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Results for “circular orbits”

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.

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At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

Evolution of wide binary millisecond pulsars in globular clusters

The evolution of a binary radio pulsar in a wide orbit about a low-mass degenerate dwarf companion in a globular cluster environment is investigated. An extensive series of numerical scattering experiments is used to study encounters between such a pulsar and passing cluster field stars. A wide range of initial orbital periods from 3 to 1000 days is considered with a corresponding range of masses for the degenerate dwarf companions. The calculative results are summarized in the form of scattering cross sections. In particular, cross sections are presented for three final 'channels' in which the end product of the encounter is an isolated neutron star. Cross sections for inducing various orbital eccentricities into initially circular orbits are also derived. The implications of the results for the formation of the single millisecond pulsars in M28 and M15 and for the nearly circular obit of the M4 binary radio pulsar are discussed.

Rappaport, S.↗
NASA NTRS1974

Structure of Saturn's rings: Optical and dynamical considerations

The photometric phase curves of Saturn's rings are considered, as well as a conflict between dynamical and photometric models of the rings. The dependence of ring brightness on angular separation of the earth and sun as viewed from Saturn is discussed. The nonlinear brightness surge is interpreted. Some quantitative calculations were carried out for bodies in and near the asteroidal belt. Predicted density profiles of the ring obtained with Mimas in an eccentric orbit and in a circular orbit are also included.

Franklin, F. A.↗
NASA NTRS1978

New formulation of de Sitter's theory of motion for Jupiter I-IV. I - Equations of motion and the disturbing function

Elliptic orbits are substituted for circular orbits in the first approximation, in an analysis of the common retrograde motion of Jupiter's satellites. A modification of the de Sitter theory, made possible by extended observations of the satellites, is presented with attention to that aspect of the theory which eliminates small divisors at all stages of the solution. The convergence problem is circumvented by use of Poincare's canonical relative coordinates. In addition, modified Delaunay variables and their associated Poincare variables are applied to the disturbing function, which is expanded by means of generalized Newcomb operators.

Aksnes, K.↗
NASA NTRS2007

Optical Observations of GEO Debris with Two Telescopes

For several years, the Michigan Orbital DEbris Survey Telescope (MODEST), the University of Michigan s 0.6/0.9-m Schmidt telescope on Cerro Tololo Inter-American Observatory in Chile has been used to survey the debris population at GEO in the visible regime. Magnitudes, positions, and angular rates are determined for GEO objects as they move across the telescope s field-of-view (FOV) during a 5-minute window. This short window of time is not long enough to determine a full six parameter orbit so usually a circular orbit is assumed. A longer arc of time is necessary to determine eccentricity and to look for changes in the orbit with time. MODEST can follow objects in real-time, but only at the price of stopping survey operations. A second telescope would allow for longer arcs of orbit to obtain the full six orbital parameters, as well as assess the changes over time. An additional benefit of having a second telescope is the capability of obtaining BVRI colors of the faint targets, aiding efforts to determine the material type of faint debris. For 14 nights in March 2007, two telescopes were used simultaneously to observe the GEO debris field. MODEST was used exclusively in survey mode. As objects were detected, they were handed off in near real-time to the Cerro Tololo 0.9-m telescope for follow-up observations. The goal was to determine orbits and colors for all objects fainter than R = 15th magnitude (corresponds to 1 meter in size assuming a 0.2 albedo) detected by MODEST. The hand-off process was completely functional during the final eight nights and follow-ups for objects from night-to-night were possible. The cutoff magnitude level of 15th was selected on the basis of an abrupt change in the observed angular rate distribution in the MODEST surveys. Objects brighter than 15th magnitude tend to lie on a well defined locus in the angular rate plane (and have orbits in the catalog), while fainter objects fill the plane almost uniformly. We need to determine full six-parameter orbits to investigate what causes this change in observed angular rates. Are these faint objects either the same population of high area-to-mass (A/M) objects on eccentric orbits as discovered by the ESA Space Debris Telescope (Schildknecht, et al. 2004), or are they just normal debris from breakups in GEO?

Seitzer, P.↗
NASA NTRS2010

GOATS - Orbitology Component

The GOATS Orbitology Component software was developed to specifically address the concerns presented by orbit analysis tools that are often written as stand-alone applications. These applications do not easily interface with standard JPL first-principles analysis tools, and have a steep learning curve due to their complicated nature. This toolset is written as a series of MATLAB functions, allowing seamless integration into existing JPL optical systems engineering modeling and analysis modules. The functions are completely open, and allow for advanced users to delve into and modify the underlying physics being modeled. Additionally, this software module fills an analysis gap, allowing for quick, high-level mission analysis trades without the need for detailed and complicated orbit analysis using commercial stand-alone tools. This software consists of a series of MATLAB functions to provide for geometric orbit-related analysis. This includes propagation of orbits to varying levels of generalization. In the simplest case, geosynchronous orbits can be modeled by specifying a subset of three orbit elements. The next case is a circular orbit, which can be specified by a subset of four orbit elements. The most general case is an arbitrary elliptical orbit specified by all six orbit elements. These orbits are all solved geometrically, under the basic problem of an object in circular (or elliptical) orbit around a rotating spheroid. The orbit functions output time series ground tracks, which serve as the basis for more detailed orbit analysis. This software module also includes functions to track the positions of the Sun, Moon, and arbitrary celestial bodies specified by right ascension and declination. Also included are functions to calculate line-of-sight geometries to ground-based targets, angular rotations and decompositions, and other line-of-site calculations. The toolset allows for the rapid execution of orbit trade studies at the level of detail required for the early stage of mission concept development.

Haber, Benjamin M.↗
NASA NTRS1994

Design of an unmanned Martian polar exploration system

The design of an unmanned Martian polar exploration system is presented. The system elements include subsystems for transportation of material from earth to Mars, study of the Martian north pole, power generation, and communications. Early next century, three Atlas 2AS launch vehicles will be used to insert three Earth-Mars transfer vehicles, or buses, into a low-energy transfer orbit. Capture at Mars will be accomplished by aerobraking into a circular orbit. Each bus contains four landers and a communications satellite. Six of the twelve total landers will be deployed at 60 deg intervals along 80 deg N, and the remaining six landers at 5 deg intervals along 30 deg E from 65 deg N to 90 deg N by a combination of retrorockets and parachutes. The three communications satellites will be deployed at altitudes of 500 km in circular polar orbits that are 120 deg out of phase. These placements maximize the polar coverage of the science and communications subsystems. Each lander contains scientific equipment, two microrovers, power supplies, communications equipment, and a science computer. The lander scientific equipment includes a microweather station, seismometer, thermal probe, x-ray spectrometer, camera, and sounding rockets. One rover, designed for short-range (less than 2 km) excursions from the lander, includes a mass spectrometer for mineral analysis, an auger/borescope system for depth profiling, a deployable thermal probe, and charge coupled device cameras for terrain visualization/navigation. The second rover, designed for longer-range (2-5 km) excursions from the lander, includes radar sounding/mapping equipment, a seismometer, and laser ranging devices. Power for all subsystems is supplied by a combination of solar cells, Ni-H batteries, and radioisotope thermoelectric generators. Communications are sequenced from rovers, sounding rockets, and remote sensors to the lander, then to the satellites, through the Deep Space Network to and from earth.

Baldwin, Curt↗
NASA NTRS1990

Use of non-geostationary orbits for a Ka-band Personal Access Satellite system

This paper discusses the use of satellites in circular orbits at altitudes high enough for continental U.S. (CONUS) visibility. This enables one satellite to relay signals between geographically separated earth stations within CONUS at any one time and thus bypasses the need for intersatellite links. System performance is examined for three circular satellite orbits at altitudes of 20,182 km, 10,353 km, and 5143 km. Inclination angles between the satellite orbit plane and the equatorial plane of 0, 45, and 90 deg are considered. The number of satellites required to provide continuous CONUS coverage is calculated.

Estabrook, Polly↗
NASA NTRS1968

Elliptic Capture Orbits for Missions to the Near Planets

Elliptic capture orbits around Mars and Venus have often been considered as means for reducing arrival and departure energy requirements for two-way missions. It had also generally been feared that the energy savings obtained by capturing a spacecraft into a highly elliptical orbit (rather than a near circular orbit of the same periapsis) would largely be offset by the penalties incurred in aligning the semi-major axis of the ellipse in such a way as to obtain the proper orientation of the departure hyperbola. This paper, presents the results of an analysis which takes into consideration the penalties arising from the requirement to match the orientation of the elliptical orbit with the asymptote of the departure hyperbola. The scientific aspects of elliptical orbits around the target planet are discussed, and it is shown that such orbits exhibit characteristics which may be considered advantageous or disadvantageous depending on the purpose of the mission. Alignment of ' the semi-major axis of the capture, ellipse relative to the, asymptote of the escape hyperbola was found not to be a critical requirement since the kinetic energy remains high over a substantial portion of the elliptical capture orbit. This 'means that the escape stage can operate efficiently even when ignited at some angle from the true periapsis point. Considerable freedom in choosing this angle is available at little propulsive cost. The resulting latitude in the choice of angles between arrival and escape asymptotes makes it possible to consider a wide variety of interplanetary transfers and planetary staytimes without the need for separate propulsive maneuvers to realign the capture ellipse before departure., Special consideration has also been g1ven to plane change maneuvers around the planet. These may be required for reasons of orbit dynamics or scientific experimentation and are not uniquely tied to elliptical captures. The sensitivity of the mass of the excursion module to the eccentricity of the capture orbit is discussed and mass-penalty diagrams are presented. It is shown that these penalties do not materially offset the large gains obtained through the use of the elliptical capture mode.

Casal, Federico G.↗
NASA NTRS1993

The Magellan aerobraking experiment - Attitude control simulation and preliminary flight results

On May 25, 1993 the Magellan spacecraft began an experiment on 'aerobraking', or use of aerodynamic drag from the Venusian atmosphere to provide deceleration. The Magellan spacecraft will strive to transition from a highly elliptical orbit into a near-circular orbit in approximately 70 days by successively dipping into the atmosphere to gain the necessary reduction in orbital velocity. Magellan will be the first three-axis stabilized spacecraft to perform this type of multipass aerobraking, and will also be attempting to achieve the largest velocity reduction to date, all with a spacecraft that was not designed for this application. This paper addresses redesign of the attitude control system onboard Magellan for use during aerobraking. The dynamics and control of the spacecraft were simulated, and these results are presented and summarized. Modifications necessary to the existing flight control system and, specifically, to the mass expulsion control scheme are discussed. Changes were minimized though, due to the low-cost, high-risk approach to Magellan's aerobraking experiment. Some initial flight results from the first week of aerobraking operations will be highlighted.

Carpenter, Anita S.↗
NASA NTRS2007

Mars Reconnaissance Orbiter Aerobraking Daily Operations and Collision Avoidance

The Mars Reconnaissance Orbiter reached Mars on March 10, 2006 and performed a Mars orbit insertion maneuver of 1 km/s to enter into a large elliptical orbit. Three weeks later, aerobraking operations began and lasted about five months. Aerobraking utilized the atmospheric drag to reduce the large elliptical orbit into a smaller, near circular orbit. At the time of MRO aerobraking, there were three other operational spacecraft orbiting Mars and the navigation team had to minimize the possibility of a collision. This paper describes the daily operations of the MRO navigation team during this time as well as the collision avoidance strategy development and implementation.

Long, Stacia M.↗
NASA NTRS2020

On the Maximum Stellar Rotation to Form a Black Hole without an Accompanying Luminous Transient

The collapse of a massive star with low angular momentum content is commonly argued to result in the formation of a black hole without an accompanying bright transient. Our goal in this Letter is to understand the flow in and around a newly-formed black hole, involving accretion and rotation, via general relativistic hydrodynamics simulations aimed at studying the conditions under which infalling material can accrete without forming a centrifugally supported structure and, as a result, generate no effective feedback. If the feedback from the black hole is, on the other hand, significant, the collapse would be halted and we suggest that the event is likely to be followed by a bright transient. We find that feedback is only efficient if the specific angular momentum of the infalling material at the innermost stable circular orbit exceeds that of geodesic circular flow at that radius by at least≈20%. We use the results of our simulations to constrain the maximal stellar rotation rates of the disappearing massive progenitors PHL293B-LBV and N6946-BH1, and to provide an estimate of the overall rate of disappearing massive stars. We find that about a few percent of single O-type stars with measured rotational velocities are expected to spin below the critical value before collapse and are thus predicted to vanish without a trace.

Scott C Noble↗
NASA NTRS2003

Propulsion Options for the Global Precipitation Measurement Core Satellite

This study was conducted to evaluate several propulsion system options for the Global Precipitation Measurement (GPM) core satellite. Orbital simulations showed clear benefits for the scientific data to be obtained at a constant orbital altitude rather than with a decay/reboost approach. An orbital analysis estimated the drag force on the satellite will be 1 to 12 mN during the five-year mission. Four electric propulsion systems were identified that are able to compensate for these drag forces and maintain a circular orbit. The four systems were the UK-10/TS and the NASA 8 cm ion engines, and the ESA RMT and RITl0 EVO radio-frequency ion engines. The mass, cost, and power requirements were examined for these four systems. The systems were also evaluated for the transfer time from the initial orbit of 400 x 650 km altitude orbit to a circular 400 km orbit. The transfer times were excessive, and as a consequence a dual system concept (with a hydrazine monopropellant system for the orbit transfer and electric propulsion for drag compensation) was examined. Clear mass benefits were obtained with the dual system, but cost remains an issue because of the larger power system required for the electric propulsion system. An electrodynamic tether was also evaluated in this trade study.

Cardiff, Eric H.↗