RU2726861C1 - Gas turbine engine operating method and gas turbine engine - Google Patents

Abstract

FIELD: machine building.SUBSTANCE: method of gas turbine engine operation, in which air, at least from one compressor, enters the first combustion chamber acting behind the compressor unit, the first turbine acting behind the first combustion chamber, the second combustion chamber acting behind the first turbine, and a second turbine, acting after the second combustion chamber, which is self-igniting, wherein in the compressor stages a cycle is realized, in which atmospheric air is compressed by the multistage cooled downstream compressor to a predetermined pressure, then at the outlet of the last stage of the compressor the air is heated, without mixing in the heat exchanger, with exhaust gases supplied from the turbine, then it is reheated in the combustion chamber, into which fuel is supplied and then through the fixed nozzle device, which forms the specified field of working gas flow rates and is directed to the turbine stage inlet, in which starting from the first and including the penultimate service, before each stage, fuel is supplied through radially directed to the turbine shaft and the units of nozzles fixed on the turbine housing and is uniformly sprayed on the flow section of the turbine in an amount which provides heating of the working gases to a predetermined temperature, said working gas rotating at an angular speed of rotation of the turbine shaft, then successively enters into each stage through rigidly installed on the shaft of the turbine cellular nozzle units, in which, due to geometrical effect, increased component of gas velocity in axial direction relative to turbine shaft and further along working nozzle outlet working gas flows over turbine blade aerodynamic profiles installed on shaft in radial direction, which rotate together with nozzle device and which transmit to said shaft mechanical energy in amount, equal to heat energy received from fuel combustion at the beginning of this stage, which provides rotation of turbine with equal amount of used mechanical energy and added heat energy in stages from fuel combustion in flow of working gas of nozzles supplied by units, installed before entrance to stage, and which maintain preset operating parameters with isothermality coefficient of not less than 0.95, and then working gas enters the last stage, in which without performing heating in the beginning of the stage, final expansion of working gas and lower temperature at the outlet, and then the working gas is directed to the heat exchanger, in which heat exchange is performed between it and the compressed air flow in the compressor, and as specified above, is directed to the heating chamber, and the said working gas is directed further into the atmosphere by closing the cycle.EFFECT: creation of such a gas-turbine engine, in which a process with a thermodynamic cycle is realized, which differs by efficiency from the Carnot isothermal cycle by less than 10 %, which enables to achieve high efficiency close to the Carnot cycle.14 cl, 8 dwg

Description

The technical solution relates to continuous combustion gas turbine engines in an open circuit on high-calorific gas turbine fuels. It can be used in transport plants, power plants and gas pumping units.

Known Ukrainian patent for invention No. 73660, published on August 15, 2005, IPC index F02C 7/08, which describes a method for increasing the thermal efficiency of the cycle of a gas turbine plant, which consists in the fact that the air in front of the combustion chamber is subjected to regenerative heating using gases exhaust in the turbine, and the gases after expansion in the high-pressure turbine are used for regenerative heating of air, and then expanded in the low-pressure turbine.

The disadvantages are that from the diagrams of the proposed cycle given in patent No. 73660 by the authors, it follows that the authors' proposal unambiguously leads to a decrease in useful work and to an increase in costs. The first is noted by the authors themselves, and the second is evident from elementary thermodynamics, where the flow rates are defined as the area under the lower broken line of the cycle. Therefore, this proposal only leads to a decrease in efficiency.

A known method for increasing the efficiency of a multistage axial compressor described in the patent of the Russian Federation No. 2529289, published on September 27, 2014, IPC indices: F04D 29/58; F04C 7/143; F02C 3/30, according to which the efficiency of the axial multistage compressor is increased by water injection, moreover, water is supplied to the air flow through calibrated outlet channels made on the surface of the guide vanes blades, while water injection is carried out at a saturation temperature corresponding to the sum of the local pressure and differential pressure in said outlet channels, and water injection begins in the compressor stages, where the temperature of the medium becomes higher than the saturation temperature of water at the local pressure in the compressor stages.

Water is supplied to the air flow through calibrated outlet channels made with the possibility of providing uninterrupted water flow and air flow, and the number of outlet channels and the sizes of their flow cross-sections are selected from the condition of uniform distribution of water concentration along the height of the blades.

In the patent of the Russian Federation No. 2529289, the disadvantages are that air cooling does not begin from the first stage, but from the seventh, eighth, and this entails a significant increase in energy consumption, and hence a decrease in efficiency. The compressor proposed for use in the proposed cycle is a series of well-known centrifugal stages connected in series, mounted on a common shaft, each of which consists of a rotating wheel with an axial inlet and a peripheral outlet, a scroll stationary relative to the housing with diffuser-mounted profiles.

In the first, the air is accelerated due to the rotation of the impeller as a result of the application of an external mechanical force to it through the compressor shaft, and in the second, when the diffuser channels pass, the air is adiabatically compressed with the release of heat, that is, with heating. In contrast to the known ones, in our version, in each stage between the outlet from the previous one and the axial inlet to the next, a heat exchanger is installed in which the compressed air is cooled to a temperature equal to the temperature of the air at the inlet to the previous stage in our diagram (Fig. 1) this is shown as an isobar with a falling temperature in the gap, which connects the exit from the previous degree with the entrance to the next.

This method of constructing a compressor ensures maximum efficiency, that is, minimum losses of mechanical energy and, accordingly, with the required minimum amount of heat that must be removed from the compressor. In the known compressor, the additivity principle is inactive, that is, the amount of heat removed from two series-connected stages with cooling only after the second stage is always greater than the amount of heat removed from each of the two stages cooled in stages.

The closest is the gas turbine unit and the method of its operation, described in the patent of the Russian Federation No. 2137935, which was published on September 20, 1999, the IPC index F02C 3/14 and is selected as a prototype of the method.

A gas turbine plant contains a compressor unit consisting of at least one compressor, a first combustion chamber acting after the compressor unit, a first turbine acting after the first combustion chamber, a second combustion chamber acting after the first turbine, a second turbine acting after the second combustion chamber , and the first and second combustion chambers have an annular configuration, moreover, the second combustion chamber is made in the form of a self-igniting combustion chamber equipped with swirlers.

The compressor unit contains two compressors interacting with an intercooler.

A diffuser is installed in front of the second combustion chamber.

The swirlers in the second combustion chamber are provided with a surface on the downstream side, generally radial to the wall of the combustion chamber.

The main fuel is injected into the second combustion chamber by swirlers. The waste fuel is injected through the fuel injectors, with the fuel injectors positioned around the circumference of the second combustion chamber.

The first combustion chamber is adapted to operate with premix burners.

The blades are mounted on a common rotor shaft. The rotor shaft is supported by two bearings.

The first combustion chamber and the second combustion chamber are annular. The annular configuration is made in the form of a plurality of separate combustion chambers located around the rotor shaft.

A method of operating a gas turbine plant containing a compressor unit consisting of at least one compressor, a first combustion chamber acting after the compressor unit, a first turbine acting after the first chamber, a second combustion chamber acting after the first turbine, and a second turbine acting after the second combustion chamber, moreover, the expansion of hot gases in the first turbine is reduced to such an extent that these partially expanded hot gases enter the second combustion chamber installed downstream of the second turbine with a temperature higher than the autoignition temperature of the main fuel injected into it, while the hot gases swirl into the second combustion chamber.

The partially expanded hot gases have an average velocity of> 60 m / s in the second combustion chamber.

Provide auxiliary measures that ensure self-ignition in the second combustion chamber also in the case when a change in the temperature of the gases in the fuel injection zone should be established.

The exhaust gases from the second turbine are used in the steam loop.

The disadvantages are that the gas turbine plant (further GTU) under the patent has a higher specific power and only 1-3% of its efficiency (further efficiency) is higher than the known ones. However, the patent lacks data that show a closed thermodynamic cycle, the efficiency of which is the efficiency of a gas turbine. For specialists, it is clear that intermediate heating between the stages of the turbine leads to an increase in the temperature of the working gases at the outlet, but this, in turn, requires additional measures to reduce heat losses in the cycle to maintain high efficiency.

From the data given in the description, it follows that the heating of the working gas is carried out between the first and second stages of a turbine of a classical design with adiabatic expansion, that is, with a pressure ratio at the inlet and outlet equal to from 2 to 3. And this, in turn, requires rather high costs fuel to restore the temperature to the level at the inlet to the first stage and, accordingly, leads to significant blocking of the flow path between the stages of structures that provide reliable fuel supply and reliable efficient combustion in the working gas flow, with the condition of uniform heating over the entire cross section, but this leads to increase in losses in the turbine, and gives reason to doubt the statements of the authors about the increase in efficiency. At the same time, an increase in the specific power with a constant efficiency is not a new useful solution, since this is achieved by a simple increase in the air flow rate, at which the efficiency is increased by 2-5% due to the improvement of the hydrodynamics of the flow path, which is known and tested on existing traditional GTU.

The general essential features are that a method of operation of a gas turbine engine is provided in which air from at least one compressor enters the first combustion chamber acting after the compressor unit, the first turbine acting after the first combustion chamber, the second combustion chamber acting after the first turbine, and a second turbine operating behind the second combustion chamber, which ignites spontaneously.

The technical result of the proposed method of operating a gas turbine engine is the creation of such a gas turbine engine, which implements a process with a thermodynamic cycle, which differs in efficiency from the isothermal Carnot cycle by less than 10%, which makes it possible to achieve high efficiency close to the Carnot cycle.

The essential features are that the method of operation of a gas turbine engine in which air from at least one compressor enters the first combustion chamber, acting after the compressor unit, the first turbine, acting after the first combustion chamber, the second combustion chamber, acting after the first turbine, and a second turbine, acting behind the second combustion chamber, which self-ignites, and, in the compressor stages, a cycle is implemented in which atmospheric air is compressed by a multistage cooled step-by-step compressor to a predetermined pressure, then at the outlet of the last compressor stage the air is heated without mixing in the heat exchanger with exhaust gases supplied from the turbine, then heats up in the combustion chamber, into which the fuel enters and then through a stationary nozzle apparatus, which forms a predetermined field of working gas flow velocities and is directed to the entrance to the turbine stages, in which, starting from the first and including the penultimate stage, before every in the next stage, fuel is supplied through the nozzle blocks radially directed to the turbine shaft and fixedly mounted on the turbine housing and is sprayed evenly over the section of the turbine flow path in an amount that provides heating of the working gases to a predetermined temperature, the specified working gas, rotating at the angular speed of rotation of the turbine shaft, then it enters successively into each stage through honeycomb nozzle devices rigidly mounted on the turbine shaft, in which, due to the geometric effect, the gas velocity component increases in the axial direction with respect to the turbine shaft and further downstream from the nozzle device, the working gas flows around the aerodynamic profiles turbine blades mounted on the shaft in the radial direction, which rotate together with the nozzle apparatus and which transfer to the specified shaft mechanical energy in an amount equal to the thermal energy obtained from fuel combustion at the beginning of this stage, which ensures the rotation of the ribs, with an equal amount of used mechanical energy and added thermal energy in the stages from fuel combustion in the flow of the working gas supplied by the injector blocks installed in front of the entrance to the stage, and which maintain the specified operating parameters with an isothermal coefficient of at least 0.95, and then the working gas enters the last stage, in which, without performing preheating at the beginning of the stage, the final expansion of the working gas and a decrease in the outlet temperature are performed, and then the working gas is directed to the heat exchanger in which heat exchange is carried out between it and the compressed air stream in the compressor, and as indicated above , is directed to the heating chamber, and the specified working gas is directed further into the atmosphere, closing the cycle.

The amount of thermal energy supplied by the combustion of fuel from each block of injectors before the stages is set taking into account the equality of the corresponding temperatures in all stages of the isothermal part of the turbine with the deviation of the corresponding temperatures from the calculated ones with an accuracy of 0.5% to 1%.

The pressure drop at each compressor stage and the amount of thermal energy removed from the compressed air in each stage are in values that ensure the equality of the corresponding inlet and outlet temperatures in each stage with a deviation from 0 5% to 1%, except for the first, in which the inlet temperature is determined ambient temperature.

The ratio of the pressure at the inlet to the pressure at the outlet of each isothermal part of the turbine operating with heating from_1.05 to 1.35 with the same isothermal coefficient, and in the final stage, with adiabatic expansion, not more than 3.

In the injector blocks, sprayed fuel is mixed with the working gases of the turbine through a system of channels to form a rich fuel mixture inside the fairings of the injector block, and the resulting fuel mixture is fed to the trailing edge of the fairing profile of the injector block, in which the resulting fuel mixture is distributed in volumes directly proportional to the distance from the axis shaft, with the ability to ensure uniform spraying over the section of the turbine flow cavity, self-ignites and enters the flow path of each stage of the isothermal turbine part operating with heating.

Distinctive essential features valid in all cases are that a cycle is implemented in the compressor stages in which atmospheric air is compressed by a multistage cooled infinitely variable compressor to a predetermined pressure, then at the outlet of the last compressor stage the air is heated without mixing in the heat exchanger by exhaust gases supplied from turbine, then it is added to the combustion chamber, into which fuel enters and then through a stationary nozzle apparatus in which a given field of working gas flow velocities is formed and directed to the entrance to the turbine stages, in which, starting with the first and including the penultimate stage, before each stage, fuel through the nozzle blocks radially directed to the turbine shaft and fixedly mounted on the turbine housing and is sprayed uniformly over the section of the turbine flow path in an amount that provides heating of the working gases to a predetermined temperature, the specified working gas rotating at an angular velocity rotation of the turbine shaft, then enters successively into each stage through the honeycomb nozzle devices rigidly mounted on the turbine shaft, in which, due to the geometric effect, the component of the gas velocity increases in the axial direction with respect to the turbine shaft and further downstream from the nozzle apparatus gas flows around the aerodynamic profiles of the turbine blades mounted on the shaft in the radial direction, which rotate together with the nozzle apparatus and which transfer to the specified shaft mechanical energy in an amount equal to the thermal energy obtained from fuel combustion at the beginning of this stage, which ensures the rotation of the turbine, with an equal amount used mechanical energy and added thermal energy in the stages from fuel combustion in the flow of working gas supplied by the injector blocks installed in front of the inlet to the stage and which maintain the specified operating parameters with an isothermal coefficient of at least 0.95, and then the gas enters the last stage, in which, without heating at the beginning of the stage, the final expansion of the working gas and a decrease in the outlet temperature are carried out, and then the working gas is directed to the heat exchanger, in which heat exchange is carried out between it and the compressed air stream in the compressor, and as indicated above, it is directed to the reheating chamber, and the specified working gas is directed further into the atmosphere, closing the cycle.

Distinctive essential features valid in some cases is that the amount of thermal energy supplied by the combustion of fuel from each block of nozzles in front of the stages is set taking into account the equality of the corresponding temperatures in all stages of the isothermal part of the turbine with a deviation of the corresponding temperatures from the calculated ones with an accuracy of 0.5% to 1%.

The pressure drop at each compressor stage and the amount of thermal energy removed from the compressed air in each stage are in values that ensure the equality of the corresponding inlet and outlet temperatures in each stage with a deviation from 0.5% to 1%, except for the first, in which the inlet temperature determined by the ambient temperature.

The ratio of the inlet pressure to the outlet pressure of each stage of the isothermal part of the turbine is from 1.05 to 1.35 with the same isothermal coefficient, and in the final stage, with adiabatic expansion, no more than 3.

In the injector blocks, sprayed fuel is mixed with the working gases of the turbine through a system of channels to form a rich fuel mixture inside the fairings of the injector block, and the resulting fuel mixture is fed to the trailing edge of the fairing profile of the injector block, in which the resulting fuel mixture is distributed in volumes directly proportional to the distance from the axis shaft, with the ability to ensure uniform spraying over the section of the turbine flow cavity, self-ignites and enters the flow path of each stage of the isothermal part of the turbine.

The known design is described in the patent of Ukraine for invention No. 103413 published on 10.10.2013, IPC index F02C 3/36, F02C 3/14, F02K 3/02.

The known gas turbine engine contains at least a low-pressure compressor, a high-pressure compressor with a high-pressure turbine, a high-pressure combustion chamber with a fuel supply device located between them, a low-pressure combustion chamber with a fuel supply device, and is configured to bypass a part of the working fluid between the outlet of the low-pressure compressor past the high-pressure compressor, the high-pressure combustion chamber, the high-pressure turbine, and the entrance to the turbine stages of the corresponding low-pressure, directly or through the low-pressure combustion chamber, with free flow and free redistribution of mass flow rates between both streams of the working fluid, the outlet of the low-pressure combustion chamber is connected to the inlet of the said stage of the corresponding low-pressure turbine, moreover, bypassing a part of the working fluid is possible only in one direction - from the compressor to the turbine.

The low-pressure compressor, the high-pressure compressor, the high-pressure turbine and the low-pressure turbine are made on the same shaft, or the low-pressure compressor and the low-pressure turbine are made on the same shaft, and the high-pressure compressor and the high-pressure turbine are made on the other shaft.

The bypass of a part of the working fluid in only one direction is carried out with the help of self-acting flaps, which prevent the movement of the working fluid in the direction from the turbine to the compressor and open the bypass channel when the working fluid moves in the direction from the compressor to the turbine.

The inlet of the low pressure combustion chamber is connected to the outlet of the low pressure compressor.

The low pressure combustion chamber is located between the stages of the low pressure turbine, the outlet of the low pressure compressor is connected directly to the inlet of the said stage of the corresponding low pressure turbine.

The low pressure combustion chamber is located between the stages of the low pressure turbine, the inlet of the low pressure combustion chamber is also connected to the outlet of the low pressure compressor.

The disadvantages are that the authors propose a technical solution for a gas turbine engine in which the process of fuel combustion with a heat of combustion of 43,000 kJ / kg and expansion of working gases with the production of external work occurs in a stepwise manner with different mass. In this case, before the first stage, the stoichiometric coefficient is about 3, and before the last stage, about 1, and the exhaust gases from the engine have a temperature of 1000 ° K. The authors claim that the efficiency reaches 80%.

From the above, it follows that in the proposed gas turbine engine, the specific power on the turbine shaft per unit of exhaust gas consumption must be at least 3-4 MW, which even at 2000 ° K requires an unrealistically high pressure. It follows from this that in this case, the statement about its performance and the more any efficiency is incorrect.

The closest in design is the well-known patent of the Russian Federation No. 2531110 for a gas turbine installation, which contains nozzle blades, published 01/10/2012 in bull. No. 1, IPC F02C 3/14, F23R 3/30.

The gas turbine unit contains a compressor configured to receive and compress a working fluid, a combustion chamber configured to receive a compressed working fluid from the compressor and fuel, and to burn a mixture of compressed working fluid and fuel to form exhaust gas, and a turbine having a first section and a second section and configured to receive exhaust gas from the combustion chamber and use it to rotate the shaft, while an annular combustion device for reheating is located between the first and second sections of the turbine, which contains a pre-mixing nozzle blade made with the possibility of mixing air and fuel with the creation of an air-fuel mixture and with the possibility of introducing this mixture into the exhaust gas coming from the first section of the turbine.

Said annular combustion device comprises an annular channel having an inner annular wall, an outer annular wall and the first and second walls of the inner part located between the inner annular wall and the outer annular wall, and the specified nozzle blade is held in place by the outer annular wall and the first and second walls the inner part of the annular channel. The specified blade-nozzle has a cylindrical part and a part with an aerodynamic profile. The specified cylindrical part of the blade-nozzle has an air inlet designed to receive air from the external air duct located between the outer annular wall and the first wall of the inner part of the annular channel or the specified part with the aerodynamic profile of the blade-nozzle has an air intake hole designed to receive air from the inner air duct located between the inner annular wall and the second wall of the inner part of the annular channel.

The specified blade-nozzle on the first side side of the aerofoil part has a first group of outlets, and on the second side of the aerofoil part opposite to the specified first side side has a second group of outlets, and the said first and second groups of outlets are intended for introducing an air-fuel mixture into the exhaust gas coming from the first section of the turbine.

Installation containing an annular combustion device for reheating, configured to be placed between the first and second sections of the turbine of a gas turbine plant and containing an annular channel having an inner annular wall, an outer annular wall and first and second walls of the inner part located between the inner and outer annular walls , and a pre-mixing nozzle blade adapted to mix air and fuel to create an air-fuel mixture and to introduce this mixture into the exhaust gas stream leaving the first section of the turbine.

Said vane-nozzle is held in place by the outer annular wall and also by the first and second walls of the inner part of the annular channel. The annular channel has a first fixing hole located in the first wall of the inner part, a second fixing hole located in the second wall of the inner part, a third fixing hole located in the outer annular wall, a first sliding annular seal located in the first fixing hole, and a second sliding annular a seal located in a second mounting hole, said vane-nozzle being held in place by a first sliding O-ring in a first mounting hole, a second sliding O-ring in a second mounting hole, and a third mounting hole.

The specified blade-nozzle has a cylindrical part and a part with an aerodynamic profile. Said cylindrical part of the nozzle blade has an air inlet for receiving air from an external air duct located between the outer annular wall and the first wall of the inner part of the annular channel. The specified part with the aerodynamic profile of the blade-nozzle has an internal volume of a substantially triangular shape, intended for receiving air from the external air duct and located on the side of the trailing edge of the specified part with the aerodynamic profile of the blade-nozzle.

The specified part with the aerodynamic profile of the nozzle blade has an air inlet for receiving air from the internal air duct located between the inner annular wall and the second wall of the inner part of the annular channel. The specified part with the aerodynamic profile of the blade-nozzle has an internal volume intended for receiving air from the internal air duct, and the specified essentially annular internal volume is located on the side of the leading edge of the part with the aerodynamic profile of the blade-nozzle.

The specified cylindrical part of the nozzle blade has a fuel supply opening for receiving fuel. The specified vane-nozzle has a cylindrical internal volume intended for receiving fuel and located between the first internal volume for air located on the side of the leading edge of the specified vane-nozzle, and the second internal volume for air located on the side of its trailing edge.

The specified blade-nozzle on the first side side of the aerofoil part has a first group of outlets, and on the second side of the aerofoil part opposite to the specified first side side has a second group of outlets, and the said first and second groups of outlet openings are intended for introducing an air-fuel mixture into the exhaust gas coming from the combustion chamber.

Installation containing an annular channel having an inner annular wall and an outer annular wall, and nozzle blades made with the possibility of mixing air and fuel with the creation of an air-fuel mixture and introducing this mixture into the central chamber located between the inner and outer annular walls.

These nozzle blades are held in place by the outer annular wall of the annular channel.

Each of the specified nozzle blades contains a cylindrical part and a part with an aerodynamic profile.

The disadvantages are that the gas turbine unit (further GTU) according to the prototype patent has a higher specific power and its efficiency (further efficiency) is higher by 1-3% than the known ones. However, the patent does not contain data that confirm a closed thermodynamic cycle, the efficiency of which is the efficiency of a gas turbine plant. It is understandable for specialists that intermediate heating between turbine stages increases the temperature of the working gases at the outlet, but this, in turn, requires additional measures to reduce heat losses in the cycle to maintain efficiency. From the data given in the description, it turns out that the heating of the working gas is carried out between the first and second stages of a turbine of a classical design, that is, with a difference from 2 to 3 atmospheres.

And this, in turn, requires a sufficiently large fuel consumption to restore the temperature to the level at the inlet to the first stage and, accordingly, leads to the need for a much more complex and cumbersome design in the flow path, which provide reliable fuel supply and reliable efficient combustion in the working gas flow , with the condition of uniform heating over the entire cross section, but this leads to an increase in losses in the turbine, and gives reason to doubt the authors' statements about the increase in efficiency. At the same time, an increase in the specific power with a constant efficiency is not a new useful solution, since it is achievable by a simple increase in the air flow rate, at which there is an increase in the efficiency by 2-5% due to the improvement of the hydrodynamics of the flow path, which is known and tested on existing traditional GTU.

It is obvious that such a one-time introduction of thermal energy, especially without taking into account the necessary pressure and temperature drops, as well as losses in a structure with a classical construction of turbine stages, does not provide any significant increase in efficiency.

The general essential features are that a gas turbine engine includes a compressor configured to receive and compress a working fluid, a combustion chamber configured to receive a compressed working fluid from the compressor and fuel, with the possibility of burning to form a working gas, and a turbine that has at least two stages and is configured to receive the working gas from the combustion chamber and use its energy to rotate the shaft, while at least one device for reheating is located between the turbine stages, which contain injector blocks.

The technical result of the proposed design of a gas turbine engine is the creation of such a gas turbine engine, which implements a process with a thermodynamic cycle, which differs in efficiency from the isothermal Carnot cycle by less than 10%.

The essential features are that a gas turbine engine includes a compressor configured to receive and compress a working fluid, a combustion chamber configured to receive a compressed working fluid from the compressor and fuel, with the possibility of burning to form a working gas, and a turbine that has at least two stages and configured to receive the working gas from the combustion chamber and use its energy to rotate the shaft, while at least one device for reheating is installed between the turbine stages, which contain nozzle blocks, and, in the working cavity, is installed a compressor cooled in stages, and a heat exchanger is installed behind the compressor in front of the main combustion chamber of reheating, and then in the flow path in the steps of the isothermal part of the turbine rotor, which has at least three stages, on the outer surface of the flow path of the engine casing at the inlet to each stage, radially directed blocks of fuel injectors, having in their outer cross-section fairings with an axisymmetric aerodynamic profile, the direction of the longitudinal chord of which coincides with the direction of the velocity vector of the incoming gas flow and have channels with the possibility of supplying a rich fuel mixture through holes in the rear part of the profile, and further in the axial direction after each block of nozzles, stages of the isothermal part of the turbine are installed, including honeycomb nozzles rigidly fixed on the shaft in each stage and behind them at least one row of turbine blades, then behind the stages of the isothermal part of the turbine, at least one stage of the turbine is installed, with the possibility of providing an adiabatic expansion of the working gas, the outlet of which is directed to the air heating heat exchanger after the compressor in front of the heating chambers due to the thermal energy of the turbine exhaust gases.

The turbine stages, with the ability to work in the mode of adiabatic expansion, are installed on a separate shaft.

The cells of the nozzle apparatus uniformly cover the flow path of the turbine stage, and each cell is formed by concentrically located shells and radially directed plates. Radial plates of the cells of the nozzle apparatus are made with a cross-section in the form of a symmetric aerodynamic profile.

The change in the cross-section along the length of the nozzle apparatus channels required to accelerate the gas flow is achieved by changing the thickness of the radially mounted plates, as well as by partially placing the aerodynamic profile of the turbine blades inside the nozzle apparatus channel with partial overlap in the axial direction.

On the outer surface of the turbine housing inside the flow path at the inlet to each stage, fuel injector blocks radially directed to the axial line of the shaft are installed, which are hollow rods with a blind end and radial holes forming channels for fuel supply, with the possibility of supplying fuel in specified quantities through lateral radial holes - nozzles in the walls of these rods, and on the rods, sections of independent fairings are hingedly mounted on the rods, with the possibility of rotation relative to the longitudinal axis of the rod within an angle of up to 80 ⁰ relative to the plane of the longitudinal axis of the turbine and with the possibility of matching the direction of the chord of the fairing profile with the vector oncoming flow velocities and have a system of channels connecting the inner volume of the fairing with the turbine flow path.

The engine has a linear arrangement in which all compressor and turbine stages are located along the same center line.

The initial part of the turbine with a reheating chamber includes the stages of the first isothermal part of the turbine serving the payload, installed along one centerline, and the second isothermal part of the turbine, together with the final adiabatic stages of the turbine and the compressor, have a shaft directed along the other axial line, rotated relative to the first at an angle up to 180 °, based on the requirements of the layout, moreover, both of these parts are connected by a gas duct with fixed nozzle devices installed in it, with the possibility of creating the necessary field of speeds of movement of the working gas, the compressor is installed on the side of the final stage of the turbine, the drive shafts of which are connected to the shaft of the specified part turbines.

In the compressor, except for the last stage, cooling heat exchangers are installed between the stages.

Distinctive essential features valid in all cases are that the gas turbine engine includes a compressor configured to receive and compress a working fluid, a combustion chamber configured to receive a compressed working fluid from the compressor and fuel, with the ability to burn to form a working gas, and a turbine, which has at least two stages and is configured to receive the working gas from the combustion chamber and use its energy to rotate the shaft, while at least one device for reheating is installed between the stages of the turbine, which contain blocks of nozzles, moreover, in the working cavity, a compressor cooled in stages is installed, and a heat exchanger is installed behind the compressor in front of the main combustion chamber of the reheating, and then in the flow path in the steps of the isothermal part of the turbine rotor, which has at least three stages, on the outer surface of the flow path of the engine housing at the entrance to each At the second stage, radially directed blocks of fuel injectors are installed, having in their outer cross-section fairings with an axisymmetric aerodynamic profile, the direction of the longitudinal chord of which coincides with the direction of the velocity vector of the incoming gas flow and have channels with the possibility of supplying a rich fuel mixture through the holes in the rear part of the profile, and further in the axial direction, after each block of nozzles, stages of the isothermal part of the turbine are installed, including honeycomb nozzles rigidly fixed on the shaft in each stage and behind them at least one row of turbine blades, further behind the steps of the isothermal part of the turbine, at least one stage of the turbine is installed , with the possibility of ensuring the adiabatic expansion of the working gas, the outlet of which is directed to the heat exchanger for heating the air after the compressor in front of the heating chambers due to the thermal energy of the turbine exhaust gases.

Distinctive essential features valid in individual cases are that the turbine stages, working with the ability to work in the mode of adiabatic expansion, are installed on a separate shaft.

The cells of the nozzle apparatus uniformly cover the flow path of the turbine stage, and each cell is formed by concentrically located shells and radially directed plates. Radial plates of the cells of the nozzle apparatus are made with a cross-section in the form of a symmetric aerodynamic profile.

The change in the cross-section along the length of the nozzle apparatus channels required to accelerate the gas flow is achieved by changing the thickness of the radially mounted plates, as well as by partially placing the aerodynamic profile of the turbine blades inside the nozzle apparatus channel with partial overlap in the axial direction.

On the outer surface of the turbine housing inside the flow path at the inlet to each stage, fuel injector blocks radially directed to the axial line of the shaft are installed, which are hollow rods with a blind end and radial holes forming channels for fuel supply, with the possibility of supplying fuel in specified quantities through lateral radial holes - nozzles in the walls of these rods, and on the rods, sections of independent fairings are hingedly mounted on the rods, with the possibility of rotation relative to the longitudinal axis of the rod within an angle of up to 80 ⁰ relative to the plane of the longitudinal axis of the turbine with the possibility of matching the direction of the chord of the fairing profile with the velocity vector oncoming flow and have a system of channels connecting the inner volume of the fairing with the turbine flow path.

The engine has a linear arrangement in which all compressor and turbine stages are located along the same center line.

The initial part of the turbine with a reheating chamber includes the stages of the first isothermal part of the turbine serving the payload, installed along one centerline, and the second isothermal part of the turbine, together with the final adiabatic stages of the turbine and the compressor, have a shaft directed along the other axial line, rotated relative to the first at an angle up to 180 °, based on the requirements of the layout, moreover, both of these parts are connected by a gas duct with fixed nozzle devices installed in it, with the possibility of creating the necessary field of speeds of movement of the working gas, the compressor is installed on the side of the final stage of the turbine, the drive shafts of which are connected to the shaft of the specified part turbines.

In the compressor, except for the last stage, cooling heat exchangers are installed between the stages.

Due to the fact that the design implements a method of operation of a gas turbine engine in which a cycle is implemented in the compressor stages in which atmospheric air is compressed by a multistage cooled compressor to a predetermined pressure, further at the outlet of the last compressor stage the air is heated, in an immiscible heat exchanger, by exhaust gases supplied from turbine, then heats up in the combustion chamber, into which fuel is supplied and then through a stationary nozzle apparatus forms a predetermined field of working gas flow velocities and is directed to the entrance to the turbine stages, in which, starting with the first and including the penultimate stage, fuel is supplied before each stage, through the nozzle blocks radially directed to the turbine shaft and fixedly mounted on the turbine housing and is sprayed uniformly over the section of the turbine flow path in an amount that provides heating of the working gases to the specified temperature, the specified working gas, rotating from the angular velocity of rotation the flow of the turbine shaft further comes sequentially into each of the feet through the honeycomb nozzle assemblies rigidly mounted on the turbine shaft, in which, as a result of geometric influence, there is an increase in the composite velocity of gas movement in the axial direction with respect to the turbine shaft and further downstream of the nozzle of the apparatus, the working gas flows around the aerodynamic profiles of the turbine blades mounted on the shaft in the radial direction, which rotate together with the nozzle apparatus and which transfer to the specified shaft mechanical energy in an amount equal to the thermal energy from fuel combustion at the beginning of this stage, which ensures the rotation of the turbine, with equal the amount of used and added thermal energy from the combustion of fuel in the working gas flow supplied by the injector blocks installed before entering this stage, and maintain the specified operating parameters with an isothermal coefficient of at least 0.95, and then the working gas enters the last stage in which, without in When heating is performed at the beginning of the stage, the final expansion of the working gas and a decrease in the outlet temperature are performed, and then the working gas is sent to the heat exchanger in which heat exchange is carried out between it and the compressed air stream in the compressor, which, as mentioned above, is sent to the heating chamber, and the specified the working gas is directed further into the atmosphere completing the cycle, and to accomplish this, the gas turbine engine includes a compressor configured to receive and compress the working fluid, a combustion chamber configured to receive the compressed working fluid from the compressor and fuel, with the possibility of burning to form a working gas, and a turbine, which has at least two stages and is configured to receive working gas from the combustion chamber and use its energy to rotate the shaft, while at least one device for reheating is installed between the turbine stages, which contain blocks of injectors, moreover, in the working floor spine, a compressor cooled in stages is installed, and a heat exchanger is installed behind the compressor in front of the main combustion chamber of reheating, and then in the flow path in the steps of the isothermal part of the turbine rotor, which has at least three stages, on the outer surface of the flow path of the engine casing at the inlet to each stage are installed radially directional blocks of fuel injectors, having in their outer cross section fairings with an axisymmetric aerodynamic profile, the direction of the longitudinal chord of which coincides with the direction of the velocity vector of the incoming gas flow and have channels with the possibility of supplying a rich fuel mixture through the holes in the rear part of the profile, and further in the axial direction, after each block of nozzles, stages of the isothermal part of the turbine are installed, including honeycomb nozzles rigidly fixed on the shaft in each stage and behind them at least one row of turbine blades, then behind the steps of the isothermal part of the turbine, installed at least more than one turbine stage, with the possibility of ensuring the adiabatic expansion of the working gas, the outlet of which is directed to the heat exchanger for heating the air after the compressor in front of the heating chambers due to the thermal energy of the turbine exhaust gases.

The proposed method and the created design of a gas turbine engine with the indicated essential features implement a process with a thermodynamic cycle, which differs in efficiency from the isothermal Carnot cycle by less than 10%, which makes it possible to achieve a high efficiency close to the Carnot cycle.

In FIG. 1 shows the thermodynamic cycle in coordinates T - S (temperature-entropy);

In FIG. 2 shows a block diagram of a gas turbine engine

In FIG. 3 shows a longitudinal section along the flow path of the turbine of a gas turbine engine;

In FIG. 4 shows a cross-section AA with elements of the nozzle blocks;

In FIG. 5 shows a cross-section BB along the block of nozzles;

In FIG. 6 shows a section B-B along the plates of the nozzle apparatus and the turbine blades;

In FIG. 7 shows a section Г-Г on which we see the location of the shells and radial plates of the nozzle apparatus in the flow path of the turbine;

In FIG. 8 shows a block diagram of a gas turbine engine with a modified layout within essential features.

The gas turbine engine includes a compressor 10, which is cooled in stages, and a heat exchanger 12 is installed behind the compressor in front of the main combustion chamber 11 of the preheating, and then (see Fig. 3) in the flow part 13 in the steps of the isothermal part of the turbine rotor 14, which has at least three stages, and in the shown case seven, on the outer surface of the engine flow path at the entrance to each stage (see Fig. 3 and Fig. 4), radially directed blocks 15 of fuel injectors are installed, which have fairings 16 in their outer cross-section which have asymmetric aerodynamic profile, the direction of the longitudinal chord of which coincides with the direction of the velocity vector of the incoming gas flow due to unfolding and have channels.

That is, on the outer surface of the turbine housing 17, inside the flow path at the inlet to each stage, fuel injector units 15 radially directed to the shaft centerline are installed, which contain hollow rods 18 with a blind end, which form channels (see Fig. 5) for supplying fuel , with the ability to supply fuel in predetermined volumes through the side holes - nozzles 19 in the walls of the marked rods 18.

Sections of independent fairings 16 are hingedly mounted on hollow rods 18, with the possibility of rotation relative to the longitudinal axis of the hollow rod 18 and relative to each other within an angle of up to 80 ° from the plane of the longitudinal centerline of the turbine with the possibility of coinciding the direction of the chord of the fairing profile with the velocity vector of the incoming flow and have channels 20 and holes, can be in the form of slots 21, which connect the inner volume of the fairing with the turbine flow part 13, with the possibility of supplying a rich fuel mixture through the holes 21 in the rear part of the fairing 16 profile.

In the axial direction, after each block 15 of the nozzles, the stages of the isothermal part of the turbine are installed, which include rigidly fixed on the shaft 22 (see Fig. 3 - Fig. 7) in each stage honeycomb nozzles 23 and behind them at least one row of blades 24 of the turbine 25 (Fig. 2), further behind all the stages of the isothermal part of the turbine, for example, in our example, seven stages, at least one stage of the turbine 26 is installed, with the possibility of providing adiabatic expansion of the working gas, the outlet of which is directed to the heat exchanger 12 for heating the air after the compressor 10 before reheating combustion chambers 11 due to thermal energy of exhaust gases.

Each cell of the nozzle apparatus 23 is formed by concentrically located shells 27 and radially directed plates 28, moreover, the cells of the nozzle apparatus uniformly cover the flow path of each of the turbine stages. Radial plates 28 cells of the nozzle apparatus are made with a cross-section in the form of a symmetrical airfoil (see Fig. 6). The change in the cross-section along the length of the nozzle apparatus channels, the confusion required to accelerate the gas flow, is provided by the change in the thickness of the radially mounted plates 28, as well as the partial placement of the aerodynamic profile of the turbine blades 24 inside the nozzle apparatus channel with partial overlap in the axial direction.

In FIG. 2 shows that the gas turbine engine has a linear arrangement in which all the stages of the compressor and the turbine are located along one center line, but taking into account all the essential features of the sixth independent claim, it is possible that the initial part 30 of the turbine with the heating chamber 31 includes a part of the isothermal stages. parts of the turbine serving the payload 32 are installed along one center line, and the second part 33 of the turbine with a part of the stages of the isothermal part of the turbine and the final adiabatic stages of the turbine 34 and the compressor 35 are directed along another center line, which is rotated relative to the first at an angle, even up to 180 ° , based on the requirements of the layout, moreover, both marked parts of the turbine are connected by a gas duct 36 with fixed nozzles 37 installed in it, with the possibility of creating the required field of velocities of the working gas, the compressor 35 is installed on the side of the final stage of the turbine 34 with adiabatic expansion, and between them heat exchanger 38 from which the exhaust gases are already directed outside.

The arrangement shown in FIG. 8 provides the possibility of three-speed operation of the gas turbine engine, optimizes the operation of both the compressor and the load, and also provides a significant reduction in losses in the hot gas ducts of the switching flows from the compressor, heat exchanger and turbine, the efficiency of which differs less than 10% from the efficiency of the Carnot cycle with the support the set operating parameters with the turbine isothermal coefficient not worse than 0.95 and the compressor isothermal coefficient 1.1.

The method of operation of a gas turbine engine will be considered starting with a consideration of the T - S diagram, which is shown in FIG. 1:

Point 1 - ambient temperature. It has a temperature lower than the temperature of the compressed air in the compressor after cooling in a stage.

Point 2 is the maximum temperature of the compressed air in the compressor stage, which is also the temperature at the outlet from the compressor and at the inlet to the heat exchanger.

Point 3 - temperature of compressed air at the outlet of the heat exchanger, it is the temperature at the inlet to the combustion chamber of reheating.

Point 4 is the temperature at the outlet of the reheating combustion chamber, it is the temperature at the inlet to the turbine, it is also the minimum temperature at the outlet of each stage of the isothermal part of the turbine.

Point 5 is the maximum temperature in the stages of the isothermal part of the turbine after heating at the inlet to the nozzle apparatus.

Point 6 - the temperature at the outlet of the last stage of the isothermal part of the turbine is equal to the temperature at point 4 and is equal to the temperature at the inlet to the adiabatic stage of the turbine.

Point 7 is the temperature at the outlet of the adiabatic stage of the turbine and the temperature at the inlet to the heat exchanger from the exhaust gas side.

Point 8 is the temperature of the exhaust gases at the exit from the heat exchanger to the environment. The difference between the temperature at point 7 and point 3 determines underrecovery and significantly affects the intensity of heat exchange between the streams and, accordingly, the dimensions of the heat exchanger.

In the operation of a gas turbine engine, a cycle is implemented that differs in efficiency from the isothermal Carnot cycle, according to my calculations, by less than 10%, which makes it possible to achieve a high efficiency close to the Carnot cycle.

That is, in the compressor stages, a cycle is implemented in which atmospheric air is compressed by a multistage cooled step-by-step compressor, for example by supplying water, as is known from the prior art, to a given pressure with a temperature between points 1 and 2, and the process is maintained taking into account the deviation of the corresponding temperatures from calculated with an accuracy of 0.5% to 1%, and the pressure drop at each compressor stage and the amount of thermal energy removed from the compressed air in each stage are in sizes that ensure the equality of the corresponding inlet and outlet temperatures in each stage with a deviation from 0.5% to 1% except for the first, in which the inlet temperature is determined by the ambient temperature.

Further, at the exit from the last stage of the compressor, the air is heated to the temperature of point 3, without mixing in the heat exchanger 12, with the exhaust gases supplied from the turbine, then it is warmed up in the combustion chamber to the temperature of point 4, to which the fuel is supplied, and then through the stationary nozzle apparatus it forms a given the field of the working gas flow velocities and is directed to the inlet to the stages of the isothermal part of the turbine 25, in which, starting from the first and including the penultimate stage, fuel is supplied before each stage, through the nozzle blocks 15 radially directed to the turbine shaft and fixedly mounted on the turbine housing, and sprayed uniformly over the section of the turbine flow path in an amount that provides heating of the working gases to a predetermined temperature, the temperature of which is maintained within the limits limited by points 4 and 5.

Moreover, the ratio of the inlet pressure to the outlet pressure of each stage of the isothermal part of the turbine is from 1.05 to 1.35, for example 1.1, and in the final stage, with adiabatic expansion, the pressure ratio is 2.5.

In the nozzle blocks 15 through a system of channels, sprayed fuel is mixed with the working gases of the turbine to form a rich fuel mixture inside the fairings 16 of the nozzle block 15, and the resulting fuel mixture is supplied to the trailing edge of the fairing profile 16 of the nozzle block 15 of which the resulting fuel mixture is distributed in volumes directly proportional to the distance from the axial line of the shaft 22, with the possibility of providing uniform atomization over the section of the flow cavity 13 of the turbine, self-ignites and enters the flow path of each stage of the isothermal part of the turbine.

The noted working gas, rotating at the angular speed of rotation of the turbine shaft, then comes sequentially into each stage through the honeycomb nozzles 23 rigidly mounted on the turbine rotor shaft 22, in which, as a result of geometric influence, an increase in the component of the gas velocity in the axial towards the turbine shaft in the direction and further along the outlet of the nozzle apparatus 23, the working gas flows around the aerodynamic profiles of the turbine blades 24 installed on the shaft radially and the rotor 14 with the turbine blades 24 rotate together with the nozzle assemblies that transmit mechanical energy, which ensures the rotation of the turbine, with equal the amount of used and added thermal energy from the combustion of fuel, which spontaneously ignites, in the flow of the working gas supplied by the nozzle blocks 15 installed in front of the entrance to each stage, and the robots maintain the specified parameters with the isothermal coefficient not worse than 0.95, and then the working gas enters the last from a turbine 26 in which, without performing heating at the beginning of the stage, the final expansion of the working gas is performed and the outlet temperature is reduced to the temperature indicated by point 7, and then the working gas is sent to the heat exchanger 12, in which heat exchange is carried out between it and the compressed stream in the compressor 10 air, which, as indicated above, is directed to the combustion chamber of the reheating 11, and the working gas is directed further into the atmosphere, completing the cycle at point 8.

In our version, a heat exchanger is installed in each compressor stage between the outlet from the previous one and the axial inlet to the next stage, in which the compressed air is cooled to a temperature equal to the temperature of the air at the inlet to the previous stage in the diagram (see Fig. 1) this is reflected in the form of an isobar with a falling temperature connecting the exit from the previous degree with the entrance to the next one. This method of constructing the compressor ensures maximum efficiency, that is, the minimum loss of mechanical energy and, accordingly, the minimum amount of heat that must be removed from the compressor.

Analysis of the essential features shown in the method of operation and confirmed by the design, taking into account theoretical engineering calculations and confirmed by the operating parameters of known gas turbine engines, indicate that the most optimal is that the amount of thermal energy supplied by the combustion of fuel from each block of nozzles in front of the stages is set taking into account the delay of the deviation corresponding temperatures from those calculated with an accuracy of 0.5% to 1%, the pressure drop at each compressor stage and the amount of thermal energy removed from the compressed air in each degree are in sizes that ensure the equality of the corresponding inlet and outlet temperatures in each stage with a deviation from 0 , 5% to 1% except for the first, in which the inlet temperature is determined by the ambient temperature, the ratio of the inlet pressure to the outlet pressure of each isothermal turbine stage is from 1.05 to 1.35 with the same isothermal coefficient, and in the final stage, with adiabats expansion, with a pressure ratio of not more than 3.

The most effective, in the example, cycle of operation of a heat engine on a single-phase working medium is the Carnot cycle. The efficiency of this ideal cycle is determined only by the temperatures of compression and expansion. Modern GTUs operating on the Brayton cycle have an efficiency 50-55% lower at equal temperatures than in the Carnot cycle, and they are the same, but supplemented by a steam cycle, which allows the use of the heat of the turbine exhaust gases, which allows to obtain the efficiency of which is 20-30% lower than in the Carnot cycle, however, the last way to increase the efficiency is possible only in a gas turbine with a sufficiently large capacity, that is, more than 100 MW.

Therefore, at present, the main direction for increasing the efficiency, despite very high costs, remains the way of increasing the temperature at the inlet to the turbine. However, the lag behind the efficiency of the Carnot cycle not only remains, but even increases due to the need to intensify the cooling of the elements of the combustion chambers and the turbine.

In fig. 1 shows a cycle that, with the elimination of the final adiabatic stage in the turbine, replacing the processes of stepwise compression and expansion with heating or cooling by isothermal processes of compression and expansion and idealization of all other processes, equivalent in energy, will have practically the same efficiency with the Carnot cycle and which will similarly be depend only on the temperatures of compression and expansion of equivalent isothermal processes. To determine them, the isothermal coefficient is used as the ratio of the temperature of the equivalent isothermal process of compression or expansion to the temperature of the beginning of the adiabatic real process of expansion or compression in a real stage of a gas turbine plant, provided that the energy received or expended in the stage is equal in both cases.

Thus, fulfilling the conditions of equality of temperatures at the beginning of the compression or expansion process and equality of the coefficient of isothermality in stages, which is performed by the process of isobaric cooling or heating in each stage of the turbomachine and taking into account the real losses and other features of the processes, based on the results obtained during the operation of real machines, the permissible compression or expansion ratio in a real gas turbine plant is determined, which provides a given approximation of its efficiency to the efficiency of the Carnot cycle.

It is obvious that with an increase in the number of stages of a real compressor and turbine, i.e. a decrease in the pressure ratio at the inlet and outlet of each stage, and a decrease in losses, i.e. as the polytropic index of real compression and expansion processes approaches the adiabatic index equal to 1.4 for air, the efficiency of the turbine and compressor will approach the efficiency of ideal isothermal machines provided by the Carnot cycle.

Since the proposed principle of operation and organization of processes in a given turbine and compressor allows, with a given accuracy, to bring the efficiency of a real gas turbine engine closer to the efficiency of an ideal Carnot cycle, which includes isothermal compressors and a turbine, and also taking into account the significant difference between the proposed turbomachines and existing ones, it seems appropriate to further denote as quasi-isothermal, i.e. close to isothermal ones with an efficiency no worse than 0.9 of the efficiency of ideal isothermal turbines and compressors provided by the Carnot cycle.

Based on this, it was determined that the minimum expansion coefficient of 1.05 corresponds to small gas turbines, and 1.35 corresponds to large gas turbines in which the losses are minimal and the compression and expansion processes are practically adiabatic and can only increase due to losses during the necessary cooling at very high temperatures. Along the lower limit of pressure drops, an unjustified increase in the number of stages occurs, when the upper limit specified by me in the formula is exceeded, there is a sharp drop in the turbine isothermal coefficient (significantly below 0.95) and, accordingly, a sharp decrease in the efficiency of the GTU.

The compressor isothermal coefficient of 1.1 corresponds to the performance of a good centrifugal stage with a compression ratio of 1.4-1.5, at which the temperature of the compressed air in the stage ranges from 320 to 380 ° K, which allows the use of water for cooling in normal heat transfer mode.

In order to reduce the heat load on the heat exchanger and the maximum temperature in it, a final adiabatic turbine stage is introduced into the cycle, which allows to reduce the temperature at the inlet to the heat exchanger from the turbine exhaust side by up to 25%, gives no more than 0.5-0.75% decrease in efficiency ...

An increase in the pressure drop at the specified stage by more than 3 leads to a sharp drop in efficiency. Hydraulic losses in the pipes and the heat exchanger, as well as losses due to underrecovery within the range of 25-75 ° C at the hot end of the heat exchanger, based on the results of the operation of real gas turbine plants, as a rule, do not exceed 1-1.5%.

To ensure the maximum compactness of the heat exchanger and increase the efficiency in this technical solution, a design of the nozzle apparatus, the impeller is used differently than in classical turbines, which makes it possible to ensure minimum losses in the specified range of step differences, and therefore to ensure operation at higher pressures, together with the implementation isothermal conditions, leads to a significant increase in the specific power of the turbine per unit of air flow, and therefore to a decrease in the size of the heat exchanger and other gas turbine units.

This method and design makes it possible to create a gas turbine unit that, in a wide power range, is capable of having an efficiency of at least 65% at a temperature in the turbine not higher than 1100-1200 ° K, and at temperatures of 1400-1500 ° K, materials for which are sufficiently well mastered by the industry to obtain an efficiency of more 70%.

For comparison, I would like to inform you that in order to achieve an efficiency equal to 65% in the national program of Japan, the task is to develop a material for turbine elements by 2020, which is able to provide a sufficient gas turbine resource at a temperature of about 2000 ° K, for their use in a combined high-power gas turbine for thermal power plants , and the method I proposed and its design features make it possible to create a GTU with a given deviation of its efficiency from the efficiency of the most efficient Carnot cycle of no more than 10% at any temperatures and provided technological capabilities of real production.

Claims (14)

1. A method of operation of a gas turbine engine, in which air from at least one compressor enters the first combustion chamber acting after the compressor unit, the first turbine acting after the first combustion chamber, the second combustion chamber acting after the first turbine and the second turbine acting behind the second combustion chamber, which self-ignites, characterized in that a cycle is implemented in the compressor stages in which atmospheric air is compressed by a multistage cooled step-by-step compressor to a predetermined pressure, then at the outlet of the last compressor stage the air is heated without mixing in the heat exchanger with exhaust gases supplied from the turbine, then heats up in the combustion chamber, into which the fuel enters and then through a stationary nozzle apparatus, in which a given field of working gas flow velocities is formed, is directed to the entrance to the turbine stages, in which, starting from the first and including the penultimate stage, before each step serves fuel through the nozzle blocks radially directed to the turbine shaft and fixedly mounted on the turbine housing and is sprayed uniformly over the section of the turbine flow path in an amount that provides heating of the working gases to a predetermined temperature, the specified working gas, rotating at the angular speed of rotation of the turbine shaft, then enters sequentially into each stage through honeycomb nozzles rigidly mounted on the turbine shaft, in which, due to the geometric effect, there is an increase in the gas velocity component in the axial direction with respect to the turbine shaft and further down the nozzle outlet, the working gas flows around the aerodynamic profiles of the turbine blades, installed on the shaft in the radial direction, which rotate together with the nozzle apparatus and which transmit mechanical energy to the specified shaft in an amount equal to the thermal energy received from the combustion of fuel at the beginning of this stage, which ensures the rotation of the turbine, with an equal number of the quality of the used mechanical energy and the added thermal energy in the stages from the combustion of the fuel supplied by the injector blocks in the flow of the working gas, installed in front of the inlet to the stage, and which maintain the specified operating parameters with the isothermal coefficient not worse than 0.95, and then the working gas enters the last a stage in which, without performing preheating at the beginning of the stage, a final expansion of the working gas and a decrease in the outlet temperature is carried out, and then the working gas is sent to a heat exchanger, in which heat is exchanged between it and the compressed air stream in the compressor, and, as indicated above, is directed into the heating chamber, and the specified working gas is directed further into the atmosphere, closing the cycle. 2. The method of operation according to claim 1, characterized in that the amount of thermal energy supplied by the combustion of fuel from each block of injectors before the stages is set taking into account the equality of the corresponding temperatures in all stages of the isothermal part of the turbine with the deviation of the corresponding temperatures from the calculated ones with an accuracy of 0.5 % up to 1%. 3. A method of operation according to claim 1, characterized in that the pressure drop at each compressor stage and the amount of thermal energy removed from the compressed air in each stage are in values that ensure the equality of the corresponding inlet and outlet temperatures in each stage with a deviation from 0, 5% to 1% except for the first, in which the inlet temperature is determined by the ambient temperature. 4. A method of operation according to claim 1, characterized in that the ratio of the inlet pressure to the outlet pressure of each stage of the isothermal part of the turbine is from 1.05 to 1.35 with the same isothermal coefficient, and in the final stage, with adiabatic expansion, no more 3. 5. The method of operation according to claim 4, characterized in that the atomized fuel is mixed with the working gases of the turbine in the injector blocks through the channel system to form a rich fuel mixture inside the fairings of the injector block, and the resulting fuel mixture is supplied to the trailing edge of the fairing profile of the injector block, in which the resulting fuel mixture is distributed in volumes directly proportional to the distance from the shaft axis, with the ability to ensure uniform spraying over the section of the turbine flow cavity, self-ignites and enters the flow path of each stage of the isothermal part of the turbine. 6. Gas turbine engine, including a compressor configured to receive and compress a working fluid, a combustion chamber configured to receive compressed working fluid from the compressor and fuel, with the possibility of burning to form a working gas, and a turbine that has at least two stages and made with the possibility of receiving the working gas from the combustion chamber and using its energy to rotate the shaft, while at least one device for reheating is installed between the turbine stages, which contain nozzle blocks, characterized in that a cooled stepwise compressor, and a heat exchanger is installed behind the compressor in front of the main combustion chamber of reheating, and then in the flow path in the stages of the isothermal part of the turbine rotor, which has at least three stages, on the outer surface of the flow path of the engine casing at the entrance to each stage, radially directed fuel injector units are installed k, having in their external cross-section fairings with an axisymmetric aerodynamic profile, the direction of the longitudinal chord of which coincides with the direction of the velocity vector of the incoming gas flow and have channels with the possibility of supplying a rich fuel mixture through the holes in the rear part of the profile, and then in the axial direction after each of the nozzle block, the stages of the isothermal part of the turbine are installed, including rigidly fixed on the shaft, in each stage, honeycomb nozzles and behind them at least one row of turbine blades, then behind the steps of the isothermal part of the turbine, at least one turbine stage is installed, with the possibility of providing an adiabatic expansion of the working gas, the outlet of which is directed to the heat exchanger for heating the air after the compressor in front of the heating chambers, due to the thermal energy of the turbine exhaust gases. 7. The engine of claim. 6, characterized in that the turbine stages, with the ability to operate in the mode of adiabatic expansion, are installed on a separate shaft. 8. The engine according to claim 6, characterized in that the cells of the nozzle apparatus uniformly cover the flow path of the turbine stage and each cell is formed by concentrically located shells and radially directed plates. 9. The engine of claim. 8, characterized in that the radial plates of the cells of the nozzle apparatus are made with a cross-section in the form of a symmetrical airfoil. 10. The engine according to claim 9, characterized in that the change in the cross-section along the length of the nozzle apparatus channels necessary to accelerate the gas flow is achieved by changing the thickness of the radially mounted plates, as well as by partially placing the aerodynamic profile of the turbine blades inside the nozzle apparatus channel with partial axial overlap. 11. The engine according to claim 6, characterized in that on the outer surface of the turbine housing inside the flow path at the inlet to each stage, fuel injector blocks radially directed to the axial line of the shaft, which are hollow rods with a blind end and radial holes forming channels for supplying fuel, with the possibility of supplying fuel in specified quantities through lateral radial holes - nozzles in the walls of these rods, and sections of independent fairings are hinged on the rods, with the possibility of rotation relative to the longitudinal axis of the rod within an angle of up to 80 ° relative to the longitudinal plane the turbine axes and with the possibility of coincidence of the chord direction of the fairing profile with the oncoming flow velocity vector and have a system of channels connecting the inner volume of the fairing with the turbine flow path. 12. The engine of claim. 6, characterized in that the engine has a linear arrangement in which all stages of the compressor and turbine are located along one center line. 13. The engine according to claim 6, characterized in that the initial part of the turbine with a heating chamber includes the stages of the first isothermal part of the turbine serving the payload, installed along one center line, and the second isothermal part of the turbine together with the final adiabatic stages of the turbine and the compressor have a shaft directed along another axial line deployed relative to the first at an angle of up to 180 °, based on the requirements of the layout, and both of these parts are connected by a gas duct with fixed nozzle devices installed in it, with the possibility of creating the required field of speeds of movement of the working gas, the compressor is installed from the side the final stage of the turbine, the drive shafts of which are connected to the shaft of the specified part of the turbine. 14. An engine according to claim 6, characterized in that, with the exception of the last stage, cooling heat exchangers are installed between the stages in the compressor.

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