Patent Application
20170234772
Turbine engines, and particularly gas or combustion turbine engines, are rotary engines that extract energy from a flow of combusted gases passing through the engine onto a multitude of turbine blades. Gas turbine engines have been used for land and nautical locomotion and power generation, but are most commonly used for aeronautical applications such as for aircraft, including helicopters. In aircraft, gas turbine engines are used for propulsion of the aircraft. In terrestrial applications, turbine engines are often used for power generation.
Gas turbine engines for aircraft are designed to operate at high temperatures to maximize engine efficiency. Temperatures in the high pressure turbine are around 1000° C. to 2000° C. and fluid from the compressor is around 500° C. to 760° C. Internal components of gas and steam turbines, for example, steam turbine blades are typically visually inspected, during a turbine outage, by inserting a borescope through an opening in the outer turbine shell and articulating the video head of the borescope to achieve the desired inspection view. Typically a waiting period is necessary after shutdown and before inspection because current borescope inspection equipment has a temperature limit of approximately 50° C. As a result of this temperature limitation, gas and steam turbine inspections cannot be performed until the turbine cools down from its normal operating temperature.
In one aspect, the invention relates to an optical imaging system, including a housing configured for mounting to a wall of a turbine engine, a camera located in the housing, a hollow probe extending from the housing and having a longitudinal axis, an image receiving device at an end of the hollow probe and communicably coupled with the camera; and at least one mechanism coupled with the housing and configured to urge the hollow probe to move along the longitudinal axis and rotate the hollow probe about the longitudinal axis.
In another aspect, the invention relates to a gas turbine engine, including a radial wall defining an interior and an exterior of the gas turbine engine and having an aperture, a set of turbine blades located in the interior and configured to rotate about a shaft, and an optical imaging system having a housing configured for mounting to the radial wall, a camera located in the housing, a hollow probe extending from the housing and having a longitudinal axis, an image receiving device at an end of the hollow probe in communication with the camera, and at least one mechanism coupled with the housing and configured to urge the hollow probe to move along the longitudinal axis through the aperture into the interior of the gas turbine engine, and configured to rotate the hollow probe about the longitudinal axis.
In yet another aspect, the invention relates to a method for operating an optical imaging system in a gas turbine engine having a rotating set of turbine blades. The method including moving an image receiving device, which is in communication with the camera, into an interior of an operating gas turbine engine, selecting a sampling frequency for imaging based at least in part on a rotational speed of the rotating set of turbine blades, and capturing a set of images of a target visually in-line with the rotating set of turbine blades.
In the drawings:
The various aspects described herein relate to an optical imaging system such as a borescope assembly and method for inspecting internal components of a turbine engine while the turbine engine is being operated. Installing optics to monitor and image hot gas path components such as airfoils and combustors, in an operating gas turbine is not a relatively easy or straight-forward task. Presently, rigid optics transmit light with higher imaging fidelity than fiber optics and thus rigid optics can be located inside a gas turbine to relay images to a convenient location where an imaging device such as an infrared (IR) camera can be placed. However, to image its interior with a fixed optics probe, an engine has to be shut down. The various aspects described herein relate to an optical imaging system with a traversing and yawing optics probe such that, while a gas turbine is operating, different regions of the hot gas path can be viewed by remotely moving the probe. The various aspects described herein improve the efficiency in testing and allow more regions to be viewed. Further, the various aspects described herein can be particularly useful in viewing a shroud above a set of rotating turbine blades in a gas turbine engine.
For purposes of illustration, the present invention will be described with respect to an aircraft gas turbine engine. It will be understood, however, that the invention is not so limited and may have general applicability in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications.
The fan section 18 includes a fan casing 40 surrounding the fan 20. The fan 20 includes a plurality of fan blades 42 disposed radially about the centerline 12.
The HP compressor 26, the combustor 30, and the HP turbine 34 form a core 44 of the engine 10 which generates combustion gases. The core 44 is surrounded by core casing 46 which can be coupled with the fan casing 40.
A HP shaft or spool 48 disposed coaxially about the centerline 12 of the engine 10 drivingly connects the HP turbine 34 to the HP compressor 26. A LP shaft or spool 50, which is disposed coaxially about the centerline 12 of the engine 10 within the larger diameter annular HP spool 48, drivingly connects the LP turbine 36 to the LP compressor 24 and fan 20.
The LP compressor 24 and the HP compressor 26 respectively include a plurality of compressor stages 52, 54, in which a set of compressor blades 56, 58 rotate relative to a corresponding set of static compressor vanes 60, 62 (also called a nozzle) to compress or pressurize the stream of fluid passing through the stage. In a single compressor stage 52, 54, multiple compressor blades 56, 58 can be provided in a ring and extend radially outwardly relative to the centerline 12, from a blade platform to a blade tip, while the corresponding static compressor vanes 60, 62 are positioned downstream of and adjacent to the rotating blades 56, 58. It is noted that the number of blades, vanes, and compressor stages shown in
The HP turbine 34 and the LP turbine 36 respectively include a plurality of turbine stages 64, 66, in which a set of turbine blades 68, 70 are rotated relative to a corresponding set of static turbine vanes 72, 74 (also called a nozzle) to extract energy from the stream of fluid passing through the stage. In a single turbine stage 64, 66, multiple turbine blades 68, 70 can be provided in a ring and extend radially outwardly relative to the centerline 12, from a blade platform to a blade tip, while the corresponding static turbine vanes 72, 74 are positioned upstream of and adjacent to the rotating blades 68, 70. It is noted that the number of blades, vanes, and turbine stages shown in
In operation, the rotating fan 20 supplies ambient air to the LP compressor 24, which then supplies pressurized ambient air to the HP compressor 26, which further pressurizes the ambient air. The pressurized air from the HP compressor 26 is mixed with fuel in the combustor 30 and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine 34, which drives the HP compressor 26. The combustion gases are discharged into the LP turbine 36, which extracts additional work to drive the LP compressor 24, and the exhaust gas is ultimately discharged from the engine 10 via the exhaust section 38. The driving of the LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP compressor 24.
Some of the ambient air supplied by the fan 20 can bypass the engine core 44 and be used for cooling of portions, especially hot portions, of the engine 10, and/or used to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are normally downstream of the combustor 30, especially the turbine section 32, with the HP turbine 34 being the hottest portion as it is directly downstream of the combustion section 28. Other sources of cooling fluid can include, but are not limited to, fluid discharged from the LP compressor 24 or the HP compressor 26.
Embodiments of the optical imaging system 100 can include a housing 106, a camera 108 located within the housing 106, a hollow probe 118 extending from the housing 108, an image receiving device 114 at the end of the hollow probe 118 and at least one mechanism 104 configured to maneuver the hollow probe 118 within the interior 115 of the gas turbine engine. The housing 106 is included and configured for mounting to the radial wall 110 of the turbine engine. The optical imaging system 100 can be manipulated to directionally control the image receiving device 114, including when inside the gas turbine engine 10. More specifically, at least one mechanism 104 can be coupled with the housing 106 and configured to urge the hollow probe 118 to move along or traverse 123 the longitudinal axis 112 through the aperture 111 into the interior 115 of the gas turbine engine. Further, the urging mechanism can be configured to rotate the hollow probe 118 about the longitudinal axis 112 to induce yaw 125. The urging mechanism 104 can include one or more motors useful for rotating and translating a shaft. For example, as shown, the urging mechanism 104 can include both a translational motor 122 and a rotational motor 124. The urging mechanism 104 can be formed from any device useful for urging or maneuvering the hollow probe 118 along the longitudinal axis 112 into a cavity in the interior 115 of the turbine engine including, but not limited to, one or more permanent magnet stepper motors, hybrid synchronous stepper motors, variable reluctance stepper motors, lavet type stepping motors, AC motors, DC motors, gearboxes, etc. and combinations thereof.
Directional control of the image receiving device 114 is provided by a controller 102 external to the gas turbine engine 10. Thus, the image receiving device 114 is directionally controlled such that a selected one or more components internal to the gas turbine engine 10 can be viewed externally of the gas turbine engine 10. Parts of the optical imaging system 100 can be cooled including, but not limited to, by flowing a cooling medium along a substantial portion of the length of the hollow probe 118 and particularly about the image receiving device 114.
As shown in
Contained within the housing 106, the camera 108 is responsive to imaging data of one or more components of a turbine engine positioned within a field of view 128 of the image receiving device 114. The camera 108 is configured to sense a temperature of a surface in the cavity or interior 115 of the turbine engine The camera 108 can be any device for recording image data correlated to surface temperatures including, but not limited to, an infrared camera, a visible camera, a pyrometer, a multi-spectral camera, a hyperspectral camera, a charge-coupled device, an active pixel sensor, a complementary metal-oxide-semiconductor (CMOS) sensor, etc.
The hollow probe 118, which can also be referred to as a borescope, extends from the housing 106 along the longitudinal axis 112 normal to the radial wall 110 towards the interior 115 of the turbine engine. The hollow probe 118 provides a conduit of optical communication from the image receiving device 114 at the end of the probe 118 to the camera 108 within the housing 106. The hollow probe 118 can include any components used in the transmission of optical data including, but not limited to, free space, one or more lenses, fiber optic cable and combinations thereof.
The image receiving device 114 located at the distal end of the hollow probe 118 redirects incoming optical data to relay along the longitudinal axis 112. As shown in
Concentric to the hollow probe 118, one or more guide tubes 116, 130 can protect and assist to maneuver the hollow probe 118. A moving guide tube 116 can traverse and rotate with the camera housing 106 along the longitudinal axis 112. A fixed or stationary guide tube 130 can be fixed to a wall of the turbine engine where the wall can be any interior structure within the turbine engine including, but not limited to, a radial wall that forms the vanes of a turbine stage.
When the hollow probe 118 or borescope is maneuvered to the correct location and yaw angle, the probe optics enable the camera 108 to image the surface of the shroud 120. Advantageously, the camera 108 attached to the traversing and yawing urging mechanism 104 and coupled to the hollow probe 118 allows the shroud 120 to be imaged while the gas turbine engine is operating. The hollow probe along with the guide tubes 116, 130 can include multiple tubes with optical elements and passages for cooling and purging of air.
Referring now to
Referring now to
Then, at step 406, the camera captures a set of images of a target visually in-line with the rotating set of turbine blades. The captured set of images can form any set of images of the interior of the turbine engine, including, but not limited to a set of images of the set of turbine blades, a set of thermal images and a set of images of the shroud.
The sequence depicted is for illustrative purposes only and is not meant to limit the method 400 in any way as it is understood that the portions of the method may proceed in a different logical order, additional or intervening portions may be included, or described portions of the method may be divided into multiple portions, without detracting from embodiments of the invention.
Benefits of the above-described embodiments include capturing two-dimensional data related to temperatures of a shroud that are located above a set rotating turbine blades in an operating gas turbine. The shrouds are located in a very high temperature and pressure environment and are proximate to rotating blades moving at very high velocity. The probe is remotely controlled in order that the probe stays in the hot gas path for the minimum time to take the required images thereby preserving the operational life of the optical imaging system components. The optical imaging system provides temperature measurements that are necessary to validate analytical designs and models needed to estimate life of these components.
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.
