Search NASA⌕ Search

SEARCH · Search NASA

Results for “Elliptical 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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Launching rockets and small satellites from the lunar surface

Scientific payloads and their propulsion systems optimized for launch from the lunar surface differ considerably from their counterparts for use on earth. For spin-stabilized payloads, the preferred shape is a large diameter-to-length ratio to provide stability during the thrust phase. The rocket motor required for a 50-kg payload to reach an altitude of one lunar radius would have a mass of about 41 kg. To place spin-stabilized vehicles into low altitude circular orbits, they are first launched into an elliptical orbit with altitude about 840 km at aposelene. When the spacecraft crosses the desired circular orbit, small retro-rockets are fired to attain the appropriate direction and speed. Values of the launch angle, velocity increments, and other parameters for circular orbits of several altitudes are tabulated. To boost a 50-kg payload into a 100-km altitude circular orbit requires a total rocket motor mass of about 90 kg.

Anderson, K. A.↗
NASA NTRS2010

Deliberate Satellite Fragmentations and their Effects on the Long-Term Space Environment

Since 1964 at least 56 spacecraft and two launch vehicle upper stages have been deliberately fragmented while in Earth orbit. Many of these events have had no long-lasting effects on the near-Earth space environment, but one represents the most devastating satellite breakup in history that will pose hazards to operational spacecraft in low Earth orbit for decades to come. International space debris mitigation guidelines now call for avoiding the creation of long-lived debris from intentional satellite fragmentations. This paper summarizes the reasons for and environmental consequences of deliberate satellite fragmentations. Contrary to popular belief, only one in five deliberate fragmentations have been related to the testing of anti-satellite weapon systems, for which only one such test has occurred during the past 25 years. Other reasons for deliberate satellite fragmentations range from engineering tests to protecting national security information. Whereas the majority of deliberate satellite fragmentations have occurred in low Earth orbits, some have involved spacecraft in highly elliptical orbits. The former Soviet Union and the current Russian Federation have been responsible for 90% of all identified deliberate on-orbit satellite fragmentations.

Johnson, N. L.↗
NASA NTRS2013

Combined Global Navigation Satellite Systems in the Space Service Volume

Besides providing position, velocity, and timing (PVT) for terrestrial users, the Global Positioning System (GPS) is also being used to provide PVT information for earth orbiting satellites. In 2006, F. H. Bauer, et. al., defined the Space Service Volume in the paper GPS in the Space Service Volume , presented at ION s 19th international Technical Meeting of the Satellite Division, and looked at GPS coverage for orbiting satellites. With GLONASS already operational, and the first satellites of the Galileo and Beidou/COMPASS constellations already in orbit, it is time to look at the use of the new Global Navigation Satellite Systems (GNSS) coming into service to provide PVT information for earth orbiting satellites. This presentation extends GPS in the Space Service Volume by examining the coverage capability of combinations of the new constellations with GPS GPS was first explored as a system for refining the position, velocity, and timing of other spacecraft equipped with GPS receivers in the early eighties. Because of this, a new GPS utility developed beyond the original purpose of providing position, velocity, and timing services for land, maritime, and aerial applications. GPS signals are now received and processed by spacecraft both above and below the GPS constellation, including signals that spill over the limb of the earth. Support of GPS space applications is now part of the system plan for GPS, and support of the Space Service Volume by other GNSS providers has been proposed to the UN International Committee on GNSS (ICG). GPS has been demonstrated to provide decimeter level position accuracy in real-time for satellites in low Earth orbit (centimeter level in non-real-time applications). GPS has been proven useful for satellites in geosynchronous orbit, and also for satellites in highly elliptical orbits. Depending on how many satellites are in view, one can keep time locked to the GNSS standard, and through that to Universal Time as long as at least one satellite is in view (the longest duration with no satellites in view is important in determining the maximum clock drift from GNSS time). Instantaneous position requires four satellites in view, but because orbital motion is predictable, it is possible to build up knowledge of the orbital position gradually through time without a need for constant four satellite coverage. However, it is desirable to have four satellite coverage when performing satellite maneuvers, since there can be significant changes in velocity, leading to large changes in orbit parameter, causing substantial divergence in position over time. The Space Service Volume has been defined as the volume between three thousand km altitude and geosynchronous altitude, and can be divided into medium orbit services between three thousand km altitude and eight thousand km altitude, and high orbit services above eight thousand km. The Terrestrial Service Volume includes the Earth s surface, the atmosphere, and space below the altitude of three thousand km. The Terrestrial Service Volume is the volume within which the GNSS systems will have very similar performance to the Earth surface, and satellites need only use the signals specified to provide terrestrial performance. Above three thousand km the use of signals passing by the Earth s limb becomes important, so it is desirable to have additional information on signal strength, phase delay, and group delay covering wider beam angles than are needed for terrestrial service (and which can be obtained by monitoring GNSS signals from the Earth s surface). This presentation will look at each of the new GNSS constellations in combination with GPS (GLONASS with GPS, Galileo with GPS, Beidou/COMPASS with GPS), and also at the combination of all four GNSS systems. The presentation will largely follow the format of GPS in the Space Service Volume , presenting data on the availability of one, two, three, or four of the various combinations of GNSS constellation satelles at approximately two thousand grid points evenly spaced and fixed in longitude and latitude, the duration of the longest single-fold outages (intervals when no satellites are available), and the duration of the longest four-fold outages (intervals when fewer than four satellites are available) at several altitudes. Following the original paper, we will use the altitudes of three hundred km (typical LEO satellite, and within the Terrestrial Service Volume), at three thousand km (border between Terrestrial Service Volume and Space Service Volume), at eight thousand km (the border between medium and high orbit service within the Space Service Volume), at fifteen thousand km (just below the GNSS constellations), at twenty five thousand km (just above the GNSS constellations), at thirty six thousand, five hundred km (limit of Space Service Volume definition, geosynchronous altitude), and at seventy thousand km (to show the potential usefulness of GNSS beyond geosynchronous altitude).

Force, Dale A.↗
NASA NTRS1984

An analytic solution for the orbital perturbations of the Venus Radar Mapper due to gravitational harmonics

Hill's variational equations are solved analytically for the orbital perturbations of a spacecraft nominally in an elliptic orbit around a non-spherical body. The rotation of the central planet about its spin-axis is not considered in the analysis. The perturbations are restricted to the planetary gravitational harmonics only. An extremely simple algorithm is derived to transform the spherical harmonic potentials to the orbital coordinate system, and the resulting accelerations are shown to be simply trigonometric functions of the true anomaly. With the principal matrix solution for the differential equations of the adjoint system given in closed form, the orthogonality of the trigonometric functions makes it possible to obtain an analytic solution for the non-homogeneous problem, at intervals of 2 pi in true anomaly. The solution for orbital perturbations can be extended over several revolutions by applying well-known results from Floquet's theory. The technique is demonstrated with results presented on the spacecraft periapsis altitude for the forthcoming Venus Radar Mapper Mission.

Vijayaraghavan, A.↗
NASA NTRS1986

An exact solution for orbit view-periods from a station on a tri-axial ellipsoidal planet

This paper presents the concise exact solution for predicting view-periods to be observed from a masked or unmasked tracking station on a tri-axial ellipsoidal surface. The new exact approach expresses the azimuth and elevation angles of a spacecraft in terms of the station-centered geodetic topocentric coordinates in an elegantly concise manner. A simple and efficient algorithm is developed to avoid costly repetitive computations in searching for neighborhoods near the rise and set times of each satellite orbit for each station. Only one search for each orbit is necessary for each station. Sample results indicate that the use of an assumed spherical earth instead of an 'actual' tri-axial ellipsoidal earth could introduce an error up to a few minutes in a view-period prediction for circular orbits of low or medium altitude. For an elliptical orbit of high eccentricity and long period, the maximum error could be even larger. The analytic treatment and the efficient algorithm are designed for geocentric orbits, but they should be applicable to interplanetary trajectories by an appropriate coordinates transformation at each view-period calculation. This analysis can be accomplished only by not using the classical orbital elements.

Tang, C. C. H.↗
NASA NTRS1992

Debris about asteroids: Where and how much?

We summarize several recent findings on the size and shape of the region within which material can stably orbit an asteroid. If the asteroid (with assumed density 2.38 g/cu cm) circles the Sun at 2.55 AU, co-planar prograde material will remain trapped whenever started on unperturbed circular orbits at less than about 220 R(sub A) (asteroid radii); co-planar retrograde particles are stable out twice as far. Our 3-D stability surface, which encloses several hundred numerically calculated orbits that start with various inclinations, is shaped like a sphere with its top and bottom sliced off; its dimensions scale like the Hill radius =(mu/3)(exp 1/3)R, where mu is the asteroid-to-solar mass ratio and R is the asteroid's orbital radius. If the asteroid moves along an elliptical orbit, a fairly reliable indicator of the dimensions of the hazard zone is the size of its Hill sphere at the orbit's pericenter. Grains with radii less than a few mm will be lost through the action of radiation forces which can induce escape or cause collisions with the asteroid on times scales of a few years; interplanetary micrometeoroids produce collisional break-up of these particles in approximately 10(exp 4) yrs. The effects of Jupiter and of asteroids that pass close to the target asteroid allow particles to diffuse from the system, again shrinking the hazard zone. None of the considered sources-primordial formation, debris spalled off the asteroid during micrometeoroid impact, captured interplanetary particles, feeder satellites, etc., seem capable of densely populating distant orbits from the asteroid. No certain detections of debris clouds or of binary asteroids have been made. Thus, it seems highly unlikely that a spacecraft fly-by targeted at 100 R(sub A) from the asteroid over its orbital pole would encounter any material.

Burns, Joseph A.↗
NASA NTRS1993

Infrared imaging spectroscopy of the Galactic center - Distribution and motions of the ionized gas

High spatial spectral resolution IR images of the Galactic center in the Br-gamma recombination line of hydrogen were taken. A coherent filament of gas extending from north of IRS 1, curving around IRS 16/Sgr A complex, and continuing to the southwest, is seen. Nine stellar sources have associated Br-gamma emission. The total Br-gamma line flux in the filament is approximately 3 x 10 exp -15 W/sq m. The distribution and kinematics of the northern arm suggest orbital motion; the observations are accordingly fit with elliptical orbits in the field of a central point of mass.

Herbst, T. M.↗
NASA NTRS1961

A Study of the Positions and Velocities of a Space Station and a Ferry Vehicle During Rendezvous and Return

A study is made of the families of non-thrusting ascent trajectories during rendezvous with an orbiting space station and the descent trajectories to the earth's atmosphere. Equations of motion are derived and results are shown for two typical orbits of the station (one circular and one elliptic orbit}. Boundaries of launch (at time of booster burnout) and rendezvous conditions are given and the effects of delays in launch time are discussed.

Eggleston, John M.↗
NASA NTRS1992

Contribution of Doppler and interferometric tracking during the Magellan approach to Venus

On May 4, 1989, the Magellan spacecraft began its 463 day and 1.3 billion km earth to Venus interplanetary cruise. Magellan's Venus approach trajectory required prediction accuracy of 6.1 seconds in arrival time and 126 km in the position of closest approach to achieve a desired Venus orbit. Data collection from Magellan's SAR required an elliptical orbit with an inclination between 84 deg and 86 deg to the Venus equator, a period between 3.1 and 3.3 hours, a periapsis altitude within 275 km and 325 km and a latitude of periapsis between 0 deg and 10 deg North. Predictions of Magellan's arrival time and closest approach to Venus were derived from cruise trajectories determined exclusively from coherent two-way Doppler using both S-band (2.3 GHz) uplink/downlink and X-band (8.4 GHz) uplink/downlink and X-band spacecraft-quasar interferometric delay data. Analysis of data reduction strategies employing various data arc lengths and data combinations shows that interferometric tracking data, used in conjunction with Doppler, improved trajectory solution accuracy and robustness.

Graat, E. J.↗
NASA NTRS1995

Method for the calculation of spacecraft umbra and penumbra shadow terminator points

A method for calculating orbital shadow terminator points is presented. The current method employs the use of an iterative process which is used for an accurate determination of shadow points. This calculation methodology is required since orbital perturbation effects can introduce large errors when a spacecraft orbits a planet in a high altitude and/or highly elliptical orbit. To compensate for the required iteration methodology, all reference frame change definitions and calculations are performed with quaternions. Quaternion algebra significantly reduces the computational time required for the accurate determination of shadow terminator points.

Ortizlongo, Carlos R.↗
NASA NTRS2003

Formation Tetrahedron Design for Phase I of the Magnetospheric Multiscale Mission

The Magnetospheric Multiscale Mission (MMS) is a NASA mission intended to make fundamental advancements in our understanding of the Earth's Magnetosphere. There are three processes that MMS will study including magnetic reconnection, charged particle acceleration, and turbulence. There are four phases in the nominal mission and this work addresses some of the outstanding issues in phase I. The nominal phase I orbit is 1.2 x 12 R(sub e) highly elliptic orbit with four spacecraft nominally forming a regular tetrahedron. In this paper we investigate the relative dynamics of the four MMS spacecraft about an assumed reference orbit. There are several tetrahedron dimensions required in Phase I of the mission and in this work we design optimal tetrahedrons for the 10 km baseline. The performance metric used in the optimization process is directly related to the science return, and is based on an extension of previous work performed by Glassmeier. The optimizer we use is a commercially available Sequential Quadratic Programming (SQP) routine. Multiple optimal solutions are found, and we characterize how the performance of the formation varies between different regions of the reference orbit.

Hughes, Steven P.↗
NASA NTRS1982

Periodic orbits of the elliptic restricted problem for the Sun-Jupiter-Saturn system

A systematic approach to generate periodic orbits in the elliptic restricted problem of three bodies is introduced. The approach is based on (numerical) continuation from periodic orbits of the first and second kind in the circular restricted problem to periodic orbits in the elliptic restricted problem. Two families of periodic orbits of the elliptic restricted problem are found by this approach. The mass ratio of the primaries of these orbits is equal to that of the Sun-Jupiter system. The sidereal mean motions between the infinitesimal body and the smaller primary are in a 2:5 resonance, so as to approximate the Sun-Jupiter-Saturn system. The lineaar stabilities of these periodic orbits are studied as functions of the eccentricities of the primaries and of the infinitesimal body. The results show that both stable and unstable periodic orbits exist in the elliptic restricted problem that are close to the actual Sun-Jupiter-Saturn system. However, the periodic orbit closest to the actual Sun-Jupiter-Saturn system is (linearly) stable.

Kwok, J. H.↗