Satellite Orbit Rotation for Eclipse Mass Reduction
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Solution Overview
Problem
Satellites in highly elliptical orbits around massive bodies face long-lasting eclipses, which require oversized power and thermal regulation subsystems, increasing mass and payload competition, and existing solutions like increasing battery size are insufficient or mass-intensive.
Innovation Solution
A method and device that calculate a second elliptical orbit by rotating the initial orbit around its apsides, determining a maneuver to reach this new orbit, and optimizing battery mass and fuel mass to reduce overall satellite mass, allowing the satellite to operate during eclipses with less battery mass and additional fuel.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Reliability
If the battery capacity is increased to maintain operation during long eclipses, then the satellite can operate during long-lasting eclipses, but the mass and volume of the satellite increase
Solution Approach 1:
The patent applies dynamics by making the orbital parameters variable rather than fixed. The satellite performs maneuvers to transition between different elliptical orbits with varying eclipse durations, allowing the system to adapt its operational characteristics dynamically. This resolves the contradiction by enabling the satellite to operate during long eclipses only when necessary, rather than requiring permanent oversizing of power subsystems.
Solution Approach 2:
The patent changes the orbital parameters (semi-major axis, eccentricity, inclination) to modify the eclipse duration. By adjusting these parameters through controlled maneuvers, the satellite can transition from orbits with long eclipses to orbits with shorter eclipses, thereby reducing the required battery capacity while maintaining operational reliability when needed.
2Reliability
If the battery capacity is increased to maintain operation during long eclipses, then the satellite can operate during long-lasting eclipses, but the volume occupied by batteries increases
Solution Approach 1:
The patent makes the orbital configuration dynamic, allowing transitions between orbits with different eclipse characteristics. This enables the satellite to minimize battery volume by only preparing for long eclipses when actually encountered, rather than permanently carrying excessive battery capacity that would occupy significant volume.
Solution Approach 2:
By changing orbital parameters through maneuvers, the satellite can reduce eclipse duration and consequently reduce the required battery volume. The system adapts its orbital parameters to match mission requirements, optimizing the balance between operational reliability and volume constraints.
3Weight of moving object
If the satellite performs a maneuver to change orbit and reduce eclipse duration, then the battery mass can be reduced, but fuel mass must be added
Solution Approach 1:
The patent applies preliminary action by performing orbit-changing maneuvers before long eclipses occur, or at strategically optimal times when the satellite's velocity is lower (near apoapsis). This timing allows the maneuvers to be executed more efficiently, minimizing the fuel required while achieving the goal of reducing eclipse duration and battery mass.
Solution Approach 2:
The patent optimizes the orbital parameters (particularly the inclination and ascending node) to achieve the desired reduction in eclipse duration. By carefully selecting the maneuver parameters and execution timing, the system minimizes the fuel mass required while achieving sufficient reduction in battery mass requirements.
Applied Scientific Principles
This section explains which scientific principles are used to turn an abstract innovation direction into a practical engineering solution.
Function Achieved in This Case
This approach reduces the satellite's mass by exchanging battery mass for fuel mass necessary for the orbit change maneuver, optimizing design and enabling operation during long eclipses while conserving mass for other scientific applications.
Implementation Method 1
a step of calculating a second elliptical orbit obtained by rotation of the first orbit around an axis connecting the periapsis and the apoapsis
Data Source
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AI summary
The invention relates to a method and a device for optimizing the mass of a satellite. The method comprises: a step (201) of calculating a second elliptical orbit (107) obtained by rotating the first orbit (103) around an axis (106) connecting the periapsis (104) and the apoapsis (105), the second elliptical orbit (107) being associated with a second maximum eclipse duration (D2) shorter than the first maximum eclipse duration (D1), a step (202) of determining a maneuver enabling the satellite to reach the second orbit (107), and a step of calculating (203) a second battery mass (Mb2) enabling the satellite to remain operational during the second maximum eclipse duration (D2), and of calculating a fuel mass (Mc) required to perform the maneuver.