RU2733682C1 - Cooling method of working blades of turbine of double-flow gas turbine engine and device for its implementation - Google Patents

Abstract

FIELD: turbines or turbomachines.

SUBSTANCE: group of inventions relates to high-temperature turbines of gas turbine engines, namely, to methods and systems of cooling of turbine blades of aircraft engines. Technical result is achieved in cooling method of turbine working blades of double-flow gas turbine engine, which includes extraction of cooling air from air cavity of combustion chamber, its transportation through an air-to-air heat exchanger installed in the air path of the second circuit, and through the inlet cavity into the spinning device, further supply of "cold" air together with "hot" air between turbine disk and deflector into inner cavities of working blades through air channels in turbine wheel and control of flow rate of both "cold" and "hot" air, characterized in that the inlet cavity is divided into two cavities, the swirling unit, respectively, into two parts, each of which is connected to several supply pipelines with cutoff valves, air pressure at the working blades inner cavities is increased in the centrifugal compressor made between the turbine disk and the deflector, and control of supply and "cold" and "hot" air modes is discrete. Technical result is achieved in a device for cooling a working blade of a turbine of a double-flow gas turbine engine, comprising an air-to-air heat exchanger installed in an air path of the second circuit, an inlet cavity into a spinning device, cold air supply system together with hot air between turbine disk and deflector into inner cavities of working blades through air channels in turbine wheel and system to control flow rate of both cold and hot air, characterized in that the inlet cavity is divided into two cavities, the swirling device – into two parts, respectively, to the input of each of which there connected are several supply pipelines with shutoff valves, and to the outlet – centrifugal compressor located between the turbine disk and the deflector.

EFFECT: higher efficiency of gas turbine engines with high-temperature turbines.

6 cl, 14 dwg

Description

The group of inventions relates to high-temperature turbines of gas turbine engines, and in particular to methods and systems for cooling rotor blades of aircraft turbines. It is on aircraft engines that a wide range of regulation in terms of power, speed and temperature level in front of the turbine is required. One more circumstance should be borne in mind that on these types of engines the maximum engine power mode is short-term, and cruising modes are long-term in the engine life cycle.

The closest invention to the present invention is a method for cooling the rotor blades of a turbine of a by-pass gas turbine engine and a device for its implementation according to RF patent No. 2196239, IPC F02C 7/12, publ. 10.01.2003

The method for cooling the rotor blades of a high-pressure turbine of a by-pass gas turbine engine according to the patent of the Russian Federation No. 2196239 includes the selection of cooling air from the air cavity of the combustion chamber, its transportation through an air-to-air heat exchanger installed in the air duct of the secondary circuit, into the swirl apparatus, the subsequent supply of cooling air to the internal cavities of the rotor blades through the air channels in the turbine rotor and regulation of its flow rate.

A device for cooling the rotor blade of a turbine of a two-circuit gas turbine engine according to RF patent No. 2196239 contains an air-to-air heat exchanger installed in series in the second circuit, connected by its input to the air cavity of the combustion chamber, and by its output to an air manifold with control valves in it, a multi-channel air duct passing through the inner cavities of the nozzle blades, the stator swirl apparatus and air channels in the impeller connected to the inner cavities of the rotor blades.

In the solution according to the patent of the Russian Federation No. 2196239, in order to obtain acceptable efficiency of the engine in the entire operating range, the maximum air flow is supplied at maximum modes, and in cruising modes, the required cooling air flow is reduced using control valves installed on the manifold after the heat exchanger. In this case, the most dangerous from the point of view of cooling of the rotor blade is the zone of its leading edge, therefore, the cooling air flow rate is reduced to a guaranteed level, which ensures reliable cooling of the leading edge in all operating modes of the engine. In this case, the "colder" zones of the blade surface receive an "excess" of cooling air in comparison with the optimal consumption of cooling air for these engine modes, which worsens the efficiency of the gas turbine engine. Therefore, for high-temperature turbines in cruising modes, one has to put up with excessive cooling of the middle and outlet parts of the airfoil of the rotor blades, which reduces its efficiency. On the other hand, an increase in the operating temperature of the gas in front of the turbine is necessary in order to obtain decent efficiency values at the turbine.

The known method and system for cooling the turbine rotor blades according to the RF patent for invention No. 2 387864 IPC F01D 5/18, publ. 04/23/2010, prototype.

This method of cooling the rotor blades of a turbine of a two-circuit gas turbine engine includes the selection of cooling air from the air cavity of the combustion chamber, its transportation through an air-to-air heat exchanger installed in the air duct of the secondary circuit, into the swirl apparatus, and the subsequent supply of cooling air to the inner cavities of the rotor blades through the air channels in the turbine rotor wheel and regulation of its flow rate, the inner cavity of each rotor blade located at the input edge is separated from the rest of the cavity by a partition directed along the input edge, the formed cavity is communicated by perforations in the wall with the turbine flow path and cooling air is supplied to it from air cavity of the combustion chamber through an additional stator swirling apparatus and through additional air channels in the impeller.

This is a device for cooling the rotor blade of a turbine of a two-circuit gas turbine engine, containing a series-installed air-to-air heat exchanger located in the second circuit, connected by its input to the air cavity of the combustion chamber, and by its output to an air manifold with control valves in it, a multi-channel air duct passing through the inner cavities of the nozzle blades, the stator swirling apparatus and air channels in the impeller connected to the inner cavities of the rotor blades, a bladeless diffuser is placed on the impeller between its air channels and the stator swirling apparatus, the inner cavity of each rotor blade, located at the input edge, is separated from the rest of the cavity by a partition directed along the input edge, the formed cavity is communicated by perforations in the wall with the turbine flow path and is connected to the air cavity of the combustion chamber through an additional stator swirl apparatus and through an additional air channels in the impeller.

Disadvantages: the use of this method and device has several disadvantages:

- the use of two concentrically installed swirling devices and variable valves significantly complicates the design of the turbine and worsens the operation of the cooling system for the following reasons. Since the air flow for cooling the impeller cannot exceed 5% ... 7% of the air flow through the primary circuit, the height of the nozzle blades will be very small. And this will lead to a relative increase in the thickness of the boundary layer and an increase in aerodynamic losses in them.

The use of adjustable valves will lead to negative consequences: namely, a sharp drop in pressure at the inlet to the swirling apparatus and a change in the air velocity triangles at the outlet of the swirl nozzles, this will lead to a shock inlet of cooling air under the deflector, i.e. loss of pressure at the inlet to the rotor blades, and as a consequence - to the flow of combustion products in the cavity of the rotor blades. The use of diffusers slightly increases the static pressure at the inlet to the air ducts, but due to the sudden narrowing of the air flow at the inlet to these ducts, significant pressure losses will occur.

All this makes it impossible to deeply regulate the flow of cooling air and, as a result, leads to a deterioration in the efficiency of operation of gas turbine engines with such cooling systems.

The objective of the invention is to improve the efficiency of gas turbine engines with high-temperature turbines by optimizing the flow of cooling air in the rotor blades of high-pressure turbines in the entire range of operation of a multi-mode engine while maintaining a satisfactory temperature state of the cooled blades, that is, while maintaining the reliability and service life of the engine.

This task is achieved by the fact that in the method of cooling the rotor blades of a turbine of a two-circuit gas turbine engine, including the selection of cooling air from the air cavity of the combustion chamber, its transportation through an air-to-air heat exchanger installed in the air duct of the secondary circuit, and through the inlet cavity into the swirl apparatus, the subsequent supply of "cold" air together with "hot" air between the turbine disk and the deflector into the inner cavities of the rotor blades through air channels in the turbine impeller and regulation of the flow rate of both "cold" and "hot" air, characterized in that the inlet cavity is divided into two, the swirling apparatus, respectively, into two parts, to which several supply pipelines with shut-off valves are connected, the air pressure at the inlet to the inner cavities of the rotor blades is increased in a centrifugal compressor made between the turbine disk and the deflector, and regulation of the supply and "cold" and "hot" air ha for modes is performed discretely.

The first and second cavities of the blade are divided into several sectors and the supply pipelines are connected to each sector, while at the maximum mode all the shut-off valves are open, and in other modes they are closed in turn.

At maximum engine operation, equality of angles can be ensured:

α ₁ = β ₀ ,

where: α ₁ - angle of outflow of cooling air from the swirling apparatus,

β ₀ - angle of installation of the leading edge of the blade,

and in other modes, provide the closest possible values of the angles α ₁ and β ₀ .

This task is also achieved by the fact that in a device for cooling the rotor blade of a turbine of a two-circuit gas turbine engine containing an air-air heat exchanger installed in the air duct of the second circuit, an inlet cavity into the swirling apparatus, a system for supplying "cold" air together with "hot" air between a turbine disk and a deflector into the inner cavities of the rotor blades through the air channels in the turbine rotor and the system for regulating the flow of both "cold" and "hot" air, by the fact that the inlet cavity is divided into two, the swirl apparatus, respectively, into two parts, to the inlet of which is connected by several supply pipelines with shut-off valves, and to the outlet there is a centrifugal compressor located between the turbine disk and the deflector.

The first and second cavities can be divided into several sectors and lead pipes are connected to each sector.

Each rotor blade can have two cavities: front and rear, with perforations of the leading edge inserted into the front cavity.

The essence of the invention is presented in the drawings, figures 1 ... 14, where:

- in Fig. 1 shows an example of a specific implementation of a device for cooling the rotor blade of a high-pressure turbine of a by-pass gas turbine engine, which allows to implement the proposed method,

- in Fig. 2 shows a diagram of the cooling air supply,

- in Fig. 3 shows a diagram of the cooling air supply, the second option,

- in Fig. 4 shows the first version of the cooling system,

- in Fig. 5 shows a diagram of the cooling air supply,

- in Fig. 6 shows a diagram of the cooling air supply,

- in Fig. 7 shows the turbine assembly, the first stage,

- in Fig. 8 shows a turbine blade, the first option,

- in Fig. 9 shows a turbine blade, the second option,

- in Fig. 10 shows the design of a centrifugal compressor, the first option,

- in Fig. 11 shows the design of a centrifugal compressor, the first option,

- in Fig. 12 shows a fragment of the drawing of the impeller,

- in Fig. 13 shows view A,

- in Fig. 14 shows a drawing of the sector layout and the supply pipeline, section B - B.

List of designations used in the description.

air-to-air heat exchanger 1,

second circuit 2,

entrance 3,

air cavity 4,

combustion chamber 5,

exit 6,

"cold" air duct 7,

shut-off valve 8,

inner cavity 9,

nozzle blades 10,

inlet cavity 11,

spinner 12,

stator 13,

air ducts 14,

impeller 15,

inner cavity 16,

working blades 17,

deflector 18,

centrifugal compressor 19,

leading edge 20,

perforation holes 21,

turbine flow path 22,

the first cavity 23,

second cavity 24,

radial partitions 25,

first part 26,

second part 27,

hot air duct 28,

shut-off valve 29,

additional partitions 30,

sector 31,

streamlined blades 32,

outer seal 33,

inner seal 34,

hub 35,

disc 36,

shaft 37,

control unit 38,

communication lines 39,

lock 40,

bandage shelf 41,

seal 42,

trailing edge 43,

exit slit 44,

partition 45,

anterior channel 46,

back channel 47,

blade 48,

the leading edge of the blade 49,

diffuser 50,

ribs 51,

α ₀ - angle of installation of streamlined blades 32

α ₁ - angle of outflow of cooling air from the swirling apparatus,

β ₀ - angle of installation of the leading edge of the blade,

The cooling device for the rotor blade of a high-pressure turbine of a two-circuit gas turbine engine (Fig. 1 ... 14) contains a series-installed air-air heat exchanger 1, located in the second circuit 2, connected by its outlet 3 with the air cavity 4 of the combustion chamber 5 (secondary air zone of the combustion chamber) , and outlet 6 - with air ducts of "cold" air 7 with shut-off valves 8 in them, passing through the internal cavities 9 of the nozzle blades 10, the inlet cavity 11, the twisting apparatus 12 of the stator 13 and the air channels 14 in the impeller 15, connected to the internal cavities 16 rotor blades 17, which are part of the impeller 15. On the impeller 15 between its air channels 14 and the twisting device 12 of the stator 13, there is a deflector 18 with a centrifugal compressor 19. Perforations 21 are made on the water edges of the rotor blades 17 for communication with the flow part of the turbine 22.

A feature of the device is that (Fig. 2) that the inlet cavity 11 is divided into two cavities: the first 23 and the second 24, separated by radial partitions 25, and the swirling apparatus 12, respectively, into two parts: the first 26 and the second 27, to the entrance to the first part 23 is connected to the "cold" air ducts 7 with shut-off valves 8, and to the second part 24 are connected to several hot air ducts 28 with shut-off valves 29. A centrifugal compressor 19 is connected to the outlet of the swirl apparatus 12, located between the impeller 15 and deflector 18.

A variant of the cooling system (FIG. 3) is possible in which the first cavity 23 and the second cavity 24 are divided into equal sectors 31 by additional partitions 30.

One “cold” air duct 7 with shut-off valves 8 is connected to the inlet to the sectors 31 of the first part 23, and one supply pipelines of “hot” air 28 with shut-off valves 29 are connected to the inlets to the sectors 31 of the second part 24 (Fig. 3) ...

The swirling apparatus 12 contains streamlined blades 32 and is sealed by external and internal seals 33 and 34 (Figs. 1 and 6). A shaft 37 is attached to the hub 35 of the disc 36 (Fig. 6).

The device contains a control unit 38. The control unit 38 is connected by communication lines 39 with shut-off valves 8 and 29 (Fig. 1).

FIG. 8 shows the first version of a turbine blade 17. It contains an inlet edge 20, perforations 21, a shroud 41, an outer seal 42, an outlet edge 43, an outlet slot 44.

FIG. 9 shows the second version of the turbine blade 17. It additionally contains a baffle 45, and the inner cavity 16 is divided by a baffle 45 into a front channel 46 and a rear channel 47.

FIG. 10 shows the design of a centrifugal compressor, the first option, the centrifugal compressor 19 contains radial blades 48.

FIG. 11 shows the design of a centrifugal compressor, the second option, the centrifugal compressor contains curved blades 48.

FIG. 12 shows a fragment of the drawing of the impeller 14, it contains blades 48 under the deflector 16 with the input edges of the blade 49.

FIG. 13 is a view A showing the curved leading edges of the blade 49 and FIG. 14 shows a drawing of the sector layout and the supply pipeline, section В-В. It can be seen that the leading edges of the blades 49 are made at an angle α ₀ = β ₀ , where:

α ₀ - the angle of installation of the streamlined blades 32 of the swirl apparatus 12,

α ₁ - the angle of outflow of the cooling air flow from the swirling apparatus 12,

β ₀ - angle of installation of the leading edge of the blade 49.

The method is implemented as follows:

The supply of cooling air from the air cavity 4 of the combustion chamber 6 for cooling the rotor blades 17 is carried out through two independent channels (Fig. 1). Through the first channel, the cooling air is taken from the air cavity 4 and transported to the inner cavity 16 of the working blade 17 through a sequentially placed air-air heat exchanger 1 installed in the air duct of the second circuit 2, “cold” air ducts 7 with shut-off valves 8 in them, twisting apparatus 12, centrifugal compressor 18, air channels 14 in the impeller 15.

Through the second channel, the cooling air is taken from the air cavity 4 and transported through the “hot” air ducts to the inner cavity 9 of the working blade 17 through sequentially placed shut-off valves 29, a swirling apparatus 12, a centrifugal compressor and air channels 14 in the impeller 17.

From the air channels 14, air through the perforations 21 in the leading edge 20 enters the flow path of the turbine 22, carrying out film cooling of the wall of the rotor blade 17 in the region of its leading edge 20, which can ensure reliable operation of this zone at very high gas temperatures. To reduce the pressure loss of the cooling air in the cavity behind the swirling apparatus 12, on the one hand, it is separated from the cavity at the inlet to the compressor 19 of the impeller 15 by an external seal 33 and an internal seal 34.

The flow rate of the cooling air entering the internal cavities 16 is reduced as the engine operation mode is throttled using the shut-off valves 8 and 29. It is possible to regulate the flow rate of the cooling air entering the cavity of the first and second cavities 23 and 24 of the inner cavity 16 of the rotor blade 17, if it is divided by a partition 25.

At maximum engine operating conditions, all shut-off valves 8 and 19 are open.

With a decrease in engine speed and gas temperature in front of the turbine, the consumption of cooling air is reduced by closing the shut-off valves 8 and 29. Thus, the consumption of cooling air in the internal cavities 16 of the blades 17 decreases and increases in the tract of the combustion chamber 5, thereby increasing the mass of the working fluid in the turbine. To maintain the engine operating mode, the fuel supply to the combustion chamber 5 is reduced, which lowers the gas temperature in front of the turbine and reduces the specific fuel consumption of the engine, i.e. improves efficiency.

The use of shut-off valves 8 and 29 for each sector 31 serving 3 ... 7 streamlined blades 32 of the swirl apparatus 12 allows maintaining optimal speeds and directions of the cooling air flow at all engine operating modes at the outlet of the swirl apparatus 12.

Diffuser 50 and ribs 51 (Fig. 14) in the cooling system reduce losses from sudden air expansion when entering the inlet cavity 11 and increase the pressure of the cooling air entering the inner cavities of the working blade 17, or into the front and rear channels 46 and 47, in the presence of partitions 45 in the rotor blades 17, which means to improve the cooling of the rotor blades 17.

The use for each sector 31 of its own shut-off valve allows at low engine operating conditions to reduce by 2 ... 4 times the consumption of cooling air, without reducing its pressure and without changing the triangles of the velocities of the outflow of cooling air from the swirling apparatus 12 (Fig. 14).

U - disk rotation speed,

V is the total speed of the cooling air flow,

V ₀ - axial component of the air outflow velocity,

V _u - the circumferential component of the air outflow velocity,

α ₀ - the angle of installation of the streamlined blades 32 of the swirl apparatus 12,

α ₁ - the angle of outflow of the cooling air flow from the swirling apparatus 12,

β ₀ - angle of installation of the leading edge of the blade 49.

Considering that approximately:

α ₁ = α ₀ ,

it is proposed to ensure equality of angles at maximum engine operation:

α ₁ = β ₀ ,

where: α ₁ - angle of outflow of cooling air from the swirling apparatus,

β ₀ is the angle of installation of the leading edge of the blade, and in other modes provide the closest possible values of the angles α ₁ and β ₀ . This will provide a shock-free intake of cooling air to the centrifugal compressor and reduce air pressure losses in the cooling system.

Thus, the invention allows, on the one hand, to ensure the reliability and specified service life of the engine, and, on the other hand, high efficiency in the designs of high-temperature turbines in a wide range of power (revolutions) control of the gas turbine engine.

This is especially true for modern high-temperature marine and aircraft engines.

Claims (10)

1. A method of cooling the rotor blades of a turbine of a two-circuit gas turbine engine, including the selection of cooling air from the air cavity of the combustion chamber, its transportation through an air-to-air heat exchanger installed in the air duct of the secondary circuit, and through the inlet cavity into the swirling apparatus, the subsequent supply of "cold" air together with the "hot" air between the turbine disk and the deflector into the inner cavities of the rotor blades through the air channels in the turbine impeller and regulation of the flow rate of both "cold" and "hot" air, characterized in that the inlet cavity is divided into two cavities, the apparatus swirls, respectively, into two parts, to which several supply pipelines with shut-off valves are connected, the air pressure at the inlet to the inner cavities of the rotor blades is increased in a centrifugal compressor made between the turbine disk and the deflector, and regulation of the supply of both "cold" and "hot »Air by modes performed discretely tno. 2. The method of cooling the rotor blade of a turbine of a by-pass gas turbine engine according to claim 1, characterized in that the first and second cavities are divided into several sectors and the supply pipelines are connected to each sector, while at the maximum mode all the shut-off valves are open, and in other modes they are are alternately closed. 3. The method of cooling the turbine rotor blade of a by-pass gas turbine engine according to claim 1 or 2, characterized in that at the maximum engine operation, equality of angles is ensured α ₁ = β ₀ , where α ₁ is the angle of outflow of the cooling air flow from the swirl apparatus, β ₀ - angle of installation of the leading edge of the blade, and in other modes, provide the closest possible values of the angles α ₁ and β ₀ . 4. A device for cooling the rotor blade of a turbine of a two-circuit gas turbine engine, containing an air-air heat exchanger installed in the air duct of the second circuit, an inlet cavity into the swirling apparatus, systems for supplying "cold" air together with "hot" air between the turbine disk and a deflector into the internal cavities of the rotor blades through the air channels in the turbine impeller and a system for regulating the flow of both "cold" and "hot" air, characterized in that the inlet cavity is divided into two cavities: the first and the second, the swirling apparatus, respectively, into two parts, to the inlet of which is connected by several supply pipelines with shut-off valves, and to the outlet there is a centrifugal compressor located between the turbine disk and the deflector. 5. A device for cooling the rotor blade of a turbine of a by-pass gas turbine engine according to claim 4, characterized in that the first and second cavities are divided into several sectors and lead pipelines are connected to each sector. 6. A device for cooling the rotor blade of a turbine of a by-pass gas turbine engine according to claim 4 or 5, characterized in that each rotor blade has two cavities: front and rear, with perforations of the leading edge being brought out into the front cavity.

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