Turbine engine interstage seal

A turbine engine includes axially adjacent compressor and turbine rotors which define a groove therebetween. Into the groove between the rotors radially extends an annular interstage seal assembly. The seal assembly includes a thermally insulative cellular wall structure which resists thermal and pressure included distortions. At its radially inner extent the cellular wall defines a seal structure in cooperation with the rotors, and also carries heat shield members on an axial fore disposed toward the turbine rotor.

BACKGROUND OF THE INVENTION 
The present invention is in the field of sealing apparatus and method for 
use in a turbine engine. More particularly the present invention relates 
to a seal structure employed in a turbine engine wherein the engine 
includes a centrifugal compressor receiving ambient air and pressurizing 
this ambient air for delivery to a combustor; and a turbine disposed in 
back-to-back relationship with the compressor and receiving combustion 
products formed by combustion of a fuel in the pressurized air supplied by 
the compressor. Historically such turbine engines have employed a 
disk-like centrifugal compressor rotor and a similar disk-like radial 
inflow turbine disposed in back-to-back relationships such that they 
cooperatively define therebetween a radially inwardly extending annular 
groove. At the radially inner extent of the annular groove the turbine and 
compressor rotors cooperatively carry a seal runner, or centering ring, 
which served as a concentricity maintaining structure with respect to the 
turbine and compressor rotors respectively. 
A housing structure of the turbine engine conventionally carries a 
disk-like annular seal member extending radially inwardly into the annular 
groove between the compressor rotor and turbine rotor, and at its radially 
inner extent carrying a seal structure. The seal structure may be perhaps, 
of a labyrinth type, sealingly engaging the seal runner, which is carried 
cooperatively by the compressor rotor and turbine rotor. Generally 
speaking, these conventional seal structures were made of sheet metal 
stampings of comparatively light gage and have been subject to a variety 
of shortcomings because of their structural nature. 
By way of example, the annular sealing structure is subjected to an axial 
force because of the fluid pressure differential thereacross existing 
between the discharge of the centrifugal compressor and the inlet pressure 
at the radial inflow turbine. This pressure differential causes the 
relatively flexible conventional sheet metal seal structure to be 
displaced axially in a direction towards the turbine rotor. Additionally 
the seal structure is subjected to a radial temperature differential 
resulting from the flow of hot gasses radially inwardly and across the 
turbine rotor. The hot gas flow provides a heating input to the annular 
seal structure which is most pronounced at its radially outer extent and 
decreases progressively radially inwardly therefrom. On the other hand the 
seal structure is subjected to a comparative cooling effect on the 
compressor side thereof as a result of the small portion of compressor 
discharge airflow which circulates in the backspace between the compressor 
rotor and the adjacent face of the annular seal structure. Because of this 
differential radial heating and cooling on opposite faces of the seal 
structure, conventional seal structures have displayed a warping or 
buckling similar to that experienced with the bottom of a frying pan which 
is heated most intensely in the center and is cooler at its outer 
periphery. 
Such seal structures, due to this differential thermal expansion 
experienced in the radial direction of the structure, will in one location 
warp in one axial direction and in an adjacent location may warp or 
displace in the opposite axial direction. In the case of the conventional 
annular seal structure such warping and displacements superimposed upon 
the axial displacements which resulted from differential pressures across 
the seal structure results in axial as well as radial movements at the 
center opening of the seal structure whereat sealing integrity is to be 
maintained. 
Further, the heat input to the seal structure may be circumferentially 
nonuniform. Thus, the seal structure will have a nonuniform 
circumferential and radial temperature distribution. This 
circumferentially nonuniform temperature distribution further contributes 
to warping of the seal structure. As a result, the labyrinth or other 
conventional seal structure which is carried at the center opening of the 
interstage seal is displaced axially as well as radially and sealing 
integrity is generally not successfully maintained by conventional 
interstage seals of the type described above. 
A result of the axial, radial, and warping displacements experienced by 
conventional seal structures has been the necessity to provide larger than 
desired clearances between the turbine back face and the seal structure. 
Consequently, a portion of the combustion products flowing radially 
inwardly onto the turbine are lost into the excessive clearance provided 
at the seal structure. Turbine efficiency is adversely affected by this 
loss of a portion of the combustion products. Also, the heat input to the 
seal structure is increased as energy input to and efficiency of the 
turbine is decreased by excessive seal assembly clearances. 
Accordingly, it is an object for the present invention to provide a seal 
assembly of the type described which resists axial, radial, and warping 
displacements; and which also allows decreased turbine back face 
clearances. 
The present invention realizes the objects thereof by providing a turbine 
engine having an improved seal structure, said turbine engine comprising; 
a compressor section inducting ambient air and delivering this air 
pressurized to a combustor, means introducing fuel to said pressurized air 
in said combustor to maintain combustion providing a flow of 
high-temperature pressurized combustion products, and a turbine section 
expanding said combustion products toward ambient to extract shaft power 
therefrom, said compressor section including a centrifugal compressor 
rotor, and said turbine section including a radial in-flow turbine rotor, 
said compressor rotor and said turbine rotor being disposed in 
back-to-back relation so that air flowing to said compressor rotor and 
combustion products flowing from said turbine rotor flow in a single axial 
direction and said rotors define an annular groove extending radially 
inwardly therebetween, said improved seal structure comprising a cellular 
annular impermeable wall member carried by said turbine engine and 
extending radially into said groove to there sealingly cooperate with said 
rotors, said cellular wall member having a circumferentially arrayed 
plurality of axially extending cells therein.

FIG. 1 provides a schematic presentation of a turbine engine 10 embodying 
the present invention. In order to gain a generalized overview of the 
operation of the turbine engine 10, it is necessary to note that the 
engine 10 includes a dynamic compressor section 12 which in operation 
ingests ambient air, as indicated by arrow 14, and delivers this air 
pressurized to a combustor section 16, as is indicated by arrow 18. A flow 
of fuel is added to the pressurized air in combustor section 16, as 
indicated by arrow 20. Combustion maintained within the combustor section 
16 provides a flow of high temperature pressurized combustion products. 
These combustion products flow to a turbine section 22 (as is indicated by 
arrow 24) wherein they are expanded toward ambient pressure and discharged 
(as is indicated by arrow 26) to produce shaft power. The turbine section 
22 drives a shaft 28 which in turn drives the dynamic compressor section 
12. The shaft 28 includes an externally extending portion 28' whereby the 
turbine engine 10 may deliver shaft power to an external power-consuming 
device (not shown). 
Between the compressor section 12 and turbine section 22, a seal apparatus 
30 is disposed about the shaft 28 and cooperates therewith to 
substantially prevent flow of pressurized air from compressor section 12 
to turbine section 22 along the shaft 28. That is, because of the seal 
apparatus 30, substantially all of the air pressurized by compressor 
section 12 flows to turbine section 22 via the combustor 16. 
In viewing FIG. 2, it will be seen that the turbine engine 10 includes a 
housing generally referenced with the numeral 32. The housing 32 defines 
an inlet 34 opening to ambient air via an inlet screen 36, and through 
which a flow of ambient air (arrow 14) is received. The housing 32 
journals a rotor member generally referenced with the numeral 33. Rotor 33 
includes a compressor rotor portion 35, a turbine rotor portion 37, and an 
elongate tie bolt 39. Cooperatively, the portions 33, 35, tie bolt 39, and 
an externally extending shaft portion 28' substantially complete the 
rotational assembly 38 of the engine 10, recalling the schematic depiction 
of FIG. 1. The rotational assembly is journaled in housing 32 by a pair of 
angular contact bearings 41 which sustain both radial and thrust loads. At 
its left end, viewing FIG. 2, the shaft portion 28' includes a splined 
drive coupling portion 43 whereby shaft power is transferred from engine 
10 to, for example, a power distribution gear box (not shown). 
The compressor rotor member 35 is rotatably received in the housing 32 and 
receives thereon the inlet flow of ambient air 14. Rotor member 35 
delivers the received air pressurized (arrow 18) via a diffuser structure 
40 into a pressurized air plenum 42. The plenum 42 surrounds combustor 16, 
which includes outer perforate walls allowing air flow inwardly thereof 
from the plenum 42. 
Carried by the housing 32 and extending inwardly through the plenum 42 and 
into combustor 16 is a plurality of fuel delivery nozzles 44, only two of 
which is visible viewing FIG. 2. The nozzles 44 deliver a spray of fuel 
(arrow 20) into the combustor 16 so that combustion is maintained therein. 
Combustion products flow from combustor 16 to turbine section 22 (arrow 
24) via a nozzle member 46. The turbine section 22 includes the radial 
in-flow turbine rotor 37 which, as noted previously, is disposed in 
back-to-back relation with the centrifugal compressor rotor 35. The 
turbine rotor 37 rotatably extracts shaft power from the combustion 
products 24, which are then exhausted from the engine 10 via an exhaust 
duct opening 50 (arrow 26). 
Viewing FIGS. 3-5, it is seen that the compressor rotor 35 and turbine 
rotor 37 are disposed in back-to-back relation. In other words, fluid flow 
approaching compressor rotor 35, and fluid flow departing turbine rotor 37 
moves in the same axial direction. The rotors 35 and 37 define an inwardly 
extending radial groove 52 therebetween, which groove leads to and 
terminates at a piloting ring 54 received between the rotors 35 and 37. 
The piloting ring 54 assists in maintaining mutual concentricity of the 
rotor members 35 and 37, and also defines a radially outwardly disposed 
outer surface 56. 
The housing 12 defines a radially inwardly disposed annular groove 58 
axially between the diffuser 40 and the turbine nozzle 46. The groove 58 
is axially coextensive with the groove 52 so that the two grooves 52 and 
58 are radially congruent. 
Carried in the groove 58 and extending radially inwardly into groove 52 is 
an annular seal assembly generally referenced with the numeral 30. 
Recalling that the source of pressurized air in the engine 10 is 
compressor section 12, it is apparent that seal assembly 30 is exposed on 
its left face to pressurized air from compressor 12, and is exposed on its 
right face to lower pressure, but higher temperature, combustion products 
from combustor 16. Consequently, the seal assembly 30 at an axially 
disposed radially outer marginal annular face 62 thereof is pressure 
biased axially into sealing engagement with the housing 32 at an axially 
disposed annular shoulder 60 of groove 58. 
In order to resiliently bias the seal assembly 30 toward the shoulder 60, 
the assembly 30 includes an annular spring member 64. The spring member 64 
is disposed on the cool side of seal assembly 30 toward compressor 12, and 
is attached at its radially inner margin to the assembly 30 by a plurality 
of rivets 66. The spring member 64 may be circumferentially continuous 
like a Belleville spring, or may include, as depicted, a plurality of 
resilient finger portions 68, only one of which is visible on FIG. 3. The 
spring member 64 extends radially outwardly to engage a shoulder 70 of 
housing groove 58, which shoulder 70 is disposed axially opposite to 
shoulder 60. 
Importantly, the seal assembly 30 includes an annular wall member, 
generally referenced with the numeral 72. At its radially outer margin, 
the wall member 72 includes an axially thickened rim portion 74 which is 
received closely but axially movably in the groove 58. Extending radially 
inwardly from the rim portion 74 the wall 72 includes an axially thick, in 
comparison to conventional seals, wall portion 76, which wall portion is 
substantially cellular, as will be seen. The wall portion 76 extends 
inwardly to piloting ring 54 whereat an axially spaced apart pair of face 
sheets 78 of the wall 72 each define one of a pair of axially aligned 
openings (80, 82). At the respective openings 80, 82, each of the face 
sheets 78 define a respective radially disposed surface 84 which is in 
close but radially spaced running relation with the outer surface 56 of 
piloting ring 54. Consequently, the wall 72 cooperates with the piloting 
ring to inhibit fluid flow from compressor 12 to turbine 22 along the 
shaft 28. The two axially spaced surfaces 84 of the openings 80, 82 
function somewhat like a knife-edge or labyrinth seal to inhibit fluid 
leakage between the wall 72 and piloting ring 54. 
Because of the cellular nature of the wall 72 intermediate of the two face 
sheets, the wall portion 76 includes a ring member 86. This ring 86 is 
generally U-shaped in cross section, and circumscribes the piloting ring 
54 radially spaced outwardly thereof. The ring 86 is sealingly secured to 
each of the face sheets 78 so that fluid pressure leakage into the cells 
of wall 72, as well as liquid permeation into these cells, is prevented. 
Attached to the wall 72 by the same rivets which secure the spring member 
64 thereto, but on the opposite side of the wall from the spring 64, is an 
annular bracket member 88. The bracket member 88 cooperates with the wall 
72 to define a radially outwardly disposed annular recess 90. Received 
movably in the recess 90, and in the groove 58 of housing 32 adjacent the 
shoulder 60, is an annular turbine back face shroud member 92. The back 
face shroud member 92 defines the annular face 62 which is urged axially 
into sealing contact with the shoulder 60 by spring 64 and by differential 
air pressure when the engine 10 is operating. The bracket 88 allows 
relatively free radial movement of the turbine back face shroud member 92 
relative to wall 72, but restricts axial relative movement to the extent 
of a relatively small clearance (not shown) therebetween. As will be 
explained, the back face shroud member 92 is also urged axially into 
engagement with the bracket member 88. 
Also received into the recess 90 and in groove 58 along with the turbine 
back face shroud member 92 is an annular radiation heat shield member 94. 
The members 92 and 94 are axially spaced apart over most of their radial 
dimension. Like the turbine back face shroud 92 the radiant shield 92 is 
relatively movable in the recess 90. However, the radiant shield member 92 
is itself cone shaped like a Belleville spring. At its radially outer 
margin the shield member 94 engages the rim portion 74. At its radially 
inner margin, the turbine back face shroud member 94 engages the inner 
extent of the back face shroud member 92 to urge the latter into 
engagement with the bracket member 88. 
Turning now to FIGS. 6-8, steps in the process of making the seal member 30 
are depicted. FIG. 6 depicts an annulus 96 of honeycomb material 98. The 
annulus 96 is formed from a larger body of honeycomb material and has an 
outer diameter slightly smaller than that of seal 30, an inner diameter 
slightly larger than ring 86, and a plurality of circumferentially spaced 
openings 100 therethrough spaced outwardly of the central opening 102 
thereof. In addition, the honeycomb material 98 itself includes a 
plurality of fine-dimension walls, generally referenced with the numeral 
104 which extend substantially perpendicularly to the plane of the annulus 
96. The fine-dimension walls 104 cooperatively define a regular repetitive 
geometric array of cells 106 extending through the annulus 96. 
Viewing FIG. 7, it is seen that the ring member 86 is received into the 
central opening 102 of the honeycomb wall 96. At each of the openings 100 
the honeycomb wall 96 receives one of a matching plurality of annular 
reinforcing grommet members 108. The annulus 96 is itself received into a 
ring member 110, which at its outer perimeter defines the rim 74. After 
the annulus 96, ring 86, grommets 108, and outer ring member 110 are 
united, for example, by brazing, then the face sheets 78 are united with 
the previously described components. This operation of uniting the face 
sheets 78 with the previously brazed components of wall 72 may be carried 
out in a brazing furnace by use of a brazing foil interposed between the 
sheets 78 and honeycomb annulus 96. Consequently each of the fine 
dimension walls 106 is structurally coupled to the face sheets 78 at its 
juncture therewith. This coupling of the fine dimension walls 106 with the 
face sheets 78 results in an extremely stiff structure for the wall 72. 
Finally, FIG. 8 depicts the assembly of spring member 64 with wall 72 and 
bracket 88 by means of rivets 66. The bracket member 88 captures the back 
face shroud member 92 and radiant heat shield member 94 so that these 
members are retained movably in cooperation with the wall member 72. The 
reinforcing grommets 108 allow the rivets 66 to be headed without crushing 
the fine-dimension wall structure 104 of the honeycomb material. 
Returning briefly to a consideration of how the seal 30 functions in engine 
10 during operation of the latter, it will be recalled that an air 
pressure differential urges the seal 30 rightwardly. This air pressure 
differential tends to bow the wall portion 72 so that the center part 
thereof moves axially toward the turbine rotor 48. However, the wall 72 
composed of the two face sheets 78 and fine-dimension honeycomb material 
98 extending therebetween is very stiff so that virtually no bowing of the 
wall 72 occurs. This high degree of stiffness of the wall 72 also results 
in the wall member 72 resisting buckling in response to circumferentially 
nonuniform temperatures which might be imposed upon the seal 30. However, 
the turbine back face shroud 92 is relatively massive so that it also 
assists in leveling any circumferentially nonuniform temperatures. 
Additionally, the back face shroud member 92 is free to move radially 
relative to the wall member 72 so that temperature cycling and 
circumferential nonuniformities which cause physical excursions of the 
back face shroud member 92 do not result in stresses being transmitted to 
wall member 72. Finally, the radiant heat shield member 94 additionally 
assists in isolating the cellular wall member 72 from the severe 
high-temperature conditions in the adjacent turbine section 22. The result 
of all of the above is that the radially inner periphery of seal assembly 
does not experience the physical dislocations and excursions in physical 
position with changing operations of the engine 10, as do conventional 
interstage seals. Consequently, the engine 10 may be designed with a 
closer clearance between the turbine rotor 37 and the back face shroud 
member 92. An improved efficiency and power output result for the engine 
10.