阅读说明：本技术 一种用于验证垂直起降技术的可重复使用火箭及验证方法 (Reusable rocket for verifying vertical take-off and landing technology and verification method ) 是由 胡智珲 罗庶 马道远 李钧 姜航 赵学光 杨跃 刘浩 梁家伟 朱佩婕 陈辰 尹 于 2021-08-17 设计创作，主要内容包括：本发明涉及一种用于验证垂直起降技术的可重复使用火箭及验证方法。包括：主舱体,动力系统,吹除系统,控制系统,辅助动力系统,固定回收支腿系统；其中,控制软件中写入的飞行弹道按以下方法设计：S1、确定约束量要求及精度要求；S2、确定待迭代的各控制量；S3、给控制量赋初值；S4、根据约束量与控制量之间的关系,确定修正系数；S5、根据修正系数,计算修正量,进行弹道迭代计算；S6、获得满足精度要求的弹道。本方法降低了产品研制周期与生产成本；降低了后续进行大型可回收液体运载火箭研制风险；可在火箭上安装整流罩与栅格舵,能适应更高的飞行速度,可作为亚轨道探空火箭。(The invention relates to a reusable rocket for verifying a vertical take-off and landing technology and a verification method. The method comprises the following steps: the main cabin body, a power system, a blowing system, a control system, an auxiliary power system and a fixed recovery supporting leg system; the flight trajectory written in the control software is designed according to the following method: s1, determining the constraint quantity requirement and the precision requirement; s2, determining each control quantity to be iterated; s3, giving an initial value to the control quantity; s4, determining a correction coefficient according to the relation between the constraint quantity and the control quantity; s5, calculating a correction amount according to the correction coefficient, and performing ballistic iteration calculation; and S6, obtaining a trajectory meeting the precision requirement. The method reduces the product development period and the production cost; the risk of developing a large recyclable liquid carrier rocket in the follow-up process is reduced; the fairing and the grid rudder can be arranged on the rocket, can adapt to higher flight speed, and can be used as a sub-orbit sounding rocket.)

1. A reusable rocket for validating vertical take-off and landing techniques, comprising:

the main cabin body, a power system, a blowing system, a control system, an auxiliary power system and a fixed recovery supporting leg system;

wherein the content of the first and second substances,

the main cabin body is a cylindrical shell, the slenderness ratio of the main cabin body is close to that of a large carrier rocket, and the pneumatic appearance required by flight is guaranteed to be maintained in the processes of hoisting, transporting, erecting, launching and recovering;

the power system is arranged in the main cabin body and provides power for accelerating or decelerating the rocket;

the control system is arranged in the main cabin body and comprises a central computer provided with control software, a camera, a plurality of sensors, an inertial measurement unit, a GPS positioning device and a corresponding antenna, and all electric appliance parts are connected in the main cabin body through cables; the control system is used for resolving the current attitude, speed and position of the rocket body in real time, ensuring that the rocket flies along a target trajectory, and taking off and landing according to a designed time sequence;

the auxiliary power system is arranged in the main cabin body and used for injecting gas outwards and providing pitching, rolling and yawing moments for rocket control;

a blowing system is arranged near a power system in the main cabin body to ensure that each single machine in the cabin works stably and normally;

the fixed recovery supporting leg system is installed at the tail of the main cabin body, a buffering effect is provided when the rocket is recovered, landing buffering is reduced mainly through a buffer inside the supporting leg, safe landing of the rocket is ensured, and meanwhile, a heat-proof material needs to be wrapped on the supporting leg, so that the supporting leg can bear burning of tail flames of the engine when the supporting leg returns.

2. The reusable rocket for validating VTOL technology of claim 1 wherein the power system comprises a cryogenic liquid methane tank and cryogenic liquid oxygen tank, a nozzle, a cryogenic liquid engine, a servo mechanism for controlling nozzle swing, liquid oxygen and methane mixed in the nozzle at the tail of the main tank after flowing out of the tank are burned, a fuel storage tank and an oxidizer storage tank are both connected with the engine, the engine is connected with the nozzle, the nozzle is arranged at the tail of the main tank, and the gas outlet of the nozzle is open to the outside of the main tank.

3. A reusable rocket for validating VTOL techniques according to claim 1, wherein the flight trajectory written in the control software is designed as follows:

the invention discloses a rocket trajectory iterative computation method for verifying a vertical take-off and landing technology, which comprises the following steps of:

step S1, determining the constraint quantity requirement and the precision requirement;

the specific content of step S1 is:

the constraint requirements comprise the requirements of the height, the speed, the range and the azimuth angle of the landing point, wherein the constraint requirement of the tail end height is 0m, the precision requirement is 0.001m-0.1m, and the optimal precision requirement is 0.01 m; the end speed requirement is 0m/s, the precision requirement is 0.001m/s-0.1m/s, and 0.01m/s is preferred; the range accuracy requirement is 0.01m-1m, preferably 0.1 m; the precision requirement of the landing point azimuth angle is 0.01-1 degree, and preferably 0.1 degree;

step S2, determining each control quantity to be iterated;

the specific content of step S2 is:

determining control quantity to be iterated, and firstly dividing the trajectory into an acceleration ascending section, a first pushing section, a deceleration ascending section, an acceleration descending section, a second pushing section and a deceleration descending section, wherein the deceleration returning section end height HH is one of the control quantity, and the HH is used for controlling the height of a drop point to meet the accuracy requirement of the height of the drop point; the attitude angle of the whole trajectory is then designed, wherein the characteristic variable characterizing the pitch angleIn order to control one of the quantities, the control unit,the device is used for controlling the landing range to meet the requirement of range precision; the firing angle A0 is a third control variable, A0 is used for controlling the landing point azimuth angle to enable the landing point azimuth angle to meet the accuracy requirement of the landing point azimuth angle, the firing angle is defined as the included angle between the emission aiming direction and the due north direction of the emission point, and the landing point azimuth angle is defined as the included angle between the direction of the landing point relative to the emission point and the due north direction of the emission point;

step S3, giving an initial value to the control quantity;

the specific content of step S3 is:

assigning an initial value to the control variable, the initial value enabling rapid convergence of ballistic iteration calculations;

step S4, determining a correction coefficient according to the relation between the constraint quantity and the control quantity;

the specific content of step S4 is:

when determining the correction coefficient, the unit of the angle of incidence and the azimuth of the landing point, the unit of the final height of the deceleration return section and the unit of the landing height are the sameA more obvious monotonous relation exists, and the correction coefficient is set to be 1; range andthe correction coefficient of (d) is calculated as follows: recording the range L and L every time a trajectory is calculatedValue, then the correction factor is

Step S5, calculating a correction amount according to the correction coefficient, and performing ballistic iteration calculation;

specifically, when determining each control quantity to be iterated when performing trajectory iteration calculation in step S5, the time of each flight segment is not used as the control quantity, but the flight segments are divided into six flight segments, and different physical quantities which are more convenient to calculate are used as the control quantities: the method comprises the following steps that an accelerating ascending section, a first pushing adjusting section, a decelerating ascending section, an accelerating descending section, a second pushing adjusting section and a decelerating descending section are adopted, the thrust of each section in the six sections is different, the pitch angle of each section is different, and the AO shooting angle is determined initially but the coordinate system transformation is carried out in the whole process for calculation; the height at which the acceleration falling section ends in the acceleration falling section is taken as the control amount HH;

and step S6, obtaining a trajectory meeting the precision requirement.

4. A reusable rocket for validating vertical take-off and landing techniques according to any one of claims 1-3, wherein said rocket is designed to take-off and land at a timing sequence designed in the following way,

the flight time is designed in sections, and the axial flight overload of each corresponding altitude section is matched, so that the axial flight overload of each corresponding altitude section is suitable for the verification flight of the vertical recovery verification rocket, and the specific design method comprises the following steps:

the thrust of the engine is adjusted according to the time sequence, so that the staggered change of the axial flying overload of the rocket body is realized, the proper axial flying overload is provided for the whole process of the vertical recovery demonstration and verification of the rocket flying, and the vertical controlled landing recovery can be finally realized;

the flight time is segmented into a takeoff ascending section, a first deceleration ascending and then acceleration descending section and a deceleration descending section, the takeoff ascending section corresponds to a flight height h1, the first deceleration ascending and then acceleration descending section corresponds to a flight height h2, the deceleration descending section corresponds to a flight height h3, thrust adjustment is performed on the engine in each stage, the takeoff ascending section, the first deceleration ascending and then acceleration descending section and the deceleration descending section are determined according to thrust-quality-height-overload iterative calculation, the magnitude of the thrust adjustment (namely the propellant quality) is performed on the engine in each stage, and staggered change of rocket body axial flight overload is achieved.

5. A sub-orbital sounding rocket, characterized in that a reusable rocket for verifying VTOL techniques according to any one of claims 1-4 is equipped with fairings and grid rudders, able to adapt to higher flight speeds.

6. A method for validating a vertical take-off and landing technique, a reusable rocket for validating the vertical take-off and landing technique according to any one of claims 1 to 4, characterized by comprising the following three experimental contents:

the reusable rocket performs static ignition work;

the reusable rocket performs mooring protection ignition work;

the reusable rocket performs vertical flight.

7. The method for validating VTOL technology of claim 6, wherein the reusable rocket performs static ignition specifically comprises: carrying out a ground static ignition test, fixing the rocket on a launching tower, and igniting the engine; the engine is circularly precooled by using the low-temperature oxygen tank and the methane tank, so that the precooling flow of the engine and the correctness of precooling program setting can be verified under a low-temperature environment; the oxygen tank and the methane tank adopt a self-generated pressurization scheme, and the design of key parameters such as pressurization scheme selection, pressurization flow, pressurization temperature and the like can be verified; and after the two storage tanks are filled, ignition is carried out, the servo mechanism is used for controlling the engine to swing and spray, and the control system is used for sending an instruction to adjust the thrust of the engine.

8. The method for validating VTOL technology of claim 6, wherein the reusable rocket is tethered protected by ignition specifically comprising: the rocket is lifted off the ground, and the engine is in a full-range starting ignition state. The method comprises the following steps that firstly, the thrust of an engine is larger than the gravity of a rocket, the rocket ascends, then the thrust of the engine is adjusted, so that the gravity of the rocket is larger than the thrust of the engine, the rocket starts to fall back after the rocket ascends in a deceleration manner, at the moment, the weight of the rocket is reduced due to fuel consumption, the thrust of the engine is larger than the gravity of the rocket again, and when the rocket descends to a takeoff height in a deceleration manner, the engine is shut down, and the rocket is lifted; the whole-course engine action is controlled by the control system; the scheme verifies the low-temperature power technology and the engine thrust adjusting technology.

9. Method for validating a VTOL technique according to claim 6, characterized in that

The specific contents of the vertical flight of the reusable rocket are as follows: placing the rocket on the ground, so that the rocket is subjected to a flight test in a free state, the height H of the top point of the trajectory is high, and the engine is not shut down in the whole process after being started; starting the thrust of the engine to be larger than the gravity of the rocket, so that the rocket ascends in an accelerated way; then adjusting the thrust of the engine to be smaller than the gravity of the rocket, and accelerating the rocket to descend after the rocket decelerates and rises to the top of a trajectory; when the calculated height is reached, the thrust of the engine is adjusted to be larger than the gravity of the rocket, and the rocket is decelerated and descends; when the rocket is near the ground, the engine is shut down, and the landing support legs are used for buffering to finish rocket landing; the whole-course engine action is controlled by a control system, and an instruction is sent out at the preset time of a program after the calculation according to the position and the speed of the rocket; the engine works in the whole course of the flight, and the engine is started, thrust is adjusted, the engine is shut down according to the designated time, and the engine thrust adjusting technology is verified; the rocket returns after flying to the height H, the control system measures the position, speed and other parameters of the rocket in the period, the rocket is stably ascended and returned by utilizing the engine swing jet and the jet of the auxiliary power system, and the high-precision navigation and high-precision guidance technology is verified; starting the engine in a low-temperature state, providing power for the engine by using liquid oxymethane in the storage tank, and verifying a low-temperature power technology; and when the landing leg is close to the ground, the engine is closed, and the landing leg is buffered and landed by using the landing leg, so that the landing leg buffering technology is verified.

Technical Field

The application relates to the technical field of carrier rockets, in particular to a reusable rocket for verifying a vertical take-off and landing technology and a verification method.

Background

Compared with the traditional disposable rocket, the reusable rocket adopting the vertical take-off and landing technology has the greatest difference in whether the first sub-stage of the rocket can be recovered or not. With the development of commercial aerospace, the market demands for low-cost rockets are urgent, and the realization of commercial launching tasks by using high-reliability, low-cost and reusable rockets has become a future development trend. The SpaceX company successfully uses the falcon 9 rocket to finish the satellite launching task for multiple times, and successfully recovers a first sub-stage of the rocket by adopting a vertical take-off and landing technology, so that the rocket launching cost is greatly reduced; the new Scherbard aircraft of the blue origin company successfully completes the sub-orbital vertical recovery test, verifies a series of key technologies such as guidance, engine variable thrust, landing buffer support legs and the like, and lays a foundation for the subsequent development of a new gren large-scale reusable rocket. The reusable rocket enables the original price advantage of the rocket launching service in China to be no longer existed, but the rocket launching service in China is late in the aspect. However, the direct development of the reusable rocket has great technical difficulty and high risk, and the key technical challenges and technical verification of vertical take-off and landing need to be developed first.

In the vertical take-off and landing technology, various modes such as parachute landing recovery, vertical landing leg recovery, winged flying return and the like can be adopted during return, but the parachute landing recovery drop point precision is low, and the requirement on the structural strength of the rocket body is high; the return section has poor pneumatic performance when the winged flying-back device flies back, the research and development cost and the system complexity are comprehensively considered, and meanwhile, the successful experience of predecessors is used for reference.

Disclosure of Invention

The application provides a reusable rocket for verifying a vertical take-off and landing technology and a verification method, so as to achieve the purpose of verifying the vertical take-off and landing key technology of the reusable rocket.

The application provides the following technical scheme:

a reusable rocket for validating vertical take-off and landing techniques, comprising:

the main cabin body, a power system, a blowing system, a control system, an auxiliary power system and a fixed recovery supporting leg system;

wherein the content of the first and second substances,

the main cabin body is a cylindrical shell, the slenderness ratio of the main cabin body is close to that of a large carrier rocket, and the pneumatic appearance required by flight is guaranteed to be maintained in the processes of hoisting, transporting, erecting, launching and recovering;

the power system is arranged in the main cabin body and provides power for accelerating or decelerating the rocket;

the control system is arranged in the main cabin (generally on the head), and comprises a central computer provided with control software, a camera, a plurality of sensors, an inertial measurement unit, a GPS positioning device and a corresponding antenna, wherein all electrical components are connected in the main cabin through cables; the control system is used for resolving the current attitude, speed and position of the rocket body in real time, ensuring that the rocket flies along a target trajectory, and taking off and landing according to a designed time sequence; specifically, a central computer, a camera and a measurement and control device (including a GPS positioning device and an antenna) are all fixed on a bulkhead through a bracket, and a buffer device is arranged between the bracket and each single-unit inertial unit; the camera is fixed at the head and the tail through a damping device; each single unit inertial unit is connected through a cable on the surface of the cabin body, and a cable cover is arranged on the surface of the cable for protection;

the auxiliary power system is arranged in the main cabin body (generally, the auxiliary power system can be arranged at the head part) and is used for injecting gas outwards to provide pitching, rolling and yawing moments for rocket control;

a blowing system is arranged near a power system in the main cabin body to ensure that each single machine in the cabin works stably and normally;

the fixed recovery supporting leg system is installed at the tail of the main cabin body, a buffering effect is provided when the rocket is recovered, landing buffering is reduced mainly through a buffer inside the supporting leg, safe landing of the rocket is ensured, and meanwhile, a heat-proof material needs to be wrapped on the supporting leg, so that the supporting leg can bear burning of tail flames of the engine when the supporting leg returns.

Furthermore, the power system comprises a low-temperature liquid methane tank, a low-temperature liquid oxygen tank, a spray pipe, a low-temperature liquid engine and a servo mechanism capable of controlling the spray pipe to swing, the control system sends an instruction to control the swing of the control system to provide rocket thrust and partial operating torque, liquid oxygen and methane flow out of the tank body and then are mixed to enter the spray pipe at the tail part of the main cabin body for combustion, the fuel storage tank and the oxidant storage tank are both connected with the engine, the engine is connected with the spray pipe, the spray pipe is arranged at the tail end of the main cabin body, and a fuel gas outlet of the spray pipe leads to the outside of the main cabin body.

Further, the flight trajectory written in the control software is designed according to the following method:

the invention discloses a rocket trajectory iterative computation method for verifying a vertical take-off and landing technology, which comprises the following steps of:

step S1, determining the constraint quantity requirement and the precision requirement;

the specific content of step S1 is:

the constraint requirements comprise the requirements of the height, the speed, the range and the azimuth angle of the landing point, wherein the constraint requirement of the tail end height is 0m, the precision requirement is 0.001m-0.1m, and the optimal precision requirement is 0.01 m; the end speed requirement is 0m/s, the precision requirement is 0.001m/s-0.1m/s, and 0.01m/s is preferred; the range accuracy requirement is 0.01m-1m, preferably 0.1 m; the precision requirement of the landing point azimuth angle is 0.01-1 degree, and preferably 0.1 degree;

step S2, determining each control quantity to be iterated;

the specific content of step S2 is:

determining control quantity to be iterated, and firstly dividing the trajectory into an acceleration ascending section, a first pushing section, a deceleration ascending section, an acceleration descending section, a second pushing section and a deceleration descending section, wherein the deceleration returning section end height HH is one of the control quantity, and the HH is used for controlling the height of a drop point to meet the accuracy requirement of the height of the drop point; the attitude angle of the whole trajectory is then designed, wherein the characteristic variable characterizing the pitch angleIn order to control one of the quantities, the control unit,the device is used for controlling the landing range to meet the requirement of range precision; the firing angle A0 is a third control variable, A0 is used for controlling the landing point azimuth angle to enable the landing point azimuth angle to meet the accuracy requirement of the landing point azimuth angle, the firing angle is defined as the included angle between the emission aiming direction and the due north direction of the emission point, and the landing point azimuth angle is defined as the included angle between the direction of the landing point relative to the emission point and the due north direction of the emission point;

step S3, giving an initial value to the control quantity;

the specific content of step S3 is:

assigning an initial value to the control variable, the initial value enabling rapid convergence of ballistic iteration calculations;

step S4, determining a correction coefficient according to the relation between the constraint quantity and the control quantity;

the specific content of step S4 is:

when the correction coefficient is determined, the unit of the angle of incidence and the azimuth of the landing point, the unit of the end height of the deceleration return section and the unit of the landing height are the same and have a more obvious monotonous relation (monotonous increasing or monotonous decreasing), and the correction coefficient is set to be 1; range andthe correction coefficient of (d) is calculated as follows: recording the range L and L every time a trajectory is calculatedValue, then the correction factor is

Step S5, calculating a correction amount according to the correction coefficient, and performing ballistic iteration calculation;

specifically, when determining each control quantity to be iterated when performing trajectory iteration calculation in step S5, the time of each flight segment is not used as the control quantity, but the flight segments are divided into six flight segments, and different physical quantities which are more convenient to calculate are used as the control quantities: the method comprises the following steps that an accelerating ascending section, a first pushing adjusting section, a decelerating ascending section, an accelerating descending section, a second pushing adjusting section and a decelerating descending section are adopted, the thrust of each section in the six sections is different, the pitch angle of each section is different, and the AO shooting angle is determined initially but the coordinate system transformation is carried out in the whole process for calculation; the height at which the acceleration falling section ends in the acceleration falling section is taken as the control amount HH;

and step S6, obtaining a trajectory meeting the precision requirement.

Furthermore, the rocket takes off and lands according to the time sequence designed by the following method,

the flight time is designed in sections, and the axial flight overload of each corresponding altitude section is matched, so that the axial flight overload of each corresponding altitude section is suitable for the verification flight of the vertical recovery verification rocket, and the specific design method comprises the following steps:

the thrust of the engine is adjusted according to the time sequence, so that the staggered change of the axial flying overload of the rocket body is realized, the proper axial flying overload is provided for the whole process of the vertical recovery demonstration and verification of the rocket flying, and the vertical controlled landing recovery can be finally realized;

the flight time is segmented into a takeoff ascending section, a first deceleration ascending and then acceleration descending section and a deceleration descending section, the takeoff ascending section corresponds to a flight height h1 (the height can be obtained according to the flight time and overload), the first deceleration ascending and then acceleration descending section corresponds to a flight height h2 (the height is obtained according to the flight time and overload level accumulation of each section), the deceleration descending section corresponds to a flight height h3, the engine is subjected to thrust adjustment in each stage, specifically, the takeoff ascending section, the first deceleration ascending and then acceleration descending section and the deceleration descending section are determined according to thrust-mass-height-overload iterative calculation, the magnitude of the thrust adjustment (namely, the mass of the propellant) is carried out on the engine in each stage, and the staggered change of the rocket body axial flight overload is realized.

The invention also provides a sub-orbit sounding rocket, and a fairing and a grid rudder are arranged on the reusable rocket for verifying the vertical take-off and landing technology, so that the rocket can adapt to higher flight speed.

The invention also provides a method for verifying the vertical take-off and landing technology, which is performed according to the reusable rocket for verifying the vertical take-off and landing technology, and comprises the following three experimental contents:

the reusable rocket performs static ignition work;

the reusable rocket performs mooring protection ignition work;

the reusable rocket performs vertical flight.

Specifically, the specific contents of the reusable rocket for static ignition work are as follows: carrying out a ground static ignition test, fixing the rocket on a launching tower, and igniting the engine; the engine is circularly precooled by using the low-temperature oxygen tank and the methane tank, so that the precooling flow of the engine and the correctness of precooling program setting can be verified under a low-temperature environment; the oxygen tank and the methane tank adopt a self-generated pressurization scheme, and the design of key parameters such as pressurization scheme selection, pressurization flow, pressurization temperature and the like can be verified; and after the two storage tanks are filled, ignition is carried out, the servo mechanism is used for controlling the engine to swing and spray, and the control system is used for sending an instruction to adjust the thrust of the engine. Therefore, the low-temperature starting technology and the engine thrust adjusting technology of the engine are verified.

Specifically, the reusable rocket for mooring protection ignition specifically comprises the following contents: the rocket is lifted off the ground, and the engine is in a full-range starting ignition state. The method comprises the following steps that firstly, the thrust of an engine is larger than the gravity of a rocket, the rocket ascends, then the thrust of the engine is adjusted, so that the gravity of the rocket is larger than the thrust of the engine, the rocket starts to fall back after the rocket ascends in a deceleration manner, at the moment, the weight of the rocket is reduced due to fuel consumption, the thrust of the engine is larger than the gravity of the rocket again, and when the rocket descends to a takeoff height in a deceleration manner, the engine is shut down, and the rocket is lifted; the whole-course engine action is controlled by the control system; the scheme verifies the low-temperature power technology and the engine thrust adjusting technology. Under the working condition of mooring protection ignition work, the flight speed of the rocket is low, and random wind has large interference on the rocket, so that a high-precision navigation technology and a high-precision guidance technology are verified;

specifically, the reusable rocket specifically comprises the following contents of vertical flight: placing the rocket on the ground, so that the rocket is subjected to a flight test in a free state, the height H of the top point of the trajectory is high, and the engine is not shut down in the whole process after being started; starting the thrust of the engine to be larger than the gravity of the rocket, so that the rocket ascends in an accelerated way; then adjusting the thrust of the engine to be smaller than the gravity of the rocket, and accelerating the rocket to descend after the rocket decelerates and rises to the top of a trajectory; when the calculated height is reached, the thrust of the engine is adjusted to be larger than the gravity of the rocket, and the rocket is decelerated and descends; when the rocket is near the ground, the engine is shut down, and the landing support legs are used for buffering to finish rocket landing; the whole-course engine action is controlled by a control system, and an instruction is sent out at the preset time of a program after the calculation according to the position and the speed of the rocket; the engine works in the whole course of the flight, and the engine is started, thrust is adjusted, the engine is shut down according to the designated time, and the engine thrust adjusting technology is verified; the rocket returns after flying to the height H, the control system measures the position, speed and other parameters of the rocket in the period, the rocket is stably ascended and returned by utilizing the engine swing jet and the jet of the auxiliary power system, and the high-precision navigation and high-precision guidance technology is verified; starting the engine in a low-temperature state, providing power for the engine by using liquid oxymethane in the storage tank, and verifying a low-temperature power technology; and when the landing leg is close to the ground, the engine is closed, and the landing leg is buffered and landed by using the landing leg, so that the landing leg buffering technology is verified.

In general, the above technical solutions contemplated by the present invention can achieve the following beneficial effects:

1. a shape design can verify the calculation method of a blowing system (normal temperature gas is used for blowing the interior of a cabin, the temperature is not too low, the calculation method needs to be calculated for long time, the calculation method is a conventional algorithm), the vertical rocket trajectory calculation method of single start-up, the low-temperature engine thrust adjustment technology, the high-precision navigation, the high-precision guidance technology and the landing leg buffering technology, so that the product development period and the production cost are reduced;

2. the rocket completes the verification of the vertical take-off and landing technology, and reduces the risk of developing a large recyclable liquid carrier rocket in the follow-up process;

3. the fairing and the grid rudder can be arranged on the rocket, can adapt to higher flight speed, and can be used as a sub-orbit sounding rocket.

Drawings

FIG. 1 is a top view of a rocket according to the present invention;

FIG. 2 is a cross-sectional view taken along line A-A of FIG. 1;

FIG. 3 is a schematic diagram of a launching system of the rocket trajectory calculation method, in which an origin of coordinates O1 is fixedly connected with a launching point, an x-axis is in a launching plane and points to a launching aiming direction, a y-axis is perpendicular to a horizontal plane of the launching point and points to a sky direction, and a z-axis and the x-axis and the y-axis form a right-hand rectangular coordinate system; a0 is the firing angle; the O point is the origin of the coordinate system of the earth;

FIG. 4 is a schematic flow chart of a rocket trajectory calculation method of the present invention, wherein t is a time variable from the takeoff time; h is a ballistic height variable from the takeoff moment and changes along with t; Vy-Y direction velocity of emission system, Vy 1-Y direction velocity judgment condition of emission system ending in acceleration ascending section, HH-height judgment condition of ending in acceleration descending section, H-altitude, H₀-landing altitude requirement, l-range, l₀-range constraint requirement, Az-landing azimuth, Az 0-landing azimuth constraint requirement;

FIG. 5 is a sectional view of the ballistic flight of the rocket according to the present invention, in which a thrust adjusting section I is a first thrust adjusting section, a thrust adjusting section II is a second thrust adjusting section, and the thrust adjusting section is a thrust adjusting section;

FIG. 6 is a flow chart of a rocket flight timing design method of the present invention;

FIG. 7 is a timing diagram illustrating rocket thrust adjustment of the present invention;

FIG. 8 is a flow chart of a method for determining the pressure condition of a buffer in the rocket destaging determination method according to the present invention;

FIG. 9 is a flow chart of a ground clearance condition determination method in the rocket landing determination method of the present invention;

FIG. 10 is a flow chart of a rocket body axial overload condition discrimination method in the rocket destage discrimination method of the present invention;

FIG. 11 is a flow chart of a rocket destaging determination method in the rocket destaging determination method of the invention;

fig. 12 is an arrow coordinate system corresponding to the rocket destaging determination method of the present invention.

Detailed Description

In order to make the objects, technical solutions and advantages of the present invention more apparent, the present invention is described in further detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative of the invention and are not intended to limit the invention. In addition, the technical features involved in the embodiments of the present invention described below may be combined with each other as long as they do not conflict with each other.

1-2, a reusable rocket for validating vertical take-off and landing techniques, comprising:

the main cabin body, a power system, a blowing system, a control system, an auxiliary power system and a fixed recovery supporting leg system;

wherein the content of the first and second substances,

the main cabin body is a cylindrical shell, the slenderness ratio of the main cabin body is close to that of a large carrier rocket, and the pneumatic appearance required by flight is guaranteed to be maintained in the processes of hoisting, transporting, erecting, launching and recovering;

the power system is arranged in the main cabin body and provides power for accelerating or decelerating the rocket;

the control system is arranged in the main cabin (generally at the head) and comprises a central computer provided with control software, a camera, a plurality of sensors, an inertial measurement unit, a GPS positioning device and a corresponding antenna, and all electrical components are connected in the main cabin through cables; the control system is used for resolving the current attitude, speed and position of the rocket body in real time, ensuring that the rocket flies along a target trajectory, and taking off and landing according to a designed time sequence; specifically, the central computer, the camera and the measurement and control device (including the GPS positioning and the antenna) are all fixed on the bulkhead through a support, and a buffer device is arranged between the support and each single-unit inertial unit. The camera is fixed at the head and the tail through a damping device. Each single unit inertial unit is connected through cables on the surface of the cabin body, and cable covers are arranged on the surfaces of the cables for protection.

The auxiliary power system is arranged in the main cabin body (generally at the head) and used for injecting gas outwards and providing pitching, rolling and yawing moments for rocket control;

a blowing system is arranged near a power system in the main cabin body to ensure that each single machine in the cabin works stably and normally;

the fixed recovery supporting leg system is installed at the tail of the main cabin body, a buffering effect is provided when the rocket is recovered, landing buffering is reduced mainly through a buffer inside the supporting leg, safe landing of the rocket is ensured, and meanwhile, a heat-proof material needs to be wrapped on the supporting leg, so that the supporting leg can bear burning of tail flames of the engine when the supporting leg returns.

Specifically, the power system comprises a low-temperature liquid methane tank, a low-temperature liquid oxygen tank, a spray pipe, a low-temperature liquid engine and a servo mechanism capable of controlling the spray pipe to swing, wherein the control system sends an instruction to control the swing of the control system and provide rocket thrust and partial operating torque, liquid oxygen and methane flow out of the tank body and then are mixed to enter the spray pipe at the tail part of the main cabin body for combustion, a fuel storage tank and an oxidant storage tank are connected with the engine, the engine is connected with the spray pipe, the spray pipe is arranged at the tail end of the main cabin body, and a fuel gas outlet of the spray pipe leads to the outside of the main cabin body.

As shown in fig. 3-5, the flight trajectory written in the control software is designed according to the following method (i.e. a single-start vertical rocket trajectory calculation method):

firstly, defining an emission system: the origin of coordinates O1 is fixedly connected with the emission point, the x axis is in the emission plane and points to the emission aiming direction, the y axis is perpendicular to the horizontal plane of the emission point and points to the sky direction, and the z axis and the x and y axes form a right-hand rectangular coordinate system.

The direction angle is as follows: the direction angle is the included angle between the emission aiming direction and the north direction of the emission point, and the clockwise direction is positive when viewed from the y axis.

The landing point azimuth angle: the included angle between the direction of the landing point relative to the emission point and the north direction of the emission point is positive clockwise when viewed from the y axis.

The ballistic iteration calculation method comprises the following steps:

step S1, determining the constraint quantity requirement and the precision requirement;

the specific content of step S1 is:

the constraint requirements comprise the requirements of the height, the speed, the range and the azimuth angle of the landing point, wherein the constraint requirement of the tail end height is 0m, the precision requirement is 0.001m-0.1m, and the optimal precision requirement is 0.01 m; the end speed requirement is 0m/s, the precision requirement is 0.001m/s-0.1m/s, and 0.01m/s is preferred; the range accuracy requirement is 0.01m-1m, preferably 0.1 m; the accuracy of the landing azimuth angle is required to be 0.01-1 degrees, and 0.1 degree is preferred.

Step S2, determining each control quantity to be iterated;

the specific content of step S2 is:

determining control quantity to be iterated, and firstly dividing the trajectory into an acceleration ascending section, a first pushing section, a deceleration ascending section, an acceleration descending section, a second pushing section and a deceleration descending section, wherein the deceleration returning section end height HH is one of the control quantity, and the HH is used for controlling the height of a drop point to meet the accuracy requirement of the height of the drop point; the attitude angle of the whole trajectory is then designed, wherein the characteristic variable characterizing the pitch angleIn order to control one of the quantities, the control unit,the device is used for controlling the landing range to meet the requirement of range precision; the firing angle A0 is a third control variable, A0 is used for controlling the landing point azimuth angle to enable the landing point azimuth angle to meet the accuracy requirement of the landing point azimuth angle, the firing angle is defined as the included angle between the emission aiming direction and the due north direction of the emission point, and the landing point azimuth angle is defined as the included angle between the direction of the landing point relative to the emission point and the due north direction of the emission point;

in practical applications, in step S2,

the condition for judging the end of the accelerated rising segment is that the Y-direction speed of the launching system reaches a specific value Vy1, and the value is set to be related to the height of a ballistic vertex: the end time of this segment is T1 (i.e., T1 in FIG. 4); vy1 is monotonically increasing in relation to ballistic apex height.

The duration Tm1 of the first thrust adjusting section is related to the performance of the engine for adjusting the thrust and the thrust adjusting amount; the end time of the segment is t 2; for example, if the engine thrust adjustment performance requires 1s of time for adjusting 10% of the thrust (assuming that the engine rated thrust (full thrust) is P, the engine can adjust 10% P per second), the thrust adjustment amount is adjusted from 70% to 50%, and the adjustment amount is 20%, the time period Tm1 requires 2 × 1s — 2 s;

the condition for judging the end of the deceleration ascending section is that the Y-direction speed is 0; the end time of the segment is t 3;

the accelerated descending section ending judgment condition is that the height reaches a control amount HH; the end time of this segment is T4 (i.e., T4 in FIG. 4);

the duration Tm2 of the second thrust adjusting section is related to the performance of the engine for adjusting the thrust and the thrust adjusting amount; the end time of the segment is t 5; for example: assuming the engine rated thrust (full thrust) is P, the engine can adjust 10% P per second, which requires 2s if adjusted from 70% P to 50% P.

The deceleration descending section ending determination condition is that the Y-direction speed is 0, and the ending time of this section is t 6.

The time when each segment ends is T1 (i.e., T1 in fig. 4), T2, T3, T4 (i.e., T4 in fig. 4), T5, and T6 in chronological order.

Step S3, giving an initial value to the control quantity;

the specific content of step S3 is:

assigning an initial value to the control variable, the initial value enabling rapid convergence of ballistic iteration calculations;

in practical applications, specifically, in the step S3, an initial value is assigned to the control variable, and when the constraint quantity (the constraint quantity includes a range, a landing azimuth, a landing speed of 0, a vertical speed returning to zero is 0 (a landing criterion is determined), a lateral speed returning to zero is 0, and an altitude is 0) is determined, since the initial value of the firing angle is consistent with the defining direction of the landing azimuth and the initial reference, the initial value of the firing angle is determined to be consistent with the target landing azimuth Azi₀Equal; the initial value of the height of the deceleration return section end is necessarily at the altitude H of the flying point₀And ballistic vertex H_mBetween altitudes, let HH be x_H(H_m-H₀) Wherein x is_HE (0,1), since the judgment condition for the end of the deceleration descending section is that the Y-direction speed is 0m/s, if x is_HToo small, a negative height at the end of the descent deceleration phase may occur, at which time no meaningful atmospheric parameters can be solved, so x_HThe value interval of (2) is set between (0.5, 1.0); characteristic amountThe initial value of (1) is (0 degree, 90 degrees), and can be determined according to the range L (i.e. L, L in figure 4)₀Is a range constraint requirement) requires a selection, for example:for a range of 50mThe problem of iterative divergence can be caused by overlarge attitude angle, and iterative convergence is facilitated by a smaller initial value.

In practical application, the control quantity is not required to be designed for meeting the speed requirements of three directions of a landing point, the ending judgment condition of the deceleration descending section is set to be that the Y-direction speed of the launching system is 0, and the pitching angle is passedAnd the yaw angle psi is designed, so that the thrust direction and the speed direction of the return section rocket are opposite, the x-direction speed and the z-direction speed of the launching system can return to 0 when landing, and the rolling angle is 0 in the whole course because the rolling angle does not influence the calculation of the three-degree-of-freedom trajectory.

The attitude angle comprises a pitch angleAnd a yaw angle psi, the specific design content being:

in the accelerated ascending section: the pitch angle is 90 degrees and the yaw angle is 0 degree;

in the first tuning and pushing section: the pitch angle is transited from 90 DEG to a uniform speedThe yaw angle is 0 degree;

in the deceleration ascending section: a pitch angle ofThe yaw angle is 0 degree;

in the accelerating descending section, the second pushing adjusting section and the decelerating descending section: from the pitch angleAt a constant speed transits toAngular velocity to maintain arrow attitude stabilityShould not be too large, then the pitch angle is maintainedFlying, calculating the trajectory inclination angle theta and trajectory deviation angle sigma whenRecord the momentFrom that moment on, orderThe return leg pitch procedure is shown as follows:

the yaw angles of the acceleration descending section and the second adjusting and pushing section are 0 degrees, the trajectory yaw angle sigma is calculated in the deceleration descending section, and the yaw angle psi is-sigma;

vx, Vy and Vz are the velocity of the emission system X, Y, Z, V is the velocity of the emission system, a has no meaning, and a is used together with sin and has the meaning of an inverse trigonometric function arcsin.

Specifically, in the deceleration descent segment, if the approaching landing speed is too low, the calculated trajectory inclination angle and trajectory deflection angle may have large jump, and the processing mode is that the current attitude angle is recorded at the moment when the speed is less than 0.1m/sψ_fFrom that moment, let the pitch and yaw angle freeze atψ_fUntil the Y-direction speed reaches 0 m/s.

Step S4, determining a correction coefficient according to the relation between the constraint quantity and the control quantity;

the specific content of step S4 is:

when the correction coefficient is determined, the unit of the angle of incidence and the azimuth of the landing point, the unit of the end height of the accelerated descent section and the unit of the landing height are the same and have a more obvious monotonous (monotonous increasing or monotonous decreasing) relationship, and the correction coefficient is set to be 1; range andthe correction coefficient of (d) is calculated as follows: recording the range L and L every time a trajectory is calculatedValue, then the correction factor is

The basic rule is that where the rocket is launched and then returns to landing, the landing azimuth monotonically increases as the firing increases. Because the height of the trajectory vertex is fixed, the higher the ending height of the acceleration descending section is, the smaller the Y-direction speed Vy of the launching system at the ending moment of the acceleration descending section is, the unchanged duration of the pushing adjusting section is, the smaller the starting moment Vy of the deceleration descending section is, and the deceleration descending section decelerates from a smaller speed

When the velocity is 0, the smaller the height difference at the beginning of the segment, the higher the altitude at the landing time (the velocity of the emission system Y is O), and the landing altitude monotonically increases with the increase of the terminal altitude of the acceleration-descent segment.

Step S5, calculating a correction amount according to the correction coefficient, and performing ballistic iteration calculation;

specifically, when determining each control quantity to be iterated when performing trajectory iteration calculation in step S5, the time of each flight segment is not used as the control quantity, but the flight segments are divided into six flight segments, and different physical quantities which are more convenient to calculate are used as the control quantities: the method comprises the following steps that an accelerating ascending section, a first pushing adjusting section, a decelerating ascending section, an accelerating descending section, a second pushing adjusting section and a decelerating descending section are adopted, the thrust of each section in the six sections is different, the pitch angle of each section is different, and the AO shooting angle is determined initially but the coordinate system transformation is carried out in the whole process for calculation; the height at which the acceleration-down section ends is taken as the control amount HH in the acceleration-down section.

And step S6, obtaining a trajectory meeting the precision requirement.

As shown in fig. 6-7, the rocket further includes a measuring system for acquiring the flight state information and navigation position information of the rocket body in real time, after the main engine is ignited after the rocket starts to fly, determining whether the rocket body is off the platform according to a predetermined off-platform detection program, and sending a continuous work or emergency shutdown signal to the main engine; judging whether the preset time for departure is generally s5-7 s; after the rocket is successfully lifted off the platform, the rocket continuously flies and lands according to the designed time sequence, and the thrust of the main engine directly acts on the rocket; the flight time sequence written in the control software of the rocket is designed according to the following method:

the flight time is designed in sections, and the axial flight overload of each corresponding altitude section is matched, so that the axial flight overload of each corresponding altitude section is suitable for the verification flight of the vertical recovery verification rocket, and the specific design method comprises the following steps:

the thrust of the engine is adjusted according to the flight time sequence, so that the staggered change of the axial flight overload of the rocket body is realized, the proper axial flight overload is provided for the vertical recovery demonstration and verification of the whole flight process of the rocket, and the vertical controlled landing recovery can be finally realized;

the flight time is divided into a takeoff ascending section, a first deceleration ascending and then acceleration descending section and a deceleration descending section, the takeoff ascending section corresponds to a flight height h1, the first deceleration ascending and then acceleration descending section corresponds to a flight height h2, the deceleration descending section corresponds to a flight height h3, the thrust of the engine is adjusted in each stage, the takeoff ascending section, the first deceleration ascending and then acceleration descending section and the deceleration descending section are determined according to the iterative calculation of thrust-quality (referring to the quality of the rocket) and height-overload, the thrust is adjusted in each stage, namely the thrust is adjusted by adjusting the combustion quality of the propellant, because the propellant in the rocket is consumed in the flying process, the change of the mass of the propellant can cause the change of the quality of the rocket, the adjustment of the thrust of the engine in each stage is determined according to the iterative calculation of the thrust-quality-height-overload, the staggered change of the axial flying overload of the rocket body is realized; the heights h1 and h2 are obtained by iterative operation by adopting the following conventional formula, wherein the thrust-mass-height-overload is as follows:

A_p＝M_FG2FSM_JT2FG[P/M；0；0]p is thrust, M is rocket body mass at solving time (rocket body mass changes due to propellant mass change), and M is_JT2FGFor converting an arrow coordinate system into a matrix of an emission coordinate system, A_pThe acceleration generated by the thrust under the emission coordinate system. AK gfs + Acof + Acf + A_p+ AN, AK is the sum acceleration that the arrow body receives under the emission coordinate system, gfs is the gravitational acceleration, Acof is the involved acceleration, Acf is the Coriolis acceleration, AN is the acceleration that aerodynamic force produces under the emission coordinate system, N ═ A_p+A_NN is an overload, V_fs(i+1)＝V_fs(i)+t_stepA_k(i +1) is the calculation period, V_fs(i) For calculating the speed of the cycle, V_fs(i +1) is the speed of the present tamper period; r_fs(i+1)＝R_fs(i)+t_stepV_fs(i+1)，R_fs(i) For the position of the last calculation cycle, R_fs(i +1) is the position of the calculation cycle; h (i +1) ═ M_FS2DX(R_fs(i+1)+R_0fs)|-R_0dx(i +1), H (i +1) is the height of the calculation period (i.e. corresponding to H1 and H2 values), R_0fsAs the position of the emission point in the emission coordinate system, M_FS2DXFor transforming the emitting coordinate system into a ground-centered coordinate system matrix, R_0dx(i +1) is the distance from the corresponding surface location to the geocentric.

However, the height of h3 judged by final landing is related to the guidance strategy and navigation error precision of the arrow body and the landing speed born by the landing leg, and finally the result of h3 value is obtained by comprehensive analysis and is related to the index level of hardware, such as the design index of the bearing capacity of the landing buffer mechanism.

Further, the specific principle of realizing the staggered change of the rocket body axial flying overload is as follows:

after the rocket is successfully lifted off the platform, the thrust of a main engine is not adjusted in the rising section of the takeoff of the rocket, the axial flying overload of the rocket body is not too large or too small, the rocket body continuously flies according to the preset overload n1, the value of n1 is 1.1g, the rising height of the rocket is too high if the value is too large, the thrust adjusting amplitude of the engine is limited, the rocket cannot be recovered, and the risk of too slow lifting exists if the value is too small; in another embodiment n1 is 1.3g and in yet another embodiment n1 is 1.2 g.

The flying height h1 of the rocket is continuously judged in the flying process, when h1 is larger than a program preset value, the thrust of a main engine is reduced, the flying overload of the rocket body is kept smaller than n2 through thrust adjustment, the rocket body is not too large or too small for the descending section of the rocket which is decelerated and ascended first and then accelerated, but is smaller than the ascending section of the rocket, the value of n2 is 0.6g, the value of n2 is not too large, if the value of n2 is too large, the rocket needs to be decelerated for a longer time, the mass of the propellant is consumed to be increased, because the lower limit of the thrust adjustment of the engine is limited, the gravity of the rocket is smaller than the lower limit of the thrust adjustment of the engine finally, the aircraft cannot descend, and if the value of h1 is too small, the thrust adjustment range of the engine is limited and cannot be realized; in another embodiment n2 is 0.9g and in yet another embodiment n2 is 0.75 g.

The flying rocket is in a deceleration ascending state firstly, flies in an acceleration descending state after the top point of the ballistic trajectory flight (the top point can be judged by the rocket body navigation information), the flying process continuously judges the flying height h2 of the rocket (the height h2 point is the thrust adjusting point in the falling process), when h2 is less than the program preset value, the flying is judged to return to the deceleration starting point, the thrust of the main engine is increased, through thrust adjustment, the flying overload of the rocket body is kept to be larger than n3, n3 is 1.1g, at the moment, the rocket body is not too large or too small for the deceleration descending section of the rocket, however, the value n3 is 1.1g, if the value is too large, the rocket decelerates faster, the descent speed returns to zero quickly, the aircraft ascends in a reverse acceleration mode, and finally cannot land, if the value is too small, the deceleration performance is weak, the speed of the aircraft cannot be reduced to 0 when the altitude returns to zero, and the requirement of the landing speed cannot be met. And while keeping the flight overload, the thrust and the vector propulsion direction of the engine are adjusted in a small amplitude, the landing speed and the landing attitude of the arrow body are controlled, and the landing speed precision and the landing attitude precision are ensured to meet the design index requirements of a landing system. And when the rocket body decelerates and descends, continuously judging the flying height h3 of the rocket, and when h3 is smaller than a program preset value, closing the main engine of the rocket, vertically landing and recovering the rocket body, and finishing vertical take-off and landing demonstration and verification flight. In another embodiment n3 is 1.4g and in yet another embodiment n3 is 1.25 g.

As shown in fig. 7, the T1-T7 are calculated from the ignition T0 at predetermined time intervals, and further, the specific timing steps are designed as follows:

step 1, after an ignition signal is sent out by a demonstration and verification rocket main engine at the time of T0, carrying out demonstration and verification rocket body departure judgment at the set time of T1 at the interval of preset time T1 and T1 of 3s, and sending a continuous working or emergency shutdown signal to the demonstration and verification rocket main engine; in another embodiment t1 takes the value 5s, and in yet another embodiment t1 takes the value 4 s.

Step 11, if the arrow body is judged not to normally leave the platform at the time T1, the interval preset time T2 is set, T2 is 5s, the main engine executes an emergency shutdown program at the time T2, and the interval preset time T1 is smaller than the interval preset time T2; in another embodiment t2 takes the value 7s, and in yet another embodiment t2 takes the value 6 s.

Step 12, if the rocket body is judged to be normally off the platform, the main engine continues to normally work, the rocket body is kept to accelerate and ascend, the flying height is judged at the T3 time within the preset time T3, T3 is 30s, if the flying height does not reach the preset height h1, the thrust of the engine is kept unchanged, the rocket continuously flies according to the condition that the overload is not less than the preset overload n1, n1 is 1.3g, if the preset height h1 is reached, the thrust is reduced at the T4 interval within the preset time T4 for 35s, and the main engine enables the demonstration and verification rocket to decelerate and ascend at the T4 time; in another embodiment t3 takes the value 32.5s, and in yet another embodiment t3 takes the value 35 s. In another embodiment t4 takes the value 40s, and in yet another embodiment t4 takes the value 37.5 s. In another embodiment n1 is 1.2g and in yet another embodiment n1 is 1.1 g.

Step 2, after the demonstration and verification rocket reaches a ballistic vertex, the rocket body starts to accelerate and descend, the rocket overload is kept smaller than n2, n2 is 0.9g, the preset time T5 is spaced, T5 is 75s, the altitude is judged at the time T5, if the preset altitude h2 is not reached, the thrust of the main engine is kept unchanged, the current overload is continued to fly, if the preset altitude h2 is reached, the preset time T6 is spaced, T6 is 80s, the thrust of the main engine is increased at the time T6, and the demonstration and verification rocket is made to decelerate and descend; when the height is judged to be smaller than the program preset value h2, the thrust of the main engine is increased, and the rocket body flying overload is kept to be larger than n3 through thrust adjustment; in another embodiment t5 takes the value 80s, and in yet another embodiment t5 takes the value 77.5 s. In another embodiment t6 takes the value 85s, and in yet another embodiment t6 takes the value 82.5 s. In another embodiment n2 is 0.75g and in yet another embodiment n2 is 0.6 g.

And 3, after the moment T6, the demonstration and verification rocket enters a vertical landing final guide section, the flying overload is kept to be larger than n3, n3 is 1.4g, the main engine adjusts the thrust magnitude and the vector direction in a small amplitude mode, the landing speed and the landing attitude of the rocket body are controlled, the landing speed precision and the landing attitude precision meet the design index requirements of a landing system, the rocket body descends in a deceleration mode, the preset time T7 is set at intervals, T7 is 100s, the flying height judgment is carried out at the moment T7, when the flying height is smaller than the program preset value h3, the main engine of the demonstration and verification rocket is closed, the rocket body is recovered in vertical landing, and the vertical take-off and landing demonstration and verification flying are completed. In another embodiment t7 takes the value 120s, and in yet another embodiment t7 takes the value 110 s. In another embodiment n3 is 1.25g and in yet another embodiment n3 is 1.1 g.

The specific maximum height of the rocket in the embodiment is 1000m, wherein the value of h1 is 350-450m, the specific value is 400m, the value of h2 is 250-350m, the specific value is 300m, the value of h3 is 0-1m, and the specific value is 0.5 m. In another possible embodiment, the rocket flying height may be any height greater than 0m, not limited to 1000m, and may be 10000m, or 100000m, 1000000m, and may be larger as required.

Specifically, as shown in fig. 8 to 12, a predetermined lift-off detection program is written in the rocket control software of the present invention to determine whether the rocket body is lifted off, and a lift-off detection method for demonstrating and verifying the rocket by using reusable technology is provided, wherein a buffer is provided on a rocket recovery leg, and the pressure value borne by the buffer can be measured, and the method includes:

s1, judging according to the pressure value of the rocket recovery leg buffer;

s2, judging according to the axial overload of the arrow body;

s3, judging according to the measured actual height of the lower end surface of the rocket body supporting leg from the ground;

s4, recording the time difference dt between the current time and the ignition time as T_t-T₀Wherein T is_tIs the current time, T₀The moment of sending an ignition instruction of the engine;

s5, when two conditions of the buffer pressure condition, the rocket body axial overload condition and the ground clearance height condition of the lower end face of the supporting leg judged in the S1-3 are simultaneously met, judging that the rocket is out of the stage, recording the current time as T1, and downloading the rocket through remote measurement;

when the three criteria are not satisfied but dt is less than or equal to Tth, continuously measuring the physical quantity, and when the three criteria are not satisfied and dt is greater than Tth, determining that the rocket is out of stage and performing a post-processing procedure; tth is the rocket destaging time calculated according to the maximum design deviation and the ballistic design condition, and then a margin of 20 percent is added;

s6, when the following events occur, the determination is stopped:

(a) the rocket takes off and downloads telemetering data of taking off zero seconds;

(b) when the rocket fails, the rocket receives an emergency shutdown remote control instruction sent by the ground.

Further, the step S1 is specifically: and measuring the pressure Pmd1-4 in a nitrogen chamber or an oil cylinder of four buffers of the recovery leg, wherein the sampling period is 50-1000 kHz. After the engine finishes ignition timing sequence control, the four counters Cmd1-4 start to count, the initial values of Cmd1-4 are all 0, when the pressure measurement value is smaller than the binding threshold value Pth, the corresponding counter is increased by one, and when the pressure measurement value is larger than or equal to the binding threshold value Pth, the counter is cleared by 0; when three terms in Cmd1-4 are greater than Kmd at the same time, it is determined that the buffer pressure condition is satisfied, and Kmd is generally 20-50, for example, 20, 25, 30, 35, 40, 45, 50, etc.

Further, the step S2 is specifically: measuring axial overload of the arrow body, namely, the arrow body overload Nx1 in the x direction under the arrow body coordinate system, wherein the sampling period is 50-100Hz,N_x1,kfor the k-th calculation cycle, the axial overload of the arrow body is generally 5-10, for example 5, 6,8, 10. Etc.; after the ignition time sequence control of the engine is completed, the counter Cnx starts counting, the initial value of Cnx is 0, when the overload measured value is larger than the binding threshold value Nth, the counter Cnx is increased by one, and when the overload measured value is smaller than or equal to the binding threshold value Nth, the counter is cleared by 0; when the Cnx value is greater than Knx, it is determined that the rocket body axial overload condition is met, and Knx generally takes a value of 5-15, for example, 5, 7, 10,12, 15, etc.

Further, the step S3 is specifically: the actual ground clearance Hgd1-2 of the lower end surface of the arrow body supporting leg is respectively measured by two height meters,wherein H_tx,kThe sampling period is 10-100Hz, and n is generally 5-10, such as 5, 7, 9,10 and the like. After the engine finishes ignition timing control, two counters Cgd1-2 start to count, the initial value Cgd1-2 is 0, when Hgd1 or Hgd2 is greater than the binding threshold Hth, the corresponding counter is increased by one, and when Hgd1 or Hgd2 is less than or equal to the binding threshold Hth, the corresponding counter is cleared by 0; when one value of Cgd1-2 is greater than Kgd, it is determined that the ground clearance condition of the lower end face of the supporting leg is satisfied, and Kgd generally takes a value of 5-15, for example, 5, 7, 10,12, 15, etc.

Further, the judgment sequence is that the pressure condition of the buffer is judged firstly; and judging the axial overload condition of the rocket body, and finally judging the ground clearance condition of the lower end surface of the rocket supporting leg.

Further, Nth is an axial overload discrimination threshold, generally takes a value of 1.05 to 1.3 times of gravity acceleration, and a specific value is related to a takeoff thrust-weight ratio. .

Specifically, in this embodiment, the determination is performed according to the pressure value measured by the buffer pressure sensor: the rocket is supported on the ground of the launching area through the recovery supporting legs, due to the action of the gravity of the rocket, the buffer in the recovery supporting legs is in a compressed state, the pressure of a nitrogen chamber and an oil cylinder in the buffer is higher, after the rocket engine is ignited, the thrust is gradually built, the compression amount of the buffer is reduced, the pressure of the nitrogen chamber and the oil cylinder in the buffer is reduced, when the rocket takes off, the buffer is not compressed and is in a free state, and the pressure is smaller than a threshold value;

when the measured pressure is smaller than the binding threshold value, the establishment of the thrust is completed, and the compression amount generated by the supporting leg for providing the supporting force for the arrow body is reduced; when the rocket is lifted off the platform, the supporting legs are in a loose free state, and the buffer is almost not compressed until the supporting legs touch the ground when landing;

judging according to arrow overload: when the thrust of the rocket is greater than the self gravity, the rocket leaves the launching platform to take off;

and (3) measuring height through an altimeter to judge: after the rocket leaves the platform, the ground clearance measured by the altimeter is rapidly increased; until the arrow ground returns to the ground, the ground clearance is reduced again;

judging the starting time and the ending time: after the ignition timing sequence control of the engine is finished, starting to judge; and when the rocket is out of the stage or fails, stopping the judgment of the out-of-stage.

The present embodiment has the following criteria:

in S1, Cmd1-4 indicates the times that the measured value of the pressure of the buffer corresponding to the label is continuously smaller than the binding threshold value Pth, Kmd is the binding value and indicates the times meeting the pressure condition of the buffer, Kmd generally takes a value of 20-50 and is related to the design value of the buffer, and the specific numerical value is obtained through simulation or experiment; the threshold value Pth is related to the design value of the buffer, the weight of the rocket body and the takeoff overload, and the specific value is determined through mathematical simulation and experiments.

When three items in Cmd1-4 are greater than Kmd at the same time, judging that the buffer pressure condition is met; due to the deviation of the machining process and the fact that the rocket body has a certain inclination angle after filling, a plurality of modes of separating the supporting legs from the ground exist in the takeoff process, and the condition of fault redundancy of the pressure sensor is considered, so that the criterion is set as 4 and 3.

In S2Representing the average value of axial overload in the latest continuous n periods; n is a radical of_x1,kThe arrow body axial overload of the k calculation period; n is generally 5-10; cnx, representing the number of times that the overload measured value is continuously greater than the binding threshold value Nth, Knx is the binding value, representing the number of times that the arrow body axial overload condition is met, and Knx generally takes the value of 5-15; nth is an axial overload judgment threshold, generally takes a value of 1.05-1.3 times of gravity acceleration, and a specific value is related to a takeoff thrust-weight ratio.

In S3Representing the average value of the height of the antenna installation position of the altimeter from the ground in the latest continuous n periods; h_tx,kThe ground clearance of the antenna installation position of the altimeter in the kth calculation period is calculated; h is_azThe axial distance from the mounting position of the altimeter antenna to the lower end face of the supporting leg is generally 5-10; cgd1-2 represents the number of times that the measured height value of the corresponding height gauge is continuously larger than the binding threshold value Hth, Kgd is the binding value and represents the number of times that the condition of the height of the lower end face of the supporting leg from the ground is met, Kgd generally takes a value of 5-15, Hth is related to the measurement precision of the height gauge, and generally Hth is 1-1.5 times larger than the measurement precision of the height gauge.

When one Cgd1-2 is larger than Kgd, judging that the ground clearance height condition of the lower end face of the supporting leg is met; due to the fact that the posture of the arrow body is inclined when the arrow body leaves the platform, the measured values of the two altimeters are deviated, and redundant judgment when the altimeters are in failure is considered.

Specifically, the judgment sequence of the embodiment is to judge the pressure first; judging overload and finally judging height; in practice, these three measurements are taken at the same time, but theoretically the pressure changes first; when the compression amount of the supporting legs is gradually reduced, because the lower end surfaces of the supporting legs still do not leave the ground, the rocket is balanced with the thrust and the compression force of the supporting legs under the supporting force of the ground, at the moment, the overload is not obviously changed, and the overload can be obviously changed only when the rocket body is not balanced by the supporting force of the ground after the rocket body leaves the platform; the height is integrated as acceleration and finally changes.

Judging a destaging fault: and when the three items are not satisfied and dt is larger than Tth, judging that the rocket is out of stage and performing a post-processing procedure.

The invention also provides a sub-orbital sounding rocket, which is characterized in that a fairing and a grid rudder are arranged on the reusable rocket for verifying the vertical take-off and landing technology, so that the rocket can adapt to higher flying speed. Particularly, a fairing is arranged at the head of the rocket, so that the aerodynamic shape of the rocket is changed, and the rocket can adapt to flight with higher Mach number. Meanwhile, a grid rudder is additionally arranged in quadrants I, II, III and IV of the tail part of the rocket, is folded when the rocket ascends and unfolded when the rocket returns, and controls the posture of the rocket together with the swinging and spraying of the engine so as to land stably.

The invention also provides a method for verifying the vertical take-off and landing technology, which is performed according to the reusable rocket for verifying the vertical take-off and landing technology, and comprises the following three experimental contents:

the reusable rocket performs static ignition work;

the reusable rocket performs mooring protection ignition work;

the reusable rocket performs vertical flight.

Specifically, the specific contents of the reusable rocket for static ignition work are as follows: carrying out a ground static ignition test, fixing the rocket on a launching tower, and igniting the engine; and (3) carrying out circulating precooling on the engine by using the low-temperature oxygen tank and the methane tank, simultaneously starting the blowing system in the cabin, receiving the transmission data of a single machine by using ground related equipment at the moment, confirming whether the blowing system works normally, and verifying the calculation method of the blowing system. The working time of the blowing system can verify the precooling flow of the engine and the correctness of precooling program setting under a low-temperature environment according to a formula; after the oxygen tank and the methane tank are pressurized in a self-generating pressurization mode, the propellant is conveyed to the engine, then ignition is carried out, the servo mechanism is used for controlling the engine to swing and spray, the control system is used for sending an instruction to adjust the thrust of the engine, whether the engine finishes thrust adjustment and swing and spray is judged according to related data sent back by the single machine in the control system, and the low-temperature engine thrust adjustment technology can be verified.

Specifically, the reusable rocket for mooring protection ignition specifically comprises the following contents: the rocket is lifted to be 5m away from the ground by using the crane, the support legs are not influenced by the back rolling of the tail flame of the engine due to the height of 5m, the engine is in a whole-course starting ignition state, the rocket has the ascending height of about 5m, the limit height of a suspender of the crane can be exceeded due to the overhigh rocket, the loading capacity of the crane is influenced due to the overhigh suspender, and the cost of renting the crane is improved. The engine thrust is larger than the rocket gravity at the beginning, the rocket ascends for about 5m, then the engine thrust is adjusted, the rocket gravity is larger than the engine thrust, the rocket starts to fall back after being subjected to deceleration and ascending, at the moment, the weight of the rocket is reduced due to fuel consumption, the engine thrust is larger than the rocket gravity again, when the rocket decelerates and lands to the takeoff height, the engine is shut down, and the rocket is lifted by a crane. Because the operating mode rocket has low flying speed and larger interference of random wind to the rocket, an auxiliary power system is additionally added at the moment and is matched with the engine for swinging and spraying to control the stability of the rocket body. And verifying the high-precision navigation technology and the high-precision guidance technology according to the feedback data of each single machine in the control system and whether the rocket stably returns to the position before takeoff.

Specifically, the reusable rocket specifically comprises the following contents of vertical flight: the rocket is placed on the ground, so that the rocket is subjected to a flight test in a free state, the height H of the top point of a trajectory is about 1000m, the height is selected to reduce research and development cost, part of pipelines of a power system are placed on the periphery of the rocket body, meanwhile, a fairing of the rocket body is omitted, and the engine is not shut down in the whole process after being started. The thrust of the engine is greater than the gravity of the rocket, so that the rocket is accelerated to rise. And then adjusting the thrust of the engine to be smaller than the gravity of the rocket, and accelerating the rocket to descend after the rocket ascends to the top point of the trajectory in a deceleration way. The single-start vertical rocket trajectory calculation method is adopted to calculate the flight time sequence of the rocket, and the flight time sequence is bound into a single machine (control software) of a control system in advance, when the rocket reaches a specified calculation height, the thrust of an engine is adjusted to be larger than the gravity of the rocket, and the rocket decelerates and descends. When the rocket is near the ground, the engine is shut down, the landing support legs are used for buffering, and meanwhile, the heat-proof layer on the surface can protect the support legs, so that the support legs can be repeatedly used, and finally the rocket is landed. The whole-course engine action is controlled by a control system, and an instruction is sent out at the preset time of a program after the rocket position and speed are calculated. The rocket successfully flies back according to the binding trajectory, and the correctness of the vertical rocket trajectory calculation method for single startup is also verified. The rocket returns after flying to 1000m, the control system measures the position, speed and other parameters of the rocket in the period, the rocket is stably ascended and returned by utilizing the engine swing jet and the auxiliary power system jet, and the high-precision navigation and high-precision guidance technology is verified. The engine is started in a low-temperature state, and liquid oxymethane in the storage tank provides power for the engine, so that the correctness of the calculation method of the blowing system is verified. When the rocket is close to the ground, the engine is closed, the supporting legs are used for buffering landing, the rocket successfully lands and finishes recovery, and the landing supporting leg buffering technology can be verified. The rocket successfully ascends and descends, the matching of interfaces of all systems is checked, the rationality of all key indexes is evaluated, and the overall design technology is verified.