deep cavity centerbody.

<CIT> discloses a sound absorbing exhaust nozzle center plug.

<CIT> discloses an acoustic zoned system for turbofan engine exhaust application.

A center plug for attenuating noise in a gas turbine engine is disclosed as claimed in claim <NUM>.

Optionally, the outer skin is connected to and positioned radially outward of the forward bulkhead, the aft bulkhead and the first perforated disk. Optionally, the outer skin includes a first plurality of perforations extending into the first sub-cavity. Optionally, the first resonator cavity is configured to attenuate noise having frequencies within a range extending from about <NUM> to about <NUM>,<NUM>.

Optionally, the angle is within a first range from about forty degrees to about eighty degrees with respect to the axial centerline. Optionally, the first perforated disk is oriented within a second range from about minus ten degrees to about plus ten degrees with respect to a radial direction extending perpendicular to the axial centerline. Optionally, the first intermediate bulkhead is configured to form a cone extending circumferentially about the axial centerline. Optionally, the first perforated disk extends perpendicularly to the axial centerline.

Optionally, a second resonator cavity is disposed within the volume defined by the inner skin, the outer skin and the forward bulkhead and the aft bulkhead, the second resonator cavity including a second perforated disk extending between the inner skin and the outer skin and forming a third sub-cavity and a fourth sub-cavity. Optionally, a second intermediate bulkhead is disposed aft of the second perforated disk and extending between the inner skin and the outer skin. Optionally, the second intermediate bulkhead is oriented at the angle with respect to the axial centerline. Optionally, the second perforated disk is oriented within the second range from about minus ten degrees to about plus ten degrees with respect to the radial direction extending perpendicular to the axial centerline.

Optionally, a third resonator cavity is disposed within the volume defined by the inner skin, the outer skin and the forward bulkhead and the aft bulkhead, the third resonator cavity including a third perforated disk extending between the inner skin and the outer skin and forming a fifth sub-cavity and a sixth sub-cavity. Optionally, the second resonator cavity is bounded by the first intermediate bulkhead and the second intermediate bulkhead and the third resonator cavity is bounded by the second intermediate bulkhead and the aft bulkhead. Optionally, the outer skin includes a second plurality of perforations extending into the third sub-cavity and a third plurality of perforations extending into the fifth sub-cavity.

Optionally, a second resonator cavity is disposed within the volume defined by the inner skin, the outer skin and the forward bulkhead and the aft bulkhead, the second resonator cavity including a second perforated disk extending between the inner skin and the outer skin and forming a third sub-cavity and a fourth sub-cavity, the second perforated disk extending perpendicularly to the axial centerline extending through the center plug. Optionally, the first resonator cavity and the second resonator cavity are separated by an intermediate bulkhead oriented at an angle within a range from about forty degrees to about eighty degrees with respect to the axial centerline.

Referring now to the drawings, <FIG> illustrates an aircraft <NUM>, in accordance with various embodiments. The aircraft <NUM> is an example of a passenger or transport vehicle in which noise attenuation systems may be implemented in accordance with various embodiments. In an illustrative embodiment, the aircraft <NUM> has a starboard wing <NUM> and a port wing <NUM> attached to a fuselage <NUM>. The aircraft <NUM> also includes a starboard engine system <NUM> connected to the starboard wing <NUM> and a port engine system <NUM> connected to the port wing <NUM>. In various embodiments, the aircraft <NUM> also includes a starboard horizontal stabilizer <NUM>, a port horizontal stabilizer <NUM> and a vertical stabilizer <NUM>. A pylon <NUM> is used to connect a gas turbine engine within the starboard engine system <NUM> to the starboard wing <NUM> and a gas turbine engine within the port engine system <NUM> to the port wing <NUM>, though, in various embodiments, the gas turbine engines may be connected to other portions of the aircraft <NUM>, such as, for example, to the port and starboard sides of the fuselage <NUM>.

Referring now to <FIG>, a side cutaway illustration of a gas turbine engine system <NUM>, such as, for example, either of the starboard engine system <NUM> or the port engine system <NUM>, is provided. The gas turbine engine system <NUM> includes the pylon <NUM> and a gas turbine engine <NUM> (e.g., a propulsion system) such as, for example, a geared turbofan engine that uses an outlet guide vane <NUM> (OGV) (or a plurality of outlet guide vanes) to structurally connect a fan module to a core engine module as well as redirect the incoming fan flow to the outlet guide vane <NUM>. The gas turbine engine <NUM> is mounted to the pylon <NUM>, which may be mounted to or otherwise configured with an aircraft airframe. Examples of an aircraft airframe include, but are not limited to, an aircraft wing (e.g., the starboard wing <NUM> or the port wing <NUM>) or an aircraft fuselage (e.g., the fuselage <NUM>).

The gas turbine engine <NUM> extends along an axial centerline A between an airflow inlet <NUM> and a core exhaust system <NUM>. The gas turbine engine <NUM> includes a fan section <NUM>, a low-pressure compressor section <NUM> (LPC), a high-pressure compressor section <NUM> (HPC), a combustor section <NUM>, a high-pressure turbine section <NUM> (HPT) and a low-pressure turbine section (LPT) <NUM>. The engine sections are typically arranged sequentially along the axial centerline A. The low-pressure compressor section <NUM> (LPC), the high-pressure compressor section <NUM> (HPC), the combustor section <NUM>, the high-pressure turbine section <NUM> (HPT) and the low-pressure turbine section <NUM> (LPT) form a core <NUM> (or an engine core) of the gas turbine engine <NUM>.

Each of the low-pressure compressor section <NUM> (LPC), the high-pressure compressor section <NUM> (HPC), the high-pressure turbine section <NUM> (HPT) and the low-pressure turbine section <NUM> (LPT) typically include a rotor having a plurality of rotor blades arranged circumferentially around and connected to one or more respective rotor disks - e.g., a low-pressure compressor rotor <NUM>, a high-pressure compressor rotor <NUM>, a high-pressure turbine rotor <NUM> and a low-pressure turbine rotor <NUM>. A fan rotor <NUM> is connected to a gear train <NUM>. The gear train <NUM> and the low-pressure compressor rotor <NUM> are connected to and driven by the low-pressure turbine rotor <NUM> through a low-speed shaft <NUM> (or a low-speed spool). The high-pressure compressor rotor <NUM> is connected to and driven by the high-pressure turbine rotor <NUM> through a high-speed shaft <NUM> (or a high-speed spool).

Air enters the gas turbine engine <NUM> through the airflow inlet <NUM> and is directed through the fan section <NUM> and into a core gas flow path C and a bypass gas flow path B. The air within the core gas flow path C may be referred to as "core air. " The air within the bypass gas flow path B may be referred to as "bypass air. " The core air is directed through the low-pressure compressor section <NUM>, the high-pressure compressor section <NUM>, the combustor section <NUM>, the high-pressure turbine section <NUM> and the low-pressure turbine section <NUM> and exits the gas turbine engine <NUM> through the core exhaust system <NUM>, which includes an exhaust center body <NUM> surrounded by an exhaust nozzle <NUM>. Within the combustor section <NUM>, fuel is injected into and mixed with the core air and ignited to provide a hot airstream that drives the turbine sections. The bypass air is directed through the bypass gas flow path B, and out of the gas turbine engine <NUM> through a bypass exhaust nozzle <NUM> to provide forward engine thrust. The bypass air may also or alternatively be directed through a thrust reverser to provide reverse engine thrust. A fan nacelle <NUM> is typically employed to surround the various sections of the gas turbine engine <NUM> and a core nacelle <NUM> is typically employed to surround the various sections of the core <NUM>.

Referring now to <FIG>, a schematic illustration of a core exhaust system <NUM>, such as, for example, the core exhaust system <NUM> described above, is provided. In various embodiments, the core exhaust system <NUM> includes an exhaust nozzle <NUM> and a center body <NUM>, similar to the exhaust nozzle <NUM> and the exhaust center body <NUM> described above. The center body <NUM> may be formed in two sections, including, for example, a center plug <NUM> and an aft cone <NUM>. The exhaust nozzle <NUM> and the center plug <NUM> cooperate to form an annulus <NUM> through which exhaust gasses from a combustor section exit the core exhaust system <NUM>. In various embodiments, the center plug <NUM> and the aft cone <NUM> are connected along a circumferential seam <NUM> at an aft end of the center plug <NUM>. In the illustrated embodiment, the aft portion of the center plug <NUM> and the aft cone <NUM> extend aft from an aft end of the exhaust nozzle <NUM>. The radially outer surfaces of the center plug <NUM> and the aft cone <NUM> combine to form a flow control surface that substantially prevents recirculation of the exiting exhaust gasses and facilitates convergence of the exhaust gasses as they exit the annulus <NUM>. The center plug <NUM> forms a transition between an aft end of a turbine rotor (not shown) located just inside the core exhaust system <NUM> and the aft cone <NUM>. In various embodiments, the center plug <NUM> and the aft cone <NUM> may have hollow center portions that permit cooling air to pass from an intake <NUM> at an aft tip of the aft cone <NUM> to internal portions of the engine or to house instrumentation, wiring, or the like.

Referring now to <FIG>, schematic illustrations of a center plug <NUM>, similar to the center plug <NUM> described above, are provided. In various embodiments, the center plug <NUM> includes an outer skin <NUM> having an aerodynamic outer contour. The outer skin <NUM> is seamlessly constructed such that the center plug <NUM> has a substantially smooth outer surface. The center plug <NUM> may have a forward flange <NUM> configured for attachment to a casing proximate an aft end of a turbine rotor and an aft flange <NUM> configured for attachment to an aft cone, such as, for example, the aft cone <NUM> described above. As illustrated, the outer skin <NUM> may include an acoustically permeable portion <NUM> located on a forward portion of the outer skin <NUM> and extending around substantially the entire circumference of the forward portion of the outer skin <NUM>. In various embodiments, the acoustically permeable portion <NUM> is formed via one or more pluralities of perforations, such as, for example, a first plurality of perforations <NUM>, a second plurality of perforations <NUM> and a third plurality of perforations <NUM>, each of which extends through the outer skin <NUM> and into an interior portion of the center plug <NUM> that is bounded by the outer skin <NUM> and an inner skin <NUM>, the latter of which has a substantially cylindrical shape, is centered along a central longitudinal axis A of the center plug <NUM> and forms a substantially open center portion <NUM> of the center plug <NUM>.

Referring now to <FIG>, a cross sectional schematic illustration of a core exhaust system <NUM>, such as, for example, the core exhaust system <NUM> described above, is provided. In various embodiments, the core exhaust system <NUM> includes an exhaust nozzle <NUM> and a center body <NUM>, similar to the exhaust nozzle <NUM> and the center body <NUM> described above. The center body <NUM> may be formed in two sections, including, for example, a center plug <NUM> and an aft cone <NUM>, similar to those described above. The exhaust nozzle <NUM> and the center plug <NUM> cooperate to form an annulus <NUM> through which exhaust gasses from a combustor section exit the core exhaust system <NUM>. Similar to the description provided above with reference to <FIG>, the center plug <NUM> includes an outer skin <NUM> having an aerodynamic outer contour. The center plug <NUM> may have a forward flange <NUM> configured for attachment to a casing proximate an aft end of a turbine rotor and an aft flange <NUM> configured for attachment to the aft cone <NUM>. As illustrated, the outer skin <NUM> may include an acoustically permeable portion <NUM> located on a forward portion of the outer skin <NUM> and extending around substantially the entire circumference of the forward portion of the outer skin <NUM>. The acoustically permeable portion <NUM> may be formed by one or more pluralities of perforations, such as, for example, a first plurality of perforations <NUM>, a second plurality of perforations <NUM> and a third plurality of perforations <NUM> that extend through the outer skin <NUM>, with the first plurality of perforations <NUM> disposed forward of the second plurality of perforations <NUM> and the second plurality of perforations <NUM> being disposed forward of the third plurality of perforations <NUM>. As illustrated, each of the first plurality of perforations <NUM>, the second plurality of perforations <NUM> and the third plurality of perforations <NUM> typically extend in a circumferential pattern about the outer skin <NUM>.

As illustrated, the acoustically permeable portion <NUM> may coincide with one or more resonator cavities <NUM>, such as, for example, a first resonator cavity <NUM>, a second resonator cavity <NUM> and a third resonator cavity <NUM>. The one or more resonator cavities is typically contained within a volume or space bounded by the outer skin <NUM>, an inner skin <NUM>, a forward bulkhead <NUM> and an aft bulkhead <NUM>, with the volume or space extending circumferentially between the outer skin <NUM> and the inner skin <NUM> and axially between the forward bulkhead <NUM> and the aft bulkhead <NUM> to form a generally annular volume or space. The first resonator cavity <NUM> is formed by the volume or space bounded by the outer skin <NUM> and the inner skin <NUM> and the forward bulkhead <NUM> and a first intermediate bulkhead <NUM>. In similar fashion, the second resonator cavity <NUM> is formed by the volume or space bounded by the outer skin <NUM> and the inner skin <NUM> and the first intermediate bulkhead <NUM> and a second intermediate bulkhead <NUM>, and the third resonator cavity <NUM> is formed by the volume or space bounded by the outer skin <NUM> and the inner skin <NUM> and the second intermediate bulkhead <NUM> and the aft bulkhead <NUM>.

Still referring to <FIG>, each of the first resonator cavity <NUM>, the second resonator cavity <NUM> and the third resonator cavity <NUM> is configured to exhibit acoustic attenuation properties of a Helmholtz resonator having a double degree of freedom (DDOF) type design. For example, the first resonator cavity <NUM> is divided into two sub-cavities (e.g., a first sub-cavity <NUM> and a second sub-cavity <NUM>) via a first perforated disk <NUM> disposed between the forward bulkhead <NUM> and the first intermediate bulkhead <NUM> and extending between the outer skin <NUM> and the inner skin <NUM>. Similarly, the second resonator cavity <NUM> is divided into two sub-cavities (e.g., a third sub-cavity <NUM> and a fourth sub-cavity <NUM>) via a second perforated disk <NUM> disposed between the first intermediate bulkhead <NUM> and the second intermediate bulkhead <NUM> and extending between the outer skin <NUM> and the inner skin <NUM>, and the third resonator cavity <NUM> is divided into two sub-cavities (e.g., a fifth sub-cavity <NUM> and a sixth sub-cavity <NUM>) via a third perforated disk <NUM> disposed between the second intermediate bulkhead <NUM> and extending between the outer skin <NUM> and the inner skin <NUM>. The forward bulkhead <NUM>, the first intermediate bulkhead <NUM>, the second intermediate bulkhead <NUM> and the aft bulkhead <NUM> may be oriented at an angle <NUM> with respect to a central longitudinal axis A of the center plug <NUM> such that a cone-like structure is formed, with the angle <NUM> optionally being in a first range from about forty degrees (<NUM>°) to about eighty degrees (<NUM>°), or from about fifty degrees (<NUM>°) to about seventy degrees (<NUM>°) or being on the order of about sixty degrees (<NUM>°). Similarly, in various embodiments, one or more of the first perforated disk <NUM>, the second perforated disk <NUM> and the third perforated disk <NUM> may be oriented in a substantially radial direction, as illustrated in <FIG>, with respect to the central longitudinal axis A, or within a second range from about minus ten degrees (-<NUM>°) to about plus ten degrees (<NUM>°) from the radial direction with respect to the central longitudinal axis A.

During operation, an exhaust stream transits and exits the annulus <NUM> formed by the exhaust nozzle <NUM> and the center plug <NUM>. Non-attenuated acoustic waves from the exhaust stream enter the first plurality of perforations <NUM> and into the first resonator cavity <NUM> where attenuation occurs. The non-attenuated acoustic waves of the exhaust stream enter the first sub-cavity <NUM> and then the second sub-cavity <NUM> via the first perforated disk <NUM>. The acoustic waves are then reflected off the non-perforated walls of the second sub-cavity <NUM>, pass back through the first perforated disk <NUM> and into the first sub-cavity <NUM> and exit the first resonator cavity <NUM> via the first plurality of perforations <NUM> as attenuated acoustic waves. Similarly, non-attenuated acoustic waves from the exhaust stream enter the second plurality of perforations <NUM> and into the second resonator cavity <NUM> where attenuation occurs. The non-attenuated acoustic waves of the exhaust stream enter the third sub-cavity <NUM> and then the fourth sub-cavity <NUM> via the second perforated disk <NUM>. The acoustic waves are then reflected off the non-perforated walls of the fourth sub-cavity <NUM>, pass back through the second perforated disk <NUM> and into the third sub-cavity <NUM> and exit the second resonator cavity <NUM> via the second plurality of perforations <NUM> as attenuated acoustic waves. And finally, non-attenuated acoustic waves from the exhaust stream enter the third plurality of perforations <NUM> and into the third resonator cavity <NUM> where attenuation occurs. The non-attenuated acoustic waves of the exhaust stream enter the fifth sub-cavity <NUM> and then the sixth sub-cavity <NUM> via the third perforated disk <NUM>. The acoustic waves are then reflected off the non-perforated walls of the sixth sub-cavity <NUM>, pass back through the third perforated disk <NUM> and into the fifth sub-cavity <NUM> and exit the third resonator cavity <NUM> via the third plurality of perforations <NUM> as attenuated acoustic waves.

Benefits of the noise attenuation structures described above include the ability to attenuate low-frequency noise (e.g., on the order of <NUM> to <NUM>,<NUM>) via DDOF cavities that would otherwise be too large to incorporate into a center plug in the form of single degree of freedom (SDOF) resonator cavities. Further, the shape of the perforated and non-perforated walls of the disclosed DDOF resonator cavities (e.g., the cone-shaped non-perforated walls and the disk-shaped perforated walls) enable relative ease of manufacture when compared to the more typical honeycomb-shaped resonator cavities. The shape of the perforated and non-perforated walls also enable the resulting resonator cavities to fit against the compound curvature often exhibited by the outer skin of the center body by attaching the radially outboard edges of a series of perforated sheet metal disks and non-perforated sheet metal cones to the radially inside surface of the forward portion of the center body and the radially inboard edges of the same structures to the radially outside surface of the inner skin of the center body.

Numbers, percentages, or other values stated herein are intended to include that value, and also other values that are about or approximately equal to the stated value, as would be appreciated by one of ordinary skill in the art encompassed by various embodiments of the present disclosure. A stated value should therefore be interpreted broadly enough to encompass values that are at least close enough to the stated value to perform a desired function or achieve a desired result. The stated values include at least the variation to be expected in a suitable industrial process, and may include values that are within <NUM>%, within <NUM>%, within <NUM>%, within <NUM>%, or within <NUM>% of a stated value. Additionally, the terms "substantially," "about" or "approximately" as used herein represent an amount close to the stated amount that still performs a desired function or achieves a desired result. For example, the term "substantially," "about" or "approximately" may refer to an amount that is within <NUM>% of, within <NUM>% of, within <NUM>% of, within <NUM>% of, and within <NUM>% of a stated amount or value.

Claim 1:
A center plug (<NUM>,<NUM>,<NUM>) for attenuating noise in a gas turbine engine (<NUM>), comprising:
an inner skin (<NUM>,<NUM>), the inner skin (<NUM>,<NUM>) having a substantially cylindrical shape and extending along an axial centerline (A);
an outer skin (<NUM>,<NUM>) positioned radially outside the inner skin (<NUM>,<NUM>);
a forward bulkhead (<NUM>) disposed proximate a forward end of the inner skin (<NUM>,<NUM>), the forward bulkhead (<NUM>) connected to and extending radially outward from the inner skin (<NUM>,<NUM>), and the forward bulkhead (<NUM>) oriented at an angle (<NUM>) with respect to the axial centerline (A);
an aft bulkhead (<NUM>) disposed proximate an aft end (<NUM>) of the inner skin (<NUM>,<NUM>), the aft bulkhead (<NUM>) connected to and extending radially outward from the inner skin (<NUM>,<NUM>), and the aft bulkhead oriented at the angle (<NUM>) with respect to the axial centerline;
a first resonator cavity (<NUM>) disposed within a volume defined by the inner skin (<NUM>,<NUM>), the outer skin (<NUM>,<NUM>) and the forward bulkhead (<NUM>) and the aft bulkhead (<NUM>), the first resonator cavity (<NUM>) including a first perforated disk (<NUM>) extending between the inner skin (<NUM>,<NUM>) and the outer skin (<NUM>,<NUM>) and forming a first sub-cavity (<NUM>) and a second sub-cavity (<NUM>),
characterised in that the center plug (<NUM>,<NUM>,<NUM>) further comprises:
a first intermediate bulkhead (<NUM>) disposed aft of the first perforated disk (<NUM>) and extending between the inner skin (<NUM>,<NUM>) and the outer skin (<NUM>,<NUM>), the first intermediate bulkhead oriented at the angle (<NUM>) with respect to the axial centerline.