RU2524483C1 - Liquid propellant rocket engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: proposed rocket engine has four chambers secured at the frame, turbopump unit secured thereto. Said unit comprises turbine, oxidiser and fuel pumps, heat protection, lines for oxidiser and fuel feed to gas generator and combustion chamber. In compliance with this invention, said frame is an all-welded 3D structure composed by fender and bottom frames interconnected by rods. Note here that crossbars with bearings are secured to bottom frame, journals of chambers being fitted in said bearing to turn about swing axis. Besides, the engine comprises four bent oxidiser gas feed lines with their common manifold connected with turbine outlet. Other eight bends are connected to appropriate heads of the chambers. Note here that sets of flexible pipes with bellows are arranged in said lines perpendicular to said swing axis. Said bellows are connected with fixed part of the line by one butt and, by opposite one, with movable part incorporated with unit of chambers swinging in one plane.

EFFECT: more efficient control over thrust vector, higher specific thrust pulse, perfected design.

12 cl, 28 dwg

Description

The invention relates to rocket technology and can be used in the manufacture of rocket launchers with a four-chamber liquid rocket engine.

In rocketry, multi-chamber liquid-propellant rocket engines are known and widely used for use in various types of launch vehicles, both manned and unmanned systems.

Known liquid rocket engine (LRE), which includes two chambers mounted on the frame, attached to the frame of a turbopump unit having a turbine, oxidizer and fuel pumps, pipelines for supplying oxidizer and fuel to the gas generator and engine chambers. The engine frame is made collapsible and contains two mirror-symmetric sections, has supports for the perception of efforts from the engine and the heel for attaching the frame to the rocket body. The engine contains a bent bent pipeline for supplying high-temperature oxidizing gas, a single end section, which is connected to the turbine outlet, and the other two elbows, by corresponding chambers through bellows expansion joints, which are the rocking nodes of the chambers (RF Patent No. 2158838, IPC F02K 9/42, publ. 10.11 .2000 g.).

The disadvantages of this technical solution are:

- asymmetric arrangement of the turbopump unit (TNA) relative to the chambers in the plan (it is shifted to the side), which in turn causes a displacement of the center of mass of the engine relative to the axis of the rocket and worsens its mass-centering characteristics;

- the openness of the engine frame along its mating plane, as a result of which its strength and rigidity decreases when the engine is located autonomously (outside the rocket) during transportation and installation works.

Known liquid rocket engine, which contains a symmetrically located four chambers 1 relative to the axis N of the engine, placed between them a turbopump unit 2 along the axis of the engine, mounted on the frame 3, pipelines for the supply of oxidizer and fuel to the gas generator 4 and chamber 1. The power element of the engine, perceiving and dynamic loads, is frame 3, which is an all-welded spatial truss, consisting of mating 5 and lower 6 frames, as well as tubular rods 7 between and. Four chambers 1 are bolted to the ring supports 8 with special bosses 9 for the brackets 10. The engine is made according to an open circuit, the thrust vector is controlled by steering nozzles 11 mounted on the heat protection frame 12 (working drawings 11D55-00-02, technical description 11D55- 00TO engine 11D55 development JSC KBHA, Voronezh, see figure 1, prototype). Figure 2 shows a top view of the engine.

The known engine is made with a large degree of pressure reduction on the turbine, because the exhaust gas after the turbine enters the thrust vector control nozzles, after which it is released into the atmosphere.

The gas is reducing, the pressure in the chamber is not high, about 70 ÷ 80 kg / cm ² , therefore, due to the above reasons, the thrust vector control efficiency and specific thrust impulse of the engine are low.

The disadvantage of this technical solution is that the engine chambers are rigidly fixed to the frame and cannot rotate, changing the direction of the thrust vector, therefore special steering nozzles are used, mounted on the thermal protection frame.

The engine frame 3 is made in the form of a spatial truss, in which the tubular rods 7 between the mating 5 and the lower 6 frames form non-rigid figures, such as trapezoids or quadrangles, therefore, the rigidity of the frame is significantly reduced. To increase the rigidity and strength of the frame of the known technical solution, it was necessary to weld a large number of box-shaped headscarves between racks and frames, as well as supporting bosses.

The turbopump unit 2 of the known engine is fixed to the rocket tank directly by the input "O" 14, there is no space for accommodating the ejector or booster pump unit and start valves. At the entrance of "G" to the turbopump, the above units are also not installed, the spatial gap for their placement on the engine is not enough. In connection with this circumstance, the pressure at the pump inlets is low: 3.5 kg / cm ² for the fuel pump and 5.5 kg / cm ² for the oxidizer pump, so the anticavitation reserves of the pumps are small. The layout of the engine - prototype has an axial overall dimension, which does not allow placing pre-pumping units at the inputs "O" and "G" in the overall dimensions of the tail section of the rocket block.

The gap between the nozzle of the chamber and the bottom heat shield of the prototype engine is closed by a seal 15 in the form of silica fabric and mulite-silica felt in a shell made of heat-resistant rubberized fabric.

The disadvantages of the used seal material are fragility and a sufficiently large weight, which makes this material inapplicable for installation between the movable and fixed parts of the thermal protection of the proposed engine.

To ignite the fuel components in the gas generator and the chambers, igniters 16, 17 with pyro cartridge are used. Due to the insufficient reliability of ignition and the power of charges, the number of igniters is doubled, which somewhat complicates the design of the launch system. In addition, access to the ignition system on the engine mounted in the rocket block is difficult through the hatches of the latter.

In the lines of fuel, oxidizer and gas supply to the steering nozzles of a known engine, flexible elements are installed containing bellows 18, which at the joints are flanges fastened together by studs to prevent their destruction from internal pressure during engine operation.

The above flexible elements do not allow bending of the bellows to an angle of 3.5 ° ÷ 5 ° or more, and therefore their use is impossible on the swinging unit of the camera.

Fixing the steering nozzles of the engine - the prototype in the zero position is possible only if electric current is supplied to the drive 19, connected by an adjustable rod 20 to the lever 21 of the steering nozzle 11.

In the absence of electric voltage on the drive, the steering nozzles are not fixed, therefore, under certain conditions, their integrity may be violated.

A fuel manifold 22 is mounted on the nozzle of the engine chamber to which a plate with two holes (not shown) is welded. When ground firing, sections of the nozzles of the chambers are fastened together by rods connected to the plates with bolts. The engine nozzles are high-altitude, made with a high degree of gas expansion, operate off-balance during ground firing, therefore they experience increased vibrational dynamic loads, however, collectors and fastenings to the bench frame are absent over the length of the nozzle, which reduces the reliability of the chambers from loss of bearing capacity and destruction .

Drainage pipelines 23 of the engine are discharged for bottom thermal protection 13, are not grouped together, drainage boxes are not provided, therefore components are drained into the atmosphere spontaneously, not by command from a certain point in time. In addition, combining the drainage of components is impractical in a firefight.

Soft thermal insulation is installed on the lines of the engine oxidizer, and cryogenic thermal insulation, which is applied to the surface of pipelines and units in the form of foam and which is more reliable, is absent, and therefore the insulating properties of the general thermal insulation are somewhat reduced.

The engine does not have a specially designed heating system for control and control units, the functioning of which depends on the operation of the engine at sub-zero temperatures in the tail section of the rocket unit, which is especially important at engine start, and warm air comes from already running engine chambers.

The disadvantage of this solution is, as already mentioned above, the lack of heating of the fuel components in the engine lines, which is an unfavorable factor for ensuring the cyclogram of its start.

In connection with the implementation of the prototype engine in an open circuit and with low pressure in the combustion chamber, the cylindrical part of the combustion chamber is shortened in it, which is unacceptable for engine chambers in a closed circuit with afterburning of the generator gas in them, requiring longer residence time of the fuel components to complete them combustion.

The objectives of the present invention are to increase the thrust vector control efficiency, increase the specific thrust impulse, improve the engine's power structure, optimize the use of the space of the tail section of the rocket block by the aggregates while the axial dimension of the engine remains constant, increase the protection of the engine units from high-temperature gases flowing out of the nozzles of the chambers, and ensure reliable starting engine and ease of loading the tubes of the gas generator and chambers with starting ampoules, providing fuel supply components, purging, draining the working media between the units mounted on the fixed and moving parts of the engine, ensuring that the chambers are in the zero position, and that they are safe during transportation of the engine under fire bench testing in terrestrial conditions and when the chambers operate in an off-design mode with under-expansion of gases in nozzles, increasing the fire safety of the working engine, improving the quality of the oxidizing agent, by eliminating its boiling at engine start and bubbles entering the BTNAO, as well as protecting ag regattas from cooling by the oxidizing agent, improving the conditions for starting the engine and normal operation at sub-zero ambient temperatures, increasing the residence time and completeness of combustion of fuel components in the combustion chambers, reducing the mass and increasing ductility and reliability of the entire engine chamber rotation system.

The tasks in the proposed technical solution are achieved by the fact that a liquid propellant rocket engine, including four chambers mounted on a frame, a turbo pump unit having a turbine, oxidizer and fuel pumps, thermal protection, oxidant and fuel supply pipelines to the gas generator and engine chambers, according to the invention, contains a frame made in the form of an all-welded spatial truss, consisting of mating and lower frames, interconnected by rods, while to the lower joint a traverse with bearings is attached to the nogout, into which the trunnions of the chambers are inserted to rotate around the swing axis, in addition, the engine contains four curved oxidizing gas supply lines, a single end collector of which is connected to the turbine outlet, and eight other nozzles are connected to the corresponding chamber heads, moreover in the lines perpendicular to the axis of swing of the chambers are blocks of flexible piping with bellows, one joint connected to the fixed part of the line, and the other with its moving part, which is included in the swing Keep the camera unit in the same plane.

An important circumstance is that the oxidizing gas lines, symmetrically located between the turbopump unit and the frame traverses, are involved in the perception of efforts from the internal traverses, which are affected by the camera’s traction force. In addition, the oxidizing gas pipelines are made with two bends of the fixed part to compensate for the temperature displacement of up to 4–5 mm caused by their heating from the passing gas.

A hole is made in the center of the engine’s thermal protection, in which a part of the turbopump assembly, protected by a conical casing mounted on the outer side of the thermal protector, is removed outside the thermal protection towards the nozzle exit of the chambers, a bellows compensator is installed between the tank flange at the O input and the exhaust manifold of the turbopump and in the lines of entry "O" and "G" are installed valves for starting the fuel components and booster turbopump units.

To ensure the axial displacement of the THA from the effects of thermal displacements of the oxidizing gas lines and to exclude the radial movements of the "G" TNA pumps from the vibrations of the working engine, lugs are mounted on the bosses of the heat shielding ring, inside of which spherical eccentric sleeves are mounted, which are in contact with the eccentric pin fixed to the threaded part on the support TNA bracket.

In addition, a composite thermal protection installed on the engine, which consists of four spherical shields mechanically mounted on the flanges of the nozzles of the chambers and swinging with them, and a flat stationary screen polished to a mirror shine, is peripherally fixed by braces to the frame, and the central part is mounted to the ring mounted on the THA, while between the spherical protection and the screen there is a sealing ring made with cuts to increase its compliance.

For the chemical ignition of fuel components in the gas generator and combustion chambers, the engine blocks of the gas generator and chambers with starting fuel are installed on the engine, which are mounted on the gas generator and stretching the heat shield to the frame, respectively, while the tubes of the ampoule blocks are tilted and directed to the periphery of the tail compartment of the block rockets.

On the camera blocks between the fixed and moving parts, flexible pipe blocks are mounted perpendicularly and symmetrically to the camera swing axis and fixed to the frame traverses on both sides of the camera’s critical section. Through flexible pipelines, fuel components and working fluids of purges, controls or drains from the units are supplied or discharged.

The gas bellows are made cooled, and the flexible elements of small internal diameters are made of steel tubes.

During transportation, installation in the rocket block, as well as at the starting position, the engine chamber blocks are installed in the zero position using technological and pyrotechnic clamps that are inserted and fixed into the holes in the frame plates and the eyes of the combustion chamber.

On the cylindrical part, the fuel manifolds and the nozzle section of the chambers, bandages with bosses are mounted for welding the rods of the engine to the bench frame.

The fuel and oxidizer drainage pipelines are grouped together by their outlets and connected to drainage boxes mounted on the bottom heat shield and the booster turbopump, respectively.

Cryogenic thermal insulation is applied on the surface of the bellows expansion joint and the start valve of the O input line, and heat-insulating covers are mounted on the bellows.

Heat-insulating covers are installed on the block of launching cylinders, the regulator, the blocks of ampoules of the gas generator and combustion chambers with starting ampoules, the fittings of which are connected to the pipelines of the heating system for supplying warm air from the rocket.

The length of the cylindrical part of the combustion chamber is maximally increased within the frame dimensions, while the gas ducts on the chamber head are made in the form of two nozzles to provide the necessary clearance with the spherical bottom of the rocket oxidizer tank.

The power supply pipelines of the steering drives of the camera rotation system are not located around the entire perimeter of the frame, but at the points of approach to the steering drives they are made of a “lyre-like” configuration, with compensation bending.

Figure 3 presents the main view of the proposed engine;

Figure 4 shows a top view of the engine;

Figure 5 shows the main view of the frame of the rocket engine;

Figure 6 shows a top view of the frame;

Figure 7 shows a view of a block of a turbopump assembly (TNA);

On Fig presents a section along the traverse with bearings;

Fig. 9 shows oxidation gas supply lines between the TNA collector and chamber heads;

Figure 10 shows the placement of the engine in the tail compartment of the PH between the flange of the tank "O" and thermal protection;

Figure 11 shows a remote element with a heat shield to the TNA;

12 shows thermal protection of the engine;

On Fig shows a section at the contact point of the movable and fixed parts of the thermal protection;

On Fig shows a view of the location of the blocks of ampoules of the gas generator and chambers on the engine;

On Fig shows a view of the gas generator and its ampoule block;

On Fig shows the installation of technological and pyrotechnic clamps of the zero position of the cameras in the transport position;

On Fig shows the installation of pyrofixer in the starting position;

On Fig, 19, 20, 21 shows the installation of bandages on the cameras for mounting to the stand;

On Fig shows a cross section of the drainage box "G" on thermal protection;

On Fig shows the location of the drainage box "O" on BTNAO thermal protection;

On Fig presents a section of the input block "O" with applied thermal insulation;

In Fig.25 shows a diagram of a heating system of units;

On Fig shows the relative position of the gas ducts of the camera head and the bottom of the tank "O" with applied thermal insulation;

On Fig presents a system of rotation of the engine chambers (SPKS).

On Fig shows a remote element with mounting piping SPKS to the frame.

The proposed LRE is presented in the indicated figures 3-28, and its main elements are the following:

1 - camera;

2 - frame;

3 - turbopump unit (TNA);

4 - turbine ТНА (in the case);

5 - oxidizer pump;

6 - fuel pump;

7 - pipeline supplying the oxidizer to the gas generator;

8 - pipeline for supplying fuel to the chamber;

9 - gas generator (GG);

10 - mating frame of the frame;

11 - lower frame of the frame;

12 - rods;

13, 14 - traverses;

15 - bearings;

16 - trunnion cameras;

17 - line supply of oxidizing gas;

18 - collector;

19 - turbine outlet;

20 - pipe;

21 - camera head;

22, 23 - blocks of flexible pipelines;

24, 25 - bellows;

26, 28 - joints of the flexible piping block;

27 - fixed part of the highway;

29 - the moving part of the highway;

30 - thermal protection;

31 - conical casing;

32 - flange;

33 - oxidizer tank;

34 - bellows compensator;

35, 36 - start valves "O" and "G";

37, 38 - booster turbopump units "O" and "G" (BTNAO, BTNAG);

39 - spherical defense;

40 - flange of the nozzle of the chamber;

41 - fixed screen;

42 - heat protection ring;

43 - sealing rings;

44 - block ampoules of a gas generator (GG);

45 - block ampoules cameras (COP);

46 - stretching;

47 - the moving part of the camera unit;

48 - fixed part of the camera unit;

49, 50 - flanges of flexible piping blocks;

51 is a critical section of the camera;

52 - technological lock;

53 - pyrotechnic retainer;

54 - plate frame with holes and fixtures for the retainers;

55 - camera eye with holes for the clips;

56 - a collector of a combustible nozzle of the chamber;

57 - cut nozzles of the chambers;

58, 59 - bandages on the nozzles of the chambers;

60 - bosses;

61, 62 — drainage boxes “G” and “O”, respectively;

63 - membrane;

64 - squib;

65 - stock;

66 - cover;

67 - cryogenic thermal insulation;

68, 69 - heat-insulating covers of the bellows compensator;

70, 71, 72, 73 - heat-insulating covers of the heating system;

74 - fittings for heat-insulating covers;

75 - pipelines;

76 - throttle washers;

77 - pipelines of the camera rotation system;

78 - steering gear;

79 - bosses of a ring of thermal protection;

80 - eye;

81 - spherical eccentric sleeve;

82 - eccentric finger;

83 - supporting bracket TNA;

84 - dead end section of pipelines SPKS;

85 - streamers attaching the input "O" to the frame;

86 - streamers attaching the entrance "G" to the frame;

87 - stretching the fastening ring 42 thermal protection to the frame;

88 - the base of the frame;

89 - thermal insulation of the oxidizer tank;

90 - "lyre-shaped" pipe sections to the steering gear;

91 - bandage on the cylindrical part of the chamber;

92 - bench frame;

93 - thrust of fastening the engine to the bench frame;

94 - pipe for cooling the bellows;

95 - steel tubes;

96 - block starting cylinders;

97 - fuel valve on the camera head;

98 - pyropatron pyrophorix;

99, 100 - hinged supports of the BGT;

101 - spacer;

102 - units of the camera rotation system (SPKS);

103 - flange of the tank "O" PH;

104 - frame bosses;

M is the axis of swing of the camera;

D - hole in the thermal protection;

L is the length of the cylindrical part of the chamber;

Δ is the gap between the gas ducts of the chamber and the bottom of the block "O";

P - mating plane of the frame frame.

A liquid rocket engine contains four chambers 1 mounted on a frame 2 (Figs. 3, 4). A turbopump unit 3 is attached to the frame 2 (see FIGS. 5, 6), having a turbine 4, oxidizer 5 and fuel 6 pumps (see FIG. 7), oxidant 7 and fuel 8 supply pipelines to the gas generator 9 (see FIG. 3, 4) and camera 1 of the engine.

Frame 2 (see Figs. 5, 6) is made in the form of an all-welded spatial truss and contains a mounting frame 10 and a lower frame 11 interconnected by rods 12, while the yoke 13, 14 with bearings 15 are attached to the lower frame 11 (see Fig. 8) into which the pins 16 of the chambers 1 are inserted in order to enable their rotation around the swing axis “M”.

The engine contains four curved oxidizing gas supply lines 17 (see Fig. 9), a single end manifold 18 of which is connected to the turbine outlet 19, and eight other nozzles 20 are connected to the corresponding heads 21 of the chambers 1, moreover, in the indicated lines perpendicular to the swing axis "M "There are blocks of flexible pipelines 22, 23 with bellows 24, 25, connected by one joint 26 to the fixed part 27 of the highway, and the other joint 28 is fixed with its movable part 29 included in the camera unit.

In the center of thermal protection 30 (see FIG. 11), a hole “D” is made, into which part of the turbopump assembly 3, namely the fuel pump 6, is removed outside the thermal protection towards the cut nozzles of the chambers 6. The fuel pump is protected by a conical casing 31 mounted on the outside thermal protection side. Such a technical solution for displacing the turbopump assembly towards the cut of the nozzles of the chambers made it possible to obtain free space between the flange 32 of the oxidizer tank 33 along the axis of the rocket and the exhaust manifold of the TNA turbine, in which the bellows expansion joint 34 was installed.

In connection with the displacement of the TNA towards the cut of the nozzles of the chambers, there was free space in the inlet lines “O” and “G”, sufficient to accommodate the start valves 35, 36 and booster turbopump units 37, 38 “O” and “G”, respectively. Simultaneously with the optimization of the use of the space of the tail section of the rocket unit, the task of unifying the tank bottom of the prototype product with the location of the docking flange 32 along the axis of the rocket was solved. However, bending moment and other loads from the engine to the tank are not transmitted in connection with the installation of the bellows compensator 34. The axial overall dimension of the proposed engine remained unchanged, equal to the axial overall dimension of the prototype engine.

To protect the engine units and the rocket block from high-temperature gases from the nozzles of the chambers, the tail section is closed with thermal protection 30 (see Fig. 12, 13) mounted on the proposed engine. Thermal protection consists of four spherical shields 39, mechanically fixed on the flanges 40 of the nozzles of the chambers and swinging with them, and a flat stationary screen 41 polished to a mirror shine, peripherally fixed by braces 42 to the frame 2, and the central part fixed to the ring 42 mounted on the TNA, while between the spherical shields 39 and the screen 41 there are sealing rings 43 made with cuts for their pliability and fixed on the screen 41.

The proposed engine uses chemical ignition of the fuel components in the gas generator 9 and the chambers 1, therefore, it contains a block of the ampoule of the gas generator 44 and the block of the ampoule of the chambers 45 with starting fuel. The block of the ampoule of the gas generator 44 is mounted on the gas generator 9 (see Fig. 14, 15) in a position close to vertical for the exit of the gas bubble, and with a slope for the convenience of loading its tube with a starting ampoule. The unit of the ampoule of the cameras 45 is fixed to a extension 46 of the fastening of the ring 46 to the engine frame 2 also with an inclination to the axis of the product. Refueling entrances to the tubes are directed to the periphery of the tail section of the rocket unit with access through hatches in the block body.

On the camera block 1 (see Fig. 9), between its movable 47 and fixed 48 parts, flexible piping blocks 22, 23 are mounted, which are fixed with their flanges 49 and 50 to the traverses 13, 14 of frame 2. The flexible piping blocks are perpendicular and symmetrical to the axis the sway of the camera "M" and on both sides of its critical section 51 and ensure the sway of the camera unit at an angle of up to 5 ° max.

The camera unit 1 is locked in the zero position when the engine is transported by the technological lock 52 (see Fig. 16, 17), and in the mounting test complex and at the starting position by the pyrotechnic lock 53. Technological 52 and pyrotechnic 53 locks are installed and fixed in the frames 54 and eye 55 of engine chamber 1. The holes in the camera eye are made when the camera is installed in the zero position in the slipway by the holes in the plates 54 and a special technological device mounted on the studs of the frame plate.

In the transport position of the engine, a spacer 101 is installed between the pyrofixer 53 and the frame plate 54.

During ground firing of the engine, significant vibrational, dynamic, lateral forces act on the nozzles of its chambers, which can lead to destruction of the chambers, therefore, in the proposed engine, on the cylindrical part, fuel collectors 56 and a cut 57 nozzles of chambers 1 (see Fig. 18) are mounted bandages 58, 59 with bosses 60 (see FIG. 20) for welding the thrust of the engine mount to the bench frame 92. This reduces the amplitude of oscillation, deformability of the nozzles and increases the reliability of operation.

The proposed engine has drainages "O", "G", nitrogen from the control lines, starting fuel, etc., which must be removed for thermal protection to prevent fire in the tail section or damage to sensors and other engine elements. Autonomous removal of each drain for thermal protection complicates the design, because drainage cavities must communicate with the atmosphere at a certain time from the start of engine operation. Therefore, the pipelines of the indicated drainages from the cavities of the units are divided into components, grouped together by their outlets and connected to the drainage boxes 61, 62 (see Figs. 22, 23), which are exited for thermal protection 30 by their outlets.

Opening of the drainage box 61 takes place upon a command to the squib 64, from the actuation of which the rod 65 moves with the cover 66, after which the drainage cavities “G” communicate with the atmosphere.

The drainage box “O” 62 is fixed by welding on the nozzle of the nitrogen outlet from the BTNAO turbine.

Before starting the proposed engine in the input line “O”, the oxidizing agent fills the internal cavity of the bellows compensator 34 and part of the start valve “O” 35 to its membrane. Due to the fact that the temperature of liquid oxygen is ~ 94 K, cooling of closely spaced units occurs, which can cause a change in their performance. On the other hand, the influence of positive temperature on the bellows compensator 34 and the start-up valve “O” 35 can cause gas formation in their internal cavities, which can negatively affect the operation of BTNAO and TNA.

Therefore, on the surface of the bellows expansion joint 34 (see Fig. 24) and part of the start-up valve “O” 35, cryogenic thermal insulation 69 up to 25 mm thick is applied, consisting of adhesive, heat-insulating and sealing layers, and heat-insulating covers 68 and 69 are mounted on the bellows. from lining, insulation and shell.

At sub-zero ambient temperatures, the viscosity of the starting and main fuel increases, the rubber in the starting cylinders is embrittled, the operation of the flow regulator is complicated, therefore, the proposed engine uses a heating system for these units (see Fig. 25). The heating system of the engine assemblies consists of heat-insulating covers 70 ... 73, to the fittings 74 of which are connected pipelines 75 for supplying warm air from the rocket. To ensure the necessary costs and temperatures between the covers and units in the piping installed throttle washers 76.

In the proposed engine, the length "L" of the cylindrical part of the chamber 1 is maximally increased within the dimensions of the frame 2, while the gas ducts on the camera head are made in the form of two nozzles 20 (see Fig. 26) to provide the necessary clearance "Δ" with the spherical bottom of the tank 33 missile oxidizer with cryogenic thermal insulation applied to it 89, because The “classic” single gas duct does not fit in its overall dimensions.

The rotation of the chambers of the proposed engine (see Fig. 28) is carried out by a hydraulic system for rotating the combustion chambers (SPKS) with units 102 mounted on the frame 2, in which the pipelines 77 of the camera turning system are not located around the entire perimeter of the frame — their dead end section 84 is excluded, not involved in the work, and in the places of approach to the steering drives, sections 90 of the drives 77 are made of a “lyre-shaped” shape, because compensation loop reduces bending stresses in pipelines, increases their compliance.

The proposed liquid rocket engine operates as follows.

The TNA block (see Fig. 7) is attached to the oxidizing gas lines 17 to the traverses 13, 14 of the engine frame 2 using nuts and studs (see Fig. 9). In addition, at the inputs "O" and "G" it is attached to the bosses 104 on the base of the frame with braces 85, 86 (see figure 4) by welding. Along the longitudinal axis of the engine, the TNA unit from radial displacements is fixed by braces 87 (see Figs. 4, 11), welded by one end to the heat protection ring 42, and the other to the bosses 104 on the frame bases.

The blocks of the chambers 1 with their pins are installed in the bearings 15 of the traverse 13, 14 (see Fig. 8), which are bolted to the bases 88 of the frame (see Fig. 5).

Thus, all engine assemblies are mounted on the frame 2, which is attached to the flange 103 of the rocket tank “O” of the frame 10 by the mating plane “P” of the frame 10 (see FIG. 10) and holds the weight of the entire engine.

According to the launch program, vacuuming and filling of the fuel cavities is performed first. Then, the drainage and discharge of fuel from the ampoule block GG 44 is stopped, nitrogen is supplied from electro-pneumatic valves to open the start valves “O” and “G” 35 and 36, respectively, the oxidizer cavities TNA 3 and GG 9 are poured, after which the chambers 1 are purged with nitrogen. Next, the turbine and rotor BTNAO 37 are unwound with nitrogen, and nitrogen tanks 96 are filled with fuel under pressure. The fuel displaced by nitrogen through the diaphragm from the launch barrels enters the BTNAG 38 rotor and rotates the membranes of the start-up ampoules in blocks 44, 45. The single-component unitary fuel starts up in the GG and KC1 ... 4, where it self-ignites from the connection with oxygen. The gas generator 9 starts, the gas spins the turbine and the rotor TNA 3. The increased pressure opens the fuel valves 97 on the chamber head, and the main fuel flows to the nozzles KC1 ... 4. The engine enters the main thrust mode. Then, the chambers 1 are released from the activation of the pyro cartridge 98 pyrofixers 53, the pyro valve (not shown) is opened to power the steering gears 78, the engine switches to the final thrust stage (KST), purge to the KST and stop by closing the fuel valve (not shown) to the gas generator 9.

The thrust from the chambers 1 is summed up when the engine is running with its weight, transmitted to the traverses 13, 14 of the frame 2. These forces are transmitted through the frame 2 and its mounting frame 10 to the rocket. To prevent excessive deformation of the frame on the working LRE, the loading frame 10 is closed (annular) made of tubular rods connected by plates through bosses. The arrangement of the frame racks in space in the form of triangles makes the frame more rigid, and it was possible to reduce its weight by using thin-walled pipelines ⌀ 30 × 1, ⌀ 24 × 1 made of high-strength steel EI-712.

If necessary, the deviations of the chamber 1 are rotated by their pins 16 in the bearings 15 of the traverse 13, 14, while bending the bellows and tubes of the flexible piping blocks 22 and 23 by a predetermined angle. The stress reduction and reliable operation of the bellows 24 with the oxidizing gas is ensured by the supply of an oxidizing agent for its cooling through the pipeline 94 (see Fig. 9). The bellows are protected from stretching by locking their joints 26 and 28 to the hinged supports 99 and 100. Flexible OSH elements, made in the form of steel tubes, withstand a large number of loading cycles without breaking and are reliable in operation to simplify the design.

When the engine is running, high temperature gas at high pressure is supplied through the oxidation gas supply line 17. As a result, its constituent pipes are heated and subjected to thermal displacements, including TNA, which is subject to vibrational loads. To reduce their amplitude in the radial direction, the TNA is fixed by an eccentric pin 82 mounted between the TNA bracket 83 and the spherical eccentric sleeve 81, which, in turn, is mounted in the eyelet 80 (see Fig. 11). The efforts from the TNA are transmitted from the eye 80 through the bosses 79 and the heat protection ring 42 through the rods 87 to the engine frame 2.

When the engine is running, in addition to thermal displacements of the TNA 3, displacements of the bottom of the oxidizer tank 33 occur under the influence of temperature and internal pressure of the boost. These mutual movements are compensated by the bellows compensator 34 (see Fig. 10), in which bending and axial tension-compression of the bellows occur, therefore, the bending moment from the engine acting on the input line “O” does not affect the rocket tank, and stresses capable of cause damage to the trunk, does not occur.

All thermal protection components 30 sheet parts on the nozzle exit side are polished to a mirror shine to reflect radiant heat fluxes and are made of lightweight OT4-O titanium alloy to reduce thermal protection weight. Between the spherical shields 39 on the nozzles of the chambers and the stationary screen 41, sealing rings 43 (see Figs. 12, 13) are installed, which have cuts (not shown) to increase their compliance and more tightly bend around the spherical shields along the perimeter, which ensures reliable entry prevention gas chambers flowing from nozzles into the tail section of a rocket.

The proposed solution provides reliable engine starting using chemical ignition (see Fig. 14) of fuel components of starting ampoules with unitary starting fuel in a sufficient volume, because high temperatures develop in the gas generator and combustion chambers, contributing to the rapid flow of the combustion process of the fuel components. In addition, in the proposed engine, the ampoule blocks GG and KS are located in a position close to vertical (see Fig. 15), in order to remove, first of all, gas bubbles and inclined to the axis of the engine, and their filling entrances are directed to the periphery of the tail section, what is achieved by the convenience of loading the tubes of blocks of ampoules GG and KS with starting ampoules through the hatches of the outer shell of the rocket block.

When transporting the engine, the chambers 1 are held in the zero position by means of technological latches 52 (see Fig. 16), which maintain their integrity. At the same time, the pyrotechnic lock 53 with the help of the spacer 101 does not enter its stem into the hole of the camera eye 55. When the engine is installed in the PH block, the spacers 101 are removed, the pyrofixers with their rods enter the eyes of the cameras 55 and are attached to the studs on the frame with nuts, after which the technological clamps 52 are removed from the engine (see Fig. 17). Such a device provides for the fixation of the chambers at all stages of operation with the engine from assembly to flight as part of the rocket, when the camera is released by triggering the pyro cartridge 98 pyrofixer 53.

A nozzle operating in an off-design mode in terrestrial conditions at the stand, the engine experiences large vibrational, dynamic and other loads that can lead to its destruction. To prevent this from happening, three bandages 58, 59, 91 are mounted on the cylindrical part and the nozzle of the chamber to secure it to the stand frame 92 (see Fig. 18).

As a result of such fastening of the engine, the destruction of its components on the stand stopped.

The combination of drainage pipelines "O" and "G" in groups and removing them for thermal protection through drainage boxes 61, 62 (see Figs. 22, 23) installed in opposite places on thermal protection increased the fire safety of the engine, as excluded the mixing of the oxidizing agent and fuel among themselves.

The positive effect of the installation at the input “O” to the engine of the input block “O”, consisting of a bellows compensator 34 and a start valve “O” 35, heat-insulating covers 68, 69 (see FIG. 24) and applying cryogenic thermal insulation 67 to its surface allowed to reduce the influence of the negative temperature of the oxidizing agent on nearby aggregates and improved the quality of the oxidizing agent by eliminating its boiling at the start and the ingress of gaseous bubbles onto the BTNAO screw.

The operability of control and regulation units at sub-zero ambient temperatures on the proposed engine is ensured by the use of a warm air heating system (see Fig. 25), supplied from the rocket through pipelines 75 to fittings 74 on heat-insulating covers 70, ..., 73 mounted on the block of launch cylinders 96, regulator, ampoule blocks GG and KS 44, 45.

The flow rate and temperature of warm air on the units are determined by the selection of throttle washers 76.

The completeness of combustion of the fuel components in the combustion chambers of the chambers 1 is achieved by increasing the time they stay in the chamber, which, in turn, is achieved by increasing the length of the cylindrical part of the combustion chambers within the frame dimensions (see Fig. 26).

The camera rotation system (SPKS) (see Fig. 27) is mounted on the engine frame 2 near the mounting frame (see Fig. 3). With its layout on the engine, a number of problems were solved:

1. A sufficiently large number of units are successfully placed above the “G” entrance to the engine, which is positive when installing the engine in the unit, because the area for reconnecting the engine flange with the flange of the rocket pipe is freed. There is no other such favorable place on the engine. At the same time, the units are located in the hatch area of the shell of the rocket body and access to their maintenance is convenient.

2. Refueling and other valves are divided into groups and assigned to the areas of two other hatch shell manholes, which is convenient for system maintenance.

3. Pipelines 77 are separated from the point of fuel extraction from the engine to the steering gears 78 symmetrically, have approximately the same lengths, while the dead end section of the pipelines 84, which is not involved in the operation of the SPKS, is excluded.

4. Pipelines 77 SPKS are rigidly fixed in the tool holders with clamps for the racks of the frame 2, while sections from the steering drives 78 are suitable from them, the longitudinal axis of which changes its position in space when moving the rods connected to the bracket on the camera. In this regard, voltage arises in the supply sections of pipelines to the steering gears. To reduce the loads in places leading pipelines to the steering gears, these sections of pipelines 90 are made “lyre-like”, i.e. with compensation bending, which reduces stress in the pipelines, increases the reliability of the camera rotation system.

The main positive effects of the proposed liquid rocket engine in comparison with the prototype engine is the following:

1. The engine has a significantly higher energy and a higher specific thrust impulse and is able to bring several hundred kilograms more payload to near-earth orbit;

2. The engine layout makes more rational use of the space of the tail section of the rocket block, which, having displaced part of the heat pump for heat protection, placed start-up valves and booster turbopump assemblies in the input lines O and G, which increased the pressure "O" at the entrance to the heat pump to 11.3 kg / cm ² , pressure "G" - up to 8.6 kgf / cm ^2, respectively, which increases the anti-cavitation reserves of TNA pumps.

Placing the bellows expansion joint in a limited space between the oxidizer tank flange and the exhaust manifold from the TNA turbine without displacing the oxidizer tank flange from the engine axis made it possible to unify the bottom of the prototype rocket tank, which significantly reduced the labor costs for upgrading the rocket.

3. controlling the thrust vector of the engine by swinging the chambers themselves, in which the oxidizing gas is completely burned after the TNA turbine, is more efficient, has a higher efficiency than by pumping nozzles through which the not completely burnt reducing gas is emitted from the TNA turbine.

Claims (12)

1. A liquid rocket engine (LRE), comprising four chambers mounted on a frame, a turbopump assembly attached to a frame, having a turbine, oxidizer and fuel pumps, pipelines for supplying oxidizer and fuel to the gas generator and engine chambers, thermal protection, characterized in that the frame made in the form of an all-welded spatial truss and contains mating and lower frames, interconnected by rods, while travers with bearings are attached to the lower frame, into which the camera trunnions are inserted to ensure the possibility of their rotation around the axis of rotation, in addition, the engine contains four curved oxidizing gas supply lines, a single end manifold of which is connected to the turbine outlet, and eight other end nozzles are connected to the respective chamber heads, and blocks are located in the indicated lines perpendicular to the axis of rotation of the chambers flexible pipelines with bellows, connected by one joint to the fixed part of the line, and the other joint secured with its movable part included in the camera unit. 2. The liquid rocket engine according to claim 1, characterized in that a hole is made in the center of its thermal protection, into which part of the turbopump assembly (TNA), protected by a conical casing mounted on the outside of the thermal protection, is removed outside the thermal protection towards the cut of the nozzles of the chambers, a bellows compensator is installed between the tank flange at the “O” input and the exhaust manifold ТНА, and the fuel component start valves and turbo-pump booster units are installed in the “O” and “G” intake manifolds. 3. The liquid rocket engine according to claim 1, characterized in that it has a composite thermal protection consisting of four spherical shields mechanically mounted on the flanges of the nozzles of the chambers and swinging with them, and a stationary screen polished to a mirror shine, on the periphery fixed by stretch marks to the frame, and the central part fixed to the ring mounted on the TNA, while between the spherical shields and the screen there are sealing rings made with cuts. 4. The liquid rocket engine according to claim 1, characterized in that it contains blocks of ampoules of a gas generator (GG) and chambers (KS) with starting fuel, mounted on a gas generator and stretching the heat shield to the frame, respectively, whose tubes are tilted and directed with their filling inputs to the periphery of the tail section of the rocket unit. 5. The liquid rocket engine according to claim 1, characterized in that on the camera block between the stationary and moving parts mounted blocks of flexible piping located perpendicularly and symmetrically to the axis of swing of the camera and fixed to the traverse of the frame on both sides of the critical section of the camera, while the gas bellows It is made cooled, and flexible elements of small internal diameters are made of steel tubes. 6. The liquid-propellant rocket engine according to claim 1, characterized in that in the plate racks of the frame and the eye of the chamber holes are made in which pyrotechnic and technological clamps are installed and fixed. 7. The liquid rocket engine according to claim 1, characterized in that bandages with bosses are mounted on the cylindrical part, the fuel manifolds and the nozzle exit of the chambers for welding the engine tie rods to the bench frame. 8. The liquid rocket engine according to claim 1, characterized in that the fuel and oxidizer drainage pipelines from the cavities of the engine assemblies are separated into components, grouped together by their outlets and connected to drainage boxes mounted on the bottom heat shield and the booster turbine pump unit, respectively. 9. The liquid rocket engine according to claim 1, characterized in that cryogenic thermal insulation is applied to the surface of the bellows compensator and the start valve in the O input line, and heat-insulating covers are mounted on the bellows. 10. The liquid-propellant rocket engine according to claim 1, characterized in that heat-insulating covers are installed on the starting cylinder block, regulator, GG and KS ampoule blocks with starting ampoules, the pipelines with throttle washers of the heating system are connected to their fittings. 11. The liquid rocket engine according to claim 1, characterized in that the length of the cylindrical part of the chamber is maximally increased within the frame dimensions, while the gas ducts on the chamber head are made in the form of two nozzles. 12. The liquid rocket engine according to claim 1, characterized in that on its frame for controlling the thrust vector, a camera rotation system is installed, in which the pipelines of the camera rotation system are located at the points of supply to the steering drives, and they are made of a “lyre-like” configuration with compensation bending .

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