RU2414615C1 - Gas turbine engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: gas turbine engine includes HP compressor, combustion chamber and HP and LP turbines with cooled nozzle diaphragm of LP turbine and cooling channel of inter-disk cavity with supply manifold at the inlet. Cooling channel of inter-disk cavity is separated from flow part of turbine by means of labyrinth seal and slot-type gap. Supply manifold of cooling channel of inter-disk cavity through inlet manifold of cooling system of LP turbine nozzle blades and heat exchanger installed in the second engine loop is connected to axial unloading cavity of compressor. Supply manifold of cooling channel of inter-disk cavity is separated from inner cavities of cooled nozzles blades, and its connection to inlet manifold of cooling system of nozzle blades is provided by means of transit tubes installed in inner cavities of nozzle blades with a gap relative to their walls. Transit tubes are equipped with calibrated sections.

EFFECT: invention allows decreasing the temperature of cooling air supplied to inter-disk cavity to cool the side surfaces of turbine disks.

4 cl, 4 dwg

Description

The invention relates to cooling systems for gas turbine engines using air.

A two-circuit gas turbine engine is known, comprising a compressor, a combustion chamber and a two-stage turbine with a cooled nozzle apparatus between the turbine stages and the cooling disk of the interdisc cavity, the supply duct of which through a heat exchanger installed in the second circuit of the engine and the internal cavities of the cooled nozzle vanes are connected to the cavity of the high-pressure compressor (see RF patent No. 459986, IPC F02C 7/18, publ. 08.12.1976 g). The disadvantage of this solution is the use for cooling the interdisc space of very expensive air taken in front of the combustion chamber.

This drawback was eliminated in another technical solution closest to our offer, namely, in a gas turbine engine containing a high pressure compressor, a combustion chamber and high and low pressure turbines with a cooled nozzle apparatus of a low pressure turbine and a cooling channel for the interdisc space with a supply manifold at the inlet moreover, the cooling channel of the interdiscal cavity is separated from the flow part of the turbine by a labyrinth seal and a slot gap, and the supply manifold of the cooling channel of the interdisk cavity through the inlet manifold cooling nozzle vanes and a heat exchanger mounted in the second engine circuit, connected to the cavity dumisnoy high pressure compressor (see. №2200859 RF patent, IPC F02C 7/12, publ. 20.03.2003 g).

However, the cooling air cooled in the heat exchanger, passing through the internal cavities of the nozzle blades, is heated there and is supplied to the cooling of the interdisc cavity heated to such an extent that it becomes unsuitable for efficient cooling of the side surfaces of the turbine disks.

The objective of the invention is to reduce the temperature of the cooling air entering the interdisc space for cooling the side surfaces of the turbine disks.

This problem is achieved in that in a gas turbine engine containing a high pressure compressor, a combustion chamber and high and low pressure turbines with a cooled nozzle apparatus of a low pressure turbine and a cooling channel for the interdisc cavity with a supply manifold at the inlet, the cooling channel for the interdisc cavity being separated from the flow part turbines with a labyrinth seal and a gap, and the supply manifold of the cooling channel of the interdisc cavity through the inlet manifold of the cooling system of nozzle vanes low-pressure bins and a heat exchanger installed in the second circuit of the engine are connected to the dummy cavity of the high-pressure compressor, in it the supply manifold of the cooling disk of the interdisc cavity is separated from the internal cavities of the cooled nozzle blades, and its connection with the inlet manifold of the cooling system of the nozzle blades is made using transit tubes installed in the internal cavities of the nozzle blades with a gap relative to their walls, while the transit tubes are equipped with calibrated sections, and the channel l cooling of the interdiscal cavity can be made in the form of sequentially located cooling cavity of the Central part of the disk of the high pressure turbine and the cooling cavity of the root part of the disk of the high pressure turbine and the disk of the low pressure turbine, separated from each other by a labyrinth seal.

In addition, for nozzle apparatuses consisting of blocks of nozzle blades, one transit tube located in the middle blade can be made in each block of nozzle blades, and deflectors with perforations for cooling the walls of the blades can be placed between the wall of the blades and the transit tube.

New in the invention is that the supply manifold of the cooling disk of the interdisc cavity is separated from the internal cavities of the cooled nozzle blades, and its connection with the input manifold of the cooling system of the nozzle blades is made using transit tubes installed in the internal cavities of the nozzle blades with a gap relative to their walls, this transit tubes are equipped with calibrated sections, and the cooling channel of the interdisc cavity can be made in the form of consecutive cooling cavities ntralnoy part cooling the high pressure turbine disk cavity and the root portion of the disc the high-pressure turbine and low pressure turbine disk, separated by labyrinth seals.

In addition, for nozzle apparatuses consisting of blocks of nozzle blades, one transit tube located in the middle blade can be made in each block of nozzle blades, and deflectors with perforation holes for cooling the walls of the blades can be placed between the wall of the blades and the transit tube.

Isolation of the supply manifold of the cooling channel of the interdiscal cavity from the internal cavities of the cooled nozzle blades allows you to get rid of the air heated in the nozzle apparatus into this manifold, and by feeding the nozzle cooling system of the cooling of the interdisc cavity to the input manifold of the cooling system of the nozzle blades using transit tubes colder cooling air along the shortest path.

The installation of transit tubes in the internal cavities of the nozzle blades with a gap relative to the walls of the blades eliminates direct contact of the cooling air going into the interdisc cavity and the hot walls of the nozzle blades, which also reduces the heating from the hot walls of the nozzle blades of air going to cool the interdisc cavity.

The calibrated sections in the transit tubes make it possible to control the flow rate of cooling air through these tubes and to ensure the minimum residence time of the cooling air in the hottest area, which also reduces the heating of the cooling air.

The implementation of the cooling channel of the interdisc space in the form of a sequentially located cooling cavity of the central part of the high-pressure turbine disk and the cooling cavity of the root part of the high-pressure turbine disk and the low-pressure turbine disk, separated from each other by a labyrinth seal, allows directing the cooling air first to cool the most heat-stressed central section high-pressure turbine disk, and only then to cool its root part and low-pressure turbine disk.

The implementation for nozzle devices, consisting of blocks of nozzle blades, one transit tube located in the middle blade, allows you to minimize the number of transit tubes and, therefore, reduce the heating of the cooling air supplied to the cooling in the interdisc space.

Placing a deflector with perforation holes between the blade wall and the transit tube for cooling the walls of the blade, in addition to cooling the walls of the nozzle blades, slightly reduces the heating of the transit air.

Figure 1 shows a longitudinal section of a gas turbine engine;

figure 2 shows a cross section of a transit tube in the region of calibrated holes;

figure 3 shows a cross section of a feather blade with a deflector in it;

figure 4 shows a cross section of a nozzle apparatus for the case when the transit tube is placed in the middle nozzle blade.

The gas turbine engine comprises a high pressure compressor 1, a dummy cavity 2 with a labyrinth 3 between it and the last stage of the high pressure compressor 1, a combustion chamber 4 and a high pressure turbine 5, a low pressure turbine 6 with a cooled nozzle apparatus 7 and a cooling channel 8 of the interdisc cavity made in the form of a sequentially located supply manifold 9, a cooling cavity 10 of the central zone of the disk 11 of the high pressure turbine 5 and the cooling cavity 12 of the root part of the disk 11 of the high pressure turbine 5 and disk 1 3 low-pressure turbines 6, separated by a labyrinth seal 14. From the flow part of the turbine 15, the cavity 10 is separated by a labyrinth seal 16, and the cavity 12 by a gap gap 17. The supply manifold 9 through the inlet manifold 18 of the nozzle blade cooling system 19 and the heat exchanger 20, installed in the second circuit 21 of the engine, is connected to the dummy cavity 2 of the high-pressure compressor 1. The supply manifold 9 is separated from the internal cavities 22 of the cooled nozzle blades 19 by a wall 23 and connected to the inlet manifold 18 of the cooling system Nia nozzle vanes 19 transit tubes 24 mounted in the internal cavities 22 of nozzle vanes 19 with a gap 25 on their walls, the transit tubes 24 are provided with grooved portions 26 at their inlet. For nozzle apparatuses 7, consisting of blocks 27 of nozzle blades 19, in each block one transit tube 24 is placed, located in the middle blade 19. Between the wall 25 of the nozzle blades 19 and the transit tube 24 can be placed deflectors 28 with perforations 29 for cooling the walls 25 nozzle blades 19. The dumis cavity 2 is connected to the inlet manifold 18 of the cooling system of the nozzle blades 19 by the air duct 30.

The engine operates as follows.

The air from the compressor 1 through the labyrinth seal 3 enters the dummy cavity 2, from where through the struts of the combustion chamber 4 through the duct 30 through the heat exchanger 20 installed in the second circuit 21 of the engine into the inlet manifold 18 of the cooling system of the nozzle vanes 19. From the inlet manifold 20, the air flow branches out in two directions: one through the internal cavities 22 of the cooled nozzle blades 19 of the cooled nozzle apparatus 7 and exits through the trailing edge of the blade 19 into the flow part 15 of the turbines 5 and 6, and the second through calibrated the inlets 26 of the transit tubes 24 to the supply manifold 9 of the cooling channel of the interdisc cavity 8 and from it to the cavity 10 for cooling the central zone of the disk 11 of the high pressure turbine 5 and then through the labyrinth seal 14 to the cavity 12 for cooling the central part of the disk 11 of the high pressure turbine 5 and the disk 13 of the low pressure turbine 6 and through the slotted gap 17 flows into the flow part of the turbine 15, washing and cooling the side surfaces of the disks 11 and 13. Cooling air flowing through the labyrinth seal 16 from the cavity 10 into the flow part of the tour ins 15, washes the peripheral zone of the disc 11, high pressure turbine 5, the cooling of the disc.

For nozzle devices 7, consisting of blocks 27, three nozzle blades 19 in each block, air in the cavities 10 and 12 of each block 27 enters only through one transit tube 24 located in the middle blade 19 of block 27.

For the nozzle apparatus 7, in the nozzle blades 19 of which the deflectors 28 are located, cooling air from the internal cavities 22 enters through the perforations 29 to the walls 25, and then flows through the outlet edges into the flow part of the turbines 15.

The transportation of cooling air into the interdiscal cavity through the transit tubes significantly reduces the contact surface of this cooling air with the walls of the scapula, as a result of which its heating decreases, its temperature decreases, and since this air washes the surface of the design of the interdisc cavity, including turbine disks, it decreases through this temperature and the actual drive. An additional decrease in the heating of cooling air in the transit tubes occurs due to a decrease in the contact surface of the heat supply from the wall of the blades to the transit tube, when the latter are located only in the middle blades of the nozzle blocks.

The next decrease in the temperature of the disks occurs when the transit tube is placed inside the deflector with perforations, since in this case the heating of the air inside the deflector is reduced due to the fact that there is a gap between the blade wall and the deflector through which cooling air circulates.

Thus, the present invention allows to reduce the heating of the air entering the interdisc cavity, to reduce the temperature of the cooling air entering the turbine structure and, in particular, the temperature of the disks from the side of the interdisc cavity.

Claims (4)

1. A gas turbine engine comprising a high pressure compressor, a combustion chamber and high and low pressure turbines with a cooled nozzle apparatus of a low pressure turbine and an interdisc cavity cooling channel with a supply manifold at the inlet, the interdisc cavity cooling channel being separated from the turbine duct through a labyrinth seal and a gap seal the gap, and the supply manifold of the cooling channel of the interdiscal cavity through the inlet manifold of the cooling system of the nozzle blades of the low pressure turbine and heat exchange installed in the second circuit of the engine is connected to the compressor dummy cavity, characterized in that the supply manifold of the interdisc cavity cooling channel is separated from the internal cavities of the cooled nozzle blades, and its connection with the input manifold of the nozzle blade cooling system is made using transit tubes installed in the internal the cavities of the nozzle vanes with a gap relative to their walls, while the transit tubes are equipped with calibrated sections. 2. The gas turbine engine according to claim 1, characterized in that the cooling disk of the inter-disk cavity is made in the form of successively arranged cooling cavities of the central part of the high-pressure turbine disk and the cooling cavity of the root part of the high-pressure turbine disk and the low-pressure turbine disk, separated by a labyrinth seal. 3. The gas turbine engine according to claim 1, characterized in that for nozzle devices, consisting of blocks of nozzle vanes, in each block of nozzle vanes there is one transit tube located in the middle blade. 4. The gas turbine engine according to claim 1, characterized in that between the wall of the blades and the transit tube there are deflectors with perforations for cooling the walls of the blades.

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