CN112128020A

CN112128020A - Double-component rocket engine without atomizing device or spraying device

Info

Publication number: CN112128020A
Application number: CN201910553230.3A
Authority: CN
Inventors: 吴凡
Original assignee: Individual
Current assignee: Individual
Priority date: 2019-06-25
Filing date: 2019-06-25
Publication date: 2020-12-25

Abstract

The patent aims to provide a design scheme of a liquid rocket engine without an atomizing device and a spraying device, introduces a flow of fuel entering a combustion process after passing through a gasification device, adds other cautions, also introduces a specific structure of the gasification device, introduces a concept of a mixing chamber, gives two possibilities of thermodynamic cycle, finally provides some functions of starting and stopping, ensures safety performance, introduces how to change the thrust magnitude and direction for the rocket engine, provides new requirements for a liquid pump and an air pump, and lists some advantages and disadvantages of the rocket engine. The technical field is a liquid double-component rocket engine; the running mode of the liquid double-component rocket engine.

Description

Double-component rocket engine without atomizing device or spraying device

Background:

most of the current mainstream rocket engines comprise an atomizer or a liquid drop generating device in a broad sense, and complicated consideration needs to be given to the generation of tiny liquid drops so as to ensure a combustion ratio of 98 percent to 99 percent, according to some reports: the combustion thrust system directly based on gasification mixing without atomization equipment is difficult to develop greatly, and the engine introduced by the patent is an air-gas hybrid engine optimized as much as possible based on basic physical theory and natural scientific common knowledge

Such as patent US 4879874; US 4901525; in the liquid engine described in US5267437, the fuel is either pressurised or partially vaporized

In the liquid engine described in patent CN1201075C, the components that are separate from the combustion chamber and the thrust chamber in many other engines are integrated and reorganized, and although the mass is reduced, they impose higher demands on the manufacturing process

US5918460 discloses a liquid rocket engine using gasified liquid oxygen to supply combustion chamber

This patent, however, not only gasifies the fuel but also the oxidant and ensures that the vast majority of the combustion in the combustion chamber is a gaseous mixture, rather than a mixture of droplets or a single component oxidant with a surrounding gaseous fuel or a single component fuel or a surrounding gaseous oxidant

The technical field is as follows:

two-component liquid rocket engine comprising gasification means, thermodynamic cycle process comprising cooling of the engine

The description part is as follows:

the double-component rocket engine does not need an atomizing device or a spraying device.

The pumps and turbines mentioned in this patent refer to the same concept in most cases.

The liquid oxygen methane supply system is mainly designed for the liquid oxygen methane supply system.

In order to fit actual needs, the user of the patent should make the engine and set specific parameters of each link according to actual conditions.

Based on the consideration of simple and convenient process, the engine does not adopt the design scheme of the internal casting channel of the combustion chamber and the thrust chamber.

Liquid methane oxide is recommended as fuel and oxidizer, and an inert gas tank for purging is also provided for multiple start-up considerations.

Because different fuel and oxidant combinations may be considered for actual manufacturing, the steady state operating temperature of the gasification chamber may also vary.

The engine operates in the manner generally shown in fig. 1, fuel (2, 13) and oxidant (1, 14) are pumped by pumps (4, 15) into a gasification chamber (5, 17), the gas gasified by the gasification chamber is accelerated by an air pump (6, 18) into a mixing chamber (9, 21, 30), the mixture in the mixing chamber is combusted (11, 23, 32) in a combustion chamber, the resulting high-temperature and high-pressure gas is discharged by a thrust chamber (12, 24, 36), on the other hand, a condenser (7, 19) collects liquid droplets condensed on the tube wall, the liquid droplets pass through a coolant (8, 22), and the liquid droplets are pumped by the pump to the combustion chamber and the thrust chamber to cool the combustion chamber and the thrust chamber, and the condensed liquid droplets are re-pumped into the fuel pump, thus achieving a thermodynamic cycle.

In addition, another major difference between this engine and the currently used engines is that it also has a gas mixing chamber above the combustion chamber, as shown in fig. 4 and 5, where fuel enters from both ends of the mixing chamber and oxidant is located at the top of the mixing chamber, thereby achieving sufficient mixing of the fuel and oxidant, and the fuel tube and mixing chamber are not specified for the angle of injection inclination when assembled.

The general layout of the engine combustion chamber (32) and the thrust chamber (36) is shown in figure 4, the arrangement corresponds to that the

flow charts

1, 31 and 34 are connected with the condensed liquid pipes (37 and 38) and the pumps (4 and 15), if the flow chart shown in figure 2 is adopted, one end is connected with the inlet of the gasification chamber, the other end is connected with the air pump, in addition, a hollow fastening ring (35 and 41) of the throat part of the combustion chamber is also arranged, and the cooling liquid flows to cool the pipe wall.

In order to ensure sufficient fuel and oxidant flow, a gasification chamber as shown in fig. 6 and 7 is selected, and the design is based on the basic physical principle that: in order to ensure that the pressure at the inlet cross-section of each gasification tube is approximately equal, the gas outlet tube need not be designed to be larger or smaller than the cross-section of the gasification tube.

The micro-tubes in the gasification chamber are of different lengths (51), the gasification chamber is provided with a concave surface (50), the concave surface can be a surface (50) that a sphere is cut by a cone passing through a vertex in geometry, and can also be a space quadric surface, the inlet of the gasification chamber is narrower than the inner diameter (45) of the pipeline where the gasification tube is positioned, the design is to increase the total cross-sectional area of the gasification tube as much as possible, the fillers (48, 52) between the gasification tubes are made of materials with good heat conduction performance as much as possible, and can be solid or liquid, and the gasification chamber shown in the figure is heated by an electric heating plate (53).

The cross-section of the chamber is shown in figure 6, and it is theoretically ensured that the cross-sectional area of the hollow is maximized while the cross-sectional area of the cylindrical individual evaporator tubes (46, 44) is as small as possible, so figure 6 is only a schematic illustration, and it should be noted that the evaporator tubes (not including the electrical heating strips) have a rotationally symmetrical structure, which is necessary for uniform distribution of the outlet gas.

Another possible operation of the engine is shown in fig. 2, with only minor modifications on the previous basis, where the engine is installed in the vicinity (16) of the rocket body, fuel (13) and oxidant (14) are pumped by a liquid pump (15) to the gasification chamber, and on the other hand, part of the fuel and oxidant flows directly through the combustion chamber wall and the thrust chamber wall, the heated liquid is returned to the gasification chamber (17) after passing through a cooling chamber (22), and further, the condensate, which has been at a higher temperature after passing through the gasification chamber, is returned directly to the gasification chamber without passing through the combustion chamber wall and the thrust chamber wall (this process may require pressurization by a pump, not shown in the figure), fuel enters the mixing chamber from both sides (20), and oxygen enters the mixing chamber from the mixing chamber axis.

The ignition device (a) is located at the lowest end of the combustion chamber, the ignition device is a loop which only needs to generate electric arc at the moment of ignition, two electrodes of the ignition device are required to be high temperature resistant and small enough, if the power supply of the ignition device is arranged on the rocket body, the engine can be started for a plurality of times theoretically, and if the power supply of the ignition device is arranged on the launching cradle for reducing the weight of the rocket body, only a single ignition can be provided theoretically.

Since the engine directly ignites the gaseous fuel-oxidant mixture, it is in most cases more difficult to control the thrust direction and stabilize the engine in turbulent conditions than an engine using an atomizing device, and therefore a mechanical arm as shown in fig. 5(39, 40, 42) is used.

The use of cooling liquid is only considered if and only if the radiation heat dissipation efficiency is low and the condensation manifold (7, 19) is not sufficiently filled, so that the user of the invention may consider cooling the walls of the tubes, and the cooling liquid itself may use an oxidant or a fuel, in order to collect the condensate.

The arrangement of the condenser tubes (7, 19) is shown in fig. 3.

There is filter screen (29) condensation collecting pipe's top, the setting of filter screen is for better collection condensate, also in order to prevent that great liquid drop from getting into the air pump (if have), need not to heat the filter screen under general conditions, the setting of filter screen is shown in the figure, the filter screen is flowed through to the gas-liquid mixture, the induced liquid drop of shape of filter screen flows to the pipe wall, the pipe wall has recess (27), arrow body's acceleration (29) lets liquid can be collected by condenser pipe (28), can think that the liquid in the condenser pipe mixes gas, consequently, it has the ability of holding a certain amount of gas to need the liquid pump.

Description of the drawings:

FIG. 1: for one possible flow path for the engine, arrow 3 refers to the engine-to-arrow body interface.

FIG. 2: for another flow path that may be employed by the engine, arrow 16 refers to the interface between the engine and the arrow body.

FIG. 3: the screen is cut at 1/4 to show the condenser inlet recess, the recess obscured by the screen is not shown, arrow 3 indicates the engine-to-arrow junction, and 25 is the flow direction of the gas-liquid mixture relative to the tube.

FIG. 4: is an overall layout of the engine, wherein the wall thickness of the pipes is not represented in the figure, and furthermore 37, 38 are condensate (outlet for condensate).

FIG. 5: for the thrust direction changing device layout, wall shading is omitted because the same structure as that of fig. 4 is expressed, 39 is a hydraulic transmission device, and 43 is a bolt joint.

FIG. 6: the cross section of the gasification tube is schematically shown, and the practice of the actual situation is more dense than that in the figure.

FIG. 7: the diameter of the vaporizing tube (gas-guide tube) is shown in FIG. 51.

The specific implementation mode is as follows:

in practice it should be ensured that the combustion generating surface is even below the interface of the mixing chamber and the combustion chamber.

The security is guaranteed: in order to ensure that the oxidant does not flow back in the fuel line or that the fuel does not flow back in the oxidant line, a pressure detection device, or a gas concentration detection device, is provided in the fuel line or the oxidant line (adjacent to the junction of the line and the mixing chamber), and furthermore a throttle valve should be provided in the fuel line in response to a corresponding safety alarm.

Air pump and liquid pump: there is not special regulation to the selection of air pump and liquid pump, only need to guarantee that can hold a small amount of gas in the liquid pump and be unlikely to go out the problem, and the entry of liquid pump will guarantee two, and one can carry fuel and oxidant by extensive high pressure, and another entry can let the condensate get into and pump with the export of carrying fuel and oxidant (flow chart 1), and the same way, the requirement to the air pump is: it is required that it can operate with a small amount of liquid droplets mixed, the air pump needs to be able to operate at a low temperature, for example below-100 c, and furthermore the air pump can be simply replaced by a fan rotating at a high speed.

The power regulation mechanism comprises: for engines already in steady state, this is achieved by varying the power of the liquid and air pumps of the fuel and oxidant tubes, noting that this cannot be achieved by adjusting the temperature of the gasification tubes

Thrust direction changing mechanism: the stabilization is achieved by a hydraulically driven lever, as shown in fig. 5, and considering that the structure is different and the total length of the mixing chamber and the combustion chamber is often longer, a scheme that one end of the lever is connected with a fastening ring and the other end of the lever is connected with the interface of an engine and an arrow body is adopted, and the lever is also provided with a section of slide rail, a slide rod is arranged in the slide rail, and the other end of the slide rod is connected with the combustion chamber, so that the whole structure strength can be enhanced while the thrust direction is changed.

The thermal cycling mechanism is as follows: the engine mainly uses radiation heat transfer and assists forced convection heat transfer and heat transfer, and heat generated by the outer walls of the combustion chamber and the thrust chamber is partially used for heating condensate.

The main advantages are that: the design does not impose higher requirements on the existing manufacturing process, and the design almost maximally ensures the sufficiency of combustion; no afterburning equipment is needed; it is supplied in a simple optimum mixing ratio (1: 2 (molecular ratio) for methane and oxygen); the gasification chamber is simpler and simpler to manufacture than a common structure.

The main disadvantages are that: the mass theory of the engine is estimated to be larger than that of a common engine with the same power under the existing manufacturing process, the power of a required air pump is correspondingly increased for the engine with higher thrust requirement (the engine at least has two air pumps more than the engine which is generally popular), and for the task requiring high thrust, the mass of an rocket body is increased if a gasification chamber is made to be large, so the initial estimation is probably heavier than that of the engine with the same power.

Turning on and turning off the machine: the engine is characterized in that the gasification chamber is preheated when the engine is started, the temperature of the gasification chamber when the engine is started is higher than that of the gasification chamber when the engine is in a steady state, and the pipe orifice pressure of fuel input and oxidant input when the engine is started is ensured to be equal for any fuel.

After preheating the gasification chamber, the liquid pump is started, and then the electric igniter is ensured to be in a running state until the engine generates fuel gas.

During shutdown, it is ensured that the fuel supply and the oxidant supply are closed simultaneously and that the pressure drops to 0 at the same rate and simultaneously with a drop in the temperature of the gasification chamber, care being taken to shut off the coolant circulation.

The method is also suitable for changing the thrust.

Claims

1. A kind of double-component liquid rocket engine, include fuel gasification chamber and oxidizing agent gasification chamber (two gasification chambers) at the same time, do not include the liquid droplet generating device, the structure of the gasification chamber is that the hollow pipeline is densely covered with the gasification tube, the gasification tube is different in size and enters the mouth and distribute on a continuous curved surface, there are media with good thermal conductivity in gasification tube and gasification tube, there are electric heating apparatuses on both sides of gasification chamber; the inlet of the gasification chamber is narrower than the outlet.

2. A rocket engine as claimed in claim 1, wherein the upper end of the combustion chamber of the engine (in the positive direction of rocket acceleration) is provided with a mixing chamber, the upper end of the mixing chamber is provided with 3 ducts, the mixing chamber has an axisymmetric structure, the top end of the mixing chamber is provided with an oxygen pump, the two sides of the mixing chamber are provided with fuel pumps, the mixing chamber is approximately cylindrical, and the diameter of the cross section of the mixing chamber is narrower than that of the combustion chamber.

3. A rocket engine as claimed in claims 1 and 2, wherein the gasified fuel and oxidant pass through a condensing unit, a condensing net is arranged in the condensing unit, the condensing net is of a continuous curved surface structure, the top of the condensing net is more protruded than the periphery of the condensing net, a condensing pipe is arranged in a column section where the condensing net is located, the condensing pipe is attached to the periphery of a pipeline, the interfaces of the condensing pipe and the air supply pipe form a certain inclination angle with the air supply pipe, the condensing pipe is an acute angle, and the inlet of the condensing pipe is located below the condensing net (the rocket acceleration direction is taken as the positive direction).

4. A rocket engine as claimed in claims 1, 2, 3, provided with or equivalent to the following work flows: the fuel and the oxidant are pumped into the gasification chamber by a pump, the gas gasified by the gasification chamber is accelerated by an air pump to be sent to a mixing chamber, the mixture in the mixing chamber is combusted in a combustion chamber, the generated high-temperature and high-pressure gas is discharged from a thrust chamber, on the other hand, a condensation pipe collects liquid drops condensed on the pipe wall, the liquid drops pass through cooling liquid (the fuel or the oxidant per se) and are accelerated by a liquid pump to be sent to the combustion chamber and the thrust chamber, and the condensate passes through the pump (the fuel pump and the oxidant pump) again.

5. An engine as claimed in claims 1, 2, 3, 4 and having the following features: an annular pipeline is arranged around the combustion chamber and the thrust chamber, condensate flows through the pipeline, and: the fastening ring is arranged at the junction of the combustion chamber and the thrust chamber, the fastening ring is hollow, and condensate flows through the hollow ring.

6. A rocket engine as claimed in claims 1, 2, 3, 4, 5 and having the following features: the fastening screw ring is connected with a mechanical transmission device and comprises: the combustion chamber is connected with the transmission device through a rigid rod.

7. A rocket motor as claimed in claims 1 and 2, wherein the fuel and the oxidizer are gasified and then fed to the mixing chamber by means of a high-speed fan or an air pump.