RU2498087C1 - Gas-turbine engine turbine - Google Patents

Abstract

FIELD: machine building.

SUBSTANCE: gas-turbine engine turbine comprises outer, inner and intermediate casings, stage with nozzle diaphragm and rotor wheel with annular insertion above the rotor wheel, cooling system for the turbine and casing, and radial clearance adjuster. The said adjuster comprises annular insertions, outer, inner and intermediate casings. The intermediate casing is of conical form with radial flange. The annular insertion above the rotor wheel is fixed on the intermediate and inner casings. The turbine and casing cooling system consists of a valve, main pipeline connected to the upper cavity of the nozzle diaphragm, inner pipeline connected to the swirler, a front baffle with holes and holes in the disk.

EFFECT: effective control of radial clearances in turbine, increasing engine thrust in takeoff and afterburner mode, increasing turbine efficiency factor and reliability.

3 cl, 3 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.

Known turbine for a gas turbine engine according to patent for invention No. 2435039 IPC F01D H / 24, published on 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 adjusting radial clearances of at least 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, such as a high-pressure turbine for a turbomachine, such as the French patent publication No. 2688539 disclosed in France, 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 coming 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, it is known that 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 smallest possible value.

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, as well as 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 creation of 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, and a radial clearance control device containing annular inserts, an external, internal and the intermediate housing, while the intermediate housing is made of a conical shape with a radial flange while the annular insert above the impeller is fixed to the intermediate and internal pusah fact that according to the invention comprises a turbine cooling valve and the 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 1-3), where:

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

- figure 2 shows the annular insert coated with a soft abradable material, type A,

- figure 3 shows the annular insert with panels of "honeycomb seals",

The design of the turbine GTE presented in the drawings of FIGS. 1 ... 3. 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. In the front deflector 9, holes 12 are made 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 the common valve 22. To the main pipe 21 after the common valve 22 is connected an additional pipe 23 connected to the hole 24 made in the outer casing 25. In addition to the external of the turbine engine turbine housing 25 comprises an inner casing 26 and an intermediate casing 27 installed between them. The intermediate casing 27 has a conical shape (the shape of a truncated cone) with a radial flange 28 connected to the flange 29 of the outer casing 25 by bolts 30. In addition, the intermediate casing 27 has a front radial baffle 31 and a rear radial baffle 32. Holes 33 are made in the rear radial baffle 32, and an annular gap 34 is made between the front radial baffle 31. and the outer casing 25.

An annular insert 35 is attached to the intermediate casing 27 and the outer casing 25. The annular insert 35 can be made split from at least 3 parts 36 (not shown in FIGS. 1 ... 3) and mounted on the rear radial partition 32 in the annular grooves 37 and 38 of the intermediate casing 27 of the annular grooves 37 and 38 of the brackets 39 (figure 1 ... 3).

On the inner surface of the annular inserts 35, a soft, easy-to-wear coating 40 may be applied, or honeycomb inserts 41 may be attached.

The operation of the turbine GTE is as follows (figure 1 ... 3).

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 δ0 has a calculated value, and in the afterburner (maximal) mode, the radial clearance δ at the initial moment in the absence of regulation would increase sharply.

The common valve 2 is opened 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. 1-3).

To compensate for this, the cooling air through the pipe 23 through the hole 24 enters between the outer and inner bodies 25 and 26 and then through the holes 33 and the gap 34. Within a few seconds, the cooling air lowers the temperature of the bodies 25 ... 27, while the annular insert 35 increases the diameter, on allowing the impeller 4 to touch the annular insert 35, pushing it to a larger diameter and at the same time maintaining the smallest possible radial clearance 6.

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 whiter rapid increase in engine power in afterburner (maximum) modes due to the simultaneous cooling of the stator and rotor.

3. Ensure reliable take-off of the aircraft with engines equipped with such systems for regulating the radial clearance without preliminary heating of the gas turbine engine or significantly reduce the time of gas turbine engine heating due to the simultaneous heating of the stator and rotor. This is especially necessary for military aircraft.

4. Ensure reliable take-off at high ambient temperatures for the same reason.

5. Almost instantly transfer the operation mode of the gas turbine engine from cruising to afterburning mode for the same reason. This is especially important for military aircraft.

Claims (3)

1. The 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 and the housing, and a radial clearance control means comprising annular inserts, an external, internal and intermediate housing, while the intermediate housing is made in a conical shape with a radial flange, while the annular insert above the impeller is mounted on the intermediate and inner housings, characterized in the turbine housing and cooling system comprises a valve main conduit connected with the upper space of the nozzle, the inner pipe connected to the unit spins, the front baffle with holes and openings in the disc. 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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