Method for evaluating carrying capacity of carrier rocket under power system fault
By collecting launch vehicle status information and performing dynamic integral calculations, the problem of assessing the launch vehicle's carrying capacity under propulsion system failure was solved, ensuring that the launch vehicle could enter its original target orbit and improving its reliability and flight safety.
Patent Information
- Application Number
- CN202410253639.4
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2024-03-06
- Publication Date
- 2026-02-06
AI Technical Summary
Existing launch vehicles lack sufficient detection, isolation, and adaptability capabilities in the event of a power system failure, making it impossible to accurately assess their carrying capacity. This could lead to deviations from the planned flight trajectory or crashes, and both misjudgments and omissions could result in launch mission failure.
By collecting launch vehicle status information, it is determined whether the launch capacity assessment start conditions are met. The required oxidizer mass for entering the original target orbit is calculated using dynamic integration. It is then determined whether there is enough oxidizer to enter the original target orbit, and a trajectory reconstruction signal is issued or not.
It enables rapid and accurate assessment of the launch vehicle's carrying capacity under power system failure, ensuring flight safety, reducing misjudgments and omissions, and improving the reliability of the launch vehicle.
Smart Images
Figure CN121474947A_ABST
Abstract
Description
TECHNICAL FIELD
[0001] The present application relates to the field of launch vehicle guidance control, in particular to a launch vehicle carrying capacity evaluation method under power system failure. BACKGROUND
[0002] Among various failures of launch vehicles, power system failure is the most common and the most serious. The existing launch vehicles have insufficient detection and isolation capabilities and adaptability for power system failure. After such failure occurs, the launch vehicle cannot reach the original target orbit due to deviation from the pre-designed flight trajectory, and even crashes eventually. Flight trajectory reconstruction will change the target orbit of the launch vehicle. Therefore, under the condition of primary power system failure, how to quickly and accurately evaluate the launch vehicle carrying capacity and determine whether the launch vehicle can still enter the original target orbit is the premise and basis for flight trajectory reconstruction. If the launch vehicle can still enter the original target orbit, flight trajectory reconstruction should not be performed, and if the launch vehicle cannot enter the original target, flight trajectory reconstruction should be started. Misjudgment and missed judgment will lead to failure of the launch mission. Launch vehicle carrying capacity evaluation under power system failure is crucial to flight safety, and is a technical problem urgently to be solved in the field. SUMMARY
[0003] To solve the above problems, the present application provides a launch vehicle carrying capacity evaluation method under power system failure. The method has clear principles, simple and direct implementation, and can evaluate the carrying capacity of the launch vehicle according to the state information of the launch vehicle under the condition of power system failure, determine whether the launch vehicle can enter the original target orbit, and whether trajectory reconstruction is needed, thereby improving the reliability of the launch vehicle.
[0004] The object of the present application can be achieved by the following technical solutions:
[0005] A launch vehicle carrying capacity evaluation method under power system failure, comprising the following steps:
[0006] Step S1: Collecting state information of the launch vehicle, including shutdown margin and speed at the time of primary pre-command shutdown, primary shutdown type, speed, position and program angle at the time of primary and secondary separation;
[0007] Step S2: Determining whether the state of the launch vehicle meets the launch capacity evaluation start condition, if the start condition is not met, the launch capacity evaluation is not started, and no trajectory reconstruction signal is sent; if the start condition is met, the launch capacity evaluation is started, and step S3 is entered;
[0008] Step S3: Taking the speed, position, program angle, oxidizer mass and total mass of the launch vehicle at the time of primary and secondary separation as the initial state, calculating the oxidizer mass consumed by the launch vehicle to enter the original target orbit through dynamics integration;
[0009] Step S4: judging whether the oxidizer mass consumed by the launch vehicle to enter the original target orbit is greater than the available oxidizer mass, if the oxidizer mass consumed is less than or equal to the available oxidizer mass, the launch vehicle can enter the original target orbit, and no trajectory reconstruction signal is sent; if the oxidizer mass consumed is greater than the available oxidizer mass, the launch vehicle cannot enter the original target orbit, and a trajectory reconstruction signal is sent.
[0010] Further, in step S2, the launch capability evaluation starting condition includes:
[0011] (1) the first shutdown type is small overload shutdown, and the first shutdown margin is greater than
[0012] (2) the first shutdown type is timing shutdown;
[0013] (3) the first shutdown type is depletion shutdown, and the first pre-order shutdown time speed deviation is less than
[0014] Any one of the above three conditions is met, that is, the launch capability evaluation starting condition is met.
[0015] Preferably, in step S2, and are the binding values designed in advance.
[0016] Further, in step S3, the oxidizer mass consumed by the launch vehicle to enter the original target orbit is calculated by dynamic integration, including:
[0017] The following calculation is repeated until the second shutdown condition is met, and the state quantity x end at the time of stopping calculation is recorded:
[0018]
[0019] In the formula, h is the integration step, x n = [V x ,V y ,V z ,X,Y,Z,m Y ,m] T is the state quantity at time nh, [V x ,V y ,V z ] is the velocity of the launch vehicle, [x,y,z] is the position of the launch vehicle, m Y is the oxidizer mass, and m is the mass of the launch vehicle.
[0020] The oxidizer mass consumed by the launch vehicle to enter the original target orbit is:
[0021] Δm Y =m Y0 -m Yend
[0022] In the formula, m Y0 For the mass of secondary oxidizer added, the binding value; m Yend =x end (7) is the mass of the oxidant at the end of the kinetic integration.
[0023] Preferably, in step S3, the second-stage shutdown condition in the dynamic integral is that the semi-major axis calculated based on the dynamic integral state is greater than or equal to the semi-major axis at the second-stage shutdown moment of the original target orbit.
[0024] Furthermore, in step S3, the function f(x) in the dynamic integral is a right-hand function of the exoatmospheric launch vehicle dynamics equations, expressed as follows:
[0025]
[0026] In the formula, T is the thrust calculated based on the thrust model. ψ and yaw angle are the pitch angle and yaw angle obtained from iterative guidance calculations, respectively. x ,g y ,g z [G] represents the Earth's gravitational acceleration calculated based on the Earth's gravitational model. Y G represents the oxidizer consumption per second, and G represents the propellant consumption per second.
[0027] Preferably, in step S3, the step size of the kinetic integration is 1 second.
[0028] Preferably, the payload capacity assessment calculation is performed one second after the second stage ignition of the launch vehicle, and the calculation is performed only once.
[0029] Furthermore, in step S4, the method for calculating the available oxidizer mass for the launch vehicle is as follows:
[0030]
[0031] In the formula, For the mass of oxidizer available for the launch vehicle, m Yy The safety margin of oxidizer required for the launch vehicle to overcome interference.
[0032] A computer-readable storage medium storing a computer program that, when executed by a processor, implements the above-described method.
[0033] By employing the above technical solutions, this invention has the following advantages and positive effects compared with the prior art:
[0034] This invention proposes a method for assessing the carrying capacity of a launch vehicle under propulsion system failure. It is the first to propose a method for determining the start of the carrying capacity assessment and a method for calculating the mass of oxidizer required for the launch vehicle to enter the target orbit. The principle of this invention is clear and the implementation is simple and direct. It can conduct real-time carrying capacity assessment based on the launch vehicle's status information under propulsion system failure, determine whether the launch vehicle can enter the original target orbit, and whether trajectory reconstruction is required, thereby improving the reliability of the launch vehicle. Attached Figure Description
[0035] To more clearly illustrate the technical solutions of the embodiments of the present invention, the accompanying drawings used in the description of the embodiments will be briefly introduced below. Obviously, the drawings described below are merely some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort. In the drawings:
[0036] Figure 1 This is a flowchart of a method for evaluating the carrying capacity of a launch vehicle under power system failure, according to the present invention. Detailed Implementation
[0037] Exemplary embodiments of the present disclosure will now be described in more detail with reference to the accompanying drawings. While exemplary embodiments of the present disclosure are shown in the drawings, it should be understood that the present disclosure may be implemented in various forms and should not be limited to the embodiments set forth herein. Rather, these embodiments are provided to enable a more thorough understanding of the present disclosure and to fully convey the scope of the disclosure to those skilled in the art. It should be noted that, unless otherwise specified, the embodiments and features described herein can be combined with each other. The present invention will now be described in detail with reference to the accompanying drawings and embodiments.
[0038] Example 1
[0039] This invention provides a method for evaluating the carrying capacity of a launch vehicle under propulsion system failure. Please refer to [link / reference]. Figure 1 , Figure 1 This is a flowchart of the launch vehicle carrying capacity assessment method under propulsion system failure according to the present invention, including the following steps:
[0040] Step S1: Collect the status information of the launch vehicle, including the shutdown margin and speed at the first stage pre-command shutdown time, the first stage shutdown type, and the speed, position and program angle at the first and second stage separation time;
[0041] Step S2: Determine whether the launch vehicle status meets the launch capacity assessment activation conditions. If the activation conditions are not met, launch capacity assessment will not be activated and trajectory reconstruction signal will not be issued. If the activation conditions are met, launch capacity assessment will be activated and proceed to step S3.
[0042] Step S3: Using the velocity, position, program angle, oxidizer mass, and total mass of the launch vehicle at the moment of first and second stage separation as the initial state, calculate the mass of oxidizer required for the launch vehicle to enter the original target orbit through dynamic integration;
[0043] Step S4: Determine whether the mass of oxidizer required for the launch vehicle to enter the original target orbit is greater than the mass of available oxidizer. If the mass of oxidizer required is less than or equal to the mass of available oxidizer, the launch vehicle can enter the original target orbit and no trajectory reconstruction signal is issued. If the mass of oxidizer required is greater than the mass of available oxidizer, the launch vehicle cannot enter the original target orbit and a trajectory reconstruction signal is issued.
[0044] Furthermore, in step S2, the conditions for initiating the carrying capacity assessment include:
[0045] (1) The first-level shutdown type is small overload shutdown and the first-level shutdown margin is greater than
[0046] (2) The first-level shutdown type is timed shutdown;
[0047] (3) The first-level shutdown type is exhaustion shutdown and the speed deviation at the time of the first-level pre-command shutdown is less than [missing information].
[0048] If any one of the above three conditions is met, the conditions for initiating the carrying capacity assessment are met.
[0049] Preferably, in step S2, and This is a pre-designed binding value. In this embodiment, and The values are 107 and -204 m / s, respectively.
[0050] Furthermore, in step S3, the mass of oxidizer required for the launch vehicle to enter the original target orbit is calculated through dynamic integration, including:
[0051] Repeat the following calculation until the second-level shutdown condition is met, then stop the calculation and record the state variable x at the time of stopping. end :
[0052]
[0053] In the formula, h is the integration step size, x n =[V x V y V z ,X,Y,Z,m Y [m] T Let [V] be the state variable at time nh. x Vy V z [x, y, z] represents the launch vehicle's velocity, [x, y, z] represents the launch vehicle's position, and m Y m is the mass of the oxidizer, and m is the mass of the launch vehicle;
[0054] The mass of oxidizer required for the launch vehicle to enter its original target orbit is:
[0055] Δm Y =m Y0 -m Yend
[0056] In the formula, m Y0 The mass of the secondary oxidant added, the binding value, in this embodiment is taken as 34800 kg; m Yend =x end (7) is the mass of the oxidant at the end of the kinetic integration.
[0057] Preferably, in step S3, the second-stage shutdown condition in the dynamic integral is that the semi-major axis calculated based on the dynamic integral state is greater than or equal to the semi-major axis at the second-stage shutdown moment of the original target orbit.
[0058] Furthermore, in step S3, the function f(x) in the dynamic integral is a right-hand function of the exoatmospheric launch vehicle dynamics equations, expressed as follows:
[0059]
[0060] In the formula, T is the thrust calculated based on the thrust model. ψ and yaw angle are the pitch angle and yaw angle obtained from iterative guidance calculations, respectively. x ,g y ,g z [G] represents the Earth's gravitational acceleration calculated based on the Earth's gravitational model. Y G represents the oxidizer consumption per second, and G represents the propellant consumption per second.
[0061] Preferably, in step S3, the step size of the kinetic integration is 1 second.
[0062] Preferably, the payload capacity assessment calculation is performed one second after the second stage ignition of the launch vehicle, and the calculation is performed only once.
[0063] Furthermore, in step S4, the method for calculating the available oxidizer mass for the launch vehicle is as follows:
[0064]
[0065] In the formula, The mass of oxidizer usable in the launch vehicle is taken as 34,800 kg in this embodiment; m YyIn this embodiment, the safety margin of oxidizer required for the launch vehicle to overcome interference is set at 300 kg.
[0066] By applying this method, in the event of a propulsion system failure, the launch vehicle's carrying capacity can be assessed in real time based on its status information to determine whether it can return to its original target orbit and whether trajectory reconfiguration is necessary. Flight test results show that the carrying capacity assessment can be completed within one second, enabling a decision on whether to issue a trajectory reconfiguration signal.
[0067] Example 2
[0068] This embodiment discloses a computer-readable storage medium storing a computer program that, when executed by a processor, implements the above-described method.
[0069] The above description is merely a preferred embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the scope of the technology disclosed in the present invention should be included within the scope of protection of the present invention. Therefore, the scope of protection of the present invention should be determined by the scope of the claims.
Claims
1. A method for evaluating the carrying capacity of a launch vehicle under propulsion system failure, characterized in that, Includes the following steps: Step S1: Collect the status information of the launch vehicle, including the shutdown margin and speed at the first stage pre-command shutdown time, the first stage shutdown type, and the speed, position and program angle at the first and second stage separation time; Step S2: Determine whether the launch vehicle status meets the launch capability assessment activation conditions. If the activation conditions are not met, the launch capability assessment will not be activated and the trajectory reconstruction signal will not be issued. If the activation conditions are met, the carrying capacity assessment will be initiated, and the process will proceed to step S3. Step S3: Using the velocity, position, program angle, oxidizer mass, and total mass of the launch vehicle at the moment of first and second stage separation as the initial state, calculate the mass of oxidizer required for the launch vehicle to enter the original target orbit through dynamic integration; Step S4: Determine whether the mass of oxidizer required for the launch vehicle to enter the original target orbit is greater than the mass of available oxidizer. If the mass of oxidizer required is less than or equal to the mass of available oxidizer, the launch vehicle can enter the original target orbit and no trajectory reconstruction signal is issued. If the mass of oxidizer required is greater than the mass of available oxidizer, the launch vehicle cannot enter the original target orbit and a trajectory reconstruction signal is issued.
2. The method for evaluating the carrying capacity of a launch vehicle under propulsion system failure as described in claim 1, characterized in that, In step S2, the conditions for initiating the carrying capacity assessment include: (1) The first-level shutdown type is small overload shutdown and the first-level shutdown margin is greater than (2) The first-level shutdown type is timed shutdown; (3) The first-level shutdown type is exhaustion shutdown and the speed deviation at the time of the first-level pre-command shutdown is less than [missing information]. If any one of the above three conditions is met, the conditions for initiating the carrying capacity assessment are met.
3. The method for evaluating the carrying capacity of a launch vehicle under propulsion system failure as described in claim 2, characterized in that, In step S2, and The binding value is pre-designed.
4. The method for evaluating the carrying capacity of a launch vehicle under propulsion system failure as described in claim 1, characterized in that, In step S3, the calculation of the mass of oxidizer required for the launch vehicle to enter the original target orbit through dynamic integration includes: Repeat the following calculation until the second-level shutdown condition is met, then stop the calculation and record the state variable x at the time of stopping. end : In the formula, h is the integration step size, x n =[V x V y V z ,X,Y,Z,m Y [m] T Let [V] be the state variable at time nh. x V y V z [x, y, z] represents the launch vehicle's velocity, [x, y, z] represents the launch vehicle's position, and m Y m is the mass of the oxidizer, and m is the mass of the launch vehicle; The mass of oxidizer required for the launch vehicle to enter its original target orbit is: Δm Y = m Y0 -m Yend In the formula, m Y0 For the mass of secondary oxidizer added, the binding value; m Yend =x end (7) is the mass of the oxidant at the end of the kinetic integration.
5. The method for evaluating the carrying capacity of a launch vehicle under propulsion system failure as described in claim 4, characterized in that, In step S3, the second-stage shutdown condition in the dynamic integral is that the semi-major axis calculated based on the dynamic integral state is greater than or equal to the semi-major axis at the second-stage shutdown moment of the original target orbit.
6. The method for evaluating the carrying capacity of a launch vehicle under propulsion system failure as described in claim 4, characterized in that, In step S3, the function f(x) in the dynamic integral is a right-hand function of the exoatmospheric launch vehicle dynamics equations, expressed as follows: In the formula, T is the thrust calculated based on the thrust model. ψ and yaw angle are the pitch angle and yaw angle obtained from iterative guidance calculations, respectively. x ,g y ,g z [G] represents the Earth's gravitational acceleration calculated based on the Earth's gravitational model. Y G represents the oxidizer consumption per second, and G represents the propellant consumption per second.
7. The method for evaluating the carrying capacity of a launch vehicle under propulsion system failure as described in claim 4, characterized in that, In step S3, the step size for the kinetic integration is 1 second.
8. The method for evaluating the carrying capacity of a launch vehicle under propulsion system failure as described in claim 1, characterized in that, The payload capacity assessment calculation is performed one second after the second stage ignition of the launch vehicle, and only once.
9. The method for evaluating the carrying capacity of a launch vehicle under propulsion system failure as described in claim 1, characterized in that, In step S4, the method for calculating the available oxidizer mass for the launch vehicle is as follows: In the formula, For the mass of oxidizer available for the launch vehicle, m Yy The safety margin of oxidizer required for the launch vehicle to overcome interference.
10. A computer-readable storage medium storing a computer program, characterized in that, When the computer program is executed by a processor, it implements the method as described in any one of claims 1-9.