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Lateral range and lift-drag ratio requirements for return to optimum or nonoptimum recovery sites from near-earth circular orbit
Regardless of one's favorite model for the origin of the Earth-Moon system (fission, coformation, tidal capture, giant-impact) the early history of lunar orbital evolution would produce significant thermal and earth and ocean tidal effects on the primitive earth. Three of the above lunar origin models (fission, coformation, giant-impact) feature a circular orbit which undergoes a progressive increase in orbital radius from the time of origin to the present time. In contrast, a tidal capture model places the moon in an elliptical orbit undergoing progressive circularization from the time of capture (for model purposes about 3.9 billion years ago) for at least a few 10(exp 8) years following the capture event. Once the orbit is circularized, the subsequent tidal history for a tidal capture scenario is similar to that for other models of lunar origin and features a progressive increase in orbital radius to the current state of the lunar orbit. This elliptical orbit phase, if it occurred, should have left a distinctive signature in the terrestrial and lunar rock records. Depositional events would be associated terrestrial shorelines characterized by abnormally high, but progressively decreasing, ocean tidal amplitudes and ranges associated with such an orbital evolution. Several rock units in the age range 3.6-2.5 billion years before present are reported to have a major tidal component. Examples are the Warrawoona, Fortescue, and Hamersley Groups of Western Australia and the Pangola and Witwatersand Supergroups of South Africa. Detailed study of the features of these tidal sequences may be helpful in deciphering the style of lunar orbital evolution during the Archean Eon.
Motion of near lunar satellite in initially circular orbit analyzed by integrating equations for time rates of change of orbit elements
Motion of near lunar satellite in initially circular orbit analyzed by integrating equations for time rates of change of orbit elements
Gravitational and magnetic torque effects on rotational motion of asymmetric Pegasus satellite in circular orbit
Computer program includes time minimizing function required for propellant constrained minimum time impulsive transfer between coplanar circular orbits or unrestricted orbits. Basic formulation is extended to include inclined circular or inclined elliptic orbits.
The navigation accuracy for tracking the orbiter prior to interim upper stage (IUS) deployment using the tracking data relay satellite system (TDRSS) was studied. The orbiter navigation accuracy for both one and two TDRSS satellites, for short and long data arcs, and for Doppler-only and Doppler-plus range solutions was examined. All test cases were run with the orbiter in a 150-n. mi. circular orbit, 28.5 degree inclination, at the time interval from the completion of the orbital maneuvering system (OMS)-2 maneuver to OMS-2 plus 2 hours (approximate time for IUS deployment). The data used were simulated by the simulation navigation (SIMNAV) program. The software tool used to process the TDRS data was the Shuttle Navigation Analysis Program (SNAP), a Kalman filter tool used to solve for the orbiter position and velocity. Results summarize the expected navigation accuracy using the TDRS system. It was concluded that: (1) data from both TDRS satellites were essential for accurate navigation results: (2) range data were essential for the short arc test case but were not needed for the long arc test case; and (3) with Doppler and range data from both TDRS satellites, the results converged to a reasonable solution after 5 to 10 minutes of data.
This paper is a presentation of some practical aspects of orbital transfer from Geosynchronous Transfer Orbit (GTO) to close, near-circular orbits of the Moon. The intent is to identify the important parameters affecting the problem and to bound (approximately) the range of required delta V for a spacecraft that has been placed in GTO. The basic geometric relationships are described and the dynamics are simulated by use of the Zero-Sphere-of Influence Patched Conic method. It is found that the inclination of the transfer orbit to the Earth-Moon plane is relatively unimportant while the position of the line of apsides with respect to the Moon's orbit is the main geometric parameter of interest. It is shown that this parameter can be controlled by selecting the time of day for launch and that two launch windows of approximately 45 minutes duration are available each day of the year if use is made of the recommended phasing orbit transfer. The phasing orbit transfer not only provides twice-daily launch windows, but also provides a mechanism for efficacious correction of GTO injection errors. Delta V penalties for out-of-plane transfer and for late launch are evaluated and the method is recommended for use as an affordable means of achieving lunar orbit.
NASA's Small Explorer Program (SMEX) is a sustained program of scientific satellites limited in mass to 200-300 kg (depending on orbital inclination) for a 500-km circular orbit. The SMEX program was undertaken to obtain the benefits of scientific yield, short development time, and high flight rate, with the goal to launch a mission every year. NASA also uses the program to train engineers and managers in designing and developing spacecraft. The phases of a SMEX mission life cycle are described including the competitive selection of missions through announcement of opportunity and evaluation of proposals, the definition of mission and system requirements, design and development of the spacecraft, testing, launch, and operations. Program content of some SMEX missions is then reviewed. The first SMEX mission was the Solar, Anomalous, and Magnetospheric Particle Explorer, which confirmed that anomalous cosmic rays are only partially ionized atoms. The Fast Auroral Snapshot Explorer was intended to measure the electric, magnetic, and time-variable fields and to record particle flow in the auroral acceleration region. A future mission is the Submillimeter Wave Astronomy Satellite to examine low-level molecular transitions.
Thermal drag, a variant of the Yarkovsky effect, may act on small asteroids with sizes from a few meters to a few tens of meters. Yarkovsky thermal drag comes from an asteroid's absorbing sunlight in the visible and reradiating it in the infrared. Since the infrared photons have momentum, by action-reaction, they kick the asteroid when they leave its surface. The reradiation, which is asymmetric in latitude over the asteroid, gives a net force along the asteroid's pole. Due to the asteroid's thermal inertia, averaging this force over one orbital period produces a net drag if the spin axis has a component in the orbital plane. Thermal drag tends to circularize orbits. It can increase or decrease orbital inclinations. An object whose spin axis points in random directions over its lifetime displays little change in orbital inclination. Thermal drag appears to have little to do with the delivery of chondrites from the asteroid belt; the thermal drag timescale (10(exp 8) years for meter-sizzed objects) is long compared with their cosmic ray exposure ages, and aphelia in the asteroid belt are not expected for mature thermal drag orbits. However, Yarkovsky thermal drag may act on the recently discovered near-Earth asteroids, which have radii of 10-30 m. Asteroid 1992 DA, for instance, might have its orbit shrunk by 0.1 AU in 3 x 10(exp 7) years, removing it from an Earth-crossing orbit. The near-Earth asteroids also tend to have small to moderate orbital eccentricities, as expected for highly evolved thermal drag objects. However, the time needed to bring them in from the asteroid belt (about 10(exp 9) years) is long compared with the collisonal and dynamical lifetimes (both about 10(exp 8) years) for Earth-crossing objects, arguing against their emplacement by thermal drag.
Since 1991, The MITRE Corporation has performed several independent evaluations of proposed mobile satellite service (MSS) systems that would operate from low Earth orbit (LEO) or medium Earth orbit (MEO), also known as intermediate circular orbit (ICO). This paper introduces a top level Risk Taxonomy tailored to summarize the technical and programmatic risks that MITRE has identified. In general, as risks are identified and addressed, a system's technical characteristics, cost and schedule are affected. This paper traces changes in three key parameters - satellite launch mass, system cost, and system schedule - for each of the five original non-GEO MSS systems for which license applications were made to the U.S. Federal Communications Commission (FCC) from November 1990 until June 1991. Finally, specific risk areas are identified using the Risk Taxonomy as a framework for discussion.
Satellite orbits close to an oblate planet precess about its equatorial plane, while distant satellites precess around the plane of the planet's heliocentric orbit. In between, satellites in nearly circular orbits precess about a warped intermediate surface called the Laplace 'plane.' Herein we derive general formulas for locating the Laplace plane. Because Uranus and Pluto have high obliquities, their Laplace planes are severely warped. We present maps of these Laplace planes, of interest in telescopic searches for new satellites. The Laplace plane of the Solar System as a whole is similarly distorted, but comets in the inner Oort cloud precess too slowly to sense the Laplace plane.
NASA's Near Earth Asteroid Rendezvous Mission began its record-setting exploration of the asteroid 433 Eros by inserting the spacecraft into orbit about Eros on February 14, 2000. This is the first spacecraft from any country to orbit an asteroid. The mission has overcome a failed insertion burn attempt on December 20, 1998, an event that would have ended most planetary missions, to return to the same target and successfully begin its science mapping a little more than a year later. Shortly after the successful insertion into orbit, the mission was renamed NEAR Shoemaker (NEAR) in memory of the late astronomer and geologist Eugene Shoemaker. NEAR will gather science data at Eros until February 14, 2001, which is the nominal end of mission. The NEAR mission is managed by the Johns Hopkins University, Applied Physics Laboratory in Laurel, Maryland. Since the initial mission concept in 1992, the design and implementation of the NEAR navigation system have been the responsibility of the Jet Propulsion Laboratory, California Institute of Technology. This presentation will show some of the unique features of navigation and mission design related to orbiting an asteroid and to designing a robust navigation system for the NEAR spacecraft. The problem of navigating a spacecraft about an asteroid is made difficult by the relative uncertainty in the asteroid physical properties which perturb the orbit: i.e., the mass, gravity field, and spin state. To help solve this problem, the navigation system for NEAR uses traditional DSN radio metric Doppler and range tracking, along with new technologies of optical landmark tracking and laser ranging to the asteroid surface. The experiences to date for each of these data types in the navigation solutions will be presented. Plans for the remainder of the NEAR mission will be presented, which include low orbits (down to 35 km radius circular orbits), and close flybys that may pass within 1 km of the surface. In addition, at the end of mission, NASA has approved a controlled descent and hovering phase that will culminate with the spacecraft impacting the surface. The maneuver planning for this final phase will also be presented.
The exploration of Ice Giants, Kuiper Belt Objects (KBOs) and their moons pose unique challenges from a mission design standpoint. NASA is currently developing a scalable 1-10 kilowatt electric class in-space fission reactor, known as Kilopower. The focus of this paper is to investigate the applicability of Kilopower Electric Propulsion to orbiting missions to Uranus, Neptune, and Pluto. This effort is broken into two pieces for each destination. First, a broad search of interplanetary trajectories with multiple gravity assists is completed to identify a range of mission opportunities from 2025 to 2045. Second, preliminary analysis is completed to understand the accessibility of various destination orbits, including elliptical orbits around the primary body and circular orbits around the largest moons. Results suggest that orbital missions to Uranus and Neptune are feasible with reasonable time of flight. Further work is necessary to achieve similar success with Pluto missions, but preliminary results are promising.
This paper describes our effort in searching different lunar relay architectures and assessing their coverage performance and other pros and cons for RF and optical links. Regarding RF links, we considered three notional lunar relay architectures, all communicating with the three sites of the Deep Space Network (DSN): Goldstone, California; Canberra, Australia; and Madrid, Spain: 1) A constellation of three relay orbiters: two at the Polar frozen elliptical orbits, and one at an equatorial circular orbit. 2) One relay orbiter in a 74-day Lissajous orbit at Earth-Moon Lagrange Point L2. 3) One relay orbiter in a 14-day Lunar distance retrograde orbit. Regarding optical link, we considered a notional multi-hop relay architecture that consists of one relay orbiter in a 74-day Lissajous orbit at Earth-Moon L2 (as in architecture #2 in the RF case), three Near-Earth relay satellites in geosynchronous orbit at longitudes 187° E, 110° E, and 347° E, and three ground stations at White Sands (US), Guam (US), and Tenerife (Spain). We introduce the additional constraints of Sun-“Earth”Probe (SEP) angle for return links, and Sun-Probe-“Earth” (SPE) angle for forward links. SPE and SPE angles determine the amount of sunlight (noise) that goes into the detector of the telescope, thus affecting the capacity of an optical link.
In this paper, we examine the effects of radiation pressure, Poynting-Robertson (PR) drag, and solar wind drag on dust grains trapped in mean motion resonances with the Sun and Jupiter in the restricted (negligible dust mass) three-body Problem. We especially examine the evolution of dust grains in the 1:1 resonance. As a first step, the Sun and Jupiter are idealized to both be in circular orbit about a common center of mass (circular restricted three-body problem). From the equation of motion of the dust particle in the rotating reference frame, the drag-induced time rate of change of its Jacobi "constant," C, is then derived and expressed in spherical coordinates. This new mathematical expression in spherical coordinates shows that C, in the 1:1 resonance, both oscillates and secularly increases with increasing time. The new expression gives rise to an easy understanding of how an orbit evolves when the radiation force and solar wind drag are included. All dust grain orbits are unstable in time when PR and solar wind drag are included in the Sun-Jupiter-dust system. Tadpole orbits evolve into horseshoe orbits; and these orbits continuously expand in size to lead to close encounters with Jupiter. Permanent trapping is impossible. Orbital evolutions of a dust grain trapped in the 1:1 resonance in the planar circular, an inclined case, an eccentric case, and the actual Sun-Jupiter case are numerically simulated and compared with each other and show grossly similar time behavior. Resonances other than 1:1 are also explored with the new expression. Stable exterior resonance trapping may be possible under certain conditions. One necessary condition for such a trap is derived. Trapping in interior resonances is shown to be always unstable.
Regularization of restricted three-body problem extended to case where three primaries of any mass revolve in circular orbits around common center of mass and fourth body of infinitesimal mass moves in their field
Analytical approaches to path-adaptive guidance functions, circular orbit trajectories, and use of Fortran-compiled programs