Note : Les descriptions sont présentées dans la langue officielle dans laquelle elles ont été soumises.
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SEAL'COOLING METHOD AND APPARATUS
Field of the Invention
This invention pertains to seal cooling structures
and methods in gas turbine engines and, more particul-
arly, to seal cooling structures and methods for
labyrinth-type seals.
De'script'ion o`f''the'Pr'ior Axt
A big factor in the performance of a gas turbine
engine is the effectiveness of seal structures that
must operate over a wide range of operational conditions.
In particular, a compressor discharge pressure seal
must operate effectively at high pressures and
usually at relatively high operating temperatures.
The compressor discharge pressure seal is employed to
prevent compressed air from leaking between a rotating
co~pres`sor section of an engine and a nonrotating
combustor section. To optimize engine performance,
tight clearance on this seal is highly desirable to
minimize air leakage. Any air that leaks through
the seal does not pass through the combustion cycle
of the engine and, therefore, does not contribute to
the power produced by the products of combustion.
7'~
While preventing excessive air leakageJ the
compressor discharge pressure seal must also permit relative
rotation between upper and lower sections of the seal and must
operate in a region of the engine near the compressor discharge
outlet where temperatures reach over 1100 F (593 C).
Current practice is to employ labyrinth-type seals in
this region. A labyrinth-type seal comprises one or more
circumferential teeth that are contiguous with a circumferential
sealing surface wherein the teeth and sealing surface are
relatively rotatable. Labyrinth seals can provide a high
restriction to gas flow, and while there is some leakage,
labyrinth seals do permit free rotation between upper and lower
sections of the seal. This type of seal has many other well
known advantages and is widely used at various sealing locations
in gas turbine engines.
The effectiveness of labyrirlth seals is a function of
the clearance between the sealing teeth and the contiguous sealing
surface. While engine parts can be accurately machined to obtain
minimum clearances and a highly effective seal, practical
operation of the engine results in seal clearance degradation due
to differential thermal growth betweerl the sealing teeth ~nd the
sealing surface. This is recognized, arid to some extent alle~ated,
by widespread use of honeycomb material or other abradahle,
readily deformable materials to form the sealing surface with
which the labyrinth teeth coact. By this approach, if the seaLing
teeth grow at a faster rate than the sealing surface, the sealing
surface will be deformed without injury to the sealing teeth. This
will automatically establish the minirnum clearance available when
the sealing surface is at its maximum growth position, and the
sealing teeth are at their minimum growth position.
To minimi~e thermal growth, methods and structures
have been developed for cooling labyrinth seals by directing cooling
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air along the outer surface of the sealing structure,
as shown in U.S. Pat.No.3,527,053 to Horn dated
September 8, 1970 assigned to the same assignee as the
present invention. Other systems have been developed
which directly inject air into the space between the
teeth of the labyrinth seal, as shown in U.S. Pat.No.
3,989,~10 to Ferrari dated Nov. 2,1976 also assigned
to the same assignee as the presen~ invention. The
use of cooling, air, as disclosed in these systems, greatly
improves the efficiency of sealing structures by
decreasing thermal expansion, thereby making it possible
to maintain tighter seal clearance to cut down on seal
leakage.
However, prior art systems involving compressor
discharse pressure seals do not fully utilize the cooling
air available in the most effective possible manner. In
addition, problems have resulted because of friction-
induced work done on the cooling air by the rotating
portlon of the sealing structure in the region where the
cooling air is dîrected against rapidly rotating portions
of the seal. This increases the cooling air temperature
and, therefore, the seal temperature. Also, in prior
art systems, cooling air has been derived from boundary
layer air at the base of the last blade of the compressor,
and this boundary layer air can be as much as 100F
(55.5 C) warmer than nonboundary layer air. Finally,
no matter where the cooling air comes from, the flow
pattern must circulate through the entire seal and around
support structures surrounding the seal. Any regions within
the seal or surrounding the seal with zero airflow will
undergo insufficient cooling, increasing the risk of
material failure due to unnecessarily high temperatures.
It is, thereforer an object of the present invention
to derive the cooling air for the compressor discharge
pressure seal from a sufficiently pressurized source with
- the lowest practical ~ir temperature.
It is another object of the present invention to inject
the cooling air into the seal in a more efficient manner which
decreases the amount of work done on the cooling air by the
internal seal structure and, therefore, which results in cooler
5 air temperature .
It is another object of the present invention to
distribute the cooling air in and around the compressor discharge
pressure seal with a flow pattern that does not create "dead
spots" with zero airflow in the seal or in the surrounding support
10 strUcture s .
These and other objects will be more fully understood
from the drawings and from the following description and example,
all of which are intended to be representative of rather than in
any way limiting on the scope of the present invention.
Briefly, in the method and apparatus of the present
in~ention, cooling air for a compressor discharge pressure seal
is derived from compressed air discharged from a compressor
dump diffuser. Air from this source is the coolest available
source of air that is suficiently compressed for use in this region.
20 This air is cooler than boundary layer air extracted frvm the base
OI a compressor blade, as is done in previous compressor seal
cooling systems.
In a specific form, the cooling air is injected through
passages in an outer ring of She seal into a space between first and
25 second teeth of the seal, taken from the upstream direction. These
passages are angled to direct the cooling air with a tangential
velocity component in the direction of rotation of the teeth and
downstream portion of the seal structure. The tangential velocity
component reduces the amount of frictional drag work done by the
30 teeth, seal rotor, and associated seal structure on the cooling air
injected into the seal. This ultimately reduces the increase in
cooling air temperature caused by friction and allol,vs the cooling
air to extract more heat energy from the seal structure.
The cooling airflow path is also provided with a forward
purge outlet into an open cavity in the support structures located
forward of the seal. This forward purge outlet is provided to avoid
zero airflow conditions from occurring in this forward cavity and
S prevent possihle overheating of the seal components and turbine
rotor structure.
DESCRIPTION OF THE DE~AWINGS
, ~
While the specification concludes with claims distinctly
claiming and particularly pointing out the invention described
herein, it is believed that the invention will be more clearly
understood by reference to the discussion below in conjunction
with the foLlowing drawings:
FIGURE 1 is a vertical cross-sectional view of a prior
art seal cooling structure;
FIGURE 2 is a vertical cross-sectional view of seal
cooling structure o~' the present invention and the surrounding
components of a gas turbine engine; and
FIGURE 3 is a cross-sectional view of the present
invention taken along line 3-3 of Figure 2.
DESC~IPTION OF THE PREFERRED E~IBODIMENT
Referring now to Figure 1, a compressor discharge
pressure seal 10 is shown in its u~ual location within a typical
gas turbine engine. This seal 10 is generally located between
compressor 11 and a combustor 16 in serial flow relation. In
the gas turbine engine, a compressor section cornpresses engi~e
intake air and a compressor discharge pressure seal retains
this compressed air in the thrust-producing flowpath of the engine
while permitting relative rotation along this flowpath of compressor
parts in relation to the nonrotating combustor 16.
In Figure 1, an aft compressor blade 12 of the compressor
section 11 is shown forward of the compressor discharge seal 10.
Intake air is compressed by compressor blades rotating about a
central axis oE rotation of the turbine engine and then directed
through an out~et guide vane 14 and a compressor dump diffuser
lS to diffuse the compressed air and direct the air into the
combustor 16. In the combustor section of the engine, the
5 compressed air is combined with fuel and ignited to form a
thrust-producing propulsive gas flowstream. For the purpose
of simplifying the description of this invention, a complete gas
turbine engine is not shown. It is believed that the reader will
fully appreciate this invention without a description of an entire
10 engine. If the reader desires an e~fplanation of the operations
within a gas turbine engine that affect ~ compressor discharge
pressure seal, the reader is referred to United Stat~s
~atent Numl~er 3, 527, 053 to Horn dated September 8,
1~70 .
The compressor discharge pressure seal 10 is
provided to prevent compressed air Erom escaping into central
regions 19 of the gas turbine engine while, at the same time,
permitting rotation of a compressor rotor 18 in respect to the
outlet guide vane 14 and combustor 16, which do not rotate.
20 Compressor blades 12, one of which is shown in Figure 1, are
attached to the compressor rotor I8, and the rotor rotates the
compressor blades to compress intake air passing through the
compressor section 11 of the engine. The outlet guide vane 14
does not rotate and rerno~es a component of rotational velocity
25 of the compressed air before it enters the combustor. The
compressor dump diffuser 15 diffuses the air, causing a decrease
in flow velocity and an increase in pressure.
The compressor discharge seal 10 is compris2d of a
series of circumferential labyrinth teeth 20 contiguous with a
30 seal outer stator 22 that defines a sealing surface. Outer edges
24 of the teeth 20 are initially assembled so as to form a very
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close fit against the stator 22. IJpon rotation of the compressor
rotor 18 and the attached labyrinth teeth 20 about the engine axis,
the outer edges 24 of the teeth create a slight groove in the inner
surface of the seal stator 22. The very close fit between the
5 teeth 20 and the seal stator 22 inside these grooves provides a
high degree of restriction to gas flow between the rotation teeth 20
and the stationary seal outer stator 22.
An object of this invention is to minimize differential
thermal growth between the interacting portions of th~s labyrinth-
10 type seal and thereby maintain a closer fit between the teeth 20and the seal stator 22 to improve seal effectiveness under operating
conditions .
In prior art systems, such thermal growth has been
decreased by passing compressor discharge air between the seal
15 teeth and seal stator to maintain the seal components at lower,
more consistent temperatures. In the prlor art system shown in
Figure 1, boundary layer air from the compressor discharge at
the base of the aft compressor blade 12 is directed radially inward
and axially aft along the compressor rotor 18 to the region of the
20 compressor discharge pressure seal 10. Some of the compressed
air then leaks through the seal along the path depicted by a wavy
arrow, and caltinues to Ilow a~t into a central region 19 of the gas
turbine engine.
The method and apparatus of the present invention is
25 shown in Figures 2 and 3. In Figure 2, the flowpath of compressed
air used to cool the seal structure 10 is shown with multiple arrows.
This air first is extracted downstream of the compressor dump
diffuser 15. In one embodiment of this invention, air in this region
is approximately 100 F (55. 55 C) cooler than the compressor
30 boundary layer air used in prior art systems, such as the system
shown in Figure 1. The cooling air is directed through an inlet 30
to an open region 32 radially surrounding the seal stator 22.
--8--
From this open region 32, the air is directed through
passages 34 in ;q seal bracket 23 and across an open slot 21 in the
seal stator 22 into the space between the first and second labyrinth
teeth 20 of the compressor discharge seal. The first and second
5 teeth are furthest upstream in respect to airflow through the
turbine. The passages 34 are uniquely oriented on an angle in
respect to a radius fromthe engine axis to impart a tangential
component of velocity in the direction of rotor rotation. The
direction of annular orientation is shown in Figure 3 wherein it
10 can be readily appreciated that the passages 34 cause the cooling
air to be injected into the seal in the direction of rotor rotation.
The labyrinth teeth 20 are attached to the rotor 18 for rotation
therewith. Thus, the labyrinth teeth 20 rotate during en~ne
operation, while the seal stator 22 does not. By orienting the
15 passages 34 on an angle, the cooling air is injected in the direction
of rotation of the teeth 20, thereby decreasing the frictional drag
between the injected air and the teeth. The tangential component
of velocity provided by this form of the present invention reduces
the work done by the frictional drag on the cooling air and,
20 therefore, decreases the resulting increase in the temperature
of the cooling air, Ultimately, the internal seal structure is
maintained at a lower temperature.
The cooling air tends to flow through passages 34
because the region downstream of the compressor dump diffuser
25 15 is at a higher static pressure than the central regions 1~ of
the gas turbine beyond the compressor discharge seal 10. The
cooling air tends to leak in the aft direction across the region
between the teeth outer edges 24 and the seal stator 22. A
small but continuous flow of leakage air is sufficient to cool
30 the seal components and maintain the internal seal structure at
a relatively low temperature. This allows the seal to maintain
a closer fit between the outer edges 24 of the seal teeth and seal
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g
stator 22 because of diminished thermal expansion and diminished
differential therrnal growth.
Another unique feature of this invention is the manner
in which a cavity 35 forward of the compressor discharge seal 10
is purged with air to avoid overheating of the rotor structure 18
as a result of zero throughflow. As can be seen in Figure 2, an
annular series of inlet holes 36, one of which is shown, is provided
to inject a small quantity of air into this forward cavity 35, Air
will flow in this direction because the static pressure downstream
of the diffuser 15 is higher than at the exit of the compressor,
upstream of the guide vane 14. These holes 36 are cut at an an~le
to impart a tangential velocity component in the direction of rotor
rotation, similar to the manner in which a tangential velocity
COInpOnent is imparted by passages 34. Tan~ential injection
reduces the amount of frictional drag between the injected air and
the rotating compressor rotor 18. This reduces the amount of
work done on the injected air, the resulting increase in temperature
of the air and, consequently, the rotor 18, and minimizes the
quantity of air required for cavity purge~
This forward cavity purge apparatus offers anothe~
advantage over the previous seal cooling system, shown in ~igure 1,
wherein a flow of air is created through the forl,vard cavity 35
because o~ leakage flow through the compressor discharge pressure
seal. In the prior art system, when thermal expansion of the
labyrinth teeth 20 of the seal is such as to cause a lesser clearance
between the outer edges 24 of the teeth and the seal stator 22, the
leakage flow of air is substantially reduced. Ca~ity 35 flow-through
can approach zero, and overheating of the rotor 18 can result. In
the present invention, because inlet holes 36 are provided, the
amount of air injected into cavity 35 for cavity purge remains
relatively constant and is not affected by clearance change in the
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compressor discharge pressure seal. Therefore, if the seal
clearance diminishes causing a temporary drGp in seal leakage
flow, the forward cavity 35 remains pur~ed, and the affected
portions of the rotor 18 do not overheat. The combined effects
5 of the cooling airflow through passages 34 and inlet holes 36
serve to maintain both the compressor rotor 18 an~ the compressor
discharge seal 10 at reasonable temperatures, thereby improving
the performance of the high pressure turbine seal and rotor cooling
circuit.
While specific embodiments of the present invention
have been described, it will be apparent to those skilled in the
art that various modifications thereto can be made without
departing from the scope of the invention as recited in the appended
claims. For example, while the invention has been described in
15 conjunction with a labyrinth-type compressor discharge pressure
seal in a gas turbine engine, it will be appreciated that various
aspects of this invention are applicable to other sealing regions in
a gas turbine engine and can be applied to sealing structures other
than labyrinth-type seals. The methods and apparatus of the
20 present invention can be used to increase the performance of
various seals in any type of turbomachinery. The scope of the
invention, therefore, is to be derived from the following claims.
