A method, apparatus, and medium for determining a landing reachable domain for an aircraft

Abstract

The application discloses a kind of methods, devices and media for determining the landing accessible domain of aircraft, suitable for aircraft technical field.The stable autorotation glide state interval is solved by dynamics model, and the virtual target point is designed according to initial point and flight parameters, to determine the constraint condition in the stable autorotation glide state interval, and then the target trajectory endpoint is determined as autorotation glide trajectory according to distance function and constraint condition, to determine autorotation glide accessible domain.The technical problem that controllable pitch rotor aircraft needs to select a safe target point on the ground to realize autorotation landing in the case of air power failure is solved, and the distance function, constraint condition and virtual target point are used to calculate the autorotation landing accessible domain and autorotation landing trajectory of rotor aircraft on the ground, reduce the technical requirements for pilots, and compared with parachute landing and traditional autorotation landing method, more safe and controllable autorotation landing can be realized when facing more complex ground environment.

Description

A method, apparatus and medium for determining the landing reachability of an aircraft Technical Field

[0001] This invention relates to the field of aircraft technology, and in particular to a method, apparatus and medium for determining the landing reachability of an aircraft. Background Technology

[0002] Currently, autogyro glide landing control technology is widely used in the autogyro landing of helicopters and rotorcraft. In practical engineering implementation, for helicopter autogyro landings, pilots can only rely on visual information and their own judgment to select a suitable landing site within a certain ground range. For other aircraft, in the event of power failure, parachute descent or autogyro glide descent are used.

[0003] Most aircraft lack the capability for autonomous autorotation landing or can only achieve simple, non-fixed-point autorotation landings. When faced with sudden loss of power in mid-air, rotorcraft must resort to parachuting. However, parachuting is uncontrollable, and in complex ground environments, the safety of the landing site and the recovery of the aircraft cannot be guaranteed. In contrast, autorotation gliding requires a higher level of pilot skill, and suitable landing sites are difficult to find, thus limiting the application scenarios of autorotation gliding landing control technology.

[0004] Therefore, ensuring the safety of an aircraft's autorotation landing is a problem that urgently needs to be solved by those skilled in the art. Summary of the Invention

[0005] The purpose of this invention is to provide a method, apparatus and medium for determining the landing reachability of an aircraft, reducing the barriers to pilot skill requirements, and enabling safer and more controllable autorotation landings when facing more complex ground environments, compared to parachute landings and traditional autorotation landing methods.

[0006] To address the aforementioned technical problems, this invention provides a method for determining the landing reachability of an aircraft, comprising:

[0007] The dynamic parameters, current position, and current flight parameters of the aircraft are obtained, and the current position is used as the initial point.

[0008] The dynamic model is invoked to input the dynamic parameters, and the output parameters of the dynamic model are obtained as the stable autorotation glide state space of the aircraft.

[0009] Each virtual target point is determined based on the initial point and the current flight parameters, and the constraints of the trajectory endpoints are determined based on the initial point, each virtual target point, and the stable autorotation glide state space.

[0010] The target trajectory endpoints corresponding to each of the virtual target points are determined according to the distance function and the constraints, and the target trajectory endpoints are connected to determine the landing reachable domain of the aircraft, wherein the distance function is determined by the distance between each of the virtual target points and the corresponding trajectory endpoints.

[0011] Preferably, obtaining the output parameters of the dynamic model as the stable autorotation glide state space of the aircraft includes:

[0012] The dynamic model is balanced according to the dynamic parameters to obtain the stable rotation constraint equation and the kinetic energy constraint equation of the rotor system under the stable rotation state of the aircraft, wherein the dynamic parameters include at least the aerodynamic parameters of the aircraft and the inertial parameters of the airframe;

[0013] The stable rotation glide state space is determined based on the stable rotation constraint equation and the kinetic energy constraint equation.

[0014] Preferably, the process of establishing the stable rotation constraint equation includes:

[0015] Obtain the velocity and angular velocity values ​​of the aircraft's mechanical system under attitude stability;

[0016] The corresponding acceleration value is obtained by differentiating each of the velocity values.

[0017] The corresponding angular acceleration values ​​are obtained by differentiating each angular velocity value.

[0018] The angular velocity value, the acceleration value, and the angular acceleration value are all 0 as the stable rotation constraint equation, wherein the resultant external force value and the external torque value of the aircraft are all 0;

[0019] Correspondingly, the process of establishing the kinetic energy constraint equations includes:

[0020] Obtain the rotational speed and rotor inertia of the rotor system of the aircraft;

[0021] The kinetic energy is determined based on the relationship between the rotational speed and the moment of inertia of the rotor system to establish the kinetic energy constraint equation;

[0022] Correspondingly, determining the kinetic energy based on the relationship between the rotational speed and the moment of inertia of the rotor system includes:

[0023] Obtain the rotor type of the rotor system of the aircraft;

[0024] Determine the target rotor based on the rotor type;

[0025] Obtain the rotor speed value corresponding to the target rotor;

[0026] The kinetic energy value corresponding to the target rotor is determined based on the relationship between the rotor speed value and the rotor inertia value.

[0027] The kinetic energy values ​​are added together to establish the kinetic energy constraint equation;

[0028] or:

[0029] Obtain the rotor type of the rotor system of the aircraft;

[0030] The kinetic energy is determined based on the relationship between the rotor type, the rotational speed, and the rotor moment of inertia.

[0031] Preferably, the current flight parameters include at least the current speed value and the current altitude value, and the step of determining each virtual target point based on the initial point and the current flight parameters includes:

[0032] The threshold distance of the aircraft is determined based on the current speed value, the current altitude value, and the current position;

[0033] A preset distance is determined based on the distance between the threshold distance and the current position;

[0034] A circle is drawn with the initial point as the center and the preset distance as the radius to determine the outer boundary of the aircraft;

[0035] Points are selected on the outer boundary as virtual target points.

[0036] Preferably, the constraint conditions for determining the trajectory endpoints based on the initial point, each of the virtual target points, and the stable rotational glide state space include:

[0037] The range of unknown values ​​for the trajectory endpoints is determined based on the initial point and each of the virtual target points;

[0038] Based on the range of values ​​of the unknowns and the stable autorotation glide state space, the first state variable constraint parameter, the second state variable constraint parameter, the first control variable constraint parameter, and the second control variable constraint parameter of the spacecraft are determined.

[0039] The process constraints of the aircraft are determined based on the first state variable constraint parameters and the first control variable constraint parameters.

[0040] The endpoint constraints of the aircraft are determined based on the second state variable constraint parameters and the second control variable constraint parameters.

[0041] Determine the corresponding state equations based on the dynamic equations of the aircraft;

[0042] Determine the corresponding process performance index and endpoint performance index based on the process constraints, endpoint constraints, and state equations.

[0043] The process performance index and the endpoint performance index are combined and processed to obtain the constraint conditions of the trajectory endpoints.

[0044] Preferably, determining the target trajectory endpoints corresponding to each of the virtual target points based on the distance function and the constraints includes:

[0045] The distance function is constructed based on the distance formula between the current virtual target point and the corresponding trajectory endpoint;

[0046] Set the process performance index within the constraints to 0;

[0047] Determine multiple trajectory endpoints based on the relationship between the current virtual target point and the constraints;

[0048] The optimal performance index function is determined based on the relationship between the distance function and the endpoint performance index within the constraints.

[0049] The optimal performance index function is invoked to input the current virtual target point and multiple trajectory endpoints;

[0050] Obtain multiple output parameters corresponding to the optimal performance index function;

[0051] Select the smallest output parameter from among the multiple output parameters;

[0052] The corresponding target trajectory endpoint is determined from among the multiple trajectory endpoints based on the minimum output parameter.

[0053] Preferably, after determining the landing reachability domain, the method further includes:

[0054] When the current flight parameters meet the preset altitude conditions, the speed of the aircraft will be adjusted to a safe range;

[0055] Adjust the propeller pitch and adjust the flight attitude of the aircraft to the target attitude for landing.

[0056] To solve the above-mentioned technical problems, the present invention provides a device for determining the landing reachability of an aircraft, comprising:

[0057] The acquisition module is used to acquire the dynamic parameters, current position, and current flight parameters of the aircraft, and to use the current position as the initial point;

[0058] The calling module is used to call the dynamic model to input the dynamic parameters, and to obtain the output parameters of the dynamic model as the stable autorotation glide state space of the aircraft;

[0059] The first determining module is used to determine each virtual target point based on the initial point and the current flight parameters, and to determine the constraint conditions of the trajectory endpoints based on the initial point, each of the virtual target points and the stable autorotation glide state space.

[0060] The second determining module is used to determine the target trajectory endpoints corresponding to each of the virtual target points according to the distance function and the constraint conditions, and connect each of the target trajectory endpoints to determine the landing reachable domain of the aircraft, wherein the distance function is determined by the distance between each of the virtual target points and the corresponding trajectory endpoints.

[0061] To solve the above-mentioned technical problems, the present invention provides a device for determining the landing reachability of an aircraft, comprising:

[0062] Memory, used to store computer programs;

[0063] A processor, used to execute the computer program to implement the steps of the method for determining the landing reachability of an aircraft as described above.

[0064] To address the aforementioned technical problems, the present invention provides a computer-readable storage medium storing a computer program, which, when executed by a processor, implements the steps of the method for determining the landing reachability of an aircraft as described above.

[0065] This invention provides a method for determining the landing reachability of an aircraft, comprising: acquiring the aircraft's dynamic parameters, current position, and current flight parameters, and using the current position as the initial point; calling a dynamic model to input the dynamic parameters, and acquiring the output parameters of the dynamic model as the stable autorotation glide state space of the aircraft; determining virtual target points based on the initial point and current flight parameters, and determining the constraints of the trajectory endpoints based on the initial point, the virtual target points, and the stable autorotation glide state space; determining the target trajectory endpoints corresponding to each virtual target point based on the distance function and the constraints, and connecting the target trajectory endpoints to determine the landing reachability of the aircraft, wherein the distance function is determined by the distance between each virtual target point and its corresponding trajectory endpoint. This method solves for the stable autorotation glide state interval through a dynamic model. The stable autorotation state is characterized by the rotor rotating under no-power conditions by utilizing the updraft generated during descent to continuously generate lift and ensure a stable descent rate. Virtual target points are designed based on the initial point and flight parameters, constraints are determined using the stable autorotation glide state interval, and the target trajectory endpoints are determined as the autorotation glide trajectory based on the distance function and constraints to determine the autorotation glide reachability. This invention solves the technical problem of autorotation landing of controllable pitch rotorcraft in the event of power failure in the air, requiring the selection of a safe target point on the ground. By using distance functions, constraints, and virtual target points, the invention can calculate the reachable domain and trajectory of autorotation landing of the rotorcraft on the ground, reducing the technical requirements on pilots. Compared with parachute landing and traditional autorotation landing methods, this invention enables safer and more controllable autorotation landing in more complex ground environments.

[0066] In addition, the present invention also provides a device and medium for determining the landing reachability of an aircraft, which has the same beneficial effects as the method for determining the landing reachability of an aircraft described above. Attached Figure Description

[0067] To more clearly illustrate the embodiments of the present invention, the accompanying drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0068] Figure 1 is a schematic diagram of the existing helicopter autorotation landing process;

[0069] Figure 2 is a flowchart of a method for determining the landing reachability of an aircraft according to an embodiment of the present invention;

[0070] Figure 3 is a schematic diagram of a dynamic model provided in an embodiment of the present invention;

[0071] Figure 4 is a schematic diagram of an outer boundary provided in an embodiment of the present invention;

[0072] Figure 5 is a schematic diagram of a landing reachable area provided by an embodiment of the present invention;

[0073] Figure 6 is a structural diagram of a device for determining the landing reachability of an aircraft according to an embodiment of the present invention;

[0074] Figure 7 is a structural diagram of another device for determining the landing reachability of an aircraft provided in an embodiment of the present invention. Detailed Implementation

[0075] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the protection scope of the present invention.

[0076] The core of this invention is to provide a method, apparatus and medium for determining the landing reachability of an aircraft, reducing the barriers to pilot technical skills, and enabling safer and more controllable autorotation landing when facing more complex ground environments, compared to parachute landing and traditional autorotation landing methods.

[0077] To enable those skilled in the art to better understand the present invention, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0078] It should be noted that rotorcraft (including helicopters, gyroplanes, multi-rotor aircraft with controllable pitch, compound rotors, and eVTOL aircraft with controllable pitch) are increasingly used in the civilian sector, especially eVTOL aircraft, which have broad application prospects and have received widespread attention from the industry and the market. When faced with a sudden loss of power in mid-air, measures need to be taken to achieve an emergency landing. In existing technologies, when a rotorcraft experiences an in-flight engine failure, it deploys its parachute at an appropriate time after detecting the danger. However, for rotorcraft, parachute landings, while increasing the aircraft's load, also carry the risk of parachute lines becoming entangled with the rotor. For autorotation landings, Figure 1 shows a schematic diagram of the existing helicopter autorotation landing process. As shown in Figure 1, helicopters and gyroplanes mainly achieve forward flight with a minimum sink rate by controlling the collective pitch. Even so, traditional remote-controlled autorotation descent requires a high level of skill from the operator, and autorotation descent has high requirements for the length, flatness, and openness of the landing site.

[0079] Figure 2 is a flowchart of a method for determining the landing reachability of an aircraft according to an embodiment of the present invention. As shown in Figure 2, the method includes:

[0080] S11: Obtain the aircraft's dynamic parameters, current position, and current flight parameters, and use the current position as the initial point;

[0081] S12: Call the dynamic model to input dynamic parameters and obtain the output parameters of the dynamic model as the stable autorotation glide state space of the aircraft;

[0082] S13: Determine each virtual target point based on the initial point and current flight parameters, and determine the constraints of the trajectory endpoints based on the initial point, each virtual target point, and the stable autorotation glide state space;

[0083] S14: Determine the target trajectory endpoints corresponding to each virtual target point based on the distance function and constraints, and connect the target trajectory endpoints to determine the landing reachable domain of the aircraft;

[0084] The distance function is determined by the distance between each virtual target point and its corresponding trajectory endpoints.

[0085] Specifically, during the autorotation landing of a manned helicopter, the pilot will use visual information and other means to make judgments and select a suitable landing site. For the autorotation landing of unmanned or autonomous rotorcraft, it is necessary to make judgments and select a landing point as well as to plan a trajectory for the target point.

[0086] It should be noted that once the aircraft enters a stable autorotation state, the rotor speed is controlled to the stable autorotation landing rotor speed by controlling the collective pitch, ensuring stable rotor speed and heave rate, and preventing further rotor energy loss. The aircraft's dynamic parameters are obtained, including at least the aerodynamic parameters of the rotor and the inertial parameters of the airframe. The aircraft's current position is determined by viewing it from above as a point in the XY coordinate system, and this current position serves as the reference point (initial point) for establishing the virtual target point. The aircraft's current flight parameters can be flight speed, flight altitude, etc., and are not limited here; both can also be considered. Correspondingly, the flight speed includes forward speed and vertical speed.

[0087] Step S12 involves calling a dynamic model, which includes a rotor dynamic model, a fuselage aerodynamic model, a six-degree-of-freedom kinematic model, and an atmospheric environment model. Figure 3 is a schematic diagram of a dynamic model provided in an embodiment of the present invention. As shown in Figure 3, wind field interference and various parameters under the six-degree-of-freedom model are input into the atmospheric environment model, which outputs the vacuum velocity atmospheric density parameter. The vacuum velocity atmospheric density parameter and control distance are input into the rotor dynamic model, and the vacuum velocity atmospheric density parameter is input into the fuselage aerodynamic model. The rotor dynamic model and the fuselage aerodynamic model output aerodynamic forces and aerodynamic moments, respectively. The aerodynamic forces and aerodynamic moments are input into the six-degree-of-freedom model to output parameters such as velocity, attitude, angular velocity, and position, and are then cyclically input into the six-degree-of-freedom model for adjustment.

[0088] It should be noted that the dynamic model in the embodiments of the present invention may include, but is not limited to, the several dynamic models mentioned above, and may also include other dynamic models, etc., which are not limited here and can be set according to the actual situation. In addition, the dynamic model provided in the embodiments of the present invention can borrow mature dynamic models and input dynamic parameters, or new dynamic models can be established, etc., which are not limited here.

[0089] The dynamic model is invoked to input dynamic parameters, and the output parameters of the dynamic model are obtained as the stable autorotation glide state space of the aircraft. Correspondingly, the dynamic model is trimmed to determine the autorotation glide state space. Maintaining a stable autorotation state requires balancing gliding constraints. Trimming involves adjusting control parameters to balance the forces and moments generated by airflow on all aerodynamic components throughout the aircraft. The corresponding trimming process is not limited; the stable autorotation glide state space is determined by balancing gliding constraints and the corresponding torque kinetic energy stabilization process.

[0090] Furthermore, virtual targets are determined based on the initial point and current flight parameters. That is, virtual target points are designed from the initial point of entering a stable autorotation descent state. Considering the possibility of hollow areas within the reachable domain, the problem is divided into inner and outer boundaries. For the outer boundary, virtual target points are set directly around the farthest point from the starting point. For the inner boundary, due to the near-symmetrical nature of the aircraft's dynamic model, the shape of its unpowered autorotation reachable domain is necessarily symmetrical along the initial heading.

[0091] The method for determining each virtual target point is not limited; the initial point can be used as the center, and the defined area for the virtual target points can be rectangular, elliptical, or circular. The constraints for the trajectory endpoints are determined based on the initial point, each virtual target point, and the stable self-rotating descent state space. Since the various parameter states of the aircraft in the stable self-rotating descent state serve as the constraint parameters for the current trajectory endpoints, the constraints for the trajectory endpoints are determined through the initial point, each virtual target point, and the constraint parameters.

[0092] The distance function is derived from the distance constraints between each virtual target point and each trajectory endpoint, and it is determined by the distance formula between two points. By determining the corresponding target trajectory endpoints through the distance function and constraints, it can be understood that one virtual target point corresponds to multiple trajectory endpoints. The optimal endpoint parameters are then selected from these multiple endpoints as the target trajectory endpoint. Finally, the target trajectory endpoints are connected to determine the final landing reachable domain.

[0093] This invention provides a method for determining the landing reachability of an aircraft, comprising: acquiring the aircraft's dynamic parameters, current position, and current flight parameters, and using the current position as the initial point; calling a dynamic model to input the dynamic parameters, and acquiring the output parameters of the dynamic model as the stable autorotation glide state space of the aircraft; determining virtual target points based on the initial point and current flight parameters, and determining the constraint conditions of the trajectory endpoints based on the initial point, the virtual target points, and the stable autorotation glide state space; determining the target trajectory endpoints corresponding to each virtual target point based on the distance function and the constraint conditions, and connecting the target trajectory endpoints to determine the landing reachability of the aircraft, wherein the distance function is determined by the distance between each virtual target point and its corresponding trajectory endpoint. This method solves for the stable autorotation glide state interval through a dynamic model. The stable autorotation state is characterized by the rotor rotating under no-power conditions by utilizing the updraft generated during descent to continuously generate lift and ensure a stable descent rate. Virtual target points are designed based on the initial point and flight parameters, constraint conditions are determined using the stable autorotation glide state interval, and the target trajectory endpoints are determined as the autorotation glide trajectory based on the distance function and constraint conditions to determine the autorotation glide reachability. This invention solves the technical problem of autorotation landing of controllable pitch rotorcraft in the event of power failure in the air, requiring the selection of a safe target point on the ground. By using distance functions, constraints, and virtual target points, the invention can calculate the reachable domain and trajectory of autorotation landing of the rotorcraft on the ground, reducing the technical requirements on pilots. Compared with parachute landing and traditional autorotation landing methods, this invention enables safer and more controllable autorotation landing in more complex ground environments.

[0094] Based on the above embodiments, step S12, obtaining the output parameters of the dynamic model as the stable autorotation glide state space of the aircraft, includes:

[0095] The dynamic model is trimmed based on the dynamic parameters to obtain the stable rotation constraint equations and the kinetic energy constraint equations of the rotor system under the stable rotation state of the aircraft. The dynamic parameters include at least the aerodynamic parameters of the aircraft and the inertial parameters of the airframe.

[0096] The stable rotation glide state space is determined based on the relationship between the stable rotation constraint equation and the kinetic energy constraint equation.

[0097] Specifically, the dynamic model is trimmed based on the dynamic parameters to obtain the stable rotation constraint equations and kinetic energy constraint equations for the stable rotation state of the aircraft. The rotation glide state is determined based on the relationship between the two constraint equations.

[0098] As one example, the process of establishing the stable rotation constraint equations includes:

[0099] Obtain the velocity and angular velocity values ​​of the aircraft's mechanical system under attitude stability;

[0100] The corresponding acceleration values ​​are obtained by differentiating each velocity value.

[0101] The corresponding angular acceleration values ​​are obtained by differentiating each angular velocity value.

[0102] The angular velocity, acceleration, and angular acceleration values ​​are all 0, which is taken as the stable rotation constraint equation, where the net external force and external torque of the aircraft are all 0.

[0103] Correspondingly, the process of establishing the kinetic energy constraint equations includes:

[0104] Obtain the rotational speed and rotor inertia of the aircraft's rotor system;

[0105] The kinetic energy is determined based on the relationship between the rotor system's rotational speed and rotor inertia to establish the kinetic energy constraint equation;

[0106] Obtain the rotor type of the aircraft's rotor system;

[0107] Determine the target rotor based on rotor type;

[0108] Obtain the rotor speed value corresponding to the target rotor;

[0109] The kinetic energy value of the target rotor is determined based on the relationship between the rotor speed and the rotor inertia.

[0110] The kinetic energy values ​​are summed to establish the kinetic energy constraint equations;

[0111] or:

[0112] Obtain the rotor type of the aircraft's rotor system;

[0113] Kinetic energy is determined based on the relationship between rotor type, rotational speed, and rotor moment of inertia.

[0114] Specifically, the stability and rotation constraint equations are established by assuming the aircraft is descending a straight line with a small incline. The equilibrium equations for stable flight are established using the velocity (uvw) and angular velocity (pqr) parameters at the stable state. Differentiating each velocity value yields the corresponding acceleration value. During the stable process, the angular velocity and acceleration are both zero. Differentiating each angle value also yields a angular acceleration value of zero. Based on these angular velocity, acceleration, and angular acceleration values, the equilibrium equations are established, with the specific formulas as follows:

[0115] p≈0

[0116] q≈0

[0117] r≈0

[0118] Where uvw represents the velocity value under the aircraft's airframe system, and pqr represents the angular velocity value under the airframe system. The acceleration value under the machine system, The value of angular acceleration under the machine system.

[0119] Furthermore, under equilibrium conditions, based on the torque relationship F=Ma, the constraint relationship between the net external force and the external torque of the aircraft is determined, as shown in the following formula:

[0120]

[0121]

[0122] In the above formulas, the first formula is when the net external force is 0, and the second formula is when the external torque is 0. These are based on the corresponding values ​​in the three-dimensional coordinate system.

[0123] The process of establishing the kinetic energy constraint equation involves obtaining the rotational speed and rotor inertia of the aircraft's rotor system, and determining the kinetic energy based on the relationship between these two parameters. As one example, the rotor inertia is calculated using the rotational inertia of the rotor system and the rotational inertia of the associated transmission gears of each rotor. The rotor inertia is determined by summing the rotational inertia of the rotor system and the rotational inertia of the associated transmission gears of each rotor. The rotor inertia is expressed by the following formula:

[0124] I sum =I rw +I gear

[0125] Among them, I rw I is the moment of inertia of the rotor. gearThe moment of inertia after disengaging the clutch, excluding the rotor, generally refers to the moment of inertia of the rotor shaft and transmission gears.

[0126] It should be noted that kinetic energy is determined by the rotational speed and moment of inertia of the rotor system. The number of rotors and their corresponding kinetic energies vary depending on the type of rotor system. Rotor types include at least multi-rotor systems sharing a single power transmission system, multi-rotor systems with individual power systems, and multi-rotor systems with a compound configuration.

[0127] Correspondingly, a multirotor sharing a single power transmission system refers to a helicopter where the main rotor and tail rotor share a single transmission system. The main rotor and tail rotor are driven by a single engine and connected via a gear shaft. Such a multirotor essentially constitutes an energy system. Calculating the kinetic energy requires calculating the kinetic energy of the main rotor and tail rotor, then summing them to obtain the kinetic energy of the rotor system. A multirotor with a separate power system refers to a multirotor where each rotor has its own power system, such as a common quadcopter. Each rotor is driven independently, and there is no direct energy relationship between the different rotors. Each rotor is considered a separate energy system, and the kinetic energy of each rotor is calculated separately and then summed to obtain the kinetic energy of the rotor system. A multirotor with a compound configuration refers to a rotor assembly that has both common and separate drives, typically found in helicopters, especially twin-rotor or multi-engine helicopters. Its kinetic energy requires calculating the kinetic energy of the entire aircraft driven by the combined power of all rotors, as well as the kinetic energy of each individual rotor, and then summing them to obtain the kinetic energy of the rotor system.

[0128] Understandably, for different rotor types, the target rotor is determined based on the rotor type. The rotor speed corresponding to the target rotor is the rotational speed of each rotor in the rotor system. The corresponding kinetic energy value is determined based on the relationship between the rotor speed and the rotor inertia. As one example, the determination of the kinetic energy value of the target rotor includes:

[0129] The squared rotational speed value is obtained by squared the rotational speed of the target rotor.

[0130] Multiply the squared rotational speed value by the rotor inertia value and divide by 2 to obtain the kinetic energy value of the target rotor.

[0131] The corresponding target rotor can be one rotor or multiple rotors, without limitation. It is possible to determine the corresponding kinetic energy value of multiple target rotors at the same time, or to calculate the kinetic energy value of one target rotor first, and then calculate the kinetic energy of the next target rotor, until all target rotors have been calculated.

[0132] The kinetic energy relationship of a rotor is as follows:

[0133]

[0134] Among them, Ω rw is the rotor speed value, and const is the kinetic energy value of a rotor.

[0135] The kinetic energy is obtained by adding the kinetic energy values ​​of the target rotors. The kinetic energy of an aircraft in a stable autorotating glide state is constant.

[0136] In addition, the anti-torque coefficient C resulting from rotor rotation in a stable state of the aircraft Q It is 0, that is, C Q =0.

[0137] The process for determining the stable autorotation glide state space provided in this embodiment of the invention involves a rotorcraft generating lift by using the upward airflow generated during descent to propel the rotor in the absence of power, thereby ensuring a stable descent rate and facilitating the subsequent determination of the autorotation glide reachable domain.

[0138] Based on the above embodiments, the current flight parameters include at least the current speed value and the current altitude value. Step S13, determining each virtual target point based on the initial point and the current flight parameters, includes:

[0139] The threshold distance for the aircraft is determined based on the current speed, current altitude, and current position.

[0140] The preset distance is determined based on the distance between the threshold distance and the current location;

[0141] Draw a circle with the initial point as the center and a preset distance as the radius to determine the outer boundary of the aircraft;

[0142] Points on the outer boundary are taken as virtual target points.

[0143] Specifically, the threshold distance is determined based on the relationship between the current speed value, the current height value, and the current position. The threshold distance is the distance to the farthest position of the current initial point. The preset distance is determined based on the relationship between the threshold distance and the current position. The preset distance is obtained by multiplying the threshold distance by a preset multiple. The preset multiple can be set to 2 times or 3 times, etc., without limitation, and can be set according to the actual situation.

[0144] Draw a circle with the initial point as the center and a preset distance as the radius to determine the outer boundary of the aircraft, and take points on the outer boundary as each virtual target point.

[0145] Figure 4 is a schematic diagram of an outer boundary provided by an embodiment of the present invention. As shown in Figure 4, a virtual target point (x0, y0) is designed around the initial point (x0, y0) of entering a stable self-rotation and sliding state. mk y mkHere, considering that the reachable domain may have hollow interiors, the problem is divided into two parts: inner boundary and outer boundary. For the outer boundary, a virtual target point is set directly around the farthest point of the starting point.

[0146] As one embodiment, the constraint conditions for determining the trajectory endpoints in step S13 based on the initial point, each virtual target point, and the stable rotational glide state space include:

[0147] The range of unknown values ​​for the trajectory endpoints is determined based on the initial point and each virtual target point;

[0148] The first state variable constraint parameters, the second state variable constraint parameters, the first control variable constraint parameters, and the second control variable constraint parameters of the aircraft are determined based on the range of unknown values ​​and the stable autorotation glide state space.

[0149] The process constraints of the aircraft are determined based on the first state variable constraint parameters and the first control variable constraint parameters.

[0150] Determine the endpoint constraints of the aircraft based on the second state variable constraint parameters and the second control variable constraint parameters;

[0151] Determine the corresponding state equations based on the dynamic equations of the aircraft;

[0152] Determine the corresponding process performance indicators and endpoint performance indicators based on process constraints, endpoint constraints, and state equations.

[0153] The process performance indicators and endpoint performance indicators are processed to obtain the constraints of the trajectory endpoints.

[0154] Specifically, the range of unknown values ​​[t0, t] for the trajectory endpoints is determined based on the initial point and each virtual target point. f Based on the range of unknown values ​​and the state parameters of the stable autorotation glide state space, the parameters of each variable of the spacecraft are determined, including the first state variable constraint parameter, the second state variable constraint parameter, the first control variable constraint parameter, and the second control variable constraint parameter.

[0155] The process constraints of the aircraft are determined based on the relationship between the first state variable constraint parameters and the first control variable constraint parameters. These first state variable parameters and first control variable constraint parameters are generally derived from the state variable constraints and control variable constraints caused by the structural and performance limitations of the rotorcraft, pilot control restrictions, airworthiness regulations, comfort, and flight safety. Specific process constraints are as follows:

[0156]

[0157] Among them, c ec is the first state variable constraint parameter. i The first control variable constraint parameter.

[0158] The endpoint constraints of the aircraft are determined based on the relationship between the second state variable constraint parameters and the second control variable constraint parameters. These parameters are derived from the state variable constraints resulting from the initial conditions of entering the autorotation phase and the characteristics of the autorotation landing deceleration phase. The specific endpoint constraints are as follows:

[0159]

[0160] in, These are the constraint parameters for the second state variable. These are the constraint parameters for the second control variable.

[0161] The corresponding state equations are determined based on the aircraft's dynamic equations, and the specific formulas are as follows:

[0162]

[0163] Based on the process constraints, endpoint constraints, and state equations, determine the corresponding process performance indices and endpoint performance indices, where L(x(t), u(t), t) are process performance indices, and φ(x(t0), t0, x(x) are process performance indices. f ), t f ) is an endpoint performance metric.

[0164] The process performance indicators and endpoint performance indicators are combined and processed to obtain the constraints of the trajectory endpoints, which is the optimal control problem, as follows:

[0165]

[0166] Where x(t) is the state variable, u(t) is the control variable, L(x(t), u(t), t) is the process performance index, and φ(x(t0), t0, x(t)) is the control variable. f ), t f ) is an endpoint performance metric.

[0167] As one embodiment, step S14, determining the target trajectory endpoints corresponding to each virtual target point based on the distance function and constraints, includes:

[0168] A distance function is constructed based on the distance formula between the current virtual target point and the corresponding trajectory endpoint;

[0169] Set the process performance index within the constraints to 0;

[0170] Determine multiple trajectory endpoints based on the relationship between the current virtual target point and the constraints;

[0171] Determine the optimal performance index function based on the relationship between the distance function and the endpoint performance index within the constraints;

[0172] Call the optimal performance metric function with the current virtual target point and multiple trajectory endpoints as input;

[0173] Obtain multiple output parameters corresponding to the optimal performance index function;

[0174] Select the smallest output parameter from multiple output parameters;

[0175] The target trajectory endpoint is determined from multiple trajectory endpoints based on the minimum output parameter.

[0176] Specifically, based on the current virtual target point (x mk y mk ) and the corresponding trajectory endpoints (x(t) f ), y(t) f The distance formula for x gives the distance function (x) mk -x(t f )) 2 +(y mk -y(t f )) 2 .

[0177] In the above embodiment, the process performance index is 0. Multiple trajectory endpoints (x(t) are determined based on the relationship between the current virtual target point and the constraints. f ), y(t) f The optimal performance index function is determined based on the relationship between the distance function and the endpoint performance index. The specific formula is as follows:

[0178] φ(x(t0), t0, x(t) f ), t f )=(x mk -x(t f )) 2 +(y mk -y(t f )) 2

[0179] Based on the aforementioned optimal performance index function, by inputting the current virtual target point and the corresponding trajectory endpoints, multiple output parameters can be obtained. The minimum output parameter is found among these multiple output parameters, and the corresponding target trajectory endpoint can be found based on the minimum output parameter.

[0180] Figure 5 is a schematic diagram of a landing reachable domain provided by an embodiment of the present invention. As shown in Figure 5, the corresponding optimized trajectory is obtained by optimizing the constraint relationship between the distance function and the virtual target point. The actual reachable domain is the inner boundary. For the inner boundary, since the dynamic model of the aircraft is basically symmetrical from left to right, the shape of its unpowered self-rotation reachable domain must be symmetrical along the starting heading direction. Therefore, if a hollow region appears inside the reachable domain, it must also be symmetrical along the central axis along the starting heading direction. It can be selected as an unreachable point of the virtual target point and can be selected along the heading direction.

[0181] The process for determining the landing reachability of an aircraft provided in this embodiment of the invention determines the target trajectory endpoint as the autorotation glide path based on the distance function and constraint conditions, thereby determining the autorotation glide reachability. By using the distance function, constraint conditions, and virtual target point, the autorotation landing reachability and autorotation landing trajectory of the rotorcraft on the ground can be calculated, reducing the barriers to pilot technical skills.

[0182] Based on the above embodiments, after determining the landing reachability domain, the method further includes:

[0183] When the current flight parameters meet the preset altitude conditions, adjust the aircraft's speed to a safe range;

[0184] Adjust the propeller pitch and adjust the aircraft's flight attitude to the target attitude for landing.

[0185] Specifically, its autorotation landing process includes the following steps:

[0186] Step 1: Enter a stable autorotation state. Control the rotor speed to the stable autorotation landing rotor speed by controlling the collective pitch to ensure that the rotor speed and sink rate are stable and the rotor energy is no longer lost.

[0187] Step 2: Calculate the reachable zone of the autorotation landing projected onto the ground based on the altitude and flight status of the aircraft. After entering a stable autorotation state, select a landing target point within the reachable zone and obtain the autorotation landing trajectory for the target point.

[0188] Step 3: By controlling the attitude, maintain the autorotation glide trajectory to track the autorotation landing target trajectory calculated by the flight control computer (that is, the above embodiment), and at the same time stabilize the rotor speed and sink rate by controlling the propeller pitch.

[0189] Step 4: Once the aircraft has stabilized at the deceleration altitude (which varies depending on the specific aircraft, but is generally 25m), use attitude control to bring the lateral and longitudinal speeds within a safe range. Simultaneously, increase the rotor pitch and adjust the pitch attitude to reduce the sink rate and forward speed. Achieve a leveling deceleration maneuver through horizontal speed control and a given vertical speed control. The deceleration altitude can be adjusted appropriately based on the specific characteristics of the rotorcraft.

[0190] Step 5: When the aircraft is at 8-10m, pull up the propeller pitch to increase the pitch angle and rapidly reduce the vertical speed and forward speed. During the final landing process, control the pitch angle to rapidly decrease, and return the attitude to near 0 degrees upon touchdown, controlling the descent rate to within 3m / s to complete the landing.

[0191] It should be noted that the specific parameters in this embodiment are not limited, but are only a preferred embodiment, and can be set according to the actual situation.

[0192] The control process of the aircraft provided in this embodiment of the invention, after determining the autorotation landing trajectory plan, achieves a safe autorotation landing by using propeller pitch control and attitude control. Compared with parachute landing and traditional autorotation landing methods, it can achieve a safer and more controllable autorotation landing when facing more complex ground environments.

[0193] The foregoing has described in detail various embodiments of the method for determining the landing reachability of an aircraft. Based on this, the present invention also discloses a device for determining the landing reachability of an aircraft corresponding to the above method. Figure 6 is a structural diagram of a device for determining the landing reachability of an aircraft provided in an embodiment of the present invention. As shown in Figure 6, the device for determining the landing reachability of an aircraft includes:

[0194] The acquisition module 11 is used to acquire the dynamic parameters, current position, and current flight parameters of the aircraft, and to use the current position as the initial point;

[0195] Module 12 is used to call the dynamic model to input dynamic parameters and obtain the output parameters of the dynamic model as the stable autorotation glide state space of the aircraft.

[0196] The first determining module 13 is used to determine each virtual target point based on the initial point and the current flight parameters, and to determine the constraint conditions of the trajectory endpoints based on the initial point, each virtual target point and the stable autorotation glide state space.

[0197] The second determining module 14 is used to determine the target trajectory endpoints corresponding to each virtual target point according to the distance function and the constraint conditions, and connect each target trajectory endpoint to determine the landing reachable domain of the aircraft, wherein the distance function is determined by the distance between each virtual target point and its corresponding trajectory endpoint.

[0198] Since the embodiments of the device part correspond to the embodiments described above, please refer to the embodiments described in the method part for the embodiments of the device part, and will not be repeated here.

[0199] For a description of the device for determining the landing reachability of an aircraft provided by the present invention, please refer to the above method embodiments. The present invention will not be described in detail here, but it has the same beneficial effects as the above method for determining the landing reachability of an aircraft.

[0200] Figure 7 is a structural diagram of another device for determining the landing reachability of an aircraft provided in an embodiment of the present invention. As shown in Figure 7, the device includes:

[0201] Memory 21 is used to store computer programs;

[0202] Processor 22 is used to implement steps of a method for determining the landing reachability of an aircraft when executing a computer program.

[0203] The processor 22 may include one or more processing cores, such as a quad-core processor or an octa-core processor. The processor 22 may be implemented using at least one of the following hardware forms: Digital Signal Processor (DSP), Field-Programmable Gate Array (FPGA), or Programmable Logic Array (PLA). The processor 22 may also include a main processor and a coprocessor. The main processor, also known as the Central Processing Unit (CPU), is used to process data in the wake-up state; the coprocessor is a low-power processor used to process data in the standby state. In some embodiments, the processor 22 may integrate a Graphics Processing Unit (GPU), which is responsible for rendering and drawing the content to be displayed on the screen. In some embodiments, the processor 22 may also include an Artificial Intelligence (AI) processor, which handles computational operations related to machine learning.

[0204] The memory 21 may include one or more computer-readable storage media, which may be non-transitory. The memory 21 may also include high-speed random access memory and non-volatile memory, such as one or more disk storage devices or flash memory devices. In this embodiment, the memory 21 is used to store at least the following computer program 211, which, after being loaded and executed by the processor 22, is capable of implementing the relevant steps of the method for determining the landing reachability of an aircraft disclosed in any of the foregoing embodiments. In addition, the resources stored in the memory 21 may also include an operating system 212 and data 213, etc., and the storage method may be temporary storage or permanent storage. The operating system 212 may include Windows, Unix, Linux, etc. The data 213 may include, but is not limited to, the data involved in the method for determining the landing reachability of an aircraft.

[0205] In some embodiments, the device for determining the landing reachability of an aircraft may further include a display screen 23, an input / output interface 24, a communication interface 25, a power supply 26, and a communication bus 27.

[0206] Those skilled in the art will understand that the structure shown in Figure 7 does not constitute a limitation on the apparatus for determining the landing reachability of an aircraft and may include more or fewer components than shown.

[0207] The processor 22 implements the method for determining the landing reachability of an aircraft provided in any of the above embodiments by calling instructions stored in the memory 21.

[0208] For a description of the device for determining the landing reachability of an aircraft provided by the present invention, please refer to the above method embodiments. The present invention will not be described in detail here, but it has the same beneficial effects as the above method for determining the landing reachability of an aircraft.

[0209] Furthermore, the present invention also provides a computer-readable storage medium storing a computer program, which, when executed by processor 22, implements the steps of the method for determining the landing reachability of an aircraft as described above.

[0210] It is understood that if the methods in the above embodiments are implemented as software functional units and sold or used as independent products, they can be stored in a computer-readable storage medium. Based on this understanding, the technical solution of the present invention, in essence, or the part that contributes to the prior art, or all or part of the technical solution, can be embodied in the form of a software product. This computer software product is stored in a storage medium and executes all or part of the steps of the methods described in the various embodiments of the present invention. The aforementioned storage medium includes various media capable of storing program code, such as USB flash drives, portable hard drives, read-only memory (ROM), random access memory (RAM), magnetic disks, or optical disks.

[0211] For an introduction to the computer-readable storage medium provided by the present invention, please refer to the above method embodiments. The present invention will not be described in detail here, but it has the same beneficial effects as the above method for determining the landing reachability of an aircraft.

[0212] The foregoing has provided a detailed description of the method, apparatus, and medium for determining the landing reachability of an aircraft provided by this invention. The various embodiments in the specification are described in a progressive manner, with each embodiment focusing on its differences from other embodiments. Similar or identical parts between embodiments can be referred to interchangeably. For the apparatus disclosed in the embodiments, since it corresponds to the method disclosed in the embodiments, the description is relatively simple; relevant parts can be referred to in the method section. It should be noted that those skilled in the art can make various improvements and modifications to this invention without departing from the principles of the invention, and these improvements and modifications also fall within the protection scope of the claims of this invention.

[0213] It should also be noted that, in this specification, relational terms such as "first" and "second" are used only to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

Claims

1. A method for determining the landing reachability region of an aircraft, characterized in that, include: The dynamic parameters, current position, and current flight parameters of the aircraft are obtained, and the current position is used as the initial point. The system invokes a dynamic model to input the dynamic parameters and obtains the output parameters of the dynamic model as the stable autorotation glide state space of the aircraft; it determines each virtual target point based on the initial point and the current flight parameters, and determines the constraint conditions of the trajectory endpoints based on the initial point, each virtual target point, and the stable autorotation glide state space; it determines the target trajectory endpoints corresponding to each virtual target point based on the distance function and the constraint conditions, and connects each target trajectory endpoint to determine the landing reachability domain of the aircraft, wherein the distance function is determined by the distance between each virtual target point and the corresponding trajectory endpoint; correspondingly, the current flight parameters include at least the current velocity value and the current altitude value, and determining each virtual target point based on the initial point and the current flight parameters includes: determining the outer boundary of the aircraft based on the current velocity value, the current altitude value, and the current position; Points are selected on the outer boundary as virtual target points.

2. The method for determining the landing reachability domain of an aircraft according to claim 1, characterized in that, The step of obtaining the output parameters of the dynamic model as the stable autorotation glide state space of the aircraft includes: balancing the dynamic model according to the dynamic parameters to obtain the stable autorotation constraint equation and the kinetic energy constraint equation of the rotor system under the stable autorotation state of the aircraft, wherein the dynamic parameters include at least the aerodynamic parameters of the aircraft and the inertial parameters of the airframe; and determining the stable autorotation glide state space according to the stable autorotation constraint equation and the kinetic energy constraint equation.

3. The method for determining the landing reachability domain of an aircraft according to claim 2, characterized in that, The process of establishing the stable rotation constraint equation includes: obtaining the velocity and angular velocity values ​​of the aircraft's mechanical system under stable attitude; differentiating each velocity value to obtain the corresponding acceleration value; differentiating each angular velocity value to obtain the corresponding angular acceleration value; and using the angular velocity value, the acceleration value, and the angular acceleration value all being 0 as the stable rotation constraint equation, wherein the net external force value and the external torque value of the aircraft are all 0. Correspondingly, the process of establishing the kinetic energy constraint equation includes: obtaining the rotational speed value and rotor inertia value of the aircraft's rotor system; and determining the kinetic energy based on the relationship between the rotational speed value and the rotor inertia value of the rotor system to establish the kinetic energy constraint equation.

4. The method for determining the landing reachability domain of an aircraft according to claim 1, characterized in that, Determining the outer boundary of the aircraft based on the current speed value, the current altitude value, and the current position includes: determining a threshold distance of the aircraft based on the current speed value, the current altitude value, and the current position; determining a preset distance based on the distance between the threshold distance and the current position; and drawing a circle with the initial point as the center and the preset distance as the radius to determine the outer boundary of the aircraft.

5. The method for determining the landing reachability of an aircraft according to claim 4, characterized in that, The step of determining the constraints of the trajectory endpoints based on the initial point, each of the virtual target points, and the stable gliding state space includes: determining the range of unknown values ​​for the trajectory endpoints based on the initial point and each of the virtual target points; determining the first state variable constraint parameter, the second state variable constraint parameter, the first control variable constraint parameter, and the second control variable constraint parameter of the aircraft based on the range of unknown values ​​and the stable gliding state space; determining the process constraints of the aircraft based on the first state variable constraint parameter and the first control variable constraint parameter; determining the endpoint constraints of the aircraft based on the second state variable constraint parameter and the second control variable constraint parameter; determining the corresponding state equation based on the dynamic equation of the aircraft; determining the corresponding process performance index and endpoint performance index based on the process constraints, the endpoint constraints, and the state equation; and performing a summation and processing on the process performance index and the endpoint performance index to obtain the constraints of the trajectory endpoints.

6. The method for determining the landing reachability domain of an aircraft according to any one of claims 1 to 5, characterized in that, After determining the landing reachable area, the method further includes: when the current flight parameters meet the preset altitude conditions, adjusting the speed value of the aircraft to a safe range; adjusting the propeller pitch; and adjusting the flight attitude of the aircraft to the target attitude for landing.

7. A device for determining the landing reachability of an aircraft, characterized in that, include: The acquisition module is used to acquire the dynamic parameters, current position, and current flight parameters of the aircraft, and to use the current position as the initial point; The calling module is used to call the dynamic model to input the dynamic parameters, and to obtain the output parameters of the dynamic model as the stable autorotation glide state space of the aircraft; The first determining module is used to determine each virtual target point based on the initial point and the current flight parameters, and to determine the constraint conditions of the trajectory endpoints based on the initial point, each of the virtual target points and the stable autorotation glide state space. The second determining module is used to determine the target trajectory endpoints corresponding to each of the virtual target points according to the distance function and the constraint conditions, and connect each of the target trajectory endpoints to determine the landing reachable domain of the aircraft, wherein the distance function is determined by the distance between each of the virtual target points and the corresponding trajectory endpoints; correspondingly, the current flight parameters include at least the current speed value and the current altitude value, and determining each virtual target point according to the initial point and the current flight parameters includes: determining the outer boundary of the aircraft according to the current speed value, the current altitude value and the current position; Points are selected on the outer boundary as virtual target points.

8. A device for determining the landing reachability of an aircraft, characterized in that, include: Memory, used to store computer programs; A processor, configured to implement the steps of the method for determining the landing reachability of an aircraft as described in any one of claims 1 to 6 when executing the computer program.

9. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores a computer program that, when executed by a processor, implements the steps of the method for determining the landing reachability of an aircraft as described in any one of claims 1 to 6.

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