Spacecraft Electrical Power Systems
All spacecraft have a variety of systems that control operations such as guidance, environmental control, and communications. In turn, each of those systems relies on a stable electrical power supply. While some spacecraft electrical power systems or EPS rely on the use of solar panels, others rely on fuel cells and batteries contained within the spacecraft. The selection of a main energy source for spacecraft depends on the required power for the mission and the duration of the mission. Regardless of the type of system, every spacecraft electrical power system employs power conditioning, distribution, and power management.
Electrical Power System for Apollo Command and Service Modules
As designed, the EPS for the Apollo spacecraft delivered a nominal 28 vdc. The Apollo Block I and Block II command and service modules utilized an electrical power system that operated from any combination of seven dc sources. Those sources included:
- Three fuel cells – 575 kilowatt-hours each
- Three silver-oxide entry batteries – 40 ampere-hours each
Inverters in the system derived three-phase 400 hz 115 vac for the operation of all the ac power required by the system. The major portion of the generated ac voltage powered the fuel cell pump motors.
After the failure of the cryogenic oxygen system during the Apollo 13 mission, a 400 ampere-hour battery was added to the service module. The battery could provide 12 kilowatt hours of emergency power through the command module main buses.
Primary Source
When considering the operation of the electrical power system in the Apollo spacecraft, the three fuel cells located in the service module provided the primary source of power. Two of the three entry batteries located in the command module supplemented the fuel cells during high energy demand periods.
Basic Distribution System
The spacecraft used two redundant buses and a single point ground connected to the spacecraft structure for distribution. Fuel cells energized two main dc buses designated as Main Bus A and Main Bus B and the entry and post landing A, B, and C batteries. The respective entry and post landing batteries powered battery buses A and B. Battery C connected to either or both buses during the failure of batteries A or B.
Both main dc buses energized the flight and post landing bus through diode pairs. Pyrotechnic batteries powered the pyrotechnic A and B buses. Both the pyrotechnic buses remained isolated from the main electrical buses by a normally-open switch.
In the Apollo spacecraft, one or two solid state inverters supplied ac power and produced 1250 volt-amps each. Inverters 1 and 2 powered through Main Bus A and B while inverter 3 powered through either Main Bus A or Main Bus B. Six motor switches operated as the ac control and either connected or disconnected the inverters from the ac buses. This operation prevented the simultaneous connection of the inverters to the same ac bus. Any over-voltage or overload condition resulted in the automatic disconnect of the inverters.
References
A Historical Overview of the Electrical Power Systems in the US Manned and Some US Unmanned Spacecraft. Fenn College of Engineering. Cleveland State University. 1985
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