Patent Application
20110206502
1. Technical Field
The present invention relates generally to gas turbine engine high pressure turbine shrouds and support thereof.
2. Background Information
A conventional gas turbine engine typically includes a compressor, combustor and turbine, both rotating turbine components such as blades, disks and retainers, and stationary turbine components such as vanes, shrouds and frames routinely require cooling due to heating thereof by hot combustion gases. Cooling of the turbine, especially the rotating components, is important to the proper function and safe operation of the engine. The high pressure turbine (HPT) stages typically maintain a very small tip clearance between turbine blade tips and shrouds surrounding the tips. Shroud supports maintain the shrouds in desired position relative to the rotating blade tips to control clearances between the shrouds and blades. The tip clearance should be made as small as possible for good efficiency, however, the tip clearance is typically sized larger than desirable for good efficiency because the blades and turbine shroud expand and contract at different rates during the various operating modes of the engine.
The turbine shroud has substantially less mass than that of the turbine blades and disk and therefore responds at a greater rate of expansion and contraction due to temperature differences experienced during operation. Since the turbines are bathed in hot combustion gases during operation, they are typically cooled using compressor discharge pressure and/or bleed air suitably channeled thereto.
In an aircraft gas turbine engine for example, acceleration burst of the engine during takeoff provides compressor discharge and/or bleed air which may be hotter than the metal temperature of the turbine shroud. Accordingly, the turbine shroud grows radially outwardly at a faster rate than that of the turbine blades which increases the tip clearance and, in turn, decreases engine efficiency. During a deceleration chop of the engine, the opposite occurs with the turbine shroud receiving compressor discharge and/or bleed air which is cooler than its metal temperature causing the turbine shroud to contract relatively quickly as compared to the turbine blades, which reduces the tip clearance and may cause rubs between the blade tip and the shroud. Accordingly, the tip clearance is typically sized to ensure a minimum tip clearance during deceleration, for example, for preventing or reducing the likelihood of undesirable rubbing of the blade tips against the turbine shrouds. This can damage and prematurely wear out the blade tip and the shroud which also decreases the engine efficiency.
The turbine shroud therefore directly affects overall efficiency or performance of the gas turbine engine due to the size of the tip clearance. The turbine shroud additionally affects performance of the engine since any compressor discharge and/or bleed air used for cooling the turbine shroud is therefore not used during the combustion process or the work expansion process by the turbine blades and is unavailable for producing useful work.
Accordingly, it is desirable to control the reduce the amount of bleed air used in cooling the turbine shroud for maximizing the overall efficiency of the engine.
In order to better control turbine blade tip clearances, some turbine shrouds and their supports are designed to grow and shrink in reacting to changes in compressor discharge pressure and/or bleed air during transient operating conditions such as acceleration bursts and deceleration chops as well as during cruise.
Cooling air used for cooling of the high pressure turbine (HPT) stages effects the operation of the shroud support in controlling the tip clearances between the shrouds and blades. A particular challenge for the HPT shroud support is to accommodate the blade transient response (time constant), such as during engine speed acceleration, while adding as little weight as possible.
It is important to prevent or minimize rubs between the blade tips and the shrouds during transients. It is desirable to minimize the effect of the turbine cooling on the transient response of the turbine shroud support during engine transients.
It is highly desirable to minimize effects on turbine shrouds and their supports due to growth and shrinkage because of changes in compressor discharge pressure and/or bleed air during transient operating conditions such as acceleration bursts and deceleration chops.
A gas turbine engine turbine shroud assembly includes an annular thermal shield circumferentially disposed around a radially outer surface of an annular shroud support. The thermal shield includes a honeycomb layer attached to the shroud support along a radially inner side of the honeycomb layer and a heat shield attached to the honeycomb layer along a radially outer side of the honeycomb layer. The honeycomb layer may be brazed to the shroud support and the heat shield may be brazed to the honeycomb layer.
A more particular embodiment of the gas turbine engine turbine shroud assembly includes a lip depending radially inwardly from a forward end of the heat shield and an axial extension extending axially beyond an aft end of the honeycomb layer. The thermal shield may be segmented.
Another particular embodiment of the gas turbine engine turbine shroud assembly includes at least axially spaced apart first and second stage thermal shields circumferentially disposed around a radially outer surface of an annular shroud support. The first and second stage thermal shields include first and second honeycomb layers attached to the shroud support along first and second radially inner sides of the first and second honeycomb layers respectively and the first and second stage thermal shields include first and second heat shields attached to the first and second honeycomb layers along first and second radially outer sides of the first and second honeycomb layers respectively. The first and second honeycomb layers may be brazed to the shroud support along first and second radially inner sides of the first and second honeycomb layers respectively and the first and second heat shields may be brazed to the first and second honeycomb layers along first and second radially outer sides of the first and second honeycomb layers respectively.
First and second lips may depend radially inwardly from first and second forward ends of the first and second heat shields. First and second axial extensions may extend axially beyond first and second aft ends of the first and second honeycomb layers respectively. The first lip may extend radially inwardly to axially cover a leading edge of the shroud support. The first and second stage thermal shields may be segmented.
A gas turbine engine high pressure turbine may incorporate the gas turbine engine turbine shroud assembly including at least axially spaced apart first and second stage thermal shields. First and second high pressure turbine stages of the high pressure turbine include first and second stage disks respectively and annular high pressure turbine first and second stage shroud assemblies radially spaced apart from and circumscribing first stage and second stage blades extending radially outwardly from the first and second stage disks respectively. First and second stage shrouds of the first and second stage shroud assemblies are coupled to first and second stage shroud hangers respectively and a shroud support include forward and aft hooks radially inwardly supporting the first and second stage shroud hangers. The first and second stage thermal shields are circumferentially disposed around a radially outer surface of the annular shroud support and generally concentric with the first and second stage shroud assemblies.
Illustrated in
Further referring to
The combustion produces hot combustion gases 54 which flow through the high pressure turbine 16 causing rotation of the high pressure rotor 12 and continue downstream for further work extraction in the low pressure turbine 34 and final exhaust as is conventionally known. In the exemplary embodiment depicted herein, the high pressure turbine 16 includes, in downstream serial flow relationship, first and second high pressure turbine stages 55, 56 having first and second stage disks 60, 62. A high pressure shaft 64 of the high pressure rotor 12 connects the high pressure turbine 16 in rotational driving engagement to the impeller 32. A first stage nozzle 66 is directly upstream of the first high pressure turbine stage 55 and a second stage nozzle 68 is directly upstream of the second high pressure turbine stage 56. An annular outer channel 74 disposed between the outer combustor casing 46 and the outer combustion liner 72 extends axially and downstream from an upstream or forward end 49 of the outer combustion liner 72 to at least the first high pressure turbine stage 55.
Referring to
Referring to
A turbine cooling system 137 is used to cool high-pressure turbine (HPT) first stage blades 92 of the first stage disk 60 with clean cooling air 97 in order to minimize sand and/or dirt ingested into HPT blade cooling passages and, thus, prevent blocking of the small blade cooling passages and consequent blade failure. The clean cooling air 97 is bled at a bleed location 95 downstream of an outlet 140 of the diffuser 42 as the CDP air 76 enters the deswirl cascade 44 along an internal radius portion 133 thereof. The clean cooling air 97 bled in this manner is substantially free of particulate matter which could clog fine cooling passages in the first stage blades 92 of the first stage disk 60.
Referring to
The first and second stage shroud assemblies 171, 271 are designed to maintain minimal tip clearances C between first stage and second stage blade tips 192, 194 of the first stage and second stage blades 92, 94 and the first and second stage shrouds 172, 272. The first and second stage shroud assemblies 171, 271 and their components are cooled by the CDP air 76 and supported by the shroud support 114 which is exposed to the CDP air 76.
The first and second stage shroud assemblies 171, 271 thermally expand and contract at relatively faster rates than that of the relatively slower responding, higher mass first stage and second stage blades 92, 94 and first and second stage disks 60, 62 illustrated in
Referring to
Otherwise, open cells 301 of the first and second honeycomb layers 284, 286 are sealed by the first and second heat shields 300, 302 and thus reduce radiant and convective heat transfer between the surrounding CDP air 76 and the shroud support 114. The first and second heat shields 300, 302 also reduce radiant and convective heat transfer between the surrounding CDP air 76 and the shroud support 114 because they prevent slowing of the flow of CDP air 76 over the honeycomb layers thus reducing time of contact and time for heat transfer to occur.
First and second lips 310, 312 on first and second forward ends 314, 316 of the first and second heat shields 300, 302 respectively also reduce heat transfer between the surrounding CDP air 76 and the shroud support 114 by reducing slowing down of the flow of CDP air 76 over the honeycomb layers and further protecting contact between the honeycomb layers and shroud support 114 and the flow of CDP air 76. The first and second lips 310, 312 depend radially inwardly from the first and second forward ends 314, 316 of the first and second heat shields 300, 302 respectively and include radii R of curvature. The first and second lips 310, 312 depend radially inwardly to substantially axially cover the first and second honeycomb layers 284, 286. The first lip 310 extends radially inwardly to axially cover a leading edge 320 of the shroud support 114.
First and second axial extensions 330, 332 extending axially beyond first and second aft ends 334, 336 of the first and second honeycomb layers 284, 286 respectively also reduce heat transfer between the surrounding CDP air 76 and the shroud support 114 by guiding the flow of CDP air 76 aftwardly away from the honeycomb layers and further preventing contact between the shroud support 114 and the flow of CDP air 76 in the area of shrouds. As illustrated in
The first and second stage thermal shields 280, 282 help match the blade and disk transient response to that of the shroud support while adding very little weight to the turbine. The heat shields are brazed to the honeycomb layer and then cut into annular segments or sectors illustrated as the thermal shield segments. The thermal shield segments are brazed onto the HPT shroud support 114. The air in the honeycomb cells of the honeycomb layers act as insulators, therefore the part can be lower in weight and have the proper clearance time constant.
One exemplary heat shield material is 0.020 inches thick and is brazed to thicker honeycomb material about 0.100 inches thick. The thermal shield is at a HPT shroud support diameter and the thermal shield segments may be a 20 or 24 degree sector. The correct number of thermal shield segments are brazed onto the shroud support to cover it 360 degrees for both first and second turbine stages. The thermal shield is brazed on in sectors or segments because of a coefficient of expansion difference between the heat shield and HPT shroud support material. HPT shroud supports are typically made from low coefficient of expansion alloys and because of cost and availability heat shields are not made from low coefficient of expansion alloys. If the heat shields were not in sectors then they might not withstand the stress caused from the low coefficient of expansion difference and temperature difference during transients.
While there have been described herein what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein and, it is therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention. Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims.
