Note: Descriptions are shown in the official language in which they were submitted.
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SEAL ASSEMBLY
The present invention relates generally to gas
seals in gas-turbine engines, and, more particularly, to
remo~able annular seals for turbine nozzles.
BAC~ROU~D OF THE INVENTIO~
Gas turbine engines typically include a fan
section, a compressor section, a combustion section, and
a turbine section in serial flow relationship. Air
~ drawn in by the fan section is compressed or pressurized
in the compressor section before being heated in the
combustion section to produce a high velocity gas
stream. Energy is extracted from the gas stream in the
turbine section by being utilized to cause rotation of a
plurality of rotor stages. Each rotor stage comprises
an annular array of rotor bladss and receives combustion
gases from an upstream adjacent annular array of nozzle
vanes. The rotor blades are secured to a rotor disX
which rotates as the gas stream passes over the rotor
blades. The nozzle vanes are stationary and turn the
gas stream in a desired direction over the rotor blades.
The temperature of the gas stream as it exits the
combustion section may be as high as, for example, Z000
degrees Fahrenheit (or 1093 degrees .Centigrade). This
stream of hot gases may attain a velocity, for example,
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in excess of 2000 feet per second (or 610 meters per
second) as i~ passes through the nozzle vanes. The
nozzle vanes are suitably attached to an inner band or
platform and an outer band or shroud, such as, for
example, by welding or being integrally cast therewith ~o
form an annular array of vanes commonly referred to ~s a
turbine nozzle. Since the turbine nozzle directs the gas
stream in predetermined directions for efficient
operation of the engine, it is desirable to assure that
all of 'che gas stream passes through the vanes of the
nozzle by reducing leakage around the platform and shroud
ends of the nozzle.
It is known in the art to provide seals to prevent
gas leakage in gas turbine engines. U.S. Patent
3,4~3,070 issued Jan. 21, 1969 and assigned to the
assignee of the present invention describes one form of
seal, i.e., a honeycomb seal. One disadvantage with
prior art seal arrangements has been the difficulty of
replacing seals during engine servicing. For example,
seals have been brazed or welded to the nozzle bands. A
further ~ disadvantage arises from other forms of
attachment by either rlvets or pins which may be required
to permit radial shifting of the seal. Such radial
shifting is utilized to compensate for differential
thermal expansion and contraction in the turbine section.
When rivets and pins are used, the support structure for
the seal is relatiqely heavy and costly and~decreases the
fuel efficiency of the engine. Another disadvan~age of
the prior art seal arrangements has been the requirement
for special tooling for installing and handling pins
and/or rivets.
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S~MMARY 0~ THE INVE~TIO~
Accordingly, it is an ob~ect of the present
in~ention to provide a new and improved seal arrang2ment.
Another object of the present invention is to
provide a seal assembly for a gas turbine engine nozzle
which accommodates differential thermal expansion and
contraction.
Another object of the present invention is to
provide a seal assembly for a gas turbine engine which i5
easily removable and replaceable.
Another object of the present invention is to
provide a seal assembly for a gas turbine engine which is
lightweight, relatively low cost and ~emovable a~d
replaceable without special tooling.
In an illustrative embodiment, the present in-7ention
comprises a unitary annular seal structure for attachment
~o a turbine nozzle in a gas turbine engine. The nozzle
includes an annular platform disposed circumferentially
about a longitudinal axis of the enqine. An annular
array of vanes is secured to the platform. The seal
structure includes an abradable annular seal memher, a
seal backing member and a seal attachment ring. The
attachment ring includes an annular, radially extending,
axially acting spring member positioned to cooperate with
a plurality of radially extending tabs on the backing
member. In use, the seal structure is positioned within
a circular opening within the turbine nozzle. The nozzle
includes a radially depending appendage formed as part of
the nozzle platform. The spring member abuts one side of
the appendage and the tabs are positioned to abut another
side of the appendage for holding the annular spring
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member in gas sealing engagem~nt with the appendage to
thereby provide a seal against gas leakage and to axially
restrain the seal structure. The spring member and tabs
comprise a radially slideable joint for the seal
structure. In order to restrict circumferential motion
of the seal structure, slots are formed in the appendage
for receiving ~he tabs. Seal replacement is easily
achieved by bending the tabs and sliding the seal
structure axially out of the nozzle. Differential
thermal expansion and contraction are accommodated by the
radially slideable seal arrangement. A primary gas seal
is maintained by circumferential sliding contact between
the abradable seal member and an annular knife edge
attached to a turbine disk.
BRIEP DESC~IPTION OF TE~ D~AWI~GS
The novel features believed characteristic of the
invention are set forth in the appended claims. The
invention, in accordance with an exemplary preferred
embodiment, together with further objects and advantages
thereof is more particularly described in the following
detailed description taken in conjunction with the
accompanying drawings in which: -
FIG. 1 is a schematic side elevation view, in
partial section, of a gas-turbine engine with whicn the
present invention may be used;
FIG. 2 is an enlarged partial schematic view of a
turbine section of the engine of FIG. 1 incorporating one
smbodiment of the present invention;
FIG. 3 is a perspective view of a section of a
turbine nozzle and seal structure, separated for clarity,
illustrating one embodiment of the present invention;
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FIG. 4 is an enlarged partial cross-sectional view
of a nozzle platform illustrating one form of the present
inventive seal in sealing relationship with the plat'orm;
and
S FIG. 5 is a view of the platform and seal
arrangement of FIG. 4 in which the seal and platform have
been displaced due to differential radial growth
therebetween.
The exemplary features set out herein illustrate
preferred embodiments of the present invention, and are
not to be construed as limiting either the scope of the
invention or the scope of the disclosure thereof in any
manner.
~ETAIL~D DESC~IPTION
FIG. 1 is a partial cutaway view of a highly
simplified schematic illustration of an exemplary gas
turbine engine lO. The engine lO is a high-bypass turbo-
fan engine arranged substantially concentrically about alongitudinal or axial axis, depicted by the dashed line
12. The ensine lO includes a fan section 14, a
compressor section 16, a combustion section 18, a high
pressure turbine section 20, and a low pressure turbine
section 22 all being, but fo~ the present invention,
conventional. ~ir pulled in ~y the fan section 14 is
compressed in the compressor section 16 and then flows
into the combustion section 18 where it is mixed with
fuel and ignited to produce a high energy (high
temperature and high pressure) gas stream. The gas
stream flows across a plurality of blades in each rotor
stage of the high pressure turbine section 20 and the low
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pressure turbine section 22, causing rotation of the
rotor stages. The high pressure turbine section 20
rotates the compressor section 16 through a shar~ 24.
The low pressure ~urbine section 22 rotates the fan
section 14 and other components through a rotor shaft 25.
Although a turbo-fan engine is shown in FIG. 1, it is to
be understood that the invention hereinafte~ described
can be effectively employed on other types of gas turbine
engines.
Referring to ~IG. 2, there is shown a portion of the
low pressure turbine section 22 of the engine 10. The
turbine section 22 includes alternating annular arrays of
nozzle vanes 26 and rotor blades 28, the vanes 26 and
blades 28 being designed as airfoils for reacting the hot
gas stream. The nozzle vanes 26 are attached to a
radially ou~er band 32 and a radially inner platform 30
to form non-rotating annular nozzle stages 36 ~best seen
in FIG. 3). The nozzle stages are suitably attached to
and supported by an annular outer engine shroud 38. The
nozzle vanes 26 are typically designed for two purposes:
to lncrease the velocity of the heated gases flowing past
the~, and to direct the flow of gases to strike the rotor
blades 28 at a desired angle. Each of the rotor blades
28 is typically attached at its radially inner end to a
platform 34. Extendins from' the platform 34 is a
conventional dovetail (not shown) which connects the
blade 28 to a disk 40 which is attached~to the rotor
shaft 25. The asse~bled blades 28 and disk 40 form a
turbine wheel.
The high velocity gas or fluid exits one of the
nozzl~ stages 36 and strike an adjacent array of rotor
blades 28 in a turbine wheel to effect its rotation and
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drive the shaft 25. As stated above, any leaXage of
fluid/gas around a circumferential periphery of one of
the annular no~zle stages 36 contributes to inefficiency
of the engine. Such inefficiency may arise from gas
leakage around the platform 30 causing a reduction of net
velocity of gas flow and increased turbul~nce. One major
cause of gas leakage is differential thermal expansion
between platform 30 and its adjacent seal or mating
surface due to high temperature gases reacting on the
no~zle vanes 26. Differential tempe~atures during engine
operation results in expansion and contraction between
the nozzle stages 36 and any adjacent mating surfaces.
Thus, it is desirable to provide a sealing structure
which accommodates movement from differential expansion
and contraction yet also effectively seals against
leakage during such movement.
Reference is now made to FIGS~ 3 through 5 in
conjunction with FIG. 2 which illustrate a sealing
; structure 42 in accordance with one embodiment of tne
present invention for reducing gas leakage around a
nozzle stage. The structure 42, within the dashed lines
in FIG. 2, is shown in greater detail in FIGS. 4-5. The
hot gases passing through the turbine section 22 heat the
various components, such as the airfoils 26 and 28, the
2~ bands 32, the platforms 30 and 34 and the disks 40 and
cause them to thermally expand. Since the components may
be made of diverse materials, may have diverse material
thicknesses, and may be subject to diverse rates of
heating, each component may expand differently.
Accordingly, the sealing structure 42 is adapted to
accommodate differential expansion in a radial direction
~nd similar contrsction as cosponents cool.
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The sealing structure seal assembly 42 comprises a
seal assembly 44 which includes a co~nentiGnal abradable
annular seal member 46 having a raZially outer
circumferential surface fixedly attached ~o a first
circumferential, radially inner surface of an annular
seal backing ring 48 by brazing, welding or other
suitable means well known in the art. The member 46 may
be a ~honeycomb type seal well known in the art. An
attachment ring 50 is suitably attached, for example by
brazing or welding, to a second circum~erential, radially
outer surface of the backing ring 4B, i.e., the surface
opposite the surface to which the seal 46 is attached.
In accordance with one embodiment of the invention the
attachment ring 50 is a generally annular U-shaped ring
having axially spaced first and second annular legs 52
and ~6, respectively, extending radially outwaraly of the
ring 48. The backing ring 48 can be seen to have an
essentially L-shaped structure with a first up-turr.ed
edge 54 against which the second leg 56 of attachment
; 20 ring 50 is positioned. A remaining or second edge 55 of
ring 48 includes a plurality of circumierentially spaced
tabs 58, which are up-turned in the installed position
~best seen in FIG. 3).
Considering specifically FIG. 3, a portion of an
annular nozzle stage 36 is illustrated perspectively and
partially exploded for clarity. This figure shows some
of the structural aspects of the outer band 32 which
provides physical support and enables attachment to the
shroud 38 (see ~IG. 2) of turbine sec~ion 22. In
accordance with one embodiment of the invention, a
radially inner surface of platform 30 (the
circumferential periphery of nozzle stage 36) is formed
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with a generally L-shaped radially inwardly e~tending
annular support member 60. The member 60 includes a
first leg 66 suitably attached to (by brazing, welding
or integral casting) and e~tending radially inwardly
fro~ a radially inner surface of the platform 30, A
second a~ially e~tending leg 62 of the ~-shaped support
member 60 e~tends generally parallel to the longitudinal
a~is 12 from the first leg 66 and has a plurality of
circumferentially spaced slots 64 formed in an a~ially
facing edge surface for receiving the tabs S8 radially
eztending from backing ring 4a.
A significant element of the present invention is
the first leg 52 of attachment ring 50 illustrated in
FIG. 4 in accordance with one embodiment. Although the
first leg 52 may be a straight member, it preferably has
a radially outwardly e~tending straight base portion 52A
and a radially outwardly e~tending straight distal end
portion 52B being integral therewith through an inclined
intermediate portion 52C. The first leg 52 is sized for
and formed of a suitable high temperature, spring metal
material, for e~ample, commercially available Hast-X, to
provide elastic resiliency in the a2ial direction for
enabling spring action. During assembly, an axially
facing surface of the first leg 52 is pressed against
and abuts a lateral (a~ially facing) surface 68 of
support member 60 to effect an initial aYial preload to
effect an annular fluid seal thereagainst. In the
preferred embodiment only distal end 52B provides a seal
against the surface 68 wi~h portions 52A and 52C
providing an offset.
The tabs 58 on backing ring 48 are bent during
assembly from an initial axial position to a radial
position to fit into the slots 64 and provide a retaining
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force against the spring action of first leg 52. It
will be noted that the spring leg 52 is formed with the
reverse curve portion 52B to establish a limited sliding
contact surface and also to enhance the spring action of
leg 52. The tabs 58 cooperate with the slots 64 to
constrain circumferential rotation of seal assembly 44
while permitting differential radial e~pansion and
contraction between the seal assembly 44 and
platform 30. In particular, the tabs 5B are free to
move radially in the slots 64, and the first lçg 52 is
free to slide radially in sealing contact with and
against the lateral surface 68 of the first leg 60.
The use of honeycomb seals is well known in the
aircraft engine art for providing a sealing relationship
between rotating and non-rotating elements~ However,
for purposes of clearly explaining the invention, as
shown in FIGS. 2 and 4, a radially inner circumferential
surface of the seal member 46 is closely positioned
adjacent to one or more ~knife~ or labyrinth edges 70 in
:~ 20 conventional sealing relationship therewith. Tne
edges 70 are carried by cantilevered arms 72 a~ially
e~tending from rotatable turbine disks 40. Since the
primary gas flow path is through nozzle vanes 26 and
rotor blades 28, the portions of the nozzle stages and
turbine wheels in the flow path are hotter tnan those
portions, including in particular the seal assembly 44,
out of the flow path. Consequently, there is
differential radial e~pansion and contraction between
and within the nozzle stages 36 and the turbine wheels.
If the sealing structure does not compensate for such
differential thermal movementj a gap may be created or
enlarged between edges 70 and an adjacent seal member ~6
which will allow gases to bypass the turbine nozzles and
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13LN-1 74 2
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reduce the efficiency of the engine. In general, the
seal between member 96 and edges 70 is referred to as a
primary seal while the seal between ring 50 and platform
30 is referred to as a secondary seal.
Referring to FIG. 5, there is shown a significant
advantage of the present invention in accommodating
differential radial movement within a gas turbine engine
adjacent the turbine nozzle stage 36. The seal assembly
44 comprising the seal member 46, backing ring 48 and
attachment ring 50 remains in sealing relationship with
the knife edge 70 while the nozzle stage 36 has expanded
radially. The sliding spring contact relationship
between attachment ring 50 and the depending leg 66 of
platform support member 50 allows the gas stream to be
lS confined to its normal flow path by virtue of the sealing
contact between spring leg portion 52B and the surface 68
of platform leg portion 66. The length of the tabs 58 is
sufficiently long to permit radial movement ~ithout
losing contact between tabs 58 and leg 62 of support
member 60.
It will be appreciated by those skilled in the art
that the present invention allows the seal assembly 44 to
freely move in a radial direction with respect to the
vanes 26 and the platform 30. Inasmuch as the vanes 26
are directly subjected to hot gases they will expand
radially outwardly more than the seal assembly 44 which
will be at a lower temperature. FIG. 5 illustrates a
position of the nozzle platform 30 due to thermal
expansion as contrasted with the reference position,
illustrated in FIG. 4. Inas~uch as the seal assembly 44
and the edge 70 and associated arm 72 are located away
from the primary hot gas flow, they expand/contract
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relatively little with respect to each other for
maintaining enhanced sealing at the edge 70.
Expansion or contraction within engine 10 in an
axial direction is also accommodated by the inventi~Je
S sealing structure. However, for axial movement, each
seal member 46 remains in its normal r~lative posi~ion
with respect to a corresponding turbine nozzle stage 36.
The member 46 is provided with sufficient width in the
axial direction to allow axial displacement of knife edge
79 without its losing contact with member 46. Thus, the
seal assembly 44 allows axial displacement between a
turbine nozzle stage 36 and the seal member 46 without
loss of seal effectiveness.
The seal backing member 48 is preferably formed of a
light weight sheet metal material. Prior to attachment
- to the turbine nozzle, the tabs 58 are parallel to a
central axis of the annular seal assembly 44 which may be
the centerline 12, for example. When the seal structure
44 is first axially positioned within a turbine nozzle
stage 36, the tabs 58 are be~t from an initial axial
position generally parallel to the surface of ring 48 to
a radially outwardly position as shown in FIGS. 3-5 and
serve to hold the seal assembly in place against axial
and circumferential movement within the turbine nozzle
stage 36. Whenever the engine 10 is disassembled for
servicing, the seal assembly 44 is easily detached from
the turbine nozzle stage by simply bending the tab~ 58
to their initial axially oriented position and axially
pulling the assembly 44 out of the nozzle stage.
Replacement requires only a simple positioning of a new
seal assembly 44 within the nozzle stage and bending of
tabs 58 for its restraint. In this manner, a seal can be
easily re~oved and replaced.
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While there have been described herein what are
considered to be preferred embodiments of the invention,
other modifications will occur to those s~illed in the
ar~ from the teachings herein, and it is, therefore,
desired to secure in the appended claims all s~ch
modifications as fall within the true spirit and scope of
the invention. ~or example, although the tabs 58 have
been disclosed as attached to ring 48 and the slots 64
are disposed in the member 60, an opposite configuration
is also possible~
