RU2769743C1 - Method for cooling the buckets of the turbine of a gas turbine engine and apparatus for implementation thereof - Google Patents

Abstract

FIELD: engine building.

SUBSTANCE: invention relates to high-temperature turbines of gas turbine engines, namely, to methods for cooling the buckets of turbines of gas turbine engines for various purposes. The air intended for cooling the buckets 10 in the impeller of the turbine 2 is collected from the air path 18 behind the rotor of the compressor 1 through inputs 17 in the blades 12 of the rectifying apparatus 6 and supplied into the air channels 14 of the heat exchange modules 11 located in the blades 12 of the rectifying apparatus 6 of the compressor, and through the outputs 19 into the twisting apparatus 8. Fuel from the fuel collector 5 is supplied to the fuel channels 13 of the heat exchange modules 11 located in the blades 12 of the rectifying apparatus 6 of the compressor through the inputs 15, and to the fuel injectors 7 of the combustion chamber 4 installed in the body 3 through the outputs 16. The air cooled in the heat exchange modules 11 is supplied to the internal cavities of the buckets 10 of the turbine from the twisting apparatus 8 through channels 9 in the impeller 2.

EFFECT: increase in the efficiency of the gas turbine engine is achieved by maintaining high efficiency of the thermodynamic cycle thereof.

2 cl, 3 dwg

Description

The invention relates to high-temperature turbines of gas turbine engines, and in particular to methods for cooling the working blades of turbines of gas turbine engines for various purposes.

Significant progress in the field of increasing the specific power and efficiency of modern gas turbine engines (GTE) has been achieved due to the improvement of their thermodynamic cycle based on an increase in the pressure level of the air entering the combustion chamber and an increase in the gas temperature at the turbine inlet.

However, an increase in the gas temperature at the turbine inlet requires intensification of the cooling of the turbine blades. For their cooling, a part of the air flow after the compressor is usually taken. In a gas turbine engine with a high pressure ratio in the compressor, the air temperature at the outlet of the compressor increases significantly, which requires an increase in the air flow required to cool the turbine, leading to a decrease in the efficiency of the thermodynamic cycle. To eliminate this disadvantage, pre-cooling of the air intended for cooling the turbine is used in heat exchangers using various cooling media.

Known method and device for cooling the turbine. The method provides for supplying cooling air to hot components in a gas turbine engine, and includes taking part of the air in the compressor in front of the combustion chamber in the form of a separate flow, cooling it and directing the cooled flow to hot engine elements for cooling them. The gas turbine engine implementing this method contains a compressor with a plurality of stages of rotor blades and stator blades for compressing air, a combustion chamber and a turbine, and also contains a fan located upstream of said compressor and having a bypass channel in which an air-to-air heat exchanger is located cooling the air flow taken after the compressor to cool the hot engine elements (US Patent No. 5581996, published 08/16/1995).

Closest to the proposed technical solution is a method of cooling the working blades of a high-pressure turbine of a bypass gas turbine turbojet engine and a device for its implementation. The method includes taking cooling air after the compressor from the air cavity of the combustion chamber, transporting it through an air-to-air heat exchanger installed in the air path of the secondary circuit, to the swirl apparatus, followed by supplying cooling air to the internal cavities of the rotor blades through the air channels in the turbine wheel. The device for implementing the method contains an air-to-air heat exchanger located in the second circuit, connected by its inlet to the air cavity of the combustion chamber, and by its outlet to the stator spin apparatus, communicated through air channels in the impeller with the internal cavities of the high-pressure turbine rotor blades (Patent RU 2387846 F01D 5/18, published 04/27/2010).

Known methods and devices for their implementation have a number of disadvantages:

, и определяется по зависимости:As you know, the efficiency of the engine working cycle (thermal efficiency), showing what part of the heat supplied in the cycle is converted into useful work, depends only on the level of air pressure when heat is supplied to it in the engine combustion chamber, determined by the degree of increase in air pressure in its compressor

, and is determined by the dependence:

,

,

- термический КПД,where

- thermal efficiency,

- степень повышения давления воздуха в компрессоре,

- the degree of increase in air pressure in the compressor,

k - adiabatic index, for air k=1.4

(see "Theory and calculation of air-breathing engines" under the editorship of S.M. Shlyakhtenko, M., Mashinostroenie, 1987).

The above known methods of cooling the rotor blades of a high-pressure turbine of a bypass turbojet engine are characterized by the removal of heat from part of the air flow in the primary circuit (after the high-pressure compressor) with a high level of its pressure, and the supply of this heat to the air flow in the second circuit with a significantly lower pressure level (by 4...7 times) in comparison with the air pressure in the primary circuit, which leads to a decrease in the thermodynamic efficiency of the engine operating cycle and its efficiency.

Devices for implementing known methods for cooling high-pressure turbine rotor blades have air-to-air heat exchangers in the second engine circuit for pre-cooling air intended for cooling high-pressure turbine rotor blades. Such an arrangement of heat exchangers leads to increased hydraulic losses in the tract of the second circuit of the engine due to its clutter and to a decrease in the efficiency of the engine.

Hydraulic losses in the communications for supplying cooling air to the heat exchanger and removing it from it reduce the pressure level and lead to the need to compensate for it by increasing the cooling air flow to ensure the required cooling efficiency of the high-pressure turbine rotor blades, which leads to a decrease in engine efficiency.

The objective of the invention is to increase the efficiency of a gas turbine engine by maintaining the high efficiency of its thermodynamic cycle by optimizing the cooling of the turbine blades.

The essence of the invention lies in the fact that the method of cooling the turbine blades of a gas turbine engine includes the selection of cooling air from the air path of the compressor, its preliminary cooling in heat exchange modules located in the compressor straightener blades with fuel entering the combustion chamber, the supply of cooled air to the spin apparatus with its subsequent supply to the internal cavities of the rotor blades through the air channels in the turbine wheel.

The problem is also solved by a device for cooling the turbine blades of a gas turbine engine, which contains heat exchange modules with fuel and air channels located in the compressor rectifier blades, the inputs of the fuel channels are connected to the fuel manifold, and their outputs are connected to the fuel injectors of the combustion chamber, the inlets of the air channels communicated with the air path of the compressor, and their outputs - with the swirl device, communicated through the air channels in the turbine wheel with the internal cavities of its rotor blades.

Preservation of the thermal stability of the heated fuel at the outlet of the heat exchange modules is confirmed by the following estimated calculation, the initial data for which are taken on the basis of publicly available information (UDK 665.73/75, A.I. Lanshin, A.A. Tsereteli “Results of an experimental study of a high-temperature gas generator in order to create a promising version of the engine with T _g ^* 1800K", Engine, No. 4, 2017, www.aviaport.ru) for the conditional air flow at the outlet of the high-pressure compressor G _B = 1 kg / s.

1. Initial data:

1.1. The total air pressure ratio in the compressor

π _kΣ \u003d 25

1.2. The amount of air required to burn 1 kg of fuel

L _o =15 kg.

1.3. Fuel parameters (RT according to GOST 10227-2013) in the fuel manifold:

- temperature t _T \u003d 100 ° C,

- pressure P _T \u003d 9 MPa (90 kgf / cm ² ),

- heat capacity С=2.38 kJ/kg/deg.

1.4. Excess air ratio in the combustion chamber

α _{in c / s} \u003d 2.5

1.5. Relative air flow rate for cooling the turbine blades on the air flow rate at the compressor outlet

g _{in cool} = 0.04 (4%).

1.6. Required cooling air temperature reduction

Δt _B =200°C.

2. According to known dependencies:

2.1. Air consumption for cooling turbine blades

G _cool =G _{B ⋅g} cool =1⋅0.04=0.04 kg/s.

2.2. Air temperature behind the compressor:

and the corresponding heat capacity С _р =1.04 kJ/kg/deg.

2.3. Fuel consumption

G _t \u003d (G _in -G _{in cool} ) / α / L _o \u003d (1-0.04) / 2.5 / 15 \u003d 0.0256 kg / s.

2.4. Heat flow into the fuel from the cooled air

Q _t \u003d Q _in \u003d C _p ⋅G _in cool ⋅Δt _in \u003d 1.04 0.04 200 \u003d 8.32 kJ / s.

2.5. Fuel heating in the heat exchanger

Δt _t \u003d Q _t / C / G _t \u003d 8.32 / 2.38 / 0.0256 \u003d 137 ° C.

2.6. Fuel temperature at the outlet of the heat exchanger

t _{t out \u003d} t _t +Δt _t \u003d 100 + 137 \u003d 237 ° С.

2.7. Saturated vapor pressure of the fuel at this temperature

(t _{t out} =237°C) P _{sat t} =1850 mm Hg (2.5 kgf/cm ² ) significantly lower fuel pressure (P _t =90 kgf/cm ² ), which indicates its thermal stability.

In FIG. 1 shows a schematic diagram of the cooling system for the turbine blades of a gas turbine engine; in fig. 2 - blade straightener compressor with heat exchange module; in fig. 3 - cross-section of the blades of the directing apparatus of the compressor with the heat exchange module.

A gas turbine engine with a device that implements the proposed method of cooling the turbine blades contains a compressor rotor 1, a turbine impeller 2, a housing 3 with a combustion chamber 4 and a fuel manifold 5. A vane straightener 6 is placed behind the compressor rotor 1. The combustion chamber 4 has fuel injectors 7. The swirling apparatus 8 is connected by means of air channels 9 in the turbine wheel 2 with the internal cavities of the rotor blades 10. The heat exchange modules 11 are located in the blades 12 of the straightener 6 of the compressor 1 and have fuel 13 and air channels 14. The inputs 15 of the fuel channels 13 are in communication with fuel manifold 5, and their outputs 16 - with fuel injectors 7 of the combustion chamber 4. The inputs 17 of the air channels 14 are connected with the air path 18 of the compressor, and their outputs 19 - with the spin apparatus 8, which communicates through the air channels 9 in the turbine wheel 2 with internal cavities of its rotor blades 10. Air channels 14 and fuel channels 13 section us walls 20.

The method is carried out using the device as follows:

The air intended for cooling the rotor blades 10 in the turbine wheel 2 is taken from the air path 18 behind the compressor rotor 1 through the inlets 17 in the blades 12 of the straightener 6 and is fed into the air channels 14 of the heat exchange modules 11 located in the blades 12 of the straightener 6 of the compressor , and through the outputs 19 to the spin apparatus 8.

Fuel from the fuel manifold 5 through the inlets 15 is fed into the fuel channels 13 of the heat exchange modules 11 located in the blades 12 of the straightener 6 of the compressor, and through the outlets 16 into the fuel injectors 7 of the combustion chamber 4 installed in the housing 3.

Through the walls 20 between the fuel channels 13 and the air channels 14 of the heat exchange modules 11 located in the blades 12 of the directing vane 6 of the compressor, heat is exchanged with heat removed from the air intended for cooling the working blades 10 of the turbine, and its corresponding supply to the fuel intended for supplying into the combustion chamber 4. Accordingly, the air temperature decreases, and the fuel temperature rises.

The air cooled in the heat exchange modules 11 is supplied from the swirler 8 through channels 9 in the impeller 2 into the internal cavities of the turbine blades 10.

The fuel heated in the heat exchange modules 11 with the help of nozzles 7 is supplied with spray into the combustion chamber 4, where, during combustion, it transfers to the working fluid of the thermodynamic cycle of the engine (gas-air mixture), along with the heat released during its combustion reaction, the heat removed into it from the air, designed to cool the turbine blades.

Thus, the cooling of the turbine blades is carried out with the preliminary cooling of the air taken for this purpose from the air path of the compressor, with the return of heat to the combustion chamber with fuel at a high level of air pressure in it, while maintaining the efficiency of the thermodynamic cycle of the gas turbine engine and increasing its efficiency.

Claims (2)

1. A method for cooling the working blades of a turbine of a gas turbine engine, characterized in that it includes taking cooling air from the air path of the compressor, pre-cooling it in heat exchange modules located in the blades of the compressor straightener with fuel entering the combustion chamber, supplying cooled air to the spin apparatus with its subsequent supply to the internal cavities of the rotor blades through the air channels in the turbine wheel. 2. A device for cooling the working blades of a turbine of a gas turbine engine, characterized in that it contains heat exchange modules with fuel and air channels located in the blades of the directing vane of the compressor, wherein the inlets of the fuel channels are connected to the fuel manifold, and the outlets are connected to the fuel injectors of the combustion chamber, the inlets of the air the channels are in communication with the air path of the compressor, and the outlets are in communication with the swirling device, which is connected through the air channels in the turbine impeller with the internal cavities of its rotor blades.

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