(variants)

Field of the Invention

The claimed group of inventions relates to the field of of power-plant engineering in particular to gas turbine engines, and can be used at thermal power plants, also as primary engines in road and waterborne transport, aviation and other fields of national economy.

Prior Art

Currently, gas turbine engines c HenpeptiBHtiM cropaHHeM TonjiHBa npn nocToaHHOM flaBJieHHHare widely used including detachable housing with two cavities with interconnected nozzle assemblies; in one cavity, the compressor wheel is installed, and the other includes turbine wheel which are rigidly fixed on the common shaft installed in the housing, combustion chamber, heat exchanger, inlet and outlet nipples (See p.101 "Scientific and Technical Dictionary", "Sovetskaya encyclopedia", Moscow, 1977, and p.200 "Brief scientific and Technical Dictionary", Gos. Izdatelstvo of technical and theoretical literature, Moscow, 1956 ).

In such gas turbine engines, compression of the ambient air is performed in special compressors; the air is supplied in the combustion chamber together with a fuel; with combustion, the fuel heats the air which forms the excessive gas pressure which in the turbine converses in mechanical power; the main portion of this is spent for the air compression in the compressor.

The distinctive structural feature of the all known turbine engines is the fact that each thermodynamic process of the cycle is realized in the separate device; the devices are arranged along a common shaft in a certain sequence, that is, the working body compression is performed in the compressor, heat feed - in the combustion chamber, and the air expansion with performing useful work— in the gas turbine.

Existence of diverse devices complicates the design of gas turbine engines and leads to INCREASE in their dimensions. In order to elevate the efficiency, in many modern gas turbine engines the sucked air after pressurization stages passes through heat exhanger, is heated, and expands. Further, the air volume has a nonlinear expansion coefficient; the maximum expansion takes place at 20/25 of the temperature in the combustion chamber. With feeding the heated air that has lost its potential energy into the combustion chamber, the main part in gas formation (increase in volume and pressure) provides the amount of the fuel being incinerated as the compressed air before it is supplied to the combustion chamber is expanded to the maximum extent, and the expansion coefficient at the temperature of the blend burning has the minimum value. This results in considerable increase in toxicity, considerable increase in the fuel consumption and considerable decrease in efficiency.

A gas turbine engine of internal combustion with the fuel mixture combustion in Otto cycle including a turbine installed on the shaft in the engine housing.; the turbine blades are situated in the combustion chamber; explosive mixture is supplied to the combustion chamber through the nozzles arranged around the blades and having continuous ignition, with synchronous cutoff; the obtained pressure is supplied to the turbine blades and reflectors (FR jNb 1538421).

The disadvantage of these engines is the rounded shape of the blades formed as deepenings in the turbine housing; this decreases the efficiency of the gas pressure towards the turbine rotation as the gases press evenly in all directions. Another disadvantage is the fact that is impossible to arrange a larger number of the blades due their shapes within such a fullmetal turbine.

A gas turbine engine is known including a housing wherein a compressor is situated including a rotor with a disk with working blades with are installed on its shaft kinematically connected with the turbine shaft, and an open diffuser, arranged concentratically relatively to the latter of the gas turbine; combustion chamber arranged uniformly along the internalsurface of the working wheel of the gas turbine at the same distance from the rotation axis and rigidly fixed on it; input is made as a diffuser , and the output is made as a nozzle, fuel burners, installed in the combustion chamber and connected with fuel tubes are characterised in that the compressor is provided with guiding device, input of each combustion chambers is located in the open diffuser over the blades of guiding device and is oriented towards its output, and output of each combustion chamber protrudes over the outer surface of the working wheel and is made as a Laval nozzle. Each combustion chamber is provided with supersonic diffuser; the engine is provided with a generator with windings, a stator combined with the engine housing, and a rotor kinematically connected though transmission with the turbine shaft, and combustion chambers are provided with electrodes installed in the output portion of the diffuser and connected electrically with windings of the generator stator (RU jV° 2078968, prototype).

The disadvantages of this engine is high fuel consumption and hgih demands on its quality; high air consumption and lubricants; high volume of toxic releases of combustion products; instability of the moment created on the shaft and the number of revolutions of the latter; low specific power (on the unit mass); high prime cost; non- perfect stream formation; low efficiency (for devices of Segner wheel type cannot exceed 50%) due to energy loss with discharge of the hot compressed working body from the nozzles in the combustion chamber into the atmosphere; non-optimum structure of the working body stream and high losses with the output velocity; possibility of the structure destruction and the jet adhesion; thermal pollution of environment with discharged combustion products.

Summary of the Invention

The aim of the invention embodiments is development of an active gas turbine engine and expansion of the gas turbine engines inventory .

The technical result which provides the solution of the problem in both embodiments lies in multiple decrease in the fuel consumption with decrease in demanda to its quality and expansion of nomenclature including biologic fuel, paraffin oil, diesel fuel;

- multiple decrease in air and lubricant consumption;

- multiple decrease in release of combustion products;

- increase in efficiency (on high revolutions);

- high specific power (to the unit mass), decrease in the prime cost;

- quick response to the fuel feed;

- durability due to the bearings location outside of the high temperature zone;

- discharge of hot gases on periphery;

- improved stream formation using a bevel locking device;

- relatively low rotation velocity;

- structural changes in the metal are absent;

and also in improvement in performance and efficiency at the cost of formation of the optimum structure of the working body stream at the input with guiding devices, radial removal of the exhaust working body out of the zone of interaction with the blades; decrease in internalenergy losses and energy losses with the output velocity, conversion of the working body energy at the maximum diameter with the minimum volume of the working body (in the nozzles) and rather high energy density excluding a possibility of the structure destruction and the jet adhesion, stabilization with the minimum rotational friction of the rotor which also represents a kinetic energy accumulator (fly wheel), and also to the optimum arrangement of the nozzles and blades in several rows, and to decrease in thermal environmental pollution.

Concept of the invention of a gas turbine engine with two-way air feed lies in the fact that the gas turbine engine a shaft where rotor of the gas turbine is installed; on two sides thereof rotors of first and second centrifugal compressors are installed symmetrically; between them an internal guiding device is installed the blades thereof are made on both sides in the housing, and from suction side of first compressor on two side in the housing a guiding device is installed ; the blade thereof are made on two sides in the housing, and the bearings for the shaft installation in the housing; with that rotor of the gas turbine is made with two rows of combustion chambers arranged in circumference and is installed so to provide a capability of rotation in the distributing stator made with the fuel feed channels and mixing chamber for fuel mixture feed into inlet windows of each combustion chamber having inclined outlet nozzles faced with their cutoffs the vanes of the support stator where fuel overheat channels are made connected with fuel feed channels with fuel burners of the mixing chambers made in the distributing stator around the blades of second compressor.

Preferably, the bearing are located outside the high temperature zone, the rotor is installed relatively to the distributing stator and support stator with the gaps along circular and side surfaces connected with the channels of the cooling air; the number of nozzles in two rows is equal to the number of blades in the support stator; the rotor is made with with two rows of combustion chambers and nozzles shifted relatively to each other in angular direction; windows for exhaust gases removal in the stator are made radially and are directly connected to atmosphere; the blades are made with the edges inclined towards the circle and are installed radially or at a slant position; the nozzles are made with a critical cross-section at the cutoff in the plane inclined to the radial plane of the rotor on which recessions are made under each nozzle; direction of the nozzle symmetry axis is selected from the condition of the maximum approximation to the tangent surface of the rotor and perpendicular action of the jet on the support stator blades; combustion chambers have a trapezoidal shape diverging to peripheral portion and inclined to against direction of rotation; on each side of the distributing stator circular grooves are made alternately alternately connected with fuel overheat channels of each other blade alternately with fuel overheat channels of other blades.

Concept of the invention as related to the gas turbine engine with one-way air feed lies in the fact that the gas turbine engine includes a shaft installed on bearings in the housing wherein a rotor of the gas turbine is mounted; on one side of the rotor of first and second centrifugal compressors between them an internal guiding device; the blades of the guiding device are made in the housing; an input guiding device is located on the suction side of first compressor; the rotor of the gas turbine is made with two rows of combustion chambers arranged in circumference and is installed so to provide rotation in the distributing stator made with fuel feed channels and mixing chambers to feed fuel mixture into inlet windows of each combustion chamber having inclined outlet nozzles faced with their cuts cpe3aMH the vanes of the support stator located in the circular gap; in the vanes, the channels for fuel overheat are made which are connected through fuel feed channels with the fuel burners of the combustion chambers made in the distributing stator around the blades of third compressor; the blades are made in the rotor

Preferably, the bearing are located outside the high temperature zone, the rotor is installed relatively to the distributing stator and support stator with the gaps along circular and side surfaces connected with the channels of the cooling air; the number of nozzles in two rows is equal to the number of blades in the support stator; the rotor is made with with two rows of combustion chambers and nozzles shifted relatively to each other in angular direction; windows for exhaust gases removal in the stator are made radially and are directly connected to atmosphere; the blades are made with the edges inclined towards the circle and are installed radially or at a slant posotion; the nozzles are made with a critical cross-section at the cutoff in the plane inclined to the radial plane of the rotor on which recessions are made under each nozzle; direction of the nozzle symmetry axis is selected from the condition of the maximum approximation to the tangent surface of the rotor and perpendicular action of the jet on the support stator blades; combustion chambers have a trapezoidal shape diverging to peripheral portion and inclined to against direction of rotation; on each side of the distributing stator circular grooves are made alternately alternately connected with fuel overheat channels of each other blade alternately with fuel overheat channels of other blades.

Description of the drawing

In the drawing Fig.l a construction diagram of the engine with two-way air feed (longitudinal section) in the Fig.2 and Fig.3 - enlarged fragments of the Fig.1 , in the Fig.4 - cross-section according to the Fig.l, in the Fig.5 - enlarged fragment of the Fig.4, in the Fig.6— construction diagram of the engine rotora, in the Fig.7 - construction diagram of the distributing stator, in the Fig.8 - the end view of the engine according to the Fig.l, in the Fig.9 - construction diagram of the engine with one-way air feed (longitudinal section), in the Fig.10 - enlarged fragment of the Fig.9, in the Fig.l 1 - cross-section according to the Fig.9, in the Fig.12 - construction diagram of the engine rotor with third compressor, in the Fig.13 - construction diagram of the distributing stator, in the Fig.14 - a view from right end of the engine according to the Fig.9 are shown.

Detailed Description of the Invention

In the drawings Fig.1 - Fig.14 the designations are made:

- shaft 1 of the gas turbine;

- active rotor 2 of the gas turbine;

- vanes 3 of first (right and left according to drawings Fig.1-3) centrifugal compressor (firt stage of the air compression);

- vanes 4 of the internal guiding device (guiding device of first compressor, right and left according to drawings Fig.1-3);

- vanes 5 of the input guiding device B (right and left according to drawings Fig.1-3);

- vanes 6 of second (right and left according to drawings Fig.1-3 ) centrifugal compressor (second stage of the air compression);

- bearings 7 of the shaft 1 (roller or slider bearings);

- rotor 8 of first centrifugal compressor; - housing (made in the form of a flange) 9 of the internal guiding device with the vanes 4 of the internal guiding device;

- direction 10 of exhaust gases removal;

- combustion chamber 1 1 ;

- flange 12;

- flange 13 of first compressor;

- housing 14 of the inlet guiding device;

- fuel inlets 15 into the channels of distributing stator 22;

- direction 16 of atmospheric air intake with first compressor;

- rotor 17 of second centrifugal compressor;

- support blades 18 of the support stator;

- support (external) stator 19;

- nozzles 20 (fuel burners) for evaporated (overheated) fuel injection;

- converging mixing chambers 21 of rectangular cross-section in the distributing 22;

- channels 23 for fuel feed to fuel burners 20;

- converging rectangular nozzles 24 on the external rotor surface 2;

- channels 25 for fuel overheat (evaporator);

- inlet windows 26 for formed fuel injection into combustion chambers 11 ; - directions 27,28 of exhaust discharge from the nozzles 24 in two rows of combustion chambers 11 ;

- zone 29 of removal the stator exhaust gases;

- windows 30 for radial removal the exhaust gases of the housing 19;

- bevel locking device 33.

In the drawings Fig.9 - Fig.13 the following is additionally designated:

- vanes 31 of the third compressor, 3aKpenjieHHbie Ha rotor 2;

- flow cross-sections 32 for intake of the compressed air in third compressor. The gas turbine engine with two-way air feed (intake) according to the Figs.l-

8 includes a shaft 1 where and active rotor 2 of the gas turbine is installed; on two sides of the rotor, the rotors 8,17 are installed of first H second centrifugal compressors with vanes 3,6, respectively; between the rotors an internal guiding device is placed; the vanes 4 thereof are made on two sides in the housing (flange) 9. From the suction side of first compressor in direction 16 from two sides (to the right and to the left according to the Figs.1-3) in the housing 14 an input guiding device is located the vanes 5 thereof are made on two sides in the housing 14, and bearings 7 for the shaft 1 installation in the housing 14. Thus, in this embodiment of the engine on each side of the rotor 2, first and second right and left compressors with the rotors 8,17 and vanes 3,6 are located symmetrically according to drawing in the Figs.1-3, the right and left guiding devices with the vanes 4,5, and also a TaK¾ce right and left bearings 7. The rotor 2 of the gas turbine is symmetrical with two rows of combustion chambers 11 arranged in circumference so to provide their capability of rotation in the distributing stator 22 made with fuel feed channels 23 and mixing chambers 21 for feeding the fuel mixture into inlet windows 26 of each combustion chamber 11 with inclined outlet nozzles 24 uniformly distributed over the circumference faced in pairs with their cutoffs the blades 18 of the support stator 19 located in the circular gap where channels 25 for fuel overheat are made connected through fuel feed channels 23 with fuel burners 20 of the mixing chambers 21 made in the distributing stator 22 around the vanes 6 of the second compressor.

The bearings 7 are located on both sides of on the suction side of first compressor, that is, outside the high temperature zone.

The rotor 2 is installed relatively the distributing stator 22 and support stator 19 with the gaps over circular and side surfaces connected with cooling air circulation channels (not shown).

The number of nozzles in two rows is equal to 24 the number of blades 18 of the support stator 19.

The rotor 2 of the gas turbine is made with two rows of combustion chambers 11 and nozzles 24 relatively to each other in angular direction.

The windows 30 for removal exhaust gases are made in the support stator 19 radially and directly connected with atmosphere.

The blades 18 in the support stator 19 are made with the edges inclined to circumference and are installed radially or at a slant position.

The nozzles 24 are made with the critical section on the cutoff in the plane inclined to the radial plane of the rotor 2 where pits are made beneath under each nozzle 24. Direction of the axis of symmetry of the nozzle 24 is selected from the condition the maximum approach to the tangent surface of the rotor 2, and perpendicular action of the jet on the blades 18 of the support stator 19.

The combustion chambers 11 are made of trapezoidal shape diverging to periphery at an inclined against direction of rotation.

On each side in the distributing stator 22 two circular grooves are made alternately (in the figure one groove is shown, not marked with item); each of them is connected with fuel overheat channels 25 of each other vane 18 alternately with fuel overheat channels of other blades 18.

The inlet windows 26 of the combustion chambers 11 are made of rectangular cross-section in the internal surface of the active rotor 2 and represent an aerodynamic stage of flash expansion in the internal volume of the combustion chamber 11 against the direction of the active rotor 2 rotation. The combustion chambers 11 are also inclined against the rotor 2 rotation. Internal distributing stator 22 gas has converging mixing chambers 21 of rectangular cross-section faced with the narrow critical portion rectangular inlet windows 26 of combustion chambers 11 . The mixing chambers 21 of the distributing stator 22 in the median cross-section have diverging nozzles (fuel burners) 20 for injecting evaporated fuel directly in to the air stream. The nozzles 24 are made with their configuration convex in direction of the rotor 2 rotation. Direction of the axis of the nozzles 24 symmetry is located at a small angle to the rotor 2 radius in the rotational plane; direction 27,28 of the combustion products exhaust from the nozzles 24 in two rows of the combustion chambers 11 are defined with the fact that direction of the nozzle 24 axis of symmetry is selected from the condition of the maximum approach to the tangent surface of the rotor 2 and perpendicular action of the jet on the blades 18 of the support stator 19.

The channels 23 for fuel feed provided with fuel burners 20 are made in the distributing stator 22 on both sides of the rotor 2.

In front of the first compressor a bevel locking device 33 is installed.

Two rows of the combustion chambers 11 and the nozzles 24 are shifted relatively to each other in angular direction by a half of the central angle between the nozzles in the same row. The support blades 18 of the stator 19 and the windows 30 are common for two rows of the combustion chambers 11 and nozzles 24 of the rotor 2. Curvature of the nozzles 24 are selected from the conditions of uniform wearing of their internal surface.

The combustion chambers 11 are formed of two neighboring detachable elements (not designated) of 1-shaped cross-section. The zone 29 for removal the exhaust gases is common for all the exhaust gases incoming from chambers 11.

The gas turbine engine with one-way air feed (intake) according to the Figs.9- 13 includes a shaft installed on bearings 7 in the housing 9,14 where the rotor 2 of the gas turbine is mounted, one one side thereof the rotors 8,17 of first H second centrifugal compressors with the vanes 3,6 are installed, respectively; between them , the internal guiding device is located, the vanes 4 thereof are made kin the housing (flange) 9. From the suction side of first compressor an input guiding device is located, the vanes 5 thereof are made in the housing 14.

Thus in this embodiment of the engine, on one side of the rotor 2 right and left according to drawings Figs.9,10, first and second compressors with rotors 8,17 and the blades nonacTaMH 3,6, guiding devices with the vanes 4,5 are installed. The rotor 2 of the gas turbine is symmetrical with two rows of combustion chambers 11 arranged on the circumference and is installed so to provide a capability of rotation in the distributing stator 22 made with the fuel feed channels 23 and mixing chambers 21 for feeding the fuel mixture into inlet windows 26 of each combustion chamber 11 having inclined outlet nozzles 24 uniformly distributed over the circumference and faced in pairs with their cutoffs the blades 18 of the support stator 19; the vanes where fuel overheat channels 25 are made are located in the circular gap; the channels are connected through fuel feed channels 23 with fuel burners 20 of the mixing chambers 21 made in the distributing stator 22 around the vanes 31 of third compressor.

The roller bearing 7 are located on both sides from the suction side of first compressor, that is, outside the high temperature zone.

The rotor 2 is installed relatively the distributing stator 22 and support stator 19 with the gaps on circumferential and side surfaces connected with the cooling air circulation channels (not shown) .

The number of the nozzles 24 in two rows is equal to the number of the blades

18 of the support stator 19. The rotor 2 of the gas turbine is made with two rows of the combustion chambers 11 and nozzles 24 shifted relatively to each other in angular direction.

The windows 30 for removal the exhaust gases are made in the support stator 19 radially and are direct connected with atmosphere.

The blades 18 of the support stator 19 are made with the edge inclined to circumference and installed radially or at a slant position.

The nozzles are made with the critical cross-section on the cutoff in the plane inclined to the radial plane of the rotor where pits (not designated) are made beneath each nozzle 24. Direction of the axis of symmetry of the nozzle 24 is selected from the condition of the maximum approximation to the tangent surface of the rotor 2 and perpendicular action of the jet on the support stator blades 18 of the support stator 19.

The combustion chambers 11 are made of trapezoidal shape diverging to periphery at an angle to direction of rotation.

On each side in the distributing stator 22 (in the Fig.13 one groove is shown, not marked with item), with two circular groove are alternately connected with fuel overheat channels of each other vane 18 alternately with fuel overheat channels 25 of other blades 18.

The inlet windows 26 of the combustion chambers 11 are made of rectangular cross-section in the internal surface of the active rotor 2 and represent an aerodynamic stage of flash expansion in the internal volume of the combustion chamber 11 against the direction of the active rotor 2 rotation. The combustion chambers 11 are also inclined against the rotor 2 rotation. Internal distributing stator 22 gas has converging mixing chambers 21 of rectangular cross-section faced with the narrow critical portion rectangular inlet windows 26 of combustion chambers 11 . The mixing chambers 21 of the distributing stator 22 in the median cross-section have diverging nozzles (fuel burners) 20 for injecting evaporated fuel directly in to the air stream. The nozzles 24 are made with their configuration convex in direction of the rotor 2 rotation. Direction of the axis of the nozzles 24 symmetry is located at a small angle to the rotor 2 radius in the rotational plane; direction 27,28 of the combustion products exhaust from the nozzles 24 in two rows of the combustion chambers 11 are defined with the fact that direction of the nozzle 24 axis of symmetry is selected from the condition of the maximum approach to the tangent surface of the rotor 2 and perpendicular action of the jet on the blades 18 of the support stator 19.

The channels 23 for fuel feed provided with fuel burners 20 are made in the distributing stator 22 on both sides of the rotor 2.

Two rows of the combustion chambers 11 and the nozzles 24 are shifted relatively to each other in angular direction by a half of the central angle between the nozzles in the same row. The support blades 18 of the stator 19 and the windows 30 are common for two rows of the combustion chambers 1 1 and nozzles 24 of the rotor 2.

Curvature of the nozzles 24 are selected from the conditions of uniform wearing of their internal surface.

The combustion chambers 11 are formed of two neighboring detachable elements (not designated) of 1-shaped cross-section. The zone 29 for removal the exhaust gases is common for all the exhaust gases incoming from chambers 11.

The gas turbine engine with two-way air feed (intake) according to the

Figs.1-8 operates as follows.

The shaft 1 of the engine with c rotor 2, rotors of 8,17 first and second centrifugal compressors is brought to rotation from any start-up source.

The first centrifugal compressor 8 with 3 intakes the atmospheric air 16 and in the guiding device 4 increases pressure and velocity on other side at the output from the blades c .zrpyrofi CTopoHw Ha Bbrxo^e c jionacTefi 3 yeejiHOiBaeT .zxaBJiemie H cKopocTB B HanpaBJifliomeM annapare 4. Further the compressed air through the radial converging channel enters second centrifugal compressor (rotor 17) with vanes 6. At the output of the vanes 6 the air pressure increases under action of the vanes 31.

The vanes 4,5 of the guiding device provide optimization of compressors operation and increasing their efficiency.

Further the compressed air enters converging rectangular mixing chambers 21 of the internal distributing stator 22 and increases pressure, density and velocity in their converging critical part. The system of air feed support system (not shown) operates independently with the maximum dosing accuracy and pressure and velocity rise in in converging mixing chambers 21 of the distributing stator 22. In the critical part of the converging chambers 21 of the distributing stator 22 fuel burners 20 for fuel injection are located. With the fuel injection directly in the air stream compressed with compressors, the completed fuel mixture enters the combustion chamber 11 through the injection windows 26. As the fuel mixture is formed directly at the input of the combustion chamber 11, and the fuel and the air are in the same aggragative state, efficiently blended fuel enters the combustion chamber 11.

With the fuel combustion, pressure in the combustion chambers increases, constantly obtainable gas volume (of gaseous combustion products) under the pressure existing in the combustion chambers 11 is directed into the intake intake part of converging nozzles 24. At the output of the critical section of each rectangular nozzle 24 at the outer diameter of the active rotor 2 rectangular (flattened) gas jet with high energy density (elevated velocity and pressure) acts on the vane Ha 18 of the stator 19 and loses its energy completely as the jet velocity drops to zero. The number of the support vanes 18 of the stator 19 is equal to the number of rotating combustion chambers 11 and the number of active nozzles 24 in left and right rows (stages) of the active rotor 2. As the combustion chambers 11 with the nozzles 24 in each stage (row) are shifted relatively another by a half of the central angle between the nozzles 24 in them same stage the gas jets in directions 27,28 with the active rotor 2 rotation alternately acts on their respective support vane 18 of the stator 19 without deformation and power pulse repetition per revolution doubles. Thus a smoothed shifted characteristic is obtained (yBejiH¾HBaeTca total number of phase of interaction between the jet with blades 18 per one complete revolution of the active rotora 2). The interval between the power contact peaks is reduced with the peaks of power contacts. Directions 27,28 of combustion products exhaust from the nozzles 24 in two rows of combustion chambers 11 is determined by the fact that the direction of axis of symmetry of the nozzle 24 is selected from the condition the maximum approach to the tangent surface of the rotor 2 and perpendicular action of the jet on the blades 18 of the support stator 19. With that, the working body (combustion products) does not transfer mechanical energy to any intermediate element (piston or turbine blades). Energy of the combustion product jet exhausting from the nozzles 24 acts on support planes of the blades 18 inducing rotation of the rotor 2.

The blades 18 of the stator 19 have through fuel overheat channels 25 and provide fuel feed to all the mixing chambers 21 of the distributing stator 22 (the number of the nozzles 24 is equal to the number of the vanes 18 of the support stator 18). Rather small amount of the exhaust gases is removed in direction 10 into the windows 30 radially creating no pressure in the secondary zone 29. The active rotor 2 of the gas turbine is cooled efficiently (no shown) of the central part and side surfaces of the chambers 11 , temperature of the chamber 11 outer walls of the active rotor 2 is considerably lower that the temperature of fuel combustion. The bearings 7 do not contact with the high temperature zone and operates under the ideal conditions.

The gas turbine engine with one-way air feed (intake) according to the Figs.9-14 operates as follows.

The shaft 1 of the engine with rotor 2, rotors 8,17 of first and second centrifugal compressors and the vanes 31 of third compressor is brought to rotation from any start-up source. The engine with one-way air intake differs from the previous one in the presence of two serial centrifugal compressors one one side and the presence of flow third compressor. As a whole, the operation algorithm almost completely corresponds to that of the engine with two-way air intake.

The first centrifugal compressor 8 with vanes 3 intakes the atmospheric air 16, and on the opposite side at the vanes 3 output increases the pressure and velocity in the guiding device 4. Then the compressed air through the radial converging channel it is sucked with second centrifugal compressor (rotor 17) with the vanes 6. At the output of the vanes 6 the air pressure increases under action of the vanes 3, further the pressure rises under action of the vanes 31.

The vanes 4,5 of the guiding device provide optimization of the compressor functioning and efficiency improvement.

Further the compressed air enters in converging rectangular mixing chambers 21 of the internal distributing 22 and increases pressure, density and velocity in their converging critical part. The system for the air (not shown) operates independently with the maximum dosing accuracy and pressure and velocity rise in in converging mixing chambers 21 of the distributing stator 22. In the critical part of the converging chambers 21 of the distributing stator 22 fuel burners 20 for fuel injection are located. As the fuel mixture is formed directly at the input of the combustion chamber 11, and the fuel and the air are in the same aggregative state, efficiently blended fuel enters the combustion chamber 11. With the fuel injection directly in the air stream 383

15

compressed with compressors, the completed fuel mixture enters the combustion chamber 11 through the injection windows 26. With the fuel combustion, pressure in the combustion chambers 11 increases, constantly obtainable gas volume (of gaseous combustion products) under the pressure existing in the combustion chambers 11 is directed into the intake intake part of converging nozzles 24. At the output of the critical section of each rectangular nozzle 24 at the outer diameter of the active rotor 2 rectangular (flattened) gas jet with high energy density (elevated velocity and pressure) acts on the vane Ha JionaTKy 18 of the stator 19 and loses its energy completely as the jet velocity drops to zero. The number of the support vanes 18 of the stator 19 is equal to the number of rotating combustion chambers 11 and the number of active nozzles 24 in left and right rows (stages) of the active rotor 2. As the combustion chambers 11 with the nozzles 24 in each stage (row) are shifted relatively another by a half of the central angle between the nozzles 24 in them same stage the gas jets in directions 27,28 with the active rotor 2 rotation alternately acts on their respective support vane 18 of the stator 19 without deformation and power pulse repetition per revolution doubles. Thus a smoothed shifted characteristic is obtained (the total number of phase of interaction between the jet with blades 18 per one complete revolution of the active rotora 2 rises). The interval between the power contact peaks is reduced with peaks of power contacts. Directions 27,28 of combustion products exhaust from the nozzles 24 in two rows of combustion chambers 11 is determined by the fact that the direction of axis of symmetry of the nozzle 24 is selected from the condition the maximum approach to the tangent surface of the rotor 2 and perpendicular action of the jet on the blades 18 of the support stator 19.

With that, the working body (combustion products) does not transfer mechanical energy to any intermediate element (piston or turbine blades). Energy of the combustion product jet exhausting from the nozzles 24 acts on support planes of the blades 18 inducing rotation of the rotor 2.

The blades 18 of the stator 19 have through fuel overheat channels 25 and provide fuel feed to all the mixing chambers 21 of the distributing stator 22 (the number of the nozzles 24 is equal to the number of the vanes 18 of the support stator 18). Rather small amount of the exhaust gases is removed in direction 10 into the windows 30 radially creating no pressure in the secondary zone 29. The active rotor 2 of the gas turbine is cooled efficiently (not shown) of the central part and side surfaces of the chambers 11 , temperature of the chamber 11 outer walls of the active rotor 2 is considerably lower that the temperature of fuel combustion. The bearings 7 do not contact with the high temperature zone and operates under the ideal conditions.

As a result, in both embodiments of the engine resistance to pulsed loads rises and operating peak widens. The flat jet acts in parallel to the axis of rotation and the plane of the rotor 2 outer diameter over its width at the maximum effective diameter in the narrow support band. With such interaction parallel combination of all forces of practically equal magnitudes at the same action radius. Aerodynamic resistance of the rotor 2 is almost equal to zero - to this aim the rotor 2 has a smooth finished surface and all the forces are directed in one direction. Further as the exhaust gases leave the vanes 18 of the support stator 19 the exhaust gases are released into the zone 29 and further through the outlet windows are removed from the engine into atmosphere. With that, the pressure in the zone 29 is equal to atmospheric pressure, and the pressure in the combustion chamber 11 is that serial air pressiruzation a can provide (many times higher than atmospheric pressure) providing large pressure difference at inlet and outlet with account of the minimum air consumption and expenditure and, as a consequence, corresponding lower minimum fuel consumption. With account of all the parameters mentioned above we obtain the maximum power with the minimum consumption and expenditure of the air and fuels and the maximum efficiency, and the minimum volume of the toxic releases with the minimum temperature per a unit of useful engine power. In the actual experimental engine being investigated the support vane 18 of the stator 19 has a certain configuration selected from structural considerations (not shown in the drawing). With the rotor 2 rotation and combination of the cutoff plane of the nozzle 24 with support plane of the vane 18 of the stator 19, we obtain a construction of the diverging nozzle 24, the gas jet velocity in direction 27,28 rises and a specific pulse energy rises which also elevates power with the same fuel consumption. In the course of further experimental investigations we obtained a capability of the engine operation in the pulsed (explosive) mode with considerably higher power, decrease in the fuel consumption with the unchanged air consumption. Calculations of the pulsed (explosive) modes provides the parameters: the rate of the fuel formation, the rate of the fuel combustion. The engine reliability does not decrease as the nozzles 24per one revolution of the rotor 2 in sequence alternately change the mode of operation. So with the rotor diameter of 300mm, the number of the support vanes 18 of the support stator 19, the number of the shifted nozzles 24 in two stages (two rows— stages in 16 nozzles 24) is 32. We find that the total number of interaction per one revolution is 1024; with the length of involute of the outer diameter of the rotora of 942 mm, pulse density and repetition rate can be easily calculated.

Example of parameters and specifications for the symmetrical gas turbine engines with two-way air feed.

I .Diameter of the active rotor 2— 300 mm.

2. Number of the active nozzles24 in two rotor rows (stages) -32

3. Number of the stator support blades 18 - 32

4. Number of force interactions per the rotor 2 revolution 2 -1024

5. Dimension of the critical cross-sections of the rectangular nozzle 24 -25*2.3 mm

6. Angle of the focusing part of the nozzle 24 - from 3°30'

7. Number of the rotor 2 revolutions taken as a constant for further investigations -21700 rpm.

8. Power on the shaft 1 of the engine -1500 kW

9. Engine weight, no more than - 24kg.

10. Relation of engine the unit power to the unit weight 20/25 g of weight to 1 kW.

I I .Air consumption depends on the pressure in combustion chambers 11 and the total square of critical cross-sections of all the nozzles 24 and the total square of critical cross-sections of the cooling nozzles (not shown) of combustion chambers 11 on the rotor 2 periphery.

12. Operating temperature in the combustion chamber 11 does not exceed 450 degrees at the temperature of the fuel mixture combustion from 735C ⁰ . 13. The bearing 7 work under ideal conditions and do not contact with the high temperature zone.

Example of dimensions and specifications for the gas turbine engine with oneway air intake.

1.Diameter of the active rotor 2 - 425 mm.

2. Number of the active nozzles 24 in two rotor rows (stages) - 48

3. Number of the stator support blades 18 - 48

4. Number of force interactions per the rotor revolution - 2304

5. Dimension of the critical cross-sections of the rectangular nozzle 30*2.7 mm

6. Angle of the focusing part of the nozzle 24 - from 3°30'.

7. In the absence of suitable technologial equipment the engine power was approximately estimated using a reducer (c κθ3φφΗΐΓΗεΗτθΜ noHHaceHHa 10.5 in three equal stages of power rotation with subsequent noc e¾yroiipiM summation of the obtained values. With the number of rotor revolution of 21700rpm under proper load the engine power with the rated power consumption was about 7000 kW.

8. Engine weight, no more than -27 kg.

9. Relation of the engine unit power to unit at a minimum (similarly to the previous engine).

10. All other parameters correspond to the design parameters of the previous engine (with two-way intake).

From the claimed parameters of the gas turbine engine of the proposed design it follows:

The claimed gas turbine engines of the new design surpass all the available modern gas turbine engines in all parameters.

With that, considerable decrease in toxicity, significant reduction of the fuel consumption and significant rise in efficiency was achieved.

The gas turbine engine can be used in aviation as the engine has the relation power/weight many times better; higher power with the minimum fuel consumption per the unit power; lower power consumption and lower release of exhaust gases. With considerably better cooling of strength elements, the engine reliability is many times higher; operational lifetime also increases. The engine can be used in road transport (under conditions of high dust concentration), and on the thermal power plants, waterborne transport and other fields of national economy.

With application of the claimed original engine, high power can be achieved with environmental purity, and multiple decrease in the fuel consumption (under parameters which are unattainable for existing design of modern original engines being used at the present.

Application of the claimed gas turbine engines on thermal and nuclear power plants will provide:

1) multiple increase in output electric power of thermal and nuclear power plants.

2) multiple decrease of the fuel consumption.

3) multiple reduction in turbine dimensions with increase in the power.

4) multiple reduction of the the turbine production costs.

5) multiple increase in mechanical strength, and duration of continuous operation.

Comparative experimental studies in power and economic feasibility of the claimed gas turbine engines and modern devices existing currently, confirmed high performance of the gas turbine engine.

Industrial Applications

The present invention is implemented using the universal modern equipment available in the industry. All parts forming the structure of turbine are bodies of revolution; all the components of required length and diameter are made of the relevant modern steels can be manufactured on modern equipment.