Patent Description:
Gas turbine instrumentation such as, for example, infrared imaging sensors may be used to evaluate engine performance under operating conditions. Many systems include the imaging sensor coupled to the engine and a penetrating optical probe inserted through the engine case into the engine gas path. The presence of the probe within the gas path tends to disrupt gas path flow for downstream stages and may tend to reduce overall engine operating efficiency.

<CIT> discloses a device for measuring parameters of an aerodynamic flow of a turbomachine.

From a first aspect of the invention, a vane assembly as claimed in claim <NUM> is provided.

In various embodiments, the probe has a sensor coupled to the probe. In various embodiments, the thickened airfoil comprises a pressure side thickened region relative to the nominal airfoil at a pressure side, and wherein the first thinned airfoil is disposed circumferentially adjacent to the pressure side thickened region. In various embodiments, the thickened airfoil comprises a suction side thickened region relative to the nominal airfoil at a suction side. In various embodiments, the first thinned airfoil is disposed circumferentially adjacent to the suction side thickened region. In various embodiments, the thickened airfoil comprises a pressure side thickened region and a suction side thickened region relative to the nominal airfoil at a respective pressure side and a suction side. In various embodiments, the first thinned airfoil is disposed circumferentially adjacent to the pressure side thickened region and a second thinned airfoil is disposed circumferentially adjacent to the suction side thickened region. In various embodiments, the thickened airfoil and the first thinned airfoil each extend between a common inner platform and a common outer platform.

From a second aspect of the invention, a gas turbine engine as claimed in claim <NUM> is provided.

In various embodiments, the thickened airfoil comprises a pressure side thickened region relative to the nominal airfoil at a pressure side. In various embodiments, the first thinned airfoil is disposed circumferentially adjacent to the pressure side thickened region. In various embodiments, the thickened airfoil comprises a suction side thickened region relative to the nominal airfoil at a suction side. In various embodiments, the first thinned airfoil is disposed circumferentially adjacent to the suction side thickened region. In various embodiments, the thickened airfoil comprises a pressure side thickened region and a suction side thickened region relative to the nominal airfoil at a respective pressure side and a suction side. In various embodiments, the first thinned airfoil is disposed circumferentially adjacent to the pressure side thickened region and a second thinned airfoil is disposed circumferentially adjacent to the suction side thickened region.

From a further aspect of the invention, a method of instrumenting a gas turbine engine as claimed in claim <NUM> is provided.

The detailed description of exemplary embodiments herein makes reference to the accompanying drawings, which show exemplary embodiments by way of illustration and their best mode. While these exemplary embodiments are described in sufficient detail to enable those skilled in the art to practice the disclosures, it should be understood that other embodiments may be realized and that logical, chemical, and mechanical changes may be made without departing from the scope of the disclosures. For example, the steps recited in any of the method or process descriptions may be executed in any order and are not necessarily limited to the order presented. Also, any reference to attached, fixed, connected or the like may include permanent, removable, temporary, partial, full and/or any other possible attachment option.

In various embodiments and with reference to <FIG>, a gas turbine engine <NUM> is provided. Gas turbine engine <NUM> may be a two-spool turbofan that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. In operation, fan section <NUM> can drive air along a bypass flow-path B while compressor section <NUM> can drive air for compression and communication into combustor section <NUM> then expansion through turbine section <NUM>. Although depicted as a turbofan gas turbine engine <NUM> herein, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including turbojet engines, a low-bypass turbofans, a high bypass turbofans, or any other gas turbine known to those skilled in the art.

Gas turbine engine <NUM> may generally comprise a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis A-A' relative to an engine static structure <NUM> via one or more bearing systems <NUM> (shown as bearing system <NUM>-<NUM> and bearing system <NUM>-<NUM>). It should be understood that various bearing systems <NUM> at various locations may alternatively or additionally be provided, including for example, bearing system <NUM>, bearing system <NUM>-<NUM>, and bearing system <NUM>-<NUM>.

Low speed spool <NUM> may generally comprise an inner shaft <NUM> that interconnects a fan <NUM>, a low pressure (or first) compressor section <NUM> (also referred to a low pressure compressor) and a low pressure (or first) turbine section <NUM>. Inner shaft <NUM> may be connected to fan <NUM> through a geared architecture <NUM> that can drive fan <NUM> at a lower speed than low speed spool <NUM>. Geared architecture <NUM> may comprise a gear assembly <NUM> enclosed within a gear housing <NUM>. Gear assembly <NUM> couples inner shaft <NUM> to a rotating fan structure. High speed spool <NUM> may comprise an outer shaft <NUM> that interconnects a high pressure compressor ("HPC") <NUM> (e.g., a second compressor section) and high pressure (or second) turbine section <NUM>. A combustor <NUM> may be located between HPC <NUM> and high pressure turbine <NUM>. A mid-turbine frame <NUM> of engine static structure <NUM> may be located generally between high pressure turbine <NUM> and low pressure turbine <NUM>. Mid-turbine frame <NUM> may support one or more bearing systems <NUM> in turbine section <NUM>. Inner shaft <NUM> and outer shaft <NUM> may be concentric and rotate via bearing systems <NUM> about the engine central longitudinal axis A-A', which is collinear with their longitudinal axes. As used herein, a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure" compressor or turbine.

The core airflow C may be compressed by low pressure compressor <NUM> then HPC <NUM>, mixed and burned with fuel in combustor <NUM>, then expanded over high pressure turbine <NUM> and low pressure turbine <NUM>. Mid-turbine frame <NUM> includes airfoils <NUM> which are in the core airflow path. Low pressure turbine <NUM>, and high pressure turbine <NUM> rotationally drive the respective low speed spool <NUM> and high speed spool <NUM> in response to the expansion.

Gas turbine engine <NUM> may be, for example, a high-bypass geared aircraft engine. In various embodiments, the bypass ratio of gas turbine engine <NUM> may be greater than about six (<NUM>). In various embodiments, the bypass ratio of gas turbine engine <NUM> may be greater than ten (<NUM>). In various embodiments, geared architecture <NUM> may be an epicyclic gear train, such as a star gear system (sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear) or other gear system. Geared architecture <NUM> may have a gear reduction ratio of greater than about <NUM> and low pressure turbine <NUM> may have a pressure ratio that is greater than about <NUM>. In various embodiments, the bypass ratio of gas turbine engine <NUM> is greater than about ten (<NUM>:<NUM>). In various embodiments, the diameter of fan <NUM> may be significantly larger than that of the low pressure compressor <NUM>, and the low pressure turbine <NUM> may have a pressure ratio that is greater than about (<NUM>:<NUM>). Low pressure turbine <NUM> pressure ratio may be measured prior to inlet of low pressure turbine <NUM> as related to the pressure at the outlet of low pressure turbine <NUM> prior to an exhaust nozzle. It should be understood, however, that the above parameters are exemplary of various embodiments of a suitable geared architecture engine and that the present disclosure contemplates other gas turbine engines including direct drive turbofans.

In various embodiments, the next generation of turbofan engines may be designed for higher efficiency which is associated with higher pressure ratios and higher temperatures in the HPC <NUM>. These higher operating temperatures and pressure ratios may create operating environments that may cause thermal loads that are higher than the thermal loads encountered in conventional turbofan engines, which may shorten the operational life of current components.

In various embodiments, low pressure compressor <NUM>, HPC <NUM>, high pressure turbine <NUM>, and low pressure turbine <NUM> may comprise alternating rows of rotating rotors and stationary stators. Stators may have a cantilevered configuration or a shrouded configuration. More specifically, a stator may comprise a stator vane, a casing support and a hub support. In this regard, a stator vane may be supported along an outer diameter by a casing support and along an inner diameter by a hub support. In contrast, a cantilevered stator may comprise a stator vane that is only retained and/or supported at the casing (e.g., along an outer diameter). In various embodiments, a stator may include a vane assembly such as vane instrumentation assembly <NUM>.

In various embodiments, rotors may be configured to compress (or expand) and spin a fluid flow. Stators may be configured to receive and straighten the fluid flow. In operation, the fluid flow discharged from the trailing edge of stators may be straightened (e.g., the flow may be directed in a substantially parallel path to the centerline of the engine and/or HPC) to increase and/or improve the efficiency of the engine and, more specifically, to achieve maximum and/or near maximum compression (or expansion) and efficiency when the straightened air is compressed (or expanded) and spun by rotor <NUM>.

In various embodiments and with additional reference to <FIG>, a vane instrumentation assembly <NUM> is illustrated. Vane instrumentation assembly <NUM> comprises a sensor <NUM> coupled to a probe <NUM>. Sensor <NUM> may be in electronic communication with a Data Acquisition (DAQ) system <NUM> configured to receive sensor data from the sensor <NUM>. Vane instrumentation assembly <NUM> comprises a plurality of airfoils <NUM> each extending into gas path <NUM> between an inner platform <NUM> and an outer platform <NUM>. The plurality of airfoils <NUM> include a nominal airfoil, a thickened airfoil, and a thinned airfoil. In various embodiments, the probe <NUM> may be inserted at penetrations <NUM> through outer case <NUM> and inner case <NUM> into one of the plurality of airfoils <NUM>. In various embodiments, the probe <NUM> may be an optical probe comprising a probe head <NUM> having a field of view <NUM>. In various embodiments, the probe head may include an optical fiber and a mirror assembly and/or a lens assembly which may orient or turn the field of view toward a region of interest.

The probe head <NUM> may be aligned with a window <NUM> of the airfoil providing the field of view <NUM> of the region of interest within the cases (<NUM>, <NUM>) such as, for example an airfoil <NUM> of downstream rotor blade <NUM>. Although field of view <NUM> is illustrated as oriented toward airfoil <NUM> of downstream rotor blade <NUM> it will be appreciated that window <NUM> and probe head <NUM> may be positioned to orient field of view <NUM> at any region of interest such as, for example, airfoil <NUM> of upstream rotor blade <NUM>, blade outer air seals <NUM>, blade platforms <NUM> and/or the like.

In various embodiments and with additional reference to <FIG>, nominal airfoils <NUM> of the plurality of airfoils <NUM> are illustrated in cross section through the gas path <NUM>. Probe <NUM> is disposed within the nominal airfoil <NUM>. The nominal airfoil <NUM> has a first chord thickness W1. A gas path flow from the upstream stage encounters the nominal airfoils <NUM> and is turned between the respectively adjacent pressure side <NUM> and suction side <NUM> of nominal airfoils <NUM> as shown by streamlines <NUM>. In various embodiments, the probe <NUM> may protrude from the suction side <NUM> of the nominal airfoil <NUM> tending thereby to disrupt the flow streamlines <NUM> and, in response, generating a separated flow as shown by arrows <NUM>. Airfoils <NUM> of the downstream rotor blades <NUM> encounter the separated flow tending thereby to induce pressure loss, aeroacoustic vibrations, sensor measurement error, and the like.

In various embodiments and with additional reference to <FIG>, a thickened airfoil <NUM> of the plurality of airfoils <NUM> is illustrated. Thickened airfoil <NUM> comprises a suction side thickened region <NUM> relative to the nominal airfoil <NUM> (shown by overlapping outline <NUM>') at a suction side <NUM>. Probe <NUM> is embedded within the suction side thickened region <NUM> of the thickened airfoil <NUM>. A thinned airfoil <NUM> (i.e., a first thinned airfoil) is disposed adjacent to the thickened airfoil <NUM> and is thinned at the pressure side <NUM> (i.e. reduced chord thickness relative to the nominal airfoil <NUM>) as illustrated by nominal airfoil outline <NUM>". In various embodiments the thickened airfoil <NUM> may be coupled circumferentially proximate (i.e., along a common arc described by an intersecting radial line from central longitudinal axis A-A') to the thinned airfoil <NUM> to extend from a common platform <NUM> (i.e., between a common inner platform and a common outer platform) to form a vane doublet.

As shown by outlines <NUM>' and <NUM>" each of the thickened airfoil <NUM> and the thinned airfoil <NUM> comprises a similar chord length and/or camber to the nominal airfoil <NUM> but differs in chord thickness. In various embodiments, each of the thickened airfoil <NUM>, the thinned airfoil <NUM>, and the nominal airfoil <NUM> may have an identical chord length. The thickened airfoil comprises a second chord thickness W2 and the thinned airfoil <NUM> comprises a third chord thickness W3. The second chord thickness W2 is greater than the first chord thickness W1 and the third chord thickness W3 is less than the first chord thickness W1. Stated another way, the second chord thickness W2 is greater than the third chord thickness W3 and the first chord thickness W1. In various embodiments, W2 may be between <NUM>% and <NUM>% greater than W1 and W3 may be between <NUM>% and <NUM>% of W1. In this regard, the gas path flow from the upstream stage encountering the thickened airfoil <NUM> and the thinned airfoil <NUM> may be turned as illustrated by streamlines <NUM> between the respective suction side <NUM> and pressure side <NUM> tending thereby to inhibit generating the separated flow. In this regard, the thickened airfoil <NUM> and the adjacent thinned airfoil <NUM> are configured to turn the gas path flow therebetween to an identical trailing edge exit angle (defined between the trailing edge exit plane <NUM> and the streamlines <NUM>) as the nominal airfoil <NUM>. In various embodiments, each of the thickened airfoil <NUM>, the thinned airfoil <NUM>, and the nominal airfoil <NUM> may have respective leading edges and trailing edges aligned along a common arc described by an intersecting radial line from central longitudinal axis A-A'. In this regard, each of the thickened airfoil <NUM>, the thinned airfoil <NUM>, and the nominal airfoil <NUM> may have an identical leading edge position and/or trailing edge position along the longitudinal axis.

In various embodiments and with additional reference to <FIG> a thickened airfoil <NUM> of the plurality of airfoils <NUM> is illustrated. Thickened airfoil <NUM> comprises features, geometries, construction, manufacturing techniques, and/or internal components similar to thickened airfoil <NUM>. Thickened airfoil <NUM> differs in comprising a pressure side thickened region <NUM> relative to the nominal airfoil <NUM> (shown by overlapping outline <NUM>') at a pressure side <NUM>. Probe <NUM> is embedded within the pressure side thickened region <NUM> of the thickened airfoil <NUM>. A thinned airfoil <NUM> (i.e., a second thinned airfoil) is disposed adjacent to the thickened airfoil <NUM> and is thinned at the suction side <NUM> (i.e. reduce chord thickness relative to the nominal airfoil <NUM>) as illustrated by nominal airfoil outline <NUM>". In various embodiments the thickened airfoil <NUM> may be coupled circumferentially proximate the thinned airfoil <NUM> to extend from the common platform <NUM> to form the vane doublet.

As shown by outlines <NUM>' and <NUM>" each of the thickened airfoil <NUM> and the thinned airfoil <NUM> may comprise an identical chord length to the nominal airfoil <NUM> but may differ in chord thickness. In this regard, the gas path flow from the upstream stage encountering the thickened airfoil <NUM> and the thinned airfoil <NUM> may be turned as illustrated by streamlines <NUM> between the respective pressure side <NUM> and suction side <NUM> tending thereby to inhibit generating the separated flow. In this regard, the thickened airfoil <NUM> and the adjacent thinned airfoil <NUM> are configured to turn the gas path flow therebetween to an identical trailing edge exit angle (defined between the trailing edge exit plane <NUM> and the streamlines <NUM>) as the nominal airfoil <NUM>.

In various embodiments and with additional reference to <FIG>, a thickened airfoil <NUM> of the plurality of airfoils <NUM> is illustrated. Thickened airfoil <NUM> comprises features, geometries, construction, manufacturing techniques, and/or internal components similar to thickened airfoil <NUM> and thickened airfoil <NUM>. Thickened airfoil <NUM> differs in comprising both the pressure side thickened region <NUM> and the suction side thickened region <NUM> relative to the nominal airfoil <NUM> (shown by overlapping outlines <NUM>'). Probe <NUM> is embedded within the pressure side thickened region <NUM> and the suction side thickened region <NUM> of thickened airfoil <NUM>. Thinned airfoil <NUM> is disposed adjacent to the pressure side thickened region <NUM> and thinned airfoil <NUM> is disposed adjacent to the suction side thickened region <NUM>. In various embodiments, the thickened airfoil <NUM> may be coupled relatively circumferentially between the first thinned airfoil and the second thinned airfoil to and extend from a common platform <NUM> and thereby form a vane triplet.

As shown by outlines <NUM>' and <NUM>" each of the thickened airfoil <NUM> and the thinned airfoils (<NUM>, <NUM>) may comprise an identical chord length to the nominal airfoil <NUM> but may differ in chord thickness. For example, a thickened airfoil may have a chord thickness between <NUM>% and <NUM>% greater than the chord thickness of the nominal airfoil and a thinned airfoil may have a chord thickness between <NUM>% and <NUM>% less than the chord thickness of the nominal airfoil. In this regard, the gas path flow from the upstream stage encountering the thickened airfoil <NUM> and the thinned airfoils (<NUM>, <NUM>) may be turned as illustrated by streamlines <NUM> between the thickened airfoil <NUM> and each of the respectively adjacent thinned airfoils (<NUM>, <NUM>) tending thereby to inhibit generating the separated flow. In this regard, the thickened airfoil <NUM> and the adjacent thinned airfoils (<NUM>, <NUM>) are configured to turn the gas path flow therebetween to an identical trailing edge exit angle (defined between the trailing edge exit plane <NUM> and the streamlines <NUM>) as the nominal airfoil <NUM>.

In various embodiments, the distance between a thickened region of a thickened airfoil (such as suction side thickened region <NUM> of thickened airfoil <NUM>) and a corresponding surface of a thinned airfoil (such as pressure side <NUM> of thinned airfoil <NUM>) define a throat area. Stated another way and with brief reference to <FIG>, a throat area may be defined by the minimum distance T between the suction side of a first airfoil (such as suction side <NUM> of thickened airfoil <NUM>) and the pressure side of a radially adjacent airfoil (such as pressure side <NUM> of thinned airfoil <NUM>). In various embodiments, the throat area between a thickened airfoil and an adjacent thinned airfoil may be about <NUM>% of a throat area between adjacent nominal airfoils <NUM>. The throat area between a thickened airfoil and an adjacent thinned airfoil is within ± <NUM>% of the throat area between adjacent nominal airfoils <NUM>. The trailing edge exit angle for a thickened airfoil and an adjacent thinned airfoil is within about ± <NUM>% of a trailing edge exit angle for adjacent nominal airfoils <NUM>.

However, the benefits, advantages, solutions to problems, and any elements that may cause any benefit, advantage, or solution to occur or become more pronounced are not to be construed as critical, required, or essential features or elements of the disclosures.

The scope of the disclosures is accordingly to be limited by nothing other than the appended claims, in which reference to an element in the singular is not intended to mean "one and only one" unless explicitly so stated, but rather "one or more.

Claim 1:
A vane assembly (<NUM>), comprising:
a plurality of airfoils (<NUM>) each extending between an inner platform (<NUM>) and an outer platform (<NUM>),
characterised by the plurality of airfoils (<NUM>) comprising a nominal airfoil (<NUM>), a thickened airfoil (<NUM>; <NUM>; <NUM>), and a first thinned airfoil (<NUM>; <NUM>) circumferentially adjacent to the thickened airfoil (<NUM>; <NUM>; <NUM>),
wherein each of the thickened airfoil (<NUM>; <NUM>; <NUM>) and the first thinned airfoil (<NUM>; <NUM>) comprise a similar chord length and/or camber to the nominal airfoil (<NUM>);
wherein the nominal airfoil (<NUM>) has a first chord thickness (W1); the thickened airfoil (<NUM>; <NUM>; <NUM>) has a thickened region relative to the nominal airfoil and a second chord thickness (W2); and the first thinned airfoil (<NUM>; <NUM>) is thinned relative to the nominal airfoil and has a third chord thickness (W3);
wherein the second chord thickness (W2) is greater than the first chord thickness (W1) and the third chord thickness (W3) is less than the first chord thickness (W1); and
a probe (<NUM>) disposed within the thickened airfoil (<NUM>; <NUM>; <NUM>),
wherein a throat area between the thickened airfoil (<NUM>; <NUM>; <NUM>) and the first thinned airfoil (<NUM>; <NUM>) is within ± <NUM>% of a throat area between the nominal airfoil (<NUM>) and a circumferentially adjacent nominal airfoil (<NUM>) of the plurality of airfoils (<NUM>), and wherein the thickened airfoil (<NUM>; <NUM>; <NUM>) and the adjacent first thinned airfoil (<NUM>; <NUM>) are configured to turn a gas path flow therebetween to within ± <NUM>% of a trailing edge exit angle as the nominal airfoil (<NUM>).