JPS6147289B2 - - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION The present invention is directed to a new and improved shroud for minimizing the air gap between a blade tip and a shroud surrounding the blade in a small gas turbine engine.

To obtain maximum turbine efficiency, it is important to minimize air gaps between the rotating blade tips and the shroud or cylinder at the various temperatures at which gas turbine engines operate. Therefore, it is desirable to design turbine components to uniform the thermal expansion gradients of the shroud and blade tips to maintain a minimum tip clearance between the blades and the shroud. It will be readily appreciated that if the gap is too narrow, rubbing may occur between the parts, whereas if the gap is too wide, there will be a loss of efficiency.

Accordingly, it is an object of the present invention to provide a new and improved shroud or cylinder for a compressor or power turbine of a gas turbine engine. The shroud introduces cooling air through the shroud and directs the cooling air into the middle of the blade tip clearance area, which is extremely important for controlling the blade tip clearance and, therefore, for efficient operation of the gas turbine engine. The cooling air introduction means is provided to concentrate the action of the cooling air.

Yet another object of the invention is to provide an integral inner ring;
a compressor having an integral outer restriction ring configured to cooperate to define a labyrinth passageway for introducing cooling air for improved turbine tip clearance control; To provide a turbine shroud. The shroud reduces the magnitude and gradient of the temperature of the metal parts of the shroud and at the same time eliminates hot spots in the combustion chamber.
shroud integrity at high gas temperatures.

Yet another object of the present invention is to provide a new and improved air-cooled system having means for providing a film cooling flow in the air gap between the blade tips and the shroud to further improve the efficiency of a gas turbine engine. It is to provide a shroud.

In the present invention, the turbine assembly is, for example, the first
A stationary air-cooled shroud is provided extending over and spaced apart from each tip of a plurality of rotor blades, such as a staged gasification turbine.
The shroud or cylinder basically includes an inner ring, an outer restraint ring, and a support ring. The inner ring has a radially inner surface juxtaposed to and spaced apart from the rotating blade tip. The matrix includes a plurality of parallel spaced circumferential grooves extending generally perpendicular to the longitudinal axis of the inner ring. A plurality of connecting grooves are also cut into the outer surface of the inner ring. Each connecting groove also extends parallel to the longitudinal axis of the inner ring and communicates between two adjacent circumferential grooves. The connecting grooves are staggered along the circumference of the inner ring. An outer restraining ring is disposed coaxially with and surrounding the inner ring so as to cover the groove arrangement of the inner ring and define an air cooling labyrinth passage. A support ring cooperates with the inner and outer rings to define a plenum chamber for cooling air provided by the turbine compressor and an inlet flow path. Cooling air from the plenum chamber enters the inlet flow passage and is forced into a labyrinth passage consisting of circumferential connecting grooves, thereby reducing the magnitude and slope of thermal expansion of the shroud and improving the blade tip. The gap can be controlled. The gap is controlled by the combined stress reaction of the inner cooled ring and the outer restraint ring. The inner and outer rings can be assembled in line contact, but they can also be assembled with some gap, depending on the desired constraints between the two parts, or in a tight fit. It may also be possible to assemble it. In addition to the benefit of being able to control the clearance, the present cooled shroud would allow an increase in turbine inlet temperature that would not be possible with an uncooled shroud.

The invention further provides that the labyrinth includes a metering opening extending between a circumferential groove and a leading edge of the inner ring to provide a film cooling effect on the inner radial surface of the inner ring. , thereby assisting in maintaining proper tip clearance and purging hot gases at the leading edge of the shroud or cylinder.

The above and other objects, features and advantages of the present invention will become more apparent from the following description taken in conjunction with the accompanying drawings.

First, the basic structure of the turbine rotor shroud, which is the so-called basis of the present invention and which the present invention aims to improve, will be explained. Figures 1 to 3 show the basic structure of a turbine rotor shroud, with the first stage of the gas producer turbine being designated by the number 10;
It is connected to a combustion chamber (not shown) by an annular combustion chamber outlet 12 . Combustion gases, indicated by arrow G, are provided to the turbine and first enter the first stage stationary vanes 14 and then the first stage blades 1 of the turbine rotor.
6. Each rotating blade 16 has a base portion 18 secured to the rotating assembly of the turbine, for example at 20, and further has an airfoil portion 22. As shown in FIG.
The tip of the blade is designated by the number 24. The airfoil portion may include air cooling passages as shown in dotted lines. Annular shroud or cylinder 3
0 extends around and surrounds the first stage rotor blade 16. The shroud basically consists of an inner annular ring 32, an outer annular restraint ring 34 and an annular support ring 36. Preferably, each ring 32, 34 and 36 is of unitary construction and extends around the entire circumference of the turbine rotor. Furthermore, preferably an inner annular ring 3
2 and the annular support ring 36 are integral and formed from a single integral member. The inner ring 32 is generally cylindrical in shape, with its longitudinal axis coinciding with the longitudinal axis of the turbine rotor, and whose longitudinal axis coincides with the longitudinal axis of the turbine rotor.
2 has a radially inner surface 40 and a radially outer surface 42. The spacing between the radial inner surface 40 and the blade tip 24 is indicated by the letter "t" and represents the air gap between the first stage rotor blade 16 and the shroud 30. A primary object of the present invention is to arrange the shroud cooling means to maintain a substantially uniform tip clearance "t" between the shroud and the first stage turbine rotor over all operating conditions of the turbine engine.

As illustrated in FIGS. 1-3, a series of groove formations, numbered in the 50s, are provided in the outer radial surface 42 of the inner ring to provide a labyrinth for cooling the shroud or cylinder 30. Forms part of a passage. The groove structure is a circumferential groove 5
2, 54, 56 and 58. The circumferential grooves are spaced parallel to each other, extend around the entire circumference of the inner ring, and are generally perpendicular to the longitudinal axis of the inner ring. The groove has a circumferentially extending barrier 51,
Bounded by 53, 55, 57 and 59, the leading edge of the groove structure is defined by barrier 51, while the trailing edge of the groove structure is defined by barrier 59. As illustrated in FIGS. 1 and 2, the width and depth of each groove 52-58 can be designed such that the cross-sectional areas of the circumferential grooves are substantially equal. Barriers 51, 53, 55, 57 and 5
By providing 9, the flow of pressurized cooling air through the groove arrangement will be substantially uniform.

Referring to FIGS. 1 and 2, the leading edge barrier 5
1 comprises a plurality of openings 60 passing through the barrier. The apertures 60 are preferably equally spaced around the entire circumference of the inner ring 32. The opening 60 provides communication between the first circumferential groove 52 and an inflow passage designated by the numeral 70. The inlet passageway is defined by outer restraint ring 34, inner ring 32 and support ring 36. The inflow passage 70
has a plenum chamber 72 communicating with opening 60 and an inlet chamber 74 communicating with plenum chamber 72 via opening 76 . Cooling air provided from a gas turbine engine compressor is delivered to inlet chamber 74 through opening 80, as shown in FIG. Cool air from the compressor may also be provided via opening 80 to a second opening 82 that communicates with the combustion gas flow path. The cooling air flow is indicated by the letter "p" and is effective to sweep away hot gases located upstream of the inner ring 32 and also between the radially inner surface 40 of the inner ring 32 and the rotor blades 16. A film of cooling air is provided in the region of the tip spacing "t" between.

As shown in FIG. 3, the arrangement of the connection grooves provided in each barrier is configured to be shifted in position with respect to the arrangement of connection grooves of adjacent barriers. Therefore, the connecting grooves 62 and 66 in the barriers 53 and 57 are arranged so as to be aligned, but are configured to be offset from the intermediate connecting groove 64. As in the case of the openings 60 in the leading edge barrier 51, the connecting grooves 62, 64 and 66 are also spaced, for example, at intervals of 30° around the entire circumference of the inner ring 32.
It is formed so that 12 connection grooves are provided. A connecting groove 68 is provided in the trailing edge barrier 59 . The groove can be arranged at an angle to the longitudinal axis of the inner ring. This configuration provides cooling air exiting the trailing edge of the groove arrangement with an initial swirl flow that substantially coincides with the hot combustion gas flow path downstream of the rotating blades 16. The trailing edge of the restraining ring 34 is provided with a notch, indicated by the number 35, so that the connecting groove 68 has a cross-sectional area sufficient to prevent pressure build-up within the labyrinth passage.

In operation, pressurized cooling air from the gas turbine engine's compressor is provided to opening 80, then to plenum chamber 72, and distributed through a series of connecting openings 60 to the labyrinth passages. As the cooling air flows through each connecting groove 60 (see FIG. 3), it impinges on the barrier 53 and thus in the circumferential groove 5.
2 to flow laterally in two opposite directions. Each airflow then impinges on one of the series of connecting grooves 62 provided in the second barrier 53. Said air flow operation is repeated in the circumferential groove 54. This configuration effectively distributes pressurized cooling air over the entire circumference of inner ring 32, thereby providing convective cooling of shroud 30. The pressurized cooling air is supplied to shroud 3.
It is discharged from the labyrinth passage through a connecting groove 68 with a predetermined angle provided on the trailing edge of the 0.
With this configuration, the shroud and rotating blade 1
Air gap control between the inner cooling ring 3, which is under tangential condition during convection cooling of the shroud and assembly,
2 and the outer restraining ring 34. In addition to the benefit of providing clearance control, the cooled shroud 30 allows for increased turbine inlet temperatures that are not possible with uncooled shrouds or cylinders. The convection cooling effect of the shroud reduces the height and gradient of the shroud metal temperature, and the thermal deformation of the shroud due to hot spots in the combustion chamber is also effectively reduced. The cooling shroud is thus able to maintain its integrity even at the high gas temperatures encountered in various operating conditions of gas turbine engines.

FIGS. 4 and 5 illustrate the basic structure of a turbine rotor shroud, which is also the basis of the so-called present invention and which the present invention seeks to improve. Illustrated herein is a gas turbine engine having an annular combustion chamber outlet, the shroud flange support connection means being disposed adjacent the trailing edge of the shroud, and the shroud extending from the flange to the shroud. It is cantilevered by a support arm projecting to the leading edge. Therefore, with this configuration,
Flange supports are disposed radially of the shroud as well as cantilevered support rings or support arms for the shroud. The flange support is therefore spaced apart from the hot combustion gases flowing through the turbine. 1-3 in which the shroud flange is disposed adjacent the leading edge of the shroud.
The shroud support structure shown is quite different.

In the structure illustrated in FIG. 5 and FIG.
It will be readily appreciated that there is a significant temperature difference (and therefore a significant thermal expansion difference) between the cantilevered arm structure compared to the shroud that extends around the turbine rotor. The thermal expansion difference is primarily compensated for by the inherent deflection of the cantilevered arm support. In the embodiment illustrated in FIGS. 5 and 6, the air-cooled shroud 130 includes an integral annular restraint ring 132 and an integral annular inner ring 13.
4, and an integral annular support ring 136. Inner ring 134 cooperates with annular restraint ring 132 to define a labyrinth passageway 138 for introducing pressurized cooling air. three rings 132,
134 and 136 cooperate to form the labyrinth passage 13
8. Defining an inlet flow path 140 leading to 8. Support ring 136 and restraint ring cooperate to define an elongated fluid flow passageway to the inlet flow passageway 140. The outer restraining ring 132 has a flange portion 133;
On the other hand, the support ring flange 136 is the flange portion 1
It has 37. Flanges 133, 137 are suitably joined to support the annular shroud structure. With this configuration, the shroud has a support flange 13
3,137, thereby providing a relatively flexible support connection for the shroud. Therefore, Labyrinth 13
8 and the components of the structure defining the support arms are compensated for by the deformation of the cantilevered support structure. Pressurized cooling air provided from the engine's compressor is introduced through conduit 142 and thus inlet flow passage 140 and then flows into labyrinth 138 . The labyrinth 138 has a circumferential groove 144;
The inner ring 132 has alternating connecting grooves extending from the leading edge to the trailing edge of the inner ring 132. As illustrated in FIG. 5, each circumferential groove 144 is configured to have substantially the same shape and cross-sectional area so that the mass of pressurized cooling air flowing through the labyrinth passage is uniform. be done.

Next, the structure of the turbine rotor shroud according to the present invention will be explained. Although the present invention can be suitably applied to any of the structures described above, a case will be described in which the present invention is applied to a structure similar to that shown in FIGS. 4 and 5.

In the embodiment of the invention illustrated in FIGS. 6 and 7, cantilevered flange supports are also provided on the inner and outer rings forming part of the air-cooled shroud. The outer restraining ring 132' and support ring 136' of the air-cooled shroud 130' are the same as elements 132 and 136 in the construction of FIGS. 4 and 5. On the other hand, the inner ring 134' has a plurality of grooves 162 to 1 extending in the annular inner circumferential direction extending from the leading edge to the trailing edge of the inner ring 134'.
68. Adjacent circumferential grooves are interconnected by connecting grooves designated by numbers 170-178. Adjacent to the leading edge circumferential groove 162 to provide pressurized cooling air flowing through the labyrinth of passageway 160 with a purge cooling air flow to film-cool the tip gap between the inner ring and the blade tip. An extension 182 is provided in the barrier extending between the circumferential groove 164 and generally aligned with the longitudinal axis of the inner ring. Barrier extension 182 connects connecting groove 170 provided between inlet flow 140' and circumferential groove 162.
extends through. In particular, in the present invention, metering apertures 184 are provided in the extension barrier 182 and in the circumferential groove 164 of the inner ring 134' and the radially inner surface 1 of the inner ring 134'.
35'. Each connecting groove 170 extending around the circumference of the inner ring and communicating between the inlet flow passage 140' and the leading edge groove 162 has a barrier extension 182 extending through the center and thus has a corresponding plurality of A metering aperture 184 is provided in the inner ring 134'. This configuration allows pressurized cooling air to flow into the leading edge groove 1.
62, through the connecting groove 172 and into the second circumferential groove 164, a portion of the cooling air passes through the next row of staggered connecting grooves 174.
while the remaining portion of the pressurized cooling air is directed through metering opening 184 to the radially inner leading edge 135' of inner ring 134'. Thus, cooling air entering the hot leading edge location of the air-cooled shroud 130' flows through the labyrinth passageway to the intermediate portion of the shroud. While some desired amount of cooling air continues to flow toward the trailing edge of the shroud, the remaining portion of the cooling air is directed back toward the leading edge of the shroud through metering openings 184 and out of the upstream recess of the shroud. The hot gases of the rotor are purged, thus providing film air cooling of the tip gap between the inner ring and the rotor blades.

Therefore, a restraining ring and means for introducing pressurized cooling air through the shroud are provided to reduce the temperature magnitude and gradient of the metal parts of the shroud and to reduce thermal distortions due to hot spots in the combustion chamber. A new and improved fixed air cooled turbine shroud is provided that is capable of maintaining the structural integrity of the shroud even at high temperatures during various operating modes of a turbine engine. The means for directing pressurized cooling air through the shroud is in the form of a labyrinth passage formed by a groove arrangement in the inner ring and a cooperating outer restraint ring in line contact with the arrangement. The flow flowing through the labyrinth passage is either a parallel flow type or a divided flow type. In the parallel flow type, the cooling air enters a groove arrangement located at the leading edge of the shroud where it is hottest, and the pressurized cooling air is arranged circumferentially to completely cool the circumference of the shroud. flow within the groove structure toward the trailing edge. As pressurized cooling air flows through the shroud, the greatest convective cooling effect occurs at the leading edge of the shroud, so that the temperature profile of the shroud from the leading edge to the trailing edge is substantially flat. Become. In a split flow configuration, as illustrated in the embodiment of FIGS. 6 and 7, pressurized cooling air enters the relatively hot leading edge portion of the shroud and flows in a downstream direction. At approximately the midpoint of the shroud, a metered amount of pressurized flow is redirected toward the leading edge of the shroud through a suitable metered opening in the labyrinth passage. The remaining portion of the pressurized air stream continues to flow in the downstream direction. The redirected pressurized airflow toward the shroud leading edge purges the upstream recess and provides film cooling of the inner shroud wall, particularly the void area between the shroud and the rotor blade. In an embodiment of the invention, the labyrinth passageway of the shroud is formed by groove formations cut into the inner ring, each groove formation being staggered to provide a labyrinth-type flow passage. It consists of a series of parallel and circumferential grooves with connecting grooves. The groove arrangement in the inner ring is covered by a restraining ring, and the leading edge portion of the inner ring is further restrained by a support ring. The support ring also cooperates with the restraint ring and inner ring to define an inlet flow recess for pressurized cooling air received from a compressor of a gas turbine engine.

The present invention has been described above in connection with embodiments, but
The present invention is not limited to the structure and form of the embodiment illustrated herein, and various embodiments are possible without departing from the spirit and scope of the present invention, and various changes and modifications may be made. It will be clear to those skilled in the art that this can be done.

[Brief explanation of the drawing]

FIG. 1 is a partial sectional view of the first stage of a gas producer turbine having an air-cooled shroud, the basic structure of which is the basis of the present invention. Figure 2 shows the first
FIG. 3 is a partial cross-sectional view of another form of an air-cooled shroud similar to that shown in the figure; FIG. 3 is a cross-sectional view taken along line 3--3 of FIG. FIG. 4 is a partial sectional view of an air-cooled shroud having a basic structure on which the present invention is based. FIG. 5 is a cross-sectional view taken along line 6--6 of FIG. FIG. 6 is a cross-sectional view of an embodiment of the invention. FIG. 7 is a cross-sectional view taken along line 8--8 of FIG. 10: gas producer turbine, 12: annular combustion chamber outlet, 14: first stage stationary vane, 16: first
Stage rotor blade, 18: Blade base portion, 2
4: Blade tip portion, 30... Annular shroud,
32: Inner annular ring, 34: Outer annular restraint ring, 36: Annular support ring, 40: Shroud radially inner surface, 42: Shroud radially outer surface, 5
0: Matrix (groove structure), 51: Leading edge barrier, 52, 54, 56, 58: Circumferential groove, 5
9: Trailing edge barrier, 60: Opening, 68: Connection groove, 7
0: Inlet flow path, 72: Plenum chamber, 74: Inlet chamber, 80: Opening.

Claims (1)

Claims: 1. A fixed air cooling shroud forming part of a turbine assembly and extending spaced apart above the tips of a plurality of rotor blades, comprising: a radially inner surface and a radially outer surface. With an inner ring,
The inner surface is juxtaposed to and spaced apart from each blade tip, and the outer surface of the inner ring has a groove formation thereon, the groove formation extending from an upstream to a downstream portion of the shroud. at least first, second and third spaced apart and parallel circumferential grooves extending generally perpendicular to the longitudinal axis of the inner ring; a plurality of connecting grooves extending generally parallel to the inner ring, each connecting groove providing communication between two adjacent circumferential grooves, and the inner ring having a membrane-like shape on the radially inner surface thereof. further comprising a metering aperture extending between said second circumferential groove and a leading edge of said inner ring to provide cooling; an outer restraint ring coaxially surrounding the inner ring and covering the groove arrangement; further cooperating with the inner ring and the outer restraint ring for cooling air provided to the labyrinth passageway; A support ring is provided that defines an inlet flow passage so that cooling air flow through the labyrinth passage reduces the magnitude and slope of shroud thermal expansion and provides effective control of turbine blade tip clearance. The air cooling shroud is characterized by: 2. The air cooling shroud according to claim 1, wherein the support ring and the inner ring are formed of an integral member. 3. The air-cooled shroud of claim 1, wherein each circumferential groove has a constant cross-sectional shape. 4. The air-cooled shroud of claim 1, wherein the cross-sectional area of each circumferential groove is substantially the same. 5. The air-cooled shroud according to claim 1, wherein the connecting grooves extending between two adjacent sets of circumferential grooves are arranged alternately over the circumference of the inner ring. 6. The air cooling shroud according to claim 1, wherein the connecting groove at the trailing edge of the labyrinth passage is arranged at a certain angle with respect to the longitudinal axis of the inner ring. 7. The air cooling shroud according to claim 1, wherein the inner ring is an integral member. 8. The air cooling shroud of claim 1, wherein the outer restraining ring is an integral member. 9. The air cooling shroud of claim 1, wherein the support ring and the outer restraint ring define a plenum chamber for receiving cooling air, the plenum chamber communicating with an inlet flow passage leading to a labyrinth passage.