U.S. patent application number 10/914077 was filed with the patent office on 2005-04-28 for turbine blade tip clearance system.
Invention is credited to Phipps, Anthony B..
Application Number | 20050089401 10/914077 |
Document ID | / |
Family ID | 28052554 |
Filed Date | 2005-04-28 |
United States Patent
Application |
20050089401 |
Kind Code |
A1 |
Phipps, Anthony B. |
April 28, 2005 |
Turbine blade tip clearance system
Abstract
A stage of gas turbine engine turbine blades (22) is surrounded
by segments (26) that in turn, are suspended from a flexible casing
(24). The inner surface of casing (24) is exposed to compressor
flow. On engine thrust being increased, compressor output also
increases and rapidly flexes casing (24) radially outwards, thus
equally rapidly moving the segments (24) away from the tips of the
turbine blades (22).
Inventors: |
Phipps, Anthony B.;
(Uttoxeter, GB) |
Correspondence
Address: |
MANELLI DENISON & SELTER
2000 M STREET NW SUITE 700
WASHINGTON
DC
20036-3307
US
|
Family ID: |
28052554 |
Appl. No.: |
10/914077 |
Filed: |
August 10, 2004 |
Current U.S.
Class: |
415/178 |
Current CPC
Class: |
Y02T 50/671 20130101;
Y02T 50/60 20130101; F01D 11/16 20130101; F01D 11/22 20130101 |
Class at
Publication: |
415/178 |
International
Class: |
F01D 005/08 |
Foreign Application Data
Date |
Code |
Application Number |
Aug 15, 2003 |
GB |
0319180.6 |
Claims
I claim:
1. A turbine blade tip clearance system comprising a rigid outer
casing and an inner casing having induced flexing capability
supported by the outer casing in radially spaced relationship
therewith, so as to define an annular space therebetween, and
wherein said inner casing supports a ring of arcuate segments
within it in fixed radial relationship therewith, such that on
placing the whole around a stage of disk mounted turbine blades in
coaxial relationship therewith, said segments will lie in radial
close spaced relationship with the tips of said turbine blades.
2. A turbine blade tip clearance system as claimed in claim 1
wherein said arcuate segments are suspended from said inner casing
by screw threaded fastening means so as to enable radial positional
adjustment of said arcuate segments with respect to a selected
datum.
3. A turbine blade tip clearance system as claimed in claim 1
wherein said outer casing is manufactured from a temperature
sensitive material.
4. A turbine blade tip clearance system as claimed in claim 3
wherein said material comprises nickel.
5. A turbine blade tip clearance system as claimed in claim 1 when
installed in a gas turbine engine.
6. A gas turbine engine including an installed turbine blade tip
clearance system as claimed in claim 5 and including ducting
connecting a compressor of said gas turbine engine in flow series
with the inner surface of said inner casing so as to cause said
inner casing to flex outwardly of the axis of said gas turbine
engine and thereby lift said segments away from said blade tips
during an appropriate stage in the engine operating regime.
7. A gas turbine engine as claimed in claim 6 including ducting
connecting a compressor of said gas turbine engine in flow series
with said outer casing so as to heat said outer casing and cause it
to expand and thereby slow the rate of return of said segments to
their original positions, during an appropriate stage in the engine
operating regime.
Description
[0001] The present invention relates to gas turbine engine blade
tip clearance systems, wherein such an engine, the outer wall of
the gas annulus that surrounds a stage of turbine blades comprises
segments that are moveable relative to the blades in directions
radially of the engine axis. The arrangement enables a reduction in
blade tip rub on the gas annulus wall when the blades and
associated turbine disk grow under the influence of heat and
centrifugal forces. Also, on slowing and cooling of the assembly,
the arrangement enables the spacing of the segments from the blade
tips by a distance that reduces performance losses.
[0002] Known art utilises active devices i.e. complicated sensing
devices for sensing the relative movement between blades and
segments, which devices, on sensing a change in gap magnitude,
develop signals that are passed to actuating mechanisms. These, in
turn actuate segment moving means, and thereby move the segments in
an appropriate direction.
[0003] Devices of the kind generally described hereinbefore have
drawbacks over and above their complicated system of operation.
They are heavy, which generates weight penalties where the
associated engine is utilised in an aircraft. They are expensive to
manufacture, and further, they cannot react with appropriate
efficiency so as to cater for both the rapid growth of the disk and
blade assembly during engine acceleration, and its much slower
reduction in size on deceleration and cooling.
[0004] The present invention seeks to provide an improved turbine
blade tip clearance system.
[0005] According to the present invention a turbine blade tip
clearance system comprises a rigid outer casing and an inner casing
having induced flexing capability supported by the outer casing in
radially spaced relationship therewith so as to define an annular
space therebetween, and wherein said inner casing supports a ring
of segments within it in fixed radial relationship therewith, such
that on placing of the whole around a stage of disk mounted turbine
blades in co-axial relationship therewith, said segments will lie
in radially close spaced relationship with the tips of said turbine
blades.
[0006] The present invention will now be described, by way of
example and with reference to the accompanying drawings in
which:
[0007] FIG. 1 is a diagrammatic sketch of a gas turbine engine
including a turbine blade tip clearance system in accordance with
the present invention.
[0008] FIG. 2 is an enlarged cross sectional part view of the
turbine blade tip clearance system of FIG. 1.
[0009] Referring to FIG. 1. A gas turbine engine indicated
generally by the numeral 10, comprises a compressor 12, combustion
equipment 14, a turbine section 16, and an exhaust duct 18. Turbine
section 16 includes a rigid outer casing 20 that surrounds a stage
of turbine blades 22. Casing 20 supports an inner casing 24, and
inner casing 24, in turn, supports a ring of arcuate segments 26,
the radially inner surfaces 28 of which lie closely adjacent the
tips of turbine blades 22.
[0010] Referring now to FIG. 2. In the example, casing 24 is
conical in form and is supported via its respective beaded ends 30
and 32 in grooves formed in internal flanges 34 and 36 on casing
20. Pairs of bosses 38 and 40, only one pair of which is shown, are
angularly spaced around inner casing 24, each pair being arranged
in axial alignment relative to the axis of engine 10. Bosses 38 and
40 are drilled and tapped so as to accept respective screw threaded
bolts 42 and 44, the heads 46 and 48 of which are shaped so as to
enable their location in forked ends of brackets 50 and 52 that
extend from the radially outer surface of each segment 26. Segments
26 will each have the same number of axially aligned brackets 50
and 52, as there are bosses. Stopper bars 54 are fixed to flanges
34 and 36, each side of bosses 38 and 40, so as to prevent
excessive radially outward flexing of inner casing 24 during
operation of engine 10, as is explained later in this
specification.
[0011] Assembly of the whole can be achieved by first screwing
bolts 42 and 44 through respective bosses 38 and 40, in a direction
radially outwardly of casing 24, followed by inserting the grooved
bolt heads 46 and 48 in the respective forked ends of segment
brackets 50 and 52. Nuts 56 and 58 are then screwed on to the
extremities of the respective projecting ends of bolts 42 and 44,
but not tightened against their respective bosses. The now loosely
juxta positioned segments 26 are slid onto the land of a disk
shaped jig 60, the diameter of which corresponds to the diameter of
the stage of turbine blades 22, plus a cold clearance margin i.e.
the required clearance 62 between the tips of blades 22 (shown by a
dashed line) and the adjacent inner surfaces of segments 26, when
associated engine 10 is inoperative.
[0012] It is preferable that jig 60 is supported for rotation about
its axis, so that the loosely fitted segments at the bottom of the
assembly can, in turn, be brought to the top to ease the
positioning of the respective parts.
[0013] The positioning of the loosely assembled parts is achieved
as follows: Spacers (not shown) of appropriate length are placed
between the undersides of the bolts heads and the opposing inner
end faces of their respective bosses. The bolts are then screwed
further through bosses 38 and 40 until the spacers are lightly
trapped between respective bolt heads and bosses inner faces. Each
spacer (not shown) is then removed. Nuts 56 and 58 are then screwed
along their respective bolts so as to engage the outer ends of
their respective bosses 38 and 40. These steps are repeated all
around the assembly and results in the clamping of all of the
segments 26 in co-axial relationship with inner casing 24 and,
nominally, in a desired spaced relationship with the tips of the
stage of turbine blades 22 around which the assembly is to be
fitted.
[0014] The assembly as described so far is now removed from jig 60,
and fitted into casing 20 by inserting the beaded edge of what will
be the upstream end of casing 24 with respect to the flow of gases
through engine 10 (FIG. 1) in the annular groove in flange 34. The
downstream beaded end of casing 24 is then inserted in the groove
in flange 36, which is then offered up to a small flange 64 within
casing 20, and bolted thereto. The completed assembly is then
fitted over the pre-assembled stage of turbine blades. Should
further adjustment prove necessary, due to manufacturing tolerances
of the various parts, this can be achieved before fitting casing 20
to the casing of combustion equipment 14 (FIG. 1) upstream thereof,
by accessing nuts 56 and 58 through normally capped apertures 66
and 68 in casing 20.
[0015] Gas turbine engine 10 (FIG. 1) is of the kind utilised to
power aircraft. The power regime of engine 10 embraces aircraft
taxiing, take off, cruise and landing, all of which require
different engine power outputs. Thus, gas and air temperatures and
pressures, and speed of revolution of rotary parts change in
concert. During taxiing of the associated aircraft (not shown), no
adverse temperature is experienced by the turbine system. However,
when the throttle is opened to obtain full power so as to enable
takeoff, there occurs an almost instantly considerable rise in
compressor output pressure, combustion gas temperature, and speed
of revolution of the turbine section.
[0016] The combination of increased temperature and speed of
revolutions of the turbine stage causes the latter to increase its
diameter through centrifugal forces and increased heat. The
magnitude of the increase is such that, if the ring of segments 26
(a common inclusion in gas turbine engine turbine systems) remains
in its cold position, or moves too slowly radially outwards with
respect to the turbine stage, severe rubbing of the tips of the
turbine blades on the segments would occur, with consequent loss of
material from blade tips and segments. The resultant permanently
increased annular gap therebetween would cause severe performance
losses over the entire operating regime of the engine. In the
present arrangement however, the increased pressure output from the
compressor 12 (FIG. 1) enables a portion of that output to be
immediately diverted to the radially inner surface of inner casing
24 by any known ducting (not shown) so as to induce flexing of
casing 24 radially outwards against the stopper bars 54, and by
this means, rapidly lift segments 26 away from the tips of the
turbine blades. The time lag between the start of turbine growth
and flexing of casing 24 is so small as to minimise any rubbing
action that may occur, and consequently reduces efficiency
losses.
[0017] When the engine 10 is throttled back e.g. when the
associated aircraft reaches its cruise altitude, the turbine stage
22 contracts more slowly than it expands. Compressor output
pressure also reduces and consequently reduces the force exerted on
casing 24, which then could return too quickly to its non flexed
shape and so cause rubbing between segments 26 and the tips of
blades 22. Again, the present arrangement provides means to avoid
rubbing through contraction, by making casing 20 from a material,
the magnitude and rate of expansion and contraction of which can be
controlled by heating and cooling. The nickel alloy marketed under
the registered trade mark "Waspaloy" is one such material.
[0018] On throttling back of engine 10, with consequent reduction
in pressure on casing 24, hot compressor air could be ducted (not
shown) onto casing 20 so as to rapidly heat it and cause it to
expand in a radially outwards direction. The movement is
transmitted via flanges 34 and 36, to the sub assembly of casing 24
and segments 26, and has the effect of slowing its rate of radially
inward movement to a rate more compatible with that of the turbine
stage. Rubbing of segments 26 on the blade tips during contraction
is thus minimised.
[0019] Reference to FIG. 2 shows that casings 20 and 24, and
flanges 34 and 36 define an annular chamber 70, which will fill
with engine leakage air. It is important that the leakage air
pressure in chamber 70 is prevented from reaching a magnitude such
that casing 24 is flexed radially inwards, or is prevented from
being flexed radially outwards at appropriate speed as and when
required. To this end, flange 36 has at least one vent hole 72
through its thickness.
* * * * *