Results for “orbits” — Page 27

NASA NTRS2012

The Effect of Geocenter Motion on Jason-2 and Jason-1 Orbits and the Mean Sea Level

We have investigated the impact of geocenter motion on Jason-2 orbits. This was accomplished by computing a series of Jason-1, Jason-2 GPS-based and SLR/DORIS-based orbits using ITRF2008 and the IGS repro1 framework based on the most recent GSFC standards. From these orbits, we extract the Jason-2 orbit frame translational parameters per cycle by the means of a Helmert transformation between a set of reference orbits and a set of test orbits. The fitted annual and seasonal terms of these time-series are compared to two different geocenter motion models. Subsequently, we included the geocenter motion corrections in the POD process as a degree-1 loading displacement correction to the tracking network. The analysis suggested that the GSFC's Jason-2 std0905 GPS-based orbits are closely tied to the center of mass (CM) of the Earth whereas the SLR/DORIS std0905 orbits are tied to the center of figure (CF) of the ITRF2005 (Melachroinos et al., 2012). In this study we extend the investigation to the centering of the GPS constellation and the way those are tied in the Jason-1 and Jason-2 POD process. With a new set of standards, we quantify the GPS and SLR/DORIS-based orbit centering during the Jason-1 and Jason-2 inter-calibration period and how this impacts the orbit radial error over the globe, which is assimilated into mean sea level (MSL) error, from the omission of the full term of the geocenter motion correction.

Melachroinos, Stavros A.↗

NASA NTRS2013

The Eccentric Behavior of Nearly Frozen Orbits

Frozen orbits are orbits which have only short-period changes in their mean eccentricity and argument of periapse, so that they basically keep a fixed orientation within their plane of motion. Nearly frozen orbits are those whose eccentricity and argument of periapse have values close to those of a frozen orbit. We call them "nearly" frozen because their eccentricity vector (a vector whose polar coordinates are eccentricity and argument of periapse) will stay within a bounded distance from the frozen orbit eccentricity vector, circulating around it over time. For highly inclined orbits around the Earth, this distance is effectively constant over time. Furthermore, frozen orbit eccentricity values are low enough that these orbits are essentially eccentric (i.e., off center) circles, so that nearly frozen orbits around Earth are bounded above and below by frozen orbits.

spacecraft↗

NASA NTRS2016

The Impact of New Trends in Satellite Launches on Orbital Debris Environment

The main goal of this study is to examine the impact of new trends in satellite launch activities on the orbital debris environment and collision risk. Starting from the launch of the first artificial satellite in 1957, space borne technology has become an indispensable part of our lives. More than 6,000 satellites have been launched into Earth orbit. Though the annual number of satellites launched stayed flat for many decades, the trend has recently changed. The satellite market has been undergoing a major evolution with new space companies replacing the traditional approach of deploying a few large, complex and costly satellites with an approach to use a multitude of smaller, less complex and cheaper satellites. This new approach creates a sharp increase in the number of satellites and so the historic trends are no longer representative. As a foundation for this study, a scenario for satellite deployments based on the publicly announced future satellite missions has been developed. These constellation-deploying companies include, but are not limited to, Blacksky, CICERO, EROS, Landmapper, Leosat, Northstar, O3b, OmniEarth, OneWeb, Orbcomm, OuterNet, PlanetIQ, Planet Labs, Radarsat, RapidEye Next Generation, Sentinel, Skybox, SpaceX, and Spire. Information such as the annual number of launches, the number of orbital planes to be used by the constellation, as well as apogee, perigee, inclination, spacecraft mass and area were included or approximated. Besides the production of satellites, a widespread ongoing effort to enhance orbital injection capabilities will allow delivery of more spacecraft more accurately into Earth orbits. A long list of companies such as Microcosm, Rocket Lab, Firefly Space Systems, Sierra Nevada Corporation and Arca Space Corporation are developing new launch vehicles dedicated for small satellites. There are other projects which intend to develop interstages with propulsive capabilities which will allow the deployment of satellites into their desired orbits beyond the restrictions of the launch vehicle used. These near future orbital injection technologies are also covered in the developed scenario. Using the above-mentioned background information, this study aims to examine how the orbital debris environment will be affected from the new dynamics of the emerging space markets. We developed a simulation tool that is capable of propagating the objects in a given deployment scenario with variable-sized time-steps as small as one second. Over the course of the run, the software also detects collisions; additional debris objects are then created according to the NASA breakup model and are fed back into the simulation framework. Examining the simulation results, the total number of particles to accumulate in different orbits can be monitored and the number of conjunctions can be tracked to assess the collision risks. The simulation makes it possible to follow the short- and long-term effects of a particular satellite or constellation on the space environment. Likewise, the effects of changes in the debris environment on a particular satellite or constellation can be evaluated. It is authors hope that the results of this paper and further utilization of the developed simulation tool will assist in the investigation of more accurate deorbiting metrics to replace the generic 25-year disposal guidelines, as well as to guide future launches toward more sustainable and safe orbits.

New Space↗

NASA NTRS2017

Small Orbital Stereo Tracking Camera Technology Development

Any exploration vehicle assembled or Spacecraft placed in LEO or GTO must pass through this debris cloud and survive. Large cross section, low thrust vehicles will spend more time spiraling out through the cloud and will suffer more impacts.Better knowledge of small debris will improve survival odds. Current estimated Density of debris at various orbital attitudes with notation of recent collisions and resulting spikes. Orbital Debris Tracking and Characterization has now been added to NASA Office of Chief Technologists Technology Development Roadmap in Technology Area 5 (TA5.7)[Orbital Debris Tracking and Characterization] and is a technical gap in the current National Space Situational Awareness necessary to safeguard orbital assets and crews due to the risk of Orbital Debris damage to ISS Exploration vehicles. The Problem: Traditional orbital trackers looking for small, dim orbital derelicts and debris typically will stare at the stars and let any reflected light off the debris integrate in the imager for seconds, thus creating a streak across the image. The Solution: The Small Tracker will see Stars and other celestial objects rise through its Field of View (FOV) at the rotational rate of its orbit, but the glint off of orbital objects will move through the FOV at different rates and directions. Debris on a head-on collision course (or close) will stay in the FOV at 14 Km per sec. The Small Tracker can track at 60 frames per sec allowing up to 30 fixes before a near-miss pass. A Stereo pair of Small Trackers can provide range data within 5-7 Km for better orbit measurements.

Bryan, Tom↗

NASA NTRSDate not supplied

TechEdSat 7, 10, 13, 15: Exo-brake Experiments on the ISS, First Virgin Orbit, and First Firefly-Alpha Test Flights.

The TechEdSat flight series (TES-n), developed by the Nano Orbital Workshop (NOW) group at NASA Ames, has been studying cube satellite re-entry technologies with a focus on Exo-Brake drag device research. An exo-atmospheric braking device, the Exo-Brake uses the tenuous upper atmosphere acting in a free molecular flow regime to enable controllable adjustment of the drag profile of the host spacecraft, currently enabling rapid disposal of small spacecraft after mission conclusion, and eventual targeted de-orbit to a desired entry point at the Von Karman line, or approximately 100km in altitude. The TES-7 mission, flown on the first successful Virgin Orbit flight January 17, 2021, was injected into a 500km, 61-degree inclination orbit, with an expected orbital lifetime of approximately ten years. After successful deployment of a unique hydrogen gas-cell inflation design ‘disposal’-type Mylar Exo-Brake intended to rapidly de-orbit the spacecraft, the maximum expected orbital lifetime of the TES-7 spacecraft was successfully reduced from ten years to 1.3 years, an 87% reduction in orbit lifetime via a passive, fuel-less system. The subsequent TES-13 mission utilized a new, less complex ‘disposal’-type Exo-Brake design made from collapsible rigid struts rather than inflated struts to reduce the engineering, construction, and safety complexities introduced by the prior hydrogen gas inflation design. The rigid strut Exo-Brake design is stowed via rotational compression and deploys using only stored spring energy once released via electronic actuator. This new rotation compression storage design has become the new Exo-Brake design standard used on TES missions thanks to its simplicity. Following on the success of this new Exo-Brake design demonstrated by TES-13, the design was further refined to enabled active manipulation of the Exo-Brake effective drag area though a winch-like system. Both the TES-10 and TES-15 spacecraft were equipped with geometrically similar Exo-Brake devices capable of such active drag modulation, with the intent of studying the operational impact of having such devices on cube satellite missions. To increase the survival time of the Exo-Brake at low altitudes and thus gain more experiment time and better guidance capability, TES-15 was equipped with a modulated Exo-Brake constructed from 3M™ Nextel™ 440 ceramic oxide fiber, the same fabric material used in the Space Shuttle TPS system, rather than Mylar. After launching on Alpha Flight 2 October 1st, 2022, Firefly’s first successful launch, TES-15 met the same modulation restrictions as TES-10, with modulation prohibited above 200km to avoid possible collision with other spacecraft. In this case, a low deployment orbit caused an extremely short mission life, which, coupled with delayed spacecraft identification, exacerbated collision concerns and caused performance data of the high-temperature Exo-Brake design to be inconclusive. As such, the Nano Orbital Workshop has taken a new approach of only conducting modulated Exo-Brake research on missions deploying below ISS altitude to avoid collision avoidance restrictions, and is working to improve spacecraft identification and location reporting techniques to increase experiment durations. Upcoming TES mission will therefore use one of two classes of Exo-Brake depending on their target altitude, as to be described in the submission.

M Murbach↗

NASA NTRSDate not supplied

TechEdSat 7, 10, 13, 15: Exo-brake Experiments on the ISS, First Virgin Orbit, and First Firefly-Alpha Test Flights

The TechEdSat flight series (TES-n), developed by the Nano Orbital Workshop (NOW) group at NASA Ames, has been studying cube satellite re-entry technologies with a focus on Exo-Brake drag device research. An exo-atmospheric braking device, the Exo-Brake uses the tenuous upper atmosphere acting in a free molecular flow regime to enable controllable adjustment of the drag profile of the host spacecraft, currently enabling rapid disposal of small spacecraft after mission conclusion, and eventual targeted de-orbit to a desired entry point at the Von Karman line, or approximately 100km in altitude. The TES-7 mission, flown on the first successful Virgin Orbit flight January 17, 2021, was injected into a 500km, 61-degree inclination orbit, with an expected orbital lifetime of approximately ten years. After successful deployment of a unique hydrogen gas-cell inflation design ‘disposal’-type Mylar Exo-Brake intended to rapidly de-orbit the spacecraft, the maximum expected orbital lifetime of the TES-7 spacecraft was successfully reduced from ten years to 1.3 years, an 87% reduction in orbit lifetime via a passive, fuel-less system. The subsequent TES-13 mission utilized a new, less complex ‘disposal’-type Exo-Brake design made from collapsible rigid struts rather than inflated struts to reduce the engineering, construction, and safety complexities introduced by the prior hydrogen gas inflation design. The rigid strut Exo-Brake design is stowed via rotational compression and deploys using only stored spring energy once released via electronic actuator. This new rotation compression storage design has become the new Exo-Brake design standard used on TES missions thanks to its simplicity. Following on the success of this new Exo-Brake design demonstrated by TES-13, the design was further refined to enabled active manipulation of the Exo-Brake effective drag area though a winch-like system. Both the TES-10 and TES-15 spacecraft were equipped with geometrically similar Exo-Brake devices capable of such active drag modulation, with the intent of studying the operational impact of having such devices on cube satellite missions. To increase the survival time of the Exo-Brake at low altitudes and thus gain more experiment time and better guidance capability, TES-15 was equipped with a modulated Exo-Brake constructed from 3M™ Nextel™ 440 ceramic oxide fiber, the same fabric material used in the Space Shuttle TPS system, rather than Mylar. After launching on Alpha Flight 2 October 1st, 2022, Firefly’s first successful launch, TES-15 met the same modulation restrictions as TES-10, with modulation prohibited above 200km to avoid possible collision with other spacecraft. In this case, a low deployment orbit caused an extremely short mission life, which, coupled with delayed spacecraft identification, exacerbated collision concerns and caused performance data of the high-temperature Exo-Brake design to be inconclusive. As such, the Nano Orbital Workshop has taken a new approach of only conducting modulated Exo-Brake research on missions deploying below ISS altitude to avoid collision avoidance restrictions, and is working to improve spacecraft identification and location reporting techniques to increase experiment durations. Upcoming TES mission will therefore use one of two classes of Exo-Brake depending on their target altitude, as to be described in the submission.

Marcus Murbach↗

NASA NTRS2006

Orbit design and optimization based on global telecommunication performance metrics

The orbit selection of telecommunications orbiters is one of the critical design processes and should be guided by global telecom performance metrics and mission-specific constraints. In order to aid the orbit selection, we have coupled the Telecom Orbit Analysis and Simulation Tool (TOAST) with genetic optimization algorithms. As a demonstration, we have applied the developed tool to select an optimal orbit for general Mars telecommunications orbiters with the constraint of being a frozen orbit. While a typical optimization goal is to minimize tele-communications down time, several relevant performance metrics are examined: 1) area-weighted average gap time, 2) global maximum of local maximum gap time, 3) global maximum of local minimum gap time. Optimal solutions are found with each of the metrics. Common and different features among the optimal solutions as well as the advantage and disadvantage of each metric are presented. The optimal solutions are compared with several candidate orbits that were considered during the development of Mars Telecommunications Orbiter.

genetic algorithms↗

NASA NTRS2013

Dawn Orbit Determination Team : Trajectory Modeling and Reconstruction Processes at Vesta

The NASA Dawn spacecraft was launched on September 27, 2007 on a mission to study the asteroid belt's two largest objects, Vesta and Ceres. It is the first deep space orbiting mission to demonstrate solar-electric ion propulsion, providing the necessary delta-V to enable capture and escape from two extraterrestrial bodies. At this time, Dawn has completed its science campaign at Vesta and is currently on its journey to Ceres, where it will arrive in mid-2015. The spacecraft spent over a year in orbit around Vesta from July 2011 through August 2012, capturing science data during four dedicated orbit phases. In order to maintain the reference orbits necessary for science and enable the transfers between those orbits, precise and timely orbit determination was required. The constraints associated with low-thrust ion propulsion coupled with the relatively unknown a priori gravity and rotation models for Vesta presented unique challenges for the Dawn orbit determination team. While [1] discusses the prediction performance of the orbit determination products, this paper discusses the dynamics models, filter configuration, and data processing implemented to deliver a rapid orbit determination capability to the Dawn project.

Vesta↗

DOE OSTI2025

Observation of Orbital-Selective Dual Modulations in an Anisotropic Antiferromagnetic Kagome Metal TbTi 3 ⁢Bi 4

Orbital selectivity is pivotal in dictating the phase diagrams of multiorbital systems, with prominent examples including the orbital-selective Mott phase and superconductivity. The intercalation of anisotropic layers represents an effective method for enhancing orbital selectivity and thereby shaping the low-energy physics of multiorbital systems. Despite its potential, related experimental studies, especially those elucidating the correlation between orbital selectivity and magnetism, remain limited. In this work, we systematically examine the interplay between orbital selectivity and magnetism in the newly discovered anisotropic kagome TbTi 3 ⁢Bi 4 single crystal, and report the coexistence of orbital-selective dual-band modulations (𝑞 1 ∼ 1/3⁢𝑎*, 𝑞 2 ∼ 0.28⁢𝑏*) within the antiferromagnetic (AFM) state. By combining soft x-ray and vacuum ultraviolet angle-resolved photoemission spectroscopy measurements, neutron powder diffraction, scanning tunneling microscopy, and density-functional-theory calculations, we identify these dual-band reconstructions as manifestations of the AFM order driven by a (approximately 1/3, 0.28, 0) nesting instability of the intercalated Tb 5⁢𝑑 𝑥⁢𝑧 orbitals. These orbital-selective modulations induce unusual momentum-dependent band folding and lead to the emergence of Dirac cones only at the $\bar{M}$ 1 point, signaling a topological phase transition in the AFM state. Importantly, the discovery of orbital-selective (approximately 1/3, 0.28, 0) AFM order offers crucial insights into the mechanism underlying the fractional magnetization plateau in this kagome AFM metal. Our findings not only underscore the essential role of both conducting and localized electrons in determining the magnetic orders of Ln⁢Ti 3 ⁢Bi 4 (Ln = lanthanide) kagome metals but also offer a pathway for manipulating magnetism through selective control of anisotropic electronic structures.

Zhang, Renjie [Shanghai Jiao Tong University (Chin↗

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 NTRS1972

Orbit selection considerations for earth observatory satellites

All the possible circular sun-synchronous orbits for earth observatory satellites are displayed, which have orbital altitudes between 740 and 1,115 kilometers (approximately 400 to 600 nautical miles) and have either a 16, 17, or 18 day repeat cycle for the ground trace. It was found that there are 9 orbits with a 16 day repeat cycle, 17 orbits with a 17 day repeat cycle, and 7 orbits with an 18 day repeat cycle meeting the requirements. For each of these, various characteristics such as ideal ground trace patterns, swathing patterns, and daily drift are displayed. The solar elevation angle along the orbit and the change in the solar elevation angle with season are given. This presentation of general orbital characteristics has application to finding approximate orbital elements and selecting orbits for many types of earth sensing satellite missions.

Cooley, J. L.↗

NASA NTRS1971

Orbiting propellant depot safety. Volume 1: Management summary report

Orbital missions are considered which require the use of orbiting vehicles, e.g., a cislunar shuttle, that is either chemically or nuclear propelled, functioning as orbit-to-orbit shuttles combined with space tugs for servicing earth-orbiting payloads. Such vehicles may be stationed in a low earth orbit from which they would initiate and terminate flights. The flight frequency of these vehicles indicates that large quantities of propellants will have to be delivered to them in orbit. Orbiting propellant depots, in both geocentric and selenocentric orbits, are being considered as candidate methods of making the required propellants readily available. An assessment of the potential safety hazards associated with the operation of such a depot is summarized.

Source record↗

NASA NTRS1976

Effects of solar radiation on the orbits of small particles

A modification of the Robertson (1937) equations of particle motion in the presence of solar radiation is developed which allows for partial reflection of sunlight as a result of rapid and varying particle rotations caused by interaction with the solar wind. The coefficients and forces in earlier forms of the equations are compared with those in the present equations, and secular rates of change of particle orbital elements are determined. Orbital dimensions are calculated in terms of time, probable sizes and densities of meteoric and cometary particles are estimated, and times of infall to the sun are computed for a particle moving in an almost circular orbit and a particle moving in an elliptical orbit of high eccentricity. Changes in orbital elements are also determined for particles from a long-period sun-grazing comet. The results show that the time of infall to the sun from a highly eccentric orbit is substantially shorter than from a circular orbit with a radius equal to the mean distance in the eccentric orbit. The possibility is considered that the free orbital kinetic energy of particles drawn into the sun may be the energy source for the solar corona.

Lyttleton, R. A.↗

NASA NTRS1978

Space Shuttle Orbiter Approach and Landing Test: Final Evaluation Report

The Approach and Landing Test Program consisted of a series of steps leading to the demonstration of the capability of the Space Shuttle orbiter to safely approach and land under conditions similar to those planned for the final phases of an orbital flight. The tests were conducted with the orbiter mounted on top of a specially modified carrier aircraft. The first step provided airworthiness and performance verification of the carrier aircraft after modification. The second step consisted of three taxi tests and five flight tests with an inert unmanned orbiter. The third step consisted of three mated tests with an active manned orbiter. The fourth step consisted of five flights in which the orbiter was separated from the carrier aircraft. For the final two flights, the orbiter tail cone was replaced by dummy engines to simulate the actual orbital configuration. Landing gear braking and steering tests were accomplished during rollouts following the free flight landings. Ferry testing was integrated into the Approach and Landing Test Program to the extent possible. In addition, four ferry test flights were conducted with the orbiter mated to the carrier aircraft in the ferry configuration after the free-flight tests were completed.

Source record↗

NASA NTRS1985

The long term behaviour of Earth orbits and the implications for debris control

GEO orbits with inclinations which remain less than 45 degrees are very stable. Although the inclination of the orbital plane may vary as much as 15 degrees over a period of about 50 years, the orbit altitude will always remain within a few hundred kilometers of geosynchronous altitude, GEO orbits with inclinations greater than 45 degrees exhibit remarkable instabilities ir the eccentricity due to gravitational resonance. Over a period of a century the eccentricity can reach such a large value that reentry is a possibility. The combined effects of the Sun, Moon, and oblate Earth play a significant role in determining the lifetime of a GEO transfer orbit. Depending on the initial orientation of the orbital plane with respect to the Sun and Moon, lifetimes may vary from under 6 months to over several hundred years. Transfer orbits with inclinations over 45 degrees show strong instabilities in the perigee altitude resulting in generally short lifetimes of less than a few years. All transfer orbits can be designed to decay within one year if the initial perigee altitude is less than 231 km. However, there are restrictions on orbit plane placement and time of year of launch.

Mueller, A. C.↗

NASA NTRS1985

Orbits: Computer simulation

In rotating machinery dynamics an orbit (Lissajous curve) represents the dynamic path of the shaft centerline motion during shaft rotation and resulting precession. The orbit can be observed with an oscilloscope connected to XY promixity probes. The orbits can also be simulated by a computer. The software for HP computer simulates orbits for two cases: (1) Symmetric orbit with four frequency components with different radial amplitudes and relative phase angles; and (2) Nonsymmetric orbit with two frequency components with two different vertical/horizontal amplitudes and two different relative phase angles. Each orbit carries a Keyphasor mark (one-per-turn reference). The frequencies, amplitudes, and phase angles, as well as number of time steps for orbit computation, have to be chosen and introduced to the computer by the user. The orbit graphs can be observed on the computer screen.

Muszynska, A.↗

NASA NTRS1987

A demonstration of high precision GPS orbit determination for geodetic applications

High precision orbit determination of Global Positioning System (GPS) satellites is a key requirement for GPS-based precise geodetic measurements and precise low-earth orbiter tracking, currently under study at JPL. Different strategies for orbit determination have been explored at JPL with data from a 1985 GPS field experiment. The most successful strategy uses multi-day arcs for orbit determination and includes fine tuning of spacecraft solar pressure coefficients and station zenith tropospheric delays using the GPS data. Average rms orbit repeatability values for 5 of the GPS satellites are 1.0, 1.2, and 1.7 m in altitude, cross-track, and down-track componenets when two independent 5-day fits are compared. Orbit predictions up to 24 hours outside the multi-day arcs agree within 4 m of independent solutions obtained with well tracked satellites in the prediction interval. Baseline repeatability improves with multi-day as compared to single-day arc orbit solutions. When tropospheric delay fluctuations are modeled with process noise, significant additional improvement in baseline repeatability is achieved. For a 246-km baseline, with 6-day arc solutions for GPS orbits, baseline repeatability is 2 parts in 100 million (0.4-0.6 cm) for east, north, and length components and 8 parts in 100 million for the vertical component. For 1314 and 1509 km baselines with the same orbits, baseline repeatability is 2 parts in 100 million for the north components (2-3 cm) and 4 parts in 100 million or better for east, length, and vertical components.

Lichten, S. M.↗

NASA NTRS1988

Planetary operations for the Pioneer Venus Orbiter mission

The Pioneer Venus Orbiter was inserted into Venus orbit on December 4, 1974 and is expected to operate until 1992. This paper discusses the scientific objectives of the Pioneer Venus Orbiter Project components, i.e., the Orbiter spacecraft, the atmosphere entry Probes, and the probe Bus, together with the design and operation characteristics of the Orbiter. Consideration is also given to the science instruments of the Orbiter, to their requirements with respect to pointing, orientation, and timing, and to the ground operations system. The Probes, the Bus, and the Orbiter were selected to study Venus in complementary ways: the Probes by sounding through most of the atmosphere at four locations, the probe Bus by sounding through the upper atmosphere during entry, and the Orbiter by in situ planet-wide observations of the ionosphere and the extreme upper atmosphere, and by examining the surface and interior remotely. In addition, the Orbiter also observes the interactions of Venus with the solar environment.

Jackson, Robert W.↗