WO1992007180A1

WO1992007180A1 - Radial turbine containment ring system

Info

Publication number: WO1992007180A1
Application number: PCT/US1991/005794
Authority: WO
Inventors: Phillip Vessa; Jesus Aguilar Garcia
Original assignee: Sundstrand Corporation
Priority date: 1990-10-22
Filing date: 1991-08-15
Publication date: 1992-04-30

Abstract

In order to avoid penetration of the housing (38) of a gas turbine (10) in the event of burst failure, the gas turbine (10) includes a containment ring system (32). The gas turbine (10) is of the type which includes a rotor (12) having a turbine wheel (14) with blades (16) and a nozzle (18) outwardly of the turbine blades (16). The nozzle (18) is adapted to direct hot gases at the turbine blades (16) to cause rotation of the rotor (12). The gas turbine (10) is also of the type having an annular combustor (20) about the rotor (12) along with an outlet (22) to the nozzle (18). The annular combustor (20) is defined by spaced inner and outer walls (24 and 26) connected by a generally radially extending wall (28). With this arrangement, the gas turbine (10) utilizes a radial turbine containment ring system (32) which has a first containment ring (34) outwardly of the rotor (12) and the nozzle (18) and a second containment ring (36) at least partially about the rotor (12) outwardly of the turbine blades (16) to deflect fragments toward the first containment ring (34) while dissipating fragment energy in the event of burst failure.

Description

RADIAL TURBINE CONTAINMENT RING SYSTEM

Field of the Invention

This invention is generally related to gas turbines and, more particularly, a radial inflow turbine and a containment ring system therefor.

Background of the Invention

One concern in the operation of gas turbine engines is the possible catastrophic failure of any portion of the turbine wheel including the hub, the blades, or both because such failures typically occur when the gas turbine engine is in operation. As a result, and because the rates of revolution of turbine wheels in gas turbine engines are quite, the resultant high angular velocity translates to high centrifugal forces acting on the turbine wheel. Should there be a failure where any part of the turbine wheel cracks, breaks or begins to disintegrate, the presence of this high centrifugal force will cause the separate component to move outwardly at high velocity with substantial kinetic energy.

To prevent damage to surrounding instrumentalities, it has long been common to provide gas turbine engines with so-called containment rings which are generally formed of metal and are disposed radially outwardly of the turbine wheel, usually to some degree axially to each side thereof. Generally speaking, the containment rings are made of material with sufficient strength to halt and arrest further radial outward movement upon catastrophic failure of the turbine wheel as to all parts moving radially outwardly due to impingement upon the containment ring.

Of course, strength is a primary consideration in containment rings particularly for land based or marine gas turbines. And, of course, cost is always, of concern. However, in the case of airborne gas turbines, it will be appreciated that weight becomes an important consideration in addition to cost and strength. Generally speaking, it has been known to place a containment ring in the plane of rotation of a turbine wheel in a position that is radially outwardly thereof. This is shown, for instance, in conjunction with axial flow turbines in commonly owned U.S. Patent No. 4,639,188 and, additionally, it is shown in U.S. Patent Nos. 4,452,565 and 3,928,963. While such structures are useful, they do not address a problem that exists, particularly for applications which require a radial inflow turbine wheel.

For such applications, it is not uncommon for in the event of catastrophic failure for the fragments to move both radially and axially. In other words, a given fragment may have both a radial velocity component and an axial velocity component causing it to travel along a path which would miss the containment ring positioned in the plane of rotation of the turbine wheel. If this occurs, the fragments can cause serious damage to surrounding instrumentalities which is potentially devastating.

By way of example, such fragments may damage or destroy such important components as hydraulic, communication, and/or power lines. In the case of airborne applications, the potential result from this need not be explained. As a result, there is a serious need to confine the trajectory of fragments to the plane of rotation of the rotor.

Of course, in any containment ring system, the containment ring or rings usually are brought very near to those areas of the gas turbine that are exposed to high temperatures. This, in turn, means that increasingly exotic materials, whose cost is a consideration, must be used in order to withstand the high temperatures and thermal cycling in the environment in which they are placed. Despite this fact, the containment ring or rings must also be able to reliably provide containment in order to avoid the dangers that are inherent in a catastrophic failure.

Of course, it has long been known in the field of gas turbine engines to provide a dilution air zone in the combustor. This zone is conventionally located directly within the combustion annulus downstream of the fuel injectors but well upstream of the outlet of the combustor. Generally speaking, dilution air is injected into the combustion annulus to control the temperature of hot gases.

More particularly, upstream of the dilution zone both fuel and air are injected and ignited in the combustion annulus. It is also conventional for there to be a cooling air film introduced along the walls of the combustion annulus upstream of the dilution zone. Of course, the hot gases that result from combustion then pass toward turbine blades to drive the turbine wheel.

As is known, it is important to be able to control the temperature of the hot gases as they enter the nozzle on their way to the turbine blades. This has conventionally been handled by means of a dilution zone within the combustion annulus well upstream of the outlet of the combustor in order to ensure mixing and cooling prior to entry into the nozzle. While effective, this means of controlling the temperature of the hot gases is not satisfactory in all respects or for every application. More recently, Sundstrand Corporation has developed a number of techniques for enhancing the ability to control the temperature of the hot gases as they enter the nozzle on their way to the turbine blades. While so doing, means have been established for cooling the walls of the combustor by reason of establishing a dilution air flow path substantially thereabout while also cooling the turbine shroud and, thus, the nozzle and turbine blades. According to the present invention, such advantageous cooling means are also employed to enhance the life span of the containment ring system within the gas turbine.

stnmnary of the Invention

It is a principal object of the present invention to provide a new and improved containment system for a gas turbine. More specifically, it is an object of the invention to provide such a containment system which utilizes a primary containment ring and a secondary containment and deflection ring which cooperatively interact to retain fragments within a plane of rotation of a turbine wheel. It is a further object of the present invention to provide a containment system that is dilution air cooled. An exemplary embodiment of the invention achieves the foregoing objects in a gas turbine which includes a rotor having a radial inflow turbine wheel with turbine blades and a turbine nozzle assembly outwardly of the turbine blades. The nozzle is adapted to direct hot gases at the turbine blades to cause rotation of the rotor. An annular combustor is disposed about the rotor and has an outlet to the nozzle. The annular combustor has spaced inner and outer walls connected by a generally radially extending wall opposite of the outlet. A radial turbine containment ring system includes a first containment ring and a second containment ring. With this arrangement, the first containment ring is outwardly of the rotor and the nozzle and the second containment ring is at least partially about the rotor outwardly of the turbine blades. Additional details of the invention include a housing defining a dilution air flow path about at least a portion of the annular combustor. The dilution air flow path advantageously includes a radially outer compressed air inlet in communication with a compressor supplying dilution air at an upstream end thereof. Preferably, the first containment ring is fixedly mounted outwardly of the housing at the upstream end of the dilution air flow path.

In addition, the gas turbine may include a diffuser leading from the compressor to the upstream end of the dilution air flow path. The first, or primary, containment ring is then advantageously fixedly mounted to the diffuser so as to be generally concentric with the outer wall of the annular combustor at a point radially inwardly of the housing, As a result, the first containment ring may advantageously be at least axially coextensive with the turbine blades in a plane of rotation of the rotor.

As will be appreciated, the first containment ring may be fixedly mounted outwardly of the annular combustor at the upstream end of the dilution air flow path. Thus, it may be exposed to cooling air from the compressor that enters the dilution air flow path which means that it may be made of less exotic materials and still have a long life span. It will also be appreciated that the gas turbine preferably includes front and rear turbine shrouds wherein at least the front turbine shroud substantially radially bounds the turbine blades in the plane of rotation thereof. Thus, the front turbine shroud preferably includes a radially extending portion and an axially extending portion radially inwardly of the housing. With this arrangement, the first containment ring is advantageously disposed between the axially extending portion of the front turbine shroud and the housing. As for the second containment ring, it is advantageously fixedly mounted relative to the housing at a point corresponding to the downstream end of the dilution air flow path. As for the rear turbine shroud, it substantially radially bounds the turbine blades in the plane of rotation thereof and generally conforms to the contour of the turbine blades and extends radially outwardly thereof to the nozzle in a most highly preferred embodiment. Preferably, the second containment ring is disposed radially outwardly of the rear turbine shroud between the shroud and the combustor at least partially axially coextensive with the turbine blades.

In a highly preferred embodiment, a housing substantially surrounds the annular combustor in spaced relation to the inner, outer and radially extending walls thereof including at least the rear turbine wheel shroud. The housing and walls together define a dilution air flow path extending from a compressed air inlet to at least one dilution air outlet where the compressed air inlet is in communication with a source of dilution air. In this connection, the dilution air outlet is in communication with the annular combustor downstream of the compressed air inlet preferably adjacent the second containment ring.

In a most preferred embodiment, the containment ring system includes a primary containment ring and a secondary containment and deflection ring. The primary containment ring is positioned outwardly of the rotor with the secondary containment and deflection ring being fixedly mounted and generally conforming to the rear turbine wheel shroud for containing fragments within the plane of rotation and deflecting such fragments toward the primary containment ring. While so doing, the containment ring system is arranged so as to be effective for dissipating fragment energy in the event of burst failure.

Other objects, advantages and features of the present invention will become apparent from a consideration of the following specification taken in conjunction with the accompanying drawings.

Brief Description of the Drawings

Fig. 1 is a somewhat schematic, fragmentary sectional view of a gas turbine in accordance with the present invention; and

Fig. 2 is a somewhat schematic, fragmentary sectional view similar to Fig. 1 but illustrating a modified embodiment thereof.

Detailed Description of the Preferred Embodiments

In the illustrations given, and with reference first to Fig. 1, the reference numeral 10 designates generally a gas turbine in accordance with the present invention. The gas turbine 10 includes a rotor generally designated 12 which defines a radial inflow turbine wheel 14 having blades 16 and a nozzle 18 outwardly of the turbine blades 16. The nozzle 18 is adapted to direct hot gases at the turbine blades 16 to cause rotation of the rotor 12. The gas turbine 10 also includes an annular combustor 20 about the rotor 12 and having an outlet 22 to the nozzle 18. As shown, the annular combustor 20 has spaced inner and outer walls 24 and 26 connected by a generally radially extending wall 28. Still referring to Fig. 1, a radial turbine containment ring system 32 is defined by a first containment ring 34 and a second containment ring 36. The first containment ring 34 is positioned outwardly of the rotor 12 and the nozzle 18. In addition, the second containment ring 36 is at least partially about the rotor 12 immediately outwardly of the turbine blades 16.

As will be appreciated, the gas turbine 10 includes a housing 38 defining a dilution air flow path 40, as shown by the arrows, about at least a portion and preferably all of the annular combustor 20. The dilution air flow path 40 includes a radially outer compressed air inlet as at 42 in communication with a compressor 44 supplying dilution air at an upstream end thereof. With this arrangement, the first containment ring 34 is fixedly mounted inwardly of the housing 38 at the upstream end of the dilution air flow path 40 in a plane of rotation of the rotor 12.

As will be seen from Fig. 1, the gas turbine 10 includes a diffuser 46 leading from the compressor 44 to the upstream end of the dilution air flow path 40. The first containment ring 34 is, in the illustrated embodiment, fixedly mounted to the diffuser 46 (see, also. Fig. 2) by any conventional means and is generally concentric with the outer wall 26 of the annular combustor 20 at a point radially inwardly of the housing 38. As illustrated in Figs. 1 and 2, the first containment ring 34 is at least axially coextensive with the turbine blades 16 at a point outwardly of the annular combustor 20.

Additional details of the gas turbine engine 10 include a pair of front and rear turbine shrouds 48 and 50 wherein the front turbine shroud 48 substantially radially bounds the turbine blades 16. In fact, the front turbine shroud 48 radially bounds the turbine blades 16 in the plane of rotation thereof inasmuch as it includes a radially extending portion 48a and an axially extending portion 48b radially inwardly of the housing 38. In the embodiments illustrated in Figs. 1 and 2, the first containment ring 34 is suitably disposed between the axially extending portion 48b of the front turbine shroud 48 and the housing 38.

Also, in the embodiments illustrated in Figs. 1 and 2, the second containment ring 36 is fixedly mounted relative to the housing 38 at a downstream end of the dilution air flow path 40. It will be seen and appreciated that the rear turbine shroud 50 substantially radially bounds the turbine blades 16 in the plane of rotation thereof and generally conforms to the contour of the turbine blades 16 while extending radially outwardly thereof to the nozzle 18. With this arrangement, the second containment ring 36 is disposed radially outwardly of the rear turbine shroud 50 substantially in the plane of rotation.

As shown, the second containment ring 36 is at least partially axially coextensive with the turbine blades 16 and is preferably disposed between the rear turbine shroud 50 and the annular combustor 20. It is, however, preferably located at a point downstream of the primary combustion zone 52 whereby dilution air flowing about the dilution air flow path 40 may be introduced through openings as at 54 to flow along the combustor-facing surface 36a of the second containment ring 36. As will be appreciated, the second containment ring 36 also shields a portion of the rear turbine shroud 50 as well as the turbine blades 16 from the heat generated in the annular combustor 20. Further, it will be appreciated that the housing 38 and walls 24, 26 and 28 together define the dilution air flow path 40. It will also be noted that the dilution air flow path 40 extends from the compressed air inlet 42 to at least one dilution air outlet such as 54, and preferably a plurality of such outlets substantially as shown in the drawings whereby air film cooling may be achieved, but it will also be appreciated that a primary dilution air outlet 56 will advantageously be provided for introduction of a large supply of dilution air into a dilution air zone 58 downstream of the combustion zone 52. Of course, the dilution air outlets 54 and 56 are all in communication with the annular combustor 20 downstream of the compressed air inlet 42. With the foregoing arrangement as described, the containment ring system 32 will be well appreciated as including a primary containment ring 34. It will also be seen that the second containment ring 36 essentially comprises a secondary containment and deflection ring inasmuch as its position and shape generally conforming to the rear turbine wheel shroud 50 and turbine blades 16 serves well to not only contain fragments within the plane of rotation but also to deflect such fragments toward the primary containment ring 34. In addition, and while so doing, the secondary containment and deflection^'ring 36 also serves to dissipate fragment energy in the event of a burst failure.

While not previously described, it will be appreciated that the secondary containment and deflection ring 36 is fixedly mounted relative to the housing 38 at the downstream end of the dilution air flow path 40 by any conventional means. This may comprise threading the ring 36 to the turbine shroud 50 as at 60, or using pins or any other type of fastening means that will maintain structural integrity even when fragments from the turbine wheel 12 make high energy contact therewith. In any event, the secondary containment and deflection ring 36 is suitably disposed at a position which is radially outwardly adjacent the rear turbine shroud 50 but well within in the plane of rotation of the rotor 12.

Also as shown in the drawings, the primary containment ring 34 is mounted so as to be exposed to dilution air at a point which is located externally of the annular combustor 20. It will be seen that the front turbine shroud 48 and the outer wall 26 are each interconnected as at 62 by any conventional means with the primary containment ring 34 being radially outwardly thereof and adjacent the upstream end of the dilution air flow path 40 where it is suitably cooled. Also as shown, the radially extending portion 48a of the front turbine shroud 48 and the radially outermost portion 50a of the rear turbine shroud 50 together define the nozzle 18 radially well inwardly of the primary containment ring 34. With the present invention, the containment ring system 32 serves to dissipate fragment energy as previously described while also serving to reduce fragmentation. The latter occurs due to the close shrouding effects of the primary containment ring 34 and the secondary containment and deflection ring 36 in relation to the turbine blades 16. Still further, the containment ring system 32 serves to deflect and channel fragments within the rotor burst plane minimizing the probability of impact on the housing 38. Referring to Fig. 2, it will be appreciated that like reference numerals designate like components in the respective views wherein the gas turbines have many common components. It will be seen, however, that Fig. 2 differs from Fig. 1 in that the inner wall 24 includes an inner wall segment 24a extending the dilution air flow path 40 whereby the secondary containment and deflection ring 36 is disposed directly in the dilution air flow path outwardly of the annular combustor 20 due to the fact that the inner wall portion 24a extends to a point immediately adjacent the outlet 22 of the combustor 20 where it is secured to the radially outermost portion 50a of the rear turbine shroud 50 adjacent the nozzle 18. As will be appreciated, this provides even additional cooling effects to the secondary containment and deflection ring 36 but while giving up a certain volume within the dilution zone 58 of the annular combustor 20.

While in the foregoing there have been set forth preferred embodiments of the invention, it will be appreciated that the details herein given may be varied by those skilled in the art without departing from the true spirit and scope of the appended claims.

Claims

CLAIMSWe Claim;

1. A gas turbine, comprising: a rotor including a turbine wheel having blades and a nozzle outwardly of said turbine blades, said nozzle being adapted to direct hot gases at said turbine blades to cause rotation of said rotor; an annular combustor about said rotor and having an outlet to said nozzle, said annular combustor having spaced inner and outer walls, said inner and outer walls being connected by a generally radially extending wall; a radial turbine containment ring system including a first containment ring outwardly of said rotor and said nozzle and a second containment ring at least partially about said rotor outwardly of said turbine blades.

2. The gas turbine as defined in claim 1 including a housing defining a dilution air flow path about at least a portion of said annular combustor, said dilution air flow path including a radially outer compressed air inlet in communication with a compressor supplying dilution air at an upstream end thereof, said first containment ring being fixedly mounted inwardly of said housing at said upstream end of said dilution air flow path.

3. The gas turbine as defined in claim 2 including a diffuser leading from said compressor to said upstream end of said dilution air flow path, said first containment ring being fixedly mounted to said diffuser and being generally concentric with said outer wall of said annular combustor at a point radially inwardly of said housing, said first containment ring being disposed in a plane of rotation of said rotor.

4. The gas turbine as defined in claim 2 including a diffuser leading from said compressor to said upstream end of said dilution air flow path, said first containment ring being fixedly mounted to said diffuser and being generally concentric with said outer wall of said annular combustor at a point radially inwardly of said housing, said first containment ring being at least axially coextensive with said turbine blades.

5. The gas turbine as defined in claim 1 including a housing defining a dilution air flow path about at least a portion of said annular combustor, said dilution air flow path including a radially outer compressed air inlet in communication with a compressor supplying dilution air at an upstream end thereof, said first containment ring being fixedly mounted outwardly of said annular combustor at said upstream end of said dilution air flow path.

6. The gas turbine as defined in claim 2 including at least a front turbine shroud substantially radially bounding said turbine blades in a plane of rotation thereof, said front turbine shroud including a radially extending portion and an axially extending portion radially inwardly of said housing, said first containment ring being disposed between said axially extending portion of said front turbine shroud and said housing.

7. The gas turbine as defined in claim 1 including a housing substantially surrounding said annular combustor in spaced relation to said inner, outer and radially extending walls thereof, said housing defining a dilution air flow path extending substantially entirely about said annular combustor to cool said inner, outer and radially extending walls thereof, said second containment ring being fixedly mounted relative to said housing at a downstream end of said dilution air flow path.

8. The gas turbine as defined in claim 1 including at least a rear turbine shroud substantially radially bounding said turbine blades in a plane of rotation thereof, said rear turbine shroud generally conforming to the contour of said turbine blades and extending radially outwardly thereof to said nozzle, said second containment ring being disposed radially outwardly of said rear turbine shroud in said plane of rotation.

9. The gas turbine as defined in claim 1 including at least a rear turbine shroud substantially radially bounding said turbine blades in a plane of rotation thereof, said rear turbine shroud generally conforming to the contour of said turbine blades and extending radially outwardly thereof to said nozzle, said second containment ring being at least partially axially coextensive with said turbine blades.

10. The gas turbine as defined in claim 1 including at least a rear turbine shroud substantially radially bounding said turbine blades in a plane of rotation thereof, said rear turbine shroud generally conforming to the contour of said turbine blades and extending radially outwardly thereof to said nozzle, said second containment ring being disposed between said rear turbine shroud and said annular combustor.

11. A gas turbine, comprising: a rotor including a radial inflow turbine wheel having blades and a nozzle adjacent said turbine blades, said nozzle being adapted to direct hot gases at said turbine blades to cause rotation of said rotor; an annular combustor about said rotor defined by spaced inner and outer walls connected by a generally radially extending wall, said annular combustor having an outlet leading to said nozzle, said annular combustor also including a combustion annulus defined by said inner, outer and radially extending walls upstream of said outlet; a housing substantially surrounding said annular combustor in spaced relation to said inner, outer and radially extending walls thereof and including at least a rear turbine wheel shroud, said housing and walls together defining a dilution air flow path extending from a compressed air inlet to at least one dilution air outlet where said compressed air inlet is in communication with a source of dilution air, said dilution air outlet being in communication with said annular combustor downstream of said compressed air inlet; and a containment ring system including a primary containment ring outwardly of said rotor and said nozzle and a secondary containment and deflection ring fixedly mounted and generally conforming to said rear turbine wheel shroud.

12. The gas turbine as defined in claim 11 including a diffuser leading from a compressor to said compressed air inlet at an upstream end of said dilution air flow path, said primary containment ring being fixedly mounted to said diffuser and being generally concentric with said outer wall of said annular combustor at a point radially inwardly of said housing, said primary containment ring being disposed outwardly of said annular combustor in a plane of rotation of and generally coextensive with said rotor.

13. The gas turbine as defined in claim 12 including a front turbine shroud substantially radially bounding said turbine blades in said plane of rotation of said rotor, said front turbine shroud including a radially extending portion and an axially extending portion radially inwardly of said housing, said primary containment ring being disposed between said axially extending portion of said front turbine shroud and said housing.

14. The gas turbine as defined in claim 11 wherein said secondary containment and deflection ring is fixedly mounted relative to said housing at a downstream end of said dilution air flow path, said secondary containment and deflection ring being disposed radially outwardly adjacent said rear turbine shroud in a plane of rotation of said rotor, said secondary containment and deflection ring being at least partially axially coextensive with said turbine blades and disposed between said rear turbine shroud and said annular combustor.

15. A gas turbine, comprising: a rotor including a radial inflow turbine wheel having blades and a nozzle outwardly of said turbine blades, said nozzle being adapted to direct hot gases at said turbine blades to cause rotation of said rotor; an annular combustor about said rotor and having an outlet to said nozzle, said annular combustor having spaced inner and outer walls, said inner and outer walls being connected by a generally radially extending wall; a pair of front and rear turbine shrouds together substantially radially bounding said turbine blades in a plane of rotation thereof, said rear turbine shroud being opposite of said radially extending wall and in close proximity to said turbine blades; and a containment ring system including a primary containment ring outwardly of said rotor, nozzle and front turbine shroud and a secondary containment and deflection ring fixedly mounted and generally conforming to said rear turbine wheel shroud for containing fragments within said plane of rotation and deflecting fragments toward said primary containment ring while dissipating fragment energy in the event of burst failure.

16. The gas turbine as defined in claim 15 including a housing defining a dilution air flow path extending at least partially about said annular combustor and a diffuser leading from a compressor to said dilution air flow path, said primary containment ring being fixedly mounted to said diffuser at a point radially inwardly of said housing and being disposed outwardly of said annular combustor and being generally coextensive with said rotor.

17. The gas turbine as defined in claim 16 wherein said front turbine shroud includes a radially extending portion and an axially extending portion radially inwardly of said housing, said primary containment ring being disposed between said axially extending portion of said front turbine shroud and said housing, said radially extending portion of said front turbine shroud and a radially outermost portion of said rear turbine shroud together defining said nozzle.

18. The gas turbine as defined in claim 16 wherein said secondary containment and deflection ring is fixedly mounted relative to said housing at a downstream end of said dilution air flow path, said secondary containment and deflection ring being disposed radially outwardly adjacent said rear turbine shroud and being generally arcuate so as to channel and deflect fragments toward said primary containment ring, said secondary containment and deflection ring being at least partially axially coextensive with said turbine blades ami disposed between said rear turbine shroud and said annular combustor.