CA1190480A - Vane structure having improved cooled operation in stationary combustion turbines - Google Patents

Abstract

ABSTRACT OF THE DISCLOSURE
A combustion turbine is provided with first stage vane structure having an airfoil portion or vane and inner and outer shrouds. The vane is formed from a spar which is covered by a shell having inner coolant channels which carry coolant for cooling the shell (vane) outer surface. The coolant is supplied through coolant chan-nels in the spar, and a limited amount of the coolant is discharged to the hot gas path through the vane trailing edge while most of the coolant is directed to coolant channels within the shrouds for shroud coolant action and ultimate discharge to the hot gas path through shroud openings.

Description

1 49,36g VANE STRUCTURE HAVING IMPROVED COOLED OPERATION
IN STATIONAR~ COMBUSTION T~RBINES
CROSS-REFERENCE TO RELATED APPLICATIONS
None BACKGROUND OF THE INVENTION
The present invention relates to combustion turbines and more particularly to air cooled vane struc-tures for stationary combustion turbines used to generate electricity or operate other industrial processes.
With the continuation of development activity to produce higher temperature operation for gas turbines, there continues to be a need to develop turbine hot part structures which can over a long life withstand the ther-mal stresses produced by the higher operating tempera-tures. Since available alloys do not solve the problem at the operating temperatures involved, structural provisions must be made for positive cooling of the hot parts.
Typically, compressor discharge air is employed for hot part cooli.ng. It has been common practice to direct coolant to the airfoil portion OI the turbine vane structure and to discharge the spent coolant into the hot gas path through openings which are somewhat susceptible to contamination. In some cases, the v~ne shrouds have been separately cooled requiring additional coolant air which reduces turbine efficiency.
SUMMARY OF THE INVENTION
In a combustion turbine, an improved vane struc-ture is arranged to provide for reduced vane surface

2 49,368 temperature operation in a high temperature gas path such that higher turbine efficiency and reduced vane airfoil maintenance are achieved. The vane structure comprises an airfoil portion and inner and outer shrouds with coolant directed first through the airfoil portion and then tllrough the shrouds followed by discharge.
BRIEF DESCRIPTION OF THE DRAWINGS
Figure 1 shows a representation of a longitudinal turbine section showing the location of vane assemblies in the hot motive gas path;
Figures 2 and 3 show pre-assembly and post-assembly perspective views of a vane assembly structured in accordance with the invention;
Figure 4 shows an exploded elevation view of the vane assembly with a schematic coolant circuit indicated thereon;
Figure 5A shows a cross-section of a vane employed in the vane assembly;
Figure 5B shows an enlarged portion of the vane cross-section to illustrate shell coolant channels pro-vided therein;
Figures 6A-6F show top and bottom and various section views of the vane shown in Figure 4;
Figures 7A-7H show a longitudinal inner shroud section and various other inner shroud sections which illustrate various inner shroud coolant and structural features;
Figures 8A-8H show a top plan view of the outer shroud and other views which similarly illustrate outer shroud features;
Figure 9 shows an assembled elevation view of the vane assembly; and Figure 10 shows a graphical representation of vane temperature distribution.

3 49,368 DESCRIPTION OF THE PREFERRED EM~ODIMENT
More partlcularly, there is shown in Figure 1, a stationary combustion turbine 10 in which the invention is embodied. The turbine 10 includes a plurality of com-bustors 12 which are disposed within a casing 1~ about thelongitudinal turbine axis.
Hot combustion products flow through transition ducts 16 which are coupled to the respective combustors 12 to direct the combustion products through an annular space 15 in which the first stage stationary turbine airEoils or vanes 18 are located. In turn, the vanes 18 are struc-tured and positioned to direct the hot gas against first stage rotor blades 20 at an angle and velocity which provide for efficient turbine performance.
In this case, the turbine 10 is designed for relatively high gas temperature operation, i.e., 2100F
and above. As a result, positive cooling must be provided at least for all fîrst stage rotor blades 20 and first stage stator vanes 18 so that blade and vane metal tem-peratures can be held at levels at which part life is acceptable with the use of available high temperature alloys.
With positive cooling, in this instance by means of bypass compressor discharge air, metal tempera-ture in the first stage or subsequent stages can be restricted to the desired value such as 1500F by control of the various cooling system parameters. Thus, the heat removal capa-city of the coolant system can be designed to limit the metal temperature as desired by appropriate specifications for coolant air entry temperature, flow rate of coolant air, metal area across which heat is directly transferred to the coolant air, and various other design parameters.
The vanes 18 are arranged in sector vane assem-blies 22 (Figure 2) which are disposed about the annular space 15. The vane assemblies 22 are supported by a ring 23 formed by sector ring segments which in turn are se-cured to the turbine casing 1~.

4 4g,368 Each vane assembly 22 includes at least one airfoil portion or vane 18 which is supported between inner and outer shrouds 24 and 26. When the vane assem-ln blies 22 are all secured ~ an annular configuration, the shrouds 24 and 26 cooperate to limit the hot gas path and define the annular space 15. It is preferred in this embodiment that each vane assembly 22 be provided with one vane 18 to achieve greater expansion flexibility. Each transition duct spans a plurality of vane assemblies.
Generally, the vane assembly structure is adapt-ed to provide improved metal cooling and operation so that higher gas temperatures can be specified for the turbine lO without shortenin~ the llfe of the vanes (especially in the first stage) due to creep or corrosion of available suitable high temperature alloys.
Cooling of turbine hot parts is usually provided by compressor discharge air which is returned (sometimes after cooling it to a reduced temperature) to the turbine hot gas path after it has provided its coolant action so as to limit the cost imposed on turbine efficiency for the t cooiant operation.
The coolant air is obtained from a suitable supply line and delivered through the ring 23 to a plenum chamber formed by the ring; it then flows through an inlet to a coolant distribution circuit which is formed in the hot parts to produce the desired metal surface temperature control on basis of expected temperature and flow condi-tions for the hot gas that drives the turbine 10.
It is possible to supply coolant air to internal vane coolant cavities and distribute the coolant about the vane shell prior to discharge of the coolant to the hot gas stream through holes in the vane shell on the vane pressure and suction sides and the vane trailing edge.
However, a coolant circuit of this type carries the dis-advantage of potential deposition plugging of the shellcoolant holes which eventually can impair the cooling function. Coolant air can also be discharged through the bottom of the vane as in U.S. Patent 3,560,107.

49,368 In the present invention, the coolant circuit preferably provides for coolant flow 1) through the vane with a portion of the coolant being discharged through the vane trailing edge and 2) the largest remaining portion being directed (in a predetermined split ratio) through ~n ~er ~ ~te r the ~ and ~wer vane shrouds and ultimately and advan-tageously through the less contaminable shroud surface opanings for discharge to the hot gas path.
In this way, the vane airfoil structure is provided with reliable internal cooling. As a further advantage, cooling air, which is normally used only to cool the airfoil portion of the vane followed by discharge to the hot gas flow path, is reused to cool the inner and outer shrouds prior to discharge. This two-pass use of cooling air reduces the need for separate cooling air for the shrouds and results in a turbine efficiency improve-ment.
Consideration will now be given to the manner in which the coolant circuit is structured to provide im-proved operation of a vane assembly in accordance with theinvention. Generally, with reference to Figures 2-4, coolant air enters several internal vane cooling cavities 28, 30 and 32 through an inlet 34 to the upper shroud from a coolant supply line and passes along the inner surface of the vane shell. Part of the coolant is discharged through a trailing edge of the vane 18.
Preferably, the vane 18 is made from a spar 38, which is cast with the coolant cavities and which provides the basic structural strength of the vane 18, and a rela-tively thin shell 40 which is diffusion bonded to the spar38. Coolant channels 42 are etched in the chordwise direction on the inner surface of the shell 40 or the outer surface of the spar to receive coolant from the coolant supply cavities 28, 30 and 32 through spar holes 44 and to return coolant to coolant return cavities 46 and 48 through spar holes 50 or through discharge openings in the trailing edge 36. The spar 38 projects beyond the ~4~
6 49,368 shell 40 at both the outer and inner ends to form tenons 52 and 54, which are constant cross-section extensions of the vane alr~oil. Exit ports 56 and 58 are casted in the sides of the tenons 52 and 54 (Figures 6A, 6D, 6E, 6F) to direct the coolant into shroud coollng cavities.
The inner shroud 24 (Figure 7A) has a casted and partitioned internal cooling cavity 55 having support pins 57 distributed thereacross. Coolant air from the vane 18 is distributed uniformly throughout the cavity and exi.ts through a pattern of canted cooling holes 59 covering the gas path surEace of the inner shroud 24. An airfoil shaped recess 60 is also casted in the shroud 24 to re-ceive the spar tenon 54 in Figure 8~ at final assembly.
The outer shroud 26 is similar to the inner shroud 24 in structure and purpose and like reference characters are employed. In addition, support lips 64 and 66 are integrally cast on the outer shroud 26 for keyed support of the vane assembly 22 on the ring 23 in the engine.
The shrouds 24 and 26 are suitably joined to the vane 18 at final assembly, as by use of a brazing filler material.
A detailed heat transfer design of the coolant circuit in the channel cooled vane 18 is based on two principal considerations: (1) coolant channel geometry and coolant ~low rates needed for the desired m~-iml1~
metal temperature (such as 1400F) and (2) a two dimen-sional heat transfer analysis of the mean section of the vane for the selected channel geometry.
In reducing the invention to actual practice, an initial heat transfer analysis may be performed on the shell/spar. This analysis may include a simulation of heat transfer by conduction through the airfoil wall while excluding conduction in the spanwise and chordwise direc-tions. Coolant temperature rise and pressure drop calcu-lations are included in the analysis. Calculations are made for conditions consistent with the intended engine 7 ~9,368 operation, say at a gas temperature of 2170~F. The design for the shell/spar vane is based on limiting the maximum airoil surface metal temperature to a predetermined value such as 1400F.
The selection of coolant channel geometry may be based on the simplified one dimensional conduction analy-sis. When the external thermal environment is specified in this analysis, specification of the supply and dis-charge pressures and coolant passage geometry is suffi-cient to permit calculation of coolant flow rates and wall and fluid temperature distributions in the direction of coolant flow. Several coolant channel geometries may be examined to establish a candidate geometry that maintains the specified maximum surface temperature for a reasonable length and is consistent with chemical channel etching processes. Once a candidate channel geometry is estab-lished, the channel length may be varied around the air-foil to match the external environment while maintaining the maximum surface temperature.
A two-dimensional heat transfer analysis of the vane mean section may be made for the maximum surface temperature configuration. Such an analysis may take into account conduction through the airfoil wall and in a chordwise direction but it may neglect conduction in the radial (spanwise) direction. The governing differential equation solved is the steady-state, two-dimensional heat conduction equation:
a kaT + a kaT = O
ax ax ay ay where T is the wall temperature, k is the ~emperature dependent thermal conductivity, and x and y are the spa-tial coordinates. For certain geometries and boundary conditions, it is possible to obtain an exact solution to the equation in the form T = T(x, y) . Unfortunately, this is impossible for an airfoil; therefore, a numerical technique is used to obtain a solution.

8 ~ 49,368 Before the above e~uation can be solved for the temperature distribution in the airfoil wall, the boundary conditions are defined. These conditions include:
1. the local gas-to-airfoil heat transfer coefficient;
2. the local adiabatic wall temperature;
3. the local airfoil-to-coolant heat transfer coefficient; and 4. the local coolant temperature adjacent to the internal surface of the airfoil.
In Figure 10, there is shown a graph of surface and spar midline temperatures calculated for a channel geometry and coolant flow rate designed for A maximum 1200F metal surface temperature.

Claims (7)

What is claimed is:

1. A vane structure for a stationary combustion turbine and comprising:
a. ring means supported within a casing for the turbine;
b. a plurality of sector vane assemblies;
c. means for supporting said sector vane assemblies relative to said ring means in an annular configuration providing an annular path through which hot gas flow is directed from the combustor means;
d. each of said sector vane assemblies comprising inner and outer shroud portions having opposed surfaces defining a segment of said hot gas path and at least one airfoil vane extending across said hot gas path between said inner and outer shrouds in the annular gas flow space;
e. said airfoil vane having an inner structural member and a shell member secured to said structural member, said structural and shell members cooperating to provide said vane with an exterior airfoil shape;
f. means for channeling supply coolant air to said vane structural member and for channeling coolant air over distributed portions of the inner surface of said vane shell member to limit the outer vane shell surface temperature;
g. means for supporting said vane airfoil relative to said inner and outer shrouds;
h. each of said shrouds having internal coolant cavity means disposed over the inward side of its hot gas path surface;
i. means for directing return coolant air from said shell member channeling means to said coolant cavity means in each of said shrouds; and j. means for discharging the coolant air from said shroud coolant cavity means to the hot gas flowing through the shrouded gas path.

2. A vane structure as set forth in claim 1 wherein means for provided for discharging a portion of the vane coolant through openings in said vane to the hot gas path.

3. A vane structure as set forth in claim 2 wherein the vane openings are provided through a trailing edge of said vane.

4. A vane structure for a stationary combustion turbine comprising:
inner and outer shroud portions having opposed surfaces defining a segment of an annular hot gas path and at least one airfoil vane extending across said hot gas path between said inner and outer shrouds;
said airfoil vane having an inner structural member and a shell member secured to said structural member, said structural and shell members cooperating to provide said vane with an exterior airfoil shape;
channel means in said structural member for directing supply coolant air from a source to said skin member; and means for channeling coolant air over distributed portions of the inner surface of said vane shell member to limit the outer vane shell surface temperature;
means for supporting said vane airfoil relative to said inner and outer shrouds;
each of said shrouds having internal coolant cavity means disposed over the inward side of its hot gas path surface;
means for directing return coolant air from said shell member channeling means to coolant cavity means in each of said shrouds; and means for discharging the coolant air from said coolant cavity means through said shroud surfaces to the hot gas flowing through the shrouded gas path.

5. A vane structure as set forth in claim 4 wherein means are provided for discharging a portion of the vane coolant through openings in said vane.

6. A vane structure as set forth in claim 5 wherein the vane openings are provided through a trailing edge of said vane.

7. A vane structure as set forth in claim 4 wherein said structural member is provided with inner and outer tenons for securance respectively to said inner and outer shrouds, said supply channel means extending through said tenons to said shell member channels, and said return directing means extending from said shell member channels through said tenons to said inner and outer shrouds.