Parachute deployment determination method for planetary lander based on multi-dimensional constraint window

Abstract

The application discloses a parachute opening determination method for a planetary lander based on a multi-dimensional constraint window and belongs to the technical field of deep space exploration. The application is realized in the following manner: a dynamics model of a lander entering stage, parachute descending stage and power descending stage is respectively established; a parachute descending stage and power descending stage switching condition is set; a fuel consumption constraint and thrust constraint of the power descending stage are considered; a maximum and minimum range optimization problem model is established; a height-velocity two-dimensional constraint parachute opening window is solved according to a Mach number, dynamic pressure and height constraint of the lander opening parachute; for each sampling point, a different initial track angle is set; the maximum and minimum range after the parachute opening is solved; a multi-dimensional constraint parachute opening window characterized by height, velocity, track angle and limit range is constructed; according to the lander state and the multi-dimensional parachute opening window, the maximum and minimum range after the lander opening parachute is predicted; a lander range control capability evaluation index is constructed; a parachute opening logic criterion is established from the index; and the parachute opening time of the lander is controlled according to the parachute opening logic criterion.

Description

Technical Field

[0001] The present invention relates to a planetary landing parachute deployment control method, in particular to a planetary lander parachute deployment determination method based on a multi-dimensional constraint window, and belongs to the field of deep space exploration technology. Background Art

[0002] Precision landing technology on planetary surfaces is a key technology for planetary exploration and a prerequisite for conducting in-situ exploration and sample return missions. The planetary landing process can generally be divided into three phases: entry, parachute descent, and powered descent. The entry phase primarily decelerates the spacecraft, but due to the complex aerodynamic environment, range control capabilities are limited. After the entry phase, a parachute is required to further reduce the flight speed. Parachute deployment is subject to multiple constraints, which form the deployment window. For example, for a Mars landing mission, three constraints are typically considered: dynamic pressure, altitude, and Mach number. A deployment window is established in the altitude-velocity plane. Once the lander reaches the deployment window, the deployment command is executed only when the parachute deployment conditions (referred to as deployment conditions) are met. The timing of deployment affects both parachute safety and final landing accuracy. Therefore, to achieve safe deployment and precise planetary landing, it is necessary to develop deployment decision methods that meet these multi-dimensional constraint windows.

[0003] There have been relevant studies on the design of parachute deployment judgment methods for planetary landers, mainly including the time series method, overload method, pressure altitude method, radar altitude method, and altitude-velocity window constraint method (Dong Jie, Rao Wei, Sun Zezhou, et al. Multidisciplinary cross-design and verification of key links for Mars landing [J]. Acta Astronautics, 2022, 43(1): 21-29.). Among them, the time series method, overload method, and pressure altitude method require accurate planetary atmospheric models and are difficult to use for landing missions on the surface of unknown extraterrestrial bodies. The radar altitude method requires the installation of an altitude radar near the bottom of the lander, which has high requirements for thermal protection capabilities. The altitude-velocity window constraint method uses parameters such as dynamic pressure, Mach number, and altitude as control targets, has strong adaptability, and has been widely used in many Mars landing missions. The essence of this method is to determine whether the lander state is within the safe parachute deployment constraint window. Although a lander in the constraint window can meet the conditions for safe parachute deployment, the lander's flight range is reduced after parachute deployment, and it may not be able to reach the desired landing area. Therefore, the parachute deployment judgment method cannot simply use safe parachute deployment conditions as the sole judgment criteria. Therefore, to meet the requirements of safe parachute deployment and precise landing in planetary exploration, it is necessary to improve the traditional parachute deployment judgment method and develop a parachute deployment judgment method that takes into account both safe parachute deployment and precise landing. Summary of the Invention

[0004] In response to the needs of safe parachute opening and precise landing in planetary exploration, the purpose of the present invention is to provide a planetary lander parachute opening judgment method based on a multi-dimensional constraint window, comprehensively considering the parachute landing stage and the powered descent stage, and establishing an optimization model for the extreme range of the planetary lander parachute after opening. In the altitude-speed two-dimensional parachute opening window, the initial track angle and range control capability are considered, and a multi-dimensional constraint parachute opening window characterized by altitude, speed, track angle, and extreme range is constructed to ensure the safety of parachute opening. A range control capability evaluation index is established, and the information on the range to be flown is incorporated into the parachute opening logic. According to the parachute opening logic, the planetary lander parachute opening judgment based on the multi-dimensional constraint window is realized. The present invention can take into account the needs of safe parachute opening and precise landing of the planetary lander, realize precise control of the range, and is suitable for planetary landing missions.

[0005] The purpose of the present invention is achieved through the following technical solutions.

[0006] The planetary lander parachute deployment determination method based on a multi-dimensional constraint window disclosed in the present invention comprises the following steps:

[0007] Step 1: In the planet-fixed coordinate system, establish the dynamic models of the planetary lander's entry phase, parachute phase, and powered descent phase respectively; set the switching conditions between the parachute phase and the powered descent phase, consider the fuel consumption constraints and thrust constraints of the powered descent phase, and establish the maximum range optimization problem model and the minimum range optimization problem model. The maximum range optimization problem model and the minimum range optimization problem model are collectively referred to as the range optimization problem model.

[0008] The specific implementation method of step one is:

[0009] Define the planet-fixed coordinate system O-XYZ: take the center of the planet as the origin O of the coordinate system, the OX axis points to the intersection of the equator and the prime meridian, the OZ axis points to the North Pole, and the OY axis obeys the right-hand rule;

[0010] The dynamic model of the lander entry phase is established in the planet-fixed coordinate system:

[0011]

[0012] Where r is the distance from the lander to the center of the planet, θ is the longitude, φ is the latitude, v is the lander speed, γ is the track angle, ψ is the heading angle, g is the planet's gravity acceleration, L is the lander lift acceleration, and D is the lander drag acceleration;

[0013] The expressions of lift acceleration and drag acceleration of the lander are as follows:

[0014]

[0015] Where m e is the mass of the lander during entry, ρ is the atmospheric density, S Ais the reference area of ​​the lander, C L with C D The lift coefficient and drag coefficient of the lander respectively;

[0016] The dynamic model of the lander parachute phase is established in the planet-fixed coordinate system:

[0017]

[0018] Where, is the lift acceleration of the lander after parachute opening, is the drag acceleration of the lander after parachute opening, and its expression is as follows:

[0019]

[0020] Where m p is the mass of the lander during the parachute landing phase, is the reference area of ​​the lander after parachute opening, and The lift coefficient and drag coefficient of the lander after the parachute is opened respectively;

[0021] The dynamic model of the lander's powered descent phase is established in the planet-fixed coordinate system:

[0022]

[0023] Where m T is the mass of the lander in the powered descent phase, T is the thrust amplitude, η is the thrust pitch angle, σ is the thrust direction angle, I sp represents the engine specific impulse, g0 represents the acceleration of gravity at sea level;

[0024] During the powered descent, the following thrust constraints are considered:

[0025] T min ≤T≤T max (6)

[0026] Where, T min and T max are the minimum and maximum thrust of the lander engine respectively;

[0027] During the power reduction phase, the following fuel consumption constraints are considered:

[0028] m T0 -m Tf ≥m fuel (7)

[0029] Where m T0 and m Tf are the lander masses at the initial and terminal moments of the powered descent phase, m fuel represents the lander fuel mass;

[0030] The switching condition between the parachute stage and the powered descent stage: when the lander altitude drops to h T When the parachute is jettisoned and the engine is ignited, the lander enters the powered descent phase from the parachute phase;

[0031] Considering the thrust constraints and fuel consumption constraints in the power descent phase shown in equations (6) and (7), the maximum range optimization model shown in equation (8) and the minimum range optimization model shown in equation (9) are established:

[0032]

[0033] Where S represents the distance traveled by the lander from parachute opening to landing, X = [r,θ,φ,v,γ,ψ] T represents the lander state vector, F p (·) and F T (·) represent the dynamic equations of the parachute landing stage and the powered descent stage, respectively, t p , t T and t f They represent the parachute opening time, engine ignition time and landing time respectively, h p 、v p and γ p They represent the altitude, speed and track angle of the lander at the initial moment of the parachute phase.

[0034] Step 2: Solve the altitude-speed two-dimensional constrained parachute opening window based on the Mach number, dynamic pressure and altitude constraints of the planetary lander parachute opening; uniformly sample in the altitude-speed window, set a different initial track angle for each sampling point, and solve the maximum and minimum ranges after parachute opening through the range optimization problem model established in step 1, and construct a multi-dimensional constrained parachute opening window characterized by altitude, speed, track angle and extreme range.

[0035] The specific implementation method of step 2 is:

[0036] In order to ensure that the parachute can be opened safely and smoothly, the Mach number, dynamic pressure and altitude constraints must be met during opening, as shown in formula (10):

[0037]

[0038] Where M min and M max Respectively represent the minimum Mach number and maximum Mach number when the parachute is opened, M p represents the parachute opening Mach number, q min and q max They represent the minimum dynamic pressure and maximum dynamic pressure when the parachute is opened, q p Indicates the dynamic pressure of parachute opening, h min Indicates the minimum altitude when the parachute is opened;

[0039] According to the constraints shown in formula (10), the traditional height-speed parachute opening window is obtained, as shown in formula (11):

[0040]

[0041] Where c(h) represents the speed of sound at height h;

[0042] Uniform sampling is performed in the height-speed window, as shown in Equation (12):

[0043]

[0044] Where, sample represents uniform sampling in the height-speed window, (h pi ,v pj ) represents the sampling point, W1 represents the height-speed window, h min and h max Respectively represent the minimum altitude and maximum altitude in altitude-speed, v min and v max Respectively represent the minimum speed and maximum speed in height-speed, n h and n v are all positive integers;

[0045] For the sampling points in the altitude-speed window, take different initial track angles:

[0046] γ pk =γ min +k(γ max -γ min ) / n γ ,k=0,1,2,…n γ (13)

[0047] Where, γ min and γ max Respectively represent the minimum and maximum values ​​of the initial track angle, n γ is a positive integer;

[0048] Combining equations (12) and (13), we get the parachute state set B p :

[0049] B p ={(h pi ,v pj ,γ pk )} (14)

[0050] Set B in the open state p The elements in are the initial state of the parachute segment. The maximum range optimization model (8) established in step 1 is used to solve the maximum range S after parachute opening.max , solve the minimum range S after parachute opening through the minimum range optimization model (9) established in step 1 min :

[0051]

[0052] Where Q max and Q min represent the maximum and minimum range optimization models, respectively;

[0053] This constructs a multi-dimensional constrained parachute opening window W2 characterized by altitude, speed, track angle, and extreme range:

[0054] W2={(h pi ,v pj ,γ pk ,S max ,S min )} (16)

[0055] Step 3: Based on the lander status and the multidimensional parachute opening window obtained in step 2, predict the maximum and minimum ranges of the planetary lander after parachute opening. Construct an evaluation index for the planetary lander's range control capability related to the maximum range, minimum range, and actual range to be flown. Use this index to establish a parachute opening logic criterion. Control the lander's parachute opening timing based on the parachute opening logic criterion, thus realizing the planetary lander parachute opening judgment based on the multidimensional constraint window.

[0056] The specific implementation method of step three is:

[0057] According to the multi-dimensional constrained parachute opening window W2 obtained in step 2, by constructing the limit range interpolation functions f1(·) and f2(·) as shown in formula (17), the maximum range S that can be reached by parachute opening in the current state is obtained: max With minimum range S min :

[0058]

[0059] Where h e 、v e , γ e represent the lander altitude, velocity, and track angle at the end of the entry segment, respectively.

[0060] Construct the planetary lander range control capability evaluation index as shown in formula (18):

[0061]

[0062] This evaluation index reflects the distance to be flown S togo and the range interval [S min ,S maxThe smaller J is, the closer the two are. At this time, opening the parachute can ensure that the lander can reach the desired landing area with the lowest fuel consumption cost to the greatest extent.

[0063] The parachute opening logic criterion is constructed as shown in formula (19):

[0064]

[0065] Where, ξ represents the parachute opening symbol, ξ = 1 means parachute opening, ξ = 0 means parachute not opening; X t and X t+1 Represent the lander state vector at the current moment and the next moment respectively, X t ∈W2 indicates that the lander state at the current moment is in the multidimensional constraint window, Indicates that the lander state at the next moment is not in the multidimensional constraint window. At the end of the entry phase, when the lander enters the multidimensional parachute opening window, each control cycle determines whether the parachute opening conditions are met according to the parachute opening logic criterion until the parachute opening conditions are met and the parachute opening command is executed.

[0066] Beneficial effects:

[0067] 1. To address the problem that the traditional two-dimensional altitude-speed parachute deployment window cannot achieve range constraints, the present invention discloses a planetary lander parachute deployment judgment method based on a multidimensional constraint window. This method considers the fuel consumption and thrust constraints during the powered descent phase, establishes a maximum and minimum range optimization model, uniformly samples within the traditional altitude-speed window, sets different initial track angles, solves the maximum and minimum ranges after parachute deployment, constructs a multidimensional constraint parachute deployment window characterized by altitude, speed, track angle, and extreme range, introduces the range to be flown into the parachute deployment judgment logic, achieves precise control of the range during the parachute deployment process, and improves the landing accuracy of the planetary lander.

[0068] 2. To address the need for both safe parachute deployment and precise landing during planetary landing, this paper discloses a planetary lander parachute deployment determination method based on a multidimensional constraint window. Based on a constructed multidimensional constraint parachute deployment window characterized by altitude, speed, track angle, and range limit, this method constructs an evaluation index for the planetary lander's range control capability and establishes parachute deployment determination logic. This method ensures smooth parachute deployment while minimizing fuel consumption, thus balancing the requirements of safe parachute deployment and precise landing.

[0069] 3. The planetary lander parachute deployment judgment method based on a multi-dimensional constraint window disclosed in the present invention can realize the safe parachute deployment and range control of the lander on the basis of achieving beneficial effects 1 and 2, improve the landing accuracy of the planetary lander, and have low fuel consumption. BRIEF DESCRIPTION OF THE DRAWINGS

[0070] Figure 1This is a flow chart of a planetary lander parachute deployment determination method based on a multi-dimensional constraint window disclosed in the present invention;

[0071] Figure 2 The maximum and minimum range flight trajectories for planetary landers;

[0072] Figure 3 Multi-dimensionally constrained parachute window profile for planetary landers;

[0073] Figure 4 This is the range change curve of the planetary lander before the parachute opens;

[0074] Figure 5 The landing trajectory comparison between the method proposed in this invention and the traditional altitude-speed window constraint method is shown. DETAILED DESCRIPTION

[0075] In order to better illustrate the purpose and advantages of the present invention, the invention is further described below with reference to the accompanying drawings and examples.

[0076] In order to verify the feasibility of the method, a planetary lander in the Mars exploration mission is used as an example to simulate the parachute deployment judgment method of the planetary lander based on the multi-dimensional constraint window. The altitude of the Mars lander at the end of the entry phase is h = 12 km, the speed is v = 500 m / s, the track angle is γ = -10°, and the range to be flown is S. togo =18km.

[0077] like Figure 1 As shown, the planetary lander parachute deployment determination method based on a multi-dimensional constraint window disclosed in this embodiment is specifically implemented in the following steps:

[0078] Step 1: In the planet-fixed coordinate system, establish the dynamic models of the planetary lander's entry phase, parachute phase, and powered descent phase; set the switching conditions between the parachute phase and the powered descent phase, consider the fuel consumption constraints and thrust constraints of the powered descent phase, and establish the maximum range and minimum range optimization problem models.

[0079] The specific implementation method of step one is:

[0080] Define the planet-fixed coordinate system O-XYZ: take the center of the planet as the origin O of the coordinate system, the OX axis points to the intersection of the equator and the prime meridian, the OZ axis points to the North Pole, and the OY axis obeys the right-hand rule.

[0081] The dynamic model of the lander entry phase is established in the planet-fixed coordinate system:

[0082]

[0083] Where r is the distance from the lander to the center of the planet, θ is the longitude, φ is the latitude, v is the lander speed, γ is the track angle, ψ is the heading angle (the angle between the flight speed and the north direction), and g = 3.72 m / s 2 is the gravitational acceleration of the Martian surface, L is the lift acceleration of the lander, and D is the drag acceleration of the lander.

[0084] The expressions of lift acceleration and drag acceleration of the lander are as follows:

[0085]

[0086] Where m e =4100kg is the mass of the lander during the entry phase, ρ is the density of the Martian atmosphere, S A =15.9m 2 is the reference area of ​​the lander, C L with C D The lift coefficient and drag coefficient of the lander are respectively.

[0087] The dynamic model of the lander parachute phase is established in the planet-fixed coordinate system:

[0088]

[0089] Where, is the lift acceleration of the lander after parachute opening, is the drag acceleration of the lander after parachute opening, and its expression is as follows:

[0090]

[0091] Where m p =3700kg is the mass of the parachute landing vehicle, is the reference area of ​​the lander after parachute opening, and The lift coefficient and drag coefficient of the lander after the parachute is opened.

[0092] The dynamic model of the lander's powered descent phase is established in the planet-fixed coordinate system:

[0093]

[0094] Where m T is the mass of the lander during the powered descent phase, T is the thrust amplitude, η is the thrust pitch angle (the angle between the thrust vector and the vertical direction), σ is the thrust direction angle (the angle between the projection of the thrust vector in the horizontal plane and the true north direction), I sp =225s represents the engine specific impulse, g0=9.80665m / s 2 is the acceleration due to gravity at sea level.

[0095] During the powered descent, the following thrust constraints are considered:

[0096] T min ≤T≤T max (6)

[0097] Where, T min =9944N and T max =26516N are the minimum thrust and maximum thrust of the lander engine respectively.

[0098] During the power reduction phase, the following fuel consumption constraints are considered:

[0099] m T0 -m Tf ≥m fuel (7)

[0100] Where m T0 =3300kg and m Tf are the lander masses at the initial and terminal moments of the powered descent phase, m fuel =500kg represents the fuel mass of the lander.

[0101] The switching condition between the parachute stage and the powered descent stage: when the lander altitude drops to h T =2km, the parachute is jettisoned and the engine is ignited, and the lander enters the powered descent phase from the parachute phase.

[0102] Considering the thrust constraints and fuel consumption constraints in the power descent phase shown in equations (6) and (7), the maximum range optimization model shown in equation (8) and the minimum range optimization model shown in equation (9) are established:

[0103]

[0104] Where S represents the distance traveled by the lander from parachute opening to landing, X = [r,θ,φ,v,γ,ψ] T represents the lander state vector, F p (·) and F T (·) represent the dynamic equations of the parachute landing stage and the powered descent stage, respectively, t p , t T and t f They represent the parachute opening time, engine ignition time and landing time respectively, h p 、v p and γ p They represent the altitude, speed and track angle of the lander at the initial moment of the parachute phase.

[0105] Step 2: Solve the altitude-velocity two-dimensional constrained parachute deployment window based on the planetary lander's parachute deployment Mach number, dynamic pressure, and altitude constraints. Sample uniformly within the altitude-velocity window, setting a different initial track angle for each sampling point. Using the range optimization model established in Step 1, the maximum and minimum ranges after parachute deployment are solved. This constructs a multi-dimensional constrained parachute deployment window characterized by altitude, speed, track angle, and extreme range.

[0106] The specific implementation method of step 2 is:

[0107] In order to ensure that the parachute can be opened safely and smoothly, the Mach number, dynamic pressure and altitude constraints must be met during opening, as shown in formula (10):

[0108]

[0109] Where M min =1.4 and M max =2.2 respectively represent the minimum Mach number and the maximum Mach number when the parachute is opened, M p represents the parachute opening Mach number, q min =300Pa and q max =850Pa respectively represents the minimum dynamic pressure and maximum dynamic pressure when the parachute is opened, q p Indicates the dynamic pressure of parachute opening, h min =6km indicates the minimum altitude when the parachute is opened.

[0110] According to the constraints shown in formula (10), the traditional height-speed parachute opening window can be obtained as shown in formula (11):

[0111]

[0112] Where c(h) is the speed of sound at height h.

[0113] Uniform sampling is performed in the height-speed window, as shown in Equation (12):

[0114]

[0115] Where, sample represents uniform sampling in the height-speed window, (h pi ,v pj ) represents the sampling point, W1 represents the height-speed window, h min =6km and h max =16.711km represents the minimum altitude and maximum altitude in the altitude-speed window, v min =309.6m / s and v max =487.0m / s represents the minimum speed and maximum speed in the height-speed, n h =30 and n v=20 are all positive integers.

[0116] For the sampling points in the altitude-speed window, different initial track angles are taken:

[0117] γ pk =γ min +k(γ max -γ min ) / n γ ,k=0,1,2,…n γ (13)

[0118] Where, γ min =-30° and γ max =-0° represents the minimum and maximum values ​​of the initial track angle, n γ =15 is a positive integer.

[0119] Combining equations (12) and (13), we get the parachute state set B p :

[0120] B p ={(h pi ,v pj ,γ pk )} (14)

[0121] Set B in the open state p The elements in are the initial state of the parachute segment. The maximum range optimization model (8) established in step 1 is used to solve the maximum range S after parachute opening. max , solve the minimum range S after parachute opening through the minimum range optimization model (9) established in step 1 min :

[0122]

[0123] Where Q max and Q min Represent the maximum and minimum range optimization models respectively.

[0124] This constructs a multi-dimensional constrained parachute opening window W2 characterized by altitude, speed, track angle, and extreme range:

[0125] W2={(h pi ,v pj ,γ pk ,S max ,S min )} (16)

[0126] Step 3: Based on the lander status and the multidimensional parachute opening window obtained in step 2, predict the maximum and minimum ranges of the planetary lander after parachute opening. Construct an evaluation index for the planetary lander's range control capability related to the maximum range, minimum range, and actual range to be flown. Use this index to establish a parachute opening logic criterion. Control the lander's parachute opening timing based on the parachute opening logic criterion, thus realizing the planetary lander parachute opening judgment based on the multidimensional constraint window.

[0127] The specific implementation method of step three is:

[0128] Assume that the height of the lander at the end of the entry phase is h = 12 km, the speed is v = 500 m / s, the track angle is γ = -10°, and the range to be flown is S togo =18km. According to the multi-dimensional constrained parachute opening window obtained in step 2, by constructing the limit range interpolation functions f1(·) and f2(·) as shown in formula (17), the maximum range S that can be reached by parachute opening in the current state can be obtained: max With minimum range S min :

[0129]

[0130] Where h e 、v e , γ e represent the lander altitude, velocity, and track angle at the end of the entry segment, respectively.

[0131] Construct the planetary lander range control capability evaluation index as shown in formula (18):

[0132]

[0133] This evaluation index reflects the distance to be flown S togo and the range interval [S min ,S max The smaller J is, the closer the two are. At this time, opening the parachute can ensure that the lander can reach the desired landing area with the lowest fuel consumption.

[0134] The parachute opening logic criterion is constructed as shown in formula (19):

[0135]

[0136] Where, ξ represents the parachute opening symbol, ξ = 1 means parachute opening, ξ = 0 means parachute not opening; X t and X t+1 Represent the lander state vector at the current moment and the next moment respectively, X t ∈W2 indicates that the lander state at the current moment is in the multidimensional constraint window, Indicates that the lander state at the next moment is not in the multidimensional constraint window. At the end of the entry phase, when the lander enters the multidimensional parachute opening window, each control cycle determines whether the parachute opening conditions are met according to the parachute opening logic criterion until the parachute opening conditions are met and the parachute opening command is executed.

[0137] Figure 2 The minimum and maximum range flight trajectories of a planetary lander with parachute opening at an altitude of 10 km, a speed of 400 m / s, and a track angle of -10° are given. The minimum and maximum ranges that the lander can reach under these conditions are 4.82 km and 12.06 km, respectively. Figure 3 A multi-dimensional constrained parachute opening window profile is given. For the convenience of presentation, only the corresponding track angle-range window in the altitude-speed two-dimensional window (400m / s, 10km) is given here. Figure 4 The curve of the lander's flight range change before the parachute opens is given. At the 7th second, the lander enters the multidimensional constraint window, but the performance indicators have not yet reached the minimum value; at the 19th second, the performance indicators reach the minimum value, and the lander has not left the multidimensional constraint window, the parachute opens, and the lander enters the parachute landing stage. Figure 5 A comparison chart of the landing trajectories of the proposed method and the traditional altitude-speed window constraint method is given. For the altitude-speed window constraint method, since only the altitude-speed window is considered and the range information is not taken into account, the lander parachute opens too early and cannot land accurately at the desired landing point. However, the lander using the proposed method delays parachute opening, taking into account both parachute opening safety and landing accuracy, thereby achieving range control.

[0138] The above specific description further illustrates the purpose, technical solutions and beneficial effects of the invention in detail. It should be understood that the above description is only a specific embodiment of the present invention and is not intended to limit the scope of protection of the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the scope of protection of the present invention.

Claims

1. A planetary lander parachute deployment determination method based on a multi-dimensional constraint window, characterized by: The following steps are included: Step 1: In the planet-fixed coordinate system, establish the dynamic models of the planetary lander's entry phase, parachute phase, and powered descent phase respectively; set the switching conditions between the parachute phase and the powered descent phase, consider the fuel consumption and thrust constraints of the powered descent phase, and establish the maximum range optimization problem model and the minimum range optimization problem model. The maximum range optimization problem model and the minimum range optimization problem model are collectively referred to as the range optimization problem model; Step 2: Solve the altitude-speed two-dimensional constrained parachute deployment window based on the Mach number, dynamic pressure, and altitude constraints of the planetary lander parachute deployment. Uniformly sample the altitude-speed window, set a different initial track angle for each sampling point, and solve the maximum and minimum ranges after parachute deployment using the range optimization problem model established in step 1. This constructs a multi-dimensional constrained parachute deployment window characterized by altitude, speed, track angle, and extreme range. Step 3: Based on the lander status and the multidimensional parachute opening window obtained in step 2, predict the maximum and minimum ranges of the planetary lander after parachute opening. Construct an evaluation index for the planetary lander's range control capability related to the maximum range, minimum range, and actual range to be flown. Use this index to establish a parachute opening logic criterion. Control the lander's parachute opening timing based on the parachute opening logic criterion, thus realizing the planetary lander parachute opening judgment based on the multidimensional constraint window.

2. The planetary lander parachute deployment determination method based on a multi-dimensional constraint window according to claim 1, characterized in that: The specific implementation method of step one is: Define the planet-fixed coordinate system O-XYZ: take the center of the planet as the origin O of the coordinate system, the OX axis points to the intersection of the equator and the prime meridian, the OZ axis points to the North Pole, and the OY axis obeys the right-hand rule; The dynamic model of the lander entry phase is established in the planet-fixed coordinate system: Where r is the distance from the lander to the center of the planet, θ is the longitude, φ is the latitude, v is the lander speed, γ is the track angle, ψ is the heading angle, g is the planet's gravity acceleration, L is the lander lift acceleration, and D is the lander drag acceleration; The expressions of lift acceleration and drag acceleration of the lander are as follows: Where m e is the mass of the lander during the entry phase, ρ is the atmospheric density, S A is the reference area of ​​the lander, C L with C D The lift coefficient and drag coefficient of the lander respectively; The dynamic model of the lander parachute phase is established in the planet-fixed coordinate system: Where, is the lift acceleration of the lander after parachute opening, is the drag acceleration of the lander after parachute opening, and its expression is as follows: Where m p is the mass of the lander during the parachute landing phase, is the reference area of ​​the lander after parachute opening, and The lift coefficient and drag coefficient of the lander after the parachute is opened respectively; The dynamic model of the lander's powered descent phase is established in the planet-fixed coordinate system: Where m T is the mass of the lander in the powered descent phase, T is the thrust amplitude, η is the thrust pitch angle, σ is the thrust direction angle, I sp represents the engine specific impulse, g0 represents the acceleration of gravity at sea level; During the powered descent, the following thrust constraints are considered: T min ≤T≤T max (6) Where, T min and T max are the minimum and maximum thrust of the lander engine respectively; During the power reduction phase, the following fuel consumption constraints are considered: m T0 -m Tf ≥m fuel (7) Where m T0 and m Tf are the lander masses at the initial and terminal moments of the powered descent phase, m fuel represents the lander fuel mass; The switching condition between the parachute stage and the powered descent stage: when the lander altitude drops to h T When the parachute is jettisoned and the engine is ignited, the lander enters the powered descent phase from the parachute phase; Considering the thrust constraints and fuel consumption constraints in the power descent phase shown in equations (6) and (7), the maximum range optimization model shown in equation (8) and the minimum range optimization model shown in equation (9) are established: Where S represents the distance traveled by the lander from parachute opening to landing, X = [r,θ,φ,v,γ,ψ] T represents the lander state vector, F p (·) and F T (·) represent the dynamic equations of the parachute landing stage and the powered descent stage, respectively, t p , t T and t f They represent the parachute opening time, engine ignition time and landing time respectively, h p 、v p and γ p They represent the altitude, speed and track angle of the lander at the initial moment of the parachute phase.

3. The planetary lander parachute deployment determination method based on a multi-dimensional constraint window according to claim 2, characterized in that: The specific implementation method of step 2 is: In order to ensure that the parachute can be opened safely and smoothly, the Mach number, dynamic pressure and altitude constraints must be met during opening, as shown in formula (10): Where M min and M max Respectively represent the minimum Mach number and maximum Mach number when the parachute is opened, M p represents the parachute opening Mach number, q min and q max They represent the minimum dynamic pressure and maximum dynamic pressure when the parachute is opened, q p Indicates the dynamic pressure of parachute opening, h min Indicates the minimum altitude when the parachute is opened; According to the constraints shown in formula (10), the traditional height-speed parachute opening window is obtained, as shown in formula (11): Where c(h) represents the speed of sound at height h; Uniform sampling is performed in the height-speed window, as shown in Equation (12): Where, sample represents uniform sampling in the height-speed window, (h pi ,v pj ) represents the sampling point, W1 represents the height-speed window, h min and h max Respectively represent the minimum altitude and maximum altitude in altitude-speed, v min and v max Respectively represent the minimum speed and maximum speed in height-speed, n h and n v are all positive integers; For the sampling points in the altitude-speed window, take different initial track angles: c pk =c min +k(γ max -c min ) / n γ ,k=0,1,2,…n γ (13) Where, γ min and γ max Respectively represent the minimum and maximum values ​​of the initial track angle, n γ is a positive integer; Combining equations (12) and (13), we get the parachute state set B p : B p ={(h pi ,v pj ,c pk )} (14) Set B in the open state p The elements in are the initial state of the parachute segment. The maximum range optimization model (8) established in step 1 is used to solve the maximum range S after parachute opening. max , solve the minimum range S after parachute opening through the minimum range optimization model (9) established in step 1 min : Where Q max and Q min represent the maximum and minimum range optimization models, respectively; This constructs a multi-dimensional constrained parachute opening window W2 characterized by altitude, speed, track angle, and extreme range: W2={(h pi ,v pj ,γ pk ,S max ,S min )} (16).

4. The planetary lander parachute deployment determination method based on a multi-dimensional constraint window according to claim 3, characterized in that: The specific implementation method of step three is: According to the multi-dimensional constrained parachute opening window W2 obtained in step 2, by constructing the limit range interpolation functions f1(·) and f2(·) as shown in formula (17), the maximum range S that can be reached by parachute opening in the current state is obtained: max With minimum range S min : Where h e 、v e , γ e represent the lander altitude, velocity, and track angle at the end of the entry segment, respectively; Construct the planetary lander range control capability evaluation index as shown in formula (18): This evaluation index reflects the distance to be flown S togo and the range interval [S min ,S max The smaller J is, the closer the two are. At this time, opening the parachute can ensure that the lander can reach the desired landing area with the lowest fuel consumption cost to the greatest extent. The parachute opening logic criterion is constructed as shown in formula (19): Where, ξ represents the parachute opening symbol, ξ = 1 means parachute opening, ξ = 0 means parachute not opening; X t and X t+1 Represent the lander state vector at the current moment and the next moment respectively, X t ∈W2 indicates that the lander state at the current moment is in the multidimensional constraint window, It indicates that the lander state at the next moment is not in the multidimensional constraint window; at the end of the entry phase, when the lander enters the multidimensional parachute opening window, each control cycle judges whether the parachute opening conditions are met according to the parachute opening logic criterion until the parachute opening conditions are met and the parachute opening command is executed.

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