RU2499893C1 - Gas turbine engine turbine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: radial clearance adjustment means is composed of ring inserts, inner, outer and mid casings. Said conical mid casing is equipped with radial flange. Ring insert above impeller is secured to mid and inner casings. Turbine cooling systems comprise common valve. Casing cooling system comprises manifold and holes in outer casing. Turbine cooling system comprises main pipeline connected with distributor upper cavity, inner pipeline connected with swirler, front deflector with holes and holes in disc. Ring insert inner surface may be provided with soft rubbing coating. Inner surface of every circular insert may be proved with cellular seal panels secured thereat.

EFFECT: efficient control over radial clearance, higher takeoff and augmenter thrust, efficiency and reliability.

3 cl, 7 dwg

Description

The invention relates to engine building, including aircraft and stationary gas turbine engines of gas turbine engines and can find application in aircraft building, shipbuilding, gas pumping stations and for peak power plants as a drive for an electric generator designed to generate electricity.

A known turbine of a gas turbine engine according to the invention patent No. 2435039 IPC F01D 11/24, published April 27, 2008. The turbine housing includes a radial wall and contains a support for fastening the ring surrounding the turbine blades from its inner surface. The support comprises a peripheral wall surrounding the ring coaxially with it. The housing includes many perforations that provide air for uniform ventilation of the outer surface of the peripheral wall. Perforations are formed through the radial wall of the housing, passing radially inward. The wall essentially encompasses the ventilation chamber, which is also formed by the inner surface of the housing and the outer surface of the peripheral wall of the support. The ventilation chamber includes a small hole between the radial rib of the support and the inner surface of the radial wall to release air from the chamber.

Disadvantages - structural complexity and the inability to regulate the radial clearance in all engine operating modes.

Known gas turbine engine according to the patent of the Russian Federation for the invention No. 2304221 IPC F01D 11/14, publ. 08/10/07, This gas turbine engine contains a compressor having several axial stages containing a housing, guiding devices and impellers, and a turbine containing a housing and at least one stage with a nozzle apparatus and an impeller, as well as a means for regulating radial clearances but smaller at least one stage of the compressor and / or turbine.

Disadvantages - low efficiency of regulation of the radial clearance, especially in transition modes, when forcing or throttling the engine, the structural complexity of the device for regulating the radial clearance.

A gas turbine, for example, a high pressure turbine for a turbomachine, such as the French patent disclosed in Patent No. 2688539, typically contains a plurality of fixed blades arranged so that they alternate with a plurality of movable blades in the path of the hot gas from the combustion chamber of the turbomachine . Moving turbine blades are surrounded around their periphery by a stationary annular assembly. The stationary annular assembly forms a passage along which hot gas flows through the turbine blades.

In order to increase the efficiency of such a turbine, as is known, the gap that exists between the vertices of the moving turbine blades and the parts of the stationary annular assembly facing them is reduced to the value that is as small as possible.

For this, tools have been developed that provide the ability to change the diameter of the stationary annular assembly.

However, this solution is considered insufficient if the support to which the ring is attached is also exposed at its periphery to uneven thermal deformation, when such deformation leads to deformation of the turbine ring.

A turbine engine with adjustable radial clearances according to the patent of the Russian Federation No. 2435039, IPC F01D 111/04, prototype is also known. This turbine contains an outer, inner and intermediate case, a stage with a nozzle apparatus and an impeller with an annular insert above the impeller, and also means for regulating the radial clearances of at least one stage of the turbine, while the annular inserts above the impellers are fixed on the intermediate and inner enclosures,

Disadvantages - a sharp increase in the radial clearance during engine boosting due to the rapid heating of the case.

The technical result achieved by the invention is to reduce radial clearances during engine boosting.

The solution of these problems has been achieved in a turbine of a gas turbine engine, comprising an external, internal and intermediate housing, a stage with a nozzle apparatus and an impeller with an annular insert above the impeller, a turbine cooling system including an enclosure, and a radial clearance control means containing annular inserts, external, internal and intermediate housings, while the intermediate housing is conical in shape with a radial flange, while the annular insert above the impeller is fixed to the intermediate SG and inner housings, that the inventive turbine cooling system comprise a common valve and a main line connected to the upper space of the nozzle, the inner pipe connected to the unit spins, the front baffle with holes and openings in the disc. On the inner surface of the annular insert, a soft, easily abrasive coating may be applied. Honeycomb panels can be fixed to the inner surface of the annular insert.

The invention is presented in the drawings (figures 1-7), where:

- figure 1 presents a diagram of a turbine and a radial clearance control system, a first embodiment,

- figure 2 shows the installation diagram of the annular insert,

- figure 3 shows the annular insert with perforation,

- figure 4 shows the means of regulating the radial clearance with turbulators,

- figure 5 shows the means of regulating the radial clearance with the ribs,

- Fig.6 shows an annular insert with a coating of soft abradable material, type A,

- Fig.7 shows an annular insert with panels of "honeycomb seals".

The design of the turbine GTE, presented in the drawings of FIGS. 1 ... 7. The turbine engine turbine contains a shaft 1, at least one stage 2. Figure 1 shows a turbine with one high pressure stage 2. Stage 2 contains a nozzle apparatus 3 and an impeller 4 with cooled working blades 5. The working blades 5 contain a locking part 6 with radial holes 7. The impeller 4 has a disk 8, on both sides of which are installed front and rear deflectors 9 and 10. In the disk 8, holes 11 are made for supplying cooling air to the working blades 5. Holes 12 are made in the front deflector 9, in front of which a swirl apparatus 13 is mounted, to which an inner tube 14 is connected, connected to the lower cavity 15 of the nozzle apparatus 3 and then through the lower hole 16, the inner cavity 17 of the nozzle apparatus 3, the upper hole 18 with the cavity 19 and then through the sleeve 20 - with the main pipe 21 containing a common valve 22. To the main pipe 21 after the common valve 22 is connected an additional pipe 23 connected to the collector 24, holes 25 made in the outer casing 26. In addition to the outer casing 26, the turbine engine includes an inner casing 27 and an intermediate casing 28 installed between them. The intermediate casing 28 has a conical shape (the shape of a truncated cone) with a radial flange 29 connected to the flange 30 of the outer casing 26 by bolts 31. In addition, the intermediate the housing 28 has a front radial baffle 32 and a rear radial baffle 33. Holes 34 are made in the rear radial baffle 33, and an annular gap 35 is made between the front radial baffle 32 and the outer casing 26.

An annular insert 36 is attached to the intermediate casing 28 and the outer casing 26. The annular insert 36 can be cut from at least 3 parts 37 (not shown in FIGS. 1 ... 7) and mounted on the rear radial partition 33 in the annular grooves 38 and 39 of the intermediate housing 28 with brackets 40 (FIGS. 1 and 2).

The proposed means of regulating radial clearances in addition to the above parts contains four cavities front 41, middle 42, rear 43 and bottom 44.

A feature of the proposed turbine is that, to intensify the cooling of the bodies 26 ... 28 of the hole 25, it communicates with the front cavity 41.

For better cooling of the intermediate housing 28 s, holes 45 can be made on the intermediate housing 28, and holes 46 on the annular insert 36. (FIG. 3).

For a more significant intensification of cooling on the intermediate housing 28 can be performed turbulators 47 (figure 4). Turbulators 47 can be made in the form of protrusions of any shape.

For more significant intensification of the cooling of the intermediate housing 28, which actually controls the radial gap 6, fins 48 can be made. The fins 48 can be made integral with the intermediate housing 28 or mounted on it, for example with rivets (Fig. 5).

On the inner surface of the annular inserts 36, a soft, easy-to-wear coating 49 can be applied (FIG. 6) or honeycomb inserts 50 are attached (FIG. 7.

The operation of the turbine GTE is as follows (figure 1 ... 7).

With a sharp change in the operating mode of the turbine of a gas turbine engine, for example, when it is forced, the temperature of the combustion products in front of the turbine increases. In the nominal mode, the radial clearance 60, has a calculated value, and in the afterburner (maximum) mode, the radial clearances δ at the initial moment in the absence of regulation would increase sharply.

Open the common valve 22 and cooling air cools the disk 8 and the blades 5, but the outer diameter of the impeller continues to increase for two reasons

- due to an increase in the rotational speed of the shaft 1 and the disk 8,

- due to an increase in the temperature of the combustion products passing between the working blades 5,

- due to the relatively high temperature of the cooled air, usually taken due to the last stage of the gas turbine compressor (the compressor is not shown in FIGS. 1-7).

To compensate for this, the cooling air through the pipe 23 through the collector 24 and the holes 25 enters between the outer and inner bodies 26 and 27 into the cavity 41 and then through the gap 35 into the cavity 42, then through the openings 34 into the cavity 43. For several seconds, the cooling air reduces the temperature of the housings 26 ... 28, while the annular insert 36 increases the diameter, allowing the impeller 4 to touch the annular insert 36, pushing it to a larger diameter and at the same time maintaining the smallest possible radial clearance 5.

As a result, the proposed system can maintain radial clearances constant in two main modes, at maximum (afterburner) and cruising.

The application of the invention allowed:

1. To ensure effective regulation of radial clearances as in a turbine of a gas turbine engine in all modes, primarily in afterburner (maximum) and cruising modes.

2. To provide a faster (without hysteresis) increase in engine power in afterburner (maximum) modes due to the simultaneous action of cooling air on the stator (housing) and turbine rotor.

Claims (3)

1. A turbine of a gas turbine engine, comprising an outer, inner and intermediate housing, a stage with a nozzle apparatus and an impeller with an annular insert above the impeller, a cooling system for the turbine, including the housing, and a radial clearance control means comprising annular inserts, an external, inner and intermediate housings, while the intermediate housing is conical in shape with a radial flange, while the annular insert above the impeller is mounted on the intermediate and internal housings, I distinguish The fact that the turbine cooling systems comprise a common valve, the housing cooling system comprising a manifold and openings in the external housing, in addition, the turbine cooling system comprises a main pipeline connected to the upper cavity of the nozzle apparatus, an internal pipeline connected to the spin apparatus baffle with holes and holes in the disk. 2. The turbine of a gas turbine engine according to claim 1, characterized in that on the inner surface of the annular insert a soft, easily abrasive coating is applied. 3. The turbine of a gas turbine engine according to claim 1, characterized in that the honeycomb seal panels are fixed on the inner surface of the annular insert.

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