A landing point control method for a reusable launch vehicle

Abstract

The present invention relates to a landing point control method for a reusable carrier rocket, including pre-landing power-on control, pre-landing height control, pre-landing horizontal position control, and pre-landing power-off control. The pre-landing power-on control is based on an iterative calculation of the landing point coordinate system. If the calculation reaches vyfs i If it is not greater than zero, the next iteration is performed; when vyfs is calculated i When it is greater than zero, judge the current yfs i Is it less than zero? If so, the engine is started; otherwise, the engine is not started; the iteration ends; the height control before landing is calculated through iterative calculation to obtain vyfs i Greater than zero or yfs i When it is less than zero, judge vyfs i Is it less than v mx If yes, the iteration ends; if not, the target apparent acceleration is updated; based on the target apparent acceleration W ax,j+1 Adjust the engine thrust and then proceed to the next iterative calculation. This method has a small amount of control instruction calculation, high guidance loop stability, and high engineering application value.

Description

Technical Field

[0001] The present invention belongs to the technical field of rocket recovery, and in particular relates to a landing point control method for a recoverable carrier rocket. Background Art

[0002] With the rapid development of commercial spaceflight, reducing launch costs for launch vehicles is a key technological development direction for rocket companies. Generally, rocket engines account for the majority of a rocket's cost, with the primary engine core comprising the largest portion. Recovering the primary stage is currently the primary technology for rocket recovery both domestically and internationally, and commercial rocket companies primarily utilize this approach.

[0003] After the first stage of the launch vehicle separates, it undergoes a deceleration flight phase outside the atmosphere, an aerodynamic deceleration flight phase inside the atmosphere, and a pre-landing deceleration flight phase. The deceleration flight phase outside the atmosphere mainly uses the flight position and speed information to ignite the engine to decelerate the flight so that the theoretical landing point meets the landing point requirements; the aerodynamic deceleration flight phase inside the atmosphere mainly uses aerodynamic drag to decelerate the flight and aerodynamic force to control the flight attitude. Generally, grid rudders or aerodynamic drag plates are used for attitude control to ensure that the first stage flies according to the standard ballistic trajectory; the pre-landing deceleration flight phase mainly uses the position and speed information during the flight to carry out timely position control and engine ignition control, correct the speed and position errors of the aerodynamic deceleration flight phase inside the atmosphere, and ensure that the first stage can achieve precise landing point control according to the theoretical landing point requirements.

[0004] During the deceleration flight phase before landing, the rocket is difficult to control in order to meet the landing requirements due to the limited engine thrust adjustment capability and adjustment range, and the influence of various air flow interferences in the atmosphere during flight. In particular, the position and speed control before landing is the control difficulty of the deceleration flight phase before landing. Summary of the Invention

[0005] In response to the above-mentioned problem of difficulty in controlling the deceleration flight phase before landing of the rocket, the present invention proposes a landing point control method for a recoverable carrier rocket. In this method, the control instruction calculation amount is small, the guidance loop is highly stable, and it has high engineering application value.

[0006] To achieve the above purpose, the specific technical solution adopted by the present invention is:

[0007] A method for controlling the landing point of a reusable carrier rocket, comprising:

[0008] Pre-landing power-on control: Establish the landing point coordinate system, perform iterative calculation based on the landing point coordinate system, and if the calculation reaches vyfs i If it is not greater than zero, the next iteration is performed; when vyfs is calculated i When it is greater than zero, judge the current yfs iIs it less than zero? If so, the engine is started; otherwise, the engine is not started; the iteration ends;

[0009] vyfs i is the speed of the rocket in the y-axis direction of the landing point coordinate system, yfs i is the height of the rocket in the y-axis direction of the landing point coordinate system; the y-axis of the landing point coordinate system is perpendicular to the launch direction, and the positive value is pointing to the sky;

[0010] Height control before landing: perform iterative calculation to obtain vyfs i Greater than zero or yfs i When it is less than zero, judge |vyfs i Is it less than v? mx If yes, the iteration ends; if not, the target apparent acceleration is updated as follows:

[0011]

[0012] W ax,j+1 is the apparent acceleration of the rocket in the y-axis direction of the landing point coordinate system, v mx is the threshold value for speed judgment, k ax is the adjustment coefficient of apparent acceleration;

[0013] Based on the target apparent acceleration W ax,j+1 Adjust the engine thrust and proceed to the next iteration.

[0014] Shutdown control before landing: judge according to the recursive result of height control before landing, set T delay The engine shutdown delay, when the iteration is completed, if the current time T i Less than T delay , the engine shuts down, otherwise continue with pre-landing altitude control.

[0015] Furthermore, the control method also includes pre-landing horizontal position control, and the pre-landing horizontal position control is controlled according to proportional guidance.

[0016] Furthermore, the speed judgment threshold v mx Take 2-3m / s, the adjustment coefficient k of the acceleration ax Take 0.5%-2%.

[0017] Furthermore, in the pre-landing startup control, performing iterative calculation based on the landing point coordinate system includes:

[0018] Assign initial value T i =0, yfs i =yfs0,vyfs i =vyfs0;

[0019] T i is the flight time in the iterative process, τ is the intermediate variable, Isp is the theoretical specific impulse of the engine, W ax is the apparent acceleration in the y-axis direction of the coordinate system of the rocket's landing point at the current moment; yfs0 is the initial value of the velocity iteration, and vyfs0 is the initial value of the height iteration;

[0020] Perform iterative calculation according to the following formula:

[0021]

[0022] yfs i+1 =yfs i +T i ×(vyfs i +0.5×DV i +0.5×T i ×g0)

[0023] vyfs i+1 =vyfs i +DV i +T i ×g0

[0024] If vyfs i is not greater than zero, then let T i+1 =T i +ΔT, proceed to the next iteration;

[0025] Among them, ΔT is the recursive step size, τ is the intermediate variable, and g0 is the acceleration of gravity.

[0026] Furthermore, in the pre-landing altitude control, the iterative calculation includes:

[0027] Assign initial value T i =0, yfs i =yfs0,vyfs i =vyfs0;

[0028] T i is the flight time in the iterative process, τ is the intermediate variable, Isp is the theoretical specific impulse of the engine, W ax is the apparent acceleration in the y-axis direction of the coordinate system of the rocket's landing point at the current moment; yfs0 is the initial value of the velocity iteration, and vyfs0 is the initial value of the height iteration;

[0029] Perform iterative calculation according to the following formula:

[0030]

[0031] yfs i+1 =yfs i +Ti ×(vyfs i +0.5×DV i +0.5×T i ×g0)

[0032] vyfs i+1 =vyfs i +DV i +T i ×g0

[0033] Determine whether vyfs is satisfied i Greater than zero or yfs i Less than zero, if not satisfied, let T i+1 =T i +ΔT, proceed to the next iteration; if satisfied, then judge |vyfs i Is it less than v? mx ;

[0034] Among them, ΔT is the recursive step size, τ is the intermediate variable, and g0 is the acceleration of gravity.

[0035] Furthermore, the recursive step length ΔT is 500ms-1s.

[0036] Furthermore, the target apparent acceleration W ax,j+1 Adjusting engine thrust includes:

[0037] The target apparent acceleration W ax,j+1 Compared with the current rocket's measured apparent acceleration W ax Compare and calculate the engine thrust adjustment percentage command K f

[0038]

[0039] Where k dh is the proportionality coefficient, K f0 The current engine thrust percentage command.

[0040] Furthermore, the proportional coefficient k dh Take 0.6-0.8.

[0041] Furthermore, the controlling according to proportional guidance includes:

[0042] Calculate the pitch acceleration command a y and yaw acceleration command a z

[0043]

[0044]

[0045] Where, is the angular velocity of the line of sight in the pitch direction between the rocket and the theoretical landing point at the current moment, is the yaw direction line of sight angular velocity between the rocket and the theoretical landing point at the current moment, q y is the pitch angle between the rocket and the theoretical landing point at the current moment; k q is the proportional guidance coefficient;

[0046] Calculate pitch attitude angle command and yaw attitude angle command ψ cx

[0047]

[0048]

[0049] In the above formula, n x It is the filtered value of the axial acceleration at the center of mass of the rocket measured and output by the rocket navigation system, and g0 is the acceleration due to gravity.

[0050] Furthermore, the cycle of the iterative calculation is 1-2 seconds, that is, one round of iterative calculation is performed every 1-2 seconds.

[0051] Compared with the prior art, the present invention has the following beneficial effects:

[0052] The present invention provides a landing point control method for a recoverable carrier rocket. Iterative calculations are performed based on position and velocity information during flight to accurately control engine ignition and thrust, thereby achieving precise landing of the rocket. This method requires relatively little control instruction calculation, and the guidance loop is highly stable, thus having high engineering application value. BRIEF DESCRIPTION OF THE DRAWINGS

[0053] Figure 1 This is a flow chart of the pre-landing startup control of the present invention;

[0054] Figure 2 This is a flow chart of altitude control before landing of the present invention. DETAILED DESCRIPTION

[0055] The present invention will now be further described with reference to the accompanying drawings.

[0056] The rocket's pre-landing deceleration phase primarily relies on timely position control and engine ignition control based on position and velocity information during flight, correcting for velocity and position errors during the aerodynamic deceleration phase within the atmosphere. The engine described in this invention is a landing reverse thrust engine, which generates upward thrust for the rocket, reducing landing velocity and achieving a soft landing.

[0057] A method for controlling the landing point of a reusable carrier rocket according to the present invention includes power-on control before landing, altitude control before landing, horizontal position control before landing, and power-off control before landing.

[0058] When the rocket's flight altitude reaches the theoretical ballistic ignition altitude, the pre-landing power-on control begins, that is, the on-board computer begins to perform iterative calculations.

[0059] The specific start-up controls before landing are as follows:

[0060] Establish a landing point coordinate system with the origin at the landing point planned before launch. The coordinate axes are parallel to the coordinate axes of the launch coordinate system. The x-axis is parallel to the launch direction and is positive when pointing in the launch direction. The y-axis is perpendicular to the launch direction and is positive when pointing toward the sky. The z-axis, x-axis, and y-axis meet the requirements of the Cartesian coordinate system.

[0061] Perform iterative calculation based on the landing point coordinate system, such as Figure 1 As shown,

[0062] First, assign the initial value T i =0, vyfs i =vyfs0;

[0063] T i is the flight time in the iterative process, the initial value is 0, τ is the intermediate variable, Isp is the theoretical specific impulse of the engine, W ax vyfs is the apparent acceleration in the y-axis direction of the coordinate system of the rocket's current landing point. Apparent acceleration refers to the acceleration caused by external forces other than gravity. In the present invention, the apparent acceleration refers to the acceleration caused by the engine thrust. i is the speed of the rocket in the y-axis direction of the landing point coordinate system, yfs i is the height of the rocket in the y-axis direction of the landing point coordinate system; yfs0 is the iterative initial value of the rocket speed, and vyfs0 is the iterative initial value of the rocket height;

[0064] Then perform iterative calculation as follows:

[0065]

[0066] yfs i+1 =yfs i +T i ×(vyfs i +0.5×DV i +0.5×T i ×g0)

[0067] vyfs i+1 =vyfs i +DV i +T i ×g0

[0068] Among them, ΔT is the recursive step size, τ is the intermediate variable, and g0 is the acceleration of gravity.

[0069] The recursive step size ΔT can generally be selected from 500ms to 1s, and the appropriate step size can be selected based on the computer capabilities on the arrow.

[0070] If calculated to vyfs i is not greater than zero, then let T i+1 =T i +ΔT, proceed to the next iteration; when vyfs is calculated i When it is greater than zero, judge the current yfs i Is it less than zero? If so, the engine is started; otherwise, the engine is not started; end this iteration, wait 1-2 seconds and then start the next cycle of iterative calculation; the specific interval period of iterative calculation can be optimized according to the computing power of the computer on the arrow.

[0071] Pre-landing altitude control: Controlled by iterative calculation, such as Figure 2 As shown,

[0072] Similarly, first assign the initial value T i =0, yfs i =yfs0,vyfs i =vyfs0;

[0073] Then perform iterative calculation as follows:

[0074]

[0075] yfs i+1 =yfs i +T i ×(vyfs i +0.5×DV i +0.5×T i ×g0)

[0076] vyfs i+1 =vyfs i +DV i +T i ×g0

[0077] When vyfs is calculated i Greater than zero or yfs i When it is less than zero, judge at this time|vyfs i Is it less than v? mx If yes, the iteration ends; if not, the target apparent acceleration is updated as follows:

[0078]

[0079] W ax,j+1 is the apparent acceleration of the rocket in the y-axis direction of the landing point coordinate system, v mx k is the threshold value for speed judgment, which is generally 2-3 m / s. The specific value can be determined based on simulation optimization; ax is the adjustment coefficient of apparent acceleration, which is generally 0.5%-2%. The specific value can be determined based on simulation optimization;

[0080] Based on the target apparent acceleration W ax,j+1 Adjust engine thrust, specifically:

[0081] The target apparent acceleration W ax,j+1 Compared with the current rocket's measured apparent acceleration W ax Compare and calculate the engine thrust adjustment percentage command K f

[0082]

[0083] In the above formula, K f0 The current engine thrust adjustment percentage instruction;

[0084] In order to reduce the fluctuation of the engine thrust adjustment command, the present invention improves the calculation formula:

[0085]

[0086] Where k dh is the proportional coefficient, which can generally be taken as 0.6-0.8. The specific value can be determined based on simulation optimization.

[0087] According to K f Adjust the engine thrust and proceed to the next iteration.

[0088] Horizontal position control before landing:

[0089] The horizontal position control before landing is carried out according to the proportional guidance commonly used in aircraft control, as follows:

[0090] First, obtain the pitch acceleration command a according to the following formula y and yaw acceleration command a z :

[0091]

[0092]

[0093] In the above formula, is the angular velocity of the line of sight in the pitch direction between the rocket and the theoretical landing point at the current moment, is the yaw direction line of sight angular velocity between the first stage and the theoretical landing point at the current moment, qy is the pitch angle between the rocket and the theoretical landing point at the current moment, q z k is the yaw angle between the rocket and the theoretical landing point at the current moment; q is the proportional guidance coefficient, which is generally 3-6. The specific value is determined based on simulation optimization.

[0094] Then obtain the pitch attitude angle command according to the following formula and yaw attitude angle command ψ cx :

[0095]

[0096]

[0097] In the above formula, n x It is the filtered value of the axial acceleration at the center of mass measured and output by the rocket navigation system.

[0098] Shutdown control before landing: judge according to the recursive result of height control before landing, set T delay The engine shutdown delay, when the iteration is completed, if the current time T i Less than T delay , the engine shuts down, otherwise it continues to perform pre-landing altitude control. delay It is a performance parameter determined through engine testing.

Claims

1. A method for controlling the landing point of a recoverable carrier rocket, characterized in that: include Pre-landing power-on control: Establish the landing point coordinate system, perform iterative calculation based on the landing point coordinate system, and if the calculation reaches vyfs i If it is not greater than zero, the next iteration is performed; when vyfs is calculated i When it is greater than zero, judge the current yfs i Is it less than zero? If so, the engine is started; otherwise, the engine is not started; the iteration ends; vyfs i is the speed of the rocket in the y-axis direction of the landing point coordinate system, yfs i is the height of the rocket in the y-axis direction of the landing point coordinate system; the y-axis of the landing point coordinate system is perpendicular to the rocket launch direction, and the positive value is pointing to the sky; Height control before landing: perform iterative calculation to obtain vyfs i Greater than zero or yfs i When it is less than zero, judge |vyfs i Is it less than v? mx If yes, the iteration ends; if not, the target apparent acceleration is updated as follows: W ax,j+1 is the apparent acceleration of the rocket in the y-axis direction of the landing point coordinate system, v mx is the threshold value for speed judgment, k ax is the adjustment coefficient of apparent acceleration; Based on the target apparent acceleration W ax,j+1 Adjust the engine thrust and then proceed to the next iterative calculation; Horizontal position control before landing; Shutdown control before landing: judge according to the recursive result of height control before landing, set T delay The engine shutdown delay, when the iteration is completed, if the current time T i Less than T delay , the engine shuts down, otherwise continue with pre-landing altitude control.

2. The method for controlling the landing point of a recoverable carrier rocket according to claim 1, wherein: The horizontal position control before landing is performed according to proportional guidance.

3. The method for controlling the landing point of a recoverable carrier rocket according to claim 1, wherein: The speed judgment threshold v mx Take 2-3m / s, the adjustment coefficient k of the acceleration ax Take 0.5%-2%.

4. The method for controlling the landing point of a recoverable carrier rocket according to claim 1, wherein: In the pre-landing startup control, performing iterative calculation based on the landing point coordinate system includes: Assign initial value T i =0, yfs i =yfs0,yfs i =yfs0; T i is the flight time in the iterative process, τ is the intermediate variable, Isp is the theoretical specific impulse of the engine, W ax is the apparent acceleration in the y-axis direction of the coordinate system of the rocket's landing point at the current moment; yfs0 is the initial value of the velocity iteration, and vyfs0 is the initial value of the height iteration; Perform iterative calculation according to the following formula: yfs i+1 =yfs i+Ti ×(vyfs i+0.5×DVi+0.5×Ti×g0 ) vyfs i+1 =vyfs i+DVi+Ti×g0 If vyfs i is not greater than zero, then let T i+1 =T i +ΔT, proceed to the next iteration; Among them, ΔT is the recursive step size, τ is the intermediate variable, and g0 is the acceleration of gravity.

5. The method for controlling the landing point of a recoverable carrier rocket according to claim 1, wherein: In the pre-landing altitude control, the iterative calculation includes: Assign initial value T i =0, yfs i =yfs0,vyfs i =vyfs0; T i is the flight time in the iterative process, τ is the intermediate variable, Isp is the theoretical specific impulse of the engine, W ax is the apparent acceleration in the y-axis direction of the coordinate system of the rocket's landing point at the current moment; yfs0 is the initial value of the velocity iteration, and vyfs0 is the initial value of the height iteration; Perform iterative calculation according to the following formula: yfs i+1 =yfs i+Ti ×(vyfs i +0.5×DV i +0.5×T i ×g0) vyfs i+1 =vyfs i +DV i+Ti×g0 Determine whether vyfs is satisfied i Greater than zero or yfs i Less than zero, if not satisfied, let T i+1 =T i +ΔT, proceed to the next iteration; if satisfied, then judge |vyfs i Is it less than v? mx ; Among them, ΔT is the recursive step size, τ is the intermediate variable, and g0 is the acceleration of gravity.

6. A method for controlling the landing point of a recoverable carrier rocket according to claim 4 or 5, characterized in that: The recursive step length ΔT is 500ms-1s.

7. The method for controlling the landing point of a recoverable carrier rocket according to claim 1, wherein: The target apparent acceleration W ax,j+1 Adjusting engine thrust includes: The target apparent acceleration W ax,j+1 Compared with the current rocket's measured apparent acceleration W ax Compare and calculate the engine thrust adjustment percentage command K f Where k dh is the proportionality coefficient, K f0 The current engine thrust percentage command.

8. The method for controlling the landing point of a recoverable carrier rocket according to claim 7, wherein: The proportionality coefficient k dh Take 0.6-0.

8.

9. The method for controlling the landing point of a recoverable carrier rocket according to claim 2, wherein: The control according to proportional guidance includes: Calculate the pitch acceleration command a y and yaw acceleration command a z Where, is the angular velocity of the line of sight in the pitch direction between the rocket and the theoretical landing point at the current moment, is the yaw direction line of sight angular velocity between the rocket and the theoretical landing point at the current moment, q y is the pitch angle between the rocket and the theoretical landing point at the current moment; k q is the proportional guidance coefficient; Calculate pitch attitude angle command and yaw attitude angle command In the above formula, n x It is the filtered value of the axial acceleration at the center of mass of the rocket measured and output by the rocket navigation system, and g0 is the acceleration due to gravity.

10. The landing point control method of a recoverable carrier rocket according to claim 1, characterized in that: The cycle of the iterative calculation is 1-2 seconds.

Images