Compressor diaphragm assembly

A compressor diaphragm assembly for combustion turbines includes a plurality of vane airfoils, each of which is formed with an integral inner shroud and an integral outer shroud, joined together by connecting bars which transfer loads between the vane airfoils and the casing slots of the turbine within which the vane airfoils are suspended.

CROSS REFERENCE TO RELATED APPLICATIONS 
This application is related to a similarly-entitled application, Ser. No. 
226,705, filed Aug. 1, 1988, now U.S. Pat. No. 4,889,470 by the inventor 
herein, assigned to the assignee of the present invention, and 
incorporated herein by reference. 
BACKGROUND OF THE INVENTION 
1. Field of the Invention 
This invention relates generally to combustion or gas turbines, and more 
particularly to the compressor diaphragm assemblies that are typically 
used in such turbines. 
2. Statement of the Prior Art 
Over two-thirds of large, industrial combustion turbines (which are also 
sometimes referred to as "gas turbines") are in electric-generating use. 
Since they are well suited for automation and remote control, combustion 
turbines are primarily used by electric utility companies for peak-load 
duty. Where additional capacity is needed quickly, where refined fuel is 
available at low cost, or where the turbine exhaust energy can be 
utilized, however, combustion turbines are also used for base-load 
electric generation. 
In the electric-generating environment, a typical combustion turbine is 
comprised generally of four basic portions: (1) an inlet portion; (2) a 
compressor portion; (3) a combustor portion; and (4) an exhaust portion. 
Air entering the combustion turbine at its inlet portion is compressed 
adiabatically in the compressor portion, and is mixed with a fuel and 
heated at a constant pressure in the combustor portion, thereafter being 
discharged through the exhaust portion with a resulting adiabatic 
expansion of the gases completing the basic combustion turbine cycle which 
is generally referred to as the Brayton, or Joule, cycle. 
As is well known, the net output of a conventional combustion turbine is 
the difference between the power it produces and the power absorbed by the 
compressor portion. Typically, about two-thirds of combustion turbine 
power is used to drive its compressor portion. Overall performance of the 
combustion turbine is, thus, very sensitive to the efficiency of its 
compressor portion. In order to ensure that a highly efficient, high 
pressure ratio is maintained, most compressor portions are of an axial 
flow configuration having a rotor with a plurality of rotating blades, 
axially disposed along a shaft, interspersed with a plurality of 
inner-shrouded stationary vanes providing a diaphragm assembly with 
stepped labyrinth interstage seals. 
A significant problem of fatigue cracking in the airfoil portion of 
inner-shrouded vanes exists, however, due to conventionally used methods 
of manufacturing such vanes. For example, in either of the rolled or 
forged methods used by the manufacturers of most compressor diaphragm 
assemblies, a welding process is used to join the vane airfoils to their 
respective inner and outer shrouds, such process resulting in a 
"heat-affected zone" at each weld joint. Crack initiation due to fatigue, 
it has been found, more often than not occurs at such heat-affected zones. 
Therefore, it would be desirable not only to provide an improved 
compressor diaphragm assembly that would be resistant to fatigue cracking, 
but also to provide a method of fabricating such assemblies that would 
minimize processes which produce heat-affected zones. 
The problems associated with fatigue cracking are not, however, resolved 
merely by eliminating those manufacturing processes that produce 
heat-affected zones. That is, it is well known that certain 
forged-manufactured vane airfoils, even after having been subjected to 
careful stress relief which reduces the effects of their heat-affected 
zones, can experience a fatigue cracking problem. It is, therefore, 
readily apparent that not only static, but also dynamic stimuli within the 
combustion turbine contribute to the problem of fatigue cracking. 
Forces that act upon the inner shroud and seal of a compressor diaphragm 
assembly are due, primarily, to seal pressure drop. Those forces, as well 
as aerodynamic forces acting normally and tangentially upon, and 
distributed over the surfaces of the vane airfoil, each contribute to the 
generation of other forces and moments that are transferred to the outer 
shroud, and subsequently to the casing of the combustion turbine via the 
weld joints which attach the vane airfoil to the outer shroud. 
It would appear that the simple alternative of using vane airfoils with 
integral outer and inner shrouds would quickly solve both causes of 
fatigue cracking. That is, the problem of heat-affected zones would appear 
to be eliminated entirely while the problems associated with instabilities 
due to static and dynamic stimuli within the combustion turbine would 
appear to be minimized. Such is not the case, however. 
For example, under the influence of the static forces and moments described 
above, the outer shroud segment of this hypothetical vane airfoil would 
not be stably engaged with the casing of the combustion turbine until such 
time that a restraining moment could be generated by contact of the 
extremities of the outer shroud segment with the walls of the slot formed 
in the casing to receive the segment. The outer shroud segment would, 
thus, rotate within the clearance gap (provided in the casing slot to 
account for thermal expansion). As a result, use of the hypothetical vane 
airfoil in a combustion turbine would lead to a great deal of stress in 
the vicinity of the outer shroud segment and excessive translational and 
rotational displacements, each of which would be further exacerbated under 
dynamic stimuli. It would also be desirable, therefore, to provide an 
improved compressor diaphragm assembly that would avoid the above 
described instabilities of engagement. 
SUMMARY OF THE INVENTION 
Accordingly, it is a general object of the present invention to provide an 
improved combustion turbine. More specifically, it is an object of the 
present invention to provide not only an improved compressor diaphragm 
assembly for use in such combustion turbines, but also an improved method 
of fabricating such compressor diaphragm assemblies. 
It is another object of the present invention is to provide a compressor 
diaphragm assembly that minimizes problems of fatigue cracking. 
It is still another object of the present invention is to provide a method 
of fabricating a compressor diaphragm assembly that substantially 
eliminates production of heat-affected zones. 
It is a further object of the present invention to provide a compressor 
diaphragm assembly that minimizes its instabilities of engagement with the 
casing of a combustion turbine due to both static and dynamic stimuli 
which may be experienced within the operational combustion turbine. 
It is yet a further object of the present invention to provide a compressor 
diaphragm assembly that is readily and inexpensively manufactured by 
existing technology. 
Briefly, the achievement of these and other objects is accomplished in a 
combustion turbine which has compressor diaphragm assemblies that include 
a plurality of vane airfoils joined together by load transfer means as 
taught herein. Each of the airfoils includes an integral inner shroud and 
an integral outer shroud, both of which have a groove that is adapted to 
receive a connecting bar. The grooves of adjacent such inner and outer 
shrouds, together with their respective connecting bars, constitute the 
load transfer means. A seal carrier with a pair of disc-engaging seals is 
suspended from the inner shroud. 
The above and other objects, advantages, and novel features according to 
the present invention will become more apparent from the following 
detailed description of a preferred embodiment thereof, considered in 
conjunction with the accompanying drawings wherein:

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
Referring now to the drawings, wherein like characters designate like or 
corresponding parts throughout each of the several views, there is shown 
in FIG. 1 the layout of a typical electric-generating plant 10 utilizing a 
well known combustion turbine 12 (such as the model W-501D single shaft, 
heavy duty combustion turbine that is manufactured by the Combustion 
Turbine Systems Division of Westinghouse Electric Corporation). As is 
conventional, the plant 10 includes a generator 14 driven by the turbine 
12, a starter package 16, an electrical package 18 having a glycol cooler 
20, a mechanical package 22 having an oil cooler 24, and an air cooler 26, 
each of which support the operating turbine 12. Conventional means 28 for 
silencing flow noise associated with the operating turbine 12 are provided 
for at the inlet duct and at the exhaust stack of the plant 10, while 
conventional terminal means 30 are provided at the generator 14 for 
conducting the generated electricity therefrom. 
As is shown in greater detail in FIG. 2, the turbine 12 is comprised 
generally of an inlet portion 32, a compressor portion 34, a combustor 
portion 36, and an exhaust portion 38. Air entering the turbine 12 at its 
inlet portion 32 is compressed adiabatically in the compressor portion 34, 
and is mixed with a fuel and heated at a constant pressure in the 
combustor portion 36. The heated fuel/air gases are thereafter discharged 
from the combustor portion 36 through the exhaust portion 38 with a 
resulting adiabatic expansion of the gases completing the basic combustion 
turbine cycle. Such thermodynamic cycle is alternatively referred to as 
the Brayton, or Joule, cycle. 
In order to ensure that a desirably highly efficient, high pressure ratio 
is maintained in the turbine 12, the compressor portion 34, like most 
compressor portions of conventional combustion turbines, is of an axial 
flow configuration having a rotor 40. The rotor 40 includes a plurality of 
rotating blades 42, axially disposed along a shaft 44, and a plurality of 
discs 46. Each adjacent pair of the plurality of rotating blades 42 is 
interspersed by one of a plurality of shrouded stationary vanes 48, 
mounted to the turbine casing 50 as explained in greater detail herein 
below with reference to FIGS. 3 and 4, thereby providing a diaphragm 
assembly in conjunction with the discs 46 with stepped labyrinth 
interstage seals 52. 
Due to conventionally used methods of manufacturing shrouded vanes 48, 
there exists a significant problem of fatigue cracking. For example (and 
referring now more specifically to FIGS. 3 and 4), in either of the 
methods that have been used by the manufacturers of most compressor 
diaphragm assemblies, a welding process is used to join an airfoil portion 
54 of the shrouded vane 48 to its respective inner shroud 56 and outer 
shroud 58. Such processes, as is well known, result in a heat-affected 
zone 60 at each weld joint 62. 
As defined by the Metals Handbook (9th ed.), Volume 6: "Welding, Brazing, 
and Soldering", American Society for Metals, Metals Park, Ohio, a 
"heat-affected zone" is that portion of the base metal which has not been 
melted, but whose mechanical properties or microstructure have been 
altered by the heat of welding, brazing, soldering, or cutting. In 
stainless steels alloys of the type utilized for the airfoils 54, inner 
shrouds 56 and outer shrouds 58, crack initiation due to fatigue more 
often than not occurs at such heat-affected zones 60. 
As noted above, however, problems associated with fatigue cracking are not 
resolved merely by eliminating those manufacturing processes that produce 
the heat-affected zones 60. For example, FIG. 3 illustrates an 
inner-shrouded vane 48 that is manufactured by the rolled constant section 
approach, while FIG. 4 illustrates an inner-shrouded vane 48 that is 
manufactured by the forged variable thickness-to-chord ratio approach. 
Forces that typically act upon the inner shroud 56 and its seal 52 of 
conventional compressor diaphragm assemblies such as those shown in FIGS. 
3 and 4 are primarily due to seal pressure drop F.sub.S. Those forces, as 
well as aerodynamic forces acting normally F.sub.A and tangentially 
F.sub.T upon airfoil portion 54, each contribute to the generation of 
other forces and moments that are transferred to the outer shroud 58, and 
subsequently to the casing 50 of the combustion turbine 12 via the weld 
joints 62 which attach the vane airfoil 54 to the outer shroud 58. 
Fatigue cracking, nevertheless, would still not be eliminated simply 
through the use of a hypothetical airfoil having an integrally formed 
inner and outer shroud, thereby doing away with the heat-affected zones 
60. Under the influence of the static forces and moments described above, 
the outer shroud segment of this hypothetical vane airfoil would not be 
stably engaged with the casing of the combustion turbine until such time 
that a restraining moment could be generated by contact of the extremities 
of the outer shroud segment with the walls of the slot formed in the 
casing to receive the segment. The outer shroud 58 would, thus, rotate 
within the clearance gap (provided in the casing slot to account for 
thermal expansion). As a result, use of the hypothetical vane airfoil in a 
combustion turbine would lead to a great deal of stress in the vicinity of 
the outer shroud segment and excessive translational and rotational 
displacements, each of which would be further exacerbated under dynamic 
stimuli. 
It has been found that one approach, as described in Ser. No. 226,705, 
filed Aug. 1, 1988, now U.S. Pat. No. 4,889,470, will substantially 
eliminate most fatigue cracking problems. Another approach that is 
somewhat more simple in its construction, however, is described herein. 
As shown in FIGS. 5-8, the compressor diaphragm assembly 64 according to 
the present invention includes a plurality of vane airfoils 66, each such 
airfoil 66 having an integrally-formed inner shroud 68 and an 
integrally-formed outer shroud 70. The inner shroud 68 and outer shroud 70 
of each of the airfoils 66 includes a groove 72 that is adapted to receive 
a connecting bar 74 to form load transfer means 76. Two or more adjacent 
ones of the plurality of airfoils 66 are coupled together by the load 
transfer means 76 and, thus, form the assembly 64. 
A seal carrier 78 comprising a plurality of segments 80, is suspended from 
the inner shroud 68, each such seal carrier segment 80 including at least 
one pair of disc-engaging seals 82, and being formed to engage the inner 
shrouds 68 of one or more vane airfoils 66. 
In accordance with one important aspect of the present invention, 
heat-affected zones are eliminated not only due to the plurality of vane 
airfoils' 66 being formed with integral inner shrouds 68 and integral 
outer shrouds 70, but also due to their being joined together by processes 
which use little or no heat at the critical airfoil to shroud junction. 
Furthermore, there are few if any instabilities of engagement between the 
vane airfoils 66 and the casing slot 75 (due either to static or dynamic 
stimuli) because of the load transfer means 76. 
The respective integrally-formed outer shrouds 70 are joined to form an 
outer ring 84 with the connecting bars 74. In such a manner, each 
integrally-formed outer shroud 70 is also formed with a generally T-shaped 
cross-section for engagement with the slot 75 formed in the casing 50 of 
the turbine 12, held in place by conventional retaining screws 90. 
In order to facilitate assembly and disassembly of the compressor diaphragm 
according to the present invention, and to minimize the cost of producing 
such an assembly, spacers 92 of varying sizes are provided to properly 
space the vane airfoils 66 one from the other. Referring now more 
specifically to FIGS. 6 and 7, however, it can be seen that the 
integrally-formed inner shrouds 68 and outer shrouds 70 are respectively 
joined to adjacent ones of such integrally-formed inner shrouds 68 and 
outer shrouds 70 in order to prevent excessive translational and 
rotational displacements of the resulting compressor diaphragm assemblies 
64 within the casing slots 75 of the turbine 12. 
Each vane airfoil 66 is connected to an adjacent vane airfoil 66, both at 
the integrally-formed inner shrouds 68 and at the integrally-formed outer 
shrouds 70, by the load transfer means 76 comprising the connecting bars 
74. The slots 72 which are provided in the integrally-formed inner shrouds 
68 and at the integrally-formed outer shrouds 70 may have substantially 
parallel sides as shown in FIG. 6 for use with rectangular-shaped 
connecting bars 74. As an alternative configuration, however, the slots 72 
may be tapered at an angle .theta. less than 90 degrees as shown in FIG. 
7. 
With such alternative configurations of forming the slots 72 of the 
integrally-formed inner shrouds 68 and the integrally-formed outer shrouds 
70, compressor diaphragm assemblies 64 in accordance with the present 
invention may be easily formed by joining a plurality of vane airfoils 66 
together, either by brazing, by electron beam welding, by laser welding 
(directions "A" or "B" shown in FIG. 6), byshrink fitting or simply by 
providing blade-type clearances (i.e., approximately 0.001 inches). 
The sides of the connecting bars 74 are defined by the angle .theta. which 
can vary from zero (i.e., for parallel-sided slots 72), suitable for 
joining by electron beam welding in the directions A and B as shown in 
FIG. 6, to a taper of less than 90 degrees, suitable for shrinking or 
fitted assembly. For example, with the tapered slot 72 as shown in FIG. 7, 
the connecting bars 74 could be "shrunk" using liquid nitrogen or other 
suitable means and inserted within the slot 72 for expansion thereafter in 
the slot 72. On the other hand, the vane airfoils 64 could be heated to 
approximately 500.degree. F., and the connecting bars 74 inserted therein, 
to provide a locked up system with low compressive and tensile stresses. 
Furthermore, blade type clearances could be provided between the sides of 
the tapered slots 72 and the connecting bars 74, with such connecting bars 
74 being joined to the slots 72 by a plurality of pins 96 fitted along its 
length. 
As explained herein above, the compressor diaphragm assembly 64 according 
to the present invention, thus, eliminates problems of fatigue cracking 
caused by heat-affected zones. This also substantially reduces stress 
concentrations that typically build up at the inner and outer shrouds. 
Integrally formed vane airfoils minimize costs associated with manufacture 
of such airfoils, while maximizing the quality of their production since 
long-established procedures that have been utilized for rotor blade 
manufacture (e.g., castings, forgings, contour millings, etc.) can be 
applied. As is readily evident, replacement of a single damaged vane 
airfoil 66 is easily accomplished, and the multiplicity of interfaces 
between the vane airfoils 66, segmented seal carrier 80, outer shrouds 70, 
and slot 75 provide for increased mechanical damping which will minimize 
dynamic response. 
Obviously, many modifications and variations are possible in light of the 
foregoing. It is, therefore, to be understood that Within the scope of the 
appended claims, the invention may be practiced otherwise than as 
specifically described herein.