WO1982001033A1

WO1982001033A1 - Turbine cooling system

Info

Publication number: WO1982001033A1
Application number: PCT/US1980/001278
Authority: WO
Inventors: K Karstensen
Original assignee: K Karstensen
Priority date: 1980-09-24
Filing date: 1980-09-24
Publication date: 1982-04-01
Also published as: BE890264A

Abstract

An improved gas turbine cooling system for a modularized gas turbine is provided by mounting a hollow nozzle vane (18) on an outer shroud (32) to extend inwardly to an inner shroud (36) in which they are slidingly engaged. Cooling air is directed through an annulus (24) into the hollow nozzle vanes (18) which are formed with vent holes (56) on the trailing edge thereof. Both the outer (32) and inner (36) shroud portions are contained in one of the engine modules (12) while the other of the engine modules (14) is sealingly associated with the one module (12) by sealing rings (26, 28) having a greater coefficient of thermal expansion than the outer shroud (32).

Description

Turbine Cooling System

Technical Field

This invention relates to the cooling of nozzle vanes and the associated supporting shrouds in the combustion section of a gas turbine engine. In particular, it relates to such cooling in a modularized gas turbine engine wherein cooling air is provided in a circumferential manner to one module of the gas turbine for introduction into a second module of the gas turbine for ultimate use in cooling the nozzle vanes and the supporting shrouds located in the combustion section of the engine.

Background Art In turbine engines, in particular gas turbine engines, the expanding hot gas coming from the combustion portion of the engine must be directed onto the turbine blades by fixed nozzle vanes in order to achieve maximum efficiency in the engine. These nozzle vanes, which ordinarily have an airfoil shape, extend into the hot gas stream in a radial fashion in order to direct the hot gas onto the various turbine blades. As is well known, the compressor turbine powers a compressor which serves to compress air for use in the various combustion chambers. Ordinarily a certain portion of the compressed air is bled off the compressor in order to cool various portions of the engine.

In recent years, modularized gas turbines for stationary installations have been developed. The modularized engine is particularly useful in the stationary environment as disassembly is eased, thereby facilitating maintenance in that the various components are more readily reached while the engine is disassembled. An example of such a modularized gas turbine engine is found in U.S. Patent No. 4,030,288 issued to Davis and Keedy on June 21, 1977. It can be seen in this type of engine that the compressor section, which includes the nozzle vanes that form a portion of this application, and the compressor turbine are contained in one module which is interconnected with the power turbine by an engine case or gasifier section. Thus, in the particular environment in which this invention is applicable, the gas turbine engine consists of two or more modularized sections.

In utilizing modularized sections, as indicated above, it is necessary to provide a seal between the various modules. These seals are particularly important where either hot gases or cooling gases pass from one section to another. Such seals are discussed in the aforementioned patent, however, the cooling air in the usual modularized engine provided to the nozzle vanes, is provided through a chamber interior of the combustor section. Thus, cooling air provided to the nozzle vanes when passing .through the combustor section, as indicated above, arrives at the nozzle vanes at a relatively high temperature. Thus, when passing through the nozzle vanes, which are positioned in the hot gas stream, the temperature is higher than desirable. It is possible to provide a greater volume of cooling air and achieve adequate cooling of the engine; however, in providing the greater volume of cooling air, the efficiency of the engine drops. This drop in efficiency is caused by two factors. The first factor is the decrease in available air to be provided to the combustor section for powering the engine and secondly, the requirement to compress and provide the cooling air to the engine.

It should be apparent that it is desirable to provide a minimum volume of cooling air to the engine in order to keep the engine temperature within a desirable temperature range. If the cooling air is kept at a relatively low temperature up to the point where heat transfer from the hot gas stream takes place, the maximum cooling efficiency with minimum air volume is readily .attained.

In addition to the problems of providing air at the nozzle vanes at a lower temperature, the modularized gas turbine, and in fact, most gas turbines suffer when either the coolant air is bled off through leaks in the coolant system or the hot gas from the combustor section is permitted to bypass the various nozzle vanes or turbine blades or leak into the coolant stream. Thus, one goal of a turbine engine designer is to seal the hot gas stream from the coolant stream. However, in utilizing air as a coolant it is normally not feasible to have a closed cycle cooling system. Therefore cooling air is usually bled off from the cooling stream into the hot gas stream at about the point the maximum cooling has been achieved by that cooling air. To compound the problem, expansion of the various engine components during the operation of the engine can aggravate bleeding of either the cooling air into the gas stream or the hot gases into the coolant stream. This type of leakage is particularly apparent where slip joints are used between various engine components. The present invention is directed to overcoming one or more of the problems as set forth above.

Disclosure of the Invention In one aspect of this invention, an improved cooling system for a gas turbine engine is disclosed. The gas turbine engine includes at least two modules which are separable. The improvement comprises an inner shroud portion fixed to one module and an outer shroud portion also fixed to the same module. Means are provided for sealingly associating the outer shroud portion to the second module to form an annular chamber therebetween. A hollow nozzle vane having an outer and inner end is affixed to the outer shroud portion at its outer end while its inner end is slidingly received in an aperture in the inner shroud. Finally, means are provided to introduce cooling fluid into the annular chamber.

In previous modularized gas turbines, there has existed a problem in providing cooling air to the nozzle vanes of the engine along with the loss of cooling air to the hot gas stream or leakage of hot gas into the cooling stream due to insufficient sealing. The present invention provides better sealing between the nozzle vane and the supporting shrouds coupled with distribution of cooler air through the nozzle vane. The invention is advantageous in that cooling air is kept away from the hot gas stream until needed to cool the nozzle vanes. The air used to cool the rotor shroud (50) is also much cooler than in previous design, where it was heated by passage through the nozzle vanes. This allows a lower shroud temperature and closer blade tip clearance (higher efficiency) to be achieved. Secondly, the invention is advantageous in that the method of mounting the nozzle vane permits relatively easy replacement of the vanes.

Brief Description of the Drawings

Figure 1 is a view partly in section of a portion of a gas turbine engine which forms an embodiment of the present invention.

Figure 2 is a sectional view of a nozzle vane envisioned for use in the embodiment of the invention shown in Figure 1.

Figure 3 is a cross-sectional view of the nozzle vane shown in Figure 2.

Best Mode of Carrying Out the Invention Referring to Figure 1, a portion 10 of a modularized gas turbine engine is shown partly in section. Portion 10 includes a turbine section 12 and a casing section 14. It is to be understood that what is shown in Figure 1 is a cross-sectional portion of a gas turbine engine which includes, in addition to the turbine section 12 and a casing section 14, a power section (not shown). Turbine section 12 includes a compressor either of the axial or radial flow type (not shown) and a combustor ring (not shown) wherein compressed air is mixed with a fuel that is ignited so that the power from the expanding gas may be extracted to drive the compressor turbine. In Figure 1, the hot gas stream is illustrated by arrow 16. The hot gas stream passes over a plurality of fixed nozzle vanes, one of which is shown at 18, and which are arranged circumferentially in turbine section 12. After passing nozzle vanes 18, which have an air foil shape as shown in Figures 2 and 3, the hot gas is directed against a plurality of turbine blades 20 which are mounted on a turbine wheel 22, that in turn drives the compressor. Casing section 14 is slidably associated with compressor section 12 to form an annular cavity 24 thereabout. Annular cavity 24, which is maintained in a relatively air tight condition by seal means such as ring seals 26 and 28, is formed by casing section 14 and means such as an outer shroud portion 32.

Outer shroud means or portion 32 carries the plurality of nozzle vanes 18 and includes a carrier assembly 35, an outer shroud 37, and means for directing cooling toward the turbine blades 20, such means including air manifolds 64 and 66 and a spacer ring 68. Casing section 14 is fixed to turbine section 12 by appropriate fastening means such as bolts 70 and 71 which are considered part of the outer shroud portion 32. Ring seals 26 and 28 are positioned in grooves 30 and 31, each respectively formed in outer shroud portion 32 of turbine section 12. Each seal 26 and 28 is preferably made of a material having a greater coefficient expansion than either .turbine section 12 or casing section 14. With the greater rate of expansion, the seals 26 and 28 are always maintained in an abutting relationship with cone- shaped wall 33 of casing section 14, wall 33 having a first and second cylindrical surfaces 27 and 29 of different diameters to accommodate seals 26 ana 28. Seals 26 and 28 effectively seal and thus form an annular chamber 24 at the boundary with casing section 14.

On the opposite side of annular chamber 24, that is, adjacent to the hot gas stream 16, each nozzle vane 18 is affixed to an outer shroud 32 by brazing or the like. The method of fixture of nozzle vane 18 to outer shroud 32 is important to ensure that cooling air provided to annular chamber 24 is not lost to the hot gas stream 16 through the junction of nozzle blade 18 and outer shroud 32. Furthermore, the brazing of nozzle vane 18 to outer shroud 32 precludes the leakage of hot gas into annular chamber 24 thereby unnecessarily raising the temperature of cooling air provided to annular chamber 24.

At its opposite end, nozzle vane 18 is slidingly disposed in an aperture 34 of an inner shroud 36. This method of fixture permits unequal expansion of nozzle blade 18 and the inner and outer shrouds 32 and 36 respectively. Thus, as nozzle vane 18 expands at a more rapid rate than inner and outer shroud 32 and 36, nozzle vane 18 moves downwardly relative aperture 34 as indicated in Figure 1. Any gas leakage passing by nozzle vane 18 through aperture 34 from the hot gas stream 16 is contained in a second annular chamber 38 formed by a covering 40 affixed to inner shroud 36. Finally, inner shroud 36 is fixed to turbine section 12 by appropriate fastening means such as bolts 39. Cooling fluid such as cooling air is supplied to annular cavity 24 through the casing section 14. In particular, cooling air is provided through a conduit 42 from the compressor section (not shown) of the gas turbine. Conduit 42, when internal the engine casing as indicated in Figure 1, is surrounded by thermal insulation 44 to ensure the temperature of the cooling air is maintained at a relatively low level. Cooling air from conduit 42 is first communicated to a plenum chamber 46 in casing section 14. Plenum chamber 46 is annular in form and has a plurality of outlets 48 leading into annular cavity 24. Air in annular cavity 24 is provided through various internal passages to turbine shroud assembly 50 for cooling thereof.

Cooling air in annular cavity 24 is likewise communicated to the plurality of nozzle vanes 18 in the manner shown in Figure 2. Air passes into the open outer end 52 of a liner 54 in each turbine nozzle blade. Air communicated to the interior of the plurality of liners 54 passes outwardly of the liner 54 through a plurality of openings 56 formed in the trailing edge of the nozzle vanes 18. Such air is communicted directly to the hot gas stream 16 as indicated above.

In addition to the thermal insulation 44 surrounding conduit 42, additional insulation 58 and 60 isolates annular cavity 24 from the combustor inlet plenum 62. Additionally, seal 26 ensures that cooling air in annular cavity 24 is not bled off into combustor inlet plenum 62, which could occur if a different type of joint were used between the casing section 14 and turbine section 12.

Industrial Applicability Referring again to Figure 1, it can be seen that this invention is applicable in particular to gas turbine engines. It is particularly applicable to the structure for providing cooling air from one module of a gas turbine engine to a second module of a modularized gas turbine engine.

As indicated in Figure 1, cooling air is provided through a conduit 42 to a first plenum chamber 46 located in one module of gas turbine engine 10. The one module, in particular casing section 14, is associated with turbine section 12 and sealed therewith by a pair of seals 26 and 28 to form an annular chamber 24 between the portions of the gas turbine engine. Cooling air is communicated from plenum chamber 46 to annular chamber 24 through a plurality of apertures 48 formed in wall 33 of gasifier casing section 14. Cooling air is then communicated through the open outer end 52 of turbine nozzle blade 18 for cooling thereof. Cooling air communicated into turbine nozzle blade 18 picks up heat from the hot gas stream 16 by conduction through the turbine nozzle blade 18. The cooling air is discharged from the nozzle blade 18 through a plurality of apertures 56 in the trailing edge thereof.

Annular chamber 24 is sealed against the introduction of hot gas from the hot gas stream 16 by affixing by brazing or welding or the like nozzle vanes 18 to outer shroud 32. The nozzle vanes 18 are slidably fixed to inner shroud 36 through slots or apertures 34, thus the thermal growth of nozzle vanes 18 may vary from the expansion and contraction of various shroud portions in which the nozzle vanes 18 are mounted.

Other aspects, objects, and advantages of this invention can be obtained from a study of the drawings, the disclosure and the appended claims.

Claims

Clai s

1. In a cooling system for a gas turbine, the gas turbine having a turbine section (12) and a casing section (14), the improvement comprising: an inner shroud portion (36) having an aperture (34) and being affixed to said turbine section (12); an outer shroud portion (32) affixed to said turbine section (12); means (26,28) for sealingly associating said outer shroud portion (32) to said casing section (14) and forming an annular chamber (24) therebetween; a hollow nozzle vane (18) having outer and inner ends, said nozzle vane (18) being affixed to said outer shroud (32) portion at its outer end, and being slidingly received in said aperture (34) of said inner shroud (36) at- its inner end; and means (42,46) for introducing cooling fluid into said annular chamber (24) from said casing section (14).

2. The cooling system of claim 1 wherein the hollow nozzle vane (18) has a leading edge and a trailing edge, said trailing edge having a plurality of vent holes (56) therein.

3. The cooling system of claim 1 wherein said inner shroud (36) includes a covering (40), forming with said inner shroud a second annular chamber (38).

4. The cooling system of claim 1 wherein said seal means (26,28) includes a pair of spaced apart seal rings (26,28) having a greater coefficient of thermal expansion than said turbine section (12) and casing section (14).

5. The improvement of claim 1 wherein said hollow nozzle vane (18) is fixed to said outer shroud portion (32).

6. In a cooling system for a gas turbine having a turbine section (12) and a separable casing section (14), the improvement comprising: a hollow nozzle vane (18); outer shroud assembly means (32) for carrying said nozzle vane (18) and defining an annular chamber (24) with said casing section (14), said outer shroud assembly, means (32) being releasably connected to said turbine section (12); and means (42,46,48) for introducing cooling fluid into said annular chamber (24) and said hollow nozzle vane (18) from said casing section (14).

7. The cooling system of claim 6 wherein said outer shroud assembly means (32) includes a pair of spaced apart annular seal rings (26,28).

8. The cooling system of claim 7 wherein each of said seal rings (26,28) has a coefficient of thermal expansion greater than said turbine section (12) and said casing section (14).

9. The cooling system of claim 7 wherein said casing section (14) includes an annular wall (33) having a plurality of outlet passages (48) therein, said seal rings (26,28) being in contact with said wall (33) with said outlet passages (48) therebetween.

10. The cooling system of claim 9 wherein said wall (33) has first and second cylindrical surfaces (27,29) of different diameters, said seal rings (26,28) being in sealing contact with said first and second surfaces (27,29).

11. The cooling system of claim 6 wherein said nozzle vane (18) is integrally connected to said outer shroud assembly means (32).

12. The cooling system of claim 6 wherein said outer shroud assembly means (32) includes a carrier assembly (35), an outer shroud (37), and fastener means (70,71) for releasably connecting said carrier assembly (35) and said outer shroud (37) to said turbine section (12).

13. The cooling system of claim 12 wherein the gas turbine has a plurality of turbine blades (20) and said outer shroud assembly means (32) includes manifold means (64,66,68) for directing said cooling fluid from said annular chamber (24) radially inwardly upon said turbine blades (20).

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14. The cooling system of claim 6 including an inner shroud portion (36) having an aperture (34) for slidingly receiving said nozzle vane (18), said inner shroud portion (36) being releasably connected to said turbine section (12).

15. The cooling system of claim 14 including a covering (40) defining with said inner shroud portion (36) a second annular chamber (38).