BACKGROUND OF THE INVENTION
[0001] The subject matter disclosed herein relates to clearance control systems for gas
turbine engines, and more particularly, to a system and method for controlling operating
clearance between a stationary shroud surface of a turbine engine and an adjacent
rotating assembly.
[0002] Gas turbine engines typically include a rotating assembly housed within a static
assembly. The rotating assembly typically includes sets of compressor blades (i.e.,
airfoils) and turbine blades. The compressor blades compress incoming air, and the
turbine blades extract power from the air, usually after the addition of heat. The
static assembly surrounds the rotating assembly and helps to define the flow path
of the engine. Typically, a static assembly includes a series of shroud segments providing
an inner surface to cooperate with outer surfaces (i.e., tips) of the blades of an
adjacent rotating assembly. Efficiency of a turbine engine depends, at least in part,
on the clearance or gap between the shroud surface and the adjacent rotating blades.
If the clearance is excessive, a correspondingly excessive fraction of engine flow
will pass through the gap rather than interacting with the rotating blades, resulting
in reduced engine efficiency. If the clearance is too small, interference can occur
between the rotating and stationary members of such assemblies, resulting in damage
to one or more of the cooperating surfaces.
[0003] As used herein, the term axial refers to the direction of the central axis of a gas
turbine engine, i.e., about which axis the turbo-machinery rotates. The term radial
refers to a direction that is substantially perpendicular to the central axis, and
the term circumferential refers to a set of locations and directions that do not intersect
the central axis but that lie in one or more radial planes.
[0004] Complicating clearance problems in such apparatus is the well known fact that clearances
between the rotor assembly and the static assembly of a turbine engine typically change
with engine operating conditions such as acceleration, deceleration, or other changing
thermal or centrifugal force conditions experienced by the cooperating members during
engine operation. Clearance control mechanisms for such assemblies, sometimes referred
to as active or passive clearance control systems, have included mechanical systems
or systems based on thermal expansion and contraction characteristics of materials
for the purpose of maintaining selected clearance conditions during engine operation.
Such systems generally require use of substantial amounts of air for heating or cooling
at the expense of such air otherwise being used in the engine operating cycle.
[0005] Accordingly, those skilled in the art seek improved means for controlling operating
clearance between a stationary shroud surface of a turbine engine and an adjacent
rotating assembly.
BRIEF DESCRIPTION OF THE INVENTION
[0006] According to one aspect of the invention, a system for passively controlling clearance
in a gas turbine engine comprises a static assembly arranged circumferentially about
an engine rotor assembly and defining a gap between a tip end of the rotor assembly
and an inner surface of the static assembly adjacent to the tip end. The system includes
a gap control member that defines the inner surface, is exposed to the engine working
fluid, and comprises a shape memory material selected and preconditioned to deform
in a pre-selected manner in response to a temperature of the engine working fluid.
[0007] The system may further comprise a rotor assembly having a plurality of airfoil blades,
each blade having a tip end. The rotor assembly is surrounded by the static assembly
comprising a plurality of shroud segments arranged circumferentially about the rotor
assembly, each shroud segment having an inner surface adjacent to the tip end, and
the inner surfaces of the shroud segments and the tip ends of the airfoil blades defining
a radial gap between the tip ends and the inner surfaces. Each airfoil blade includes
the gap control member that forms the tip end, the gap control member comprising a
shape memory material selected and preconditioned to deform in a pre-selected manner
in response to a temperature of the engine working fluid.
[0008] According to yet another aspect of the invention, a method for passively controlling
clearance in a gas turbine engine comprises assembling the turbine engine so as to
define an initial set of build clearances between a stationary shroud surface of the
turbine engine and an adjacent rotor assembly of the turbine engine. The assembled
engine is operated throughout a range of engine operating conditions, and an operating
clearance is observed at one or more of the engine operating conditions. A gap control
member is formulated and configured comprising a shape memory material selected and
preconditioned to deform in a pre-selected manner in response to a temperature of
the engine working fluid. The engine is re-assembled with the gap control member so
as to define a revised set of build clearances between the stationary shroud surface
and the adjacent rotor assembly.
[0009] These and other advantages and features will become more apparent from the following
description taken in conjunction with the drawings.
BRIEF DESCRIPTION OF THE DRAWING
[0010] Embodiments of the present invention will now be described, by way of example only,
with reference to the accompanying drawings in which:
FIG. 1 is a partially sectional view of a gas turbine engine including a static assembly
and an adjacent blade, the view including one exemplary embodiment of a gap control
member;
FIG. 2 is a partially sectional view of a gas turbine engine including a static assembly
and an adjacent blade, the view including another exemplary embodiment of a gap control
member with an abradable coating;
FIG. 3 is a partially sectional view of a gas turbine engine including a static assembly
and an adjacent blade, the view including another exemplary embodiment of a gap control
member;
FIG. 4 is an enlarged view depicting a portion of a gas turbine engine wherein an
exemplary shroud with a gap control member and an abradable coating is in an open-clearance
condition relative to the tip of an adjacent rotating blade;
FIG. 5 is an enlarged view depicting a portion of a gas turbine engine wherein an
exemplary shroud with a gap control member and an abradable coating is in a closed-clearance
condition relative to the tip of an adjacent rotating blade;
FIG. 6 is an enlarged view depicting a portion of a gas turbine engine wherein exemplary
shroud segments include gap control members configured to compensate for eccentricity
or other non-circularity in the assembled static structure; and
FIG. 7 is a flow chart of an exemplary method for reduced operating clearance between
a stationary shroud surface of a turbine engine and an adjacent rotating assembly.
[0011] The detailed description explains embodiments of the invention, together with advantages
and features, by way of example with reference to the drawings.
DETAILED DESCRIPTION OF THE INVENTION
[0012] FIG. 1 shows a portion of a gas turbine engine 100 comprising a rotating assembly
170 housed within a static assembly 160. Rotor assembly 170 carries a rotating blade
110, which has a tip end 112 and an apposing hub end 118. Rotating blade 110 also
has a leading edge 114 and a trailing edge 116. As shown in FIG. 1, rotating blade
is a turbine blade, but it should be appreciated that the features shown could be
applied to a compressor.
[0013] Static assembly 160 includes stator 180, which guides a working fluid, such as air
or steam or air mixed with fuel, toward the leading edge 114. Static assembly also
includes shroud segments 120 that guide the working fluid through rotating blade 110
so that rotating blade 110 can extract energy (i.e., torque) from the fluid (or, in
the case of a compressor, so that the blade can perform work on the fluid). Each shroud
segment 120 has an inner shroud surface 122 on which gap control member 130 is attached.
Gap control member 130 is exposed to the working fluid and has an inner controlled
surface 132 facing radially inward toward tip end 112.
[0014] As rotor assembly 170 rotates about its central axis, tip end 112 travels adjacent
to inner controlled surface 132, defining clearance gap 150. To reduce the size of
clearance gap 150, static assembly 160 includes means for adjusting the radial position
of shroud segment 120, including radial adjustment member 124, and shims 126.
[0015] Gap control member 130 comprises a shape memory material. A suitable shape memory
material may comprise an alloy or a polymer or another material known in the art for
providing a desired shape memory behavior characteristic. For example, a metallic
shape memory alloy (SMA) is a metal alloy that changes from an initial shape to a
second shape upon exposure to a transition temperature and changes back to the initial
shape upon re-cooling. SMA materials that exhibit such shape changes with temperature
typically undergo a solid state micro-structural phase change. This characteristic
enables an article made from SMA to change from one physical shape to at least another
physical shape and to return to the original shape. These changes in shape are much
more dramatic than simple thermal expansion and contraction. In addition, with SMA,
most or all of the changes in shape occur over a relatively small temperature range
known as the transition temperature of the material. One example of a metallic SMA
material is a titanium nickel alloy, also known as Nitinol alloy. Other metallic SMA
materials may comprise ruthenium alloyed with niobium and/or tantalum. Transition
temperatures of exemplary shape memory materials depend upon the particular composition
of the material and can be configured to occur at temperatures between approximately
25 degrees C and about 1400 degrees C, with the transition temperature depending upon
the specific formulation of the material.
[0016] In the manufacture (from such a metallic SMA or other shape memory material) of an
article intended to change during operation from one shape to at least one other shape,
the article is provided in a first shape intended for operating use at or above the
transition temperature. Such first shape is developed by working and annealing an
article comprising the alloy or other material at or above the transition temperature,
at which the solid state micro-structural phase change occurs. However, below that
critical temperature, such an alloy or other material may be malleable such that the
article of the first shape can be deformed into a desired second shape, for example,
to facilitate inclusion at substantially room temperature in an assembly. Thereafter,
for example in service operation of the article, when the article in the second shape
is heated at or above its critical temperature, it undergoes a micro-structural phase
change that results in it returning to the first shape.
[0017] As noted above, gap control member 130 comprises a shape memory material and is exposed
to the working fluid at its axial location in the flow-path. Therefore, while gap
control member 130 may exchange some heat with shroud segment 120, the temperature
of gap control member 130, under steady-state conditions, will approximate the temperature
of the working fluid at its axial location. Thus, by formulating the material used
to make gap control member 130, its shape can be programmed to change depending upon
the flow-path temperature without requiring parasitic extraction of working fluid.
[0018] Contrariwise, radial adjustment member 124, which may also comprise shape memory
material, is not typically directly exposed to the working fluid at its axial location.
Instead, radial adjustment member 124 may be exposed to a mixture of fluid sources,
enabling the temperature of the fluid to be actively controlled, and thereby enabling
the shape of radial adjustment member 124 to be controlled. Yet, while the shape of
radial adjustment member 124 may thus be controlled, doing so requires parasitic extraction
of working fluid, which may mitigate performance gains that would otherwise be realized
through the active clearance control scheme.
[0019] As discussed above, the size of clearance gap 150 depends upon a number of factors
including initial build clearance, thermal expansion and/or contraction of static
assembly 160 and rotor assembly 170, centrifugal stresses resulting from the rotational
speed of the rotor assembly 170, external loads, aerodynamic loads, and other effects.
These factors can cause the size of clearance gap 150 to change throughout the operational
envelope of engine 100. In addition, the size of clearance gap 150 adjacent to leading
edge 114 may not be equal to the size of clearance gap 150 adjacent to trailing edge
114. Moreover, as the shape of static assembly 160 may not be round, and thus the
circumferential surface defined by the combination of inner controlled surfaces 132
may also not be round, the size of clearance gap 150 may vary from one shroud segment
to another.
[0020] To compensate for these variations in the size of clearance gap 150, both the first
shape and the second shape of gap control member 130 and the transition temperature
of the shape memory material can be configured to contribute to a system for reducing
the size of clearance gap 150. Other elements of an exemplary system may optionally
include one or more active clearance control mechanisms such as radial adjustment
member 124. Other passive elements may also be included such as shims 126. In practice,
an engine may be assembled with relatively open clearances, and then operated throughout
a range of operating conditions while detecting the operating clearances. Then, based
on the observed data, one or more clearance adjustment mechanisms may be implemented
so as to achieve a desired level of clearances.
[0021] In an exemplary embodiment, inner controlled surface 132 of gap control member 130
may exhibit a planar shape. In another embodiment, inner controlled surface 132 of
gap control member 130 may exhibit a non-planar shape. In an exemplary embodiment,
gap control member 130 may be configured to retain a first shape at temperatures less
than 100 degrees C. In another exemplary embodiment, gap control member 130 may be
configured to retain a first shape at temperatures less than 200 degrees C. In another
exemplary embodiment, gap control member 130 may be configured to retain a first shape
at temperatures less than 300 degrees C. Other embodiments of gap control member 130
may be formulated to change shape at temperatures of approximately 400 degrees C,
500 degrees C, 600 degrees C, 700 degrees C, 800 degrees C, or any other operating
temperature where it is advantageous to change the shape of gap control member 130.
[0022] As shown in FIG. 2, static assembly 260 may include an abradable layer 240 disposed
between gap control member 230 and blade 210. As systems and methods are implemented
to reduce the size of clearance gap 250, risk increases that an inadvertent rub may
occur between tip end 212 and any adjacent structure that is positioned radially outwardly
from tip end 212. Abradable layer 240 may comprise a coating applied to a radially
inward surface of gap control member 230 and may comprise a material that can deform
or be abraded in the event of contact with tip end 212 without damaging tip end 212
or blade 210. Incorporation of abradable layer 240 will allow for closer clearances
and offsetting the need to account for thermal expansion as well and changes in concentricity
due to shock loading events.
[0023] Abradable layer 240 may be applied through thermal spraying, sintering, casting or
any other suitable means known in the art. Thermal spraying involves sprayed application
of melted or heated material. Sintering involves application of powdered metal followed
by heating of the composite article. As shown in FIG. 3, a gap control member 330
may also be applied to a tip end 312 of blade 310.
[0024] FIG. 4 shows an enlarged view of the region of a gas turbine engine between a static
assembly 420 and a rotor assembly 410. Gap control member 430 is attached to an inner
shroud surface 422 of static assembly 420, and an abradable layer 440 is attached
to gap control member 430 adjacent to tip end 412. As shown in FIG. 4, clearance gap
450 is relatively open, corresponding to a first shape of gap control member 430 that
is relatively thin. A first shape of gap control member 430 occurs when the operating
temperature of the working fluid is below the transition temperature of gap control
member 430.
[0025] In juxtaposition, FIG. 5 shows an enlarged view of the same region as Fig. 4, wherein
clearance gap 550 is relatively closed, corresponding to a second shape of gap control
member 530 that is relatively thick. A second shape of gap control member 530 occurs
when the operating temperature of the working fluid is above the transition temperature
of gap control member 530.
[0026] FIG. 6 shows an enlarged view of a portion of a gas turbine engine wherein exemplary
shroud segments 620 include gap control members 630 configured to compensate for eccentricity
or other non-circularity in the assembled static assembly 660. As shown in Fig. 6,
a relatively thin and constant clearance gap 650 is provided between blade 610 and
inner shroud surface 622 by incorporation of gap control members 630. It should be
noted that gap control members at 632 are relatively thin compared to gap control
members at 634.
[0027] FIG. 7 is a flow chart showing an exemplary method for reduced operating clearance
between a stationary shroud surface of a turbine engine and an adjacent rotating assembly.
As shown in FIG. 7, an engine is assembled (step 710) comprising a static assembly
and a rotor assembly. The engine is operated (step 720) throughout a range of operating
conditions, and clearances are measured (step 730) at those operating conditions.
Based on those measurements, a clearance control strategy is devised (step 740) considering
available clearance control methods. Then, shape memory materials are formulated and
configured (step 750) so as to configure customized gap control members that are capable
of achieving desired shape changes at defined engine operating temperatures. The strategy
is then implemented (step 760) and may comprise adjusting control schedules so as
to maintain a desired level of clearances without rebuilding the engine or otherwise
re-shimming or adjusting the static assembly of the engine. Once the engine has been
re-assembled (step 770) with the clearance control strategy, clearances can again
be evaluated (step 780) to determine the effectiveness of the implemented strategy.
Finally, steps 740 through this step may be repeated (step 790) until a desirable
clearance profile has been achieved.
[0028] Thus, the invention provides an improved system and method for reducing clearances
and thereby improving gas turbine performance and efficiency. In accordance with the
invention, shape memory materials are preconditioned to deform only upon achieving
a predetermined temperature level, such as steady-state operating temperatures. As
the gap control members that comprise the shape memory materials are positioned in
or near to the working fluid, there is no external actuation medium required to actuate
the gap control members. The invention provides a simple method for addressing eccentricity
or other non-circularity in static assemblies and in operation of rotor assemblies
and can be applied to compressor and turbine sections. In addition, the invention
can be applied to address transient differences in dimensions of static assemblies
and rotor assemblies and to address variations in manufacturing.
[0029] While the invention has been described in detail in connection with only a limited
number of embodiments, it should be readily understood that the invention is not limited
to such disclosed embodiments. Rather, the invention can be modified to incorporate
any number of variations, alterations, substitutions or equivalent arrangements not
heretofore described, but which are commensurate with the spirit and scope of the
invention. Additionally, while various embodiments of the invention have been described,
it is to be understood that aspects of the invention may include only some of the
described embodiments. Accordingly, the invention is not to be seen as limited by
the foregoing description, but is only limited by the scope of the appended claims.
1. A system for passively controlling clearance in a gas turbine engine (100) comprising:
a static assembly (160),
the static assembly (160) being arranged circumferentially about an engine rotor assembly
(170) and defining a gap (150) between a tip end (112) of the rotor assembly (170)
and an inner surface (132) of the static assembly (160),
at least one gap control member (130) defining the inner surface (132) adjacent to
the tip end (112) and comprising a shape memory material selected and preconditioned
to deform in a pre-selected manner in response to a temperature of a working fluid
of the engine (100), and
the at least one gap control member (130) being exposed to the working fluid.
2. The system of claim 1, wherein the at least one gap control member (130) comprises
an abradable coating.
3. The system of claim 1 or claim 2, wherein the shape memory material comprises an alloy.
4. The system of claim 3, wherein the alloy comprises one of ruthenium, niobium or tantalum.
5. The system of any of claims 1 to 3, further comprising:
a rotor assembly (170) comprising a plurality of airfoil blades (210), each of the
airfoil blades (210) having a tip end;
the rotor assembly (170) being surrounded by the static assembly (160) and comprising
a plurality of shroud segments (120) arranged circumferentially about the rotor assembly
(170);
each shroud segment (120) having an inner surface (132) adjacent to the tip end (212);
each of the airfoil blades (210) comprising the gap control member (230) at the tip
end (212);
the inner surfaces (132) of the shroud segments (170) and the tip ends (112) of the
airfoil blades (210) defining the radial gap (250) between the tip ends (212) and
the inner surfaces (132);
6. The system of claim 5, wherein the airfoil blade (210) is one of an axial compressor
blade, an axial turbine blade, a centrifugal compressor blade, or a radial turbine
blade
7. A method for passively controlling clearance in a gas turbine engine (100) comprising:
assembling (710) the turbine engine (100) so as to define an initial set of build
clearances (250) between a stationary shroud surface (260) of the turbine engine (100)
and an adjacent rotor assembly (260) of the turbine engine (100);
operating (720) the assembled turbine engine (100) throughout a range of engine operating
conditions;
observing (730) an operating clearance (250) at one or more of the engine operating
conditions;
formulating and configuring (750) a gap control member (230)comprising a shape memory
material selected and preconditioned to deform in a pre-selected manner in response
to a temperature of an engine working fluid; and
re-assembling (770) the turbine engine (100) with the gap control member (230) so
as to define a revised set of build clearances between the stationary shroud surface
(260) and the adjacent rotor assembly (170).
8. The method of claim 7, further comprising, subsequently operating (790) the assembled
turbine engine throughout a range of engine operating conditions and observing a revised
operating clearance at one or more of the engine operating conditions.
9. The method of claim 7 or 8, further comprising formulating and configuring (790) an
additional gap control member comprising a shape memory material selected and preconditioned
to deform in a pre-selected manner in response to a temperature of the engine working
fluid.
10. The method of any of claims 7 to 9, wherein the gap control member (230) comprises
an abradable coating.
11. The method of any of claims 7 to 10, wherein the rotor assembly (230) is an axial
compressor assembly.
12. The method of any of claims 7 to 10, wherein the rotor assembly (260) is an axial
turbine assembly.
13. The method of any of claims 7 to 10, wherein the rotor assembly (260) is a centrifugal
compressor assembly.
14. The method of any of claims 7 to 10, wherein the rotor assembly (260) is a radial
turbine assembly.