Patent Description:
Gas turbine engines include compressor sections to compress an airflow, combustor sections that combine the airflow with fuel for combustion and generate exhaust, and turbine sections that convert the exhaust into torque to drive the compressor sections. The combustor sections may experience relatively great temperatures. It may be desirable to identify characteristics within a combustion chamber of the combustor section, such as an air/fuel ratio or a temperature within the combustion chamber for various purposes. However, due to the relatively great temperatures within the combustor sections, it may be relatively difficult to detect this information.

<CIT> discloses a fuel injector for a gas turbine engine combustor that includes a fuel nozzle for injecting fuel into the gas turbine engine combustor and a fiber optic microphone operatively associated with the fuel nozzle for measuring acoustic pressure differentials within a combustion chamber of the gas turbine engine combustor.

<CIT> discloses systems and methods for providing optical interrogation sensors for combustion control.

<CIT> discloses a device for diagnosing a combustion state of a gas turbine.

According to the invention as claimed herein, a sensor system for use in a gas turbine engine as claimed in claim <NUM> is provided.

In any of the foregoing embodiments, the reflective surface is suspended in at least one of the inner shroud or the outer shroud.

Any of the foregoing embodiments may further include a fuel nozzle having a fuel nozzle stem, wherein the optical sensor is coupled to the fuel nozzle stem and located outside of the combustion chamber.

Any of the foregoing embodiments may further include an optical decoder configured to receive a signal from the optical sensor corresponding to the reflected optical features and to decode the reflected optical features.

Any of the foregoing embodiments may further include a controller coupled to the optical decoder and configured to receive a decoded signal from the optical decoder corresponding to the reflected optical features and to determine characteristics of the combustion chamber based on the decoded signal.

Any of the foregoing embodiments may further include a fuel nozzle and a trim valve configured to adjust an amount of fuel injected by the fuel nozzle, wherein the controller is further configured to control the trim valve to adjust the amount of fuel injected by the fuel nozzle based on the characteristics of the combustion chamber.

In any of the foregoing embodiments, the reflected optical features include infrared light waves, and the characteristics of the combustion chamber correspond to a temperature within the combustion chamber.

Any of the foregoing embodiments may further include a second optical sensor and a second reflective surface located axially forward or axially aft, as well as in a different circumferential or radial position, relative to the first reflective surface, wherein: the second reflective surface is configured to reflect second optical features of the combustion chamber received via the second opening towards the fuel nozzle; and the second optical sensor is configured to receive the second reflected optical features from the second reflective surface.

Any of the foregoing embodiments may further include an igniter configured to ignite the fuel in the combustion chamber, wherein the optical sensor is coupled to the igniter.

Any of the foregoing embodiments may further include an airfoil located in a high pressure turbine section of the gas turbine engine, wherein the reflected optical features correspond to the airfoil.

In any of the foregoing embodiments, the combustor component includes at least one of a fuel nozzle or an igniter.

The foregoing features and elements are to be combined in various combinations without exclusivity, unless expressly indicated otherwise.

The subject matter of the present invention as claimed herein is particularly pointed out in the concluding portion of the specification. A more complete understanding of the present invention as claimed herein, however, is best be obtained by referring to the detailed description and claims when considered in connection with the drawing figures, wherein like numerals denote like elements.

The detailed description of exemplary embodiments of the invention as claimed 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 inventions, 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 invention. For example, any reference to singular includes plural embodiments, and any reference to more than one component or step may include a singular embodiment or step. Also, any reference to attached, fixed, connected or the like may include permanent, removable, temporary, partial, full and/or any other possible attachment option. Where used herein, the phrase "at least one of A or B" can include any of "A" only, "B" only, or "A and B.

With reference to <FIG>, a gas turbine engine <NUM> is provided. As used herein, "aft" refers to the direction associated with the tail (e.g., the back end) of an aircraft, or generally, to the direction of exhaust of the gas turbine engine. As used herein, "forward" refers to the direction associated with the nose (e.g., the front end) of an aircraft, or generally, to the direction of flight or motion. As utilized herein, radially inward refers to the negative R direction and radially outward refers to the R direction. An A-R-C axis is shown throughout the drawings to illustrate the relative position of various components.

The 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, the fan section <NUM> drives air along a bypass flow-path B while the compressor section <NUM> drives air along a core flow-path C for compression and communication into the combustor section <NUM> then expansion through the 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 three-spool architectures, geared turbofan architectures, and turboshaft or industrial gas turbines with one or more spools.

The gas turbine engine <NUM> generally comprises a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis X-X' relative to an engine static structure <NUM> via several bearing systems <NUM>, <NUM>-<NUM>, and <NUM>-<NUM>. It should be understood that various bearing systems <NUM> at various locations may alternatively or additionally be provided, including for example, the bearing system <NUM>, the bearing system <NUM>-<NUM>, and the bearing system <NUM>-<NUM>.

The low speed spool <NUM> generally includes an inner shaft <NUM> that interconnects a fan <NUM>, a low pressure (or first) compressor section <NUM> and a low pressure (or second) turbine section <NUM>. The inner shaft <NUM> is connected to the fan <NUM> through a geared architecture <NUM> that can drive the fan shaft <NUM>, and thus the fan <NUM>, at a lower speed than the low speed spool <NUM>. The geared architecture <NUM> includes a gear assembly <NUM> enclosed within a gear diffuser case <NUM>. The gear assembly <NUM> couples the inner shaft <NUM> to a rotating fan structure.

The high speed spool <NUM> includes an outer shaft <NUM> that interconnects a high pressure (or second) compressor section <NUM> and the high pressure (or first) turbine section <NUM>. A combustor <NUM> is located between the high pressure compressor <NUM> and the high pressure turbine <NUM>. A mid-turbine frame <NUM> of the engine static structure <NUM> is located generally between the high pressure turbine <NUM> and the low pressure turbine <NUM>. The mid-turbine frame <NUM> supports one or more bearing systems <NUM> in the turbine section <NUM>. The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate via the bearing systems <NUM> about the engine central longitudinal axis X-X', 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 is compressed by the low pressure compressor section <NUM> then the high pressure compressor <NUM>, mixed and burned with fuel in the combustor <NUM>, then expanded over the high pressure turbine <NUM> and the low pressure turbine <NUM>. The mid-turbine frame <NUM> includes airfoils <NUM> which are in the core airflow path.

The gas turbine engine <NUM> is a high-bypass ratio geared aircraft engine. The bypass ratio of the gas turbine engine <NUM> may be greater than about six (<NUM>). The bypass ratio of the gas turbine engine <NUM> may also be greater than ten (<NUM>:<NUM>). The 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. The geared architecture <NUM> may have a gear reduction ratio of greater than about <NUM> and the low pressure turbine <NUM> may have a pressure ratio that is greater than about five (<NUM>). The diameter of the fan <NUM> may be significantly larger than that of the low pressure compressor section <NUM>, and the low pressure turbine <NUM> may have a pressure ratio that is greater than about five (<NUM>:<NUM>). The pressure ratio of the low pressure turbine <NUM> is measured prior to an inlet of the low pressure turbine <NUM> as related to the pressure at the outlet of the low pressure turbine <NUM>. 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 turbine engines including direct drive turbofans.

The next generation turbofan engines are designed for higher efficiency and use higher pressure ratios and higher temperatures in the high pressure compressor <NUM> than are conventionally experienced. These higher operating temperatures and pressure ratios create operating environments that cause thermal loads that are higher than the thermal loads conventionally experienced, which may shorten the operational life of current components.

In various embodiments and referring to <FIG> and <FIG>, the combustor section <NUM> may include an annular combustor <NUM>. The annular combustor <NUM> may include multiple fuel nozzles <NUM>, which each include their own trim valve <NUM> (only one of which is illustrated in <FIG>). In various embodiments, each fuel nozzle <NUM> delivers fuel to a respective section of the combustion chamber <NUM>. The fuel nozzles <NUM> may be arranged circumferentially around an axis within a combustor <NUM>. The fuel nozzles <NUM> may include stems <NUM> that extend from a diffuser case <NUM> to openings <NUM>.

Although a single fuel nozzle <NUM> (and other components) is shown in the drawings, one skilled in the art will realize that the combustor section <NUM> may include multiple openings <NUM> circumferentially around the combustor section <NUM> that receive fuel nozzles <NUM>.

Referring now to <FIG>, additional features of the combustor section <NUM> are shown. In particular, the combustor section <NUM> includes a system <NUM> for detecting optical features corresponding to combustion in the gas turbine engine <NUM> of <FIG>.

The combustor section <NUM>, or system <NUM>, includes a diffuser case <NUM>. The diffuser case <NUM> surrounds or encloses a liner and shell assembly <NUM> that defines a combustion chamber <NUM>. A fuel source <NUM> provides fuel to the fuel nozzle <NUM> for delivery to the combustion chamber <NUM>. The fuel nozzle <NUM> extends through an aperture <NUM> in the diffuser case <NUM>. An end of the fuel nozzle <NUM> may be arranged at an inlet <NUM> of the combustion chamber <NUM>. A swirler <NUM> may provide desired airflow motion from the compressor section <NUM> of <FIG> to achieve a desired air/fuel mixture. The liner and shell assembly <NUM> typically includes one or more igniters used to begin combustion of the air/fuel mixture.

The liner and shell assembly <NUM> includes or defines openings <NUM> that may receive air to increase an air/fuel ratio within the combustion chamber <NUM>. In embodiments according to the invention as claimed herein, the openings <NUM> are formed in the liner and shell assembly <NUM> to provide a view of features, such as optical features, within the combustion chamber <NUM> and may fail to provide a significant amount of air to the combustion chamber <NUM>.

The fuel nozzle <NUM> includes a fuel nozzle stem <NUM> terminating in a fuel exit <NUM> that delivers fuel to the combustion chamber <NUM>. A sensor <NUM> may be coupled to, or at least partially arranged within, the fuel nozzle stem <NUM>. Fuel provided by the fuel source <NUM> to the fuel nozzle <NUM> may cool the sensor <NUM>, or fuel within the fuel nozzle <NUM> may transfer heat from the sensor <NUM> to the fuel via the fuel nozzle stem <NUM>.

The diffuser case <NUM> and the liner and shell assembly <NUM> define an outer shroud <NUM> and an inner shroud <NUM> therebetween. In various embodiments, air may flow through at least one of the inner shroud <NUM> or the outer shroud <NUM>. In various embodiments, the fuel nozzle <NUM> may extend through the outer shroud <NUM> and may be extended into the inner shroud <NUM>. In various embodiments, the sensor <NUM> may be located in at least one of the inner shroud <NUM> or the outer shroud <NUM>. In various embodiments, the sensor <NUM> may be exposed to temperatures that are less than temperatures experienced within the combustion chamber <NUM> due to the location of the sensor <NUM> within the inner shroud <NUM> or the outer shroud <NUM>.

A reflective surface <NUM> is defined by one or both of the diffuser case <NUM> or the liner and shell assembly <NUM>. The reflective surface <NUM> is designed to reflect optical features from the combustion chamber <NUM> that are received via the opening <NUM> towards the sensor <NUM>. The reflective surface <NUM> is designed to reflect the optical features within the combustion chamber <NUM> through one or both of the inner shroud <NUM> or the outer shroud <NUM> towards the sensor <NUM>.

The reflective surface <NUM> may include, for example, a metal, a glass, or the like. According to the invention as claimed herein, the reflective surface <NUM> is formed by the diffuser case <NUM> or the liner and shell assembly <NUM> and may include the same material as the diffuser case <NUM> or the liner and shell assembly <NUM>. For example, the material of at least one of the diffuser case <NUM>, the liner and shell assembly <NUM>, or the reflective surface <NUM> may include a nickel-based alloy, a steel, or the like. In various embodiments, the reflective surface <NUM> may be resistant to relatively high temperatures such as those exceeding <NUM> degrees Fahrenheit (<NUM> degrees Celsius). In various embodiments, the reflective surface <NUM> may include a platinum coating or another reflective coating designed to increase reflection of the optical features. In various embodiments, the reflective surface <NUM> may be polished to increase reflection of the optical features.

The sensor <NUM> is an optical sensor. The sensor <NUM> may be provided by sapphire or quartz fibers or any other suitable material. There may be lenses or reflective surfaces or prisms to direct the light from the combustor through passages or fiber optics internal to the fuel nozzle <NUM> and stem to an optical decoder <NUM>. The sensor <NUM> may communicate with an optical decoder <NUM>. In various embodiments, the fuel nozzle <NUM>, trim valve <NUM>, sensor <NUM>, and optical decoder <NUM> may be integrated to provide an easily replaceable unit. The sensor <NUM> may detect a fuel delivery parameter associated with the combustor system <NUM>. In various embodiments, the sensor may detect a wavelength spectrum capable of detecting an amount of products of combustion or gasses associated with combustion such as oxygen, nitrogen, water, carbon monoxide, and/or carbon dioxide in the combustion chamber <NUM>. In various embodiments, carbon dioxide and water may be the preferred gasses. For example, the sensor <NUM> may detect optical data corresponding to a temperature within the combustion chamber <NUM>. For example, the sensor <NUM> may detect infrared light corresponding to one or more wavelengths emitted by the high temperature products of combustion such as water, carbon dioxide, or the like that corresponds to a temperature of combustion within the combustion chamber <NUM>.

The fuel nozzle <NUM> may include a first section <NUM> (located radially outward from the diffuser case <NUM> and mounted to the diffuser case <NUM>), a second section <NUM> (located between the diffuser case <NUM> and the liner and shell assembly <NUM>), and a third section <NUM> (located within the liner and shell assembly <NUM> and mounted to the liner and shell assembly). The second section <NUM> may extend from the first section <NUM> to the third section <NUM>. In various embodiments, the sensor <NUM> may be positioned on the second section <NUM> of the fuel nozzle <NUM>. This may be advantageous as the second section <NUM> may be exposed to lesser temperatures than the third section <NUM>, thus increasing a lifespan of the sensor <NUM>.

A controller <NUM> may communicate with the sensor <NUM> via the optical decoder <NUM>. In various embodiments, the optical decoder <NUM> may be included on a single chip with the controller <NUM> and, in various embodiments, the sensor <NUM> may likewise be integrated on the single chip. The optical decoder <NUM> may decode the reflected optical features received from the sensor <NUM>. The controller <NUM> may receive a decoded signal from the optical decoder <NUM> that corresponds to the reflected optical features. The controller <NUM> may then determine characteristics within the combustion chamber <NUM> based on the decoded signal. The characteristics within the combustion chamber <NUM> may correspond to a temperature within the combustion chamber <NUM>, and may be referred to as a detected fuel delivery parameter as the amount of fuel and oxygen provided to the combustion chamber <NUM> may change at least one of the temperature or the optical features. For example, the controller <NUM> may determine temperatures at various circumferential locations of the combustion chamber <NUM> through sensors <NUM> on fuel nozzles <NUM> circumferentially arranged around the combustor section <NUM>, and a system level controller may perform an algorithm or other method to determine adjustments to the fuel supply for each of the individual fuel nozzles.

Based on the detected optical features, the controller <NUM> may determine an air/fuel ratio within a given combustion chamber <NUM>. The controller <NUM> may be programmed to actuate the trim valve <NUM> to achieve a desired air/fuel ratio within the combustion chamber <NUM> and/or, balance the air/fuel ratio and/or fuel delivery to that of another sector of the combustion chamber <NUM> based on the detected fuel delivery parameter. The controller <NUM> may be used, for example, to minimize differences between the various sectors within the annular combustor <NUM> in a combustion system thereby reducing the cooling requirement on the turbine. Smoke, particulates and other engine emissions may also be reduced.

The combustor section <NUM> and system <NUM> provide several benefits over previously disclosed approaches. For example, the sensor <NUM> may fail to be exposed to fuel, combustion, or exhaust, thus increasing a lifespan of the sensor <NUM>. Additionally, a temperature within the environment of the sensor <NUM> is lower than in the combustion chamber <NUM> and is more uniform, thus increasing ease of mechanical design and thermal protection of the sensor <NUM>. Additionally, the sensor <NUM> may detect the optical features without looking through a flame within the combustion chamber <NUM>, which may affect the detected information. This increases the likelihood of the detected optical features accurately reflecting the conditions within the combustion chamber <NUM>. Additionally, because the sensor <NUM> is coupled to the fuel nozzle stem <NUM>, the sensor <NUM> may be cooled from the fuel. Additionally, installation of the sensor <NUM> and the fuel nozzle <NUM> may be relatively easy due to the integration of the sensor <NUM> and the trim valve <NUM> in a single fuel nozzle <NUM>.

Referring now to <FIG>, another cross-sectional view of the combustor section <NUM> is shown. The combustor section <NUM> may further include an igniter <NUM>. The igniter may extend through the diffuser case <NUM> and the liner and shell assembly <NUM> and may ignite a mixture of fuel and oxygen in the combustion chamber <NUM>. In various embodiments, a reflective surface <NUM> may be coupled to the diffuser case <NUM>, and a sensor <NUM> may be coupled to the igniter <NUM>. The sensor <NUM> and the reflective surface <NUM> may be oriented in such a manner that the reflective surface <NUM> reflects light from the combustion chamber <NUM> towards the sensor <NUM> such that the sensor <NUM> may detect optical data corresponding to an interior of the combustion chamber <NUM>.

Referring now to <FIG>, another combustor section <NUM> may include similar features as the combustor section <NUM> of <FIG>. However, the combustor section <NUM> may include two reflective surfaces <NUM>, <NUM> located within an outer shroud <NUM> and suspended within the outer shroud <NUM> from a diffuser case <NUM>. The first reflective surface <NUM> may receive optical features from a first opening <NUM> in a liner and shell assembly <NUM>, and the second reflective surface <NUM> may receive optical features from a second opening <NUM> in the liner and shell assembly <NUM>. A first optical sensor <NUM> is oriented in such a manner as to receive the reflected optical features from the first reflective surface <NUM>, and a second optical sensor <NUM> is oriented to receive the reflected optical features from the second reflective surface <NUM>. In various embodiments, a single optical sensor <NUM> may be designed to receive the reflected optical features from both the first reflective surface <NUM> and the second reflective surface <NUM>. This <NUM> arrangement may be more particularly focused on combustor lighting and may be used to control the amount and duration of sparking.

An optical decoder <NUM> may be designed to receive the detected optical features from the first optical sensor <NUM> and the second optical sensor <NUM>. The optical decoder <NUM> may decode the reflected optical features from both of the optical sensors <NUM>, <NUM>. A controller <NUM> may receive a signal corresponding to decoded optical features from the first optical sensor <NUM> and the second optical sensor <NUM> and may determine characteristics within multiple locations of a combustion chamber <NUM> based on the decoded optical features from both of the optical sensors <NUM>, <NUM>. In various embodiments, a harness may be designed to receive the determined characteristics from multiple controllers <NUM> oriented about a circumference of the combustor section <NUM> and to determine control of trim valves of the various fuel nozzles based on the data received from the multiple controllers <NUM>.

Another reflective surface <NUM> may be coupled to the liner and shell assembly <NUM>. The reflective surface <NUM> may reflect optical features received from the opening <NUM> towards one or more sensor <NUM>, <NUM>. The controller <NUM> may determine optical characteristics of combustion based on the received data.

The combustor section <NUM> may include an airfoil <NUM>. The airfoil <NUM> may include, for example, a turbine vane. It may be desirable to identify a temperature of the airfoil <NUM>.

In that regard, the combustor section <NUM> may include a reflective surface, such as the reflective surface <NUM>, that reflects optical features from a leading edge of the airfoil <NUM> through the outer shroud <NUM> towards a sensor, such as the sensor <NUM> or the sensor <NUM>. The controller <NUM> may be coupled to the sensor <NUM> or <NUM> and may receive information corresponding to the reflected optical features. The controller <NUM> may determine a temperature of the airfoil <NUM> based on the reflected optical features received via the sensor <NUM>.

Referring to <FIG>, a method <NUM> for controlling a trim valve of a fuel nozzle based on detected optical features corresponding to combustion in a combustor section of a gas turbine engine is shown. In various embodiments, turbine vanes (such as those shown in the drawings) may be considered to be part of a combustor section. The method includes, at <NUM>, detecting optical characteristics corresponding to a combustor section of a gas turbine engine from a reflective surface. For example, an optical sensor may be coupled to a stem of a fuel nozzle or an igniter within the combustor section and may be situated to receive one or more reflections of optical characteristics within the combustor section.

In block <NUM>, an optical decoder may decode the one or more reflected optical characteristics. For example, the optical decoder may be coupled to a fuel nozzle or igniter stem, may receive the reflected optical characteristics from the sensors, and may decode the reflected optical characteristics. An algorithm may analyze the decoded optical characteristics and may determine how controllers should be controlled based on the analysis.

In block <NUM>, a controller may receive the decoded optical characteristics and may determine characteristics of the combustion chamber based on the decoded reflected optical characteristics. For example, the controller may identify a temperature within the combustion chamber or a sector of the combustion chamber, an amount of a particular gas or other element, a ratio of air to fuel within the combustion chamber, or the like based on the decoded reflected optical characteristics.

In block <NUM>, a system level controller may perform an algorithm or other method to determine adjustments to the fuel supply for each of the individual fuel nozzles.

In block <NUM>, the individual fuel nozzle controllers may control trim valves to adjust an air to fuel ratio provided by the fuel nozzle based on the determined characteristics and the data from the system level controller. For example, the controller may control the flow of fuel from the fuel nozzle for various reasons. These may include determining that ignition has occurred in the combustion chamber, ensuring complete combustion during spool up to prevent overly-rapid spool up, confirm an uneven distribution of fuel during part power up points to eliminate or reduce combustion noise, or controlling the flow of fuel so that all sectors have the same temperature at full power.

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 inventions. The scope of the invention 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 sensor system for use in a gas turbine engine (<NUM>), comprising:
a diffuser case (<NUM>), and a liner and shell assembly (<NUM>) that defines a combustion chamber (<NUM>);
a reflective surface (<NUM>; <NUM>);
an optical sensor (<NUM>, <NUM>) configured to receive reflected optical features from the reflective surface (<NUM>; <NUM>); and
the liner and shell assembly (<NUM>; <NUM>) and the diffuser case (<NUM>; <NUM>) defining an inner shroud (<NUM>) and an outer shroud (<NUM>) between the diffuser case (<NUM>; <NUM>) and the liner and shell assembly (<NUM>; <NUM>), wherein
the reflective surface (<NUM>; <NUM>) is at least partially located in at least one of the inner shroud (<NUM>) or the outer shroud (<NUM>), and
the reflective surface (<NUM>; <NUM>) is configured to reflect optical features corresponding to combustion in the combustion chamber (<NUM>) that exit the combustion chamber (<NUM>) via at least one opening (<NUM>; <NUM>) therein and to reflect the optical features through at least one of the inner shroud (<NUM>) or the outer shroud (<NUM>) towards the optical sensor (<NUM>, <NUM>),
wherein the at least one opening (<NUM>; <NUM>) is defined by the liner and shell assembly (<NUM>; <NUM>);
characterized in that
the reflective surface (<NUM>; <NUM>) is defined by at least one of the diffuser case (<NUM>; <NUM>) or the liner and shell assembly (<NUM>; <NUM>).