RU2749235C2 - Rocket engine - Google Patents

Abstract

FIELD: rocket science.SUBSTANCE: invention relates to rocket science and can be applied to rockets and other aircraft for space and air applications. A rocket engine (RE) is described, containing a nozzle and a combustion chamber connected by means of a pump to a tank in which the fuel is located, according to the invention, the fuel is in a solid dispersed bulk state, the tank is a hopper, and the pump is an airlock feeding mechanism containing a housing with inlet and outlet pipes and a movable element mounted in the housing connected to the motion drive with a cavity capable of being combined with these pipes alternately. The shell of the hopper is made up of annular sections interconnected by snap ties, and the bottom of the hopper is made in the form of a piston, on which the combustion chamber is fixed with a nozzle and the specified airlock feeding mechanism.EFFECT: invention provides a simplification of the design, increasing the reliability and safety of the engine in case of damage to its structure.2 cl, 2 dwg

Description

The invention relates to rocketry and relates to the design of chemical rocket engines, i.e. using chemical fuel, used in air and space vehicles for transport and military purposes.

The design of a rocket engine is determined by the state of aggregation of the fuel used (liquid, solid, or a combination of both types of components) (see for example: Wikipedia: "Rocket engine"). Let's consider the problems here from a general point of view.

The function of the rocket engine is to create a flow of impulse (i.e. thrust) due to the consumption of the mass stored on board. In this case, the engine operation time is limited by the mass reserve, and the total final impulse that the aircraft will acquire in this case will be the greater, the more the mass is thrown away at a higher speed. Velocity in a chemical engine is created in the process of adiabatic expansion in a nozzle of gaseous products of exothermic chemical transformation (combustion) of fuel that occurs in a combustion chamber (hereinafter CC), into which the fuel, or its individual components, are pumped in condensed (i.e., dense) condition. Moreover, this injection is carried out against the pressure gradient arising automatically due to the manifestation of the law of balance of the inflow and outflow of mass.

If the fuel is liquid, it is usually pumped into the combustion chamber by a pump. In this case, it is usually necessary to pump two components of the liquid fuel, which react with each other, mixing in the combustion chamber. Premixing of liquid fuel components is problematic due to the high chemical activity of the oxidation reaction of a homogeneous combustible mixture. It is also possible to use a one-component liquid fuel consisting of molecules with internal free potential energy, for example, hydrogen peroxide. However, this energy is much less than that of a pair of separate fuel and oxidizer molecules.

The principle of operation of a liquid-propellant rocket engine (LPRE) described above causes the following disadvantages.

1. The need to have two pumps for supplying fuel components to the compressor station, which complicates the design of the rocket engine.

2. The need to have two separate tanks for the fuel components.

3. The need to ensure the tightness of the tanks, which is problematic from the point of view of the existing probability of their damage by meteorites or enemy countermeasures.

4. Problems associated with the prevention of evaporation of liquid components and their reaction with the structure, which complicates the operation of engines.

5. Fire hazard of a fueled rocket engine.

The pumps can be eliminated by pressurizing the tanks. But this increases the mass of the tank structure many times over and, consequently, reduces the final speed of the aircraft.

It is also possible to replace the pumps with a vapor injection system for liquid propellants using heat from the liquid cooling jacket of the nozzle. This system is viable, but it is no simpler than a turbo pump system.

Despite the above disadvantages, liquid-propellant rocket engines are widely used in rocketry, tk. allow the use of the most energy efficient fuel components, as well as provide the ability to control thrust and turn the engine on and off in flight.

Solid chemical fuels can also be composed of several separate molecular high-performance components - a fuel and an oxidizer. However, unlike liquid fuels, they may not be in a homogeneously mixed state, but represent a heterogeneous finely dispersed mixture. In this case, the chemical activity is restrained by a low diffusion coefficient in the solid phase and a relatively large diffusion distance between the fuel and oxidizer particles. Such fuel can be in a stable state indefinitely if the critical mass of the storage volume is not exceeded and heat is removed from the surface of the stack. However, solid fuel is introduced into the combustion chamber prior to starting the engine. In this case, the engine operating time is determined by the volume of fuel introduced into the combustion chamber.

The advantages of a solid-propellant rocket engine (solid propellant rocket engine) over liquid-propellant rocket engines are:

1. No need for pumps.

2. Practical single-component fuel (from a macroscopic point of view).

3. Better protection against damaging factors, tk. the fuel is located in a sufficiently strong compressor station.

4. Possibility of long-term storage of the "filled" solid rocket motor.

5. Less fire hazard.

However, this is provided at the cost of the appearance of the following disadvantages in solid propellant rocket engines compared to liquid rocket engines:

1. Limitation of the operating time of one stage by the volume of the space in the KS. fuel.

2. An increase in the mass of the CC structure, due to an increase in its volume due to the need to accommodate a large mass of solid fuel in it.

3. Impossibility of traction control.

4. Limited service life of the nozzle and COP.

5. The problem of premature engine shutdown in flight and its re-inclusion.

6. Reduced chemical efficiency of solid fuel in comparison with liquid fuel, which is associated with the need to introduce additional molecular bonds in order to ensure its stability with a close mutual arrangement of reactive components.

7. The need to keep the rocket engine constantly fueled. This makes it necessary to transport heavy missiles over long distances, while separate transportation of fuel and empty missiles with refueling at the place of use would be practically more convenient.

In connection with the above, the problem arises of obtaining a rocket engine of a fairly simple design, safe and resistant to damage to the tanks, with unlimited (in principle) operating time, with adjustable thrust and the ability to turn off and turn on the engine in flight. At the same time, the problem of ensuring a sufficient service life of the nozzle and the combustion chamber can still be solved through the use of heat-shielding endothermic and ablative composite materials, as has been proven by the practice of using solid propellants.

However, for the formulation of the essence of the invention, we will choose a certain prototype. The most suitable prototype in terms of the number of coinciding essential features and efficiency will be a liquid-propellant rocket engine with two-component fuel (see, for example, Alemasov V.E. et al. "Theory of Rocket Engines", third edition. Revised and supplemented under the editorship of Academician V.P. M, Mechanical Engineering, 1980. Page 196, Fig. 16.6 and Fig. 16.7.).

Then the purpose of the invention will be - to simplify the design, reduce operating costs, increase the reliability and safety of the engine in case of damage to its structure.

The proposed rocket engine contains a nozzle and a combustion chamber connected by means of a pump to a container in which the fuel or one of its components is located. The purpose of the invention is achieved in that the fuel, or each of its components, is in a solid dispersed free-flowing state, the container is a bunker, and the pump is a sluice feed mechanism, i.e. a mechanism comprising a housing with inlet and outlet nozzles, and an element movably installed inside the case with a cavity, which, when the movable element moves, can alternately connect either with the inlet or with the outlet nozzles. The sluice mechanism provides the ability to supply solid dispersed bulk material to the high pressure zone in the combustion chamber and operates due to the specified pressure drop, which is a property of any sluice mechanism, unlike pumps. The sluice mechanism needs a drive only to compensate for frictional forces. Known pumps cannot provide the supply of bulk material against a large pressure drop due to the presence of dry friction forces in bulk solids (as opposed to liquid solids). Replacing the liquid with bulk fuel allows you to make the fuel bunker leaky, with the size of the slots determined by the size of the granules of solid fuel. This simplifies operation, increases fire safety and resistance to structural damage. The design of the engine is also simplified, since eliminates the need for two fuel storage facilities and two fuel delivery systems for the engine.

It also allows the bunker to be ejected in chunks of structural mass as fuel is consumed. For this, in a particular embodiment, the shell of the bunker is made up of annular sections interconnected by interlocking links, and the bottom of the bunker is made in the form of a piston, on which a combustion chamber with a nozzle and the specified sluice mechanism is fixed. Due to the thrust of the engine, the piston presses the fuel located in the bunker upwards. In this case, the annular sections of the shell, as the fuel is consumed, are alternately released and can be detached due to the lock connections.

The annular section of the hopper shell, in particular, can be made with a cut along the generatrix and is spring loaded for compression of the ring. In addition, adjacent annular sections have grooved engagement that functions as lock connections between themselves, which are held in an engaging state due to the expansion of the annular sections with the fuel in the bunker. When the next section descends below the bottom of the hopper, the section ring is compressed and protrudes from the grooves of the overlying section with its protrusions, after which it freely moves away from the aircraft, i.e. discarded, thus decreasing. useless load on the aircraft. Thus, the principle of discarding the mass of the structure is implemented in proportion to the mass of fuel consumed. Thus, in this embodiment, the rocket engine is equivalent to a multistage composite engine, which allows one to achieve high, including possibly space, speeds with one engine. Those. the design of the rocket propulsion system as a whole is simplified, since it is made one-stage.

The invention is illustrated by the following description of structural examples and two figures.

FIG. 1. schematically depicts the proposed rocket engine with the simplest single-rotor sluice power mechanism.

FIG. 2 shows the configuration of the annular section of the fuel hopper shell after it has been separated from the engine.

The proposed rocket engine (Fig. 1) contains a compact combustion chamber 1 (hereinafter referred to as CC) and a nozzle 2, made in the form of a single block of low heat-conducting fire-resistant composite materials, cooled by endothermic reaction of chemical decomposition and ablation of the inner layer. In the bottom of the CC there is an opening through which it communicates with the feed sluice mechanism 3, which acts as a pump for supplying fuel to the CC against the pressure gradient from the hopper 4, which is made leaky. The feed mechanism 3 is fixed on the bottom 5 of the bunker 4, made in the form of a piston capable of moving along the shell 6 of the bunker.

The shell 6 of the hopper is made up of annular sections 7, interconnected by means of an annular protrusion 8 and a cavity 9 complementary to it, belonging to the overlying section. The protrusion 8 and the depression 9 perform the function of the locking connections of the annular sections with each other. The ring of each section has a cut along the generatrix and a spring 10 acting on the compression of the ring up to the stop formed by the corresponding folds of the ring (see Fig. 2). At the top, the hopper 4 is equipped with a lid 11.

The sluice power mechanism 3, in its simplest embodiment, contains one gear wheel 12, the teeth of which fit snugly against the body 13. In the upper part of the body 13 there is an inlet pipe 14 communicating with the hopper 4, and in the lower part - an outlet pipe 15 communicating with COP. The gear wheel is driven in rotation by an electric motor 16 with a controlled number of revolutions, powered by an on-board power source 17, for example from an electric accumulator.

The gear wheel of the feed mechanism must have sharply sharpened edges that form cutting edges. This eliminates the possibility of jamming of the mechanism, because propellant pellets that impede movement will cut easily like scissors.

Solid fuel can be a conventional blended fuel. In this case, flowability is ensured due to its granularity. The size of the granules is chosen in a compromise: according to the condition of not hanging in the chutes, where the diameter of the granules should be approximately 10 times less than the minimum diameter of the streams, and also according to the condition of sufficiently fast burnout of the granules, where the diameter of the granules should be small. The approximate diameter of the granules is estimated at 1 mm. The microstructure of the heterogeneity of the mixed fuel granules, accordingly, will be an order of magnitude smaller than the diameter of the granules. In this case, in contrast to conventional mixed solid propellant charges, there is no need to achieve high bond strength of the dispersed composition inside the granule. On the contrary, the granule should, if possible, quickly disintegrate when it enters the combustion chamber in order to have time to burn during its flight to the nozzle. This eliminates the need to strictly maintain the fuel parameters in terms of the pellet burning rate. This speed should only be high enough - more than the burning rate of composite fuel in solid propellant rocket motors, but less than the burning rate of ballistic powder. The absence of high requirements for the exact observance of the pellet burning rate here is due to the fact that it is controlled by the rate of its supply to the combustion chamber.

The proposed rocket engine works as follows. Turn on the power motor 16 and at the same time an ignition initiator, for example a powder (not shown). By changing the speed of rotation of the electric motor 16, you can change the thrust. The bottom 5 of the fuel bunker 4 made in the form of a piston, due to the thrust created by the nozzle 2, presses the bulk solid fuel in the bunker, preventing the formation of voids during fuel consumption. In this case, the fuel expands the annular sections, overcoming the force of the spring 10 and maintaining the mutual engagement of the sections. In this case, the bottom 5 slides upward along the shell 6, like a piston, guided and oriented by the shell 6, like a cylinder. When the bottom 5 rises above the next joint of the annular sections, the ring of the lower section is compressed under the action of the spring 10, and its protrusions 8 come out of the depressions 9 of the overlying section (see Fig. 1). Thus, the lock connections with the overlying section are disabled. As a result, the accelerating engine goes up, leaving the separated section of the hopper in free fall. That is, there is a timely automatic separation of the spent part of the structure of the fuel bunker. At the same time, due to the periodic discarding of the spent part of the bunker structure, a sufficiently high ratio of the initial mass of the rocket to the final mass (Tsiolkovsky number), which is necessary to achieve hypersonic speeds and also (not excluded) the first space speed, can be ensured.

After reaching the required speed (or altitude), this engine can be turned off for the ballistic phase of flight, stopping the fuel supply, and then turned on (for example, with the help of an additional combustion initiator (not shown) and performing the orbit correction phase, maneuver when aiming at the target or dialing the finish speed.

Typically, these functions require two or three stages with different motors and control systems, which complicates the design.

The proposed design is not afraid of depressurization. Also, it is not critical to changes in the magnitude or uniformity of the burning rate of fuel pellets during storage, because the combustion rate is controlled by the rate of its supply to the combustion chamber and can be corrected by changing the feed rate. The engine can be refueled immediately before starting. In view of the lower requirements for the density of the connections of the parts of the bunker and the power supply system, as well as in view of the non-volatility of solid fuel, this engine is less fire hazardous than the liquid-propellant engine, both during storage or refueling, and during the active phase of flight. For example, a bullet or splinter of an interceptor projectile will lead to a fuel leak from the liquid-propellant engine tank, and to a fire with catastrophic consequences. Damage to the bulk fuel bunker will only lead to fuel leakage, which, in a short time of the active flight segment, will not lead to a large loss of the final impulse. Also, the advantage of the proposed engine is that it can be disassembled by disassembling the hopper into separate rings for the purpose of long-term storage or transportation separately from solid fuel, which can be transported in vehicles designed for ordinary bulk cargo. In addition, the volume of the fuel tank is easily varied for the required task. To do this, it is enough to change the number of annular sections. A quick replacement of the actual motor part is also provided. So the principle of modularity and flexibility is provided for the formation of an aircraft for the required speed, range, payload and grade of fuel.

Claims (2)

1. A rocket engine (RD) containing a nozzle and a combustion chamber connected by means of a pump to a container in which fuel is located, characterized in that the fuel is in a solid dispersed free-flowing state, the container is a bunker, and the pump is a sluice power mechanism, containing a housing with inlet and outlet pipes and a movable element with a cavity installed in the housing and connected to the motion drive, capable of being aligned with the said pipes alternately. 2. RD according to claim 1, characterized in that the bunker shell is made up of annular sections interconnected by interlocking links, and the hopper bottom is made in the form of a piston on which a combustion chamber with a nozzle and said power gate mechanism is fixed.

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