Interstage seal assembly for a turbine

A seal assembly positioned between forward and aft rotor stages of a turbine separates a forward cavity and an aft cavity on each side thereof. The seal assembly includes a seal body connected to a platform portion of a stator nozzle located between the forward and aft rotor stages, wherein an inner seal cavity in flow communication with an insert in the stator nozzle is formed therebetween. A static seal member is arranged in sealing relationship with a rotary toothed member between the forward and aft cavities, with the static seal member being connected to the seal body. An intermediate member is positioned between the seal body and the static seal member to define a plenum therebetween in flow communication with the inner seal cavity, the intermediate member having a plurality of angled passages formed therein in which a first end thereof is in flow communication with the plenum and a second end thereof is in flow communication with the aft cavity. Accordingly, air entering the stator nozzle insert flows into the inner seal cavity and then into the plenum so that it exits through the angled passages into the aft cavity at an angle with respect to an axis of rotation of the aft rotor stage. In this way, such air will acquire a tangential velocity component in the direction of rotation of the aft rotor stage.

BACKGROUND OF THE INVENTION 
1. Field of the Invention 
The present invention relates generally to a gas turbine engine and, in 
particular, to the cooling of the forward and aft cavities in the 
interstage region of the turbine. 
2. Description of Related Art 
It is well known that the turbine portion of a gas turbine engine is 
provided to extract energy from a hot gas stream as it impinges on the 
turbine blades thereof. The turbine then functions to cause a rotary 
action in an associated rotary apparatus. The turbine blades are in the 
form of airfoils and, due to the environment caused by the hot gas stream, 
are manufactured from materials capable of withstanding extreme 
temperatures. On the other hand, the mountings of such turbine blades are 
designed to withstand the high mechanical loads and stresses imposed 
thereon. Accordingly, it is important for the mounting or shank portions 
of the turbine blades to be protected from the direct impact of high 
temperatures stemming from the hot gas stream. This is principally 
accomplished by providing the blade and vane elements of the turbine with 
platforms which axially combine to define a boundary for the mounting 
shank portions, thereby isolating this temperature sensitive area from the 
hot gas stream. 
The need to protect other areas from exposure to high temperatures is 
equally important throughout the rotor cavity. It becomes even more 
pronounced, however, in the interstage region of the high pressure portion 
of the turbine since the boundary of expanding hot gases is adjacent to 
the forward and aft rotor cavities bounded by the disk post for the 
forward rotor stage, the platform for the aft stationary nozzle assembly, 
and the disk post of the aft rotor stage. According to present practice, 
labyrinth-type seals are used between the forward and aft cavities. Such 
seals are well known in the art and include a plurality of circumferential 
teeth which are contiguous with a circumferential sealing surface made 
from a high temperature resistant abradable material or other deformable 
materials to form the sealing surface with which the labyrinth teeth 
coact. Due to the deformability of the honeycomb material generally used, 
the sealing surface becomes deformed without injury to the teeth and 
thereby establishes a minimum clearance required by the operating 
conditions. 
When such a labyrinth seal is installed in the high pressure interstage 
region of the high pressure or HP turbine between the forward and aft 
cavities, cooling air passes through the aft nozzle and purges the forward 
cavity behind the forward rotor disk. This air then leaks through the 
labyrinth seal to purge the aft cavity in front of the disk post for the 
aft rotor. With such an arrangement, disk creep of the aft rotor has been 
experienced due to a temperature rise in the cooling air as it passes 
through the labyrinth seal and, during some operating conditions, inflow 
of the hot gas stream into the aft cavity because of insufficient purge 
flow. In order to remedy the above-noted deficiency, axial slots have been 
incorporated into the honeycomb portion of the interstage seat to supply 
additional cooling air directly to the aft cavity so as to reduce the net 
aft cavity air temperature. It was found, however, that the air stream 
leaving the axial slots required significant energy input to be 
accelerated to the rotor speed. This had the effect of increasing the air 
temperature relative to the aft rotor stage. 
In order to increase the efficiency of the system, a plurality of angled 
slots were formed in the honeycomb material so that air flowing 
therethrough would be directed into the aft cavity with an angular 
velocity imparted thereto. This particular design is shown and disclosed 
in U.S. Pat. No. 5,215,435 to Alan L. Webb et al., which is also owned by 
the assignee of the present invention. This Webb et al. configuration is 
an improvement over the previous design since less energy is required to 
accelerate the air from the slot to rotor speed than if the slot is 
directed axially and the air temperature relative to the rotor is less 
than that which an axial slot could accomplish. Nevertheless, it still has 
been found that certain disadvantages are inherent to this design. These 
disadvantages stem from the preswirl slots being incorporated into the 
open cells of the honeycomb material, as temperature of the air is 
increased and significant loss in total pressure are experienced. 
Consequently, the angular momentum of the preswirled air entering the aft 
cavity is reduced. 
In light of the foregoing, it would be desirable for an interstage seal 
assembly to be developed through which preswirled cooling air can be 
provided directly to an aft cavity of a turbine without the disadvantages 
of prior art designs. 
SUMMARY OF THE INVENTION 
In accordance with the present invention, a seal assembly positioned 
between forward and aft rotor stages of a turbine is disclosed as 
separating a forward cavity and an aft cavity on each side thereof. The 
seal assembly includes a seal body connected to a platform portion of a 
stator nozzle located between the forward and aft rotor stages, wherein an 
inner seal cavity in flow communication with an insert in the stator 
nozzle is formed therebetween. A static seal member is arranged in sealing 
relationship with a rotary toothed member between the forward and aft 
cavities, with the static seal member being connected to the seal body. An 
intermediate member is positioned between the seal body and the static 
seal member to define a plenum therebetween in flow communication with the 
inner seal cavity, the intermediate member having a plurality of angled 
passages formed therein in which a first end thereof is in flow 
communication with the plenum and a second end thereof is in flow 
communication with the aft cavity. Accordingly, air entering the stator 
nozzle insert flows into the inner seal cavity and then into the plenum so 
that it exits through the angled passages into the aft cavity at an angle 
with respect to an axis of rotation of the aft rotor stage. In this way, 
such air will acquire a tangential velocity component in the direction of 
rotation of the aft rotor stage.

DETAILED DESCRIPTION OF THE INVENTION 
Referring now to the drawings in detail, wherein identical numerals 
indicate the same elements throughout the figures, FIG. 1 depicts a high 
pressure section of a turbine 10 and the flow of cooling air through an 
interstage seal assembly. It is seen that turbine 10 is disposed in a 
casing immediately downstream of a combustion chamber (not shown) which 
emits hot combustion gases into a hot gas passage 12 of turbine 10. In 
this way, the hot combustion gases are brought into impinging contact with 
the a forward (or stage one) section and an aft (or stage two) section of 
turbine 10, where only a forward (stage one) turbine rotor 14, an aft 
(stage two) stator vane or nozzle 16, and an aft (stage two) turbine rotor 
18 are shown. The engine operates in the conventional manner wherein a 
fuel is burned in the combustion chamber and the products of combustion 
are guided by the first stage stator vanes or nozzles (not shown) through 
forward turbine rotor 14 and by a second set of stator vanes or nozzles 16 
through aft turbine rotor 18. The stator vanes are utilized to direct the 
air into an optimized angle of attack for energy transfer to rotor blades 
20 and 22 of forward and aft rotors 14 and 18, respectively. Energy thus 
transferred is utilized to drive a shaft (not shown) coupled to the rotor 
wheels and by which a compressor and/or a fan upstream of the combustor 
and various accessories of the engine are operated. 
Each blade 20 of forward turbine rotor 14 includes a shank and dovetail 
region for attachment to a forward disk 24 by means of a like number of 
disk posts 26. It will be seen that individual platforms 28 are disposed 
between adjacent blade shanks 30 of blades 20. Similarly, each stator vane 
or nozzle 16 includes a vane platform 32. A cooperating blade shank and 
dovetail region 34 of each aft rotor blade 22 likewise attaches to a 
corresponding disk post 36 of aft disk 38 and includes an individual 
platform 40. It will be understood that platforms 28, 32, and 40 of 
forward rotor blades 20, aft stator nozzles 16, and aft rotor blades 22, 
respectively, form an inner radial boundary, and, in conjunction with 
forward turbine shroud 42, stator vane outer platforms 44, and aft turbine 
shroud 46, define hot gas passage 12 through turbine 10. 
The rotating structure of turbine 10, including forward and aft turbine 
rotor blade disks 24 and 38, their respective disk posts 26 and 36, and 
their respective blade shank mountings and blade shaft members (not 
shown), lie within a rotor cavity 48 which is disposed in the radial 
interior of hot gas passage 12. The requirements for stress tolerance to 
the mechanical forces imposed upon disk posts 26 and 36, as well as other 
rotating elements disposed within the rotor cavity 48, prevent the 
utilization of materials having extreme thermal resistance. Hence, it 
becomes necessary to substantially isolate rotor cavity 48 from the 
temperatures of hot gas passage 12 and to provide for special cooling 
measures to avoid a decrease of the performance efficiency. 
The present invention is directed to solve the above-noted cooling problems 
associated with the interstage section of the high pressure part of 
turbine 10 and, more particularly, of a forward cavity 50 formed behind 
forward disk post 26 and by one side of a labyrinth seal 52 and of an aft 
cavity 54 formed in front of aft disk post 36 and by the other side of 
labyrinth seal 52. It will be noted that forward and aft cavities 50 and 
54 are bounded on their radially inward side by a thermal shield 55. 
The general structure of labyrinth seal 52 is well known in the art and 
known to include a honeycomb outer sealing member 56 (see FIG. 2) arranged 
annularly within a support member 53 in a single strip or otherwise. 
Honeycomb member 56 cooperates with a central member 58 having a plurality 
of teeth thereon, central toothed member 58 being arranged in linear or 
stepped fashion and mounted so as to be contiguous with abradable 
honeycomb member 56 to provide the sealing function. Either honeycomb 
member 56 or central toothed member 58 can be the static or rotating 
member depending on the particular application, but both sealing members 
are preferably made from high temperature resistant special metals or 
alloys. 
In the particular application disclosed herein, honeycomb member 56 is the 
static member and is attached to a seal body 60 connected to aft stator 
nozzle platform 32 by means of forward and aft hooks 57 and 59, 
respectively. While a previous design involves providing angled slots or 
passages in honeycomb member 56, it has been found that air flowing 
through such slots is subjected to temperature pick up from the open cells 
of the honeycomb material and significant loss in total pressure due to 
high friction losses as the air passes through the irregular surface of 
the slots. Accordingly, the angular momentum of the air flowing into aft 
cavity 54 is reduced. 
In order to overcome the disadvantages associated with flowing cooling air 
through slots in honeycomb member 56, the present invention presents an 
interstage seal assembly for turbine 10 in which a separate preswirl 
plenum 62 is provided within an intermediate member 64 positioned between 
seal body 60 and support member 53 of honeycomb member 56. It will be 
understood from FIG. 2 that air flowing into a nozzle insert 66 within a 
hollow portion of stator nozzle 16 is in flow communication with an inner 
seal cavity 68 between platform 32 of stator nozzle 16 and seal body 60 by 
means of at least one dump hole 70 provided in the nozzle insert 66 and at 
least one dump hole 85 provided in the nozzle 16. Thereafter, the air 
enters preswirl plenum 62 since it is in flow communication with inner 
seal cavity 68 by means of at least one or a plurality of dump holes 72 in 
seal body 60 and at least one or a plurality of corresponding openings 73 
in intermediate member 64 aligned therewith. 
As best seen in FIG. 3, intermediate member 64 has a plurality of angled 
passages 74 formed therein in which a first end 76 thereof is in flow 
communication with preswirl plenum 62 and a second or exit end 78 is in 
flow communication with aft cavity 54. It will be appreciated that the air 
entering aft cavity 54 from preswirl plenum 62 will be at a predetermined 
angle (preferably 10.degree.-30.degree.) with respect to an axis of 
rotation 80 of aft rotor 18 (see FIG. 1) so that it will acquire a 
tangential velocity component in the direction of rotation of aft rotor 
18. In this way, less energy is required to accelerate the air from angled 
passages 74 to rotor speed than if such passages would be directed 
axially, which in turn improves the overall efficiency and reliability of 
the engine. Additionally, the air temperature relative to aft rotor 18 is 
reduced below that which an axial slot could accomplish due to the reduced 
energy required to accelerate the cooling flow to rotor speed. Such 
reduced air temperature results in a cooler aft disk post 36 and, 
consequently, an increase in the creep life capability of disk post 36. 
Because there is less heat pick up by the air through smooth-walled 
intermediate member 64 forming preswirl plenum 62 and the air is not in 
direct communication with the hotter air residing inbetween the teeth of 
central toothed member 58 of labyrinth seal 52, the air entering aft 
cavity 54 is even cooler than that provided by angled slots formed in 
honeycomb member 56 as in the previously mentioned design. 
It will be understood from FIGS. 2 and 3 that intermediate member 64 is 
substantially U-shaped in cross-section so that an open end thereof is 
located adjacent seal body 60. Accordingly, a closed end of intermediate 
member 64 is located adjacent support member 53 of honeycomb member 56 so 
as to isolate air entering and exiting preswirl plenum 62 therefrom. 
It will also be understood that angled passages 74 have a length which 
enables the air to pass therethrough in a substantially uniform flow. 
Generally, this involves a correlation between both the length and the 
diameter of angled passages 74, where the length is at least twice the 
hydraulic diameter. 
It will be understood that air entering aft cavity 54 will naturally be 
pushed radially outward due to the centrifugal forces imposed thereon by 
rotation of aft rotor 18. Accordingly, it is preferred that seal body 60 
include a rear leg member 82 which extends into the path of air exiting 
angled passages 74. In this way, the air flows into rear leg member 82 and 
is directed radially inward so as to cool thermal shield 55 and aft disk 
post 36. Seal body 60 preferably includes spline seals 84 or the like in 
the ends thereof (see FIG. 1) to prevent air from leaking from inner seal 
cavity 68 inbetween seal body 60 and honeycomb member 56. 
Having shown and described the preferred embodiment of the present 
invention, further adaptations of the interstage seal assembly, and 
particularly the preswirl plenum formed between seal body 60 and honeycomb 
member 56, can be accomplished by appropriate modifications by one of 
ordinary skill in the art without departing from the scope of the 
invention.