Dual-Battery Spacecraft Power Architecture for Peak Payload Loads
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Solution Overview
Problem
Conventional spacecraft electrical power subsystems are often undersized for high peak power payloads, lack adaptability with different payload components, and struggle with heat rejection and short-term power sourcing, especially during intermittent operations in various orbital conditions.
Innovation Solution
A dual battery configuration with a low rate discharge battery for the spacecraft bus and a high rate discharge battery for the payload module, managed by a battery charge management unit, allowing for optimal charging and discharging based on mission requirements, along with an independent thermal management system for the payload module using thermal radiators and phase change materials.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Power
If a conventional undersized electrical power subsystem is used, then the spacecraft bus can maintain simpler architecture, but it cannot provide sufficient power for high peak power payloads with high discharge currents
Solution Approach 1:
The electrical power subsystem is segmented into two independent battery circuits: a first battery circuit for the spacecraft bus and a second battery circuit for the payload module. This segmentation allows each circuit to be optimized independently - the second circuit can provide high discharge currents for peak power payloads without compromising the stability of the first circuit, thereby resolving the contradiction between power supply capability and system complexity.
2Adaptability or versatility
If a self-contained battery without spacecraft monitoring or control is used, then the system can operate independently, but it lacks modularity and cannot be adapted with different payload components
Solution Approach 1:
The second battery circuit is designed with universal interfaces and control systems that can accommodate different payload components. The battery management unit can monitor and control various payload types, and the circuit can be coupled to different solar array configurations, providing adaptability across multiple payload scenarios while maintaining manageable system integration through standardized interfaces.
3Temperature
If conventional subsystems are used, then the architecture can remain simple, but they cannot reject the heat generated by the power subsystem effectively
Solution Approach 1:
The thermal management system is segmented into separate thermal control paths for each battery circuit. The first battery circuit uses the spacecraft bus thermal management system, while the second battery circuit has its own dedicated thermal management capabilities. This segmentation allows high-power payload operations to generate heat that can be managed independently without compromising the thermal stability of the spacecraft bus, resolving the contradiction between heat rejection capability and system complexity.
4Power
If conventional subsystems are used, then the system can maintain fixed configuration, but it cannot source short-term electrical power for increased payload operation during mission
Solution Approach 1:
The electrical power subsystem implements dynamic power allocation through independent control of the two battery circuits. The second battery circuit can be dynamically activated to provide short-term high power bursts for increased payload operation when needed, while the first circuit maintains stable bus power. This dynamic capability allows the system to adapt power delivery to mission requirements, resolving the contradiction between short-term power sourcing and mission flexibility.
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
Enables flexible and efficient power supply to both low and high-discharge payloads, supports various mission demands, and allows for modular and independent testing of the payload module, improving adaptability and reliability in spacecraft operations.
Implementation Method 1
a heat sink or transient thermal absorber, such as a phase change material
Implementation Method 2
The thermal management system may include at least one thermal radiator
Data Source
AI summary
An electrical power system has a dual battery configuration that enables sufficient power supply for a spacecraft bus and a payload module being carried by the spacecraft. During a sunlight power mode, power is drawn from a solar array of the bus to power a low-discharge payload of the spacecraft and a high-discharge payload of a payload module. During the sunlight power mode, a low rate discharge battery and a high rate discharge battery are charged by a battery charge management unit of the spacecraft bus. During an eclipse power mode, the low rate discharge battery powers the low-discharge payload of the spacecraft and the high rate discharge battery powers the high-discharge payload of the payload module. The high-rate discharge battery may also be used to power the high-rate discharge payload in the sunlight power mode to meet its high current demands to meet a flexible mission operations.