Patent Application
20160237851
The present disclosure relates generally to gas turbine engines and, more particularly, to turbine blade sealing structures. Specifically, the disclosure concerns probes within turbine blade sealing structures.
A gas turbine engine typically includes a high-pressure spool, a combustion system, and a low-pressure spool disposed within an engine case to form a generally axial, serial flow path about the engine centerline. The high-pressure spool includes a high-pressure turbine, a high-pressure shaft extending axially forward from the high-pressure turbine, and a high-pressure compressor connected to a forward end of the high-pressure shaft. The low-pressure spool includes a low-pressure turbine, which is disposed downstream of the high-pressure turbine, a low-pressure shaft, which typically extends coaxially through the high-pressure shaft, and a low-pressure compressor connected to a forward end of the low-pressure shaft, forward of the high-pressure compressor. The combustion system is disposed between the high-pressure compressor and the high-pressure turbine and receives compressed air from the compressors and fuel provided by a fuel injection system. A combustion process is carried out within the combustion system to produce high-energy gases to produce thrust and turn the high- and low-pressure turbines, which drive the compressors to sustain the combustion process.
Turbine blade outer air seals (BOAS) seal turbine gas path air between the rotating blades of the turbine and the gas path surface of a casing. The distance between the tip of the blades and the surface of the BOAS greatly impacts turbine efficiency. Reducing that distance as much as possible improves engine efficiency. Accordingly, efforts have been made to monitor blade tip clearance during engine operation. However, extremely high temperatures in the gas path often make such monitoring difficult.
The operating temperature of a typical probe is lower than that of the gas path of a turbine engine. Current methods of cooling probe tips exposed to the gas path include supplying cooling air from an external cooling source. Cooling a probe tip in this manner necessitates a coolant, such as liquid nitrogen, and plumbing to deliver the coolant to the probe tip.
A probe cooling system for a gas turbine engine includes a probe housing and a bushing disposed between an end of the probe housing and a gas path, the bushing extending from the end of the probe housing to form a cavity. The system also includes a first plenum defined within the probe housing and configured to direct air from an internal cooling air supply towards the cavity. The system also includes a second plenum defined between the bushing and the probe housing and configured to direct air from the internal cooling air supply towards the cavity.
A method of cooling a probe includes positioning a probe having a probe tip proximate a gas path, and directing air from an internal cooling air supply through a plurality of plenums towards the probe tip.
In
In the illustrated embodiment, gas turbine engine 10 includes a dual-spool turbofan engine in which the advantages of the present disclosure are particularly well-illustrated. Gas turbine engine 10, of which the operational principles are well-known in the art, comprises fan 12, LPC 14, HPC 16, combustor section 18, HPT 20, and LPT 22, which are each concentrically disposed around longitudinal engine centerline axis CL. Fan 12 is enclosed at its outer diameter within fan case 24. Likewise, the other engine components are correspondingly enclosed at their outer diameters within various engine casings, including LPC case 26, HPC case 28, HPT case 30 and LPT case 32. Fan 12 and LPC 14 are connected to LPT 22 through low-pressure shaft 34. Together fan 12, LPC 14, LPT 22, and low-pressure shaft 34 make up the low-pressure spool. HPC 16 is connected to HPT 20 through high-pressure shaft 36. Together HPC 16, HPT 20, and high-pressure shaft 36 make up the high-pressure spool.
During normal operation, inlet air A enters gas turbine engine 10 where it is divided into streams of primary air AP and secondary air AS after passing through fan 12. Fan 12 is rotated by LPT 22 through low-pressure shaft 34 (either directly as shown or through a gearbox, not shown) to accelerate secondary air AS through exit guide vanes 38, thereby producing a major portion of the thrust output of gas turbine engine 10. Primary air AP (also known as gas path air) is directed first into LPC 14 and then into HPC 16. LPC 14 and HPC 16 work together to incrementally increase the pressure of primary air A. HPC 16 is rotated by HPT 20 through low-pressure shaft 34 to provide compressed air to combustor section 18. The compressed air is delivered to combustors 18A and 18B, along with fuel through injectors 40, such that a combustion process can be carried out to produce the high-energy gases necessary to turn HPT 20 and LPT 22. Primary air AP continues through gas turbine engine 10 whereby it is typically passed through an exhaust nozzle to further produce thrust.
After being compressed in LPC 14 and HPC 16 and participating in a combustion process in combustors 18A and 18B to increase pressure and energy, primary air AP flows through HPT 20 and LPT 22 such that blades 41 and blades 42 extract energy from the flow of primary air A. Primary air AP impinges on HPT blades 41 to cause rotation of high-pressure shaft 36, which turns HPC 16. Primary air AP also impinges on LPT blades 42 to cause rotation of support rotor 44 and low-pressure shaft 34, turning fan 12 and LPC 14.
The components and configuration of gas turbine engine 10 as shown in
Depicted in
In embodiments of the present disclosure, probe 52 is a microwave sensor for measuring the distance between BOAS 50 and HPT blades 41. In alternative embodiments, probe 52 can be any type of probe. For example, probe 52 can be a non-intrusive stress measurement probe, or a temperature sensor. In the present embodiment, probe 52 is installed in BOAS 50. In alternative embodiments, probe 52 can be used for monitoring an engine component or condition in any actively cooled environment within a hot environment such as within a turbine. Thus, probe 52 can be installed in a turbine environment having a higher temperature than the operating temperature of probe 52. For example, probe 52 can be installed in a vane platform or turbine duct.
In embodiments of the present disclosure, a single probe 52 is housed within housing 58. In alternative embodiments, housing 58 can include multiple probes for monitoring different engine components or conditions. For example, housing 58 can hold both a microwave sensor probe and a non-intrusive stress measurement probe. In the present embodiment, probe housing 58 and bushing 66 are manufactured as separate parts and installed in BOAS 50. In alternative embodiments, BOAS 50 can be manufactured to include structures that function similarly to housing 58 and bushing 66 of the present disclosure.
Housing 58 can tightly hold probe 52 by applying force F to the exterior walls of probe 52. Probe 52 can also include flange 54 to prevent probe 52 from extending into cavity 70. Probe 52 can also include an additional structure or structures to seat probe 52 against housing 58. Housing 58 also includes housing flange 60 to prevent housing 58 from extending radially past bushing tip 68. In this manner, probe tip 56 is held close to flush with the radially inner end of housing tip 62, and the distance between probe tip 56 and primary air AP (cavity 70) created by bushing tip 68 is preserved. In one embodiment of the present disclosure, cavity 70 has a radial depth of about 0.050 inches (0.127 centimeters). In alternative embodiments, cavity 70 can be of any depth necessary to create a barrier of cooling air (described below in greater detail) between probe tip 56 and primary air A.
Pressurizing plenums 64A and 64B and cooling plenums 72 deliver air from an internal cooling air supply (not shown) to probe housing 58, bushing 66 and cavity 70. In embodiments of the present disclosure, the internal cooling air supply supplies air for cooling BOAS 50. In other embodiments, the internal cooling air supply can be air used to cool any number of engine components. Pressurizing plenums 64A and 64B divert air from the internal cooling air supply between housing 58 and probe 52 and between housing 58 and bushing 66, respectively. In one embodiment of the present disclosure, pressurizing plenums 64A receive air from the internal cooling air supply through one or more openings 65 in housing 58. The air fills spaces between probe 52 and housing 58. Pressurizing plenums 64B receive air from the internal cooling air supply between housing 58 and bushing 66 through passage 67 and cooling channel 69 (described in greater detail and shown in
Holes 73 are arranged around the perimeter of cavity 70. In the embodiment shown in
In some embodiments, housing 58 can include radial splits to provide a tight fitting between probe 52 and housing 58.
Discussion of Possible Embodiments
The following are non-exclusive descriptions of possible embodiments of the present disclosure.
A probe cooling system for a gas turbine engine can include a probe housing and a bushing disposed between an end of the probe housing and a gas path, the bushing extending from the end of the probe housing to form a cavity. The system can also include a first plenum defined within the probe housing and configured to direct air from an internal cooling air supply towards the cavity. The system can also include a second plenum defined between the bushing and the probe housing and configured to direct air from the internal cooling air supply towards the cavity.
The system of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
A further embodiment of the foregoing system can include a third plenum extending through a wall of the bushing and into the second plenum.
A further embodiment of any of the foregoing systems can include a fourth plenum extending through a wall of the bushing and into the cavity.
A further embodiment of any of the foregoing systems can include that the probe housing is at least partially located in a blade outer air seal of a gas turbine engine.
A further embodiment of any of the foregoing systems can include that the internal cooling air supply supplies air for cooling the blade outer air seal.
A further embodiment of any of the foregoing systems can include a microwave sensor probe positioned within the probe housing.
A further embodiment of any of the foregoing systems can include a temperature sensor positioned within the probe housing.
A further embodiment of any of the foregoing systems can include a plurality of probes positioned within the probe housing.
A further embodiment of any of the foregoing systems can include that the first plenum is configured to pressurize the probe housing.
A further embodiment of any of the foregoing systems can include a radial split on the probe housing.
A further embodiment of any of the foregoing systems can include that the first and second plenums are configured such that an air pressure in the first plenum is greater than an air pressure in the second plenum, and the air pressure in the second plenum is greater than an air pressure in the cavity.
A further embodiment of any of the foregoing systems can include that an inner portion of the probe housing is cutaway to provide a contact and sealing surface between the probe and the probe housing.
A method of cooling a probe can include positioning a probe having a probe tip proximate a gas path and directing air from an internal cooling air supply through a plurality of plenums towards the probe tip.
The method of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
A further embodiment of the foregoing method can include securing the probe within a housing.
A further embodiment of any of the foregoing methods can include that directing air from the internal cooling air supply through the plurality of plenums pressurizes the probe housing.
A further embodiment of any of the foregoing methods can include that the housing is at least partially located in a blade outer air seal of a gas turbine engine.
A further embodiment of any of the foregoing methods can include creating a bather of cooling air between the probe tip and the gas path by directing cooling air through a wall of a bushing and into a cavity of the blade outer air seal.
A further embodiment of any of the foregoing methods can include that a first plenum is located between the probe and a probe housing, and wherein a second plenum is located between the probe housing and a bushing.
While the disclosure has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the disclosure not be limited to the particular embodiment(s) disclosed, but that the disclosure will include all embodiments falling within the scope of the appended claims.
