Aerospace vehicle orbit transfer method for returning to atmosphere based on energy conservation

Abstract

The invention belongs to the technical field of orbit transfer design for an aerospace vehicle to return to the atmosphere, and particularly relates to an orbit transfer method for the aerospace vehicle to return to the atmosphere based on energy conservation, and the method comprises the steps: 1, calculating a dimensionless preset entry speed ue; step 2, calculating a relative dimensionless distance lambda of a deorbit point; 3, judging whether a tangential orbital transfer feasibility critical condition is met or not, and if the tangential orbital transfer feasibility critical condition is met, executing tangential orbital transfer; and if the tangential orbital transfer feasibility critical condition is not met, executing a grazing incidence strategy.

Description

Technical Field

[0001] This application belongs to the field of aerospace vehicle reentry orbit change design technology, specifically involving an aerospace vehicle reentry orbit change method based on energy conservation. Background Technology

[0002] After completing its mission, the spacecraft needs to change its orbit from a non-intersecting orbit to an orbit that intersects with the atmosphere, re-enter the atmosphere, and be recovered.

[0003] Currently, methods for maneuvering aerospace vehicles re-entering the atmosphere typically involve adjusting velocity via rocket pulses. While this involves orbital energy analysis, it lacks a clear minimum impulse optimization strategy for a given entry velocity scenario, resulting in the following shortcomings:

[0004] High fuel consumption and failure to systematically utilize the principle of energy conservation to optimize the direction of velocity pulses lead to unnecessary fuel loss.

[0005] The parameter calculations are complex, and the original physical quantities, such as meters and seconds, are used directly for orbit calculations without introducing a dimensionless model to simplify the analysis.

[0006] The lack of critical conditions and the absence of a feasibility assessment for tangential trajectory change may lead to trajectory change failure or exceeding the entry angle limit.

[0007] This application is made in view of the aforementioned technical deficiencies. Summary of the Invention

[0008] The purpose of this application is to provide a method for re-entering the atmosphere and changing orbit based on energy conservation of aerospace vehicles, so as to overcome or mitigate at least one of the known technical defects.

[0009] The technical solution of this application is:

[0010] A method for orbital maneuvering of a spacecraft re-entering the atmosphere based on energy conservation includes:

[0011] Step 1: Calculate the dimensionless preset entry velocity u e :

[0012]

[0013]

[0014] Among them, V e Preset entry speed; V cir r is the local circular orbital velocity; μ is the Earth's gravitational constant; r D The distance from the deorbit point D to the Earth's center;

[0015] Step 2: Calculate the dimensionless distance λ between the off-track point and the orbital point:

[0016]

[0017] Where R is the distance from the atmospheric inlet E to the Earth's center;

[0018] Step 3: Determine whether the critical conditions for tangential trajectory change feasibility are met. The critical conditions for tangential trajectory change feasibility are: or or Where γ1 is the initial flight path angle; α1 is the initial orbital semi-major axis; and e1 is the initial orbital eccentricity.

[0019] If the critical conditions for the feasibility of tangential trajectory change are met, then the tangential trajectory change shall be performed;

[0020] If the critical condition for tangential trajectory change feasibility is not met, then the grazing incidence strategy will be executed.

[0021] According to at least one embodiment of this application, in the above-described energy conservation-based aerospace vehicle reentry orbit change method,

[0022] Where a1 is the initial semi-major axis of the orbit; θ D The true perimeter angle of the derailment point.

[0023] According to at least one embodiment of this application, in the above-described energy conservation-based aerospace vehicle reentry orbit change method,

[0024] According to at least one embodiment of this application, in the above-described energy-conservation-based aerospace vehicle reentry orbit change method, when performing a tangential orbit change, γ2 = γ1, the dimensionless pulse amplitude Δu = u1 - u2, and the velocity pulse... Where u1 is the initial dimensionless velocity; u2 is the preset entry dimensionless velocity;

[0025] or Where V1 is the initial velocity;

[0026] or, Where V2 is the derailment speed.

[0027] According to at least one embodiment of this application, in the above-described energy-conservation-based aerospace vehicle reentry orbit change method, a grazing incidence strategy is executed to force the orbit to be tangent to the atmosphere, thereby making the critical flight path angle γ e =0, Where γ2 is the deorbit flight path angle. Attached Figure Description

[0028] Figure 1This is a schematic diagram of an aerospace vehicle re-entering the atmosphere and changing its trajectory, as provided in an embodiment of this application.

[0029] Figure 2 This is a schematic diagram of the velocity space geometry provided in the embodiments of this application;

[0030] Figure 3 This is a schematic diagram of the energy conservation-based aerospace vehicle reentry orbit change method provided in the embodiments of this application.

[0031] To better illustrate this embodiment, some content in the accompanying drawings may be omitted, enlarged, or reduced. They are for illustrative purposes only and should not be construed as limiting the scope of this application. Detailed Implementation

[0032] To make the technical solution and advantages of this application clearer, the technical solution of this application will be described in a clearer and more complete manner below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of this application, and are only used to explain this application, not to limit this application. It should be noted that, for ease of description, only the parts related to this application are shown in the accompanying drawings, and other related parts can be referred to the general design.

[0033] Furthermore, unless otherwise defined, the technical or scientific terms used in this application description shall have the ordinary meaning understood by one of ordinary skill in the art to which this application pertains. The word "comprising" as used in this application description indicates that the concept preceding the word encompasses the concepts listed following the word and their equivalents, without excluding other related concepts.

[0034] Spacecraft re-entering the atmosphere and changing orbit, such as Figure 1 As shown, the velocity space geometric relationship is as follows: Figure 2 As shown.

[0035] Energy conservation equation for spacecraft:

[0036]

[0037] Where V2 is the deorbit velocity of the spacecraft; μ is the Earth's gravitational constant; r D V is the distance from the Earth's center to the deorbit point D of the spacecraft; e The preset entry velocity is given to the spacecraft; R is the atmospheric entrance; E is the distance from the Earth's center; and E is the Earth's radius.

[0038] Introducing local circular orbit velocity Perform dimensionless processing:

[0039]

[0040] Where V1 is the initial velocity of the spacecraft; u1, u2, ue The dimensionless velocity corresponding to the initial velocity, deorbit velocity, and preset entry velocity of the spacecraft.

[0041] Introducing dimensionless distance Simplify:

[0042] Further organization is needed:

[0043] have:

[0044]

[0045] Wherein, the dimensionless distance λ is the dimensionless distance between the off-track point and the other point.

[0046] Formula (3) shows that the amplitude of the dimensionless derailment velocity u2 is determined solely by the dimensionless preset entry velocity u. e The distance is determined by the dimensionless distance λ between the deorbit point and the orbital direction, and is independent of the orbital direction.

[0047] Velocity pulse calculation and minimum impulse optimization conditions for tangential trajectory change strategy:

[0048] Geometric relationship of velocity triangle:

[0049] In the velocity space Dxy, the x-axis is the tangential direction of the orbit, the y-axis is the radial direction, and the initial velocity is dimensionless. With derailment speed Satisfy vector relationship:

[0050]

[0051] The dimensionless pulse amplitude Δu is calculated using the law of cosines:

[0052]

[0053] Where γ1 is the initial flight path angle, and γ2 is the angle between the velocity and tangential direction at the deorbit point; γ2 is the deorbit flight path angle.

[0054] For aerospace vehicles to perform tangential trajectory changes, only the speed needs to be altered, resulting in minimal pulse amplitude. When γ2 = γ1:

[0055] Δu min =|u1-u2|(6)

[0056] For tangential orbit change to be performed, the critical conditions for its feasibility must be met:

[0057]

[0058] Combining the initial orbital flight path angle formula:

[0059]

[0060] in, a1 is the semi-major axis of the initial orbit; e1 is the initial orbital eccentricity.

[0061] Simplifying equation (7), we obtain the critical condition:

[0062] The results were:

[0063]

[0064] Compared to conventional experience, the minimum impulse strategy can save fuel and extend the life of the aircraft. For example, in a tangential trajectory change with a given entry velocity, the velocity impulse ΔV is simply the dimensionless velocity difference. |u1-u2| avoids the extra consumption of the normal component.

[0065] In extreme cases, if the critical condition (9) for the feasibility of tangential trajectory change is not satisfied, i.e., the orbital intersection cannot be achieved by simply changing the velocity magnitude, then a grazing incidence strategy is required to force the orbit to be tangent to the atmosphere, i.e., the critical flight path angle γ. e =0, at this point, the velocity pulse ΔV reaches the minimum value of this constraint, preventing the spacecraft from jumping over the atmosphere or burning up due to excessive angle. At this point:

[0066]

[0067] The orbital eccentricity is:

[0068]

[0069] At this moment, the orbit is exactly tangent to the atmospheric boundary, and the eccentricity is determined by the dimensionless preset velocity u. e The dimensionless distance λ between the point of departure and the point of departure is uniquely determined.

[0070] When efficient tangential maneuvering is not feasible, grazing incidence is used as a fallback strategy. By forcing orbital intersection through geometric critical conditions, the safe return of the spacecraft can be ensured. Therefore, the grazing incidence strategy is an extreme but crucial technical means for spacecraft to re-enter the atmosphere. Critical entry is achieved through precise control of the flight path angle, balancing thermal management requirements and fuel efficiency. Its core lies in the critical conditions of angular momentum and energy conservation. Trajectory optimization is achieved through mathematical modeling and real-time control, making it one of the key technologies for reusable spacecraft.

[0071] Based on the above, this application provides a method for orbital maneuvering of aerospace vehicles upon atmospheric reentry based on energy conservation, such as... Figure 3As shown, by using a dimensionless velocity model and a tangential orbit change strategy, the required velocity pulse for orbit change is minimized under a given entry velocity. The deorbit velocity amplitude is derived through the energy conservation equation, and the optimal pulse direction is determined using critical conditions. This approach balances efficiency and safety and is suitable for orbit change of aerospace vehicles such as satellite recovery and manned spacecraft reentry vehicles.

[0072] Step 1: Calculate the dimensionless preset entry velocity u e .

[0073]

[0074]

[0075]

[0076] Among them, V e Preset entry speed; V cir r is the local circular orbital velocity; μ is the Earth's gravitational constant; r D θ is the distance from the de-orbit point D to the Earth's center; a1 is the initial semi-major axis of the orbit; e1 is the initial orbital eccentricity; θ D The true perimeter angle of the derailment point.

[0077] Step 2: Calculate the dimensionless distance λ between the off-track point and the orbital point.

[0078]

[0079] Where R is the distance from the atmospheric inlet E to the Earth's center.

[0080] Step 3: Determine whether the critical conditions for tangential trajectory change feasibility are met. The critical conditions for tangential trajectory change feasibility are: or or Where γ1 is the initial flight path angle; The initial orbit is relative to the semi-major axis.

[0081] If the critical condition for tangential trajectory change feasibility is met, then the tangential trajectory change is performed, with γ2 = γ1, dimensionless pulse amplitude Δu = u1 - u2, and velocity pulse. Where u1 is the initial dimensionless velocity; u2 is the preset entry dimensionless velocity.

[0082] or Where V1 is the initial velocity.

[0083] or, Where V2 is the derailment speed.

[0084] Update track parameters, semi-major axis

[0085] If the critical conditions for tangential trajectory change are not met, a grazing incidence strategy is employed to force the orbit to be tangential to the atmosphere, thus setting the critical flight path angle. Where γ2 is the deorbit flight path angle.

[0086] Update orbital parameters, orbital eccentricity

[0087] In a specific example, a1 = R, a circular track, e1 = 0, θ D =π, the point of departure is the apogee, R = 6371km, λ = r D / R=1.2,V e =8km / s, the energy conservation-based aerospace vehicle reentry orbit change method disclosed in the above embodiments is implemented as follows.

[0088] Dimensionless calculation:

[0089]

[0090]

[0091]

[0092] Critical condition judgment:

[0093] Taking the absolute value, it is actually 1, the circular orbit γ1=0, the left side of equation (9) If the value on the right is much greater than 1, the condition is met, and a tangential trajectory change is executed.

[0094] Δu = 0.894 - 0.775 = 0.119, corresponding to the actual pulse.

[0095] The above-described method for reentry orbit change of a spacecraft based on energy conservation derives a dimensionless velocity relationship through the energy conservation equation, establishes critical conditions for tangential orbit change by combining the velocity triangle geometry principle, and minimizes the velocity pulse required for orbit change of the spacecraft under a given atmospheric entry velocity constraint by using the energy conservation principle and tangential orbit change strategy, thereby achieving efficient fuel utilization and precise orbit control.

[0096] The above-described embodiment discloses an energy-conservation-based method for aerospace vehicles to re-enter the atmosphere. Through a tangential trajectory change strategy, it directly employs the minimum pulse Δu = |u1 - u2| when critical conditions are met. This significantly reduces fuel consumption compared to existing methods and introduces a dimensionless model (u... iThe calculation of orbital parameters is simplified to algebraic operations, which is suitable for real-time calculation in embedded control systems. It can also automatically switch orbit change modes through critical conditions, which can balance the efficiency of tangential orbit change and the safety of grazing incidence, and avoid recovery failure caused by exceeding the entry angle limit.

[0097] The technical solution of this application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of this application is obviously not limited to these specific embodiments. Without departing from the principles of this application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will all fall within the scope of protection of this application.

Claims

1. A method for orbit change during atmospheric reentry of a spacecraft based on energy conservation, characterized in that, include: Step 1: Calculate the dimensionless preset entry velocity u e : Among them, V e Preset entry speed; V cir r is the local circular orbital velocity; μ is the Earth's gravitational constant; r D The distance from the deorbit point D to the Earth's center; Step 2: Calculate the dimensionless distance λ between the off-track point and the orbital point: Where R is the distance from the atmospheric inlet E to the Earth's center; Step 3: Determine whether the critical conditions for tangential trajectory change feasibility are met. The critical conditions for tangential trajectory change feasibility are: or or Where γ1 is the initial flight path angle; α1 is the initial orbital semi-major axis; and e1 is the initial orbital eccentricity. If the critical conditions for the feasibility of tangential trajectory change are met, then the tangential trajectory change shall be performed; If the critical condition for tangential trajectory change feasibility is not met, then the grazing incidence strategy will be executed.

2. The method for re-entering the atmosphere and changing orbit of a spacecraft based on energy conservation according to claim 1, characterized in that, Where a1 is the initial semi-major axis of the orbit; θ D The true perimeter angle of the derailment point.

3. The method for re-entering the atmosphere and changing orbit of a spacecraft based on energy conservation according to claim 2, characterized in that, 4. The method for re-entering the atmosphere and changing orbit of a spacecraft based on energy conservation according to claim 3, characterized in that, When performing a tangential trajectory change, γ2 = γ1, the dimensionless pulse amplitude Δu = u1 - u2, and the velocity pulse... Where u1 is the initial dimensionless velocity; u2 is the preset entry dimensionless velocity; or Where V1 is the initial velocity; or, Where V2 is the derailment speed.

5. The method for re-entering the atmosphere and changing orbit of a spacecraft based on energy conservation according to claim 4, characterized in that, By employing a grazing incidence strategy, the orbit is forced to be tangent to the atmosphere, resulting in a critical flight path angle γ. e =0, Where γ2 is the deorbit flight path angle.

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