The Development of a Low Cost, Modular Attitude Determination and Control System

In an attempt to reduce the cost of future satellites, new technologies are being pursued to develop a modular attitude determination and control system that will provide three-axis control and cost less than ten percent of present systems. The low cost and modularity of this system make it especially attractive to a wide variety of small satellites. This paper will present the design and developmental status of this plug and play attitude control system. The general idea is to provide a complete attitude determination and control system, including sensors, actuators, and software, that can be used for a wide variety of satellites. A single, pressurized box will contain all the electronics for the system, including the command computer, power supply, reaction wheel motor drivers, and star tracker electronics. Pressurization of the box (1atm BOL, 12,000 ft. EOL) enables the use of terrestrial components, which greatly reduces cost. The box will provide its own thermal control through the use of heaters, fans, and radiators. The power for the attitude control system will be provided by the spacecraft bus, suitably conditioned by the system\u27s onboard power supply. There will be one other interface to the spacecraft to allow for downloading of telemetry data and uploading of ground commands. This independence of the attitude control system from the satellite allows for ease of integration. The system will be controlled by a PowerPC 603e processor running flight software that is already commercially available. This use of already developed software also helps to reduce cost. The processor will communicate with the star trackers and reaction wheels via a commercial serial bus. Attitude determination will be provided by custom designed CCD or CID star trackers. Unlike most other sensors, star trackers can be used for any type of mission, from Earth-orbiting to interplanetary. Gyroscopes will not be used in the first generation of this system because they are not necessary for low bandwidth control. As already mentioned, the three-axis attitude control will be provided by reaction wheels, which are powered by polyphase AX induction motors to reduce mass. In addition to presenting the detailed design of the attitude control system, the paper will also address the issues that must be overcome, including radiation tolerance, particularly susceptibility of single event upsets and latch-up. Finally, the paper will discuss the results of the in-house testing of components and the production schedule

Spacecraft attitude and velocity control system

A spacecraft attitude and/or velocity control system includes a controller which responds to at least attitude errors to produce command signals representing a force vector F and a torque vector T, each having three orthogonal components, which represent the forces and torques which are to be generated by the thrusters. The thrusters may include magnetic torquer or reaction wheels. Six difference equations are generated, three having the form ##EQU1## where a.sub.j is the maximum torque which the j.sup.th thruster can produce, b.sub.j is the maximum force which the j.sup.th thruster can produce, and .alpha..sub.j is a variable representing the throttling factor of the j.sup.th thruster, which may range from zero to unity. The six equations are summed to produce a single scalar equation relating variables .alpha..sub.j to a performance index Z: ##EQU2## Those values of .alpha. which maximize the value of Z are determined by a method for solving linear equations, such as a linear programming method. The Simplex method may be used. The values of .alpha..sub.j are applied to control the corresponding thrusters

Control Analysis of flexible Solar Sails

Future solar sail missions will require sails with dimensions on the order of 100 m to l km. At these sizes, given the gossamer nature of the sail supporting structures, flexible modes may be low enough to interact with the control system. This paper develops a practical analysis of the flexible interactions using state-space systems and modal data from standard finite element models of the sail sub- system. The modal data is combined with a rigid core bus to create a modal coordinate state-space plant, which can be analyzed for stability with a state-space controller. Results are presented for an 80 m sail for both collocated actuation and control by actuators mounted at the sail tips

Propellantless AOCS Design for a 160-m, 450-kg Sailcraft of the Solar Polar Imager Mission

An attitude and orbit control system (AOCS) is developed for a 160-m, 450-kg solar sail spacecraft of the Solar Polar Imager (SPI) mission. The SPI mission is one of several Sun- Earth Connections solar sail roadmap missions currently envisioned by NASA. A reference SPI sailcraft consists of a 160-m, 150-kg square solar sail, a 250-kg spacecraft bus, and 50-kg science payloads, The 160-m reference sailcraft has a nominal solar thrust force of 160 mN (at 1 AU), an uncertain center-of-mass/center-of-pressure offset of +/- 0.4 m, and a characteristic acceleration of 0.35 mm/sq s. The solar sail is to be deployed after being placed into an earth escaping orbit by a conventional launch vehicle such as a Delta 11. The SPI sailcraft first spirals inwards from 1 AU to a heliocentric circular orbit at 0.48 AU, followed by a cranking orbit phase to achieve a science mission orbit at a 75-deg inclination, over a total sailing time of 6.6 yr. The solar sail will be jettisoned after achieving the science mission orbit. This paper focuses on the solar sailing phase of the SPI mission, with emphasis on the design of a reference AOCS consisting of a propellantless primary ACS and a microthruster-based secondary (optional) ACS. The primary ACS employs trim control masses running along mast lanyards for pitch/yaw control together with roll stabilizer bars at the mast tips for quadrant tilt (roll) control. The robustness and effectiveness of such a propellantless primary ACS would be enhanced by the secondary ACS which employs tip-mounted, lightweight pulsed plasma thrusters (PPTs). The microPPT-based ACS is mainly intended for attitude recovery maneuvers from off-nominal conditions. A relatively fast, 70-deg pitch reorientation within 3 hrs every half orbit during the orbit cranking phase is shown to be feasible, with the primary ACS, for possible solar observations even during the 5-yr cranking orbit phase

Spacecraft design project: High latitude communications satellite

The spacecraft design project was part of AE-4871, Advanced Spacecraft Design. The project was intended to provide experience in the design of all major components of a satellite. Each member of the class was given primary responsibility for a subsystem or design support function. Support was requested from the Naval Research Laboratory to augment the Naval Postgraduate School faculty. Analysis and design of each subsystem was done to the extent possible within the constraints of an eleven week quarter and the design facilities (hardware and software) available. The project team chose to evaluate the design of a high latitude communications satellite as representative of the design issues and tradeoffs necessary for a wide range of satellites. The High-Latitude Communications Satellite (HILACS) will provide a continuous UHF communications link between stations located north of the region covered by geosynchronous communications satellites, i.e., the area above approximately 60 N latitude. HILACS will also provide a communications link to stations below 60 N via a relay Net Control Station (NCS), which is located with access to both the HILACS and geosynchronous communications satellites. The communications payload will operate only for that portion of the orbit necessary to provide specified coverage