Robust, Radiation Tolerant Command and Data Handling and Power System Electronics from NASA Goddard Space Flight Center

In today\u27s rapidly advancing technology roadmap for space applications there is an emphasis on completing missions faster and cheaper than previous large-scale missions at the National Aeronautics and Space Administration (NASA) such as the Lunar Reconnaissance Orbiter (LRO) or the Magnetospheric Multiscale (MMS) mission. As part of this effort, focus has shifted from using mostly radiation-tolerant or radiation-hardened parts to more commercial-off-the-shelf (COTS) components for missions that can last at least one year in orbit. However, there are some portions of a spacecraft\u27s avionics, such as the command and data handling (C&DH) system and the Electrical Power Systems (EPS) that need to have some level of predictable reliability that goes beyond the capabilities of currently available COTS parts. While there are a number of COTS components that can withstand a total ionizing dose (TID) of tens or hundreds of kilorads, there is still a great deal of concern about tolerance to and mitigation of single-event effects (SEE). The Goddard Modular Smallsat Architecture (GMSA) is based on an initiative at NASA Goddard Space Flight Center (GSFC) to address this reliability issue along with minimizing cost and schedule challenges. The goal is to develop a modular, flexible, and extensible small satellite implementation approach that can accommodate spacecraft subsystems that are designed both internally within NASA and externally. There will also be an emphasis placed on maximizing the volume available for science payloads. This paper provides details a new technology development effort that will add a miniaturized board-level capability that will fulfill the requirements of both the C&DH and EPS within a 6U (10cm x 20cm x 30cm) satellite. This effort will dramatically transform spaceflight systems capabilities at GSFC with the goal of accomplishing a variety of ambitious science goals that are increasingly challenged by constraints on cost, mass, power, volume, and schedule. The selected topology for the EPS is a Direct Energy Transfer (DET) system with the battery connected directly to the bus. The shunt control technique is a linear sequential full shunt which provides a simple solar array interface and can support both 3-axis stabilized and spinner satellites. The EPS includes all the circuits needed to perform telemetry and command function using I2C interface with the C&DH. In addition, the EPS will be designed, tested, and verified to meet launcher vehicle safety requirement. The C&DH functionality will be implemented using the Smallsat Common Electronics Board (SCEB) and its adapter board. The SCEB can be configured to implement a variety of serial communication interfaces including RS-422, I2C, SPI, and SpaceWire. There will also be a number of available general purpose input/output (GPIO) signals. The SCEB will house a reprogrammable FPGA that supports a soft-core processor and flight software (FSW) which will be reused from NASA GSFC’s Core Flight System (cFS) project. Lastly, the SCEB will interface with an adapter board that will contain the analog circuitry that converts temperature, voltage, and current data collected from multiple points within 6U satellite to a digital format that can be processed, stored, and downlinked using the front end Comm interface