Horizontal take-off and landing two-stage entry orbit vehicle thrust failure decision and guidance method, device and storage medium

Abstract

The application discloses a horizontal take-off and landing two-stage entry orbit vehicle thrust failure decision and guidance method and device and a storage medium, relates to the field of aircraft guidance and control technology, and comprises the following steps: solving the trajectory from the failure point to the separation point by adopting sequential convex optimization, and obtaining a feasible sign bit representing the dynamics feasibility of the separation point; if the feasible sign bit is a first preset value, cruise parameters are determined according to the maximum reachable domain area, and a reachable domain after separation is generated, and the reachable domain after separation is used to determine a first feasible airport subset; if the feasible sign bit is not the first preset value or the reachable domain after separation does not cover the airport, the inter-stage separation is cancelled, and an emergency reachable domain is generated, and the emergency reachable domain is used to determine a second feasible airport subset; and target airports are determined from the corresponding feasible airport subsets according to the lexicographic strategy based on airport priorities and required flight ranges. The application realizes online decision and guidance after sudden thrust loss failure of the ascending stage of the vehicle, and improves the survival probability after the failure.

Description

Technical Field

[0001] This application relates to the field of aircraft technology, and in particular to a thrust failure decision-making and guidance method, device and storage medium for a horizontal takeoff and landing two-stage-to-orbit launch vehicle. Background Technology

[0002] HTHL-TSTO launch vehicles, with their potential for full reusability and low launch costs, have become a core direction for the development of future space transportation systems. However, the propulsion systems of this type of launch vehicle (such as deep air-cooled combined cycle engines) are highly complex, facing a more severe risk of thrust loss during the climb phase than traditional rocket propulsion. Once thrust loss occurs, the launch vehicle's flight state will rapidly deviate from the nominal flight profile, easily inducing aerodynamic load overruns and threatening the structural integrity of the airframe, resulting in an extremely narrow window for emergency response.

[0003] Research on decision-making and guidance for launch vehicle ascent phase failures has largely focused on traditional vertical launch rockets, primarily targeting vacuum or near-vacuum flight environments, making it difficult to adapt to the complex flight characteristics of the HTHL-TSTO launch vehicle. HTHL-TSTO requires prolonged maneuvering flight within the atmosphere, encompassing multiple phases including climb, cruise, and unpowered gliding. Following a thrust loss failure, it is necessary not only to reconstruct the continuous flight trajectory but also to simultaneously complete multiple discrete mission decisions, such as stage separation determination, degradation mode switching, and alternate landing airport selection. Current decision-making and guidance for continuous-discrete hybrid systems, primarily relying on pre-set abort modes such as the Space Shuttle's RTLS (Return-to-Launch-Site) and TAL (Trans-Atlantic Landing) modes, lacks sufficient adaptive capability to cope with sudden, unpredictable failures.

[0004] Therefore, how to achieve online decision-making and guidance after a sudden thrust loss failure during the ascent phase of a launch vehicle, and improve the survival probability of the launch vehicle after the failure, is an urgent problem to be solved.

[0005] The above content is only used to help understand the technical solution of this application and does not represent an admission that the above content is prior art. Summary of the Invention

[0006] The main objective of this application is to provide a method, device, and storage medium for thrust failure decision-making and guidance of a two-stage horizontal takeoff and landing launch vehicle, aiming to solve the technical problem of how to achieve online decision-making and guidance after a sudden thrust loss failure during the ascent phase of the launch vehicle, and improve the survival probability of the launch vehicle after the failure.

[0007] To achieve the above objectives, this application proposes a thrust failure decision-making and guidance method for a horizontal takeoff and landing two-stage-to-orbit launch vehicle, the method comprising: Based on the current state vector of the launch vehicle, a convex optimization subproblem is constructed. A sequential convex optimization algorithm is used to iteratively solve the trajectory between the fault point and the separation point in the convex optimization subproblem, and the feasibility flag of the separation point dynamics is obtained. If the feasible flag of the separation point dynamics is the first preset value, then the reachable domain area function is obtained based on the maximum theoretical longitudinal range with power, the range of the unpowered gliding segment, the extreme turning angle, and the minimum turning radius of the launch vehicle in cruise level flight. Based on the area function of the reachable region and the polar equation of the polar boundary of the reachable region, the polar boundary of the reachable region is obtained. If there is an intersection between the preset airport set and the reachable domain, the intersection is taken as the first feasible airport subset, and the required flight distance corresponding to each feasible airport in the first feasible airport subset is obtained. Based on airport priority levels and required flight distance, the target airport is obtained from the feasible airport subset.

[0008] In one embodiment, the step of constructing a convex optimization subproblem based on the current state vector of the carrier, and using a sequential convex optimization algorithm to iteratively solve the trajectory between the fault point and the separation point to obtain the feasible flag of the separation point dynamics includes: Obtain the current state vector of the launch vehicle, and construct a convex optimization subproblem based on the current state vector, wherein the current state vector includes the launch vehicle's geocentric distance, velocity, local ballistic inclination angle, and mass; A sequential convex optimization algorithm is used to iteratively solve the trajectory between the fault point and the separation point in the convex optimization subproblem. During the solution iteration process, the trust region radius of the state variables, the trust region radius of the control variables, and the trust region radius of the terminal time in the convex optimization subproblem are adaptively updated. If the sequential convex optimization algorithm converges during the iterative solution, then the feasible flag for determining the separation point dynamics is the first preset value. If the sequential convex optimization algorithm fails to converge when the iteration is completed, the feasible flag for determining the separation point dynamics is set to the second preset value.

[0009] In one embodiment, the step of obtaining the reachable domain area function based on the maximum theoretical longitudinal range with power, the range during unpowered gliding, the extreme turning angle, and the minimum turning radius of the launch vehicle in cruise level flight includes: If the reachable domain does not intersect with the negative x-axis, the area function of the reachable domain is obtained based on the turning resistance multiplication factor, the maximum theoretical longitudinal range with power, the range of the unpowered gliding segment, the extreme turning angle, and the minimum turning radius.

[0010] In one embodiment, when the reachable domain intersects the negative x-axis, the limiting turning angle includes the turning angle at the first intersection of the outer boundary and the negative x-axis and the turning angle at the second intersection of the inner boundary and the negative x-axis. The step of obtaining the area function of the reachable domain based on the maximum theoretical longitudinal range with power, the range of the unpowered gliding segment, the limiting turning angle, and the minimum turning radius of the vehicle in cruise level flight includes: Based on the turning angle at the first intersection, the minimum turning radius, the turning resistance multiplication factor, the maximum theoretical longitudinal range with power, and the range of the unpowered gliding segment, the outer boundary area function is obtained. Based on the turning angle at the second intersection point, the minimum turning radius, and the unpowered gliding distance, obtain the inner boundary area function; The reachable domain area function is obtained based on the outer boundary area function and the inner boundary area function.

[0011] In one embodiment, the step of obtaining the polar coordinate boundary of the reachable region based on the area function of the reachable region and the polar coordinate equation of the polar coordinate boundary of the reachable region includes: An optimization model is constructed based on the reachable domain area function, and feasible speeds that satisfy the constraints in the optimization model are obtained. The area set corresponding to feasible speeds is obtained based on the reachable domain area function, and the global optimal cruising speed is obtained based on the maximum value in the area set; The polar coordinate boundary of the reachable domain is obtained based on the polar coordinate equation of the globally optimal cruise speed and the polar coordinate boundary of the reachable domain.

[0012] In one embodiment, the step of obtaining the required flight distance corresponding to each feasible airport in the first feasible airport subset includes: For a target feasible airport in the first feasible airport subset, based on the turning radius corresponding to the optimal cruise speed of the launch vehicle and the polar coordinate boundary, the straight flight segment range and the actual turning center angle corresponding to the target feasible airport are obtained. Based on the straight-line flight distance and the actual turning center angle, the required flight distance to the target airport is obtained.

[0013] In one embodiment, the step of obtaining the straight-line flight segment distance and the true turning center angle corresponding to the target feasible airport based on the turning radius corresponding to the optimal cruise speed of the carrier and the polar coordinate boundary includes: Based on the turning radius corresponding to the optimal cruising speed of the carrier and the polar coordinate boundary, the geometric straight-line distance is obtained; The straight-line flight range is obtained based on the geometric straight-line distance and the turning radius corresponding to the optimal cruising speed of the launch vehicle. The true turning center angle is obtained based on the turning radius, geometric straight-line distance, and polar coordinate boundary corresponding to the optimal cruising speed of the launch vehicle.

[0014] In one embodiment, the thrust failure decision and guidance method for a horizontal takeoff and landing two-stage orbital launch vehicle further includes: If the preset airport set and the reachable domain do not intersect, or if the feasibility flag of the separation point dynamics is the second preset value, then the optimal emergency cruise parameters are obtained through the reachable domain area function based on the initial state of the fault occurrence point. Based on the optimal emergency cruise parameters and polar coordinate equations, the polar coordinate boundary of the emergency reachability domain envelope is obtained; If there is an intersection between the preset airport set and the emergency reachability envelope, the intersection will be used as the second feasible airport subset, and the required flight distance for each feasible airport in the second feasible airport subset will be obtained. Based on airport priority levels and required flight distance, the target airport is obtained from the feasible airport subset.

[0015] Furthermore, to achieve the above objectives, this application also proposes a thrust failure decision and guidance device for a horizontal takeoff and landing two-stage-to-orbit launch vehicle. The device includes: a memory, a processor, and a computer program stored in the memory and executable on the processor. The computer program is configured to implement the steps of the thrust failure decision and guidance method for a horizontal takeoff and landing two-stage-to-orbit launch vehicle as described above.

[0016] In addition, to achieve the above objectives, this application also proposes a storage medium, which is a computer-readable storage medium, on which a computer program is stored. When the computer program is executed by a processor, it implements the steps of the thrust failure decision and guidance method for a horizontal takeoff and landing two-stage orbital launch vehicle as described above.

[0017] One or more technical solutions proposed in this application have at least the following technical effects: A convex optimization subproblem is constructed based on the current state vector of the launch vehicle. A sequential convex optimization algorithm is used to iteratively solve the trajectory between the fault point and the separation point, obtaining the feasibility flag of the separation point dynamics. If the feasibility flag of the separation point dynamics is a first preset value, the reachable domain area function is obtained based on the launch vehicle's maximum theoretical longitudinal range under power, unpowered gliding range, extreme turning angle, and minimum turning radius under cruise level flight conditions. The polar coordinate boundary of the reachable domain is obtained based on the reachable domain area function and the polar coordinate equation of the reachable domain's polar coordinate boundary. If there is an intersection between the preset airport set and the reachable domain, the intersection is taken as the first feasible airport subset, and the required flight range corresponding to each feasible airport in the first feasible airport subset is obtained. The target airport is obtained from the feasible airport subset based on the airport priority level and the required flight range. Based on the acquired launch vehicle fault status, an reachability envelope is generated through online trajectory reconstruction and parsing. Based on the reachability envelope, the required flight range corresponding to feasible airports in the preset airport set is determined, and the landing airport, i.e. the target airport, is selected according to the required flight range. Compared with fixed nominal cruise parameters, this effectively expands the reachability envelope range after the fault, realizes online decision-making and guidance after the launch vehicle experiences a sudden thrust loss fault during the ascent phase, and improves the survival probability of the launch vehicle after the fault. Attached Figure Description

[0018] The accompanying drawings, which are incorporated in and form part of this specification, illustrate embodiments consistent with this application and, together with the description, serve to explain the principles of this application.

[0019] To more clearly illustrate the technical solutions in the embodiments of this application or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, for those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0020] Figure 1 This is a flowchart illustrating an embodiment of the thrust failure decision-making and guidance method for a two-stage horizontal takeoff and landing launch vehicle in this application. Figure 2 This is a schematic diagram of the reachable domain in the thrust failure decision and guidance method for a two-stage horizontal takeoff and landing launch vehicle in this application. Figure 3 This is a schematic diagram of the turning angle in the thrust failure decision and guidance method for a horizontal takeoff and landing two-stage orbital launch vehicle provided in Embodiment 2 of this application; Figure 4 This is a schematic diagram of the state quantity curves of an embodiment of the thrust failure decision and guidance method for a horizontal takeoff and landing two-stage orbital launch vehicle according to this application. Figure 5This is a schematic diagram of the equipment structure of the hardware operating environment involved in the thrust failure decision and guidance method for a horizontal take-off and landing two-stage orbital launch vehicle in the embodiments of this application.

[0021] The purpose, features, and advantages of this application will be further explained in conjunction with the embodiments and with reference to the accompanying drawings. Detailed Implementation

[0022] It should be understood that the specific embodiments described herein are merely illustrative of the technical solutions of this application and are not intended to limit this application.

[0023] To better understand the technical solution of this application, a detailed description will be provided below in conjunction with the accompanying drawings and specific implementation methods.

[0024] The main solution of this application embodiment is as follows: Based on the current state vector of the launch vehicle, a trajectory planning problem from the fault point to the expected separation point is constructed. A sequential convex optimization algorithm is used to iteratively solve the convex optimization subproblems corresponding to the trajectory planning problem, obtaining a feasible flag characterizing the dynamic feasibility of the separation point. If the feasible flag is a first preset value, then based on the launch vehicle's maximum theoretical longitudinal range with power, range during unpowered gliding, extreme turning angle, and minimum turning radius in cruise level flight, an area function of the reachable domain for cruise parameters is constructed, and the target cruise parameters are determined with the goal of maximizing the area of ​​the reachable domain. Based on the target cruise parameters, a sequential convex optimization algorithm is used to reconstruct the transition trajectory from the separation point to the cruise start point, and based on the cruise start point as a reference, the reachable domain after separation is analytically generated based on the polar coordinate boundary equation of the reachable domain. If there is an intersection between the preset airport set and the reachable domain, then the intersection is taken as the first feasible airport subset, and the mission mode is determined to be landing after separation. If the feasible flag is not the first preset value, or there is no intersection between the preset airport set and the reachable domain, then the inter-level separation is cancelled, and emergency landing trajectory reconstruction and emergency reachable domain generation are performed based on the fault point status; if there is an intersection between the preset airport set and the emergency reachable domain, then the intersection is taken as the second feasible airport subset, and the mission mode is determined to be direct emergency landing; according to the determined mission mode, the required flight range corresponding to each feasible airport in the first feasible airport subset or the second feasible airport subset is obtained; based on the airport priority level and the required flight range, the target airport is determined from the first feasible airport subset or the second feasible airport subset.

[0025] In this embodiment, for ease of description, the thrust failure decision and guidance device of a two-stage horizontal takeoff and landing launch vehicle will be used as the main execution subject for the following description.

[0026] HTHL-TSTO launch vehicles, with their potential for full reusability and low launch costs, have become a core direction for the development of future space transportation systems. However, the propulsion systems of this type of launch vehicle (such as deep air-cooled combined cycle engines) are highly complex, facing a more severe risk of thrust loss during the climb phase than traditional rocket propulsion. Once thrust loss occurs, the launch vehicle's flight state will rapidly deviate from the nominal flight profile, easily inducing aerodynamic load overruns and threatening the structural integrity of the airframe, resulting in an extremely narrow window for emergency response.

[0027] Research on decision-making and guidance for launch vehicle ascent phase failures has largely focused on traditional vertical launch rockets, primarily targeting vacuum or near-vacuum flight environments, making it difficult to adapt to the complex flight characteristics of the HTHL-TSTO launch vehicle. HTHL-TSTO requires prolonged maneuvering flight within the atmosphere, encompassing multiple phases including climb, cruise, and unpowered gliding. Following a thrust loss failure, it is necessary not only to reconstruct the continuous flight trajectory but also to simultaneously complete multiple discrete mission decisions, such as stage separation determination, degradation mode switching, and alternate landing airport selection. Current decision-making and guidance for continuous-discrete hybrid systems, primarily relying on pre-set abort modes such as the Space Shuttle's RTLS (Return-to-Launch-Site) and TAL (Trans-Atlantic Landing) modes, lacks sufficient adaptive capability to cope with sudden, unpredictable failures.

[0028] Therefore, how to achieve online decision-making and guidance after a sudden thrust loss failure during the ascent phase of a launch vehicle, and improve the survival probability of the launch vehicle after the failure, is an urgent problem to be solved.

[0029] This application provides a solution that, based on the acquired launch vehicle fault state, generates an reachability envelope through online trajectory reconstruction and analysis, determines the required flight range corresponding to feasible airports in a preset airport set based on the reachability envelope, and selects the landing airport, i.e., the target airport, according to the required flight range. Compared with fixed nominal cruise parameters, this effectively expands the reachability envelope range after a fault, realizes online decision-making and guidance after a sudden thrust loss fault during the launch vehicle's ascent phase, and improves the survival probability of the launch vehicle after a fault.

[0030] It should be noted that the executing entity in this embodiment can be a computing service device with data processing, network communication, and program execution functions, such as a tablet computer, personal computer, or mobile phone, or an electronic device capable of performing the above functions, such as a thrust fault decision and guidance device for a horizontal takeoff and landing two-stage orbit insertion vehicle. The following description uses a thrust fault decision and guidance device for a horizontal takeoff and landing two-stage orbit insertion vehicle as an example to illustrate this embodiment and the subsequent embodiments.

[0031] Based on this, embodiments of this application provide a thrust failure decision-making and guidance method for a horizontal takeoff and landing two-stage-to-orbit launch vehicle, referring to... Figure 1 , Figure 1 This is a flowchart illustrating the first embodiment of the thrust failure decision-making and guidance method for a two-stage orbital launch vehicle with horizontal takeoff and landing according to this application.

[0032] In this embodiment, the thrust failure decision and guidance method for a horizontal takeoff and landing two-stage orbital launch vehicle includes steps S110~S180: Step S110: Based on the current state vector of the launch vehicle, construct a trajectory planning problem from the fault point to the expected separation point, and use a sequential convex optimization algorithm to iteratively solve the convex optimization subproblems corresponding to the trajectory planning problem to obtain a feasible flag bit characterizing the dynamic feasibility of the separation point. Step S120: If the feasible flag is a first preset value, then based on the maximum theoretical longitudinal range with power, the range of the unpowered gliding segment, the extreme turning angle and the minimum turning radius of the carrier in cruise level flight state, construct the reachable domain area function for the cruise parameters, and determine the target cruise parameters with the goal of maximizing the reachable domain area. Step S130: Based on the target cruise parameters, the transition trajectory from the separation point to the cruise start point is reconstructed using a sequential convex optimization algorithm, and the reachable domain after separation is generated analytically based on the reachable domain polar coordinate boundary equation, with the cruise start point as the reference. Step S140: If there is an intersection between the preset airport set and the reachable domain, then the intersection is taken as the first feasible airport subset, and the mission mode is determined to be landing after separation. Step S150: If the feasible flag is not the first preset value, or there is no intersection between the preset airport set and the reachable domain, then cancel the inter-level separation, and perform emergency landing trajectory reconstruction and emergency reachable domain generation based on the fault point status. Step S160: If there is an intersection between the preset airport set and the emergency reachability domain, then the intersection is taken as the second feasible airport subset, and the mission mode is determined to be direct emergency landing; Step S170: Based on the determined mission mode, obtain the required flight range corresponding to each feasible airport in the first feasible airport subset or the second feasible airport subset; Step S180: Based on the airport priority level and the required flight range, determine the target airport from the first subset of feasible airports or the second subset of feasible airports.

[0033] In this embodiment, after a sudden thrust loss failure during the ascent phase of the launch vehicle, the current state vector of the launch vehicle is obtained. Based on the current state vector of the launch vehicle, a trajectory planning problem from the failure point to the expected separation point is constructed. The convex optimization subproblem corresponding to the trajectory planning problem is solved iteratively using a sequential convex optimization algorithm to obtain a feasible flag characterizing the dynamic feasibility of the separation point. Under the premise of considering dynamics and engine working corridor constraints, the flight trajectory from the failure point to the expected separation point is solved using sequential convex optimization. The dynamic feasibility of the separation point is evaluated based on the convergence of the algorithm.

[0034] The feasible flag includes a first preset value and a second preset value.

[0035] In one feasible implementation, step S110 may include steps A111 to A114: Step A111: Obtain the current state vector of the launch vehicle, and construct a trajectory planning problem from the fault point to the separation point based on the current state vector. The current state vector includes the launch vehicle's distance from the Earth's center, velocity, local ballistic inclination angle, and mass. Step A112: The sequential convex optimization algorithm is used to iteratively solve the convex optimization subproblem corresponding to the trajectory planning problem, and the trust region radius corresponding to the state variable, the trust region radius corresponding to the control variable, and the trust region radius corresponding to the terminal time are adaptively updated during the iterative solution process. Step A113: If the sequential convex optimization algorithm converges during the solution iteration, then the feasible flag of the separation point dynamics is determined to be the first preset value. Step A114: If the sequential convex optimization algorithm does not converge when the iteration is completed, then the feasible flag of the separation point dynamics is determined to be the second preset value.

[0036] Get the current state vector And construct a convex optimization subproblem based on the current state vector. The convex optimization subproblem constructed based on the current state vector for the k-th iteration is as follows: ; Wherein, min and st refer to the state and control variables that minimize the performance index under all constraints specified by st.

[0037] N refers to the normalized time domain being uniformly discretized into N segments, i.e., containing nodes 0, 1, ..., N.

[0038] w m , , w u , w ηThese are the weighting coefficients for terminal quality performance indicators, dynamic slack variable penalties, control law penalties, and deviation from the reference trajectory penalties, respectively. t f Let the time be the final time of the current convex optimization subproblem. The terminal time is obtained after solving the k-th convex optimization subproblem; "free" means that there are no equality constraints on the terminal quality. x i Let be the state variable at the i-th discrete time point; x i-1 Let be the state variable at the (i-1)th discrete time point; The state quantity at the i-th discrete time point is obtained after solving the k-th convex optimization subproblem; x i - It refers to the deviation between the state quantity at the i-th discrete time point and the state quantity at the k-th time point in the current subproblem; u i Let be the control quantity at the i-th discrete time point; The control quantity at the i-th discrete time point is obtained after solving the k-th convex optimization subproblem; u i - This represents the deviation between the control quantity at the i-th discrete time point and the control quantity at the k-th time point in the current subproblem. The mass of the carrier at the terminal node. These are dynamic relaxation variables used to suppress the initial "pseudo-infeasibility" phenomenon. To control the rate of change, it is defined as ; They are respectively the first one i The first node j One state variable and control variable; The launch vehicle's geocentric distance, velocity, local ballistic inclination, and mass, where superscript values ​​are indicated. k The term represents the first item. k The reference value for the next iteration; These are the trust region radii of the state variables, the trust region radii of the control variables, and the trust region radii of the terminal time, respectively. These represent the launch vehicle's geocentric distance, velocity, local ballistic inclination angle, and mass, respectively. The item with subscript 0 indicates the initial state, and the subscript... f The item represents the terminal state; This refers to the dry weight of the launch vehicle.

[0039] These are the angle of attack and throttle opening of the launch vehicle, respectively. This is the normalized time step, with a value of 1 / N; It is a fourth-order Runge-Kutta linearized recurrence term derived based on the previous generation reference trajectory; The process constraints are respectively in the first... k The values ​​taken on the next iteration trajectory and the Jacobian matrix of process constraints on the state and control variables.

[0040] The first i Angle of attack and throttle opening at the node; These are the upper and lower limits of the angle of attack, respectively; These represent the upper and lower limits of the throttle opening, respectively.

[0041] A sequential convex optimization algorithm is used to iteratively solve the trajectory between the fault point and the separation point in the convex optimization subproblem. During the iterative solution process, the trust region radii of the state variables, the control variables, and the terminal time in the convex optimization subproblem are adaptively updated. The formula for the adaptive update is: ; in, , This is a trust region contraction factor used to reduce oscillations. This is an amplification factor used to accelerate convergence. >1.

[0042] During the iterative solution process, it is determined in real time whether the sequential convex optimization algorithm converges. If the sequential convex optimization algorithm converges, the feasibility flag F of the breakaway point dynamics is determined. SCP The first preset value is, for example, the first preset value is 1, i.e., F SCP =1.

[0043] If the sequential convex optimization algorithm fails to converge when the iteration is completed, the feasibility flag of the breakaway point dynamics is determined to be the second preset value, where the first preset value is 0, i.e., F SCP =0.

[0044] In this embodiment, if the feasible flag of the separation point dynamics is the first preset value, i.e., F SCP =1, then based on the maximum theoretical longitudinal range with power, the range of the unpowered gliding segment, the extreme turning angle, and the minimum turning radius of the launch vehicle in cruise level flight, the reachable domain area function is obtained; among them, the reachable domain area function is determined for the two cases of intersection and non-intersection, depending on whether the reachable domain area intersects with the negative x-axis.

[0045] In one feasible implementation, step S120 may include step B111: Step B111: If the reachable domain does not intersect with the negative x-axis, then construct the area function of the reachable domain for the cruise parameters based on the turning drag multiplication factor, the maximum theoretical longitudinal range with power, the range of the unpowered gliding segment, the extreme turning angle, and the minimum turning radius.

[0046] In one feasible implementation, when the reachable domain intersects with the negative x-axis, the limiting turning angle includes the turning angle at the first intersection of the outer boundary and the negative x-axis and the turning angle at the second intersection of the inner boundary and the negative x-axis. Step S110 may include steps B112~B114: Step B112: Based on the turning angle at the first intersection, the minimum turning radius, the turning resistance multiplication factor, the maximum theoretical longitudinal range with power, and the range of the unpowered gliding segment, obtain the outer boundary area function; Step B113: Based on the turning angle at the second intersection point, the minimum turning radius, and the unpowered gliding distance, obtain the inner boundary area function; Step B114: Based on the outer boundary area function and the inner boundary area function, construct the reachability domain area function for the cruise parameters.

[0047] In this embodiment, if the reachable area does not intersect the negative x-axis, then based on the turning drag multiplication factor, the maximum theoretical longitudinal range with power, the range of the unpowered gliding segment, the extreme turning angle, and the minimum turning radius, a reachable area function with respect to the cruise parameters is constructed. The formula for the reachable area function is: ; If the reachable domain area intersects the negative x-axis, based on the turning angle at the first intersection point, the minimum turning radius, the turning drag multiplication factor, the maximum theoretical longitudinal range with power, and the range of the unpowered gliding segment, obtain the outer boundary area function A2; based on the turning angle at the second intersection point, the minimum turning radius, and the range of the unpowered gliding segment, obtain the inner boundary area function A3; based on the outer boundary area function and the inner boundary area function, construct the reachable domain area function with respect to the cruise parameters. The formula for the reachable domain area function is: ; ; ; ; ; in, A ( V cr ) is the area function of the reachable region. V cr The target cruising speed; A 2 represents the area of ​​the outer boundary. A3 represents the internal area; s long This represents the maximum theoretical powered longitudinal range of the launch vehicle in cruise level flight. s glide The range of the unpowered gliding segment; S aug ( δ () refers to the steering angle δ Then, the total effective straight-line extension distance of the launch vehicle, S aug The value of (δ) is the sum of the remaining cruise range with power and the gliding range without power; like Figure 3 As shown, δ max This is the extreme turning angle. δ max2 The turning angle at the first intersection. δ max3 Let be the turning angle at the second intersection point, where can be solved... Obtain the turning angle at the first intersection δ max2 ; R turn This is the minimum turning radius of the vehicle at a given cruise speed and maximum permissible roll angle. k D As a factor that multiplies turning resistance, k D It is the ratio of turning resistance to straight-line cruising resistance. k D The value is the ratio of the cornering angle to the induced drag. The function. γ Vmax This is the maximum tilt angle.

[0048] When obtaining the reachable domain area function, the goal is to maximize the reachable domain area. The target cruise parameters include the globally optimal cruise speed. Based on the reachable domain area function and the polar equation of the polar coordinate boundary of the reachable domain, the polar coordinate boundary of the reachable domain is obtained. Specifically, an optimization model is constructed based on the reachable domain area function. Multiple feasible speeds are obtained according to the constraints of the optimization model. The globally optimal cruise speed is obtained based on the feasible speeds.

[0049] In one feasible implementation, step S120 may include steps S121-S122: Step S121: Construct an optimization model based on the reachable domain area function and obtain the feasible speed that satisfies the constraints in the optimization model; Step S122: Obtain the area set corresponding to the feasible speed based on the reachable domain area function, and obtain the target cruise parameters based on the maximum value in the area set.

[0050] In this embodiment, an optimization model is constructed based on the reachable domain area function. The formula for the optimization model is: ; Where max,st refers to the cruising speed that maximizes the reachable domain area under all constraints specified by st.

[0051] For optimal speed, For the corresponding height; V min ,V max These are the minimum and maximum cruising speeds, respectively. h cr This refers to the balanced cruise altitude at the corresponding cruise speed. cruise speed V cr The corresponding minimum and maximum heights of the engine working corridor; Fuel required to transition from the current speed to cruising speed; Total available fuel.

[0052] h ( V cr ( ) is the cruising speed V cr The optimal cruising altitude is below; T avail ( V cr , h cr ) is at cruising speed V cr Cruise altitude h cr Available thrust below; D ( V cr , h cr At cruising speed V cr Cruise altitude h cr The resistance below.

[0053] It can construct an optimization model, traverse discrete velocity nodes, and lock the optimal velocity that maximizes the area. and corresponding height , The feasible speed that satisfies the constraints of the optimization model is discretized into a subset containing...M An ordered set of nodes The area set corresponding to the reachable region is calculated based on the formula of the area function of the reachable region. .

[0054] The globally optimal cruising speed is obtained by finding the maximum value in the area set. Locking onto the node containing the maximum value in the discrete sequence: ; After obtaining the target cruise parameters, the transition trajectory from the separation point to the cruise start point is reconstructed based on the target cruise parameters using a sequential convex optimization algorithm. That is, the sequential convex optimization algorithm is called to reconstruct the transition trajectory from the separation point to the cruise start point, and the reachable domain after separation is generated analytically based on the reachable domain polar coordinate boundary equation using the cruise start point as a reference.

[0055] In one possible implementation, step S130 may include step S131: Step S131: Based on the target cruise parameters and the polar coordinate equation of the polar coordinate boundary of the reachable domain, obtain the polar coordinate boundary of the reachable domain to generate the separated reachable domain.

[0056] Based on the target cruise parameters and the polar coordinate equation of the reachable region's polar coordinate boundary, the polar coordinate boundary of the reachable region is obtained. The polar coordinate boundary includes the line-of-sight distance. and the angle of view ,like Figure 2 As shown.

[0057] by For example, an auxiliary length is introduced. The polar coordinate equation is: ; ; in, Line-of-sight distance; This refers to the angle of deflection of the line of sight. d For auxiliary length, ; Substituting the globally optimal cruising speed into the polar coordinate equation, we iterate through the turning angles. parse and generate reachable domain The polar coordinate boundary.

[0058] After obtaining the reachable domain after separation, if there is an intersection between the preset airport set and the reachable domain, the intersection is taken as the first feasible airport subset, and the mission mode is determined to be landing after separation, i.e., the preset airport set is determined. With reachable domain Does an intersection exist? If an intersection exists, then... Therefore, the mission mode is determined to be separation and landing, i.e. At this point, the intersection is taken as the first feasible airport subset, that is, the feasible airport subset within the reachable domain is extracted. And obtain the required flight range corresponding to each feasible airport in the first feasible airport subset.

[0059] In this embodiment, if the preset airport set and the reachability domain do not intersect, or if the feasibility flag of the separation point dynamics is a second preset value, then the inter-level separation is cancelled, and emergency landing trajectory reconstruction and emergency reachability domain generation are performed based on the fault point status.

[0060] Specifically, based on the initial state of the fault location, the optimal emergency cruise parameters are obtained through the reachable domain area function. The Sequential Convex Optimization (SCP) algorithm is then used to reconstruct the emergency transition trajectory from the fault location to the cruise start point. The process for obtaining the optimal emergency cruise parameters is similar to that for obtaining the global optimal cruise speed, and will not be elaborated further here. Based on the optimal emergency cruise parameters and the polar coordinate equation, the polar coordinate boundary of the emergency reachable domain envelope is obtained. The process for obtaining the polar coordinate boundary of the emergency reachable domain envelope is similar to that for obtaining the polar coordinate boundary of the reachable domain, and will not be elaborated further here.

[0061] If there is an intersection between the preset airport set and the emergency reachability domain, the intersection is taken as the second feasible airport subset, and the mission mode is determined to be direct emergency landing. Specifically, if there is an intersection between the preset airport set and the emergency reachability domain envelope, the intersection is taken as the second feasible airport subset, and the required flight distances corresponding to each feasible airport in the second feasible airport subset are obtained. The process of obtaining the required flight distances corresponding to each feasible airport in the second feasible airport subset is similar to that of obtaining the required flight distances corresponding to each feasible airport in the first feasible airport subset, and will not be described in detail here.

[0062] Next, based on the determined mission mode, the required flight distance for each feasible airport in the first feasible airport subset or the second feasible airport subset is obtained.

[0063] In one feasible implementation, step S140 may include steps C110~C120: Step C110: For the target feasible airport in the first feasible airport subset or the second feasible airport subset, based on the turning radius corresponding to the optimal cruise speed of the carrier and the polar coordinate boundary, obtain the straight flight segment range and the actual turning center angle corresponding to the target feasible airport. Step C120: Based on the straight flight segment distance and the actual turning center angle, obtain the required flight distance to the target airport.

[0064] In one feasible implementation, step C110 may include steps C111 to C113: Step C111: Based on the turning radius corresponding to the optimal cruising speed of the carrier and the polar coordinate boundary, obtain the geometric straight-line distance; Step C112: Based on the geometric straight-line distance and the turning radius corresponding to the optimal cruising speed of the launch vehicle, obtain the straight-line flight segment range; Step C113: Based on the turning radius, geometric straight-line distance, and polar coordinate boundary corresponding to the optimal cruising speed of the carrier, obtain the true turning center angle.

[0065] In this embodiment, for a target feasible airport in the first feasible airport subset or the second feasible airport subset, the straight flight segment range and the actual turning center angle corresponding to the target feasible airport are obtained based on the turning radius corresponding to the optimal cruise speed of the carrier and the polar coordinate boundary.

[0066] Specifically, the geometric straight-line distance is obtained based on the turning radius corresponding to the carrier's optimal cruising speed and the polar coordinate boundary. Geometric straight-line distance The formula is: ; Based on the geometric straight-line distance and the turning radius corresponding to the optimal cruising speed of the launch vehicle, the straight-line flight segment range is obtained. The formula for the straight-line flight segment range is: ; Based on the turning radius, geometric straight-line distance, and polar coordinate boundary corresponding to the optimal cruising speed of the launch vehicle, the true turning central angle is obtained. The formula for the true turning central angle is: ; After obtaining the true turning center angle, the required flight distance to the target airport is obtained based on the straight flight segment distance and the true turning center angle. The formula for the required flight distance is: ; in, For the straight-line flight segment, This is the turning radius corresponding to the vehicle's optimal cruising speed. This represents the geometric straight-line distance from the center of the turning circle to the target waypoint. For the actual turning center angle, The required flight distance to the target airport.

[0067] After obtaining the required flight range, the target airport is selected from the subset of feasible airports based on the airport priority level and the required flight range, i.e., according to the preset airport priority level. and the calculated The optimal airport number is determined using a lexicographical order strategy. , , in, argLexMin represents the airport number with the highest priority and shortest required flight distance, obtained by lexicographical order strategy; argLexMin represents the lexicographical minimization operator, which strictly follows the comparison logic of priority level as the primary factor and the numerical value of required flight distance as the secondary factor. To be assigned to the i A Boolean selection command for alternate airports, with a value of 1 indicating selection; The total number of feasible alternate airports covered within the reachable region; P level,j Let j be the airport priority level of the target airport. s total,j Let be the required flight distance for the j-th target airport.

[0068] It should be noted that the closed-loop tracking guidance law for the launch vehicle's ascent, transition, and cruise phases is as follows: ; After entering the unpowered gliding phase, switch to tracking altitude-dynamic pressure profile, with the guidance law as follows: ; in, α、 T γ V These are angle of attack, throttle opening, and roll angle, respectively. These are the desired angle of attack, desired throttle opening, desired altitude, desired speed, and desired dynamic pressure during the unpowered gliding phase, given by the reference trajectory. For reference, the rate of change of height; These are the vehicle's current actual altitude, actual altitude change rate, actual speed, actual heading angle, and actual dynamic pressure, respectively. This is the geometric line-of-sight angle between the current position of the launch vehicle and the target landing airport. k ph , k dh , k pV , k pΨ , k p These are, respectively, altitude proportional gain, altitude differential gain, velocity feedback gain, heading line-of-sight angle feedback gain, and dynamic pressure feedback gain.

[0069] It should be noted that the hardware conditions, inter-stage separation conditions, and alternate airport parameters on which the embodiments of this application are based are set as follows: (1) Hardware requirements: Intel Core i9-13900K processor.

[0070] (2) Interstage separation conditions: The nominal separation height is set at 25 km, and the separation speed is... Ma = 5.

[0071] (3) Alternate Airport Coordinates and Priority: A set of 16 airports is defined, including the nominal departure airport, the nominal primary landing airport, and alternate airports around the route. The coordinates of the nominal departure airport (AP01) are as follows: It was assigned the second-highest priority, 2; the nominal primary landing airport (AP02) coordinates are... The airport was assigned the highest priority of 1 (the smaller the number, the higher the priority); the other 14 alternate airports were assigned priority of 3 or 4 respectively.

[0072] Table 1. Alternate Airport Coordinates and Priorities

[0073] Example 1: like Figure 4 As shown, input fault parameters: Mach number Thrust loss amplitude .

[0074] The launch vehicle's energy is determined to meet the conditions for reaching the predetermined separation point, and a flag indicating the dynamic feasibility of the separation point is output. Optimal separation cruise parameters are obtained based on area maximization optimization. A sequential convex optimization algorithm is invoked to reconstruct the transition trajectory from the separation point to the cruise start point online, and the separation reachability domain envelope is generated parsed. After joint verification, the envelope... The highest priority nominal primary landing airport AP02 is covered, and the degradation condition has not been triggered. Output the mission mode. The total calculation time for this decision and trajectory reconstruction was 0.37 seconds. After receiving the command, the bottom-level closed-loop guidance system guided the launch vehicle to climb to 25 km to complete interstage separation and was guided to the approach and landing handover point of AP02.

[0075] Example 2: like Figure 4 As shown, input fault parameters: Mach number Thrust loss amplitude .

[0076] The algorithm failed to converge and output the result because the remaining thrust did not meet the dynamic conditions for reaching the predetermined separation point. This triggers the dynamic failure early degradation logic. Based on the state at the fault location, emergency cruise parameters are optimized. A sequential convex optimization algorithm is used to reconstruct the emergency transition trajectory from the fault location to the cruise start point, and the emergency reachability domain envelope is generated. After baseline verification, the envelope... The primary landing airport AP02 is not covered, but the second-highest priority departure airport AP01 is covered. The system outputs an emergency return mode. The total calculation time for this decision and trajectory reconstruction was 0.64 seconds. After receiving the command, the bottom-level closed-loop guidance aborted the separation climb and guided the launch vehicle to perform a turn maneuver at a 30° tilt angle, returning to the AP01 approach and landing handover point.

[0077] Example 3: like Figure 4 As shown, input fault parameters: Mach number Thrust loss amplitude .

[0078] The vehicle is determined to have used its initial mechanical energy to meet the constraint of gliding to the predetermined separation point, and the output is... The system performs cruise parameter optimization and calls a sequential convex optimization algorithm to reconstruct the transition segment trajectory, generating the separated reachable domain envelope. After joint verification, the envelope The highest priority nominal primary landing airport AP02 is covered, and the degradation condition has not been triggered. Output the mission mode. The total calculation time for this decision and trajectory reconstruction was 0.55 seconds. After receiving the command, the bottom-level closed-loop guidance system controlled the launch vehicle to complete inter-stage separation and guided it to the approach and landing handover point of AP02.

[0079] This embodiment provides a thrust failure decision-making and guidance method for a two-stage takeoff and landing (TSL) launch vehicle. It constructs a convex optimization subproblem based on the launch vehicle's current state vector, and uses a sequential convex optimization algorithm to iteratively solve the trajectory between the failure point and the separation point, obtaining a feasible flag for the separation point dynamics. If the feasible flag for the separation point dynamics is a first preset value, the reachable domain area function is obtained based on the launch vehicle's maximum theoretical powered longitudinal range, unpowered gliding range, extreme turning angle, and minimum turning radius during cruise level flight. The polar coordinate boundary of the reachable domain is obtained based on the reachable domain area function and the polar coordinate equation of the reachable domain's polar coordinate boundary. If there is an intersection between the preset airport set and the reachable domain, the intersection is taken as the first feasible airport subset, and the required flight range corresponding to each feasible airport in the first feasible airport subset is obtained. The target airport is obtained from the feasible airport subset based on the airport priority level and the required flight range. Based on the acquired launch vehicle fault status, an reachability envelope is generated through online trajectory reconstruction and parsing. Based on the reachability envelope, the required flight range corresponding to feasible airports in the preset airport set is determined, and the landing airport, i.e. the target airport, is selected according to the required flight range. Compared with fixed nominal cruise parameters, this effectively expands the reachability envelope range after the fault, realizes online decision-making and guidance after the launch vehicle experiences a sudden thrust loss fault during the ascent phase, and improves the survival probability of the launch vehicle after the fault.

[0080] It should be noted that the above examples are only for understanding this application and do not constitute a limitation on the thrust failure decision and guidance method for horizontal take-off and landing two-stage orbital launch vehicles of this application. Any simple modifications based on this technical concept are within the protection scope of this application.

[0081] This application provides a thrust failure decision and guidance device for a horizontal takeoff and landing (FTLR) two-stage-to-orbit (LSO) launch vehicle. The FTLR two-stage-to-orbit (LSO) launch vehicle thrust failure decision and guidance device includes: at least one processor; and a memory communicatively connected to the at least one processor; wherein the memory stores instructions executable by the at least one processor, and the instructions are executed by the at least one processor to enable the at least one processor to execute the FTLR two-stage-to-orbit (LSO) launch vehicle thrust failure decision and guidance method in the above embodiment 1.

[0082] The following is for reference. Figure 5This document illustrates a structural schematic diagram of a thrust failure decision and guidance device suitable for implementing the embodiments of this application for a two-stage takeoff and landing (TSL) launch vehicle. The thrust failure decision and guidance device for a TSL launch vehicle in the embodiments of this application may include, but is not limited to, mobile terminals such as mobile phones, laptops, digital radio receivers, PDAs (Personal Digital Assistants), PADs (Portable Application Descriptions), PMPs (Portable Media Players), and in-vehicle terminals (e.g., in-vehicle navigation terminals), as well as fixed terminals such as digital TVs and desktop computers. Figure 5 The thrust failure decision and guidance device for a horizontal takeoff and landing two-stage orbital launch vehicle shown is merely an example and should not impose any limitations on the functionality and scope of use of the embodiments of this application.

[0083] like Figure 5 As shown, the thrust failure decision and guidance device for a horizontal takeoff and landing (LTO) vehicle may include a processing unit 1001 (e.g., a central processing unit, a graphics processing unit, etc.), which can perform various appropriate actions and processes according to a program stored in a read-only memory 1002 or a program loaded from a storage device 1003 into a random access memory 1004. The random access memory 1004 also stores various programs and data required for the operation of the LTO thrust failure decision and guidance device. The processing unit 1001, the read-only memory 1002, and the random access memory 1004 are interconnected via a bus 1005. An input / output interface 1006 is also connected to the bus. Typically, the following systems can be connected to the input / output interface 1006: input devices 1007 including, for example, a touchscreen, touchpad, keyboard, mouse, image sensor, microphone, accelerometer, gyroscope, etc.; output devices 1008 including, for example, a liquid crystal display (LCD), speaker, vibrator, etc.; storage devices 1003 including, for example, magnetic tape, hard disk, etc.; and communication devices 1009. Communication device 1009 allows the Horizontal Take-Off and Landing Two-Stage Orbiter (LTPO) thrust fault decision and guidance system to exchange data wirelessly or via wired communication with other devices. Although the figure shows an LTPO thrust fault decision and guidance system with various systems, it should be understood that implementation or possession of all the systems shown is not required. More or fewer systems may be implemented alternatively.

[0084] Specifically, according to the embodiments disclosed in this application, the processes described above with reference to the flowcharts can be implemented as computer software programs. For example, embodiments disclosed in this application include a computer program product comprising a computer program carried on a computer-readable medium, the computer program containing program code for performing the methods shown in the flowcharts. In such embodiments, the computer program can be downloaded and installed from a network via a communication device, or installed from storage device 1003, or installed from read-only memory 1002. When the computer program is executed by processing device 1001, it performs the functions defined in the methods of the embodiments disclosed in this application.

[0085] The thrust failure decision and guidance device for a horizontal takeoff and landing (FTLR) two-stage-to-orbit launch vehicle provided in this application adopts the thrust failure decision and guidance method for FTLR two-stage-to-orbit launch vehicles described in the above embodiments. It solves the technical problem of how to achieve online decision-making and guidance after a sudden thrust loss failure during the launch vehicle's ascent phase, thereby improving the survival probability of the launch vehicle after a failure. Compared with the prior art, the beneficial effects of the thrust failure decision and guidance device for FTLR two-stage-to-orbit launch vehicles provided in this application are the same as those of the thrust failure decision and guidance method for FTLR two-stage-to-orbit launch vehicles provided in the above embodiments. Furthermore, other technical features of this thrust failure decision and guidance device are the same as those disclosed in the previous embodiment method, and will not be repeated here.

[0086] It should be understood that the various parts disclosed in this application can be implemented using hardware, software, firmware, or a combination thereof. In the description of the above embodiments, specific features, structures, materials, or characteristics can be combined in any suitable manner in one or more embodiments or examples.

[0087] The above description is merely a specific embodiment of this application, but the scope of protection of this application 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 this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

[0088] This application provides a computer-readable storage medium having computer-readable program instructions (i.e., a computer program) stored thereon, the computer-readable program instructions being used to execute the thrust failure decision and guidance method for a two-stage horizontal takeoff and landing launch vehicle in the above embodiments.

[0089] The computer-readable storage medium provided in this application may be, for example, a USB flash drive, but is not limited to, electrical, magnetic, optical, electromagnetic, infrared, or semiconductor systems or devices, or any combination thereof. More specific examples of computer-readable storage media may include, but are not limited to: electrical connections having one or more wires, portable computer disks, hard disks, random access memory (RAM), read-only memory (ROM), erasable programmable read-only memory (EPROM or flash memory), optical fiber, portable compact disk read-only memory (CD-ROM), optical storage devices, magnetic storage devices, or any suitable combination thereof. In this embodiment, the computer-readable storage medium may be any tangible medium containing or storing a program that can be used by or in conjunction with an instruction execution system or device. The program code contained on the computer-readable storage medium may be transmitted using any suitable medium, including but not limited to: wires, optical cables, RF (Radio Frequency), etc., or any suitable combination thereof.

[0090] The aforementioned computer-readable storage medium may be included in the thrust failure decision and guidance device of a horizontal takeoff and landing two-stage orbiter; or it may exist independently and not be assembled into the thrust failure decision and guidance device of a horizontal takeoff and landing two-stage orbiter.

[0091] The aforementioned computer-readable storage medium carries one or more programs. When these programs are executed by the thrust fault decision and guidance device of the horizontal takeoff and landing two-stage orbital launch vehicle, the thrust fault decision and guidance device performs the following actions: It constructs a convex optimization sub-problem based on the launch vehicle's current state vector; it iteratively solves the convex optimization sub-problem using a sequential convex optimization algorithm to obtain the feasible flag of the separation point dynamics; if the feasible flag of the separation point dynamics is a first preset value, then based on the launch vehicle's current state vector... The maximum theoretical longitudinal range under powered flight, the range under unpowered gliding flight, the extreme turning angle, and the minimum turning radius under level flight conditions are used to obtain the area function of the reachable region. Based on the area function of the reachable region and the polar equation of the polar boundary of the reachable region, the polar boundary of the reachable region is obtained. If there is an intersection between the preset airport set and the reachable region, the intersection is taken as the first feasible airport subset, and the required flight range corresponding to each feasible airport in the first feasible airport subset is obtained. Based on the airport priority level and the required flight range, the target airport is obtained from the feasible airport subset.

[0092] Computer program code for performing the operations of this application can be written in one or more programming languages ​​or a combination thereof, including object-oriented programming languages ​​such as Java, Smalltalk, and C++, and conventional procedural programming languages ​​such as the "C" language or similar programming languages. The program code can be executed entirely on the user's computer, partially on the user's computer, as a standalone software package, partially on the user's computer and partially on a remote computer, or entirely on a remote computer or server. In cases involving remote computers, the remote computer can be connected to the user's computer via any type of network—including a Local Area Network (LAN) or a Wide Area Network (WAN)—or can be connected to an external computer (e.g., via the Internet using an Internet service provider).

[0093] The flowcharts and block diagrams in the accompanying drawings illustrate the architecture, functionality, and operation of possible implementations of systems, methods, and computer program products according to various embodiments of this application. In this regard, each block in a flowchart or block diagram may represent a module, segment, or portion of code containing one or more executable instructions for implementing a specified logical function. It should also be noted that in some alternative implementations, the functions indicated in the blocks may occur in a different order than those indicated in the drawings. For example, two consecutively indicated blocks may actually be executed substantially in parallel, and they may sometimes be executed in reverse order, depending on the functions involved. It should also be noted that each block in the block diagrams and / or flowcharts, and combinations of blocks in the block diagrams and / or flowcharts, can be implemented using a dedicated hardware-based system that performs the specified function or operation, or using a combination of dedicated hardware and computer instructions.

[0094] The modules described in the embodiments of this application can be implemented in software or hardware. The names of the modules do not necessarily limit the functionality of the unit itself.

[0095] The readable storage medium provided in this application is a computer-readable storage medium that stores computer-readable program instructions (i.e., a computer program) for executing the above-described thrust failure decision-making and guidance method for a two-stage horizontal takeoff and landing launch vehicle. This method solves the technical problem of how to achieve online decision-making and guidance after a sudden thrust loss failure during the launch vehicle's ascent phase, thereby improving the launch vehicle's survival probability after a failure. Compared with the prior art, the beneficial effects of the computer-readable storage medium provided in this application are the same as those of the thrust failure decision-making and guidance method for a two-stage horizontal takeoff and landing launch vehicle provided in the above embodiments, and will not be repeated here.

[0096] This application also provides a computer program product, including a computer program that, when executed by a processor, implements the steps of the above-described thrust failure decision and guidance method for a horizontal takeoff and landing two-stage-to-orbit launch vehicle.

[0097] The computer program product provided in this application can solve the technical problem of how to realize online decision-making and guidance after a sudden thrust loss failure during the ascent phase of a launch vehicle, thereby improving the survival probability of the launch vehicle after the failure. Compared with the prior art, the beneficial effects of the computer program product provided in this application are the same as those of the thrust failure decision-making and guidance method for horizontal takeoff and landing two-stage-to-orbit launch vehicles provided in the above embodiments, and will not be repeated here.

[0098] The above description is only a part of the embodiments of this application and does not limit the patent scope of this application. All equivalent structural transformations made under the technical concept of this application and using the contents of the specification and drawings of this application, or direct / indirect applications in other related technical fields, are included in the patent protection scope of this application.

Claims

1. A method for thrust failure decision-making and guidance of a horizontal takeoff and landing two-stage-to-orbit launch vehicle, characterized in that, The method includes: Based on the current state vector of the launch vehicle, a trajectory planning problem from the fault point to the expected separation point is constructed. The convex optimization subproblem corresponding to the trajectory planning problem is solved iteratively using a sequential convex optimization algorithm to obtain a feasible flag bit characterizing the dynamic feasibility of the separation point. If the feasible flag is the first preset value, then based on the maximum theoretical longitudinal range with power, the range of the unpowered gliding segment, the extreme turning angle and the minimum turning radius of the carrier in cruise level flight state, a reachable domain area function of cruise parameters is constructed, and the target cruise parameters are determined with the goal of maximizing the reachable domain area. Based on the target cruise parameters, a sequential convex optimization algorithm is used to reconstruct the transition trajectory from the separation point to the cruise start point. Then, based on the cruise start point, the reachable domain after separation is analytically generated using the polar coordinate boundary equation of the reachable domain. If there is an intersection between the preset airport set and the reachable domain, the intersection is taken as the first feasible airport subset, and the mission mode is determined to be landing after separation. If the feasible flag is not the first preset value, or there is no intersection between the preset airport set and the reachable domain, then the inter-level separation is cancelled, and emergency landing trajectory reconstruction and emergency reachable domain generation are performed based on the fault point status. If there is an intersection between the preset airport set and the emergency reachability domain, the intersection is taken as the second feasible airport subset, and the mission mode is determined to be direct emergency landing; Based on the determined mission mode, obtain the required flight range corresponding to each feasible airport in the first feasible airport subset or the second feasible airport subset; Based on airport priority levels and required flight distances, the target airport is determined from either the first subset of feasible airports or the second subset of feasible airports.

2. The thrust failure decision-making and guidance method for a horizontal takeoff and landing two-stage orbital launch vehicle as described in claim 1, characterized in that, The steps for constructing the trajectory planning problem from the fault point to the expected separation point based on the current state vector of the carrier, and iteratively solving the convex optimization subproblems corresponding to the trajectory planning problem using a sequential convex optimization algorithm to obtain the feasible flag bits characterizing the dynamic feasibility of the separation point include: Obtain the current state vector of the launch vehicle, and construct a trajectory planning problem from the failure point to the separation point based on the current state vector. The current state vector includes the launch vehicle's geocentric distance, velocity, local ballistic inclination angle, and mass. The sequential convex optimization algorithm is used to iteratively solve the convex optimization subproblem corresponding to the trajectory planning problem, and the trust region radius corresponding to the state variable, the trust region radius corresponding to the control variable, and the trust region radius corresponding to the terminal time are adaptively updated during the iterative solution process. If the sequential convex optimization algorithm converges during the iterative solution, then the feasible flag for determining the separation point dynamics is the first preset value. If the sequential convex optimization algorithm fails to converge when the iteration is completed, the feasible flag for determining the separation point dynamics is set to the second preset value.

3. The thrust failure decision-making and guidance method for a horizontal takeoff and landing two-stage orbital launch vehicle as described in claim 1, characterized in that, The steps for constructing the reachability area function of the cruise parameters based on the maximum theoretical longitudinal range under power, the range during unpowered gliding, the extreme turning angle, and the minimum turning radius of the launch vehicle in cruise level flight include: If the reachable domain does not intersect with the negative x-axis, then based on the turning drag multiplication factor, the maximum theoretical longitudinal range with power, the range of the unpowered gliding segment, the extreme turning deflection angle, and the minimum turning radius, a reachable domain area function with respect to the cruise parameters is constructed.

4. The thrust failure decision-making and guidance method for a horizontal takeoff and landing two-stage orbital launch vehicle as described in claim 1, characterized in that, When the reachable domain intersects the negative x-axis, the limiting turning angle includes the turning angle at the first intersection of the outer boundary and the negative x-axis, and the turning angle at the second intersection of the inner boundary and the negative x-axis. The steps for constructing the area function of the reachable domain with respect to the cruise parameters, based on the maximum theoretical longitudinal range with power, the range of the unpowered gliding segment, the limiting turning angle, and the minimum turning radius of the vehicle in cruise level flight, include: Based on the turning angle at the first intersection, the minimum turning radius, the turning resistance multiplication factor, the maximum theoretical longitudinal range with power, and the range of the unpowered gliding segment, the outer boundary area function is obtained. Based on the turning angle at the second intersection point, the minimum turning radius, and the unpowered gliding distance, obtain the inner boundary area function; Based on the area functions of the outer and inner boundaries, an area function of the reachable domain with respect to the cruise parameters is constructed.

5. The thrust failure decision-making and guidance method for a horizontal takeoff and landing two-stage orbital launch vehicle as described in claim 1, characterized in that, The step of determining the target cruise parameters with the objective of maximizing the reachable area includes: An optimization model is constructed based on the reachable domain area function, and feasible speeds that satisfy the constraints in the optimization model are obtained. The area set corresponding to the feasible speed is obtained based on the reachable domain area function, and the target cruise parameters are obtained based on the maximum value in the area set.

6. The thrust failure decision-making and guidance method for a horizontal takeoff and landing two-stage orbital launch vehicle as described in claim 1, characterized in that, The steps for analytically generating the separated reachable domain based on the polar coordinate boundary equation of the reachable domain include: Based on the target cruise parameters and the polar coordinate equation of the reachable domain's polar coordinate boundary, the polar coordinate boundary of the reachable domain is obtained to generate the separated reachable domain.

7. The thrust failure decision-making and guidance method for a horizontal takeoff and landing two-stage orbital launch vehicle as described in any one of claims 1 to 6, characterized in that, The step of obtaining the required flight distance corresponding to each feasible airport in the first feasible airport subset or the second feasible airport subset includes: For a target feasible airport in the first feasible airport subset or the second feasible airport subset, based on the turning radius corresponding to the optimal cruise speed of the launch vehicle and the polar coordinate boundary, obtain the straight flight segment range and the actual turning center angle corresponding to the target feasible airport. Based on the straight-line flight distance and the actual turning center angle, the required flight distance to the target airport is obtained.

8. The thrust failure decision-making and guidance method for a horizontal takeoff and landing two-stage orbital launch vehicle as described in claim 7, characterized in that, The steps for obtaining the straight-line flight segment distance and the true turning center angle corresponding to the target feasible airport based on the turning radius corresponding to the optimal cruise speed of the carrier and the polar coordinate boundary include: Based on the turning radius corresponding to the optimal cruising speed of the carrier and the polar coordinate boundary, the geometric straight-line distance is obtained; The straight-line flight range is obtained based on the geometric straight-line distance and the turning radius corresponding to the optimal cruising speed of the launch vehicle. The true turning center angle is obtained based on the turning radius, geometric straight-line distance, and polar coordinate boundary corresponding to the optimal cruising speed of the launch vehicle.

9. A thrust failure decision-making and guidance device for a horizontal takeoff and landing two-stage orbital launch vehicle, characterized in that, The device includes: a memory, a processor, and a computer program stored in the memory and executable on the processor, the computer program being configured to implement the steps of the thrust failure decision and guidance method for a horizontal takeoff and landing two-stage-to-orbit vehicle as described in any one of claims 1 to 8.

10. A storage medium, characterized in that, The storage medium is a computer-readable storage medium, and a computer program is stored on the storage medium. When the computer program is executed by a processor, it implements the steps of the thrust failure decision and guidance method for a horizontal takeoff and landing two-stage orbital launch vehicle as described in any one of claims 1 to 8.

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