Carrier rocket control method and device, medium and program product

Abstract

The invention discloses a carrier rocket control method which comprises the following steps: acquiring actual output signals of rate gyroscopes corresponding to rocket levels on a carrier rocket at a target moment, and acquiring preset weights corresponding to the rate gyroscopes of the carrier rocket at the target moment; determining the overall elastic vibration angular rate of the carrier rocket at the target moment according to the actual output signals and the preset weights respectively corresponding to the rate gyroscopes at the target moment; and controlling the attitude of the carrier rocket at the target moment according to the overall elastic vibration angular rate of the carrier rocket at the target moment. The problems that in the prior art, when elastic vibration is controlled by reducing the slenderness ratio of a carrier rocket or adjusting materials and structures, attitude control flexibility is poor, and the dead weight is too large are solved, in addition, the complexity of elastic vibration control is greatly simplified, and the implementation cost of elastic vibration control is greatly reduced.

Description

Technical Field

[0001] The present invention belongs to the field of launch vehicle control, and specifically, belongs to a launch vehicle vibration control method, equipment, medium and program product. Background Art

[0002] As a key component of the aerospace system, launch vehicles will experience elastic vibrations during flight due to factors such as the fuel they carry, mechanical structure, and aerodynamics, which will cause the flight attitude error of the launch vehicle to gradually increase and the rocket structure to be damaged. This problem is more likely to occur in launch vehicles with a large aspect ratio. Currently, the elastic vibration control of launch vehicles is mainly improved by changing the aspect ratio and materials, but these methods are costly to implement. Therefore, how to improve the elastic vibration control effect of launch vehicles at a lower cost is an urgent problem to be solved. Summary of the invention

[0003] The embodiment of the present application provides a method for controlling a launch vehicle, thereby solving the technical problem in the prior art of improving the elastic vibration control of a launch vehicle mainly by changing the aspect ratio and materials, etc., which leads to high control costs. The method achieves the determination of the overall elastic vibration angular rate through a rate gyro and preset weights corresponding to different moments, thereby controlling the attitude of the launch vehicle at the target moment at a low cost.

[0004] In a first aspect, the present application provides a launch vehicle control method, which is applied during the flight of the launch vehicle, and the method comprises:

[0005] Obtaining actual output signals of rate gyros corresponding to respective rocket stages on the launch vehicle at a target time, and obtaining preset weights corresponding to respective rate gyros of the launch vehicle at the target time;

[0006] Determine the overall elastic vibration angular rate of the launch vehicle at the target time according to the actual output signals respectively corresponding to the rate gyros at the target time and the preset weights;

[0007] The attitude of the carrier rocket at the target moment is controlled according to the overall elastic vibration angular rate of the carrier rocket at the target moment.

[0008] Optionally, before the launch vehicle flies, the method further includes:

[0009] Predicting a first rocket mass corresponding to the target time during the flight of the carrier rocket;

[0010] Determining the target mode slope corresponding to the first rocket mass according to the corresponding relationship between the rocket mass and the mode slope of the carrier rocket;

[0011] According to the target vibration mode slopes corresponding to the respective rate gyros at the target time, the corresponding preset weights are determined respectively.

[0012] Optionally, the target vibration mode slopes corresponding to the respective rate gyros at the target moments are determined by ground tests and / or modeling simulations of the launch vehicle.

[0013] Optionally, the method further includes:

[0014] Determine the preset weights of X moments; X is a positive integer;

[0015] The preset weights at the X moments are interpolated to obtain a continuous correlation relationship between each moment and the preset weight.

[0016] Optionally, the sum of the preset weights of the launch vehicle corresponding to the continuous association relationship is 1.

[0017] Optionally, controlling the attitude of the carrier rocket at the target moment according to the overall elastic vibration angular rate of the carrier rocket at the target moment includes:

[0018] Performing filtering processing on the overall elastic vibration angular frequency of the launch vehicle at the target time to filter out the elastic vibration existing in the overall elastic vibration angular rate of the launch vehicle at the target time;

[0019] The attitude of the launch vehicle at the target moment is controlled according to the overall elastic vibration angular frequency after filtering.

[0020] Optionally, the rate gyro is installed at an adjacent antinode in the vibration field generated by the elastic vibration of the launch vehicle.

[0021] In a second aspect, the present application provides an electronic device, including:

[0022] processor;

[0023] a memory for storing instructions executable by the processor;

[0024] Wherein, the processor is configured to execute to implement a launch vehicle control method as described in any one of the first aspects.

[0025] In a third aspect, the present application provides a non-temporary computer-readable storage medium. When the instructions in the storage medium are executed by a processor of an electronic device, the electronic device is enabled to execute a launch vehicle control method as described in the first aspect.

[0026] In a fourth aspect, the present application provides a computer program product, characterized in that it includes computer instructions, which are executed by a processor to implement a launch vehicle control method as described in any one of the first aspects.

[0027] One or more technical solutions provided in the embodiments of the present application have at least the following technical effects or advantages:

[0028] The embodiment of the present application sets a rate gyro on each rocket stage of the launch vehicle, and then during the flight of the launch vehicle, collects the actual output signal of the rate gyro at any target time, and determines the preset weight corresponding to the target time. Based on the corresponding actual output signal and the preset weight, the overall elastic vibration angular velocity of the launch vehicle at the target time can be obtained, and then based on the overall elastic vibration angular velocity, the launch vehicle is controlled. It can be seen that the embodiment of the present application abandons the method of improving the elastic vibration control of the launch vehicle by changing the aspect ratio and materials in the related art, but instead improves the elastic vibration control of the launch vehicle by arranging the rate gyro, relying on the actual output signal collected by the rate gyro at the target time and the corresponding preset weight. On the basis of being able to realize the elastic vibration control of the launch vehicle, the complexity of the elastic vibration control is greatly simplified, and the implementation cost of the elastic vibration control is greatly reduced. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the accompanying drawings required for use in the description of the embodiments will be briefly introduced below. Obviously, the accompanying drawings in the following description are some embodiments of the present invention. For ordinary technicians in this field, other accompanying drawings can be obtained based on these accompanying drawings without paying creative work.

[0030] Figure 1 A flow chart of a launch vehicle control method disclosed in an embodiment of the present invention;

[0031] Figure 2 A schematic diagram of the structure of an elastic vibration control system for a launch vehicle disclosed in an embodiment of the present invention;

[0032] Figure 3 A schematic diagram of an electronic device for a launch vehicle elastic vibration control system disclosed in an embodiment of the present invention;

[0033] Figure 4 A first-order vibration mode distribution diagram of the entire rocket of a first-order launch vehicle elastic vibration control system disclosed in an embodiment of the present invention;

[0034] Figure 5 The present invention discloses a second-order vibration mode distribution diagram of the entire rocket of a second-order launch vehicle elastic vibration control system. DETAILED DESCRIPTION

[0035] The embodiment of the present application provides a method for controlling a launch vehicle, thereby solving the technical problem in the prior art of improving the elastic vibration control of a launch vehicle mainly by changing the aspect ratio and materials, resulting in high control costs.

[0036] The technical solution of the embodiment of the present application is to solve the above technical problems, and the overall idea is as follows:

[0037] A launch vehicle control method is applied to the flight process of the launch vehicle, and the method comprises:

[0038] Obtaining actual output signals of rate gyros corresponding to respective rocket stages on the launch vehicle at a target time, and obtaining preset weights corresponding to respective rate gyros of the launch vehicle at the target time;

[0039] Determine the overall elastic vibration angular rate of the launch vehicle at the target time according to the output signals respectively corresponding to the rate gyros at the target time and the preset weights;

[0040] The attitude of the carrier rocket at the target moment is controlled according to the overall elastic vibration angular rate of the carrier rocket at the target moment.

[0041] The embodiment of the present application sets a rate gyro on each rocket stage of the launch vehicle, and then during the flight of the launch vehicle, collects the actual output signal of the rate gyro at any target time, and determines the preset weight corresponding to the target time. Based on the corresponding actual output signal and the preset weight, the overall elastic vibration angular velocity of the launch vehicle at the target time can be obtained, and then based on the overall elastic vibration angular velocity, the launch vehicle is controlled. It can be seen that the embodiment of the present application abandons the method of improving the elastic vibration control of the launch vehicle by changing the aspect ratio and materials in the related art, but instead improves the elastic vibration control of the launch vehicle by arranging the rate gyro, relying on the actual output signal collected by the rate gyro at the target time and the corresponding preset weight. On the basis of being able to realize the elastic vibration control of the launch vehicle, the complexity of the elastic vibration control is greatly simplified, and the implementation cost of the elastic vibration control is greatly reduced.

[0042] In order to better understand the above technical solution, the above technical solution will be described in detail below in conjunction with the accompanying drawings and specific implementation methods.

[0043] First of all, the term "and / or" in this article is only a description of the association relationship of the associated objects, indicating that there can be three relationships. For example, A and / or B can mean: A exists alone, A and B exist at the same time, and B exists alone. In addition, the character " / " in this article generally indicates that the associated objects before and after are in an "or" relationship.

[0044] As a key component of the aerospace system, launch vehicles will experience elastic vibrations during flight due to factors such as the fuel they carry, mechanical structure, and aerodynamics, which can cause the flight attitude error of the launch vehicle to gradually increase and the rocket structure to be damaged. This problem is especially more likely to occur in launch vehicles with a large aspect ratio. There are currently two common methods for elastic vibration control of launch vehicles. One method is to adjust the aspect ratio of the launch vehicle, mainly by reducing the aspect ratio parameter. At this time, even if a certain degree of elastic vibration occurs, the launch vehicle has a better resistance to elastic vibration due to its large diameter. Another method is to replace the rocket body material with a higher strength, and to set up reinforcement structures and vibration absorption structures for some structural weak points to avoid structural damage caused by elastic vibration.

[0045] However, the inventors of this application found that the above technology has at least the following technical problems in the process of implementing the technical solution of the invention in the embodiment of this application:

[0046] 1. When adjusting the aspect ratio, the aerodynamic shape of the launch vehicle becomes unsuitable for attitude adjustment. Overall, this will lead to a decrease in the sensitivity of the launch vehicle's attitude control system. At the same time, the attitude control system also needs to be adjusted and optimized to a large extent. While the cost of control system development increases, it is still difficult to achieve higher attitude control sensitivity.

[0047] 2. The solution of replacing high-strength rocket body materials and adding other strengthening or vibration-absorbing structures will lead to a significant increase in the rocket body's own weight, and the payload that the launch vehicle can carry will inevitably decrease.

[0048] In order to solve the above problems, the present application provides a launch vehicle control method, which is applied to the flight process of the launch vehicle. Figure 1 As shown, it specifically includes steps S110 to S130.

[0049] Step S110, obtaining the actual output signal of the rate gyro corresponding to each rocket stage on the carrier rocket at the target time, and obtaining the preset weight corresponding to each rate gyro of the carrier rocket at the target time;

[0050] Step S120, determining the overall elastic vibration angular rate of the launch vehicle at the target time according to the output signals respectively corresponding to the rate gyros at the target time and the preset weights;

[0051] Step S130: Controlling the attitude of the carrier rocket at the target moment according to the overall elastic vibration angular rate of the carrier rocket at the target moment.

[0052] Regarding step S110, the actual output signal of the rate gyro corresponding to each rocket stage on the carrier rocket at the target time is obtained, and the preset weight corresponding to each rate gyro of the carrier rocket at the target time is obtained.

[0053] The launch vehicle includes a plurality of rocket stages, and the embodiment of the present application pre-arranges a corresponding rate gyro on each rocket stage of the launch vehicle. For example, the rate gyro is installed at adjacent antinodes in the vibration field generated by the elastic vibration of the launch vehicle.

[0054] During the flight of the launch vehicle, the actual output signal of the rate gyro is collected according to a preset frequency. The preset frequency can be set according to actual needs.

[0055] The target time can be any time during the flight of the launch vehicle. It should be noted that the flight process of the launch vehicle is a continuous process, and each time corresponding to the preset frequency in the continuous process is sequentially used as the target time to execute steps S110 to S130. The embodiment of the present application only takes one of the target times as an example for subsequent description.

[0056] The preset weights corresponding to different moments may be different, and the preset weights corresponding to each moment may be pre-calibrated. First, the preset weights of several moments are calibrated, and the specific method includes steps S111-S113. Then, the preset weights of these moments are combined for processing to obtain the preset weights corresponding to each moment, and the specific method includes steps S114-S115.

[0057] Step S111, predicting a first rocket mass corresponding to the target time during the flight of the carrier rocket;

[0058] Step S112, determining the target mode slope corresponding to the first rocket mass according to the corresponding relationship between the rocket mass and the mode slope of the carrier rocket;

[0059] Step S113, determining the corresponding preset weights according to the target vibration mode slopes corresponding to the respective rate gyros at the target time.

[0060] Step S114, determining the preset weights of X moments, where X is a positive integer;

[0061] Step S115, interpolating the preset weights at the X moments to obtain a continuous association relationship between each moment and the preset weight. The sum of the preset weights of the launch vehicle corresponding to the continuous association relationship is 1.

[0062] That is to say, the preset weights corresponding to multiple discrete moments are first estimated, and then interpolation processing is performed based on the preset weights corresponding to the multiple discrete moments, so that a continuous correlation relationship between each moment and the preset weight can be obtained, and the continuous correlation relationship represents the relationship between any moment and the corresponding preset weight. In some embodiments, for the obtained weight curve, all values ​​on the weight curve are set in a weight data table, so that the elastic vibration control equation can directly find the corresponding weight from the weight data table to achieve weight value adjustment in the elastic vibration control equation.

[0063] By further interpolating the weight parameters that have been obtained, a higher precision correspondence between weight and launch vehicle flight time can be obtained, so as to ensure that during the actual flight of the launch vehicle, the preset weight adjustment can be achieved in a shorter time, thereby improving the control accuracy of the launch vehicle.

[0064] Among them, the sum of the preset weights of each rocket stage on the carrier rocket is 1.

[0065] Predicting the first rocket mass corresponding to the target time during the flight of the carrier rocket. When predicting the rocket mass, it can be determined by ground testing or modeling simulation.

[0066] According to the correspondence between the rocket mass and the mode slope of the launch vehicle, the target mode slope corresponding to the first rocket mass is determined. As the prior art, the relationship between the rocket mass and the mode slope of the launch vehicle is not further elaborated in this application.

[0067] According to the target vibration mode slopes corresponding to the respective rate gyros at the target time, the corresponding preset weights are determined. The target vibration mode slopes of the launch vehicle can be confirmed based on ground tests or modeling simulations.

[0068] Based on the rigid body angular velocity, vibration mode slope and flight speed of the rocket body, the rate gyro output signal is calculated, specifically:

[0069]

[0070] in, Indicates the vibration and angular velocity of the arrow body output by the rate gyro; It represents the rigid body angular velocity of the launch vehicle; represents the elastic vibration angular velocity of the rocket body, W i ′ (X i ) represents the slope of the i-th order vibration mode at the position of the inertia group, X i represents the axial coordinate of the inertia group, and n represents the number of inertia groups; Indicates the flight speed of the launch vehicle; q i (t) represents the generalized coordinates of the elastic vibration of the i-th mode.

[0071] Among them, the calculation method for confirming the vibration mode slope of the launch vehicle is as follows:

[0072] Determine the rocket mass of the target rocket at different times.

[0073] During the flight of the launch vehicle, the fuel will gradually be consumed, which will inevitably cause the mass of the launch vehicle to change, and accordingly, the vibration mode slope of the launch vehicle will also change.

[0074] In the prior art, the relationship between the modal slope of a launch vehicle and its mass can be confirmed by equation deduction. Therefore, by measuring the mass of the launch vehicle in different flight cycles, the modal slope corresponding to the mass can be calculated respectively.

[0075] Based on the function of the rocket mass and the mode slope, the mode slope corresponding to different rocket masses is calculated. The purpose of this step is to directly calculate the corresponding mode slope after determining the mass parameters of the launch vehicle during flight.

[0076] Among them, for the mass measurement of the launch vehicle, a ground measurement test is adopted to obtain the mass of the launch vehicle in different flight time periods during the operation of the launch vehicle engine in different time periods and record them.

[0077] Among them, for the mass of the rocket body in different flight time periods, the interpolation method can also be used to determine the mass of the launch vehicle, thereby obtaining a correlation curve between the mass of the launch vehicle and the flight time.

[0078] Among them, during ground tests, the overall mass of the carrier rocket can also be determined based on the engine ignition time of the carrier rocket, so as to better simulate the mass changes of the carrier rocket during actual flight.

[0079] In some embodiments, the mass change of the launch vehicle during the actual flight process is determined by the launch vehicle itself based on its flight status, specifically:

[0080] Identify the flight mass of the launch vehicle. Specifically, the mass of the launch vehicle is identified based on the added table data, the number of engines, the startup data, etc. The equation is as follows:

[0081]

[0082] Among them, C A It represents the axial force coefficient, which is a known binding amount. During the flight of the launch vehicle, it can be calculated according to zero angle of attack and zero sideslip, and the rudder angle and Mach number are obtained based on the real-time difference solution of the flight control; ΔC A represents the axial force coefficient correction term, which is determined by the known binding amount, including the bottom resistance term occupied by the engine nozzle and the friction resistance term affected by the flight altitude (Reynolds number); q represents the dynamic pressure, which is calculated in real time based on the density and the speed of the rocket body; S represents the aerodynamic reference area, which is the known binding amount; I s represents the engine specific impulse, which is a known binding quantity; P0 is the atmospheric pressure at sea level, in Pa, and is taken as 100660Pa; P H is the atmospheric pressure, which varies with altitude and is the known binding volume; M k Indicates the current arrow mass, in kg; M k-1 Indicates the mass of the arrow body in the previous cycle, in kg; W x1 It represents the axial overload under the arrow system, which is measured in real time by adding a meter; T represents the sampling period, which is 0.01s.

[0083] Based on this scheme, the mass of the launch vehicle during flight can be determined in real time, and based on the mass parameter, the vibration mode slope of the launch vehicle can be calculated.

[0084] Among them, for the obtained modal slope of the launch vehicle, an interpolation algorithm is also used to interpolate the modal slope to obtain a distribution curve of the modal slope of the target rocket. The purpose of this process is that the modal slope of the launch vehicle is usually in a continuously changing operating state. In this case, it can be fully guaranteed that in the elastic vibration control process of the launch vehicle, the modal slope can be more accurately confirmed based on the smooth curve.

[0085] Among them, for the interpolation algorithm adopted, based on the launch vehicle vibration mode slope parameters obtained at different time nodes, the specific vibration mode slope parameters at the previous and next time nodes are determined, and the changing slope of the vibration mode slope is calculated, so that within the slope parameter, based on the time node and the mass parameter, the two vibration mode slopes are interpolated to obtain the interpolation result of the vibration mode slope, and then the vibration mode slope curve is generated.

[0086] In some embodiments, for the generated mode slope curve, a mode slope data table is generated based on the values ​​in the curve, so as to facilitate directly obtaining the corresponding mode slope data for calculation during the weight calculation process.

[0087] The elastic vibration coordinates of the rocket body are quantities that vary with the flight time of the rocket body and are determined by the following equation:

[0088]

[0089] Among them, q i (t) represents the generalized coordinates of elastic vibration of the i-th mode; represents the acceleration of the i-th vibration mode; represents the velocity of the i-th vibration mode; ω i represents the natural frequency of the i-th vibration mode; ξ i represents the damping coefficient of the i-th order vibration mode; Q i represents the generalized force of the i-th vibration mode; M i represents the generalized mass of the i-th vibration mode.

[0090] The above equation only explains the relationship between the velocity, acceleration, generalized coordinates, natural frequency and mass of the elastic vibration of the launch vehicle, which is used to illustrate that the modal slope has a functional relationship with the mass of the launch vehicle, and the modal slope W of the launch vehicle can be determined based on this functional relationship. i ′ (X i ). In addition, considering that the change in the mass of a launch vehicle during flight can be accurately determined based on ground tests or modeling simulations, and the corresponding relationship between the mass of the rocket and the engine running time is also fixed, the modal slope of the launch vehicle and the preset weight calculated by the modal slope can establish a modal slope and / or preset weight-flight time relationship data table.

[0091] Take the two-stage launch vehicle as an example, where two rate gyros are set. The equations of the two rate sink gyros are as follows:

[0092]

[0093] in, and Respectively represent the arrow body vibration and angular velocity output by the first rate gyro and the second rate gyro; W1 represents the rigid angular velocity of the rocket body. This parameter is the overall parameter of the launch vehicle, that is, the rigid angular velocity of the rocket body of each stage of the launch vehicle is the same; ′ 1 and W1 ′2 respectively represent the vibration mode slopes of the first rate gyro and the second rate gyro. This parameter is related to the mass of the launch vehicle, so the output signals of the rate gyro of different rocket stages can be judged in specific processing; It represents the flying speed of the rocket body. Obviously, for the launch vehicle, the flying speed of each stage on the rocket body is the same.

[0094] After confirming the output signal of each rate gyro of the launch vehicle, a weight is set for each rate gyro output signal, and the weight and the rate gyro output signal are fused to obtain the following equation:

[0095]

[0096] Among them, K1 and K2 represent the weights of the output signals of the two rate gyros, Represents the angular velocity of the elastic vibration of the overall structure of the launch vehicle. Expanding the output equation of each rate gyro, we can get the above equation.

[0097] It can be understood that for all rate gyros, the final weights set are such that the sum of all weights is 1, that is:

[0098] K1+K2=1,

[0099] Wherein, based on the superposition equation of each rate gyro signal, the weight relationship between the mode slope and each rate gyro output signal is calculated, and the following result can be obtained:

[0100]

[0101] Among them, the rigid angular velocity of the rocket body is an inherent parameter in the launch vehicle, which is obviously impossible to eliminate. It is also obviously continuous during the flight of the launch vehicle. Therefore, it is only necessary to consider that the last two parameters in the above equation can be offset. Then, under this technical idea, the final equation is:

[0102]

[0103] This parameter is the weight of the second rate gyro output signal, and the weight is independent of other parameters in the launch vehicle. The weight parameter can be determined based only on the vibration mode slope.

[0104] Correspondingly, for K1, the equation is:

[0105]

[0106] The weight of the output signal of each rate gyro is calculated based on the mode shape slope of each rate gyro.

[0107] During the flight of the launch vehicle, the mass of the rocket body gradually decreases and the frequency of the rocket body gradually increases. The rate gyro senses this change and can be based on the relationship between the vibration slope and the mass, that is, W i ′ (X i )=f(M), where M is the real-time mass of the rocket body, and the corresponding K1 and K2 are also functions of the mass. Note that K1=f1(M), K2=f2(M). If K1 and K2 can be bound in advance according to the mass interpolation function and the flight mass can be identified in real time during the flight, K1 and K2 can be interpolated.

[0108] Among them, interpolation processing is also required. In the interpolation processing of weights, since the relationship between the mass of the launch vehicle and the weight is not linear, it cannot be processed by using linear interpolation method. At the same time, when using quadratic function for interpolation processing, there will be a problem of excessive calculation amount. Therefore, the quadratic Newton difference quotient formula is used for interpolation processing to make it easier for the attitude control computer of the launch vehicle to perform calculation processing. Specifically:

[0109] K1 consists of discrete points, namely (x1,f(x1)), (x2,f(x2))···(x n ,f(x n ), define f[x1 x2···x n ] represents the polynomial x n-1 The coefficient of .

[0110] Take three points (x1,f(x1)), (x2,f(x2)), (x3,f(x3)), and the interpolation formula is:

[0111] F(x)=f[x1]+f[x1 x2](x-x1)+f[x1 x2 x3](x-x1)(x-x2).

[0112] in:

[0113] f[x1]=f(x1);

[0114]

[0115] Take K1=F(x), the calculation of K2 is similar to that of K1.

[0116] When this method is adopted, it can be ensured that in the calculation of the difference in weights, interpolation can be performed based on the weight values ​​that have been calculated, thereby reducing the interval between the weight values ​​and improving the control accuracy of the launch vehicle.

[0117] In some embodiments, after the weights are determined, during the flight of the launch vehicle, the weights are interpolated based on the control computer of the launch vehicle, and the launch vehicle is attitude controlled according to the weight values ​​obtained by the interpolation calculation.

[0118] The purpose of this step is that for each rate gyro, the final result can be obtained based on all the parameters obtained by the target, so as to obtain the elastic control equation of the launch vehicle for the overall structure of the launch vehicle. In addition, since the specific value of the vibration mode slope has been obtained, the corresponding weight can be set for each rate gyro output signal during the flight of the launch vehicle.

[0119] After obtaining the actual output signal at the target moment and the preset weight, step S120 may be performed.

[0120] Regarding step S120, the overall elastic vibration angular rate of the launch vehicle at the target moment is determined according to the actual output signals corresponding to each of the rate gyroscopes at the target moment and the preset weights.

[0121] After the previous step, the weights of all rate gyro output signals can be determined. During the flight period of the launch vehicle, the control system adjusts the preset weights of the rate gyro output signals based on the current flight time of the launch vehicle, and multiplies the preset weights with the output signals of the corresponding rate gyros to achieve weighted fusion, which can suppress the elastic vibration of the launch vehicle.

[0122] After the weight of the rate gyro output signal has been determined, and the preset weight of each rate gyro at the target time has been determined, the elastic vibration generated in the launch vehicle is controlled by adjusting the preset weight of the rate gyro and obtaining the output signal of the rate gyro.

[0123] Among them, a weight curve can also be obtained based on the distribution curve of the target rocket vibration mode slope, and a weight data table can be generated. The purpose of this step is to directly calculate the weights of each rate gyro that can be obtained during the flight of the launch vehicle based on all the vibration mode slopes that have been obtained, so as to perform calculations. This method belongs to a method for real-time calculation of preset weights during the flight of a launch vehicle. It is essentially the same as the method of setting a preset weight data table. The effect is that there is no need to obtain preset weights based on ground tests or simulations, there are fewer steps in data processing, and based on recalculation, the accuracy of the weights calculated in real time is guaranteed to be higher than the results obtained from ground tests or simulations.

[0124] In some embodiments, all calculated values ​​for the launch vehicle mode slopes are set in a chart, and the weight of the calculation results obtained for each mode slope is calculated, and based on the obtained weight value, a weight curve of each rate gyro output signal is drawn.

[0125] Based on the weight values ​​in the weight curve or the weight values ​​in the weight data table, the weight of each rate gyro on the target rocket is adjusted. The purpose of this step is to directly adjust the weight set in the elastic control equation of the launch vehicle based on the flight time of the target during the flight of the launch vehicle.

[0126] In some embodiments, a correlation relationship between a weight parameter and a weight during the launch vehicle flight is established to adjust the weight in the launch vehicle elastic control equation.

[0127] In some embodiments, the weight parameters are interpolated based on the weight parameters that have been obtained. Specifically, based on two adjacent weight values ​​and the overall trend of all weight values, the amount of weight values ​​inserted during the weight parameter processing of the target is determined, and then a more accurate weight curve is obtained, and all weight values ​​generated in the weight curve are confirmed.

[0128] In some embodiments, after performing interpolation operations on the acquired weight parameters, the correlation between the weights and the launch vehicle flight time is reformulated, so that during the flight of the launch vehicle, the weights of each rate gyro in the elastic control equation of the launch vehicle can be more accurately adjusted based on the flight time.

[0129] Among them, after obtaining the preset weight of the rate gyro and the output signal of the rate gyro, based on the flight time of the launch vehicle, the flight control system of the launch vehicle directly selects the preset weight corresponding to the flight time period based on the current flight time, calculates the product of each rate gyro output signal and the preset weight, and sums the products. This method can offset the elastic vibration generated by each rocket stage in the launch vehicle.

[0130] Regarding step S130, the attitude of the carrier rocket at the target moment is controlled according to the overall elastic vibration angular rate of the carrier rocket at the target moment.

[0131] The principle of the elastic vibration suppression method during the flight of a launch vehicle is to weight and fuse the output signals of each rate gyro to offset the elastic vibration signal.

[0132] The output signals of each rate gyro are weighted and fused to obtain the superposition equation of each rate gyro signal. The purpose of this step is that in the processing of the rate gyro, various parameters can be known through calculation or ground testing of the launch vehicle, which means that in the specific processing process, the output equations of all rate gyros are weighted and fused to obtain the result.

[0133] The final output signal is used for the attitude control process of the launch vehicle to ensure that the launch vehicle can fly according to the set attitude control requirements.

[0134] However, considering that the installation position of the rate gyro is deeply related to the accuracy of its output signal during the flight of the launch vehicle, it is necessary to determine the installation position of the rate gyro.

[0135] Among them, the installation position of each rate gyro in the launch vehicle also needs to be configured based on the overall structure and processing method of the launch vehicle. The purpose of this step is to ensure that the elastic vibration of the rocket body can be controlled based on the rate gyro by reasonably determining the installation position of the rate gyro.

[0136] The rate gyro is installed at adjacent antinodes in the vibration field generated by the elastic vibration of the launch vehicle. The purpose of this step is to fully reduce the disturbance suffered by the rate gyro when acquiring the measurement signal by selecting the position with the smallest vibration mode slope, and at the same time, to control the elastic vibration of the launch vehicle by using relatively smaller output parameters during operation.

[0137] In some embodiments, for the same type of launch vehicle, the installation positions of the rate gyros on different stages of the launch vehicle are confirmed through ground tests to achieve appropriate configuration of the installation positions of the rate gyros.

[0138] For the control of the launch vehicle, the main purpose of the above steps S110-S130 is to eliminate the elastic vibrations generated during the flight of the launch vehicle. However, the environment during the flight of the launch vehicle is relatively complex, and the elastic vibrations of the launch vehicle may not be completely offset. Therefore, it is also necessary to process the possible residual elastic vibration signals.

[0139] The overall elastic vibration angular frequency of the carrier rocket at the target time is sent to the filter to filter out the elastic vibration in the overall elastic vibration angular rate of the carrier rocket at the target time. The purpose of this step is that after the elastic vibration of the carrier rocket has been controlled, there may be a large amount of interference signals in the actual output of the equation due to the existence of some other factors. By setting a filter, the output of the elastic control equation of the target rocket can be filtered, and then the final output signal is guaranteed to have a better processing result.

[0140] Among them, in the elastic vibration control equation of the launch vehicle, the rate gyro is used to obtain the vibration signals between different stages of the launch vehicle, and after processing, the rate gyro outputs the signal. Then, based on the weight parameters set in the elastic control equation of the launch vehicle, the elastic vibration of the launch vehicle is controlled, and the final value that can be obtained has a better offset effect.

[0141] Among them, the mathematical form of the filter used is:

[0142]

[0143] Among them, G lb (s) represents the output of the filter, ω0 represents the corresponding characteristic point of the launch vehicle body and the corresponding order natural circular frequency; ξ1 represents the corresponding characteristic point of the rocket body and the corresponding order damping coefficient; ξ2 represents the damping coefficient, which needs to be selected in the adjustment of the filter. Specifically, it is necessary to consider the stability and effect of the entire attitude control loop in the specific selection and make the final determination. In addition, under normal circumstances, the relationship between ξ1 and ξ2 can be determined, and ξ2 is about 15 times of ξ1.

[0144] Among them, the two parameters ω0 and ξ1 are determined by the test frequency and damping coefficient of the rocket body, so as to ensure that the control quality of the elastic vibration of the set filter can be fully improved.

[0145] After filtering, the control equation for the elastic vibration of the launch vehicle body is:

[0146]

[0147] Among them, U lb Represents the elastic vibration control output signal during the launch vehicle flight; is the output after processing based on the elastic vibration control equation of the launch vehicle; G lb(s) is the mathematical equation of the filter. Based on this equation, the elastic vibration that cannot be completely suppressed in the elastic vibration control equation of the launch vehicle can be filtered out, so that the final output result has higher accuracy.

[0148] Based on the above content, the launch vehicle is controlled, specifically: during the flight of the launch vehicle, based on the flight time of the launch vehicle, the preset weight is adjusted, and the actual output signal of the rate gyro at the target time of the flight time of the launch vehicle is obtained to suppress the elastic vibration of the launch vehicle. After the elastic vibration is suppressed and filtered, the output signal is the control signal of the launch vehicle.

[0149] By setting a rate gyro for each stage of the launch vehicle and performing measurements during the flight of the launch vehicle, the elastic vibration signals of each stage of the launch vehicle are obtained. At the same time, before the actual flight of the launch vehicle, the weights of the elastic vibration signals of the launch vehicle at different flight times are confirmed based on ground tests and modeling simulations. After obtaining these two types of parameters, the elastic vibration signals of each stage on the launch vehicle are weighted and offset.

[0150] When determining the installation position of the launch vehicle rate gyro, the following requirements need to be considered:

[0151] 1. Distribution of the vibration mode slope of the rocket body. The vibration mode is close to cosine or sine distribution, and the vibration mode slope of the rocket body (the derivative of the vibration mode with respect to position) corresponds to the sine (or cosine) distribution. The vibration mode slope of the rocket body is generally the smallest at the antinode (peak and trough) and the largest at the node. Therefore, it is necessary to determine the vibration mode slope and distribution curve of the rocket body through the simulation analysis of the rocket body elasticity modeling and the modal full load and no-load tests.

[0152] 2. Restrictions on the installation position on the launch vehicle. The installation position on the launch vehicle is generally not the head or tail of the rocket body. The head will generally follow the rocket until the separation of the rocket and the satellite. Placing a rate gyro on the head will result in a large loss of carrying capacity. The tail of the rocket body is generally used to place actuators, such as the nozzle of the swing jet engine. When the engine is working, the temperature here is high, the vibration is violent, and the environment is complex, which is not suitable for the placement of rate gyros. Rate gyros cannot be installed on oxidizer or fuel tanks. In addition, it must be considered that after the separation of the first stage, the first stage wreckage will be thrown away, and the rate gyro cannot be installed on the wreckage.

[0153] 3. Requirements of attitude control system for measuring angular velocity. Taking the commonly used liquid rocket as an example, the first-order rocket body frequency (1Hz~2Hz) is generally low, which is easy to interweave with the control loop crossing frequency (1rad~3rad). Phase stability is usually adopted, and high requirements are placed on the measurement of angular velocity. It is hoped that the amplitude of the elastic component of the angular velocity measured by the inertial group is as small as possible.

[0154] Based on the above three points, in the installation of the rate gyro of the launch vehicle, for the first-stage rocket, such as Figure 4 As shown, the rate gyro needs to be set at the antinode, that is, Figure 4 At point C in .

[0155] For the second-stage rocket, the frequency of the second-stage body is relatively high, generally above 5Hz, which is far from the control loop crossing frequency. The control loop adopts amplitude stability, does not require high angular velocity measurement, and has a flexible installation position. The gyro needs to be installed at the antinode, such as Figure 5 At points B and D shown.

[0156] Among them, the installation position of the gyroscope is basically determined according to the above three requirements, that is, under the first-order rocket body frequency condition, the gyroscope is installed near point C, one on each side, and arranged symmetrically. Under the second-order rocket body frequency condition, the gyroscope is installed near point B (or point D), one on each side, and arranged symmetrically. It is also possible to place one on the left (or right) side of point B and one on the left (or right) side of point D to ensure that the two gyroscopes are at the same distance from point B and point D respectively.

[0157] Among them, in the actual operation process, due to the uneven distribution of the rocket mass, the position of the gyroscope is too close to the peak or trough, which may cause the elastic angular rates measured by the two gyroscopes to have the same sign, and the above method will fail. Therefore, it is necessary to further adjust the position of the gyroscope based on the distribution of the rocket body vibration mode slope obtained from the modal test to ensure that the gyroscopes are located at positions with opposite vibration mode slopes. At the same time, the vibration mode slope is as small as possible.

[0158] The embodiment of the present application sets a rate gyro on each rocket stage of the launch vehicle, and then during the flight of the launch vehicle, collects the actual output signal of the rate gyro at any target time, and determines the preset weight corresponding to the target time. Based on the corresponding actual output signal and the preset weight, the overall elastic vibration angular velocity of the launch vehicle at the target time can be obtained, and then based on the overall elastic vibration angular velocity, the launch vehicle is controlled. It can be seen that the embodiment of the present application abandons the method of improving the elastic vibration control of the launch vehicle by changing the aspect ratio and materials in the related art, but instead improves the elastic vibration control of the launch vehicle by arranging the rate gyro, relying on the actual output signal collected by the rate gyro at the target time and the corresponding preset weight. On the basis of being able to realize the elastic vibration control of the launch vehicle, the complexity of the elastic vibration control is greatly simplified, and the implementation cost of the elastic vibration control is greatly reduced.

[0159] Based on the same inventive concept, the present application embodiment provides the following Figure 2 A schematic diagram of a launch vehicle elastic control system is shown.

[0160] The whole system includes a vibration mode slope calculation module 21, a rate gyro output signal weight calculation module 22, an elastic vibration control equation establishment module 23 and a filtering module 24, wherein all modules are executed by an electronic device 300 corresponding to the launch vehicle control method.

[0161] The vibration mode slope calculation module 21 is used to calculate the vibration mode slope of the launch vehicle, specifically:

[0162] The vibration mode slope calculation module 21 is configured to obtain the mass measurement result of the launch vehicle based on the ground test and / or the modeling simulation module.

[0163] The mode slope calculation module 21 determines the mode slope corresponding to the target mass of the launch vehicle based on the obtained launch vehicle mass measurement result and the calculated relationship between the mass of the launch vehicle and the mode slope, and stores it.

[0164] The mode shape slope calculation module 21 further performs interpolation processing on the mode shape slope after calculating the mode shape slope. Based on the values ​​of two adjacent mode shape slopes obtained, the slope of the curve between the two mode shape slopes is obtained and interpolated.

[0165] In some embodiments, after the mode slope is interpolated, a correlation between the interpolated mode slope and the launch vehicle flight is established, thereby improving the correlation between the mode slope and the launch vehicle flight time, thereby further improving the accuracy of the calculated weight.

[0166] The rate gyro output signal weight calculation module 22 is configured to obtain the calculated mode shape slope and calculate the weight of the rate gyro output signal based on the mode shape slope.

[0167] Among them, the rate gyro output signal weight calculation module 22, wherein for each rate gyro output signal weight calculation equation, taking the carrier rocket as a two-stage rocket as an example, wherein specifically:

[0168]

[0169] Among them, W1 ′ 1 and W1 ′ 2 respectively represent the vibration mode slopes of the first rate gyro and the second rate gyro, and this parameter is related to the mass of the launch vehicle. Therefore, in the specific processing, the output signals of the rate gyros of different rocket stages can be judged; K1 and K2 respectively represent the weights of the output signals of the two rate gyro.

[0170] Among them, for the calculated weight parameters, the weights are interpolated, based on two adjacent weight values, the slopes of the two weight values ​​are calculated, and based on the parameter curve of the weight values, two connected weight values ​​are interpolated to obtain more weight values, so that the weight curve has higher smoothness.

[0171] Among them, after all the weights are calculated, all the weight values ​​are set in a weight data table, so that based on the data table and the weight parameters recorded in the data table, the weight values ​​in the elastic vibration control equation of the launch vehicle are replaced.

[0172] Among them, the elastic vibration control equation establishment module 23 is used to establish the vibration equation for the launch vehicle, specifically:

[0173] Among them, the elastic vibration control equation establishing module 23 is configured to obtain weight values, adjust the weights in the elastic vibration control equation of the launch vehicle, and also obtain the output signal of the rate gyro, so as to control the elastic vibration.

[0174] Among them, according to the weight curve or weight data table that has been set, based on the information of the weight curve or weight data table, in the established elastic control equation, the weight is adjusted according to the weight curve or weight data table and the flight time of the launch vehicle.

[0175] In some embodiments, a weight data table is generated for the calculated curve to ensure that in the established launch vehicle elastic control equation, corresponding weight values ​​are selected from the generated weight table to adjust the elastic control equation.

[0176] During the flight of the launch vehicle, the rate gyro calculates the output signal of the rate gyro based on the operating parameters of the launch vehicle, and multiplies the calculated output signal by the corresponding weight.

[0177] In some embodiments, based on ground tests and / or modeling simulations, the output signal of the rate gyro during different flight processes of the launch vehicle is determined, and the output signal is stored in a data table, and the output signal in the data table is multiplied by the corresponding weight and then fused.

[0178] The filtering module 24 is used to filter the launch vehicle control signal output by the launch vehicle elastic control equation after filtering, so as to ensure the accuracy of the output signal.

[0179] Among them, the filtering module 24 is configured to receive the output result of the elastic vibration equation establishment module, and the signal after filtering the control signal of the launch vehicle, so as to ensure that the signal that has not been filtered in the signal can be finally processed.

[0180] The filter in the filter module 24 is configured as follows:

[0181]

[0182] Among them, G lb (s) represents the output of the filter, ω0 represents the corresponding characteristic point of the launch vehicle body and the corresponding order natural circular frequency; ξ1 represents the corresponding characteristic point of the rocket body and the corresponding order damping coefficient; ξ2 represents the damping coefficient, which needs to be selected in the adjustment of the filter. Specifically, it is necessary to consider the stability and effect of the entire attitude control loop in the specific selection and make the final determination. In addition, under normal circumstances, the relationship between ξ1 and ξ2 can be determined, and ξ2 is about 15 times of ξ1.

[0183] Among them, after filtering processing, it can be guaranteed that the elastic vibration signal of the launch vehicle can be completely filtered out in the output signal.

[0184] like Figure 3 As shown, it is a schematic diagram of an electronic device 300 of a launch vehicle elastic vibration control system provided in an embodiment of the present invention, which includes a processor 320 and a memory 310 for storing executable instructions of the processor 320, and the processor 320 is configured to execute a launch vehicle elastic control method of any one of the embodiments.

[0185] Among them, the memory 310 stores all computer programs 311 involved in the present invention, all stored data, including mass information of the launch vehicle obtained based on ground tests and / or obtained, and then the vibration mode slope calculated based on the mass information, or weight data, to ensure that the weight parameters finally obtained can be used in the control equation of elastic vibration.

[0186] The processor 320 obtains the relevant signals in the elastic control equation of the launch vehicle to achieve effective control of the elastic vibration of the launch vehicle. In addition, the processor 320 is also configured to obtain the output signal of the rate gyro and adjust the weight parameter in the elastic vibration control equation to output the signal.

[0187] Among them, the present invention also discloses a non-temporary computer-readable storage medium, the instructions in which are executed by the processor 320 of the electronic device 300 to execute a launch vehicle control method as provided above.

[0188] Since the electronic device 300 introduced in this embodiment is an electronic device 300 used to implement the information processing method in the embodiment of the present application, based on the information processing method introduced in the embodiment of the present application, a person skilled in the art can understand the specific implementation of the electronic device 300 of the present embodiment and its various variations, so how the electronic device 300 implements the method in the embodiment of the present application is not described in detail here. As long as a person skilled in the art implements the electronic device 300 used in the information processing method in the embodiment of the present application, it belongs to the scope of protection of the present application.

[0189] The technical solutions in the above embodiments of the present application have at least the following technical effects or advantages:

[0190] Since a corresponding weight is set for each rate gyro, the problems of reduced carrying capacity and reduced control flexibility of the rocket body caused by the adjustment of the rocket body structure itself can be effectively solved.

[0191] In addition, based on the functional relationship between the mass and the modal slope during the flight of the launch vehicle, the modal slope at different flight times is determined directly based on actual measurement or simulation, and the weight of the output signal of each rate gyro of the launch vehicle is calculated based on the modal slope to achieve real-time control of the elastic vibration of the launch vehicle.

[0192] It will be appreciated by those skilled in the art that embodiments of the present invention may be provided as methods, systems, or computer program products. Therefore, the present invention may take the form of a complete hardware embodiment, a complete software embodiment, or an embodiment combining software and hardware. Furthermore, the present invention may take the form of a computer program 311 product implemented on one or more computer-usable storage media (including but not limited to disk storage, CD-ROM, optical storage, etc.) containing computer-usable program code.

[0193] The present invention is described with reference to flowcharts and / or block diagrams of methods, devices (systems), and computer program 311 products according to embodiments of the present invention. It should be understood that each process and / or block in the flowchart and / or block diagram, as well as a combination of processes and / or blocks in the flowchart and / or block diagram, can be implemented by computer program 311 instructions. These computer program 311 instructions can be provided to a processor 320 of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable data processing device to produce a machine, so that the instructions executed by the processor 320 of the computer or other programmable data processing device generate instructions for implementing the processes in the flowchart and / or block diagram. Figure 1 A process or multiple processes and / or boxes Figure 1 A device that provides the functions specified in a block or multiple blocks.

[0194] These computer program 311 instructions may also be stored in a computer readable memory 310 that can direct a computer or other programmable data processing device to operate in a specific manner, so that the instructions stored in the computer readable memory 310 produce an article of manufacture including an instruction device that implements the process Figure 1 A process or multiple processes and / or boxes Figure 1 A function specified in one or more boxes.

[0195] These computer program 311 instructions may also be loaded onto a computer or other programmable data processing device so that a series of operating steps are executed on the computer or other programmable device to produce a computer-implemented process, thereby providing instructions for executing on the computer or other programmable device to implement the process. Figure 1 A process or multiple processes and / or boxes Figure 1 The steps for the functions specified in one or more boxes.

[0196] Although the preferred embodiments of the present invention have been described, those skilled in the art may make other changes and modifications to these embodiments once they have learned the basic creative concept. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments and all changes and modifications that fall within the scope of the present invention.

[0197] Obviously, those skilled in the art can make various changes and modifications to the present invention without departing from the spirit and scope of the present invention. Thus, if these modifications and variations of the present invention fall within the scope of the claims of the present invention and their equivalents, the present invention is also intended to include these modifications and variations.

Claims

1. A launch vehicle control method, characterized in that: Applied to the flight process of a launch vehicle, the method comprises: Obtaining actual output signals of rate gyros corresponding to respective rocket stages on the launch vehicle at a target time, and obtaining preset weights corresponding to respective rate gyros of the launch vehicle at the target time; Determine the overall elastic vibration angular rate of the launch vehicle at the target time according to the actual output signals respectively corresponding to the rate gyros at the target time and the preset weights; The attitude of the carrier rocket at the target moment is controlled according to the overall elastic vibration angular rate of the carrier rocket at the target moment.

2. The launch vehicle control method according to claim 1, characterized in that: Before the launch vehicle flies, the method further comprises: Predicting a first rocket mass corresponding to the target time during the flight of the carrier rocket; Determining the target mode slope corresponding to the first rocket mass according to the corresponding relationship between the rocket mass and the mode slope of the carrier rocket; According to the target vibration mode slopes corresponding to the respective rate gyros at the target time, the corresponding preset weights are determined respectively.

3. The launch vehicle control method according to claim 2, characterized in that: The target vibration mode slopes corresponding to the respective rate gyros at the target time are determined by ground tests and / or modeling simulations of the launch vehicle.

4. The launch vehicle control method according to claim 1, characterized in that: The method further comprises: Determine the preset weights of X moments; X is a positive integer; The preset weights at the X moments are interpolated to obtain a continuous correlation relationship between each moment and the preset weight.

5. The launch vehicle control method according to claim 3, characterized in that: The sum of the preset weights of the launch vehicle corresponding to the continuous association relationship is 1.

6. The launch vehicle control method according to claim 1, characterized in that: The controlling the attitude of the carrier rocket at the target moment according to the overall elastic vibration angular rate of the carrier rocket at the target moment comprises: Performing filtering processing on the overall elastic vibration angular frequency of the launch vehicle at the target time to filter out the elastic vibration existing in the overall elastic vibration angular rate of the launch vehicle at the target time; The attitude of the launch vehicle at the target moment is controlled according to the overall elastic vibration angular frequency after filtering.

7. The launch vehicle control method according to claim 1, characterized in that: The rate gyro is installed at adjacent antinodes in the vibration field generated by the elastic vibration of the launch vehicle.

8. An electronic device, characterized in that: include: processor; a memory for storing instructions executable by the processor; Wherein, the processor is configured to execute to implement a launch vehicle control method as described in any one of claims 1 to 7.

9. A non-temporary computer-readable storage medium, when the instructions in the storage medium are executed by a processor of an electronic device, enables the electronic device to implement a launch vehicle control method as described in any one of claims 1 to 7.

10. A computer program product, characterized in that The method comprises computer instructions, wherein the computer instructions are executed by a processor to implement a launch vehicle control method as claimed in any one of claims 1 to 7.

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