RU2609547C1 - Return stage of rocket vehicle and method of its operation - Google Patents

Abstract

FIELD: transportation; aviation.

SUBSTANCE: method of operation of the return stage of the rocket vehicle includes its speed-up in the active area of the trajectory with the help of the LPS, control with the help of control chambers and return with the help of two GTEs. GTEs are started in the rarefied atmospheric slices using an auxiliary gas generator to compensate lack of atmospheric air for GTE operation. The return stage of a rocket vehicle comprises a body, oxidant and fuel tanks, wings, and at least one liquid propellant system. Two side units with GTE are attached to the body. Air intakes are available in the upper part of the side units. The GTE may contain a ring header upstream the main combustion chamber. The main combustion chamber may contain at least one ignition device. GTEs may be equipped with a nozzle with a controlled traction vector.

EFFECT: increased reliability of stages return and GTE operability at higher altitudes.

9 cl, 18 dwg

Description

The group of inventions relates to rocket technology and can be applied to reusable returnable space rocket systems capable of performing a manned flight in the atmosphere.

In aerospace engineering, the reusable orbiter Buran is known which contains the fuselage, the wing with two consoles, the left and right blocks of control engines located in the rear of the fuselage, and the bow block of control engines located in the nose of the fuselage ([1], p. 40, 41, 193). At the orbital site, the orbital ship is a payload for the launch vehicle (there are no cruise engines on the Buran orbital ship). After performing a space flight, the orbital ship makes a non-motor descent in the atmosphere (there are no jet engines), while the motion of the orbital ship around its center of mass during flight in the upper atmosphere is controlled by control engines located in the left and right blocks of the tail control engines parts of the fuselage. In this case, the axis of the nozzles of the pitch and roll control engines are perpendicular to the longitudinal axis (ОХ axis) of the orbital ship, form angles of 30 (with the normal axis (OY axis) of the orbital ship, and the axis of the nozzles of the yaw control engines are parallel to the transverse axis (OZ axis) of the orbital ship.

The disadvantage of this project is the inability to use its layout for a reusable missile unit. The control engine blocks cannot be placed either in the rear or in the nose of the fuselage, because in the rear part of the fuselage of the rocket block there is a marching propulsion system of the first stage of the launch rocket operating in the launch area, and in the nose of the fuselage of the return rocket block there are air-jet engines operating in the area of return of the return rocket block to the airfield in the launch area of the launch vehicle . The placement of control engine blocks in the middle of the fuselage is impractical, because in this case, the control motors will be ineffective due to the small shoulders of the control forces.

Another drawback of this project is the strong influence of the air flow on the gas jets of the control engines, especially on the jets of the yaw engines, the axis of the nozzles of which are oriented along the transverse axis OZ of the orbital ship perpendicular to the direction of flight. Finally, another drawback is the influence of the yaw force of the yaw engines that occur when they are triggered on the values of the transverse overload control system and sliding angle measured by the sensors. The control engines are inoperative in space and at high altitudes in a rarefied atmosphere, and the control moments at high altitudes are small due to the low thrust of gas turbine engines (GTE) at altitude.

Known launch vehicle according to the patent of the Russian Federation for the invention No. 2482030, IPC B64G 1/14, publ. 05/10/2013

This launch vehicle contains a reusable accelerator with a reactive stabilization system connected to the second stage, consisting of a rocket unit with liquid rocket engines (LRE), connected to an aircraft kit made in the form of a glider with variable sweep wings with aerodynamic control elements connected to the rocket block according to the “low-wing” scheme, stabilizer, landing gear, jet engines with their fuel tank, bow compartment, and also contains reusable elements, at The nose compartment mounted on the rocket block is equipped with a pilot's cabin and is equipped with controllable rotary visors, the number of which is equal to the number of points of connection of the nose section with the second stage, pockets are made in the places of connection with the second stage, air-propelled engines are mounted on the upper surfaces of the variable wings sweeps and equipped with controlled protective shields, the stabilizer is made in the form of two keels mounted on the wings, accelerate the reusable missile block I about its longitudinal axis and symmetrically with respect to its transverse axis parallel wings installed even number throttled rocket engines.

Disadvantages: poor aerodynamic performance of the rocket at launch due to the bulkiness of the fuselage and the presence of bulky wings, the inoperability of a gas turbine engine in high-altitude conditions and in space in the absence of air necessary for their operation, the uncontrollability of the return stage at high altitudes. The use of bulky, heavily weighted sweep wings is not justified due to the fact that the only task of creating a return stage is to land it, and not to perform complex maneuvers at subsonic and supersonic speeds.

Known launch vehicle with a return stage according to the patent of the Russian Federation for invention No. 2495799, IPC B64G 1/14, publ. 10/20/2013, the prototype of the device and method.

This launcher contains a reusable returnable missile unit, which, in turn, contains a fuselage, a wing with two consoles, left and right blocks of gas turbine control engines, left and right blocks of control engines are located in nacelles at the tips of the wing consoles. In addition, the return stage contains nozzles for pitch and roll control engines and yaw.

When the stage returns, the gas turbine engines are started at a relatively low altitude, for example 15,000 ... 20,000 m. The flight to this altitude is completely uncontrollable.

The disadvantage of this technical solution is also the low aerodynamic quality of the returned fuselage, i.e. the first stage, due to the placement of the gas turbine engine on the wing consoles, which have a significant thickness for the transmission of jet thrust and control torque. This leads to an unjustified deterioration in the characteristics of the launch vehicle at launch. In addition, the launch vehicle has a very sophisticated control system for pitch, yaw and roll angles.

The objectives of the invention are to improve the launch characteristics of the launch vehicle and simplify the control system by pitch, yaw and roll angles, ensure its operability at any altitudes and ensure a reliable landing of the return stage in any weather.

Technical results achieved - ensuring the operability of gas turbine engines at all heights and ensuring the landing of the return stage.

The solution to these problems has been achieved in the way the return stage of the launch vehicle operates, including its acceleration in the active part of the trajectory by means of a liquid-propellant rocket engine, control by means of steering chambers and return by means of two gas-turbine engines, in that gas-turbine engines are launched in rarefied atmospheric layers using an auxiliary gas generator, compensating for the lack of atmospheric air for gas turbine engine operation, and when flying in dense atmospheric layers, auxiliary gas generators are turned off.

The solution of these problems was achieved in the return stage of the launch vehicle containing the fuselage, oxidizer and fuel tanks, wings and at least one liquid rocket engine, in that two side blocks are attached to the fuselage, in which gas turbine engines, gas turbine engines, and the main chamber are mounted combustion and auxiliary gas generator connected to the main combustion chamber, and air intakes are made in the upper part of the side blocks.

A gas turbine engine may contain in front of the main combustion chamber an annular manifold to which a gas duct is connected, and the cavity of the annular manifold communicates with the air path through openings or nozzles. The gas turbine engine may comprise, in front of the main combustion chamber, an annular perforated manifold connected to the gas duct and installed inside the air duct. The gas turbine engine may comprise an annular manifold in front of the main combustion chamber. The main combustion chamber may contain at least one ignition device, the gas generator comprises at least one ignition device. The gas generator can be connected by pipelines of the oxidizer and fuel with a turbopump unit having fuel pumps, oxidizer and a turbine.

Gas turbine engines can be equipped with a nozzle with a controlled thrust vector.

The invention is illustrated in FIG. 1 ... 18, where:

- in FIG. 1 shows the appearance of the launch vehicle on the launch pad,

- in FIG. 2 shows the appearance of the returned 1st stage in the projection from the bottom, view A, in the process of landing,

in FIG. 3 shows the appearance of the returned 1st stage in the projection from the bottom, view A, during the flight,

- in FIG. 4 shows the appearance of a four-chamber rocket engine,

- in FIG. 5 shows a view of a single-chamber rocket engine

- in FIG. 6 shows a section aa,

- in FIG. 7 shows a section bb,

- in FIG. 8 shows the design of the main engine - LRE,

- in FIG. 9 shows the design of a gas turbine engine,

- in FIG. 10 shows a diagram of the supply of fuel components to a gas turbine engine,

in FIG. 11 shows a diagram of the supply of gas generating gas into the main combustion chamber, the first option,

- in FIG. 12 shows a diagram of the supply of gas generating gas into the main combustion chamber, the second option

- in FIG. 13 shows a diagram of the supply of gas generating gas into the main combustion chamber, the third option,

- in FIG. 14 shows a schematic diagram of a gas generator,

- in FIG. 15 shows view C,

- in FIG. 16 shows a diagram of a gas generator of the engine NK-33,

- in FIG. 17 shows a jet nozzle with an adjustable thrust vector,

- in FIG. 18 shows the oxidizer tank pressurization system.

The launch vehicle (Fig. 1 ... 18) may contain at least one stage. In the following, the description of the launch vehicle is compiled using the example of a two-stage rocket with one (first) return stage.

CARRIER ROCKET

The launch vehicle (Fig. 1 ... 18) contains a return stage 1 with a central fuselage 2, and two side blocks 3 connected coaxially to it.

The return stage 1 contains wings 4 attached to the side blocks 3, which are arrow-shaped and are installed in the middle part of the return stage 1. The return stage 1 (Figs. 1 and 2) contains two sets of propulsion systems: the main one, which is one or several liquid- rocket engines (LRE) 6 and at least two steering engines 7 in each side block 3 (engine of the type of small LRE) and two gas turbine engines (short-form gas turbine engines) - 8, installed one at a time in the lateral stages 3.

Liquid rocket engines 6 in the return stage 1 can be installed either one or more. It is possible to install one or more four-chamber rocket engines 6 (Fig. 3) or several single-chamber, for example 4 single-chamber rocket engines 6 (Fig. 2-6). Examples of four-chamber and single-chamber rocket engines 6 are shown in FIG. 4 and 5.

The return stage 1 contains the front chassis 9 and two rear chassis 10. The front chassis 9 at the time of start is inside the housing 2, and the rear chassis 10 is in the side blocks 3 and are released when landing (Fig. 2 and 3).

In the central fuselage installed one or more liquid propellant rocket engines 6 and steering engines 7 (Fig. 6)

A feature of the gas turbine engine 8 is the presence of an auxiliary gas generator 11, which is connected by a gas duct 12 to the gas turbine engine 8 (Fig. 7).

The auxiliary gas generator 11 is an important and necessary element of the gas turbine engine 8 to ensure its operability at any altitudes, its design and possible embodiments are described below.

The LRE 6 (Fig. 8) contains a combustion chamber 13 and TNA 14. The LRE 6 is mounted on the power frame 15 on the hinge 16. An oxidizer tank 17 and a fuel tank 18 are installed inside the central fuselage 2 (Fig. 7). An oxidizer pipe 19 is connected to the oxidizer tank 17, and a fuel pipe 20 is connected to the fuel tank 18.

Also, in the upper part of the side steps 3, air intakes 21 are made, connected by an air path to the entrances to the gas turbine engine 8 (Fig. 6).

LIQUID ROCKET ENGINE

The liquid-propellant rocket engine 6 (FIGS. 7 and 8) comprises a combustion chamber 13 and TNA 14, coupled to the combustion chamber 12 by means of a gas duct 22. A gas generator 23 is installed on the TNA 14.

TNA 14 contains, in turn, a turbine 24, an oxidizer pump 25, a fuel pump 26. A turbopump assembly 14 may include an additional fuel pump 27.

The output from the fuel pump 26 is connected by a pipe 28 to the entrance to the additional fuel pump 27 (if any). The combustion chamber 13 comprises a head 29, a cylindrical part 30 and a nozzle 31.

A possible pneumohydraulic scheme of the LRE is shown in FIG. 8 and contains a fuel pipe 32 connected at one end to the outlet of the fuel pump 24 containing a start-off valve 33, the outlet of this pipe is connected to the main manifold 34 of the combustion chamber 12. The output from the oxidizer pump 23 by the oxidizer pipe 35 containing the oxidizer start-up valve 36 is connected with a gas generator 21. Also, the output from the additional fuel pump 25 is a fuel pipe 37 containing a start-off valve 38 of the fuel connected to the gas generator 21. On the gas generator 21, on the combustion chamber 12 is installed The at least one ignition device 39 and 40. The ignition device 39 and 40 are connected to power connections 41 with power unit 42.

The engine is equipped with a control unit 43, which is connected by electrical connections 44 to the power unit 42 and to the shutoff valves 33, 36 and 38. A feature of the liquid propellant rocket engine 6 is that the TNA 14 is rigidly fixed to the combustion chamber 13.

To control the pitch, yaw and roll angles at launch and in flight of the launch vehicle, at least two steering engines 7 installed in the side blocks 3 are used. That is, no less than 4 steering engines 7 are used on the launch vehicle, which is sufficient for management. In fact, the steering engines 7 are also marching when the returning stage 1 returns to the ground.

The design and power scheme of the oxidizer and fuel steering engines 7 are shown in FIG. 8. The main feature of the steering engines 7 is that each of them has its own small-sized TNA 45, to which an oxidizer pipe 46 and a fuel pipe 47 are connected. An oxidizer pipe 48 with an oxidizer shut-off valve 49 and a pipe are connected to the outlet from the small-sized TNA 45. high-pressure fuel 50 with a fuel shut-off valve 51. The steering engine 7 is mounted on an articulated suspension 52 with the possibility of swinging in the same plane to control the pitch and roll angles. At the yaw angles, the control of the launch vehicle in the active flight section is carried out by the mismatch of the steering engine rods 7.

The steering engines 7 have their own ignition devices 53 for re-activation when returning the return stage 1 and are used simultaneously as cruising in a vacuum and to control the flight and stabilize the position of the return stage 1 when it returns to the ground for reuse.

GAS TURBINE ENGINE

Gas turbine engines 8 (Fig. 9) are installed in the side blocks 3 and includes a housing 54, an inlet device 55, a compressor 56, an air duct 57, a main combustion chamber 58, a turbine 59, and a jet nozzle 60. The compressor 56 comprises guide vanes 61 and impellers 62, the turbine 59 contains nozzle apparatuses 63 and impellers 64. The compressor 56 and turbine 59, more precisely their impellers 62 and 64 are connected by a shaft 65. The shafts 65 may be two or three depending on the design of the gas turbine engine 8. The shaft 65 is mounted on supports 66 and 67.

The main combustion chamber 58 (Fig. 9) contains a flame tube 68, a nozzle plate 69 with fuel nozzles 70 and a fuel manifold 71. Under the flame tube 68, an inner casing 72 is installed, between which and the flame tube 68 there is an internal channel 73. Between the flame tube 68 and an external channel 74 is made by the housing 54. The internal and external channels 73 and 74 are for introducing air (or gas-generating gas) from the air path 57 into the flame tube 68 through openings 75 made therein, as well as for cooling the flame tube 68 itself.

The gas turbine engine 8 (Fig. 10) has a fuel supply system comprising a low pressure fuel line 76, a fuel pump 77 having a drive 78, a high pressure fuel line 79, the input of which is connected to the fuel pump 77, and the output is connected to the fuel manifold 71, which is connected to fuel nozzles 70 of the main combustion chamber 58.

The gas turbine engine 8 is equipped with fuel and oxidizer supply systems from the oxidizer tank 16 and the fuel tank 17 and an additional turbopump unit 80. The turbopump unit 80 may be absent, and the fuel and oxidizer can be supplied to the turbine engine 8 using a propellant supply of rocket fuel components (this is possible by very high altitudes). In addition, the gas turbine engine 8 is necessarily equipped with an auxiliary gas generator 11, which gas duct 12 is connected to the air path 57 to the main combustion chamber 58.

The additional turbopump assembly 80 includes a housing 81, a fuel pump 83, an oxidizer pump 84, and a turbine 85 installed on the shaft 82. The outlet of the oxidizer tank 16 is connected by an oxidizer pipe 86 containing the oxidizer valve 87 to the oxidizer pump 81, and the exit from the fuel tank 17 by the fuel pipe 86 containing a fuel valve 89 is connected to the inlet to the fuel pump 84. The exit from the oxidizer pump 83 by the oxidizer high pressure pipe 90 containing the shut-off valve of the oxidizer 91 is connected to the inlet to the gas generator 12. The output from the fuel pump its 84 high-pressure fuel pipe 92 containing a shut-off valve 93 and a flow regulator 94 is connected to the inlet of the gas generator 11. The output of the auxiliary gas generator 11 is connected to the inlet to the turbine 85, and the outlet from the turbine 85 by the gas duct 12 is connected to the mixing manifold 95 with gas nozzles 96 installed in the air duct 57 in front of the main combustion chamber 58.

Several options for the connection of the gas duct 12 with the air path 57 are possible (Fig. 11 ... 12).

In FIG. 11 shows the first embodiment of connecting the gas duct 12 to the air path 57. An annular manifold 97 is made on the housing 54 of the gas turbine engine 8 in the area of the air path 57, the cavity 98 of which is connected by holes 99 to the air path 57.

In FIG. 12, the second variant is also provided with an annular manifold 97, the cavity 98 of which is connected by openings 99 to the radial nozzles 100, which are perforated along the entire height with openings 101 for more uniform introduction of the generator gas into the air passing through the air duct 57.

In FIG. 13 shows the third option. According to this option, an internal annular manifold 102 having a cavity 103 and openings 104 is installed in the air path 57. A gas duct 12 is connected to the inner annular manifold 102. This option provides a more uniform supply of generator gas to the main combustion chamber 58. This is necessary to ensure uniform temperature field at the entrance to the turbine 59 and prevent local overheating of the parts of the impellers 66.

Auxiliary gas turbine engine generator

For the proposed gas turbine engine 8, the auxiliary gas generator 11 may be specially designed or used brought the gas generator of liquid rocket engines NK-33 or RD-170. A schematic diagram of the auxiliary gas generator 11 is shown in FIG. 14 and 15. The auxiliary gas generator 11 is designed to burn fuel components (fuel and oxidizer), while one of them is an excess component, and the second is an additional component. It is most preferable to use kerosene as fuel, and oxygen as an oxidizing agent.

The auxiliary gas generator 11 comprises (Fig. 14) a head 105, a chamber 106, an oxidizer (excess component) distributor 107 mounted along the axis of the chamber 106.

The chamber 106 contains two zones: a combustion zone 108 and a mixing zone 109. The first of them is intended for the combustion of two components at the optimal ratio, and the second for mixing the oxidizing agent.

The head 105 comprises a front bottom 110 with a fuel supply nozzle 111, a middle bottom 112, a fire bottom 113, an oxidizer nozzle 114, a fuel nozzle 115. A cavity 116 is formed between the front 110 and the middle 112 bottom for supplying fuel to the fuel nozzles 115, and between the fire bottom 113 and a middle bottom 112, a cavity 117 is formed for supplying the oxidizing agent to the oxidizer nozzles 114. On the average bottom 112, grooves 118 are made for supplying the oxidizing agent to the cavity 117. The chamber 106 of the auxiliary gas generator 11 includes an outer casing 119 and an inner shell 120 between there is a gap 121 for the passage of the oxidizing agent. Openings 122 are made on the oxidizer distributor 107 for supplying an excess component to the mixing zone 108. Oxidizer pipe 123 is made along the axis of the chamber 106. The main combustion chambers 58 have ignition devices 124 (Fig. 11 ... 13), and auxiliary gas generators 11 have ignition devices 125 ( Fig. 14).

Description of the gas generator engine NK-33, used as an auxiliary

As mentioned earlier, for the proposed gas turbine engine 8 can be applied to the gas generator of the engine NK-33. A detailed description of the auxiliary gas generator 11 of the NK 33 engine is given in the RF patent for the invention No. 2179256, IPC A02K 9/24, publ. 02/10/2002

The following is a description of the gas generator of the NK-33 engine (Fig. 14 ... 16). The auxiliary gas generator 11 has a head 105. The oxidizer distributor 107, located along the axis of the auxiliary gas generator 11, comprises a cylinder 126 with an oxidizer cavity 127, mixing elements 128 and 129 in the form of hollow cylinders 130, closed by tent heads 131 and perforated by openings 132. In front of each mixing element Apertures 133 are made in holes 128 and 129. The mixing elements 128 and 129 are staggered and their height decreases in gas flow.

Between the firing bottom 113 and the mixing elements 128 and 129, radial perforated plates 134 with oxidizer supply channels 135 from the cavity 127 into the cavity of the chamber 108 of the auxiliary gas generator 11 can be arranged.

The oxidizer distributor 107 is closed by the bottom 136 in the form of a truncated cone facing the apex toward the firing bottom 113, and holes 137 and 138 are made at the transition point of the cylinder 126 in the bottom 135 and at the top of the truncated cone.

At least 105 ignition device 125 is mounted on the head 105 at an angle to the axis of the auxiliary gas generator 11, which is connected by energy connections 41 to the energy block 42 (or to the pumping unit for laser ignition candles).

Jet nozzle

On a gas turbine engine 8, a jet nozzle 60 with a controlled thrust vector can be used (FIG. 17). Such nozzles are known, for example, from patents of the Russian Federation for utility models No. 21220, IPC F02K 1/05, publ. December 27, 2001 and No. 105683, IPC F02K 1/12, publ. 12/27/2010, however, their use in rocketry was proposed for the first time. The jet nozzle 60 comprises flaps 140, hydraulic cylinders 141 connected to them, and a cooling channel 142 for cooling the hydraulic cylinders 141 controlling the thrust vector of the jet nozzle 60.

TANK CHARGING SYSTEM

The oxidizer tank pressurization system 17 is shown in FIG. 18 and comprises a boost pipe 143 with a boost valve 144. The boost pipe 143 is connected to the outlet of the auxiliary gas generator 11, i.e. pressurization is carried out by a gas-generating gas containing 90 ... 95% oxygen. Fuel tank pressurization is carried out by helium. The pressurization system of the fuel tank 18 in FIG. 1 ... 18 is not shown.

GAS-TURBINE ENGINE OPERATION AT STARTING AND RETURNING THE RETURNED STAGE OF THE ROCKER-CARRIER

When the launch vehicle is launched, all engines are simultaneously launched: the liquid propellant rocket engine 6, the steering engines 7 and the gas turbine engine 8. However, the auxiliary gas generators 11 serving the gas turbine engine 8 do not work (Fig. 1).

When the launch vehicle enters the rarefied atmosphere (at an altitude of 10 ... 30 km), auxiliary gas generators 11 are included to compensate for the lack of air entering the gas turbine engine 8.

When the launch vehicle enters the vacuum, the GTE 8 is turned off. LRE 6 and steering engines 7 continue to work until the completion of the flight program in the active section of the trajectory, after which they turn off. The higher stages or the payload are disconnected, after which the steering engines 7 are turned on and the return stage 1 is directed to the ground. At the entrance to the rarefied layers of the atmosphere (at an altitude of 30 to 10 km), the gas turbine engine 8 is turned on with the simultaneous inclusion of auxiliary gas generators 11 serving them.

To do this, launch two GTE 8 by supplying electricity to the starter from an energy source (not shown). At the same time, the drive 78 of the fuel pump 77 is turned on (Fig. 91), and the fuel pump 77 delivers fuel to the fuel manifold 71 of the main combustion chamber 58 and then through the fuel nozzles 70 into the flame tube 68, where it is ignited by the ignition device 125. The impellers 64 turbines 59 are untwisted and untwisted through the shaft 65 by the impellers 62 of the compressor 56. The jet nozzle 60 generates thrust.

Almost simultaneously or after 20-30 s, auxiliary gas generators 11 serving the gas turbine engine are started 8. For this, the shut-off valves 87 and 89 are opened and oxidizer and fuel are supplied to the auxiliary gas generators 11, where they are ignited using an ignition device 125 (Fig. 9). The gas-generating gas passes through the turbine 85 and enters through the mixing manifold 95 and nozzle 96 into the air duct 57 in front of the main combustion chamber 58 to compensate for the lack of air for normal operation of the turbine engine 8. In this mode, the steering engines 7 are controlled by the return stage 1.

Upon transition of the return stage 1 to the dense layers of the atmosphere (at an altitude of 3 ... 4 km), the auxiliary gas generators 11 are turned off and the turbine engine 8 continues to operate using only the high-pressure head of atmospheric air entering through the air intakes 21 (Figs. 7 and 9).

If a thrust vectoring nozzle 60 is used (FIG. 17), it can be used to control the return stage. In this case, the steering engines 7 are turned off.

Changing the operating mode of the gas turbine engine 8 in high-altitude conditions is carried out by the flow regulator 93 (Fig. 10), and when flying an aircraft equipped with such an engine, in dense atmospheric layers using the drive 78 of the pump 77. The fuel and oxidizer can be supplied to the auxiliary gas generator 11 be significantly reduced or completely discontinued.

When the first stage moves to denser atmospheric layers, the auxiliary gas generator 11 is turned off, for this, shut-off valves 87 and 89 are turned off and the oxidant and fuel supply is cut off and the gas turbine engine 8 switches to the use of atmospheric air as an oxidizer, which is more economical.

To control in flight and partially in the mode of returning the first stage 1 of the launch vehicle, steering engines 7 are used (Fig. 18).

Landing of the return stage 1 is performed on the runway at the aerodrome. To do this, produce the chassis 9 and 10 (Fig. 2). For the final shutdown of the gas turbine engine 8 after landing the first stage 1, the fuel supply to the pump 77 is stopped.

Naturally, in the case of the use of multi-stage return rockets, not only the first stage, but also the subsequent stages can be performed.

The application of the invention allowed:

1. Ensure the reliable return of one or more stages of launch vehicles for reuse.

2. Ensure the operability of a gas turbine aircraft engine at very high altitudes (more than 30,000 m in space).

3. Significantly increase the thrust of the gas turbine engine through the use of auxiliary gas generator.

4. To improve the reliability of starting a gas turbine engine, especially in high-altitude conditions due to the use of hot gas-generating gas and reusable ignition devices when starting up.

4. To ensure multiple launch of the rocket engine and gas turbine engine through the use of reusable ignition devices on them (electric or laser).

5. Ensure control of the launch vehicle in pitch, yaw and roll angles both in the active section of the trajectory and upon return.

LITERATURE

1. Yu.P. Semenov, G.E. Lozino-Lozinsky, V.L. Lapygin, V.A. Timchenko and others. Reusable orbiter "Buran". M., "Engineering", 1995, 448 S.

Claims (9)

1. The method of operation of the return stage of the launch vehicle, including its acceleration in the active part of the trajectory by means of a liquid fuel rocket engine and control by means of steering chambers and return by two gas turbine engines, characterized in that the gas turbine engine is launched in rarefied atmospheric layers using an auxiliary gas generator that compensates for the shortage atmospheric air for gas turbine engine operation, and when flying in dense atmospheric layers, auxiliary gas generators are turned off. 2. The return stage of the launch vehicle containing the fuselage, oxidizer and fuel tanks, wings and at least one liquid rocket engine, characterized in that two side blocks are attached to the fuselage, in which gas turbine engines are installed, gas turbine engines have a main combustion chamber and a gas generator connected to the main combustion chamber, and air intakes are made in the upper part of the side blocks. 3. The return stage of the launch vehicle according to claim 2, characterized in that the gas turbine engine contains an annular manifold in front of the main combustion chamber, to which the gas duct is connected, and the annular manifold cavity communicates with the air path through openings or nozzles. 4. The return stage of the launch vehicle according to claim 2, characterized in that the gas turbine engine comprises, in front of the main combustion chamber, an annular perforated manifold connected to the gas duct and installed inside the air duct. 5. The return stage of the launch vehicle according to claim 2, characterized in that the gas turbine engine comprises an annular collector in front of the main combustion chamber. 6. The return stage of the launch vehicle according to claim 2, characterized in that the main combustion chamber contains at least one ignition device. 7. The return stage of the launch vehicle according to claim 2, characterized in that the gas generator comprises at least one ignition device. 8. The return stage of the launch vehicle according to claim 2, characterized in that the auxiliary gas generator is connected by pipelines of the oxidizer and fuel to a turbopump unit having fuel pumps, an oxidizer and a turbine. 9. The return stage of the launch vehicle according to claim 2, characterized in that the gas turbine engines are equipped with a nozzle with a controlled thrust vector.

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