Rocket engine for realizing attitude control by using gas swing

Abstract

The invention discloses a rocket engine capable of achieving attitude control through fuel gas swing. The rocket engine comprises an injector, a combustion chamber and an exhaust nozzle. An X exhaust pipe is arranged outside the combustion chamber, and the first end of the X exhaust pipe penetrates through the wall face of the combustion chamber and communicates with the interior of the combustion chamber; a Y exhaust pipe is arranged outside the combustion chamber, and the first end of the Y exhaust pipe penetrates through the wall face of the combustion chamber and communicates with the interior of the combustion chamber. The four output pipes are controlled to be switched on and switched off through the two flow dividing valves correspondingly, pressure difference vector superposition is generated when the three output pipes are switched on at the same time, high-speed mainstream fuel gas accurate deflection is completed, and rapid and accurate attitude correction or yawing of a rocket is achieved; the outlet ends of the four output pipes are all flat, when the rocket needs to yaw or posture correction, only one output pipe needs to be conducted, ejected gas forms oblique shock waves with the wide coverage characteristic, and therefore thrust vector control over the rocket is achieved, loss of mainstream gas energy is avoided, and precise posture control and propelling performance of the rocket are both considered.

Description

TECHNICAL FIELD

[0001] The present application relates to a rocket engine using gas swing to realize attitude control. BACKGROUND

[0002] With the development of technology, some attitude-controllable rocket engines appear in the prior art: the first stage of the Long March 7 carrier rocket adopts double-machine parallel helium pressurized bidirectional swing engines, and the second stage adopts four-machine parallel engines, of which two are fixed and two are double-swing, which realizes the attitude control of the rocket through the swing of the engine; Qingqiu 12 is a pump-after-swing gas generator circulating liquid oxygen and methane rocket engine independently developed by Blue Arrow Aerospace, each single-engine engine is unidirectional swing, the swing direction is the tangential direction of the rocket body, and the swing angle is ±8°, which can provide attitude control force in all directions for the rocket flight.

[0003] For example, Deep Blue Aerospace Technology proposes patent number 202010329068.X, which discloses a rocket multi-engine parallel structure, including a rack and a plurality of engines, and further including a swing mechanism, the swing mechanism includes a rotating shaft, a swing arm and a servo motor, the rotating shaft is connected with the servo motor, one end of the swing arm is arranged on the rotating shaft, and the other end of the swing arm is connected with the engine. During the rocket recovery process, the swing of the engine is controlled to adjust the heading and attitude of the rocket body, and stable recovery is ensured.

[0004] For example, Beijing University of Aeronautics and Astronautics proposes patent number 201710853057.X, which discloses an attitude control device using rocket engine gas, including a propellant tank, a heat exchanger, a main engine, a buffer gas tank and an attitude control device using rocket engine gas. The attitude control device using rocket engine gas includes at least two groups of engine groups; two groups of the engine groups are oppositely arranged.

[0005] The swing mechanism of the above-mentioned prior art scheme is complex, the purchase of multiple swing engines and their supporting servo motors increases the research and manufacturing cost of the rocket, and the reaction speed of the swing engine is slow, which cannot meet the demand of attitude adjustment and correction during the recovery of the rocket. The complexity of the control system and the swing system increases the failure rate of the rocket, which restricts the competitiveness of the carrier rocket. The technical scheme proposed by Beijing University of Aeronautics and Astronautics uses the horizontal driving force of two groups of auxiliary engines to flexibly control the attitude of the rocket, realizes the functions of pitching, rolling and yawing, however, the two groups of auxiliary engines cannot generate thrust for the rocket, and are only used for attitude control, so that the two groups of auxiliary engines are carried additionally, which not only increases the additional load of the rocket, but also increases the manufacturing cost of the rocket. SUMMARY

[0006] The present application is aimed at providing a technical solution to solve the above problems.

[0007] A rocket engine for attitude control by gas swing, comprising: an injector, a combustion chamber, and a tail nozzle; An X gas extraction pipe is arranged outside the combustion chamber, and a first end of the X gas extraction pipe penetrates a wall surface of the combustion chamber and communicates with the inside of the combustion chamber; The other end of the X gas extraction pipe is connected with an X flow divider valve, two output ends of the X flow divider valve are respectively connected with an Xa output pipe and an Xb output pipe, and outlet ends of the Xa output pipe and the Xb output pipe oppositely extend into the inside of the tail nozzle; Preferably, the X flow divider valve can independently adjust the on-off state of the Xa output pipe and the Xb output pipe, and independently control the gas flow into the Xa output pipe and the Xb output pipe; Preferably, a Y gas extraction pipe is arranged outside the combustion chamber, a first end of the Y gas extraction pipe penetrates a wall surface of the combustion chamber and communicates with the inside of the combustion chamber, the other end of the Y gas extraction pipe is connected with a Y flow divider valve, two output ends of the Y flow divider valve are respectively connected with a Ya output pipe and a Yb output pipe, and outlet ends of the Ya output pipe and the Yb output pipe oppositely extend into the inside of the tail nozzle; Preferably, the Y flow divider valve can independently adjust the on-off state of the Ya output pipe and the Yb output pipe, and independently control the gas flow into the Ya output pipe and the Yb output pipe; Preferably, the valve core of the X flow divider valve and the valve core of the Y flow divider valve are both hollow structures, and the valve core of the X flow divider valve and the valve core of the Y flow divider valve both have two through holes; Preferably, the outlet end of the Xa output pipe, the outlet end of the Xb output pipe, the outlet end of the Ya output pipe and the outlet end of the Yb output pipe all tilt towards the outlet direction of the tail nozzle, and the tilt angle is 30°-45° with respect to the central axis of the tail nozzle; Preferably, the attitude control module is further included, the attitude control module is electrically connected with the X flow divider valve and the Y flow divider valve, the attitude control module can receive the attitude detection signal of the rocket, calculate the required thrust vector adjustment amount according to the signal, and send a control instruction to the X flow divider valve and the Y flow divider valve, so as to realize independent attitude control or cooperative attitude control in the X direction and the Y direction; Preferably, the inside of the side wall of the combustion chamber and the tail nozzle is a cavity, forming a cooling channel.

[0008] Compared with the prior art, the advantages of the present application are: The present application controls the on-off of four output pipes through two flow divider valves, and when three output pipes are simultaneously turned on, a pressure difference vector is superimposed to complete the precise deflection of high-speed main flow gas, so that the rocket can be quickly and accurately attitude corrected or yawed; The outlet ends of the four output pipes are flat, when the rocket needs to yaw or attitude correction, only one output pipe needs to be turned on, the sprayed gas forms a "wide coverage characteristic" oblique shock, thereby realizing the rocket thrust vector control, avoiding the loss of main stream gas energy, and giving consideration to the attitude accurate control and propulsion performance of the rocket; The outlet ends of the four output pipes are flat, when the rocket needs to yaw or attitude correction, only one output pipe needs to be turned on, the sprayed gas forms a "wide coverage characteristic" oblique shock, thereby realizing the rocket thrust vector control, avoiding the loss of main stream gas energy, and giving consideration to the attitude accurate control and propulsion performance of the rocket;

[0009] Additional aspects and advantages of the application will be described in the following description, some of which will become apparent from the following description, or will be learned by practice of the application. BRIEF DESCRIPTION OF DRAWINGS

[0010] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the following will briefly introduce the drawings needed to be used in the embodiments or the prior art description. Obviously, the drawings in the following description only show some embodiments of the present application, and other drawings can be obtained by those skilled in the art without creative labor under the premise of the drawings.

[0011] Fig. 1 It is a structural perspective view of the present application.

[0012] Fig. 2 It is an engine working principle diagram in embodiment one.

[0013] Fig. 3 It is a split valve structure exploded view in embodiment one.

[0014] Fig. 4 It is an engine working principle diagram in embodiment two.

[0015] Fig. 5 It is a split valve structure exploded view in embodiment two.

[0016] Fig. 6 It is a working principle diagram of the present application. DETAILED DESCRIPTION

[0017] The technical solutions in the embodiments of the present application will be described clearly and completely below, obviously, the described embodiments are only some of the embodiments of the present application, not all. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative labor are within the protection scope of the present application.

[0018] In the description of the present application, it should be noted that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer" and the like indicate the orientation or positional relationship shown in the drawings, and are only for the convenience of describing the present application and simplifying the description, and do not indicate or imply that the device or element referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore cannot be understood as a limitation on the present application.

[0019] In addition, in the description of the present application, unless otherwise explicitly specified and limited, the terms "mounting", "connection", "connection" and the like should be broadly understood, for example, it can be fixedly connected, or it can be detachably connected, or it can be integrally connected; it can be mechanically connected, or it can be electrically connected; it can be directly connected, or it can be indirectly connected through an intermediate medium; it can be the communication between two elements, it can be wireless connection, or it can be wired connection. For those skilled in the art, the specific meaning of the above terms in the present application can be understood according to the specific circumstances.

[0020] In addition, the technical features involved in the subsequent description of different embodiments of the present application can be combined with each other as long as they do not conflict with each other.

[0021] Please refer to Figs. 1-5 In the embodiment of the present application, a rocket engine for realizing attitude control by gas swing includes: an injector 100, a combustion chamber 200, and an exhaust nozzle 300. An X gas extraction pipe 400 is arranged outside the combustion chamber 200, and a first end of the X gas extraction pipe 400 penetrates the wall surface of the combustion chamber 200 and communicates with the inside of the combustion chamber 200. The other end of the X gas extraction pipe 400 is connected to an X flow divider valve 500, two output ends of the X flow divider valve 500 are respectively connected to an Xa output pipe 510 and an Xb output pipe 520, and the outlet ends of the Xa output pipe 510 and the Xb output pipe 520 oppositely extend into the inside of the exhaust nozzle 300. The X flow divider valve 500 can independently adjust the off and on states of the Xa output pipe 510 and the Xb output pipe 520, and can independently control the gas flow into the Xa output pipe 510 and the Xb output pipe 520. A Y gas extraction pipe 600 is arranged outside the combustion chamber 200, and a first end of the Y gas extraction pipe 600 penetrates the wall surface of the combustion chamber 200 and communicates with the inside of the combustion chamber 200. The other end of the Y gas extraction pipe 600 is connected to a Y flow divider valve 700, two output ends of the Y flow divider valve 700 are respectively connected to a Ya output pipe 710 and a Yb output pipe 720, and the outlet ends of the Ya output pipe 710 and the Yb output pipe 720 oppositely extend into the inside of the exhaust nozzle 300. The Y-flow diverter valve 700 can independently adjust the on / off state of the Ya output pipe 710 and the Yb output pipe 720 , and can independently control the gas flow rate flowing into the Ya output pipe 710 and the Yb output pipe 720 .

[0022] The valve cores of the X-diverter valve 500 and the Y-diverter valve 700 are both hollow structures, and both the valve cores of the X-diverter valve 500 and the Y-diverter valve 700 have through holes.

[0023] When the rocket's attitude needs to be controlled, the conduction state and gas flow rate of the valve core of the X-diverter valve 500 or the Y-diverter valve 700 are controlled so that the gas ejected from the Xa output pipe 510, the Xb output pipe 520, the Ya output pipe 710, and the Yb output pipe 720 produce different pressure distributions in the tail nozzle 300, thereby pushing the high-speed gas flow in the tail nozzle 300 to be deflected, thereby realizing thrust vector control and achieving the purpose of adjusting the rocket's attitude. Example 1

[0024] like Figs. 2-3 As shown, the outlet ends of the Xa output tube 510, the Xb output tube 520, the Ya output tube 710, and the Yb output tube 720 are all circular tubes, and the valve cores of the X diverter valve 500 and the Y diverter valve 700 each have two through holes. When the rocket needs to yaw or adjust its attitude in the Xb direction, the valve core of the X-diverter valve 500 rotates, aligning and connecting one of the through holes with the Xa output pipe 510. The high-temperature gas expanded in the gas chamber flows from the X-exhaust pipe 400 into the valve core of the X-diverter valve 500, then passes through the through hole and the Xa output pipe 510, and is ejected from the outlet end of the Xa output pipe 510 to form an oblique shock wave. The oblique shock wave interacts with the inner wall of the tail nozzle 300 and the mainstream gas, causing the local gas pressure on one side of the Xa output pipe 510 to rise sharply, forming a pressure difference with the Xb direction. This pressure difference generates lateral thrust on the mainstream gas in the tail nozzle 300, forcing the mainstream gas injection direction to deviate toward the Xb direction. Typically, a unidirectional oblique shock wave causes a Y-direction deviation in the ejection direction of the high-speed mainstream gas. In this embodiment, while the Xa output pipe 510 is connected, the two through-holes within the valve core of the Y diverter valve 700 are aligned and connected to the Ya output pipe 710 and the Yb output pipe 720, respectively. The high-temperature gas expanding within the gas chamber forms two additional oblique shock waves as it is ejected from the outlets of the Ya output pipe 710 and the Yb output pipe 720, respectively. These two additional oblique shock waves superimpose vectorially on the high-speed mainstream gas, preventing some Y-direction tilt of the high-speed mainstream gas and enabling precise yaw and attitude correction of the rocket. The simultaneous connection of the three pipes and the vectorial superposition of pressure differences simultaneously achieve precise deviations in the ejection direction of the high-speed mainstream gas, enabling thrust vector control of the rocket and rapid and precise attitude correction or yaw. Example 2

[0025] The outlet end of the Xa output pipe 510, the outlet end of the Xb output pipe 520, and the outlet end of the Ya output pipe 710 and the outlet end of the Yb output pipe 720 are flat, the long axis of the flat outlet extends radially along the tail nozzle 300, and the oblique shock wave of the "fan-shaped jet" formed by the ejected gas is formed, so that the oblique shock wave of the flat outlet has a wide coverage characteristic, and only a single oblique shock wave is needed to realize the precise deflection or precise attitude correction of the rocket. Figs. 4-6 As shown in the figure, in this embodiment, when the rocket needs to be deflected or the attitude needs to be corrected, only one output pipe needs to be turned on, so as to achieve the purpose of precise deflection of the high-speed main flow gas. This single wide-coverage oblique shock wave realizes the control of the rocket thrust vector, avoids the loss of high-speed main flow gas energy, and balances the precise attitude control and propulsion performance of the rocket.

[0026] In the embodiment of the present application, the outlet end of the Xa output pipe 510, the outlet end of the Xb output pipe 520, and the outlet end of the Ya output pipe 710 and the outlet end of the Yb output pipe 720 are inclined towards the outlet direction of the tail nozzle 300, and the inclination angle is 30°-45° with the central axis of the tail nozzle 300; if the outlet is perpendicular to the axis of the tail nozzle 300, the ejected gas will collide with the main flow gas at 90°, most of the energy will be converted into impact loss, and only a small part will form an effective oblique shock wave; when inclined at 30°-45°, the ejected gas and the main flow gas form an "angle of attack", which can more smoothly blend into the main flow field, form a stable oblique shock wave, and the "deflection force" of the oblique shock wave on the main flow gas is more concentrated, so that the deflection angle of the main flow gas can be improved under the same gas flow, and the response speed of the attitude control can be greatly improved.

[0027] In the embodiment of the present application, the present application also includes an attitude control module (not shown in the figure), which is electrically connected with the X flow divider 500 and the Y flow divider 700; the attitude control module can receive the attitude detection signal of the rocket, calculate the required thrust vector adjustment amount according to the signal, and send a control command to the X flow divider 500 and the Y flow divider 700 to realize independent attitude control or cooperative attitude control in the X direction and the Y direction. The X flow divider 500 and the Y flow divider 700 are filled with a cooling agent to prevent overheating of the system in the valve body.

[0028] The side wall of the combustion chamber 200 is a double-layer sleeve structure, and the cavity between the two layers forms an annular cooling channel. The side wall of the tail nozzle 300 also adopts a double-layer structure, and the cooling channel is in communication with the cooling channel of the combustion chamber 200. The cooling channel is spiral-shaped, which can effectively exchange heat, and the cooling medium is a propellant.

[0029] It will be apparent to those skilled in the art that the application is not limited to the details of the above-exemplified embodiments and that the present application can be implemented in other particular forms without departing from the spirit or essential characteristics of the present application. The embodiments should therefore be considered in all respects as illustrative and not restrictive, the scope of the application being indicated by the appended claims rather than by the above description, and all changes which come within the meaning and range of equivalency of the claims are therefore intended to be embraced therein.

Claims

1. A rocket engine that uses gas swing to achieve attitude control, characterized in that: include: Injector, combustion chamber, tail nozzle; An X-exhaust pipe is provided on the outside of the combustion chamber, and a first end of the X-exhaust pipe passes through the wall of the combustion chamber and communicates with the interior of the combustion chamber; The other end of the X exhaust pipe is connected to the X diverter valve, and the two output ends of the X diverter valve are respectively connected to the Xa output pipe and the Xb output pipe. The outlet end of the Xa output pipe and the outlet end of the Xb output pipe extend into the interior of the tail nozzle towards each other.

2. The rocket engine for achieving attitude control by utilizing gas swing according to claim 1, characterized in that: The X-diverter valve can separately adjust the shutoff and conduction states of the Xa output pipe and the Xb output pipe, and can independently control the gas flow rate flowing into the Xa output pipe and the Xb output pipe.

3. The rocket engine for achieving attitude control by utilizing gas swing according to claim 1, characterized in that: A Y exhaust pipe is provided on the outside of the combustion chamber, and the first end of the Y exhaust pipe passes through the wall of the combustion chamber and is connected to the interior of the combustion chamber; the other end of the Y exhaust pipe is connected to a Y diverter valve, and the two output ends of the Y diverter valve are respectively connected to a Ya output pipe and a Yb output pipe, and the outlet end of the Ya output pipe and the outlet end of the Yb output pipe extend into the interior of the tail nozzle opposite to each other.

4. The rocket engine for achieving attitude control by utilizing gas swing according to claim 3, characterized in that: The Y-flow diverter valve can separately adjust the shutoff and conduction states of the Ya output pipe and the Yb output pipe, and can independently control the gas flow rate flowing into the Ya output pipe and the Yb output pipe.

5. The rocket engine for achieving attitude control by utilizing gas swing according to claim 3, characterized in that: The valve cores of the X-diverter valve and the Y-diverter valve are both hollow structures, and the valve cores of the X-diverter valve and the Y-diverter valve each have two through holes.

6. The rocket engine for achieving attitude control by utilizing gas swing according to claim 3, characterized in that: The outlet ends of the Xa output pipe, the Xb output pipe, the Ya output pipe, and the Yb output pipe are all inclined toward the outlet direction of the tail nozzle, and the inclination angle is 30°-45° with the central axis of the tail nozzle.

7. The rocket engine for achieving attitude control by utilizing gas swing according to claim 3, characterized in that: It also includes an attitude control module, which is electrically connected to the X-diverter valve and the Y-diverter valve respectively; the attitude control module can receive the rocket's attitude detection signal, calculate the required thrust vector adjustment amount based on the signal, and send control instructions to the X-diverter valve and the Y-diverter valve to achieve independent attitude control or coordinated attitude control in the X and Y directions.

8. The rocket engine for achieving attitude control by utilizing gas swing according to claim 3, characterized in that: The interiors of the combustion chamber and the side walls of the tail nozzle are hollow, forming cooling channels.

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