SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS

A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 43, ISSUE 19 - SEPTEMBER 23, 2006

13 ASTRODYNAMICS
Includes powered and free flight trajectories; orbital and launching dynamics.

20050212364 Massachusetts Inst. of Tech., Cambridge, MA USA

A New Guidance Method for a Delta V and Re-entry Constrained Orbit Transfer Problem

Van Beusekom, Craig J.; Jun. 2005; 157 pp.; In English; Original contains color illustrations Report No.(s): AD-A436655; No Copyright; Avail: CASI; A08, Hardcopy

This thesis proposes a spacecraft guidance system designed for a unique class of orbit transfer problems. It considers a vehicle that undertakes a maneuver with the objective of precisely flying through a point in space at a particular time. The spacecraft must automatically determine a transfer orbit that will take it from a circular, low-earth parking orbit to a velocity-unconstrained rendezvous with a Keplerian trajectory. A constraint exists that both the final transfer orbit and the ultimate paths of any additional stages must lead rapidly to atmospheric reentry, typically within one revolution. Constrained to a fixed delta V resulting from a two stage thrust profile, the spacecraft must execute a burn maneuver that can effectively dissipate energy to place it on a transfer orbit with previously unknown velocity requirements. Finally, the guidance strategy should be robust to the uncertainties typically encountered in real spacecraft orbit transfer problems. In order to meet these constraints, this thesis first develops new analytic analysis of the relationship between reentry, perigee, and delta V. Next, a framework is developed for selecting a favorable transfer orbit while considering the various hard and soft constraints in the problem. Following transfer orbit selection, a plane of maneuver is calculated that maximizes likelihood of first stage reentry. Then traditional guidance strategies are adapted to the problem and hypothetical spacecraft design to produce a closed loop guidance solution. Results are presented that demonstrate the effectiveness of the new method. DTIC

Atmospheric Entry; Guidance (Motion); Orbits; Reentry; Spacecraft Maneuvers; Trajectories

20050212415 Naval Research Lab., USA

Perturbed Equations of Motion for Formation Flight Near the Sun-Earth L2 Point

Segerman, Alan M.; Zedd, Michael F.; [2005]; 159 pp.; In English; 2005 AAS/AIAA Astrodynamics Specialist Conference, 7-11 Aug. 2005, Lake Tahoe, CA, USA Contract(s)/Grant(s): NNG04CB01I; No Copyright; Avail: CASI; A08, Hardcopy

This Memorandum Report consists of a compilation of three individual reports, of increasing complexity, describing investigations of formation flight of spacecraft in the vicinity of the L2 Sun-Earth 1ibration point. The individual reports form the following parts of this compilation: - Introduction to the relative motion of spacecraft about the Sun-Earth L2 Point -Linear and quadratic modelling and solution of the relative motion - Modelling the Perturbations - Elliptical Earth Orbit, Lunar Gravity, Solar Radiation Pressure, Thrusters. The three parts are self-contained, with somewhat, varying notation and terminology. After fair1y significant literature searches: this new work (of Parts 2 and 3) is deemed to be unique because it describes the primary perturbations to the description of relative motion between nearby spacecraft. The effect of the elliptical motion of the Earth about the Sun was verified to be the dominant perturbation to the circular restricted three body problem. Contributions due to lunar gravity and solar radiation pressure are seen to have much smaller effect. Derived from text

Earth Orbits; Equations of Motion; Perturbation; Spacecraft Motion; Three Body Problem; Formation Flying; Satellite Constellations; Lagrangian Equilibrium Points; Orbital Mechanics

Source: NASA.