A rocket offload control method based on multi-source sensors

Abstract

This invention provides a rocket load reduction control method based on multi-source sensors, relating to the aerospace field. The method includes: considering multiple rocket parameters, designing a rocket attitude controller, and establishing a rocket load reduction control mission model; obtaining the rocket's normal acceleration using a load reduction accelerometer; the rocket's normal acceleration includes the normal acceleration caused by aerodynamic forces and the normal acceleration caused by engine thrust; during rocket flight, measuring the chamber pressure in real time using an engine chamber pressure sensor to determine the real-time engine thrust; determining the normal acceleration caused by aerodynamic forces based on the rocket's normal acceleration and engine thrust; incorporating feedback of the normal acceleration corresponding to the aerodynamic force into the rocket attitude controller; and determining the reduced total airflow angle of attack based on the rocket load reduction control mission model to achieve rocket load reduction control. This invention achieves precise load reduction control of the rocket during the ascent phase.

Description

Technical Field

[0001] This invention relates to the aerospace field, and in particular to a rocket load reduction control method based on multi-source sensors. Background Technology

[0002] During the ascent phase of a rocket, when passing through the high dynamic pressure region, wind disturbances can cause uncertain changes in the rocket's angle of attack and sideslip angle, resulting in normal aerodynamic loads exceeding the maximum load that the rocket structure can withstand, leading to deformation or even fracture of the rocket structure. Therefore, while ensuring guidance accuracy, reducing the aerodynamic loads caused by wind disturbances is an important task of guidance and control during the rocket's ascent phase.

[0003] Existing rocket load reduction control methods can be broadly categorized into two types: passive load reduction based on predicted wind fields and active load reduction based on accelerometer feedback. Passive load reduction requires wind field forecasting before launch and adjusts the trajectory attitude angle command accordingly to reduce normal aerodynamic overload. This method relies heavily on the accuracy of wind field forecasting, cannot account for rapidly changing wind shear, and exhibits strong conservatism in trajectory correction. Active load reduction, on the other hand, incorporates normal apparent acceleration feedback in the attitude controller based on accelerometer measurements. It adjusts the rocket's attitude in real-time when passing through high dynamic pressure zones to reduce normal aerodynamic overload. However, the normal acceleration information sensitively measured by accelerometers includes not only acceleration caused by normal aerodynamic forces but also acceleration caused by engine normal thrust, making precise load reduction control impossible. Summary of the Invention

[0004] The purpose of this invention is to provide a rocket load reduction control method based on multi-source sensors to solve the problem of inaccurate load reduction control.

[0005] To achieve the above objectives, the present invention provides the following solution:

[0006] A rocket load reduction control method based on multi-source sensors includes:

[0007] Considering multiple rocket parameters, a rocket attitude controller is designed, and a rocket load reduction control mission model is established; the rocket parameters include aerodynamic forces along the fuselage normal, pitch damping moment, engine thrust, and disturbance moment;

[0008] The rocket's normal acceleration is obtained by using a load-reduction accelerometer; the rocket's normal acceleration includes the normal acceleration caused by aerodynamic forces and the normal acceleration caused by engine thrust.

[0009] During rocket flight, the engine chamber pressure is measured in real time using an engine chamber pressure sensor to determine the real-time engine thrust of the rocket.

[0010] The normal acceleration caused by the aerodynamic force is determined based on the rocket's normal acceleration and the engine thrust.

[0011] By incorporating the normal acceleration feedback corresponding to the normal acceleration caused by the aerodynamic force into the rocket attitude controller, and based on the rocket load reduction control mission model, the reduced total airflow angle of attack is determined to achieve rocket load reduction control.

[0012] Optionally, considering multiple rocket parameters, an attitude controller is designed, and a rocket load reduction control mission model is established, specifically including:

[0013] Considering multiple rocket parameters, establish the rocket's longitudinal attitude motion equation;

[0014] Based on the aforementioned rocket longitudinal attitude motion equation, determine the linear time-varying approximate equation for rocket longitudinal motion;

[0015] The rocket attitude controller is designed based on the linear time-varying approximate equation of rocket longitudinal motion, the total airflow angle of attack of the rocket is derived, and a rocket load reduction control mission model is established.

[0016] Optionally, the rocket's longitudinal attitude motion equations for:

[0017]

[0018] in, J is the pitch angle; z F is the moment of inertia of the pitch channel; AY The aerodynamic force along the normal direction of the aircraft body; x A M is the distance from the rocket's center of mass to its aerodynamic center of pressure. DZ For pitch damping moment; F P For engine thrust; x P This is the distance from the rocket's center of mass to the point of action of the engine. M is the engine pitch angle; BZ This refers to the pitch disturbance torque caused by installation deviation.

[0019] Optionally, the approximate equation for the linearly time-varying longitudinal motion of the rocket is:

[0020]

[0021] in, This is the actual pitch angle acceleration; The first coefficient; This refers to the actual pitch rate; The second coefficient; Δα is the deviation between the actual angle of attack and the nominal ballistic angle of attack; It is the third coefficient; The deviation between the actual engine sway angle and the trim sway angle; M BZb α is the pitch direction disturbance angular acceleration; w The angle of attack is caused by the wind field.

[0022] Optionally, the rocket normal acceleration for:

[0023]

[0024] Where m is the rocket mass; x W This is the distance from the rocket's center of mass to the location where the accelerometer is installed.

[0025] Optionally, during rocket flight, the engine chamber pressure is measured in real time using an engine chamber pressure sensor to determine the real-time engine thrust of the rocket, specifically including:

[0026] Before rocket launch, an engine test is conducted to generate an internal ballistic thrust curve; the internal ballistic thrust curve is the curve of nominal engine thrust changing over time.

[0027] During rocket flight, the chamber pressure is measured in real time using an engine chamber pressure sensor. A second-order Butterworth filter is designed to perform low-pass filtering on the real-time measured chamber pressure values.

[0028] Based on the internal ballistic thrust curve, an approximate formula for calculating engine thrust is determined according to engine characteristics.

[0029] The measured chamber pressure after low-pass filtering is substituted into the approximate engine thrust calculation formula to determine the real-time engine thrust of the rocket.

[0030] Optionally, the reduced total airflow angle of attack is:

[0031]

[0032] in, Nominal angle of attack; K P K is the pitch angle feedback coefficient; g This is the normal acceleration feedback coefficient; It is the fourth coefficient.

[0033] According to specific embodiments provided by the present invention, the present invention discloses the following technical effects: The present invention provides a rocket load reduction control method based on multi-source sensors, which makes full use of multi-source information from accelerometers and chamber pressure sensors for rocket load reduction, calculates engine thrust in real time to calculate the normal acceleration caused by aerodynamic forces in the rocket's normal acceleration, and then adds the normal acceleration feedback corresponding to the normal acceleration caused by aerodynamic forces to the rocket attitude controller, thereby subtracting the normal acceleration caused by aerodynamic forces, realizing precise load reduction control of the rocket in the ascent phase, and improving the reliability of rocket load reduction control. Attached Figure Description

[0034] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0035] Figure 1 This is a flowchart of the rocket load reduction control method based on multi-source sensors provided by the present invention;

[0036] Figure 2 This is a schematic diagram of the rocket load reduction control method based on multi-source sensors provided by the present invention;

[0037] Figure 3 This is a schematic diagram of the filtering result of the chamber pressure sensor measurement value provided by the present invention;

[0038] Figure 4 This is a schematic diagram of the engine thrust estimation results provided by the present invention;

[0039] Figure 5 This is a schematic diagram illustrating the rocket load reduction effect provided by the present invention. Detailed Implementation

[0040] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0041] The purpose of this invention is to provide a rocket load reduction control method based on multi-source sensors, which can realize precise load reduction control of rockets during the ascent phase and improve the reliability of rocket load reduction control.

[0042] To make the above-mentioned objects, features and advantages of the present invention more apparent and understandable, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0043] like Figure 1 As shown, this invention provides a rocket load reduction control method based on multi-source sensors, including:

[0044] Step 101: Considering multiple rocket parameters, design a rocket attitude controller and establish a rocket load reduction control mission model; the rocket parameters include aerodynamic forces along the normal direction of the rocket body, pitch damping torque, engine thrust, and disturbance torque.

[0045] In practical applications, based on the rocket's longitudinal attitude motion equations, taking the classic PD (proportional-derivative) attitude controller as an example, considering aerodynamic forces, pitch damping torque, engine thrust, and disturbance torque, the total airflow angle of attack is derived, resulting in the rocket load reduction control mission model.

[0046] rocket longitudinal attitude motion equations It can be represented as:

[0047]

[0048] in, J is the pitch angle; z F is the moment of inertia of the pitch channel; AY The aerodynamic force along the normal direction of the aircraft body; x A M is the distance from the rocket's center of mass to its aerodynamic center of pressure. DZ For pitch damping moment; F P For engine thrust; x P This is the distance from the rocket's center of mass to the point of action of the engine. M is the engine pitch angle; BZ This refers to the pitch direction interference torque caused by installation deviations, etc.

[0049] Aerodynamic force F along the normal direction of the body AY It can be represented as:

[0050]

[0051] Where α is the angle of attack of the ballistic trajectory; α w The angle of attack is caused by the wind field. q is the partial derivative of the normal aerodynamic coefficient with respect to the angle of attack; q is the dynamic pressure; S M This is the reference area for the rocket.

[0052] Pitch damping moment M DZ It can be represented as:

[0053]

[0054] Where V is the rocket velocity; The pitch rate; is the partial derivative of the aerodynamic damping coefficient with respect to the pitch rate; l is the characteristic length of the rocket.

[0055] Considering the short-period motion of the rocket, expanding equation (1) near the nominal trajectory yields...

[0056]

[0057]

[0058] in, This is the actual pitch angle acceleration; The actual pitch angle Compared with the nominal ballistic elevation angle Deviation between; The actual pitch rate; Δα = The actual angle of attack α and the nominal ballistic angle of attack Deviation between; Actual engine sway angle With the flat angle Deviation between; coefficient With M BZb The expression is as follows:

[0059]

[0060]

[0061]

[0062]

[0063] Among them, J Z M is the moment of inertia of the pitch channel; BZb This is the normal acceleration feedback coefficient.

[0064] The task of the rocket attitude control system is to ensure the actual pitch angle Tracking nominal ballistic elevation angle Even if To keep the engine sway angle near zero, taking a classic PD controller as an example, the control law for the engine sway angle is designed as follows:

[0065]

[0066] Among them, K P K is the pitch angle feedback coefficient; D This is the pitch rate feedback coefficient.

[0067] When the system is stable Equations (4) and (10) can be expressed as:

[0068]

[0069]

[0070] Further results were obtained:

[0071]

[0072] The steady-state deviation of the angle of attack can then be obtained as:

[0073]

[0074] At this point, the total angle of attack is:

[0075]

[0076] Formula (15) is the established rocket load reduction control mission model. The mission of rocket load reduction control is to reduce the normal aerodynamic load by reducing the total airflow angle of attack when the rocket passes through the high dynamic pressure zone, so as to keep the normal aerodynamic load within the range that the rocket structure can withstand.

[0077] Step 102: Use the unloaded accelerometer to obtain the rocket's normal acceleration; the rocket's normal acceleration includes the normal acceleration caused by aerodynamic forces and the normal acceleration caused by engine thrust.

[0078] In practical applications, step 102 is the design of a load reduction control method based on accelerometers.

[0079] The measurement equation for the rocket normal accelerometer is:

[0080]

[0081] in, The normal acceleration is measured by the accelerometer; m is the rocket mass; x W This is the distance from the rocket's center of mass to the location where the accelerometer is installed.

[0082] This can be further expressed as:

[0083]

[0084] coefficients in the formula and The expression is as follows:

[0085]

[0086]

[0087] The accelerometer-based load reduction control method incorporates normal acceleration feedback into the attitude controller, namely:

[0088]

[0089] Among them, K g This is the normal acceleration feedback coefficient.

[0090] When the system is stable Equations (4) and (20) can be expressed as:

[0091]

[0092]

[0093] Further results were obtained:

[0094]

[0095] At this point, the total angle of attack of the airflow is:

[0096]

[0097] In equation (24), it is noted that since the aerodynamic pressure center is always closer to the rocket's center of mass than the engine's point of action, i.e., x P <0 and x A >x P ,thereby It always holds true.

[0098] Comparing equation (24) with equation (15), as long as K g If the normal acceleration is greater than 0, then adding normal acceleration feedback to the attitude controller can reduce the total airflow angle of attack. However, since the engine thrust is unknown, the normal acceleration information obtained from the sensitive accelerometer measurement... In addition to the normal aerodynamic force F AY The resulting acceleration also includes the engine's normal thrust. The resulting acceleration leads to an increase in engine thrust F P This will affect the steady-state value of the total airflow angle of attack, meaning that precise load reduction control cannot be achieved. Therefore, this invention determines the real-time engine thrust of the rocket through step 103 to determine the normal acceleration caused by aerodynamic forces.

[0099] Step 103: During rocket flight, the engine chamber pressure is measured in real time using an engine chamber pressure sensor to determine the real-time engine thrust of the rocket.

[0100] In practical applications, step 103 is the design of a thrust calculation method based on a chamber pressure sensor.

[0101] Before rocket launch, an engine test can be conducted to obtain a nominal engine thrust curve, known as the internal ballistic thrust curve. During rocket flight, the actual engine thrust will deviate from the nominal thrust. The actual thrust can be calculated by measuring the combustion chamber pressure using a chamber pressure sensor in the engine combustion chamber.

[0102] The formula for calculating engine thrust can be expressed as:

[0103] F P =η n P c A t C Fth (25)

[0104] Where, η n For nozzle efficiency; P c A is the engine combustion chamber pressure. t C is the cross-sectional area of ​​the nozzle throat. Fth This is the theoretical thrust coefficient.

[0105] Due to the measured value of the chamber pressure sensor There is noise, and according to the above thrust calculation formula, P... c With thrust F P The relationship between them is a non-linear function, therefore we should first consider... Perform signal filtering.

[0106] Design a second-order Butterworth filter with the optimal damping ratio ξ = 0.707, and its transfer function G(s) is shown below:

[0107]

[0108] Where s is a complex variable in the Laplace transform.

[0109] In the above filter, the design parameters a0 and a1 are:

[0110]

[0111]

[0112] Where, ω a This is the cutoff angular frequency.

[0113] The following design incorporates a second-order Butterworth digital filter. Let the sampling frequency of the digital filter be f. s The desired cutoff frequency is f c Then the cutoff angular frequency of the analog filter is:

[0114]

[0115] Discretize using a bilinear transform, let:

[0116]

[0117] Where z is a complex variable in the z-transform.

[0118] Substituting the above equation into equation (26), we get:

[0119]

[0120] Where G(z) is the transfer function of the discrete system; B0 = Hc 2 ;B1=2B0;B2=B0;A0=1;A1=H(2c 2 -2);

[0121] The formulas for calculating c and H are as follows:

[0122]

[0123]

[0124] Among them, B0, B1, B2, A0, A1, A2, c, and H have no actual physical meaning; they are only used to simplify the algorithm representation. After passing through a second-order Butterworth filter, the low-pass filtered room pressure measurement value is obtained as follows:

[0125]

[0126] in, This is the measured value from the chamber pressure sensor.

[0127] For solid rocket motors, the cross-sectional area of ​​the nozzle throat gradually increases with ablation, which can be fitted based on ground test data as follows:

[0128]

[0129] Among them, D t0 D is the diameter of the nozzle throat. t1 t represents the ablation rate; t represents the combustion time.

[0130] The formula for calculating the theoretical thrust coefficient is:

[0131]

[0132]

[0133] Where Γ has no specific physical meaning and is only used for simplification, k is the specific heat ratio; A e P is the nozzle exit cross-sectional area; a The ambient atmospheric static pressure; Pe This is the nozzle outlet pressure.

[0134] The nozzle exit cross-sectional area does not change over time; the calculation formula is as follows:

[0135]

[0136] Among them, D e This is the nozzle exit diameter.

[0137] The nozzle exit pressure can be calculated using the following formula:

[0138]

[0139] The final calculated value of the engine thrust is:

[0140] F P =η n P cm A t C Fth (40)

[0141] Step 104: Determine the normal acceleration caused by the aerodynamic force based on the rocket's normal acceleration and the engine thrust.

[0142] Step 105: Add the normal acceleration feedback corresponding to the normal acceleration caused by the aerodynamic force to the rocket attitude controller, and determine the reduced total airflow angle of attack based on the rocket load reduction control mission model to realize rocket load reduction control.

[0143] In practical applications, the rocket load reduction control method based on accelerometers and chamber pressure sensors is designed as follows:

[0144] First, considering the error in the engine thrust formula, a Kalman filter is designed to estimate the deviation between the actual thrust and the nominal thrust by combining the internal ballistic thrust curve and the calculated engine thrust value, thereby obtaining the thrust estimate.

[0145] The actual thrust of an engine can be expressed as:

[0146]

[0147] in, The nominal engine thrust for internal ballistics; ΔF P This represents the thrust deviation between the actual thrust and the nominal thrust.

[0148] Assuming thrust deviation ΔF P The dynamic changes are relatively slow, that is:

[0149]

[0150] in, is the derivative of the thrust deviation; w is the error value of the derivative of the thrust deviation.

[0151] The Kalman filter is designed as follows:

[0152]

[0153] P(0)=P0(44)

[0154] K = PR -1 (45)

[0155]

[0156]

[0157] in, For thrust deviation ΔF P The initial estimate; P(0) is the initial value of the autocovariance of the estimation error; For thrust deviation ΔF P The estimated value; Q is the autocovariance of w; R is the value obtained using P cm Measuring engine thrust F P The autocovariance of the error; P0 is ΔF P The initial autocovariance of the estimation error; P is ΔF P The autocovariance of the estimation error; K is the Kalman optimal feedback gain.

[0158] Therefore, the real-time estimated value of engine thrust can be obtained using the engine chamber pressure sensor:

[0159]

[0160] Then, based on the thrust estimate, the normal acceleration caused by the thrust is subtracted from the normal acceleration sensitive to the accelerometer to obtain the normal acceleration caused by aerodynamic force:

[0161]

[0162] Finally, combining the real-time thrust estimate and the normal acceleration sensitive by the accelerometer, the load shedding control is designed as follows:

[0163]

[0164] get:

[0165]

[0166] When the system is stable Equations (4) and (51) can be expressed as:

[0167]

[0168]

[0169] in, This represents the estimation error of engine thrust.

[0170] Further results were obtained:

[0171]

[0172] At this point, the total angle of attack of the airflow is:

[0173]

[0174] If the engine thrust estimation error The above equation then becomes:

[0175]

[0176] Compared with equation (24), this method significantly reduces the impact of... The resulting total airflow angle of attack eliminates the inaccurate load reduction caused by the engine thrust component included in the accelerometer, thus achieving precise load reduction control. Its framework diagram is shown below. Figure 2 As shown.

[0177] This invention has good engineering feasibility, high credibility, and strong reliability. The rocket load reduction control method based on multi-source sensors provided by this invention is scientific, has good manufacturability, and has broad application value.

[0178] Specific simulation examples are as follows:

[0179] This section will demonstrate the methodology using a digital simulation case, not an actual flight mission. The simulation scenario is a flight mission during the ascent phase of a certain type of rocket. The simulated chamber pressure sensor readings are based on the actual chamber pressure with the addition of random white noise. The noise amplitude is 0.2 MPa and the noise frequency is 50 Hz. The Butterworth digital filter has a sampling frequency of 100 Hz and a filter cutoff frequency of 1 Hz.

[0180] Using the load reduction control method of this invention, the filtered results of the room pressure sensor measurements were obtained through digital simulation, as shown below. Figure 3 As shown, the engine thrust estimation results are as follows: Figure 4 As shown, the rocket's load reduction effect is as follows: Figure 5 As shown.

[0181] The various embodiments in this specification are described in a progressive manner, with each embodiment focusing on the differences from other embodiments. The same or similar parts between the various embodiments can be referred to each other.

[0182] This document uses specific examples to illustrate the principles and implementation methods of the present invention. The descriptions of the above embodiments are only for the purpose of helping to understand the method and core ideas of the present invention. Furthermore, those skilled in the art will recognize that, based on the ideas of the present invention, there will be changes in the specific implementation methods and application scope. Therefore, the content of this specification should not be construed as a limitation of the present invention.

Claims

1. A rocket load reduction control method based on multi-source sensors, characterized in that, include: Considering multiple rocket parameters, design a rocket attitude controller and establish a rocket load reduction control mission model, specifically including: Considering multiple rocket parameters, establish the rocket's longitudinal attitude motion equations; the rocket's longitudinal attitude motion equations for: ； in, The pitch angle; The moment of inertia of the pitch channel; The aerodynamic force along the normal direction of the aircraft body; This is the distance from the rocket's center of mass to its aerodynamic center of pressure. This is the pitch damping moment; For engine thrust; This is the distance from the rocket's center of mass to the point of action of the engine. This refers to the engine pitch angle; This refers to the pitch disturbance torque caused by installation deviation; Based on the aforementioned rocket longitudinal attitude motion equations, a linearly time-varying approximate equation for the rocket's longitudinal motion is determined; the linearly time-varying approximate equation for the rocket's longitudinal motion is as follows: in, This is the actual pitch angle acceleration; The first coefficient; This refers to the actual pitch rate; The second coefficient; This represents the deviation between the actual angle of attack and the nominal ballistic angle of attack. It is the third coefficient; This is the deviation between the actual engine sway angle and the trim sway angle; The pitch direction interference angular acceleration; The angle of attack is caused by the wind field. The rocket attitude controller is designed based on the linear time-varying approximate equation of rocket longitudinal motion, the total airflow angle of attack of the rocket is derived, and a rocket load reduction control mission model is established; the rocket parameters include aerodynamic forces along the normal direction of the fuselage, pitch damping moment, engine thrust, and disturbance moment; The rocket's normal acceleration is obtained using a deceleration accelerometer; this normal acceleration includes both the normal acceleration caused by aerodynamic forces and the normal acceleration caused by engine thrust; the rocket's normal acceleration... for: ； in, For rocket mass; This is the distance from the rocket's center of mass to the location where the accelerometer is installed; During rocket flight, the engine chamber pressure is measured in real time using an engine chamber pressure sensor to determine the real-time engine thrust of the rocket. The normal acceleration caused by the aerodynamic force is determined based on the rocket's normal acceleration and the engine thrust. By incorporating the normal acceleration feedback corresponding to the normal acceleration caused by the aerodynamic force into the rocket attitude controller, and determining the reduced total airflow angle of attack based on the rocket load reduction control mission model, rocket load reduction control is achieved; the reduced total airflow angle of attack is: ； in, This refers to the nominal angle of attack. This is the pitch angle feedback coefficient; This is the normal acceleration feedback coefficient; It is the fourth coefficient.

2. The rocket load reduction control method based on multi-source sensors according to claim 1, characterized in that, During rocket flight, engine chamber pressure sensors are used to measure chamber pressure in real time to determine the rocket's real-time engine thrust, specifically including: Before rocket launch, an engine test is conducted to generate an internal ballistic thrust curve; the internal ballistic thrust curve is the curve of nominal engine thrust changing over time. During rocket flight, the chamber pressure is measured in real time using an engine chamber pressure sensor. A second-order Butterworth filter is designed to perform low-pass filtering on the real-time measured chamber pressure values. Based on the internal ballistic thrust curve, an approximate formula for calculating engine thrust is determined according to engine characteristics. The measured chamber pressure after low-pass filtering is substituted into the approximate engine thrust calculation formula to determine the real-time engine thrust of the rocket.

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