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Kia, T.

Publications and source records attributed to Kia, T..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2002

Ultra long-life avionics architecture

For survival and achieving reliability in ultra long-life missions, fault tolerant design techniques need to handle the predominant failure mode, which is the wear-out of components. Conventional design methodologies will need excessive redundancy to achieve the required reliability. The objective of this paper is to present a new approach to design a more efficient fault-tolerant avionics system architecture that requires significantly fewer redundant components.

Long-life missions avionics↗
NASA NTRS2001

Micro Navigator

Miniature high-performance low-mass space avionics systems are desired for planned future outer planetary exploration missions (i.e. Europa Orbiter/Lander, Pluto-Kuiper Express). The spacecraft fuel and mass requirements enabling orbit insertion is the driving requirement. The Micro Navigator is an integrated autonomous Guidance, Navigation & Control (GN&C)micro-system that would provide the critical avionics function for navigation, pointing, and precision landing. The Micro Navigator hardware and software allow fusion of data from multiple sensors to provide a single integrated vehicle state vector necessary for six degrees of freedom GN&C. The benefits of this MicroNavigator include: 1) The Micro Navigator employs MEMS devices that promise orders of magnitude reductions in mass power and volume of inertial sensors (accelerometers and gyroscopes), celestial sensing devices (startracker, sun sensor), and computing element; 2) The highly integrated nature of the unit will reduce the cost of flight missions. a) The advanced miniaturization technologies employed by the Micro Navigator lend themselves to mass production, and therefore will reduce production cost of spacecraft. b) The integral approach simplifies interface issues associated with discrete components and reduces cost associated with integration and test of multiple components; and 3) The integration of sensors and processing elements into a single unit will allow the Micro Navigator to encapsulate attitude information and determination functions into a single object. This is particularly beneficial for object-oriented software architectures that are used in advanced spacecraft. Additional information is contained in the original extended abstract.

Blaes, B. R.↗
NASA NTRS1994

Highly-Autonomous Event-Driven Spacecraft Control

Future JPL missions will continue to be scientifically and technically more ambitious, and will demand more autonomy to accomplish complex tasks in uncertain environments and in close proximity to extraterrestrial surfaces. A prime example is small body rendezvous and sample return.

attitude and control subsystem (ACS) rendezvous ta↗
NASA NTRS1989

Annihilation of angular momentum bias during thrusting and spinning-up maneuvers

During spinning-up and thrusting maneuvers of rockets and spacecraft, undesired transverse torques (from error sources such as thruster misalignment, center-of-mass offset and thruster mismatch) perturb the angular momentum vector from its original orientation. In this paper a maneuver scheme is presented which virtually annihilates the angular momentum vector bias, even though the magnitude and direction of the perturbing body-fixed torques are unknown. In the analysis it is assumed that the torques are small and constant and that the spacecraft or rocket can be approximated by a rigid body, which may be asymmetric. Typical maneuvers of the Galileo spacecraft are simulated to demonstrate the technique.

Longuski, J. M.↗
NASA NTRS1989

On-board attitude determination for the Topex satellite

This paper presents an overall technical description of the on-board attitude determination system for The Ocean Topography Experiment (Topex) satellite. The stellar-inertial attitude determination system being designed for the Topex satellite utilizes data from a three-axis NASA Standard DRIRU-II as well as data from an Advanced Star Tracer (ASTRA) and a Digital Fine Sun Sensor (DFSS). This system is a modified version of the baseline Multimission Modular Spacecraft (MMS) concept used on the Landsat missions. Extensive simulation and analysis of the MMS attitude determination approach was performed to verify suitability for the Topex application. The modifications to this baseline attitude determination scheme were identified to satisfy the unique Topex mission requirements.

Dennehy, C. J.↗
NASA NTRS1988

Attitude determination and control subsystem for the TOPEX satellite

The Ocean Topography Experiment satellite will carry a modular Attitude Determination and Control Subsystem (ADCS) which contains all equipment required for attitude determination, stabilization, and control, as well as hydrazine thruster firing control, during all mission phases. Attention is presently given to the ADCS's architecture, constituent hardware components, performance requirements, and predicted on-orbit performance compliance, with emphasis on the design and analysis of the Normal Mission Mode control algorithm furnishing the primary scientific data-acquisition operational mode. This mode's attitude determination and control of on-orbit performance is predicted to better than 43 arcsec.

Dennehy, C. J.↗
NASA NTRS1988

Nuclear reactor power as applied to a space-based radar mission

The SP-100 Project was established to develop and demonstrate feasibility of a space reactor power system (SRPS) at power levels of 10's of kilowatts to a megawatt. To help determine systems requirements for the SRPS, a mission and spacecraft were examined which utilize this power system for a space-based radar to observe moving objects. Aspects of the mission and spacecraft bearing on the power system were the primary objectives of this study; performance of the radar itself was not within the scope. The study was carried out by the Systems Design Audit Team of the SP-100 Project.

Jaffe, L.↗
NASA NTRS1987

Nuclear reactor power for an electrically powered orbital transfer vehicle

To help determine the systems requirements for a 300-kWe space nuclear reactor power system, a mission and spacecraft have been examined which utilize electric propulsion and this nuclear reactor power for multiple transfers of cargo between low earth orbit (LEO) and geosynchronous earth orbit (GEO). A propulsion system employing ion thrusters and xenon propellant was selected. Propellant and thrusters are replaced after each sortie to GEO. The mass of the Orbital Transfer Vehicle (OTV), empty and dry, is 11,000 kg; nominal propellant load is 5000 kg. The OTV operates between a circular orbit at 925 km altitude, 28.5 deg inclination, and GEO. Cargo is brought to the OTV by Shuttle and an Orbital Maneuvering Vehicle (OMV); the OTV then takes it to GEO. The OTV can also bring cargo back from GEO, for transfer by OMV to the Shuttle. OTV propellant is resupplied and the ion thrusters are replaced by the OMV before each trip to GEO. At the end of mission life, the OTV's electric propulsion is used to place it in a heliocentric orbit so that the reactor will not return to earth. The nominal cargo capability to GEO is 6000 kg with a transit time of 120 days; 1350 kg can be transferred in 90 days, and 14,300 kg in 240 days. These capabilities can be considerably increased by using separate Shuttle launches to bring up propellant and cargo, or by changing to mercury propellant.

Jaffe, L.↗
NASA NTRS1987

Earth observing system instrument pointing control modeling for polar orbiting platforms

An approach to instrument pointing control performance assessment for large multi-instrument platforms is described. First, instrument pointing requirements and reference platform control systems for the Eos Polar Platforms are reviewed. Performance modeling tools including NASTRAN models of two large platforms, a modal selection procedure utilizing a balanced realization method, and reduced order platform models with core and instrument pointing control loops added are then described. Time history simulations of instrument pointing and stability performance in response to commanded slewing of adjacent instruments demonstrates the limits of tolerable slew activity. Simplified models of rigid body responses are also developed for comparison. Instrument pointing control methods required in addition to the core platform control system to meet instrument pointing requirements are considered.

Briggs, H. C.↗
NASA NTRS1984

Error analysis of analytic solutions for self-excited near-symmetric rigid bodies - A numerical study

Analytic error bounds are presented for the solutions of approximate models for self-excited near-symmetric rigid bodies. The error bounds are developed for analytic solutions to Euler's equations of motion. The results are applied to obtain a simplified analytic solution for Eulerian rates and angles. The results of a sample application of the range and error bound expressions for the case of the Galileo spacecraft experiencing transverse torques demonstrate the use of the bounds in analyses of rigid body spin change maneuvers.

Kia, T.↗