Patent Description:
Gas turbine engines typically include a compressor, a combustor, and a turbine, with an annular flow path extending axially through each. Initially, air flows through the compressor where it is compressed or pressurized. The combustor then mixes and ignites the compressed air with fuel, generating hot combustion gases. These hot combustion gases are then directed from the combustor to the turbine where power is extracted from the hot gases by causing blades of the turbine to rotate.

The compressor and turbine sections include multiple rotors and stators configured to enable optimal operation. Gas turbine engines maintain an optimal clearance (distance) between the tips of the rotors and an outside diameter of a gas path within the turbine engine, and thereby provide the conditions necessary to achieve a desired performance.

<CIT> discloses a method and an apparatus for determining a clearance between relatively movable components using a plurality of optical fibres or a plurality of electrical circuits arranged within an abradable material. <CIT> discloses systems for monitoring wear of a component using a conductor embedded in the component.

Viewed from a first aspect the present invention provides in combination a wear indicator and a component of a gas turbine engine according to claim <NUM>.

In addition to the features described in claim <NUM>, optionally embodiments may include that the first component is configured to delaminate when impacted by a blade of the gas turbine engine.

In addition to one or more of the features described above, or as an alternative, further embodiments may include a second component of the wear indicator having a blind hole partially enclosing the first component, wherein the wear indicator is secured to a surface of the component of the gas turbine engine through the second component of the wear indicator.

In addition to one or more of the features described above, or as an alternative, further embodiments may include that the component is a blade outer air seal.

In addition to one or more of the features described above, or as an alternative, further embodiments may include a measurement device electrically connected to the first plate through a first lead line and electrically connected to the second plate through second lead line, wherein the measurement device is configured to determine the resistance of the first plate, the second plate, and the plurality of wires.

In addition to one or more of the features described above, or as an alternative, further embodiments may include that the second component of the wear indicator further comprises: a first side that delaminates when impacted by a blade of the gas turbine engine; a second side parallel to the first side, the second side being secured to the component of the gas turbine engine; and a mid-section interposed between the first side and the second side; wherein the mid-section is conical frustum in shape.

In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first component has a cylindrical shape.

In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first plate and the second plate compose a substantial portion of a curved surface of the cylindrical shape, wherein the first plate and the second plate are separated by a gap in the curved surface.

In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first plate further includes a plurality of first orifices and the second plate further includes a plurality of second orifices, and wherein each of the plurality of wires extends from a first orifice to a second orifices.

Viewed from a second aspect the present invention provides a method of detecting blade clearance in a gas turbine engine according to claim <NUM>.

The detailed description explains embodiments of the present disclosure, together with advantages and features, by way of example with reference to the drawings.

In one disclosed embodiment, the engine <NUM> bypass ratio is greater than about ten (<NUM>:<NUM>), the fan diameter is significantly larger than that of the low pressure compressor <NUM>, and the low pressure turbine <NUM> has a pressure ratio that is greater than about five <NUM>: <NUM>. The geared architecture <NUM> may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM>. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.

Referring now to <FIG>, which shows a cross-sectional view of a rub button or wear indicator <NUM> installed in a gas turbine engine <NUM>, in accordance with an embodiment of the disclosure. As seen in <FIG>, the wear indicator <NUM> is attached to an inner surface <NUM> of the gas turbine engine <NUM> opposite a blade <NUM> of the gas turbine engine <NUM>. The blade <NUM> rotates along a blade path BP1. In the illustrated embodiment, the wear indicator <NUM> is attached to a blade outer air seal <NUM>, the outer air seal <NUM> is the inner surface <NUM>. One or more wear indicators <NUM> may be affixed to the inner surface <NUM> of the gas turbine engine <NUM> in order to monitor the clearance between the blade <NUM> and the inner surface <NUM>, a method discussed further below in <FIG>. In the embodiment of <FIG>, the wear indicator <NUM> has been installed in the high pressure turbine <NUM> of the gas turbine engine <NUM>. It is understood that the wear indicator <NUM> may be located in other sections of the gas turbine engine <NUM> having rotating blades <NUM>. An abradable coating <NUM> may be applied on the inner surface <NUM> of the gas turbine engine <NUM> and the wear indicator <NUM> may be covered by the abradable coating <NUM> on the inner surface <NUM>. The abradable coating <NUM> is designed to provide protection for the inner surface <NUM> against a blade <NUM> strike. If a blade <NUM> were to extend towards the inner surface <NUM> then the abradable coating <NUM> shall be struck first and absorb the impact of the blade <NUM> to prevent damage to the inner surface <NUM>. The wear indicator <NUM> may be attached to the inner surface <NUM> using an adhesive (not shown) that may or may not need a curing to adhere the wear indicator <NUM> to the inner surface <NUM>.

As seen in <FIG>, the wear indicator <NUM> may comprise a first component <NUM> and a second component <NUM> configured to partially enclose the first component <NUM>. The first component includes a plurality of wires <NUM> extending from a first plate <NUM> to a second plate <NUM>, as seen in <FIG>. The first plate <NUM> includes a plurality of first orifices <NUM> and the second plate <NUM> includes a plurality of second orifices <NUM>. The plurality of second orifices <NUM> are complimentary to the plurality of first orifices <NUM>. Each wire <NUM> extends from a first orifice <NUM> to a second orifice <NUM>. The plurality of wires <NUM> extend across a cavity <NUM> between the first plate <NUM> and the second plate <NUM>. The first component <NUM> may be partially filled with a potting material <NUM>. The potting material <NUM> is configured to fill voids <NUM> between each of the plurality of wires <NUM>, such that the plurality of wires <NUM> are electrically insulated from each other by the potting material <NUM>. The potting material <NUM> is non-conductive and capable of withstanding the high temperatures of a gas turbine engine. The potting material <NUM> may only partially fill the cavity <NUM> a portion D1. The first component <NUM> may have a round shape and/or a cylindrical shape, as seen in <FIG>. It is understood that the round and/or cylindrical shape of the first component <NUM> show in <FIG> is not intended to be limiting and the first component <NUM> may have a variety of different shapes. The first plate <NUM> and the second plate <NUM> compose a substantial portion of a curved surface <NUM> of the cylindrical shape of the first component <NUM>. It is also understood that the curved shape of the plates <NUM>, <NUM> show in <FIG> is not intended to be limiting and the plates <NUM>, <NUM> may have a variety of different shapes. The first plate <NUM> and the second plate <NUM> are separated by the gap D2 in the curved surface <NUM>. Advantageously, the gap D2 keeps the first plate <NUM> electrically separate from the second plate <NUM>.

The first plate <NUM> may be connected to a first lead line <NUM> and the second plate <NUM> may be connected to a second lead line <NUM>. The lead lines <NUM>, <NUM> may be connected to a measurement device <NUM> configured to measure the resistance through the plates <NUM>, <NUM> and plurality of wires <NUM>. The measurement device <NUM> may include a processor and a memory. For ease of illustration, the processor and memory are not shown in <FIG>. The processor can be any type or combination of computer processors, such as a microprocessor, microcontroller, digital signal processor, application specific integrated circuit, programmable logic device, and/or field programmable gate array. The memory is an example of a non-transitory computer readable storage medium tangibly embodied in or operably connected to the path determination system including executable instructions stored therein, for instance, as firmware.

Each of the plates <NUM>, <NUM> have a known resistance and each of the plurality of wires <NUM> have a known resistance. Advantageously, since the plurality of wires <NUM> and the plates <NUM>, <NUM> each have a known resistance then as the blade <NUM> cuts into the wear indicator <NUM> removing some of the plurality of wires <NUM> then the depth that the blade <NUM> cut into the wear indicator <NUM> may be determined by measuring the resistance after the cut and comparing to the original resistance prior to the cut.

The second component <NUM> includes a blind hole <NUM> configured to partially enclose the first component <NUM>. The blind hole <NUM> includes a blind hole base <NUM>. The first component <NUM> is inserted into the blind hole <NUM> such that the second component <NUM> partially encloses the first component <NUM> and the plurality of wires <NUM> are located proximate the base <NUM> of the blind hole <NUM>. In an embodiment, the blind hole <NUM> substantially matches a shape of the first component <NUM>. In an embodiment, the blind hole <NUM> may be a round shape and/or a cylindrical shape. The first component <NUM> may be securely attached to the second component <NUM> by an epoxy capable of withstanding the high temperatures of a gas turbine engine <NUM>. The second component <NUM> may be composed of a ceramic material capable of withstanding the high temperatures of a gas turbine engine <NUM>.

The second component <NUM> may have a first side <NUM> and a second side <NUM> parallel to the first side <NUM>. The first side <NUM> shares a common wall <NUM> with the blind hole base <NUM>. The second component <NUM> also comprises a mid-section <NUM> interposed between the first side <NUM> and the second side <NUM>. In an embodiment, a second diameter D4 of the second side <NUM> may be larger than a first diameter D3 of the first side <NUM>. The second side <NUM> may be affixed to the blade outer air seal <NUM> and/or the abradable coating <NUM>. The second side <NUM> may be affixed to the blade outer air seal <NUM> and/or the abradable coating <NUM> by an epoxy in a non-limiting example. If the blade <NUM> strikes the second component <NUM> then a layer of the second component <NUM> will be removed from the first side <NUM>. Thus, the first side <NUM> delaminates when impacted by a blade <NUM> of the gas turbine engine <NUM>. Delaminate may be understood to mean the removal of material from the second component <NUM> in layers. In an embodiment, the first component <NUM> is configured to delaminate when impacted by a blade <NUM> of the gas turbine engine <NUM>. The second component <NUM> will continue to delaminate in layers until the first component <NUM> is exposed to the blade <NUM> and then the first component <NUM> and the second component <NUM> will delaminate together. In an embodiment, the mid-section <NUM> of the second component <NUM> between the first side <NUM> and the second side <NUM> may have conical frustum shape, as seen in <FIG>. Advantageously, since the mid-section <NUM> is within the flow path of the gas turbine engine <NUM>, a conical frustum shape is aerodynamic and may provide reduce disturbance to airflow through the gas turbine engine <NUM>. Also advantageously, a conical frustrum shape may minimize major material loss when struck by the blade <NUM>. As may be appreciated by one of skill in the art, the second component <NUM> may include various shapes, sizes and reference dimensions not disclosed herein.

The second component <NUM> may also include an extrusion <NUM> projecting out from the second side <NUM>. The extrusion <NUM> may provide additional support to the first component <NUM>. The extrusion <NUM> may protrude into the blade outer air seal <NUM>. A portion of the first component <NUM> may protrude past the blade outer air seal <NUM>, such that the plates <NUM>, <NUM> may be connected to the lead lines <NUM>, <NUM> without interfering with the blade outer air seal <NUM>.

Referring now to <FIG> with continued reference to <FIG>. <FIG> shows a flow chart illustrating a method <NUM> for manufacturing a wear indicator <NUM> in accordance with an embodiment of the present disclosure. Blocks <NUM>, <NUM>, and <NUM> illustrate the fabrication/assembly of a first component <NUM> of the wear indicator <NUM>. At block <NUM>, a first plate <NUM> is formed having a plurality of first orifices <NUM>. Also at block <NUM>, a second plate <NUM> is formed having a plurality of second orifices <NUM>. Further at block <NUM>, the second plate <NUM> is oriented in relation to the first plate <NUM> such that the plurality of second orifices <NUM> are complimentary to the plurality of first orifices <NUM>, as seen in <FIG>. The term complimentary means that the orifices <NUM>, <NUM> are generally in line with each other.

At block <NUM> a plurality of wires <NUM> are welded to the first plate <NUM>. Each of the wires <NUM> is located in a first hole <NUM> of the first plate <NUM>. Also at block <NUM>, each of the plurality of wires <NUM> extend across a cavity <NUM> between the first plate <NUM> and the second plate <NUM>. Further at block <NUM>, the plurality of wires <NUM> are welded to the second plate <NUM>. Each of the wires <NUM> is located in a second hole <NUM> of the second plate <NUM>.

At block <NUM>, a portion D1 of the cavity <NUM> is filled with a potting material <NUM> such that the potting material <NUM> fills voids <NUM> between each of the plurality of wires <NUM>. The potting material <NUM> may enter the cavity as a liquid flowing in between the wires <NUM> and then harden to a solid. The plurality of wires <NUM> are electrically insulated from each other by the potting material <NUM>.

At block <NUM>, a second component <NUM> having a blind hole <NUM> with a blind hole base <NUM> is obtained. Also at block <NUM>, the first component <NUM> is inserted into the blind hole <NUM> such that the second component <NUM> partially encloses the first component <NUM> and the plurality of wires <NUM> are located proximate the base <NUM> of the blind hole <NUM>. Block <NUM> may also include forming the second component <NUM>. As mentioned above, a formed second component <NUM> may include a first side <NUM> that delaminates when impacted by a blade <NUM> of the gas turbine engine <NUM>; and a second side <NUM> parallel to the first side <NUM>. The first side <NUM> shares a common wall <NUM> with the blind hole base <NUM>. Block <NUM> may also include attaching the first component <NUM> to the second component <NUM>. The first component <NUM> may be attached to the second component <NUM> using an epoxy.

At block <NUM>, the second side <NUM> is attached to a blade outer air seal <NUM>. At block <NUM>, first lead line <NUM> is electrically connected to the first plate <NUM> and a second lead line <NUM> is electrically connected to the second plate <NUM>. At block <NUM>, a portion of the first side <NUM> of the second component <NUM> is removed after the first component <NUM> has been inserted into the second component <NUM>. The portion may be removed by grinding the first side <NUM>. Advantageously, a portion of the first side <NUM> may be grinded away in order to achieve a desired starting size for the wear indicator <NUM>.

Referring now to <FIG> with continued reference to <FIG>. <FIG> is a flow chart illustrating a method <NUM> for detecting blade clearance in a gas turbine engine <NUM>, in accordance with an embodiment. At block <NUM>, the wear indicator <NUM> is attached to an inner surface <NUM> of a gas turbine engine <NUM> opposite a blade <NUM> of the gas turbine engine <NUM>. At block <NUM>, a first resistance of the wear indicator <NUM> is measured. At block <NUM>, the gas turbine engine <NUM> is operated at a first selected speed for a first period of time to remove material from the wear indicator <NUM>. At block <NUM>, a second resistance of the wear indicator <NUM> is measured after material is removed from the wear indicator <NUM>. At block <NUM>, a change in resistance between the second resistance and the first resistance is determined. At block <NUM>, an amount of material removed from the wear indicator <NUM> by the blade <NUM> is determined in response to the change in resistance. Method <NUM> may also include: determining a clearance between the blade <NUM> and the inner surface <NUM> in response to the amount of material removed from the wear indicator <NUM>.

Technical effects of embodiments of the present disclosure include using a wear indicator to determine blade tip clearance through detecting a change in electrical resistance.

Claim 1:
In combination a wear indicator (<NUM>) and a component (<NUM>) of a gas turbine engine, wherein the wear indicator (<NUM>) is secured to a surface (<NUM>) of the component (<NUM>) of the gas turbine engine, the wear indicator (<NUM>) comprising:
a first component (<NUM>) including:
a first plate (<NUM>);
a second plate (<NUM>) opposite the first plate (<NUM>);
a plurality of wires (<NUM>) extending from first plate (<NUM>) to the second plate (<NUM>), wherein the first plate (<NUM>) is electrically connected to the second plate (<NUM>) through the plurality of wires (<NUM>); and
a non-conductive potting material (<NUM>) configured to partially fill the first component (<NUM>) and fill voids (<NUM>) between the plurality of wires (<NUM>), such that the plurality of wires (<NUM>) are electrically insulated from each other by the non-conductive potting material (<NUM>);
wherein the first plate (<NUM>) and the second plate (<NUM>) are separated by a gap (D2), the gap (D2) keeping the first plate (<NUM>) electrically separate from the second plate (<NUM> ).