SEARCH · Search NASA
Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Reduced computer precision effect on midcourse navigation and guidance system using optimal filtering and linear prediction
Simplified guidance procedures have been developed which are based on preflight determination of the characteristics of perturbed trajectories. The results are devoted primarily to planetary approach guidance; however, some considerations for midcourse guidance are included. The methods are studied for an Earth-to-Mars trajectory but would be applicable to Grand Tour and other types of missions. Generally requiring only a single onboard optical angular measurement, the approach procedure predicts guidance corrections for the control of periapsis radius as well as orbital plane orientation. An error analysis has shown that the periapsis radius at Mars can be controlled to a 1 sigma accuracy of about 20 km if the guidance is performed 1/2 day before periapsis passage. If the guidance is performed at the Martian sphere of influence (2.2 days before periapsis), the error doubles because of the increased effect of measurement and maneuvering errors.
Prediction of velocity requirements for minimum time response abort trajectories for midcourse guidance of lunar flight
Computation of transition matrices or error coefficients in midcourse guidance studies of lunar trajectories
Sextant simulator for midcourse guidance of manned spacecraft during interstellar travel
Midcourse guidance analysis and optimization with discrete measurements for mission logic development
Midcourse guidance, lunar parking and descent orbits, lunar landing phase, lunar ascent, and lunar and earth rendezvous
Navigator sighting accuracy in midcourse guidance system for manned space flight, using simulated vehicle mounted space sextant
A midcourse guidance and navigation system for continuous low thrust vehicles is described. A set of orbit elements, known as the equinoctial elements, are selected as the state variables. The uncertainties are modelled statistically by random vector and stochastic processes. The motion of the vehicle and the measurements are described by nonlinear stochastic differential and difference equations respectively. A minimum time nominal trajectory is defined and the equation of motion and the measurement equation are linearized about this nominal trajectory. An exponential cost criterion is constructed and a linear feedback guidance law is derived to control the thrusting direction of the engine. Using this guidance law, the vehicle will fly in a trajectory neighboring the nominal trajectory. The extended Kalman filter is used for state estimation. Finally a short mission using this system is simulated. The results indicate that this system is very efficient for short missions.
Manual abort guidance from midcourse region of lunar mission
A midcourse guidance and navigation system for continuous low thrust vehicles was developed. The equinoctial elements are the state variables. Uncertainties are modelled statistically by random vector and stochastic processes. The motion of the vehicle and the measurements are described by nonlinear stochastic differential and difference equations respectively. A minimum time trajectory is defined; equations of motion and measurements are linearized about this trajectory. An exponential cost criterion is constructed and a linear feedback quidance law is derived. An extended Kalman filter is used for state estimation. A short mission using this system is simulated. It is indicated that this system is efficient for short missions, but longer missions require accurate trajectory and ground based measurements.
Mathematical techniques used in minimum variance orbital navigation and guidance, midcourse phase, lunar parking, descent orbits, landing, ascent, and rendezvous
Manual procedures for midcourse and terminal guidance, discussing onboard optical measurements and calculations
Development of procedures for midcourse and approach guidance of moon to earth trajectories based on optical angular measurements and onboard calculations
Concepts from statistical filter theory are applied to the problem of in-flight determination of the position and velocity of a space vehicle for the purposes of midcourse guidance. The source of information is assumed to be an arbitrary sequence of measurements of any desired set of "observables" (e.g., space angles), the measurements being corrupted by additive errors so that the position and velocity are never known perfectly. A state transition approach is employed which leads naturally to a computational scheme that is readily implemented by a digital computer. The scheme can be regarded as a dynamical time-varying filter which weights the incoming observations in an optimal sense for use in producing an up-to-date optimal estimate of position and velocity. The advantages of the scheme are (1) it provides the best possible estimate (minimum error) based upon ensemble statistics of injection conditions and measurement errors; (2) it is extremely versatile, not requiring adherence to a predetermined observation schedule or reference trajectory; and (3) the computations are sufficiently simple to be practical in an on-board computer. A digital computer simulation of the proposed system is employed to demonstrate the feasibility of an all on-board system and to illustrate the performance attainable in a hypothetical physical situation.
Optimum control theory formulations for solving problems in optimum guidance for interplanetary manned space flight missions
Linear stochastic systems optimal feedback control with application to spacecraft minimum fuel midcourse guidance problems
Hybrid computer simulation for Apollo midcourse guidance from elliptical preabort orbit to another ellipse, using coplanar/noncoplanar orbit change guidance scheme