Patent Description:
Certain combustors on aircraft engines may be particularly susceptible to combustion dynamics. Under certain engine operating conditions, significant transient pressure waves ("pings") can be present, particularly in an annular combustor. These pressure waves, if of sufficient magnitude, may cause high cycle fatigue of combustor components, long before the hardware would need to be replaced under normal operations.

One known approach to combustor dynamic issues can involve careful mapping of problem regimes using test engines with multiple combustor instrumentation pressure sensors. Aircraft fuel schedules developed from this process and subsequently programmed into engine control were expected to avoid all problem areas. Despite this mapping, however, subtle differences between engines may still adversely affect combustion dynamics behavior. These changes may be due to parameters including manufacturing variations, engine deterioration, fuel composition, or unexpected flight conditions.

Therefore, it may be beneficial to monitor combustor dynamics during fielded operation of the gas turbine engine such that modifications to one or more control parameters may be made to reduce the combustor dynamics in the event they progress above a certain threshold. However, it can be difficult to accurately measure combustor dynamics during operation given the relatively harsh conditions within a combustion chamber of the gas turbine engine.

Accordingly, features for monitoring combustion dynamics in aircraft engines are desired. Specifically, features for more accurately measuring a pressure within a combustion section of a gas turbine engine would be particularly useful. <CIT> relates to a dynamic pressure probe holder and a method of obtaining a dynamic pressure signal. <CIT> relates to a combustion chamber dynamic pressure transducer tee probe holder and related method. <CIT> relates to a probe for measuring pressure oscillations. <CIT> relates to measuring pressure in a combustor (gas turbine) damping tube without moisture condensation and bundled tubes. <CIT> relates to an apparatus for monitoring the combustion oscillation of a gas turbine.

According to the present invention a pressure sensor assembly for a gas turbine engine is provided as set forth in claim <NUM>, and a gas turbine engine defining a radial direction and a circumferential direction is provided as set forth in claim <NUM>. The gas turbine engine includes a liner positioned within a compressor section or a turbine section of the gas turbine engine and at least partially defining a core air flowpath through the gas turbine engine. The liner defines a liner opening. The gas turbine engine also includes a casing at least partially enclosing the liner, the casing defining first, inner side along the radial direction, a second, outer side opposite the first side, and a casing opening. The gas turbine engine also includes a pressure sensor assembly comprising a body, an extension member, and a pressure sensor, the pressure sensor position at least partially within the body and the body positioned at least partially on the second side of the casing, the extension member extending from the body through the casing opening in the casing and towards the liner opening in the liner, the extension member defining a continuous sense cavity exposing the pressure sensor to the core air flowpath.

In certain exemplary embodiments the liner is an outer liner of a combustor assembly of the gas turbine engine, and wherein the casing is a combustor casing.

The pressure sensor includes a diaphragm exposed to the sense cavity of the extension member, and the pressure sensor is an optical-based sensor for measuring a deflection of the diaphragm. The optical-based sensor may include an optical laser directed to the diaphragm.

Further, in certain exemplary embodiments the extension member defines one or more cooling holes located inward of the casing along the radial direction and outward of the liner along the radial direction.

Additionally, in certain exemplary embodiments the liner includes a ferrule, and the extension member extends into the ferrule.

Moreover, in certain exemplary embodiments the pressure sensor assembly is removably coupled to the casing.

Further, in certain exemplary embodiments the pressure sensor assembly further includes a clamp nut, wherein the clamp nut is removably coupled to the body to hold the sensor in position.

Additionally, in certain exemplary embodiments the gas turbine engine further includes a plurality of pressure sensor assemblies arranged along the circumferential direction of the gas turbine engine.

Moreover, in certain exemplary embodiments the casing opening of the casing and the liner opening of the liner are each configured as part of a borescope inspection port.

Further, in certain exemplary embodiments the liner is a liner of a high pressure turbine in the turbine section of the gas turbine engine, and wherein the casing is a turbine casing.

In another exemplary embodiment of the present disclosure, a pressure sensor assembly is provided for a gas turbine engine. The gas turbine engine defines a radial direction and includes a liner at least partially defining a core air flowpath through a compressor section or a turbine section and a casing at least partially enclosing the liner. The pressure sensor assembly includes a body configured for positioning adjacent to the casing of the gas turbine engine, and a pressure sensor positioned at least partially within the body. The pressure sensor assembly further includes an extension member extending from the body and configured to extend at least partially through a casing opening in the casing and towards a liner opening in the liner, the extension member defining a continuous sense cavity to expose the pressure sensor to the core air flowpath of the gas turbine engine.

The pressure sensor includes a diaphragm exposed to the sense cavity of the extension member, wherein the pressure sensor is an optical-based sensor for measuring a deflection of the diaphragm. The optical-based sensor may include an optical laser directed to the diaphragm.

Moreover, in certain exemplary aspects the extension member defines one or more cooling holes located inward of the casing along the radial direction and outward of the liner along the radial direction when the pressure sensor assembly is installed in the gas turbine engine.

Further, in certain exemplary aspects the pressure sensor assembly is configured to be removably coupled to the casing.

Additionally, in certain exemplary aspects the pressure sensor assembly further comprises a clamp nut, wherein the clamp nut is removably coupled to the body to hold the sensor in position.

Moreover, in certain exemplary aspects the liner is an outer liner of a combustor assembly of the gas turbine engine, wherein the casing is a combustor casing.

The terms "forward" and "aft" refer to relative axial positions within a gas turbine engine, with forward referring to a position closer to an engine inlet and aft referring to a position closer to an engine nozzle or exhaust.

Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures, <FIG> is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of <FIG>, the gas turbine engine is a high-bypass turbofan jet engine <NUM>, referred to herein as "turbofan engine <NUM>. " As shown in <FIG>, the turbofan engine <NUM> defines an axial direction A (extending parallel to a longitudinal centerline <NUM> provided for reference), a radial direction R, and a circumferential direction (i.e., a direction extending about the axial direction A; not depicted). In general, the turbofan <NUM> includes a fan section <NUM> and a core turbine engine <NUM> positioned downstream from the fan section <NUM>.

The exemplary core turbine engine <NUM> depicted generally includes a substantially tubular outer casing <NUM> that defines an annular inlet <NUM>. The outer casing <NUM> encases, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor <NUM> and a high pressure (HP) compressor <NUM>; a combustion section <NUM>; a turbine section including a high pressure (HP) turbine <NUM> and a low pressure (LP) turbine <NUM>; and a jet exhaust nozzle section <NUM>. A high pressure (HP) shaft or spool <NUM> drivingly connects the HP turbine <NUM> to the HP compressor <NUM>. A low pressure (LP) shaft or spool <NUM> drivingly connects the LP turbine <NUM> to the LP compressor <NUM>. The compressor section, combustion section <NUM>, turbine section, and jet exhaust nozzle section <NUM> together define a core air flowpath <NUM> through the core turbine engine <NUM>.

For the embodiment depicted, the fan section <NUM> includes a variable pitch fan <NUM> having a plurality of fan blades <NUM> coupled to a disk <NUM> in a spaced apart manner. As depicted, the fan blades <NUM> extend outwardly from disk <NUM> generally along the radial direction R. Each fan blade <NUM> is rotatable relative to the disk <NUM> about a pitch axis P by virtue of the fan blades <NUM> being operatively coupled to a suitable actuation member <NUM> configured to collectively vary the pitch of the fan blades <NUM> in unison. The fan blades <NUM>, disk <NUM>, and actuation member <NUM> are together rotatable about the longitudinal axis <NUM> by LP shaft <NUM> across a power gear box <NUM>. The power gear box <NUM> includes a plurality of gears for stepping down the rotational speed of the LP shaft <NUM> to a more efficient rotational fan speed.

Referring still to the exemplary embodiment of <FIG>, the disk <NUM> is covered by rotatable front nacelle <NUM> aerodynamically contoured to promote an airflow through the plurality of fan blades <NUM>. Additionally, the exemplary fan section <NUM> includes an annular fan casing or outer nacelle <NUM> that circumferentially surrounds the fan <NUM> and/or at least a portion of the core turbine engine <NUM>. The nacelle <NUM> is supported relative to the core turbine engine <NUM> by a plurality of circumferentially-spaced outlet guide vanes <NUM>. Moreover, a downstream section <NUM> of the nacelle <NUM> extends over an outer portion of the core turbine engine <NUM> so as to define a bypass airflow passage <NUM> therebetween.

During operation of the turbofan engine <NUM>, a volume of air <NUM> enters the turbofan <NUM> through an associated inlet <NUM> of the nacelle <NUM> and/or fan section <NUM>. As the volume of air <NUM> passes across the fan blades <NUM>, a first portion of the air <NUM> as indicated by arrows <NUM> is directed or routed into the bypass airflow passage <NUM> and a second portion of the air <NUM> as indicated by arrow <NUM> is directed or routed into the LP compressor <NUM>. The ratio between the first portion of air <NUM> and the second portion of air <NUM> is commonly known as a bypass ratio. The pressure of the second portion of air <NUM> is then increased as it is routed through the high pressure (HP) compressor <NUM> and into the combustion section <NUM>, where it is mixed with fuel and burned to provide combustion gases <NUM>.

The combustion gases <NUM> are routed from the combustion section <NUM>, through the HP turbine <NUM> where a portion of thermal and/or kinetic energy from the combustion gases <NUM> is extracted via sequential stages of HP turbine stator vanes <NUM> that are coupled to the outer casing <NUM> and HP turbine rotor blades <NUM> that are coupled to the HP shaft or spool <NUM>, thus causing the HP shaft or spool <NUM> to rotate, thereby supporting operation of the HP compressor <NUM>. The combustion gases <NUM> are then routed through the LP turbine <NUM> where a second portion of thermal and kinetic energy is extracted from the combustion gases <NUM> via sequential stages of LP turbine stator vanes <NUM> that are coupled to the outer casing <NUM> and LP turbine rotor blades <NUM> that are coupled to the LP shaft or spool <NUM>, thus causing the LP shaft or spool <NUM> to rotate, thereby supporting operation of the LP compressor <NUM> and/or rotation of the fan <NUM>.

The combustion gases <NUM> are subsequently routed through the jet exhaust nozzle section <NUM> of the core turbine engine <NUM> to provide propulsive thrust. Simultaneously, a pressure of the first portion of air <NUM> is substantially increased as the first portion of air <NUM> is routed through the bypass airflow passage <NUM> before it is exhausted from a fan nozzle exhaust section <NUM> of the turbofan <NUM>, also providing propulsive thrust. The HP turbine <NUM>, the LP turbine <NUM>, and the jet exhaust nozzle section <NUM> at least partially define a hot gas path <NUM> for routing the combustion gases <NUM> through the core turbine engine <NUM>.

It should be appreciated, however, that the exemplary turbofan engine <NUM> depicted in <FIG> is by way of example only, and that in other exemplary embodiments, the turbofan engine <NUM> may have any other suitable configuration. Additionally, or alternatively, aspects of the present disclosure may be utilized with any other suitable aeronautical gas turbine engine, such as a turboshaft engine, turboprop engine, turbojet engine, etc. Moreover, aspects of the present disclosure may further be utilized with any other land-based gas turbine engine, such as a power generation gas turbine engine, or any aeroderivative gas turbine engine, such as a nautical gas turbine engine.

Referring now to <FIG>, a close-up, side, cross-sectional view is provided of a combustor assembly <NUM> and turbine in accordance with an exemplary embodiment of the present disclosure. In at least certain exemplary aspects, the combustor assembly <NUM> of <FIG> may be positioned in the combustion section <NUM> of the exemplary turbofan engine <NUM> of <FIG>, and similarly, the turbine of <FIG> may be positioned in the turbine section of the exemplary turbofan engine <NUM> of <FIG>.

As shown, the combustor assembly <NUM> generally includes an inner liner <NUM> extending between an aft end <NUM> and a forward end <NUM> generally along the axial direction A, as well as an outer liner <NUM> also extending between an aft end <NUM> and a forward end <NUM> generally along the axial direction A. The inner and outer liners <NUM>, <NUM> together at least partially define a combustion chamber <NUM> therebetween. The inner and outer liners <NUM>, <NUM> are each attached to or formed integrally with an annular dome. More particularly, the annular dome includes an inner dome section <NUM> formed integrally with the forward end <NUM> of the inner liner <NUM> and an outer dome section <NUM> formed generally with the forward end <NUM> of the outer liner <NUM>. Further, the inner and outer dome section <NUM>, <NUM> may each be formed integrally (or alternatively may be formed of a plurality of components attached in any suitable manner) and may each extend along the circumferential direction C (see <FIG>) to define an annular shape.

For the embodiment depicted, the inner liner <NUM> and the outer liner <NUM> are each formed of a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability. Exemplary CMC materials utilized for such liners <NUM>, <NUM> may include silicon carbide, silicon, silica or alumina matrix materials and combinations thereof. However, in other exemplary embodiments, one or both of the inner liner <NUM> and outer liner <NUM> may instead be formed of any other suitable material, such as a suitable metal material.

Additionally, it should be appreciated that in other embodiments, the combustor assembly <NUM> may not include the inner and/or outer dome sections <NUM>, <NUM>; may include separately formed inner and/or outer dome sections <NUM>, <NUM> attached to the respective inner liner <NUM> and outer liner <NUM>; or may have any other suitable configuration.

Referring still to <FIG>, the combustor assembly <NUM> further includes a plurality of fuel air mixers <NUM> spaced along the circumferential direction C (see <FIG>, below) and positioned at least partially within the annular dome. More particularly, the plurality of fuel air mixers <NUM> are disposed at least partially between the outer dome section <NUM> and the inner dome section <NUM> along the radial direction R. Compressed air from the compressor section of the turbofan engine <NUM> flows into or through the fuel air mixers <NUM>, where the compressed air is mixed with fuel and ignited to create the combustion gases <NUM> within the combustion chamber <NUM>. The inner and outer dome sections <NUM>, <NUM> are configured to assist in providing such a flow of compressed air from the compressor section into or through the fuel air mixers <NUM>. For example, the outer dome section <NUM> includes an outer cowl <NUM> at a forward end and the inner dome section <NUM> similarly includes an inner cowl <NUM> at a forward end. The outer cowl <NUM> and inner cowl <NUM> may assist in directing the flow of compressed air from the compressor section <NUM> into or through one or more of the fuel air mixers <NUM>. Again, however, in other embodiments, the annular dome may be configured in any other suitable manner.

Additionally, as is discussed above, the combustion gases <NUM> flow from the combustion chamber <NUM> into and through the turbine section of the turbofan engine <NUM>, where a portion of thermal and/or kinetic energy from the combustion gases <NUM> is extracted via sequential stages of turbine stator vanes and turbine rotor blades. Notably, the turbine depicted in <FIG> is configured as an HP turbine <NUM>, located immediately downstream of the combustion chamber <NUM> defined by the combustor assembly <NUM> of the combustion section <NUM>.

As is depicted, the exemplary HP turbine <NUM> of <FIG> includes a first stage of turbine nozzles <NUM> positioned at a forward end of the HP turbine <NUM>, at a location downstream of, or rather immediately downstream of, the combustion chamber <NUM> of the combustor assembly <NUM>. Additionally, the first stage of turbine nozzles <NUM> is positioned immediately upstream of a first stage of turbine rotor blades <NUM>. As will be described in greater detail below, the first stage of turbine nozzles <NUM> is configured to orient the combustion gases <NUM> from the combustion chamber <NUM> in a desired flow direction to increase a performance of the HP turbine <NUM>. For the embodiment depicted, the first stage of turbine nozzles <NUM> includes a plurality of individual turbine nozzles spaced along the circumferential direction C (see <FIG>) and extending generally along the radial direction R from an inner turbine liner <NUM> to an outer turbine liner <NUM>. The inner and outer turbine liners <NUM>, <NUM> at least partially define a portion of the core air flowpath <NUM> extending through the HP turbine <NUM> of the turbine section. The outer turbine liner <NUM> is, for the embodiment depicted, coupled to the outer liner <NUM> of the combustor assembly <NUM> at a forward end and extends aftwardly/downstream past the first stage of turbine rotor blades <NUM>.

As is also depicted in <FIG>, the gas turbine engine further includes a casing at least partially enclosing the outer liner <NUM> of the combustor assembly <NUM> and the outer turbine liner <NUM> of the HP turbine <NUM>. More specifically, the gas turbine engine further includes a combustor casing <NUM> at least partially enclosing the outer liner <NUM> of the combustor assembly <NUM>, as well as a turbine casing <NUM> at least partially enclosing the outer turbine liner <NUM> of the HP turbine <NUM>. Each of the combustor casing <NUM> and turbine casing <NUM> defines a first side <NUM> proximate to and facing the respective liners <NUM>, <NUM> (i.e., a radially inner side), a second side <NUM> opposite the first side <NUM> (i.e., a radially outer side), and a casing opening <NUM>. Additionally, for the embodiment depicted, the outer liner <NUM> of the combustor assembly <NUM> and the outer turbine liner <NUM> of the HP turbine <NUM> also similarly include a liner opening <NUM>. The liner openings <NUM> of the liners <NUM>, <NUM> are substantially aligned with the casing openings <NUM> of the combustor casing <NUM> and turbine casing <NUM>, respectively. It should be appreciated, that for the embodiment depicted, the casing openings <NUM> of the combustor casing <NUM> and turbine casing <NUM>, as well as the liner openings <NUM> of the outer liner <NUM> and outer turbine liner <NUM> are configured as part of respective borescope openings, also referred to as borescope inspection ports. For example, the gas turbine engine depicted includes a borescope plug <NUM> positioned in the casing opening <NUM> of the turbine casing <NUM> and in the liner opening <NUM> of the outer turbine liner <NUM>.

As will be appreciated, it may be beneficial during operation of the gas turbine engine to monitor a dynamic pressure within the core air flowpath <NUM> of the gas turbine engine. More specifically, it may be beneficial during operation of the gas turbine engine to monitor a dynamic pressure within the combustion chamber <NUM> of the combustor assembly <NUM> and/or within the core air flowpath <NUM> at a forward end of the HP turbine <NUM>. Monitoring the dynamic pressure within these sections of the gas turbine engine may allow for the gas turbine engine to monitor any combustor dynamics therein, and if necessary, modify operation of the gas turbine engine to minimize such combustor dynamics.

Accordingly, for the embodiment depicted, the gas turbine engine further includes a pressure sensor assembly <NUM> configured to monitor a pressure within the combustion chamber <NUM> of the combustor assembly <NUM> or within the core air flowpath <NUM> in the HP turbine <NUM> (e.g., at a forward end of the HP turbine <NUM>). More specifically, for the embodiment depicted in <FIG>, the pressure sensor assembly <NUM> is configured to monitor a dynamic pressure within the combustion chamber <NUM> of the combustor assembly <NUM>.

Referring now also to <FIG>, providing a close-up, cross-sectional view of the exemplary pressure sensor assembly <NUM> of <FIG>, the pressure sensor assembly <NUM> generally includes a body <NUM> and an extension member <NUM>. The body <NUM> is positioned at least partially on the second side <NUM> of the combustor casing <NUM>. More specifically, for the embodiment depicted, the body <NUM> is positioned completely on the second side <NUM> of the combustor casing <NUM> (i.e., does not extend into or through the casing opening <NUM>). By contrast, the extension member <NUM> extends from the body <NUM> at least partially through the casing opening <NUM> in the combustor casing <NUM> and towards the liner opening <NUM> and the outer liner <NUM> of the combustor assembly <NUM>.

Additionally, for the embodiment depicted, the pressure sensor assembly <NUM> is removably coupled to the combustor casing <NUM>. More specifically, the extension member <NUM> of the pressure sensor assembly <NUM> defines a threaded outer surface <NUM>. The threaded outer surface <NUM> of the pressure sensor assembly <NUM> is configured to engage a threaded section <NUM> of the casing opening <NUM> in the combustor casing <NUM>. Moreover, for the embodiment depicted, the pressure sensor assembly <NUM> includes a secondary retention and anti-rotation member <NUM> to ensure the pressure sensor assembly <NUM> does not come unscrewed/ unattached during engine operation. For the embodiment depicted, the member <NUM> is configured as a tab-washer. However, in other embodiments, any other suitable member may be provided. Notably, however, in other exemplary embodiments, an outer surface of the body <NUM> may instead define a threaded section configured to removably attach the pressure sensor assembly <NUM> to a correspondingly threaded section of the casing (e.g., casing <NUM>). Additionally, or alternatively, the pressure sensor assembly <NUM> may be removably coupled to the casing (e.g., casing <NUM>) in any other suitable manner, or alternatively still, may be permanently coupled to the casing (e.g., casing <NUM>).

Additionally, for the embodiment depicted, the extension member <NUM> is moveably engaged with the outer liner <NUM> of the combustor assembly <NUM>. More specifically, the outer liner <NUM> of the combustor assembly <NUM> includes a ferrule <NUM> and the extension member <NUM> extends into the ferrule <NUM>. The ferrule <NUM> includes a first radially outer member <NUM> and a second radially inner member <NUM>. The extension member <NUM> is slidably received through the radially outer member <NUM> such that the extension member <NUM> may move generally along the radial direction R relative to the radially outer member <NUM>. Additionally, for the embodiment depicted, the radially outer member <NUM> is slidably connected to the radially inner member <NUM>, allowing the radially outer member <NUM> to move generally along the axial direction A and in the circumferential direction C relative to the radially inner member <NUM>. Such a configuration effectively gives the extension member <NUM> six degrees of freedom relative to the outer liner <NUM> of the combustor assembly <NUM>.

As may also be seen in <FIG>, the pressure sensor assembly <NUM> further includes a pressure sensor <NUM>. The pressure sensor <NUM> is positioned at least partially within the body <NUM> and, for the embodiment depicted, removably attached thereto. More specifically, the pressure sensor assembly <NUM> includes a clamp nut <NUM>, the clamp nut <NUM> rotatably engaged with the body <NUM> through a threaded connection <NUM> to removably couple the pressure sensor <NUM> at least partially within the body <NUM>. Again, however, the clamp nut <NUM> may additionally, or alternatively, be removably coupled to the body <NUM> in any other suitable manner.

In an illustrative example not falling within the scope of the invention, the pressure sensor <NUM> is configured as a piezoelectric sensor. More specifically, the pressure sensor <NUM> includes a diaphragm <NUM> and a piezoelectric material <NUM> positioned against the diaphragm <NUM>. Further, as is shown, the extension member <NUM> defines a continuous sense cavity <NUM> for exposing the pressure sensor <NUM> to the core air flowpath <NUM>, or more specifically, for exposing the pressure sensor <NUM> to the combustion chamber <NUM> of the combustor assembly <NUM>. Additionally, the extension member <NUM> defines one or more cooling holes <NUM> to reduce a temperature of the air within the sense cavity <NUM> of the extension member <NUM>. The cooling holes <NUM> are positioned inward of the combustor casing <NUM> along the radial direction R and outward of the outer liner <NUM> along the radial direction R. Accordingly, the cooling holes <NUM> are exposed to a compressor discharge air flowing between the combustor casing <NUM> and outer liner <NUM>. A pressure of the compressor discharge air may generally be higher than a pressure of the air/combustion gases within the combustion chamber <NUM>. Additionally, a temperature of the compressor discharge air may generally be lower than a temperature of the air/combustion gases within the combustion chamber <NUM>. Accordingly, with the configuration depicted in <FIG>, the sense cavity <NUM> may fill with compressor discharge air, maintaining a temperature of the pressure sensor <NUM> at a lower temperature relative to a temperature within the combustion chamber <NUM>, and below a maximum operating temperature threshold.

Referring still to <FIG>, the diaphragm <NUM> is positioned adjacent to and exposed to the sense cavity <NUM> of the extension member <NUM> and the piezoelectric material <NUM> is positioned against the diaphragm <NUM> opposite the sense cavity <NUM> of the extension member <NUM>. The diaphragm <NUM> depicted is configured as a single layer of material, however, the diaphragm <NUM> may instead be formed of a plurality of layers of materials. Accordingly, during operation of the gas turbine engine, the pressure fluctuations within the combustion chamber <NUM> are translated through the compressor exit air within the sense cavity <NUM> of the extension member <NUM> to the diaphragm <NUM> of the pressure sensor <NUM>, and further to the piezoelectric material <NUM>.

More specifically, during operation of the gas turbine engine, pressure fluctuations within the combustion chamber <NUM> may compress air positioned within the sense cavity <NUM> of the extension member <NUM>, translating such pressure fluctuations to the diaphragm <NUM> of the pressure sensor <NUM>. The pressure fluctuations translated to the diaphragm <NUM> cause the diaphragm <NUM> to deform. The deformation of the diaphragm <NUM>, in turn, causes the piezoelectric material <NUM> to deform, and the deformation of the piezoelectric material <NUM> generates an electrical signal. The electrical signal is provided through an electrical connection line <NUM> of the pressure sensor <NUM> to, e.g., a controller (not shown) which may correlate the electrical signal to a pressure within the combustion chamber <NUM> of the gas turbine engine. As will be appreciated, such a pressure sensor <NUM> may be capable of differentiating between different frequencies and amplitudes of pressure changes/dynamics to which the diaphragm <NUM> is exposed, allowing the pressure sensor <NUM> to sense a specific frequency and amplitude of combustor dynamics within the combustion chamber <NUM> of the gas turbine engine, if desired.

Moreover, including a pressure sensor within a pressure sensor assembly may allow for the use of piezoelectric material to measure a pressure within the combustion chamber <NUM> of the gas turbine engine despite the elevated temperatures within the combustion chamber <NUM> the gas turbine engine that would otherwise prevent use of piezoelectric material.

It should be appreciated, however, that the exemplary pressure sensor assembly <NUM> depicted in <FIG> and <FIG> is provided by way of example only. In other exemplary embodiments, the pressure sensor assembly <NUM> may have any other suitable configuration, and further may be positioned at any other suitable location within the gas turbine engine.

For example, reference will now be made to <FIG>, providing a gas turbine engine including a pressure sensor assembly <NUM> in accordance with another exemplary embodiment of the present disclosure. It should be appreciated that the exemplary embodiment of <FIG> may be configured in substantially the same manner as the exemplary embodiment described above with reference to <FIG> and <FIG>. Accordingly, the same or similar numbers may refer to the same or similar part.

As is depicted, for the exemplary embodiment of <FIG>, the gas turbine engine includes a pressure sensor assembly <NUM>, the pressure sensor assembly <NUM> including a body <NUM>, an extension member <NUM>, and a pressure sensor <NUM>. The pressure sensor <NUM> is positioned at least partially within the body <NUM>, and the body <NUM> is positioned at least partially on a second, a radially outer side <NUM> of a casing. Additionally, the extension member <NUM> extends from the body <NUM> through a casing opening <NUM> in the casing and towards a liner opening <NUM> in a liner. The extension member <NUM> defines a continuous sense cavity <NUM> exposing the pressure sensor <NUM> to a core air flowpath <NUM>.

However, for the embodiment depicted, the casing is instead configured as a turbine casing <NUM> and the liner is instead configured as an outer turbine liner <NUM>. Notably, the casing opening <NUM> in the turbine casing <NUM> and the liner opening <NUM> in the outer turbine liner <NUM> are each positioned proximate a first stage of turbine nozzles <NUM> (i.e., radially outward of the first stage of turbine nozzles <NUM>, and aligned with the first stage of turbine nozzles <NUM> along the axial direction A). Additionally, for the exemplary embodiment depicted the pressure sensor assembly <NUM> has replaced a borescope plug <NUM>, and a borescope plug <NUM> has been positioned in the casing and liner openings <NUM>, <NUM> of the combustor casing <NUM> and outer liner <NUM>, respectively (compare with <FIG>).

Accordingly, the pressure sensor <NUM> of the pressure sensor assembly <NUM> of <FIG> may still be configured to detect combustor dynamics within the combustion chamber <NUM>. It should be appreciated, however, in other exemplary embodiments the pressure sensor assembly <NUM> may instead be positioned at any other suitable location within the turbine section, such as at any location forward of the first stage of turbine rotor blades <NUM> of the HP turbine <NUM>, or aft of the first stage of turbine rotor blades <NUM> of the HP turbine <NUM>.

Additionally, referring now to <FIG>, a side, cross-sectional view of yet another exemplary embodiment of a pressure sensor assembly <NUM> is provided. As with the exemplary pressure sensor assembly <NUM> described above with reference to <FIG> and <FIG>, the exemplary pressure sensor assembly <NUM> of <FIG> includes a body <NUM>, an extension member <NUM>, and a pressure sensor <NUM>. The pressure sensor <NUM> is positioned at least partially within the body <NUM> and generally includes a diaphragm <NUM> positioned adjacent to and exposed to the sense cavity <NUM> of the extension member <NUM>.

However, for the exemplary embodiment depicted, the pressure sensor <NUM> is an optical-based sensor configured to measure a deflection of the diaphragm <NUM> to, in turn, measure pressure within the combustion chamber <NUM>. More specifically, for the embodiment depicted, the diaphragm <NUM> defines an enclosed, interior cavity <NUM>. Additionally, the optical-based sensor includes an optical laser <NUM> directed at the diaphragm <NUM>, or rather, directed at the interior cavity <NUM> of the diaphragm <NUM>. The optical-based sensor is configured to measure light from the optical laser <NUM> reflected from the internal cavity <NUM> of the diaphragm <NUM>. Deformation of the diaphragm <NUM>, and deformation of the internal cavity <NUM> of the diaphragm <NUM>, changes characteristics of the light from the optical laser <NUM> reflected from the internal cavity <NUM> the diaphragm <NUM>. Accordingly, the optical-based sensor may determine an amount of deflection/deformation of the diaphragm <NUM> based on the sensed characteristics of the light from the optical laser <NUM> reflected from the cavity <NUM>, therefore determining a pressure within the core air flowpath <NUM>.

Moreover, it should be appreciated that in other exemplary embodiments, the pressure sensor assembly <NUM> including an optical-based sensor may be configured in any other suitable manner. For example, referring to <FIG>, a side, cross-sectional view of yet another exemplary embodiment of a pressure sensor assembly <NUM> is provided. The pressure sensor assembly <NUM> of <FIG> may be configured in substantially the same manner as the pressure sensor assembly <NUM> of <FIG>. However, for the embodiment of <FIG>, a diaphragm <NUM> of the pressure sensor assembly <NUM> is positioned in a sense cavity <NUM> of the extension member <NUM>, such that an edge of the diaphragm <NUM> is substantially flush with an inner surface <NUM> of the casing <NUM>. Notably, in still other embodiments, the diaphragm <NUM> may further be positioned even closer to the combustion chamber <NUM>. For example in still other exemplary embodiments the diaphragm <NUM> may be positioned radially inward of the casing <NUM>, e.g., such that an edge of the diaphragm <NUM> is substantially flush with a radially outer surface of the outer liner <NUM>.

Furthermore, although for the embodiments described above the gas turbine engine is depicted including a single pressure sensor assembly <NUM>, it should be appreciated that in other exemplary embodiments, the gas turbine engine may additionally include a plurality pressure sensor assemblies <NUM>. For example, referring now briefly to <FIG>, a schematic, axial view of a combustion section <NUM> of a gas turbine engine in accordance with another exemplary embodiment of the present disclosure is provided. The combustion section <NUM> may be configured in substantially the same manner as the exemplary combustion section <NUM> described above with reference to <FIG>. For example, the combustion section <NUM> generally includes an outer combustor casing <NUM> enclosing a combustor assembly <NUM>, the combustor assembly <NUM> including an outer liner <NUM> at least partially defining a combustion chamber <NUM>. For the exemplary embodiment of <FIG>, the gas turbine engine includes a plurality of pressure sensor assemblies <NUM> arranged along the circumferential direction C of the gas turbine engine. More specifically, for the embodiment depicted, the gas turbine engine includes four pressure assemblies <NUM> arranged substantially evenly along the circumferential direction C of the gas turbine engine. Such a configuration may allow for the gas turbine engine to measure combustor dynamics within the combustion chamber <NUM> that may vary along the circumferential direction C.

It should be appreciated, however, that in other exemplary embodiments, the gas turbine engine may include any other suitable number of pressure sensor assemblies <NUM> arranged in other suitable manner. Additionally, each of the plurality pressure sensor assemblies <NUM> depicted in <FIG> may be configured in substantially the same manner as one or more of the exemplary pressure sensor assemblies <NUM> described above with reference to <FIG>. Accordingly, in certain exemplary embodiments, the plurality of pressure sensor assemblies <NUM> may instead be positioned within the turbine section of the gas turbine engine.

Inclusion of one or more pressure sensor assemblies in accordance with an exemplary embodiment of the present disclosure with a gas turbine engine may allow for the gas turbine engine to more accurately monitor combustor dynamics within e.g., a combustion chamber of a combustion section of the gas turbine engine, or at a forward end of the turbine section of the gas turbine engine. More specifically, inclusion of one or more pressure sensor assemblies in accordance with an exemplary embodiment of the present disclosure may allow for use of more accurate pressure sensors which may otherwise be incapable of withstanding the relatively elevated temperatures within the combustion chamber of the gas turbine engine.

Claim 1:
A pressure sensor assembly (<NUM>), for a gas turbine engine defining a radial direction (R) and a circumferential direction (C) and comprising a liner (<NUM>, <NUM>) at least partially defining a core air flowpath (<NUM>) through a compressor section or a turbine section and a casing (<NUM>, <NUM>) at least partially enclosing the liner, the pressure sensor assembly comprising (<NUM>):
a body (<NUM>) configured for positioning adjacent to the casing of the gas turbine engine;
a pressure sensor (<NUM>) positioned at least partially within the body and comprising a diaphragm (<NUM>); and
an extension member (<NUM>) extending from the body (<NUM>) and configured to extend at least partially through a casing opening in the casing and towards a liner opening in the liner,
wherein the extension member (<NUM>) defines a continuous sense cavity (<NUM>) extending between the pressure sensor (<NUM>) and the core air flowpath (<NUM>), to expose the pressure sensor (<NUM>) to the core air flowpath (<NUM>) of the gas turbine engine,
wherein the diaphragm (<NUM>) is positioned closer to an outer surface of the casing (<NUM>, <NUM>) in the radial direction than to an inner surface of the casing (<NUM>, <NUM>); characterised in that:
the diaphragm (<NUM>) is exposed to the sense cavity (<NUM>) of the extension member (<NUM>), and wherein the pressure sensor is an optical based sensor for measuring a deflection of the diaphragm (<NUM>).