Isolated turbine shroud

A turbine shroud includes an annular shroud support joined to an annular outer casing and spaced radially inwardly therefrom to define an annular flow duct for receiving compressor bleed air. A plurality of shroud panels are joined to the shroud support and have a radially inner surface positionable radially above turbine rotor blades to define a tip clearance therebetween. An isolation shield is joined to the shroud support radially above the panels and includes a plurality of open-ended cells facing radially outwardly in the duct toward the outer casing for inducing drag in the bleed air flowable axially through the duct to thermally isolate the shroud support from the bleed air for controlling the tip clearance.

The present invention relates generally to gas turbine engines, and, more 
specifically, to clearance control between turbine rotor blade tips and a 
stator shroud spaced radially thereabove. 
BACKGROUND OF THE INVENTION 
A gas turbine engine includes in serial flow communication one or more 
compressors followed in turn by a combustor and high and low pressure 
turbines disposed axisymmetrically about a longitudinal axial centerline 
within an annular outer casing. During operation, the compressors are 
driven by the turbine and compress air which is mixed with the fuel and 
ignited in the combustor for generating hot combustion gases. The 
combustion gases flow downstream through the high and low pressure 
turbines which extract energy therefrom for driving the compressors and 
producing output power either as shaft power or thrust for powering an 
aircraft in flight, for example. 
Each of the turbines includes one or more stages of rotor blades extending 
radially outwardly from respective rotor disks, with the blade tips being 
disposed closely adjacent to a turbine shroud supported from the casing. 
The tip clearance defined between the shroud and blade tips should be made 
as small as possible since the combustion gases flowing therethrough 
bypass the turbine blades and therefore provide no useful work. In 
practice, however, the tip clearance is typically sized larger than 
desirable since the rotor 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 rotor 
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 bleed air suitably channeled 
thereto. In an aircraft gas turbine engine for example, acceleration burst 
of the engine during takeoff provides compressor bleed air which is 
actually 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 
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. 
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. 
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 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 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, active clearance 
control systems are known in the art and are relatively complex for 
varying during operation the amount of compressor bleed air channeled to 
the turbine shroud. In this way the bleed air may be provided as required 
for minimizing the tip clearances, and the amount of bleed air may 
therefore be reduced. However, in order to minimize the complexity and 
cost of providing clearance control, typical turbine shrouds are 
unregulated in cooling the various components thereof. 
SUMMARY OF THE INVENTION 
A turbine shroud includes an annular shroud support joined to an annular 
outer casing and spaced radially inwardly therefrom to define an annular 
flow duct for receiving compressor bleed air. A plurality of shroud panels 
are joined to the shroud support and have a radially inner surface 
positionable radially above turbine rotor blades to define a tip clearance 
therebetween. An isolation shield is joined to the shroud support radially 
above the panels and includes a plurality of open-ended cells facing 
radially outwardly in the duct toward the outer casing for inducing drag 
in the bleed air flowable axially through the duct to thermally isolate 
the shroud support from the bleed air for controlling the tip clearance.

DESCRIPTION OF THE PREFERRED EMBODIMENT(S) 
Illustrated in FIG. 1 is an exemplary embodiment of a turbine shroud 10 
which is axisymmetrical about an axial centerline axis 12 in an aircraft 
gas turbine engine. The aircraft engine also includes one or more 
conventional compressors one of which is represented schematically by the 
box 14, with compressed air being channeled to a conventional combustor 
(not shown) in which the air is mixed with fuel and ignited for generating 
hot combustion gases 16 which are discharged axially therefrom. 
Disposed downstream from the combustor is a conventional high pressure 
turbine (HPT) 18 which receives the combustion gases 16 for extracting 
energy therefrom. In this exemplary embodiment, the HPT 18 includes at 
least two stages, with the first stage including the turbine shroud 10 and 
the second stage including a second stage turbine shroud 10b. The first 
and second stages include conventional first and second stage stationary 
turbine nozzles 20,20b each having a plurality of circumferentially spaced 
apart stator vanes extending radially between outer and inner annular 
bands. Disposed downstream from the nozzles 20,20b are respective 
pluralities of circumferentially spaced apart first and second stage 
turbine rotor blades 22,22b extending radially outwardly from first and 
second stage rotor disks 24,24b axisymmetrically around the centerline 
axis 12. 
The first stage turbine shroud 10 illustrated in FIG. 1 is an assembly 
including a corresponding portion of an annular outer stator casing 26 
which provides a stationary support for the several components thereof. 
The outer casing 26 is axially split at a pair of adjacent first and 
second radial flanges 26a and 26b which complement each other and are 
formed as respective integral ends of the casing 26 at the splitline. An 
annular, one-piece shroud ring or support 28 is suspended from the casing 
first and second flanges 26a,b and is used in both the first and second 
stage shrouds 10,10b, as well as supports the second stage nozzle 20b. The 
shroud support 28 is generally L-shaped in transverse section and has an 
annular radial support flange 30 and an integral annular forward support 
leg 32 which extends axially forwardly from a radially inner end of the 
support flange 30. The forward support leg 32 extends axially forwardly 
for supporting respective components of both the first and second stage 
turbine shrouds and the nozzle 20b. 
An annular, multi-segmented first stage shroud hanger 34 is suspended from 
the forward, free distal end of the support 28. An annular one-piece 
second stage shroud ring or hanger 34b is also suspended from the aft end 
of the support 28 at the casing first and second flanges 26a,b and is 
disposed with the shroud support 28 coaxially about the centerline axis 
12. The second stage shroud hanger 34b is generally Y-shaped in transverse 
section and has an annular radial hanger flange, and integral annular 
forward and aft hanger legs at a radially inner end thereof. The forward 
and aft legs extend axially oppositely to each other, with the forward leg 
having a forward hook being conventionally supported on a corresponding 
aft hook 32a of the forward support leg 32. The first stage shroud hanger 
34 is generally H-shaped in transverse section with top forward and aft 
hooks being conventionally supported on a corresponding pair of forward 
hooks 32b and 32c at the distal forward end of the support leg 32. 
A plurality of arcuate first stage shroud panels 36 are conventionally 
removably fixedly joined to the forward hanger hooks 32b,c by 
corresponding hooks. And, a plurality of arcuate second stage shroud 
panels 36b are conventionally removably fixedly joined to the second stage 
shroud hanger 34b. Any suitable mounting arrangement of the respective 
shroud panels 36, 36b may be used for joining them to the respective 
shroud hangers 34, 34b and in turn to the shroud support 28 and outer 
casing 26. 
Each of the shroud panels 36,36b has an outer surface which faces radially 
outwardly towards the bottom surface of the respective shroud hanger 
34,34b. Each panel 36,36b also includes a radially inner surface 38,38b 
which is positionable radially above tips of the respective first and 
second stage rotor blades 22,22b to define respective tip clearances C 
therebetween. 
The support flange 30 and the second stage hanger radial flange are axially 
positioned or sandwiched between the first and second casing flanges 26a,b 
in abutting or sealing contact with each other, with all four flanges 
having a plurality of circumferentially spaced apart, axially extending 
common or aligned bolt holes, with each bolt hole receiving a respective 
bolt 40 (and complementary nut 40a) for axially clamping together the four 
flanges to support the first and second stage shrouds 10,10b and the 
second stage nozzle 20b from the casing 26. 
As shown in FIG. 1, the forward leg 32 of the common shroud support 28 
extends generally axially and parallel to the casing 26 and is spaced 
radially inwardly therefrom to define an annular flow channel or duct 42 
for axially receiving compressor bleed air 14a from the compressor 14. 
The bleed air 14a is suitably channeled to both first and second stage 
shrouds 10, 10b and the second stage nozzle 20b for providing cooling 
thereof against the heating caused by the hot combustion gases 16 during 
operation. Since all of these components are cantilevered or suspended 
from the outer casing 26 at the common casing flanges 26a,b they thermally 
expand and contract at relatively faster rates than that of the relatively 
slower responding, higher mass rotor blades 22, 22b and their respective 
rotor disks 24, 24b. Accordingly, the respective tip clearances c vary in 
size during operation. 
For example, during an acceleration burst typically occurring in take-off 
operation of the gas turbine engine being used for powering an aircraft in 
flight, the bleed air 14a is actually hotter than the stator metal 
temperature of the shrouds 10, 10b. As a result, the stator components 
react and grow outwardly away from the rotor blades 22, 22b relatively 
quickly and at a faster rate than that of the slower expanding blades 22, 
22b and disks 24, 24b, which increases the tip clearances C. During a 
deceleration chop, the opposite occurs, with the stator components 
receiving cooler bleed air than the metal temperature thereof, with 
corresponding thermal contraction at a greater rate than that of the 
respective blades 22, 22b and disks 24, 24b. This reduces the size of the 
tip clearances C and may lead to undesirable tip rubbing unless the tip 
clearances C are initially made relatively large, or the respective 
shrouds 10, 10b are designed for better matching the thermal expansion and 
contraction of the rotor blades 22, 22b. 
In accordance with the present invention as illustrated in FIG. 1, a 
thermal isolation shield 44 is suitably fixedly joined to the shroud 
support 28 at least in part radially above the first stage shroud panels 
36 for reducing the thermal response time of the first stage shroud 10 for 
in turn reducing the variation in the blade tip clearance C for improving 
performance. The isolation shield 44 as illustrated in more particularity 
in FIGS. 2 and 3 and includes a plurality of radially extending open-ended 
cells 46 facing radially outwardly in the flow duct 42 toward the outer 
casing 26 for inducing drag and boundary layer turbulence in the bleed air 
14a flowable axially through the duct 42 to thermally isolate at least the 
distal end portion of the shroud support 28 from the thermal affect of the 
bleed air 14a for controlling the tip clearance C above the first stage 
blades 22. 
As shown in FIG. 2, the outer casing 26 preferably includes an 
aerodynamically smooth or turbulence free inner surface 26c facing 
radially inwardly toward the shield 44. The shield 44 is disposed 
preferably generally parallel to the casing inner surface 26c and is 
spaced radially inwardly therefrom so that the flow duct 42 is effective 
for channeling the bleed air 14a axially aft or downstream through the 
duct 42. The shield 44 thereby is effective for biasing or diverting the 
bleed air 14a to flow radially outwardly therefrom and radially toward the 
casing inner surface 26c. The shield cells 46 preferably extend radially 
perpendicularly to the shroud support 28, and are preferably fully empty, 
although in alternate embodiments they could be partially filled with a 
suitable insulating material. 
Since the cells 46 are open-ended and face radially outwardly, they provide 
a substantial disruptive flow surface for the bleed air 14a being 
channeled axially thereacross. A fluid such as the bleed air 14a will 
follow the path of least resistance, and therefore the drag induced by the 
outer surface of the shield 44 will divert the bleed air 14a to flow more 
closely adjacent to the casing inner surface 26c which is aerodynamically 
smooth in comparison to the shield 44 and correspondingly has less drag. 
Accordingly, the heat transfer between the bleed air 14a and the forward 
support leg 32 of the shroud support 28 is correspondingly reduced which 
reduces the thermal response time of the forward support leg 32 and 
decreases the variation in the tip clearance C. 
In addition to diverting the bleed air 14a radially outwardly toward the 
outer casing 26, the shield 44 also effects a relatively large buffer or 
wake region over the outer surface of the shield 44 which has reduced heat 
transfer capability, which further reduces the heat transfer into the 
forward support leg 32. 
In the preferred embodiment of the invention as illustrated in FIGS. 2 and 
3, the shield cells 46 are defined by relatively thin gauge sheet walls 
for further reducing heat transfer from the bleed air 14a to the forward 
leg 32 of the shroud support 28. The thickness of the cell walls is 
substantially less than the width of the individual cells 46, at least one 
order of magnitude for example. Accordingly, the thin cell walls provide 
minimal material for heat conduction radially inwardly to the forward 
support leg 32. And, the preferably empty cells 46 provide relatively 
stagnant voids which reduces convection heat transfer radially inwardly. 
In the exemplary embodiment of the invention illustrated, the shield 44 
may be made of a suitable metal such as Hastalloy-X, although other 
suitable metals or insulating non-metal materials may be used. The open 
cell structure of the shield 44 therefore results in a relatively 
lightweight and inexpensive shield which can significantly reduce the 
otherwise quick thermal response time of the first stage turbine shroud 
10. Accordingly, the tip clearance is better controlled and may be 
maintained relatively small for maximizing efficiency of operation of the 
engine. 
As shown in FIG. 1, the shield 44 extends axially aft or downstream from 
the free distal end of the forward support leg 32, and continues axially 
downstream over most of the second stage turbine nozzle 20b to adjacent 
the second stage shroud hanger 34b along the outer surface of the forward 
support leg 32. The shield 44 may extend axially over the forward support 
leg 32 as desired for controlling the thermal response thereof due to the 
bleed air 14a being channeled downstream through the flow duct 42. In the 
exemplary embodiment illustrated in FIG. 1, the shield 44 is axially 
coextensive with the first stage shroud hanger 34 and the shroud panels 36 
supported therefrom, and additionally extends axially coextensively with a 
major portion of the second stage turbine nozzle 20b. In this way, thermal 
response of both the first stage shroud 10 and the second stage turbine 
nozzle 20b may be controlled for reducing thermal mismatch relative to the 
various rotor components. 
As shown in FIG. 1 and 2, the second stage turbine nozzle 20b includes at 
least one and preferably several radially extending inlet ports 48 
disposed in flow communication with the flow duct 42 for receiving a 
portion of the bleed air 14a therefrom for conventionally internally 
cooling the turbine nozzle 20b. Correspondingly, the shield 44 as 
illustrated in FIG. 2 and 3 includes a respective access hole 50 therein 
for receiving the respective nozzle inlet port 48. 
Accordingly, the shield 44 may extend axially as desired, and the shield 44 
may be a fully annular structure disposed about the centerline axis 12 
(see FIG. 1) and may be configured as a full 360.degree. continuous ring 
or as discrete arcuate segmented portions. In the exemplary embodiment 
illustrated in the Figures, the shield 44 comprises an opened-celled 
honeycomb with hexagonal cells 46 conventionally brazed at their radially 
inner ends to the outer surface of the forward support leg 32. As shown in 
FIG. 2, the forward support leg 32 may have a suitable recess in which the 
shield 44 may be positioned and brazed. Since the forward distal end of 
the forward support leg 32 is spaced radially inwardly from the inner 
surface 26c of the outer casing 26 as shown in FIG. 2, it provides an 
inlet in the form of an annulus to the flow duct 42. Since the forward 
support leg 32 may transiently thermally expand during operation, the 
shield 44 preferably includes a stepped leading edge 44a for increasing 
the radial depth of the flow duct 42 at the shield leading edge for 
reducing closing of the flow duct 42 thereat due to transient thermal 
expansion of the forward support leg 32 relative to the outer casing 26. 
The shield leading edge 44a has a finite axial length which is parallel to 
the remaining downstream portion of the shield, with an arcuate ramp 
therebetween. This ramp configuration assists in diverting the bleed air 
14a radially outwardly away from the shield 44 and toward the casing inner 
surface 26c. 
According, the isolation shield 44 not only provides an effective thermal 
insulator around the forward support leg 32 of the shroud support 28, but 
also effects a flow diverter for directing bleed air 14a radially 
outwardly away from the shield 44 itself. Thermal insulation is therefore 
effected by the shield 44 by its thin and open celled structure which 
reduces conduction and convection heat transfer; diverts the bleed air 14a 
literally away from the shield 44; and by the buffer region or layer over 
the outer surface of the shield 44 which itself decreases heat transfer 
therethrough. The collective effects of the isolation shield 44 decrease 
the thermal response of the first stage shroud hanger 34 and the shroud 
panels 36 supported therefrom, to in turn reduce the variations of the tip 
clearance C for improving operating efficiency of the high pressure 
turbine 18. 
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.