Oscar Hsu

Space Technology 7 disturbance reduction system - precision control flight validation

The NASA New Millennium Program Space Technology 7 (ST7) project validates technology for precision spacecraft control. The disturbance reduction system (DRS) is part of the European Space Agencys LISA Pathfinder project. The DRS controls the position of the spacecraft relative to a reference to an accuracy of one nanometer over time scales of several thousand seconds. To perform the control, the spacecraft use a new colloid thruster technology. The thrusters operates over the range of 5 to 30 micro-Newtons with precision of 0.1 micro-Newton. The thrust is generated by using a high electric field to extract charged droplets of a conducting colloid fluid and accelerating them with a precisely adjustable voltage. The control reference is provided by the European LISA Technology Package, which includes two nearly free-floating test masses. The test mass positions and orientations are measured using a capacitance bridge. The test mass position and attitude is adjustable using electrostatically applied forces and torques. The DRS controls the spacecraft position with respect to one test mass while minimizing disturbances on the second test mass. The dynamic control system covers eighteen degrees of freedom: six for each of the test masses and six for the spacecraft. After launch in late 2009 to a low Earth orbit, the LISA Pathfinder spacecraft is maneuvered to a halo orbit about the Earth-Sun LI Lagrange point for operations

Control Modes of the ST7 Disturbance Reduction System Flight Validation Experiment

The Space Technology 7 (ST7) experiment will perform an on-orbit system-level validation of two specific Disturbance Reduction System technologies: a gravitational reference sensor employing a free-floating test mass and a set of micronewton colloidal thrusters. The ST7 Disturbance Reduction System (DRS) is designed to maintain the spacecrafts position with respect to a free-floating test mass to less than 10 nm/√Hz over the frequency range of 1 to 30 mHz. This paper presents the overall design and analysis of the spacecraft drag-free and attitude controllers. These controllers close the loop between the gravitational sensors and the micronewton colloidal thrusters. There are five control modes in the operation of the ST7 DRS, starting with the attitude-only mode and leading to the science mode. The design and analysis of each of the control modes as well as the mode transition strategy are presented.

Solar Dynamics Observatory Guidance, Navigation, and Control System Overview

Angela M. Russo, Scott R. Starin, and Melissa F. VessNASA Goddard Space Flight Center Code 591, Greenbelt, Maryland 20771AbstractThe Solar Dynamics Observatory (SDO) was designed and built at the Goddard Space Flight Center, launchedfrom Cape Canaveral on February 11, 2010, and reached its final geosynchronous science orbit on March 16, 2010.The purpose of SDO is to observe the Sun and continuously relay data to a dedicated ground station. SDO remainsSun-pointing throughout most of its mission for the instruments to take measurements of the Sun. The SDO attitudecontrol system (ACS) is a single-fault tolerant design. Its fully redundant attitude sensor complement includessixteen coarse Sun sensors (CSSs), a digital Sun sensor (DSS), three two-axis inertial reference units (IRUs), andtwo star trackers (STs). The ACS also makes use of the four guide telescopes included as a part of one of the scienceinstruments. Attitude actuation is performed using four reaction wheels assemblies (RWAs) and eight thrusters, witha single main engine used to provide velocity-change thrust for orbit raising. The attitude control software has fivenominal control modes, three wheel-based modes and two thruster-based modes. A wheel-based Safehold running inthe attitude control electronics box improves the robustness of the system as a whole. All six modes are designed onthe same basic proportional-integral-derivative attitude error structure, with more robust modes setting their integralgains to zero. This paper details the final overall design of the SDO guidance, navigation, and control (GNCAtmospheric Imaging Assembly (AIA), led by Lockheed Martin Space and Astrophysics Laboratory; and ExtremeUltraviolet Variability Experiment (EVE), led by the University of Colorado. The basic mission is to observe theSun for a very high percentage of the 5-year mission (10-year goal) with long stretches of uninterrupted observationsand with constant, high-data-rate transmission to a dedicated ground station to be located in White Sands, NewMexico. These goals guided the design of the spacecraft bus that will carry and service the three-instrument payload.Overarching design goals for the bus are geosynchronous orbit, near-constant Sun observations with the ability to flythrough eclipses, and constant HGA contact with the dedicated ground station. A three-axis stabilized ACS isneeded both to point at the Sun accurately and to keep the roll about the Sun vector correctly positioned with respectto the solar north pole. This roll control is especially important for the magnetic field imaging of HM I.The mission requirements have several general impacts on the ACS design. Both the AIA and HMI instrumentsare very sensitive to the blurring caused by jitter. Each has an image stabilization system (ISS) with some ability tofilter out high frequency motion, but below the bandwidth of the ISS the control system must compensate fordisturbances within the ACS bandwidth or avoid exciting jitter at higher frequencies.Within the ACS bandwidth, the control requirement imposed by AIA is to place the center of the solar disk nomore than 2 arc

Mode Transitions for the ST7 Disturbance Reduction System Experiment

The Space Technology 7 Disturbance Reduction System experiment will perform an on-orbit system-level validation of two technologies: a gravitational reference sensor employing a free-floating test mass and a set of colloidal micronewton thrusters. The Disturbance Reduction System is designed to maintain the spacecraft s position with respect to a free floating test mass to less than 10 nm/& over the frequency range of 1 to 30 m= mi paper presents the modes that compose the Disturbance Reduction System spacecraft control as well as the strategy used to transition between modes. A high-fidelity model of the system, which incorporates rigid-body models of the spacecraft and two test masses (18 degrees of freedom), is developed and used to evaluate the performance of each mode and the efficacy of the transition strategy.

Control of the ST7 Disturbance Reduction System Flight Experiment

The Space Technology 7 (ST7) experiment will perform an on‐orbit system‐level validation of two specific Disturbance Reduction System technologies: colloidal micronewton thrusters and drag‐free control. The ST7 Disturbance Reduction System (DRS) is designed to maintain the spacecraft’s position with respect to a free‐floating test mass while limiting the residual accelerations of that test mass over the frequency range of 1 to 30 mHz. This paper presents the overall design and analysis of the spacecraft drag‐free and attitude controllers, with particular attention given to its primary mission mode. These controllers close the loop between the drag‐free sensors and the colloidal micronewton thrusters.

Space Technology 7 -- Micropropulsion and Mass Distribution

The NASA New Millennium Program Space Technology 7 (ST7) project will validate technology for precision spacecraft control. The ST7 distrubance reduction system (DRS) will contain new micropulsion technology to be flown as part of the European Space Agencys LISA (laser interferometer space antenna) Pathfinder project. After launch into a low Earth orbit in early 2010, the LISA Pathfinder spacecraft will be maneuvered to a halo orbit about the Earth-Sun LI Lagrange point for operations. The DRS will control the position of the spacecraft relative to a reference to an accuracy of one nanometer over time scales of several thousand seconds. To perform the control the spacecraft will use a new colloid thruster technology. The thrusters will operate over the range of 5 to 30 micro-Newtons with precision of 0.1 micro-Newton. The thrust will be generated by using a high electric field to extract charged droplets of a conducting colloid fluid and accelerating them with a precisely adjustable voltage. The control position reference will be provided by the European LISA Technology Package, which will include two nearly free-floating test masses. The test mass position and attitude will be sensed and adjusted using electrostatic capacitance bridges. The DRS will control the spacecraft position with respect to one test mass while minimizing disturbances on the second test mass. The dynamic control system will cover eighteen degrees of freedom, six for each of the test masses and six for the spacecraft. In the absence of other disturbances, the test masses will slowly gravitate toward local concentrations of spacecraft mass. The test mass acceleration must be minimized to maintain the acceleration of the enclosing drag-free spacecraft within the control authority of the micropropulsion system. Therefore, test mass acceleration must be predicted by accurate measurement of mass distribution, then offset by the placement of specially shaped balance masses near each test mass. The acceleration is characterized by calculating the gravitational effect of over ten million modeled points of a nearly 500-kg spacecraft. This paper provides an overview of the mission technology and the process of precision mass modeling of the DRS equipment.

Integration and Testing of the Lunar Reconnaissance Orbiter Attitude Control System

Prior to the successful launch of the Lunar Reconnaissance Orbiter (LRO) on June 18, 2009, the Attitude Control System (ACS) team completed numerous Integration and Testing (I&T) tests on each hardware component in ever more flight like environments. 12The ACS utilizes a select group of attitude sensors and actuators. This paper will chronicle the evolutionary steps taken to verify each component was constantly ready for flight as well as providing invaluable trending experience with the actual hardware. The paper will include a discussion of each ACS hardware component, lessons learned of the various stages of I&T, a discussion of the challenges that are unique to the LRO project, as well as a discussion of work for future missions to consider as part of their I&T plan