RU2490474C1 - Turbine of gas-turbine engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: turbine of a gas-turbine engine includes outer, inner and intermediate housings, a stage with a nozzle assembly and an impeller with an annular insert above the impeller, as well a control device of radial gaps at least of one turbine stage. Annular inserts above impellers are fixed on intermediate and inner housings. The radial gap control device is made in the form of control rods from material with shape memory, which are borne with one of their end faces against the outer housing and with the other one against the intermediate housing. A swirling unit connected to a cooling air supply line is installed before the impeller. A valve is installed in the cooling air supply line. A hole is made in the outer housing, to which the cooling air supply line with the valve is connected.

EFFECT: effective control of radial gaps in a turbine; increasing an engine thrust in takeoff and boosting modes; improving a turbine efficiency and reliability.

4 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 04/27/08, the turbine housing includes a radial wall and contains from its inner surface a support for attaching a ring surrounding the movable blades of the turbine. 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. August 10, 07. This gas turbine engine contains a compressor having several axial stages containing a housing, guiding devices and working stakes, 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, 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 11/24, 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 creation of the invention is to reduce radial clearances during engine boosting.

The solution of these problems was achieved in a turbine of a gas turbine engine, comprising 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 adjusting 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 housings, and the means for regulating the radial clearance is made in the form of control rods of material "with shape memory" resting against one end into the outer casing, and the other into the intermediate casing. In front of the impeller there is a swirling apparatus connected to the cooling air supply line. A valve is installed in the cooling air line. An opening is made in the outer casing, to which a cooling air supply line with a valve is connected.

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 presents a diagram of the formation of a radial clearance in the turbine of the second option,

- figure 3 presents a diagram of the formation of a radial clearance in the turbine of the third option

- figure 4 shows the annular insert,

- figure 5 shows the annular insert coated with a soft abradable material, type A,

- figure 6 shows the annular insert with panels of "honeycomb seals",

- figure 7 shows a view of A.

The design of the turbine GTE shown 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. In the front deflector 9, holes 12 are made in front of which a spin apparatus 13 is installed, to which a cooling air line 14 is connected (Fig. 1).

The means for regulating the radial clearance 15 contains three cases: the outer 16, the inner 17 and the intermediate 18. The intermediate case 18 is made conical and has a radial flange 19, which is bolted 20 to the flange 21 of the outer case 16. The intermediate case 18 has a front radial wall 22, and the outer casing 16 is the rear radial partition 23. An annular insert 24 is attached to the intermediate casing 18. The annular insert 24 can be made split from at least 3 parts 25 (Figs. 4 ... 6) and mounted on the intermediate casing 18 kobami 26 (Figures 1, 4 and 5). Also included in the radial clearance control system are control rods 27 made of “shape memory” material, such as nitinol. The annular insert 24 has two annular grooves 28 and 29 (figure 4 ... 6) for its fastening.

Openings 30 are made in the rear radial partition 23, and a gap 31 is provided between the front radial partition 22 and the outer casing 16 for the passage of cooling air.

The means for regulating the radial clearance, as indicated previously, contains control rods 27 made of material with shape memory (FIG. 2). These control rods at one end abut against the intermediate casing 18, and the other against the outer casing 16. The number and shape of the control rods 27 can be any. For example, a variant of the control rods 27 is shown, which have a straight-line shape at low temperature and take a straight-line shape at high temperatures (temperature of body parts at the maximum operation mode of a gas turbine engine).

In the second embodiment (FIG. 2), the turbine comprises a valve 32 installed in the cooling air line 14.

In the third embodiment of the turbine engine of the turbine engine (Fig. 3), the turbine contains an opening 33 in the outer casing 16, to which a cooling air supply line 34 with a valve 35 is connected.

On the inner surface of the annular inserts 24, a soft, easy-to-abrade coating 36 may be applied, or honeycomb inserts 37 may be attached.

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 80, has a calculated value, and in the afterburner (maximum) mode, the radial clearance 6 at the initial moment in the absence of regulation would increase sharply.

For the first version of the turbine engine

The control rods 27 of the material “shape memory”, for example nitinol, are straightened and move the intermediate housing 18 together with the annular insert 24, which radially moves to the axis of the TBG xx (figure 1). The valve 14 is opened almost simultaneously and the 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.

For the second option of the turbine engine

In the case of applying the second option, the valve 30 is additionally opened and the cooling air flows through the cooling air supply line through a swirling apparatus and the holes 12 and 11 enter the radial holes 7 and then into the working blades 5. (Fig. 2). Due to the increase in temperature, the cooling air itself, usually taken due to the compressor (the gas turbine compressor is not shown in FIGS. 1 ... 3), the cooling of the disk 8 and rotor blades 5 is not effective enough to reduce the diameter of the impeller 4. The outer diameter of the impeller 4 continues to increase, but slightly.

To compensate for this phenomenon, i.e. the fact that after some time after heating of the disk 8 of the turbine engine (cooling the disk 8 does not reduce its temperature, but only reduces the level of heating) open valve 35 (figure 3) and cooling air through line 31 through the hole 33 enters between the outer and inner cases 16 and 17 and further through the holes 30 and the gap 31. For several seconds, the cooling air lowers the temperature of the control rods 27 and they bend, while the annular insert increases the diameter, preventing the impeller 4 from touching the annular insert 24, move Separated it to a larger diameter and at the same time maintaining the lowest-possible radial clearance δ.

On the inner surface of the intermediate insert can be applied soft abrasive coating 36 (Fig.6) or panel honeycomb seal 37 (Fig.7).

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 an increase in engine power in afterburner (maximum) modes.

3. Ensure reliable take-off of the aircraft with engines equipped with such systems for regulating the radial clearance without preliminary heating of the turbine engine or significantly reduce the warm-up time of the gas engine. This is necessary for military aircraft.

4. Ensure reliable take-off at high ambient temperatures.

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

6. Simplify the design of the elements of the radial clearance control system, reduce its weight and place the gas turbine engine outside the tract in the low temperature zone, which will increase the reliability of the turbine.

Claims (4)

1. A turbine of a gas turbine engine comprising 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 adjusting the radial clearances of at least one stage of the turbine, with annular inserts above the impellers mounted on the intermediate and inner housings, and the means for regulating the radial clearance is made in the form of control rods of material "with shape memory" resting against one end in the outer case, and the other - the intermediate casing. 2. The turbine of a gas turbine engine according to claim 1, characterized in that a swirling apparatus is installed in front of the impeller connected to the cooling air supply line. 3. The turbine of a gas turbine engine according to claim 2, characterized in that a valve is installed in the cooling air supply line. 4. The turbine of a gas turbine engine according to claim 1 or 2, characterized in that a hole is made in the outer casing to which a cooling air supply line with a valve is connected.

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