Electric Propulsion Orbital Transfer Thrust Control
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Solution Overview
Problem
Existing methods for electrically propelled orbital transfer of spacecraft are complex and do not effectively address propellant savings, mission life prolongation, and radiation exposure management.
Innovation Solution
A method involving alternating nominal and reduced thrust steps during orbital transfer, using electric propulsion means with fixed and deployable thrusters, to optimize propellant consumption and transfer duration, while managing radiation exposure.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Productivity
If continuous nominal thrust is applied during orbital transfer, then transfer duration is reduced, but propellant consumption increases
Solution Approach 1:
The patent applies periodic action by alternating between nominal thrust steps and reduced thrust steps during orbital transfer. The method divides the transfer trajectory into multiple orbital arcs, applying nominal thrust during arcs where the spacecraft is at higher altitudes (lower orbital velocity) and reduced thrust during arcs at lower altitudes (higher orbital velocity), creating a periodic thrust pattern that optimizes both time and propellant usage
Solution Approach 2:
The patent implements dynamics by making the thrust level variable rather than constant. The control system dynamically adjusts between nominal and reduced thrust based on the spacecraft's position along the orbital arc, specifically switching thrust levels according to whether the spacecraft is traversing the ascending or descending portion of its orbit, thereby adapting the propulsion force to the instantaneous orbital conditions
2Loss of substance
If electric propulsion means are used for orbital transfer, then propellant consumption is reduced, but transfer duration increases
Solution Approach 1:
The periodic alternation between nominal and reduced thrust allows the electric propulsion system to achieve transfer durations that are acceptable for mission requirements while maintaining the propellant efficiency benefits of electric propulsion. The method finds an optimal balance by applying maximum thrust only when most beneficial for reducing transfer time
Solution Approach 2:
The patent changes the thrust parameter dynamically during the transfer process. By varying the thrust level between nominal and reduced states based on orbital position, the system optimizes the trade-off between transfer duration and propellant consumption, allowing electric propulsion to achieve both time and mass efficiency
3Loss of time
If thrust is increased to minimize transfer time, then radiation exposure increases, but propellant savings decrease
Solution Approach 1:
The periodic thrust pattern allows the spacecraft to spend portions of its orbit coasting without thrust, thereby reducing cumulative radiation exposure while still achieving acceptable transfer times. The method alternates between active propulsion phases and coasting phases, minimizing the time spent in high-radiation environments
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 method reduces propellant consumption, prolongs mission life, and minimizes radiation exposure by optimizing thrust patterns and energy management during the orbital transfer process.
Implementation Method 1
a nominal thrust step during which the propulsion means generate a nominal thrust and a reduced thrust step during which the propulsion means are partly stopped or slowed
Data Source
AI summary
A method for transferring a spacecraft (10), such as an artificial satellite, from an initial elliptical orbit (30) to a final geostationary orbit (50), the spacecraft taking at least one intermediate elliptical orbit (40) propelled by electric propulsion means (12, 13), the method includes: when the spacecraft is in an intermediate orbit, a nominal thrust step (410) in which the propulsion means generate nominal thrust while the spacecraft is on at least part of a first orbital arc (41) passing through the apogee A of the intermediate orbit, and a minimum thrust step (420), in which the propulsion means are partly stopped or slowed while the spacecraft is on at least part (43) of a second orbital arc (42) passing through the perigee P of the intermediate orbit, the two orbital arcs being complementary.